Test Techniques for Flight Control Systems of Large Transport Aircraft 012822990X, 9780128229903

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Test Techniques for Flight Control Systems of Large Transport Aircraft

Test Techniques for Flight Control Systems of Large Transport Aircraft Yakui Gao Xi’an Aircraft Design and Research Institute, Shaanxi, P.R. China

Gang An Xi’an Aircraft Design and Research Institute, Shaanxi, P.R. China

Chaoyou Zhi Xi’an Aircraft Design and Research Institute, Shaanxi, P.R. China

Academic Press is an imprint of Elsevier 125 London Wall, London EC2Y 5AS, United Kingdom 525 B Street, Suite 1650, San Diego, CA 92101, United States 50 Hampshire Street, 5th Floor, Cambridge, MA 02139, United States The Boulevard, Langford Lane, Kidlington, Oxford OX5 1GB, United Kingdom Copyright © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved. No part of this publication may be reproduced or transmitted in any form or by any means, electronic or mechanical, including photocopying, recording, or any information storage and retrieval system, without permission in writing from the publisher. Details on how to seek permission, further information about the Publisher’s permissions policies and our arrangements with organizations such as the Copyright Clearance Center and the Copyright Licensing Agency, can be found at our website: www.elsevier.com/permissions. This book and the individual contributions contained in it are protected under copyright by the Publisher (other than as may be noted herein). Notices Knowledge and best practice in this field are constantly changing. As new research and experience broaden our understanding, changes in research methods, professional practices, or medical treatment may become necessary. Practitioners and researchers must always rely on their own experience and knowledge in evaluating and using any information, methods, compounds, or experiments described herein. In using such information or methods they should be mindful of their own safety and the safety of others, including parties for whom they have a professional responsibility. To the fullest extent of the law, neither the Publisher nor the authors, contributors, or editors, assume any liability for any injury and/or damage to persons or property as a matter of products liability, negligence or otherwise, or from any use or operation of any methods, products, instructions, or ideas contained in the material herein. British Library Cataloguing-in-Publication Data A catalogue record for this book is available from the British Library Library of Congress Cataloging-in-Publication Data A catalog record for this book is available from the Library of Congress ISBN: 978-0-12-822990-3 For Information on all Academic Press publications visit our website at https://www.elsevier.com/books-and-journals

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Contents Preface................................................................................................................. xv Introduction ..........................................................................................................xix Chapter 1: Introduction .......................................................................................... 1 1.1 Overview ............................................................................................................... 1 1.2 Current situation and development trend of flight control system test verification technologies ........................................................................................ 5 1.2.1 Development of the verification test ...........................................................6 1.2.2 Current development of flight control system test verification technologies ................................................................................................7 1.2.3 Development trends of the verification test ...............................................10 1.3 Features and evaluation management of a flight control system verification test .................................................................................................... 13 1.3.1 Basic features of a verification test ...........................................................13 1.3.2 Functions of a verification test ..................................................................15 1.3.3 Principles for verification test design ........................................................16 1.3.4 Technical management of a verification test .............................................17 1.4 Requirements definition of flight control system verification............................... 18 1.4.1 Requirements definition ............................................................................19 1.4.2 Requirements verification and assignment ................................................22 1.4.3 Coverage and traceability..........................................................................24 1.4.4 Requirements management tools ...............................................................24 1.4.5 System verification supports aircraft-level verification .............................24 1.5 Engineering practice of the flight control system test verification ....................... 25 1.6 Flight control system test verification .................................................................. 28 1.6.1 Introduction to the background aircraft flight control system ....................28 1.6.2 Content of a verification test .....................................................................34 1.6.3 Procedures of a verification test ................................................................37

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Chapter 2: Unit test of the flight control system .................................................... 41 2.1 Overview ........................................................................................................... 42 2.1.1 Basic process of unit development .......................................................... 42 2.1.2 Categories of unit tests............................................................................ 43 2.1.3 Basic principles for the selection of test items ........................................ 43 2.1.4 Classification of the flight control system unit ........................................ 44 2.1.5 Basis and requirements for the preparation of the unit test outline .......... 44 2.2 Basic functional/performance test ...................................................................... 46 2.2.1 General requirements of test ................................................................... 46 2.2.2 Control and display unit .......................................................................... 47 2.2.3 Computer unit ......................................................................................... 51 2.2.4 Manipulator unit ..................................................................................... 54 2.2.5 Mechanical drive unit ............................................................................. 58 2.2.6 Mechanical transmission unit .................................................................. 62 2.2.7 Mechanical actuating unit ....................................................................... 66 2.2.8 Sensor unit .............................................................................................. 79 2.3 Strength and stiffness test .................................................................................. 92 2.3.1 Test objectives and test requirements ...................................................... 92 2.3.2 Test items and test methods .................................................................... 93 2.3.3 Judging criteria and results handling ....................................................... 93 2.4 Power characteristic test .................................................................................... 94 2.4.1 Test items and reduction principles ......................................................... 94 2.4.2 Test objectives and test requirements ...................................................... 99 2.4.3 Test items and test methods .................................................................. 101 2.4.4 Judging criteria and results handling ..................................................... 101 2.4.5 Test example: Power characteristic test of flight control computer ....... 101 2.5 Dielectric strength test ..................................................................................... 112 2.5.1 Test objectives and test requirements .................................................... 113 2.5.2 Test items and test methods .................................................................. 113 2.5.3 Judging criteria and results handling ..................................................... 114 2.6 Mechanical environment test ........................................................................... 114 2.6.1 Acceleration test ................................................................................... 114 2.6.2 Vibration test ........................................................................................ 115 2.6.3 Noise test .............................................................................................. 116 2.6.4 Shock test ............................................................................................. 117 2.7 Natural environment test .................................................................................. 119 2.7.1 Low pressure (altitude) test ................................................................. 119 2.7.2 High temperature test .......................................................................... 124 2.7.3 Low temperature test........................................................................... 128

Contents vii 2.7.4 Temperature shock test ....................................................................... 130 2.7.5 Temperature altitude test ................................................................... 133 2.7.6 Temperature humidity altitude test .................................................. 137 2.7.7 Single event test .................................................................................. 140 2.7.8 Solar radiation test .............................................................................. 142 2.7.9 Rain test .............................................................................................. 145 2.7.10 Icing test ............................................................................................. 148 2.7.11 Damp heat test .................................................................................... 151 2.7.12 Mold test ............................................................................................. 153 2.7.13 Salt spray test...................................................................................... 156 2.7.14 Sand and dust test ............................................................................... 159 2.8 Electromagnetic environment protection test ................................................... 162 2.8.1 Electromagnetic emission and susceptibility test ................................... 162 2.8.2 Lightning direct effect test .................................................................... 167 2.8.3 Lightning-induced transient susceptibility test ...................................... 167 2.8.4 High-intensity radiated field protection test .......................................... 171 2.8.5 Electrostatic discharge protection test ................................................... 176 2.9 Reliability test .................................................................................................. 177 2.9.1 Environmental stress screening test ....................................................... 177 2.9.2 Reliability preexposure test ................................................................... 178 2.9.3 Reliability growth test ........................................................................... 179 2.9.4 Reliability qualification test .................................................................. 179 2.9.5 Reliability acceptance test ..................................................................... 180 2.10 Endurance test ................................................................................................. 180 2.10.1 Test objectives and test requirements .................................................. 180 2.10.2 Test items and test methods ................................................................ 181 2.10.3 Judging criteria and results handling ................................................... 181 2.11 Testability test ................................................................................................. 181 2.11.1 Test objectives and test requirements .................................................. 182 2.11.2 Test items and test methods ................................................................ 182 2.11.3 Judging criteria and results handling ................................................... 183 2.12 Test piece selection and test sequence of the unit qualification test ................. 183 2.13 Organization and implementation of the unit qualification test ........................ 184 References ................................................................................................................. 184

Chapter 3: Verification and validation of flight control system airborne software......... 187 3.1 Overview ........................................................................................................... 187 3.1.1 Purpose and significance of verification and validation ..........................187 3.1.2 Basic requirements of verification and validation ...................................190

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Contents 3.1.3 Basic process of verification and validation ............................................194 3.2 Software testing ................................................................................................. 199 3.2.1 Unit testing .............................................................................................200 3.2.2 Component testing ..................................................................................204 3.2.3 Configuration item testing.......................................................................208 3.2.4 System testing .........................................................................................210 3.3 Model-based flight control system airborne software development and testing methods .................................................................................................. 212 3.3.1 Overview of model-based development methods ....................................212 3.3.2 SCADE model testing and verification features ......................................214 3.3.3 SCADE software testing process.............................................................215 3.4 Software whole life cycle support environment ................................................. 222 3.4.1 Basic requirements of environment .........................................................222 3.4.2 Environment architecture ........................................................................223 3.4.3 Environment composition and functions .................................................224 3.4.4 Construction process of software whole life cycle support environment ............................................................................................229 3.5 Software safety and reliability test ..................................................................... 230 3.5.1 Safety and reliability ...............................................................................230 3.5.2 Safety analysis and testing ......................................................................232 3.5.3 Reliability analysis and testing ................................................................232

Chapter 4: Flight control system control law and the flying quality evaluation test ...... 239 4.1 Overview ........................................................................................................... 240 4.1.1 Design requirements for flying quality of large transport aircraft............242 4.1.2 Design requirements for flight control system control law of large transport aircraft......................................................................................243 4.2 Stage division and objectives of the evaluation test ........................................... 244 4.3 Design requirements for the engineering simulator ............................................ 247 4.3.1 Composition............................................................................................247 4.3.2 Main functions ........................................................................................247 4.3.3 Design requirements ...............................................................................248 4.4 Test items and methods ..................................................................................... 253 4.4.1 Planning for test tasks .............................................................................254 4.4.2 Selection of test state points ....................................................................254 4.4.3 Test items ...............................................................................................255 4.4.4 Test control action ..................................................................................257 4.4.5 Test task list ............................................................................................259

Contents ix 4.4.6 Preparation of test report .........................................................................261 4.4.7 Test analysis report .................................................................................261 4.5 Data collection, processing, and evaluation methods ......................................... 262 4.5.1 Overview ................................................................................................262 4.5.2 Requirements for data collection .............................................................262 4.5.3 Requirements for data processing ...........................................................263 4.5.4 Objective evaluation methods .................................................................263 4.5.5 Subjective evaluation methods ................................................................265 4.6 Management of control law and the flying quality evaluation test ..................... 266 4.6.1 Planning of test .......................................................................................267 4.6.2 Preparation for test ..................................................................................268 4.6.3 Control of test process ............................................................................269 4.6.4 Summary of test ......................................................................................270

Chapter 5: Combined test of the flight control subsystem ......................................271 5.1 Overview ........................................................................................................... 271 5.2 Combined test of the pilot control units ............................................................. 273 5.2.1 System introduction ................................................................................273 5.2.2 Test objectives ........................................................................................276 5.2.3 Test requirements ....................................................................................276 5.2.4 Test items and methods...........................................................................278 5.2.5 Criteria for the assessment of test results ................................................282 5.3 Combined test of the fly-by-wire flight control system ...................................... 287 5.3.1 System introduction ................................................................................287 5.3.2 Test objective..........................................................................................288 5.3.3 Test requirements ....................................................................................290 5.3.4 Test items and test methods ....................................................................297 5.3.5 Test results and judging criteria ..............................................................315 5.4 Combined test of the high-lift system ................................................................ 318 5.4.1 System introduction ................................................................................318 5.4.2 Test objective..........................................................................................318 5.4.3 Test requirements ....................................................................................320 5.4.4 Test items and methods...........................................................................322 5.4.5 Test results and judging criteria ..............................................................332 5.5 Combined test of the automatic flight control system ........................................ 334 5.5.1 System introduction ................................................................................334 5.5.2 Test objectives ........................................................................................334 5.5.3 Test requirements ....................................................................................336

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Contents 5.5.4 Test items and methods...........................................................................338 5.5.5 Test results and judging criteria ..............................................................342 5.6 Combined test of the machinery control system................................................. 343 5.6.1 System introduction ................................................................................343 5.6.2 Test objectives ........................................................................................344 5.6.3 Test requirements ....................................................................................344 5.6.4 Test items and methods...........................................................................346 5.6.5 Test results and judging criteria ..............................................................352

Chapter 6: “Iron bird” integration test of the flight control system ........................355 6.1 Overview ......................................................................................................... 355 6.2 Test environment and test support equipment .................................................. 358 6.2.1 “Iron bird” integrated test bed ............................................................. 359 6.2.2 Aircraft simulator cockpit ................................................................... 361 6.2.3 Vision system...................................................................................... 364 6.2.4 Sound system ...................................................................................... 366 6.2.5 Sensor and test analysis system ........................................................... 367 6.2.6 Flight test interface ............................................................................. 371 6.2.7 Flight simulation system ..................................................................... 372 6.2.8 Flight control system tester ................................................................. 378 6.2.9 Avionics system exciter ...................................................................... 383 6.2.10 Mechanical displacement signal generator .......................................... 385 6.2.11 Ground hydraulic energy and ground power supply ............................ 391 6.2.12 Comprehensive test management system ............................................ 394 6.2.13 Aircraft motion sensor driver .............................................................. 397 6.3 Debugging and preparation for the flight control system “iron bird” integration test ................................................................................................. 403 6.3.1 Static adjustment and inspection of the flight control system ................ 405 6.3.2 Debugging and technical status of the cross-linking system .................. 408 6.3.3 Potential problems in the flight control system debugging process and cause analysis ....................................................................................... 412 6.4 “Iron bird” integration test of the cockpit control system................................. 414 6.4.1 Overview .............................................................................................. 414 6.4.2 Test principle ........................................................................................ 415 6.4.3 Static evaluation of man machine ergonomics of the cockpit control system ................................................................................................... 415 6.4.4 Static performance testing of the cockpit control system ...................... 417 6.4.5 Dynamic performance testing of the cockpit control system ................. 422

Contents xi bird” integration test of the machinery control system ............................ 423 Overview .............................................................................................. 423 Test principle ........................................................................................ 423 Evaluation of man machine ergonomics of the machinery control system ................................................................................................... 424 6.5.4 Static performance testing of the machinery control system ................. 425 6.5.5 Dynamic performance testing of the machinery control system ............ 426 6.5.6 Fault mode verification of the machinery control system ...................... 428 6.5.7 Study on effects of mechanism support stiffness on system dynamic (static) performance .............................................................................. 429 “Iron bird” integration test of the fly-by-wire flight control system ................. 429 6.6.1 Overview .............................................................................................. 429 6.6.2 Basic status inspection and testing ........................................................ 430 6.6.3 Zero position and stroke inspection....................................................... 432 6.6.4 Testing of servo actuator system ........................................................... 433 6.6.5 Logic function inspection ...................................................................... 437 6.6.6 Built-in-test functional inspection and testing ....................................... 440 6.6.7 Redundancy management function inspection ....................................... 441 “Iron bird” integration test of the high-lift system ........................................... 442 6.7.1 Overview .............................................................................................. 442 6.7.2 Test principle ........................................................................................ 443 6.7.3 Interface inspection ............................................................................... 444 6.7.4 Control function and logic inspection ................................................... 445 6.7.5 Modal conversion function inspection ................................................... 447 6.7.6 Safety protection function inspection .................................................... 449 6.7.7 Display and fault warning function test................................................. 451 6.7.8 Built-in-test and redundancy management function inspection.............. 452 6.7.9 Failure effect test .................................................................................. 452 “Iron bird” integration test of the automatic flight control system ................... 454 6.8.1 Overview .............................................................................................. 454 6.8.2 Test principles....................................................................................... 455 6.8.3 Interface inspection ............................................................................... 456 6.8.4 Polarity and transmission ratio inspection ............................................. 457 6.8.5 Control logic and display function inspection ....................................... 459 6.8.6 Control function and performance test .................................................. 461 6.8.7 Built-in-test ........................................................................................... 462 6.8.8 Failure effect test .................................................................................. 464 “Iron bird” integration test of the flight control system.................................... 466 6.9.1 Overview ............................................................................................ 466 6.9.2 Test principles..................................................................................... 467

6.5 “Iron 6.5.1 6.5.2 6.5.3

6.6

6.7

6.8

6.9

xii Contents 6.9.3 Interface inspection ............................................................................. 468 6.9.4 Polarity and transmission ratio inspection ........................................... 471 6.9.5 Stability margin test ............................................................................ 473 6.9.6 Closed-loop frequency response test ................................................... 476 6.9.7 Time-domain characteristic test .......................................................... 479 6.9.8 Boundary limit and protection function inspection .............................. 481 6.9.9 State and alarm display verification test .............................................. 483 6.9.10 Failure effect test ................................................................................ 487 6.10 “Iron bird” man machine combined test ......................................................... 489 6.10.1 Overview ............................................................................................ 489 6.10.2 Test principle ...................................................................................... 490 6.10.3 Man machine combined test of takeoff and landing and free flight ............................................................................................ 493 6.10.4 Man machine combined test of mode conversion .............................. 495 6.10.5 Man machine combined test of failure effect .................................... 496 6.10.6 Test task list ........................................................................................ 497 6.11 Test results evaluation of the flight control system “iron bird” integration test ................................................................................................. 498 6.11.1 Test results evaluation of the machinery control system ..................... 498 6.11.2 Test results evaluation of the fly-by-wire flight control system ........... 499 6.11.3 Results evaluation of the man machine combined test ...................... 507 6.12 Management of the flight control system “iron bird” integration test ............... 508 6.12.1 Test management requirements ........................................................... 509 6.12.2 Test measurement requirements .......................................................... 514 6.12.3 Test process ........................................................................................ 517

Chapter 7: Onboard ground test of the flight control system .................................521 7.1 Overview ........................................................................................................... 521 7.1.1 Installation and power-on inspection .......................................................521 7.1.2 Functional and performance test .............................................................522 7.1.3 Cross-linking performance inspection between the flight control system and other airborne systems ..........................................................522 7.1.4 Structural mode coupling test..................................................................522 7.1.5 Electromagnetic compatibility test of the flight control system ...............522 7.2 Onboard ground test of the flight control system ............................................... 523 7.2.1 Test principle ..........................................................................................523 7.2.2 Debugging and preparation before test ....................................................525 7.2.3 Test items, test methods, and judging criteria .........................................528

Contents xiii 7.2.4 Organization and implementation of the onboard ground test of the flight control system .........................................................................576 7.3 Structural mode coupling test ............................................................................ 577 7.3.1 Test objectives ........................................................................................577 7.3.2 Test principle and methods .....................................................................578 7.3.3 Implementation of the structural mode coupling test...............................582 7.3.4 Organization and implementation of the structural mode coupling test............................................................................................587 7.4 Electromagnetic compatibility test ..................................................................... 589 7.4.1 Test items ...............................................................................................589 7.4.2 Test methods ...........................................................................................590 7.4.3 Judging criteria .......................................................................................592 7.4.4 Organization and implementation of the electromagnetic compatibility (E3) test ............................................................................593

Chapter 8: Flight test of the flight control system .................................................595 8.1 Overview ........................................................................................................... 595 8.1.1 Requirements and objectives of the flight test .........................................596 8.1.2 Basis of the flight test .............................................................................597 8.1.3 Objects of the flight test..........................................................................598 8.1.4 Stages and content of the flight test ........................................................598 8.1.5 Methods and requirements of the flight test ............................................604 8.1.6 Ground support facilities .........................................................................606 8.1.7 Organization and management of the flight test ......................................607 8.1.8 Team training of the flight test ...............................................................608 8.1.9 Flight test plan ........................................................................................608 8.2 Requirements for the flight test of the flight control system .............................. 610 8.2.1 Basis of preparation ................................................................................610 8.2.2 Items and requirements of the flight test .................................................610 8.2.3 Requirements of the monitoring system ..................................................610 8.2.4 Requirements of the testing system .........................................................612 8.3 Outline of the flight test of the flight control system ......................................... 612 8.3.1 Categories of test outlines .......................................................................613 8.3.2 Basis for preparation of the flight test outline .........................................614 8.3.3 Selection of flight test items ...................................................................615 8.3.4 Selection of flight test status ...................................................................616 8.3.5 Examples of flight test items...................................................................619 8.4 Test system of the flight test of the flight control system .................................. 623

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Contents 8.5 Data acquisition, processing, and analysis of the flight test of the flight control system .......................................................................................... 625 8.6 Organization and implementation of the flight test of the flight control system.................................................................................................... 628 Further reading........................................................................................................... 630

Chapter 9: Airworthiness verification test of the flight control system ....................631 9.1 Overview ........................................................................................................... 631 9.2 Airworthiness verification test certification requirements .................................. 633 9.2.1 Engineering verification test certification requirements ..........................633 9.2.2 Flight verification test certification requirements ....................................639 9.2.3 Practices of airworthiness verification test of military aircraft ................642 9.3 Technical requirements for the airworthiness verification test ........................... 644 9.3.1 Laboratory test ........................................................................................644 9.3.2 Onboard ground test................................................................................649 9.3.3 Flight test ................................................................................................653 9.3.4 Engineering simulator test.......................................................................660 9.3.5 Unit qualification test..............................................................................665 Chapter 10: Development and expectations on test techniques of the flight control system ........................................................................673 10.1 Role of aircraft conceptual design technology in promoting the flight control system test ........................................................................................... 674 10.2 Role of flight control system design technology in promoting test techniques ........................................................................................................ 675 10.3 Role of test techniques in promoting the flight control system test .................. 677

Index ..................................................................................................................681

Preface With the continuous improvement of safety requirements for modern large transport aircraft, the flight control system has become one of the key systems to ensure flight safety. Measures such as the aerodynamic layout of multiple control surfaces of aircraft, redundancy design of the flight control system, and the application of highly reliable and advanced units are adopted to effectively improve the safety of aircraft. The aerodynamic layout of multiple control surfaces of an aircraft complicates the aerodynamic characteristics and load calculation and there are several redundant aircraft control solutions. The redundancy and reconfiguration design of the flight control system brings geometric growth to the operating mode, control law, and software workload of the system. The application of advance units such as the dissimilar redundancy control computer, precise motion mechanism under large deformation, sensor under special environment, and large-power actuation system has greatly improved the system’s maturity. Therefore the test of the flight control system, as the main process of the verification and confirmation of a complex system and new technologies and products, is an indispensable link for the successful R&D of aircraft. Based on the R&D practice of the flight control system of large transport aircraft, the book introduces the test purposes, methods, content, results analysis, and test process organization and management of the flight control system of large transport aircraft in a systematic way. While highlighting the verification of new technologies such as the airborne software and airworthiness of the flight control system, the book presents coordinated, complete, and practical content that reflects the development trend of the test techniques for the flight control system at home and abroad. The book mainly aims to make the designers of the flight control system fully understand the verification process of the flight control system of large transport aircraft and their units during R&D and impart the basic principles of flight control system verification and validation. Chapters of this book include the unit test of the flight control system, airborne software verification and validation, the control law and flight quality evaluation test, the subsystem combined test, the “iron bird” integration test, the onboard ground test, the flight test, and the airworthiness verification test.

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Chapter 1, Introduction, introduces the development and current situation, requirements definitions, and evaluation management of the test techniques for the flight control system. It introduces the engineering practice of the verification test by taking the B777 aircraft as an example and briefly describes the flight control system of this aircraft, providing the background to this book. Chapter 2, Unit Test of the Flight Control System, according to the classification of the functions and physical properties of the flight control system, introduces the test items’ selection, test requirements, test methods, and test results judgment criteria of verification tests of units, such as the functional performance test, environmental test, and life test. Chapter 3, Verification and Validation of Flight Control System Airborne Software, defines the importance of each module of the flight control system software and proposes some methods for airborne software verification and validation, as well as requirements according to the airborne software development process and stage division. Chapter 4, Flight Control System Control Law and Flying Quality Evaluation Test, introduces the flight quality and control law design requirements of large transport aircraft, as well as the content, methods, evaluation criteria, and organization management of the control law and flight quality evaluation. Chapter 5, Combined Test of the Flight Control Subsystem, introduces the development idea of “layer-by-layer integration and step-by-step verification” of the flight control system, as well as the purpose, content, and test judgment criteria of the subsystem comprehensive verification. Chapter 6, “Iron Bird” Integration Test of the Flight Control System, introduces the environment of the “iron bird” integration test of the flight control system and the purpose, content, and method of the integration test and man machine combined test. Chapter 7, Onboard Ground Test of the Flight Control System, introduces the test content, methods, and procedures of the onboard ground test, structural mode coupling test, and electromagnetic compatibility test. Chapter 8, Flight Test of the Flight Control System, introduces the flight verification of the flight control system in each stage of the flight test and emphasizes the importance of the flight control system in supporting the aircraft’s continuous safe trial flight. Chapter 9, Airworthiness Verification Test of the Flight Control System, describes the approval requirements and technical requirements for the airworthiness verification test of the flight control system. Chapter 10, Development and Expectations on Test Techniques of the Flight Control System, summarizes the development of the general aerodynamics system, flight control system, and test technology, and forecasts the development of the flight control system test. This book was compiled within 2 years by the Writing Committee composed of personnel in the discipline of flight control of No. 1 Aircraft Design and Research Institute of Aviation Industry Corporation of China (AVIC) under the unified organization, planning, and guidance of Gao Yakui, a senior technical expert of AVIC. This book is prepared based on the engineering practices of all authors, featuring extensive and novel content, and reflecting the new technologies, new methods, and new

Preface xvii achievements for the flight control system test of modern large transport aircraft. The book contains 10 chapters in total. Yakui Gao serves as the director of the Writing Committee, and Gang An and Chaoyou Zhi serve as the deputy directors of the Writing Committee. Members of the Writing Committee include Sun Jianquan, Shu Yongling, Lu Lichuan, Li Yu, Lin Hao, with Zhu Yan serving as the writing secretary. More details are presented as follows: Chapter 1, Introduction, was written by Yakui Gao, Chaoyou Zhi, Gang An, and Yongling Shu, Chapter 2, Unit Test of Flight Control System, was written by Lichuan Lu, Mingang Mou, Jie Yu, and Gui Wang, and Chapter 3, Verification and Validation of Flight Control System Airborne Software, by Yu Li, Hongfang Cheng, Jiahang He, Jia Liu, Lvyuan Wu, Xiaoyong Wei, and Xin Zhang. Chapter 4, Flight Control System Control Law and Flying Quality Evaluation Test, was written by Gang An, Wenjing Hei, Jiang Zhu, Hao Lin, and Li Ma. Chapter 5, Combined Test of Flight Control Subsystem, was written by Yongling Shu, Fei Zhang, Xiantong Chen, Ruiqin He, Yuan Zhang, and Zhenyun Shi. Chaoyou Zhi, Xia Li, Jianhui Han, Wei Li, Yahong Wang, Fen Zhang, Yan Zhang, Wei Chen, Honggang Li, Li Ma, and Shengjun Shi wrote Chapter 6, “Iron Bird” Integration Test of Flight Control System. Gang An, Fei Zhang, Yan Zhang, Lihong Wang, and Lichuan Lu wrote Chapter 7, Onboard Ground Test of Flight Control System. Hao Lin, Wei Jian, Li Ma, and Pengxuan Zhao wrote Chapter 8, Flight Test of Flight Control System. Chapter 9, Airworthiness Verification Test of Flight Control System, was written by Jianquan Sun, Xinhui Zhang, Yahong Wang, Yueheng Qiu, and Lvyuan Wu, and Chapter 10, Development and Expectations on Test Techniques of Flight Control System, was written by Yakui Gao, Gang An, Chaoyou Zhi, Jiang Zhu, Yongling Shu, Hao Lin, and Wenjing Hei. The book serves as a reference for engineering technicians engaged in the flight control system and other relevant disciplines or as a scientific research and education reference for university teachers and students. Here, we would like to express our sincere thanks to the engineering technicians from the Flight Control and Hydraulic System Research Institute of No. 1 Airplane Design and Research Institute of AVIC for their great support toward the manuscript and case study materials. The chief technical expert and researcher Chaoxu Yang and researcher Yanming Fan of AVIC are the main reviewers of the book and we highly appreciate their efforts. Due to the limited time and knowledge of the author, there might be some mistakes and errors in the book. Thus we welcome all readers to offer your kind criticism and suggestions. Authors

Introduction The flight control system of large transport aircraft is one of the core systems that affect the safety, functions, and performance of the aircraft. Therefore the scientific, standard, reasonable, and comprehensive experimental verification of the flight control system is the key to the successful R&D of the aircraft. This book describes the unit test, control law and flight quality evaluation test, combined test, “iron bird” integration test, onboard ground test, structural mode-coupling test, electromagnetic compatibility test, and flight test of the flight control system in detail. This book is a summary of the authors’ theoretical accomplishments and engineering practices over the years, with novel and systematic content, and combines theory with practice, thus showing great value both theoretically and practically. This book is intended for the reference of scientific and technical personnel in the field of flight control, especially those who are engaged in systems integration development and experimental verification research and application, and also for teachers and graduates of related majors in colleges and universities.

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Introduction 1.1 Overview As a perfect product of modern high technology, aircraft have become an important symbol of national science and technological progress as well as an embodiment of overall national strength. As an important airborne system necessary to ensure the safe task execution of aircraft, the flight control system plays a crucial role in improving the overall performance of the aircraft, improving flight quality, and enhancing aircraft safety. Especially with regard to the fly-by-wire (FBW) flight control system, it has become an important symbol of modern civilian aircraft. Of course, many aircraft, especially transport aircraft, are mostly equipped with machinery control systems with certain control capabilities to improve security. With the upgrading of the design philosophy of aircraft and the comprehensive and intelligent development of flight control systems, modern aircraft design no longer features the traditional triad of “aerodynamics 1 structure 1 power,” but the quaternity of “aerodynamics 1 structure 1 power 1 flight control” instead. The flight control system has become more important in aircraft design and designers have placed greater emphasis on the functions of the flight control system. For example, it has active control (ACT) functions including relaxing static stability, improving riding quality, boundary protection and restrictions, and reducing gust loads. The flight control system improves the coupling with the aerodynamic, power, and other airborne systems of aircraft and has addressed many problems that cannot be solved in the past by merely relying on aerodynamics, structure, and power, for example, flutter suppression reduces operating costs and carefree handling. The more functional and better performing flight control systems of modern aircraft leads to complex internal and external cross-linking relations and the increase of LRUs (local remove unit) that comprise the flight control system. Due to the strict security requirements for large transport aircraft, redundancy design should be fully considered, including the redundancy of control surface, energy (hydraulic system, power supply system), and system architecture. This is especially pertinent for the flight control system that is critical to aircraft security, which complicates the composition and cross-linking relations of flight control systems of modern large aircraft. According to statistics, the number of LRUs in the flight control system of large aircraft is eight times higher than that of conventional mechanical aircraft, which leads to a sharp increase of the design and Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00001-2 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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verification workload and the development cost of the flight control system of modern large aircraft. There are two types of flight control systems of civil aircraft (ATA100) according to its method of piloting, that is, a manual flight control system and an automatic flight control system (AFCS). The manual flight control system of large aircraft is categorized into the main flight control system and auxiliary flight control system, based on the role of the control function in flight control. For conventional aircraft, the control plane of the main flight control system is the aileron, elevator, rudder, and horizontal stabilizer and the control plane of the auxiliary flight control system is the spoiler (including ground spoiler and multifunctional spoiler), flap, and slats. In order to emphasize the role of the flap and slats of large aircraft in increasing aircraft lift and improving aircraft takeoff and landing performance, the control system of the flap and slats is called the high lift control system. With the successful application of the FBW flight control system, the multifunctional control surface becomes possible. For example, the multifunctional spoiler is used for auxiliary rolling and the aileron can be used for lift enhancement in the takeoff and landing process. A reasonable and scientific division of labor and organization is particularly important when it comes to the massive systematic project of the research and development of the flight control system of large aircraft, which involves a wide range of disciplines, a large number of participating units and personnel, and even dozens of suppliers and tens of thousands of designers. The flight control subsystem division should consider the independence of functions as well as the verifiability of performance. Meanwhile, the interface between subsystems should be as simple as possible, so that the level-bylevel integration under the subsystem test environment and the integration under the system-wide “iron bird” test bench environment can be realized to ensure development quality, improve development efficiency, save development costs, and speed up the development. Given this, the division and development of the flight control system of large aircraft are generally carried out according to cabin control subsystem, FBW flight control subsystem, automatic flight control subsystem, and high lift control subsystem. For aircraft with a mechanical control function, the flight control system also includes a mechanical control subsystem. To put simply and to avoid misunderstanding, “sub” will be omitted when describing the subsystems in the text below, that is, cabin control system, machinery control system, FBW flight control system, high lift control system (HLCS), and AFCS. The design of the flight control system is generally based on the overall design requirements of the aircraft and the design requirements for the aircraft’s maneuverability and stability, as well as relevant standards and specifications formulated for development (e.g., the requirements for natural environment adaptability, requirements for mechanical environment adaptability, requirements for electromagnetic environment adaptability, and

Introduction 3 the requirements for power supply system characteristics and adaptability). On this basis, the design codes of the flight control system as well as the design requirements for the cabin control system, design requirements for the FBW flight control system, design requirements for the AFCS, design requirements for the HLCSs, design requirements for the flight control system control law, and the design requirements for the flight control system software are formulated. For aircraft with a mechanical control function, their design should also follow the requirements for the machinery control system. The basis of the flight control system design includes the system design specifications and the design requirements for subsystems. The most basic units of the flight control system are airborne units and airborne software necessary for the operation, management, and control of flight control tasks. Airborne units of the flight control system generally include electronic products, electromechanical products, and mechanical products. Electronic products include computers, control panels, and display devices. Electromechanical products include power drive devices, balancing drive devices, sensors, and actuating devices. Mechanical products include flap and slats drive devices and cable control and drive devices. In the traditional development process, the chief design company of the aircraft should define the functions, performance, and other design requirements of each airborne unit and airborne software. This process should respect design specification of flight control system and the design requirements of the subsystems. The output of the process is the design requirement and technology agreement of the airborne unit and airborne software. Given the abovementioned design requirements for airborne units and the development task book (technical agreement), the airborne unit development unit of the flight control system will, on the basis of fully understanding relevant requirements and detailed design, refer to some special design codes and standards for the airborne unit and form the specifications or technical conditions of the airborne unit. The whole airborne unit development will follow the product specifications and technical conditions mentioned above. The purpose of testing the flight control system is to verify that the developed flight control system, subsystem, airborne unit, control law, and airborne software meet the specifications, requirements, or technical conditions above. In view of the importance and complexity of the flight control system of modern large civil aircraft, it is necessary to conduct a scientific verification testing of the flight control system. The verification testing of the flight control system also includes the effective confirmation that the flight control system has realized its functions, reached its performance, and is able to avoid faults. In fact, with the growing demand for modern aircraft, the continuous improvement of aerodynamics, flight mechanics, fluid mechanics, control theory, computer technology, electronic technology, and electromechanical technology and the successful application of new thinking, new technology, new tools, and new process, the design of flight control

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systems caters better to the aircraft’s demands and plays a positive role in improving the overall performance of the aircraft. Besides, the verification testing of flight control system is becoming more comprehensive and meticulous, while the development cost has been effectively controlled, and thus the development efficiency is enhanced. Nevertheless, the development of the flight control system is still huge and complex system engineering—one of the key and difficult points in the development of all aircraft—which requires a lot of manpower, time, and money. The engineering development of a complete set of flight control systems needs years of time and hundreds of millions of dollars. It is worth noting that almost half of the funds and workload are concentrated on the system verification test and go throughout the development process, covering the flight control system airborne unit qualification test, subsystem combined test, system integration test, engineering simulator test, onboard ground test, and flight test. The interfaces which the aircraft makes with the flight control system such as function, performance, safety, reliability, and related aircraft performance and flight quality, as well as the information, energy, alarm display, maintenance, and guarantee required for the operation safety of the flight control system, must be verified and validated through the verification test above. The verification test is comprehensive, from the airborne unit components and subsystems to the system. The verification environment is also the airborne unit-level test environment (e.g., it involves hundreds of tests of equipment in natural environment, mechanical environment, electromagnetic environment, and endurance test environment), subsystem comprehensive laboratory, “iron bird” integrated test bench, and real aircraft. As it is very large and complex system engineering, the verification and validation of the flight control system cannot be solved by any kind of design or test tools independently, but requires the combination of a series of related tools and equipment. To develop a complex flight control system, it is crucial to ensure the high quality and efficiency of the verification test. The test design is of particular importance, and should cover test requirements analysis, test planning (plan) preparation, test items determination, test documents preparation, test process formulation, test environment construction, test organization and implementation, and test results evaluation. In the process of aircraft development, the test design and test environment construction should be carried out simultaneously with the design of flight control system to ensure that the system design can be verified in time. In this chapter, the background of the verification test and technology is summarized. On the basis of the brief introduction of the current situation and the development trend of the flight control system verification test technology, the characteristics and evaluation management system of the flight control system verification test are summarized. Taking the typical civil aircraft Boeing 777 as a research object, this chapter defines the requirements of flight control system verification and the engineering practices of the flight

Introduction 5 control system verification test. Finally, the composition and working principles of the flight control system of the background aircraft studied in the book are introduced. The basic content and processes suitable for the flight control system verification test of large aircraft are proposed based on engineering practices, including the requirements and methods for the airborne unit qualification test, subsystem combined test, flight control system “iron bird” integration test, engineering simulator test, onboard ground test, flight test, and airworthiness verification test, which will be discussed in the following chapters.

1.2 Current situation and development trend of flight control system test verification technologies The flight control system technology has had many development stages such as mechanical control (hard, soft or mixed type), augmentation, control augmentation, full-authority FBW control, and integrated control. It is especially worth noticing that since the emergence of the two epoch-making flight control concepts FBW control and ACT technology in the second half of the 20th century, the development process of aircraft has undergone tremendous changes, which are reflected in the following aspects: 1. Traditional aircraft design concept and methodology have changed, and the flight control technology, together with aerodynamic, structural, and power devices for the first time, have become the four pillars that guarantee the performance of advanced aircraft platform. 2. The traditional method, where mechanical chain control was used as the main control system of aircraft since the first flight of the Wright brothers, has changed and the closed-loop control system feeding back the aircraft state is used as the main flight control system of modern aircraft. 3. The restriction principle of center of gravity configuration in aircraft layout design is broken to allow the aircraft to be designed into neutral stable and static unstable architecture. 4. The flight control system is not only used to enhance the rigid body motion performance of aircraft, but also to solve the control problem of aircraft elastic mode. 5. To make the integration of the main control systems (flight control system thrust control system and fire control system) possible. The performance of the aircraft is significantly improved by the development of the system intergraded technology centered on FBW control system. At present, digital FBW and various ACT functions have been widely used in advanced civil aircraft and the integrated control technology has become one of the typical symbols of the modern aircraft. With the development of aviation technology, control theory, and computer technology, especially the development and verification of ACT, the flight control system has become a powerful means to get rid of the design and use constraints of

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traditional aircraft and also one of the key technologies to guarantee the performance of the advanced aircraft platform and meet the requirements for completing new tasks. With the continuous development of aviation technology, the flight control system is no longer an airborne system that is considered the icing on the cake, but an indispensable key system that directly affects the flight performance and security of modern aircraft. Over the past 20 years, under the rapid development of relevant disciplines, many creative new theories, new methods, and new products have emerged in the field of flight control technology and some of the new technologies have realized engineering validation and application. It is very likely that these developing new technologies will provide technical support for the future revolutionary development of aircraft.

1.2.1 Development of the verification test The verification test technology constantly develops as the flight control system technology evolves. Based on the development process of the flight control system verification test technology in China, it can be divided into three stages. The first stage is from the 1970s to the 1980s. In this stage, the flight control system is mainly a machinery control system with the function of augmentation or control augmentation, which is normally called the flight control system. At that time, the flight control system was only used to improve the flight quality, not a key system to ensure flight security, and it was relatively simple. In this phase, the flight control system verification test mainly included machinery control system verification with the load simulator as the center, the augmentation, control augmentation, and autopilot function verification with augmentation computer as the core, and the flight control system verification integrated with the flight control system, with the “iron bird” integrated test bench as the center. Engineering simulator tests, onboard ground tests, and flight tests were also carried out. During this period, China had sufficient flight control system verification tests for aircraft with FBW control systems and established relatively exemplary test equipment, such as the quality simulator and the “iron bird” integrated test bench. These verification tests ensured the successful development of many aircraft at the time. The second stage was from the 1990s to the end of the 20th century. In this stage, new aircraft in China generally adopted FBW systems. The tight coupling between the flight control system and the aircraft body, aerodynamics, structure, and the other airborne systems of aircraft prompted designers to put more emphasis on the study of verification test technology. The concept of comprehensive verification of the aircraft system while centering on the “iron bird” integrated test bench was basically formed and the engineering simulator and “iron bird” integrated test bench were established for each model. However, the technological progress of the flight control system was mainly concentrated on small

Introduction 7 aircraft and the flight control system technology of transport aircraft was still a machinery control system. In addition, the system design and verification were mainly based on static performance, such as the machinery control system of Y-7 series aircraft. The flight control system tests of transport aircraft were still mainly the airborne unit qualification test and “iron bird” integrated test, and improvement was only seen in the automation of test equipment. The third stage is from the start of the 21st century to the present. In this stage, various aircraft flight control system technologies have made great progress. The aircraft system is no longer a separate flight control system, avionics system, fuel oil system, hydraulic system, and power supply system, but an aircraft integrated system with vehicle management system (VMS) as an integrated equipment. As a result, the aircraft system integration and the overall performance of aircraft are highly improved. The corresponding aircraft system verification test technology is also developing and a sound environment for comprehensive development and verification is in place for a large-scale aircraft system. A distinctive feature of the flight control system of transport aircraft is FBW control. ARJ21 aircraft and MA700 aircraft adopt open-loop FBW flight control systems; C919 aircraft apply a full-authority FBW system and other transport aircraft adopt a full-authority FBW system with a machinery control system as a backup. The progress of the flight control system technology has led to higher requirements for the verification test. The concept of the classified integrated test mainly reflected by the “iron bird” integration test has been introduced and the subsystem combined test and engineering simulator test are strengthened. Meanwhile, advances in computer technology, sensor technology, bus technology, and software technology have also enhanced the efficiency of the test.

1.2.2 Current development of flight control system test verification technologies The model development in the field of aviation and aerospace generally applies the top-down “V”-shape design process. On this basis, the development of the flight control system can be divided into the following stages: requirement demonstration stage, scheme definition stage, detailed design stage, airborne unit design stage, airborne unit pilot stage, airborne unit qualification stage, subsystem integration stage, full-system integration and ground test stage, flight test stage, and finalization stage. All stages have a spiral iteration and approximation with the ultimate goal of achieving the flight control system, as shown in Fig. 1.1. The left side of the “V”-shape development process is a top-down process from the aircraft requirement breakdown to the design requirements of the airborne unit, which is actually a confirmation of the requirements. The right side of the “V”-shape development process is a bottom-up process from airborne unit verification to full-system integration verification, which is actually verification. At the bottom of the “V”-shape development process is the

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Figure 1.1 Flight control system design and verification test procedure.

manufacturing and verification of the airborne unit, and the whole system development results are shown here in the form of the airborne unit. As the control law design and verification is one of the important features that distinguish the development of the flight control system from other aircraft systems, it is the key to realizing the function of the flight control system and meeting the flight quality requirements. The control law design and verification also follow the “V”-shape development process, from confirmation of requirements to flight verification. The development result of the control law is the control law software code and the verification of the control law is actually the verification of control law software. It can be seen from Fig. 1.1 that in the course of the development of a flight control system, several cyclic iteration loops are formed in a bottom-up manner, from the bottommost airborne unit design and verification of “V”-shape development process, including whether airborne unit meets the requirements for airborne unit, whether the subsystem meets the subsystem design requirements, whether the system meets the system design specifications, and whether the aircraft meets the aircraft development requirements. All verifications should be performed in the most real possible environment as per relevant existing standards and specifications. For example, the airborne unit qualification test requires that designers only need to establish or borrow a test environment as required. However, there still are a lot of test items that should be tested in the so-called real environment, such as the comprehensive verification of subsystems, all-system integrated verification, and engineering simulator verification. All of these are based on the experience and ability of

Introduction 9 the designers, the level of human resources, support funds, and the development cycle, etc. and they are sometimes not just technical issues. But designers should know them fairly well. Statistics show that the typical system verification test of the flight control system test accounts for 34% of the whole system development cost, taking the largest part in system development cost. In contrast, the software development costs (20%), hardware development costs (17%), and control law development costs (9%) take much fewer resources. It should be noted that in the statistics, the internal verification test costs of software, hardware, and control law in the design and development process are not included in the calculation. If specified further, the verification test will undoubtedly take a larger proportion of cost in the system development. Fig. 1.2 shows the cost model of a typical flight control system development process. On the whole, the flight control system verification test features a heavy and wide ranging workload, lacks sufficient extraction and summarization, and is without consistent management and technical specifications at present. In engineering practices, some inspiring and successful applications and practices exist in system engineers’ minds in the form of engineering experience and some research achievements are mostly distributed in various literatures, although these are far from forming a complete system of theory and practice. With the continuous improvement of the status of flight control systems and verification test technology, relevant research has gradually attracted the attention of the theoretical and engineering community, and verification test technology is bound to become an increasingly important research direction of the flight control system technology. The sudden changes in the software code amount of the flight control system in the example show the great challenges in the flight control system verification test. Fig. 1.3 shows the software code amount of the flight control system of T-50, F-22, Darkstar, JSF, URAV, and UCAV aircraft and it gives a direct prediction of the complexity of the flight control system from the software code amount. In the foreseeable future, the software code amount of unmanned combat aircraft vehicle (UCAV) will be dozens of times that of the

Figure 1.2 Cost model of typical flight control system development process.

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Figure 1.3 Software complexity (code amount, 10,000 lines).

most advanced fighter aircraft and unmanned aerial vehicles (UAV) in service. It is conceivable that for such a large and complex system, the cost of its verification test must be considerable. In the face of the growing number of tasks, the flight control system is able to solve more complex problems as its technologies develop continuously. Sophisticated and reliable system architecture and software functions will definitely multiply the system verification workload and the challenges will be mounting even under the constant improvement and progress of the test methods and test environment.

1.2.3 Development trends of the verification test In recent years, the flight control system verification test has presented many new features in concept and technology, mainly in the following aspects. 1. Design for validation: design verifiability should be considered at the beginning of design, that is, “design for validation,” which requires planning and management of system verification and validation at the beginning of system development. 2. Same system, same model, and same tools for design and validation: the development of test equipment should be emphasized. Advanced verification devices and

Introduction 11

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development tools should ensure the visualization of system growth (let the user see his system grow in front of his eyes) and the same platforms and tools should be applied extensively in the development and verification process. For example, in the system development of F-22 aircraft, the concept of a vertical test is proposed. In other words, direct testing with same equipment is conducted by the factory (design) at site. That is to say, in F-22 development and testing, the same system, same model, and same tools and even common Ada language are used in the design verification and acceptance maintenance. All-state and all-environment verification: verification and validation should be done under all states and all environments. The accuracy and validity of the verification should be guaranteed, that is, ensuring “what you prove is what you get.” It should be noted that for subsystem-level verification, there must be complete test equipment, all links and parameters must be strictly tested, test items under dynamic environment should be emphasized, and items should not be deleted due to test difficulty. Ensure the integrity of the tests at each stage and do not leave problems to the next stage. Application of virtual prototyping technology and emulation technology: the application of virtual prototyping technology and emulation technology and strengthening of mathematical simulation can reduce the time of the large-scale combined test. Problems found in the mathematical simulation stage should not be left to the ground test and problems that can be solved in the ground test should not be left to the test flight. Standard verification specifications and environment: the existing expertise and standard knowledge base should be leveraged to establish a standard verification environment; the ability to quickly modify system design should be improved to meet the new requirements of customers; acceptance specifications of test items should not be determined by relying on subjective experience and judgment; necessary content and reviews should not be omitted due to insufficient time, and the “reusability” of system development should be improved. Application of new development and verification methods: in order to solve the problem with a complex system while pursuing low-cost development, a new development and verification method should be adopted to complete the transition from the “V”-shape development mode to the “Y”-shape development mode. For example, Ines Fey et al. of the Daimusuoli Kreisler Research Center proposed building a software development environment with commercial MATLAB and SCADE to achieve the automatic generation of software code that complies with RTCA/DO-178B and shortens the verification cycle. Application of advanced verification means such as testing machine: a testing machine and simulator should be used to carry out early verification. For example, before the maiden flight of the F-22, a F-16 variable-stability aircraft was used to verify the maneuverability and stability (1 year) of the F-22 and a 1400-hour test flight of the avionics system was conducted on a Boeing 757 flight test platform (FTB). For another

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The complexity of the aircraft’s mission and the worsening flight environment give rise to an aircraft integrated system (aircraft management system) that takes the flight mission as the top priority. The flight control system no longer involves the system design cored with the pilot’s control of the aircraft in traditional sense, but involves intelligent control and autonomous control centered on the mission. Besides, nowadays with the sustainable growth of the global economy and the fast pace of social activities, time and human cost have become the biggest challenge for all industries. The traditional verification method which consumes massive manpower, financial resources, and time can hardly meet the requirements. It is imperative that more efficient and economical system verification technologies and methods be put in place. The continuous improvement of the flight control system design method also puts forward higher requirements for the verification test method and environment. For example, the flight control law design based on the linear small perturbation equation of an aircraft is very different from the control law design based on the six-degree-of-freedom total nonlinear equation of an aircraft in verification methods. Another example would be that from the frequency-domain flight quality criteria design based on single-input single-output system to the optimization design based on multiinput multioutput system, which urgently demands the development of new effective criteria and evaluation methods. There are also great challenges in the management of the verification process, as detailed below. 1. User requirements: the key to validation lies in the description of requirements, yet most user requirements are incomplete and vague, and the complete description of requirements comes from validation itself. 2. User confidence level: it must be acknowledged that some performance of complex systems is unknowable or unforeseeable from a physical or economic perspective. For example, when the system environment is not fully controllable, it does not mean all system changes are foreseeable. Given these factors, users may want more validation options. 3. Validation and system life cycle: for better validation, the system should be validated again if it is changed. The validation should also be conducted at the design improvement stage. 4. Validation methods: there are many methods for validation and they have their advantages and disadvantages. It is difficult to determine when and how to use these methods. In addition, emulation technology plays an increasingly important role in flight control system verification and validation. Various requirements on authenticity, complexity,

Introduction 13 real-time performance, and interactivity in emulation practices are also constantly improving. During verification and validation, building an integrated modeling, emulation, and verification environment with complete functions, complete information and data, complete structure, and powerful performance is an inevitable demand for future flight control systems.

1.3 Features and evaluation management of a flight control system verification test Practice is the sole criterion for testing accuracy. Similarly, testing is the sole method for verifying if the design results meet design requirements. For such a complex system as the flight control system, test results often cannot directly reflect the design results. For example, the high-order dynamic data obtained from the test can only get the corresponding flight quality design requirements through equivalent matching methods. That is to say, the verification test should answer not only how to get correct design results through tests, but also how to evaluate the design results from the tests. The answer to the first question is the test process and the answer to the second question is the test analysis process. They are constantly an integral part of the design and verification process of the flight control system. The verification test of the flight control system is a process to determine whether the flight control system can accurately meet the design requirements, whether the system can ensure that design requirements are improved, and whether the system can meet the user requirements in reality. It is a part of the flight control system and the entire aircraft development process. The verification test starts from the requirement analysis of the system, runs through the whole development process of the flight control system, and finally validates if the system meets the design requirements determined at the beginning of the system design. The flight control system verification test has become a key activity in the whole system development process. It has a great impact on aircraft flight safety, mission completion, and the full life cycle, as well as research and development cost. The challenges to this work are self-evident.

1.3.1 Basic features of a verification test The verification test of a flight control system is a complicated process both in terms of technology and management and its features are described as follows. 1. Failure verification test is the core of the flight control system test. The flight control system is the most critical system to ensure the successful completion of an aircraft’s mission and its safe operations. Regulations for Civilian Aircrafts stipulates that the probability of catastrophic failure of aircraft caused by function failure of the flight

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control system shall be lower than 1 3 10 9 and the flight control system realizing this security index shall have a complex system architecture of redundancy, an airborne unit of high reliability, and Class A or Class B airborne software. Therefore the verification of this system must be sufficient and the failure mode verification test becomes one of the most important contents of the flight control system test. 2. Verification and validation is a series of technical activities. The flight control system verification test is completed through a series of technical activities. The airborne unit qualification test shall be jointly organized by the airborne unit supplier and the customer representatives (airworthiness representatives of civil aircraft) stationed in the airborne unit development unit and attended by a chief engineer unit. The subsystem integrated test shall be jointly organized by the subsystem supplier and the customer representative of the subsystem development unit (airworthiness representatives of civil aircraft) and attended by a chief engineer unit. The system integration test shall be organized by a chief engineer unit and a customer representative stationed in the chief engineer unit and carried out on a special flight control system “iron bird” integrated test bench. A great deal of the tests requires the attendance of the test flight crew. The onboard ground test shall be conducted on an aircraft, wherein the chief engineer unit takes charge, the chief manufacturing unit coordinates to complete the test, and customer representatives stationed in the chief engineer unit and of the manufacturing unit attend the test. With regard to the flight test, the test flight unit shall take charge, the chief engineer unit and machine manufacturing unit shall coordinate to complete the test, and customer representatives shall attend the whole test process. In the test process, there are dozens to hundreds of times of the test task book review and validation, test outline review and validation, quality and safety inspection before test, test result analysis and report compilation, and review meetings at different levels related to the flight control system verification test, involving a high degree of complexity. 3. The verification test is a technical work that demands high input. The high cost can be seen from the technical activities above for the flight control system verification test. The airborne unit development unit should set up a special test bench for hundreds of airborne units; the subsystem development unit should set up a special test bench for the subsystem; the chief engineer unit should set up a combined test bench for the engineering simulator, flight control system “iron bird” integrated test, and devices for the onboard ground test, and the flight test unit should set up a comprehensive test flight system covering ground monitoring, telemetry and telecontrol, and onboard testing. Meanwhile, the environment required for the testing in different stages of the flight control system software, the control, testing, and analysis software required for the test, and the test pieces required in tests in different stages all require a great input of funds. It is roughly estimated that the input for the software and hardware related to large aircraft development and flight control system verification testing will be in the

Introduction 15 billions of RMB. According to statistics, the verification cost of a certain type of flight control system software accounts for 15.3% of its total life cycle cost, which is 84% higher than its programming cost. The human resources, energy, and material resources required are also a considerable expense.

1.3.2 Functions of a verification test Fig. 1.4 compares errors found in different development stages of a type of aircraft with or without a system verification test. Fig. 1.5 compares relevant costs with or without a system verification test. It can be seen from the comparison that when there is a system verification test, despite a large input required at the initial stage of system development, potential errors of the system can be detected at an early stage, which would greatly reduce the cost compared to the scenario when system errors are found at a later stage of system development or even after product delivery. At the same time, from the perspective of the entire development process, the time cost is also much lower when there are system verification and validation. From the development process of the flight control system, it can be seen that the system design and verification test are closely coupled, repeatedly iterative, and interactive. There are corresponding verification test activities at each stage of the design validation to ensure that the system development has good quality and the system design results meet the design requirements.

Figure 1.4 Verification test at different stages and comparison of errors found.

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Figure 1.5 Verification test at different stages and comparison of costs.

1.3.3 Principles for verification test design The following three principles should be considered when designing the verification test of newly developed flight control systems: 1. Safety first: in a complex environment for the flight control system the verification test is complex, and with a lot of participants in the test any negligence will undermine the safety of personnel, equipment, and even the aircraft. Therefore the test safety should be put in the first place before performing any test activities. For example, technical activities such as the pretest inspection and pretest review (quality, safety) and the maiden flight review (aircraft status, air service preparation, and ground service guarantee) all focus on test safety. 2. Higher efficiency and lower cost: as mentioned above, the flight control system verification test is a costly technical activity, but reasonable and scientific organization and the use of advanced tools will minimize the test cost. For example, the analysis and reference of previous experience, the full use of simulation tools, the reasonable allocation of test items in different stages, and the generalization of test equipment, etc. 3. Prepare the plan beforehand: scientific and reasonable planning of tests play an important role in improving test quality, saving test cost, and shortening the development cycle. For example, the plan risk assessment, the system classification synthesis, and the step-by-step implementation of the general planning of test environment construction. Temporary test tasks can be minimized with careful planning in advance.

Introduction 17

1.3.4 Technical management of a verification test It is acknowledged that, whether it is a country, a company, or a department, management is of crucial importance. For such a complex technical activity as flight control system verification testing, in addition to strong administration, technical management must be strengthened before strict test standards are formed. 1. Requirements analysis should be strengthened and top-level design should be detailed. Previous failures were usually caused by the psychological issues of managers, deadlines, the lack of in-depth requirements analysis, and insufficient top-level design. All of these factors led to a large amount of rework and waste of resources and time. Therefore designers and managers should be more aware of design inputs, and verify and validate the rationale and completeness of the design inputs from the aircraft to the system and the subsystem, and to the airborne unit through simulation. The airborne unit, subsystem, and system scheme that have been formed should have their function, performance, reliability, and safety verified in a simulation environment. 2. The document system should be formulated in a scientific way and text should be formed in a timely manner. Lessons gained from previous experience were that the vague technical plan and chaotic organization of documents would lead to massive temporary documents, loss of regular documents, and weak traceability of design. Therefore the document system, document title, content, and writing period, etc. for the entire system development should be specific while the work breakdown structure (WBS) and the statement of work (SOW) are being formulated. Besides, the design inputs and outputs inside the system and among the upstream and downstream disciplines should be available, so that the required design inputs can be obtained from upstream disciplines and the required design outputs can be provided to downstream disciplines in a timely manner. 3. The test process should be standardized and personnel in charge should be specified. The flight control system test management department should, according to relevant regulations for model development, formulate the process suitable for the test team, including responsibilities at different stages for customer representatives, test management personnel, quality management personnel, test undertaking team, and units participating in the test. Any arbitrary work may lead to concealed or neglected problems. 4. Lessons should be learnt from experience and continuous innovation should be made. Predecessors and partner units have accumulated a lot of experience in the development of numerous models and systems. It is imperative that we make a good technical reserve for system verification test design by way of going abroad, introducing, and even importing technologies, while thoroughly digesting the technical data and literature of predecessors. Under the constraints of technology, funds, schedule, and

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1.4 Requirements definition of flight control system verification Two key terms are defined based on the understanding of the Guidelines for Development of Civil Aircraft and Systems (SAE-ARP4754A): 1. Validation: the process of ensuring the accuracy and completeness of requirements (i.e., the design process) and ensuring that the design meets system and aircraft-level requirements. 2. Verification: the process of ensuring the developed aircraft, system, and airborne unit meet the corresponding design requirements. This section introduces the requirements definition of large aircraft flight control system verification. The flight control system verification process for aircraft mainly consists of three parts: requirements definition, system verification, and problem tracking. The relationship between them is shown in Fig. 1.6. Defining system requirements according to the aircraft-level and understanding the requirements level by level will ensure the accuracy and completeness of the system (airborne unit) level requirements. The system (airborne unit) level requirements drive the design and manufacture of systems and airborne units directly. System-level verification includes airborne unit-level verification, subsystem verification, and system verification to ensure the system and subsystem meet the defined design requirements. Problem tracking

Figure 1.6 Flight control system verification process for aircraft.

Introduction 19 provides closed-loop feedback for requirements definition, system design and implementation, and ensures traceability for the relationship between system requirements, verification activities, and verification states.

1.4.1 Requirements definition Requirements definition is a process of continuous refining and repeated iteration, mainly including requirements formation, document formation, validation, and approval. 1. Requirements definition process The requirements definition takes a top-down approach, from the top-level requirements definition to the main design, and then from the main design definition to bottom-level requirements and design. It is mainly completed through trade study and technical coordination. The trade study of the flight control system architecture of aircraft is shown in Fig. 1.7. Boeing takes customers’ requirements, its own requirements, FAA’s requirements, and JAR (EASA)’s requirements and other model experience as design requirements and goals. The candidate architecture of the flight control system must meet these design requirements and goals and the reliability, cost, weight, and aerodynamic configuration are taken as the elements to have architectural selection and design. For example, the setting of relaxed longitudinal static stability can reduce the weight and flight resistance of the aircraft, but it requires a FBW system that can easily achieve pitch stability augmentation. The weight of the structure can also be reduced by

Figure 1.7 Flight control system architecture design inputs and determination process of aircraft.

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using the FBW system. Through the trade study and comparison, one can draw a conclusion that the reliability of the FBW system is comparable to that of the machinery control system. Some FBW system architectures were lighter but they were not selected due to their low reliability. The trade study finally selects a system architecture that satisfies the design requirements and goals and has the perfect combination of cost, weight, and reliability. In addition to the inputs for the system architecture above, safety requirements are also one of the system architecture design inputs. After the preliminary architecture of the system is selected and aircraft requirements are defined, the requirements of the next level of the system (e.g., subsystem, control law, software) and the airborne unit shall be further defined. The flight control system coordination team continues to conduct the trade study and technical coordination, allocate the functional and performance requirements of each LRU, and define detailed equipment requirements to ensure that they meet the top-level requirements. According to this requirements definition, the requirements and object files of the primary flight control system (PFCS) and autopilot flight director system (AFDS) are formed. These files cover the design requirements, objectives, concepts, definitions, and design decisions of the flight control system and describe system functions, performance, availability, security, isolation, unit operation, and maintenance. 2. Requirements validation process Requirements validation is mainly completed through technical activities such as formal review, analysis, simulation, verification, and approval and system problems in the validation process are continuously fed back to ensure the accuracy and completeness of requirements, as shown in Fig. 1.8. In the late implementation stage of the flight control system project of aircraft, requirements validation and system verification are carried out alternately. Before most of the hardware and software is available for use, all requirements should be validated and they will be taken as important inputs to the detailed design. 1. Formal review The review of the model in the early stage is mainly carried out for requirements validation. Formal review activities include system design review (SDR), preliminary design review (PDR), and detailed critical design review (CDR). Peer representatives from airlines, airworthiness authorities, manufacturers, suppliers, aircraft/system interface teams, and other projects are invited to participate in important reviews. Designated engineering representatives (DER) will participate in reviews in all stages. The review is mainly conducted to further evaluate the accuracy and completeness of the defined requirements from the aspects of engineering, operation, project, and customer and to provide feedback on problems. Any subsequent changes must also be reviewed.

Introduction 21

Figure 1.8 Requirements validation process for flight control system design of aircraft.

SDR is mainly concentrated on general system requirements, system architecture, basic design, and the project development plan. PDR mainly proposes detailed system requirements and design and shows how to prove the design meets the requirements in the initial stage of system verification. CDR covers all changes from PDR to final system review, including assembly maintainability and accessibility. The review of system performance and safety analysis state should be included in the review of each stage. 2. Analysis Analysis is a process to ensure the accuracy of requirements through performance analysis, tolerance analysis, safety analysis, functional risk assessment, and interface analysis. Through performance analysis of the flight control system, the requirements related to the stability of the servo loop and the whole system under normal and fault conditions are validated, including the influence of tolerance on the system. The influence of assembly level tolerance is validated through the analysis on system level total tolerance. In the early stage of the project, the requirements related to system safety are validated through failure and safety analysis. These analysis methods will help avoid high costs caused by changes in the late stage of the project. The functional risk assessment analysis may have risk events that may lead to system failure. The electrical interface analysis of the system evaluates the compatibility of relevant software and hardware of the system’s electrical interface with lightning and high-energy magnetic field to ensure that signals in PFCS and AFDS are consistent with LRU hazard levels.

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Chapter 1 3. Simulation The desktop simulation is used to validate the control law-related requirements and validate the system performance from cabin input to control surface response under normal and degraded operating conditions. Specialized redundancy management simulation is carried out to validate the asynchronous and redundant operation of the system, including the requirements related to fault detection, isolation, and transient prevention. Simulation with the specialized electrical interface analysis software is carried out to validate whether the system’s electrical interface and flow are appropriate. The engineering simulator is also used to evaluate and validate the requirements related to the control quality, man machine interface, and system control under normal and failure conditions. 4. Test In the early stage of the project, various existing test beds and test methods were used for the newly developed products to validate relevant requirements. For example, the primary flight control computer (PFC) and actuator control electronics (ACE) replaceable units of Boeing 777 aircraft were newly developed and were tested on the test bed at the beginning of the project. The new actuator used on Boeing 777 aircraft has been tested on the flight control system “iron bird” integrated test bed of B757 aircraft and the requirements were verified. Special adaptive changes were made to Boeing 757 aircraft to realize the control law of Boeing 777 aircraft and the single-side flight control cabin interface. FBW system flight test was conducted to validate that the system performance and control quality meet relevant requirements, including the function of gust suppression and asymmetric compensation of thrust. 5. Requirements approval The flight control system working group and relevant external departments pertaining to aerodynamics, hydraulic system, and power supply system jointly approve the requirements and objectives of the preliminary design. Changes to the design requirements and objectives at a later stage should also be controlled through requirements review.

1.4.2 Requirements verification and assignment The first step of the verification process is to assign each requirement to one or several verification methods and stages, as shown in Fig. 1.9. All requirements verification methods and stages should be assigned to different disciplines, which are responsible for grouping the requirements and assigning them to corresponding documents. The basic requirements verification and assignment methods mainly include test, analysis, and supplier’s test and analysis. 1. Test Test is the most ideal method to show conformance to the requirements. Test methods should be assigned according to the impacts of other systems or aircraft

Introduction 23

Figure 1.9 Design requirements verification methods and assignment process.

environments to test results. For example, the redundancy management in the system should be evaluated in an independent laboratory, while the degraded operation (including the pilot’s response) of the system should be evaluated under the loss of a single-round state, and the flight requirements of the aircraft should be evaluated in an engineering simulator or test flight. 2. Analysis Due to the limitation of project economics and time, some requirements cannot be proved by tests and the analysis method is mainly adopted. For example, the performance, reliability, and safety estimation of the system are analyzed according to the system redundancy level and estimated assembly failure rate. Analysis methods are often used to reduce the requirements and the test range of requirements are verified according to test conditions. Analysis is also made to verify some requirements of the flight test that cannot be carried out due to their high risk level, such as landing on an icy runway under heavy weight, stuck control surface under extreme conditions, and all-engine failure driftdown. 3. Supplier’s test and analysis Airborne unit and system suppliers conduct extensive testing and analysis to verify that the design meets the specifications and drawing requirements.

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1.4.3 Coverage and traceability Complete verification coverage should ensure that each design requirement and objective is assigned to one or several verification activities. By appointing a project representative to approve the verification document, it should be ensured that the verification activities of corresponding requirements are acceptable and complete and completed verification documents should be filed. Nonconforming requirements will be managed through a strict feedback process to determine whether the system needs to be changed.

1.4.4 Requirements management tools PFCS and AFDS are used to verify the verification methods, stages, and documents assigned for the requirements of the conformance matrix traceability system. By saving these conformance matrix databases with requirements management tools, all design requirements, top-level security events, airworthiness plans, and related verification data can be obtained.

1.4.5 System verification supports aircraft-level verification The verification of the aircraft system aims to confirm that the system supports the requirements of other systems in different flight configurations and meets the aircraft-level requirements in terms of control, performance, and safety. Aircraft-level verification includes various aircraft-level review and analysis, aircraft-level integrated laboratory tests, and flight tests. Aircraft-level verification is carried out for the flight control system directly to provide accurate and timely feedback for the system verification process. 1. Design review Aircraft-level review is carried out for common-cause failures, including rotor burst, tire burst, bird strike, and crash. The review mainly considers the isolation of components, wires, and the hydraulic system and checks the detailed installation of the flight control system through the CATIA model. A review group consisting of representatives of reliability, cabin, pilot, customer service, and maintenance training reviews the warning information and component maintenance information of various units. 2. Analysis Power supply interruption and power-on analysis is carried out for the power supply system to ensure that the aircraft’s response to power supply system changes is predictable and acceptable, including ground warm-start, cold-start, BIT (built-in test), in-flight voltage reduction, and power supply system conversion.

Introduction 25 Various security analyses are updated according to the final system verification, including functional risk assessment, interface analysis, failure mode impact analysis, and failure probability analysis. 3. Aircraft-level test An integrated test bed is used for the final flight hardware and software verification of the flight control system, including the system integration laboratory, engineering simulator, “iron bird” integrated test bed, and flight test aircraft. The system integration laboratory is mainly used to integrate and verify the systems inside the whole aircraft. Since the power supply system used in the system integration laboratory and many LRU cannot represent real aircraft, only the impact of these LRU on the operation of the flight control system can be assessed. The engineering simulator with pilot interface-related hardware can perform the assessment of control quality, unit procedures, system response, and warnings, including the performance test of the system under the condition of wind shear or testing that cannot be carried out in flight or on aircraft due to the risks. One example is when the cabin control assembly is jammed and releases off. Aircraft ground test and flight test are the final aircraft system integration, pilot control quality evaluation, control impact, and failure impact verification. For example, test flights are conducted under a variety of failure conditions that affect the pilot’s control quality, such as single-engine and two-engine shutdowns, single-set and two-set hydraulic system losses, generator failures, and various backup modes. 4. Problem tracking and system conformance evaluation Problem tracking is a process mainly to complete project tracking and system compliance evaluation through the central database of a flight control system and problem tracking at all levels, including aircraft level, system level, and unit level. Problems raised by engineers are at all levels, including document problems found in tests, which are divided into disputes, test problem reports, and flight test action items.

1.5 Engineering practice of the flight control system test verification The verification test of the flight control system of Boeing 777 mainly aims to ensure that the system designed meets the functional, performance, and safety requirements defined by the design requirements and objectives. The verification process is shown in Fig. 1.10. Verification is a process that mainly adopts formal design review, special review and analysis, supplier verification and management, and other verification methods and tests in different stages to verify whether the system designed according to the requirements meets

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Figure 1.10 Verification process of flight control system of Boeing 777 aircraft.

the design requirements and objectives and whether it meets the top-level security requirements. 1. Formal design review Formal design review is part of the requirements validation process and system verification process. For PFCS, AFDS, and assembly, SDR, PDR, and CDR should be completed in the design and the degree to which the system meets the corresponding requirements should be proved. Similar design reviews as well as LRU-level review are part of the supplier verification process. 2. Special review and analysis System function and performance are verified and the definitions of intrasystem interface and intersystem interface are analyzed, including signal definition, signal content, performance characteristics, transmission frequency, delay, and failure mode. The aircraft-level system interface is analyzed and the compatibility among system signals in the aircraft is verified, including the signal name, refresh rate, range, and resolution defined by the interface control file.

Introduction 27 3. Supplier verification and management Equipment suppliers complete the assembly design review, analysis, and testing to support the verification of the flight control system. 4. Verification methods Tests (“iron bird” integration test, onboard ground test, flight test, etc.) and analysis (static analysis, safety analysis, error analysis, stability analysis, etc.) are two main parts of the verification methods. According to the research and development, the verification methods are divided into aircraft level, system level, subsystem level, and unit level. For each detailed requirement, the most suitable method to prove the conformance of the system should be selected based on efficiency and difficulty. 5. Tests A large number of test facilities are used in verification testing. The aircraft system integration laboratory is equipped with these test facilities and is close to the flight test facilities. The PFCS test process is shown in Fig. 1.11. After the arrival of LRU of each flight control system, the appropriate operation and verification of LRU requirements will be carried out on a special test bed first. A system-level integration test is carried out in the Boeing 777 aircraft flight control system integration laboratory and on the “iron bird” integrated test bed. The flight control system test bed is used for flight control system and hydraulic system verification tests, as well as limited aircraft-level verification activities. The test bed is composed of flight control system LRU, actuator, control surface, hydraulic system, power supply system, and other aircraft system LRUs which are important to the flight control system.

Figure 1.11 Flight control system integration test process of Boeing 777 aircraft.

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The flight control system integration laboratory and engineering simulator are mainly used for the evaluation of the aircraft-level flight control system. The flight control system integration laboratory is composed of all aviation system LRUs, the power supply system, and cabin systems related to the flight control system. The engineering simulator, consisting of all vision systems and all LRUs that are critical to the pilot’s operation, is used by the pilot to evaluate the flight quality and maneuverability of the aircraft. The top-level integration of the flight control system is completed in the ground test and flight test of the test aircraft. After an assembly is replaced, the same test procedures should be finished on a special test bed, and then the test should be finished in the next integrated level laboratory, and finally the test should be finished on an aircraft. Then the test analysis report, test summary report, and formal documents should be formed. 6. Analysis There are three types of analysis: performance analysis, availability analysis, and security analysis. Performance analysis mainly evaluates the performance and operation of systems under typical acceptable environment limits and failure conditions. Mainly linear and nonlinear time domain and frequency-domain analysis methods are adopted. With regard to the automatic landing function of AFCS, the Monte Carlo method is used to estimate the grounding performance of the automatic landing system. Availability analysis is primarily static analysis, which evaluates the ability of a system to meet nonsecurity requirements (i.e., commitments to dispatch and maintenance). Security analysis is primarily used to demonstrate the required security level that a system can provide under normal and abnormal operational failure conditions. 7. System test and analysis Some analyses and tests are arranged for the verification of systems beyond the flight control system, such as the structural mode coupling test and the electromagnetic compatibility test, but these verified design requirements must also support the whole flight control system with regard to its requirements verification.

1.6 Flight control system test verification 1.6.1 Introduction to the background aircraft flight control system Large aircraft include civil transport aircraft and military transport aircraft. As they have fundamentally the same basic function, performance, and safety requirements for flight control system but obvious differences in the basic architecture of a flight control system,

Introduction 29

Figure 1.12 Composition and mutual relationships of the background aircraft flight control system.

it is difficult to find a flight control system that can cover all aircraft and to carry out research on the test techniques for a flight control system. For the sake of research convenience and generality, the author of the book has constructed an aircraft flight control system for this study, which is called the background aircraft flight control system. The background aircraft flight control system is a three-axis four-redundancy “fly-by-wire 1 analog backup 1 mechanical backup” control system consisting of pilot control units (PCU), FBW flight control system (EFCS), mechanical flight control system (MFCS), HLCS, and AFCS. By controlling 28 aircraft control surfaces (including 2 elevators, 1 horizontal stabilizer, 1 rudder, 2 ailerons, 10 spoilers, 8 leading edge slats, and 4 trailing edge flaps), the lift control is realized at stages of aircraft attitude control, flight path control, boundary limit/protection, ground drag increase control, and aircraft takeoff and landing. The composition and mutual relationships of the background aircraft flight control system are shown in Fig. 1.12. 1. PFCS adopts the architecture of “full-authority fly-by-wire control 1 mechanical backup control” to provide normal, degraded, simulated backup, and mechanical backup control. 2. EFCS adopts the architecture of “digital 4 3 2 redundant primary flight computer (PFC) 1 actuator control electronics (ACE)” to provide normal, degraded, and simulated backup control.

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3. MFCS adopts a cable transmission device to control ailerons and horizontal stabilizers to realize the emergency mechanical backup control of aircraft. 4. HLCS adopts the architecture of “digital 2 3 2 redundant Flap-Slat Control Unit (FSECU) 1 backup control” to provide normal, degraded, and backup control. 5. AFCS adopts the architecture of “digital 2 3 2 redundant computer control (AFCC)” to provide functions including automatic pilot, flight director, autothrottle, automatic approach, and automatic navigation. 6. The flight control system adopts MIL-STD-1553B as the internal bus, PFC as the BC (bus controller), and ACE, FSECU, and AFCC as RTs (remote terminals). 7. Ailerons, elevators, rudders, and other control surfaces are driven by an electrohydraulic servoactuator (an integrated control valve 1 two actuator cylinders) and the horizontal stabilizers are driven by ball screws controlled by a dual hydraulic motor. 8. The multifunctional spoiler is driven by an electrohydraulic servoactuator and the ground spoiler is driven by a hydraulic actuator cylinder controlled by a solenoid valve. 9. The flaps are driven by “dual-motor power drive unit (PDU) 1 central coaxial torsion bar 1 ball screw” and the slats are driven by “dual-motor power drive unit (PDU) 1 central coaxial torsion bar 1 gear/rack.” 10. The flight control system cross-links with the air data computer (hereinafter referred to as air equipment), inertial/satellite integrated navigation equipment (hereinafter referred to as inertial navigation equipment), integrated processor flight control computer, central maintenance system (CMS), central warning system, landing gear control system, antiicing/deicing system, autothrottle actuator, engine parameter acquisition and processing equipment (hereinafter referred to as engine parameter), flight parameter acquisition unit, and the mechanical and electrical management computer to realize control, display, warning, and maintenance functions. The architecture of the background aircraft flight control system is shown in Fig. 1.13. The electrohydraulic actuator of the background aircraft flight control system is powered by three sets of hydraulic sources. When No.1 or No.2 hydraulic system fails, all primary control surface of the aircraft will work and the flight quality of the aircraft will not be affected. When No.1 and No.2 hydraulic system fail, the critical control surface of the aircraft will still work but the flight quality is degraded. When No.1 and No.2 hydraulic system and hydraulic pump of No.3 hydraulic system fail, the aircraft’s critical control surface actuator will be powered by a ram-air turbine (RAT) drive pump and work, but the flight quality is degraded. To ensure the flight safety and mission reliability of the aircraft, the power supply system setting of the electric equipment of the flight control system has enough redundancy. Each computer/controller is equipped with two power modules and is powered by a two-circuit or three-circuit power supply system.

Introduction 31

Figure 1.13 Architecture of background aircraft flight control system.

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Through the cross-link with air equipment and inertial navigation equipment through the ARINC429 bus, EFCS receives altitude, speed, and other information to complete the control law parameter adjustment of the flight control system. EFCS receives aircraft attitude and other information to complete the attitude maintaining and protection functions of bar after release. EFCS sends the pilot control command, flight control system status and failure information to the flight parameter acquisition unit through RS422 bus and hard wire. It cross-links with the avionic integrated processor through AFDX, receives the aircraft weight, center of gravity, maintenance information, airdrop command, and cabin door opening information, and uploads the flight control system BIT information and status information to the CMS and display processing unit. In addition, it receives engine speed, throttle position, and thrust reverser information through ARINC429 to complete the logical judgment of takeoff configuration and the logical judgment of ground automatic deceleration of spoiler. It also receives the landing gear control, wheel load and wheel speed information, antiicing/deicing system status and icing information through hard wire to complete the logical judgment of the ground automatic deceleration of spoiler and ensure the safety protection when there is aircraft icing. The external interface relationship of EFCS is shown in Fig. 1.14.

Figure 1.14 Relationships between EFCS and other systems.

Introduction 33 HLCS cross-links with the avionics system through the ARINC429 bus by the remote data concentrator (RDC) of the avionics system and completes the flight control system BIT, maintenance inspection, recording of flight parameter acquisition unit, failure warning, information display, etc. Other required information is obtained from PFC through the MIL-STD-1553B bus. The HLCS external interface relationship is shown in Fig. 1.15. AFCS cross-links with air equipment, inertial navigation equipment, radio altimeter, central warning computer, and flight parameter acquisition unit through the ARINC429 bus and cross-links with the flight management computer, CMS, display processing unit, and engine parameter acquisition unit through a RDC, so as to complete functions such as automatic pilot, automatic navigation, automatic approach, and to realize system failure warning, status display, and recording. AFCS control command is sent to EFCS via the MIL-STD-1553B bus to realize the threeaxis control of the aircraft. The cabin lighting system provides lighting signals to the automatic flight control panel. The automatic flight control panel sends the pilot control command to the flight parameter acquisition unit. The external interface relationship of AFCS is shown in Fig. 1.16.

Figure 1.15 High lift system external interface relationship.

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Figure 1.16 Interface relationship between automatic flight control system and other airborne system.

1.6.2 Content of a verification test Verification test activities of the flight control system are reflected in all stages of development. These activities promote reverse information feedback in the development process, contribute to technical or engineering changes, improvements, and optimization, and finally cause the loop iteration of the work within the system development cycle. Generally, the flight control system-level verification test has five stages: validation of users’ requirements, validation of definition stages, validation of design and implementation, integration and test verification, and onboard ground test and flight test. 1.6.2.1 Validation of users’ requirements The requirements of users should be described and validated from the perspective of the whole aircraft and corresponding specifications of user requirements should be formulated. In most cases, the initial requirements from users are not very clear and they should be

Introduction 35 described in a complete way after a strict validation process. In general, users’ requirements that are directly related to the flight control system and need to be validated include: 1. 2. 3. 4. 5. 6. 7.

definition of flight envelope; requirements on control quality and flight quality; division of working modes; requirements on reliability, security, availability, and assurance; requirements on maintenance and testability; potential upgrade and expansion requirements; and design methods and standards that must be used in design and implementation.

In general, users’ requirements should be validated from the aspects of design and manufacturing. The integrity, consistency, and design traceability of requirements should be investigated in a systematic way. Meanwhile, the developing simulation analysis tools and methods at this stage can be fully used to have modeling and analysis of the users’ requirements so as to ensure that the requirements have the above performances. 1.6.2.2 Validation in the definition stage In the definition stage, the flight control system should fully refer to the related experience of existing models, combined with specifications and protocols in other relevant fields (such as aerodynamic configuration, propulsion unit, aerodynamics, vehicle structure, electronic system) and specifications and protocols of the flight control system and its subsystems (such as control law, computer, sensor, actuator, and other subsystems), and also consider the cross-links between them. Adequate and full consideration should be given to various factors and constraints and compromises need to be made in case of conflicts, so that system requirements can be effectively validated, analyzed, and defined. According to the requirements analysis, the flight control system scheme demonstration can be generally divided into system-level and subsystem-level scheme demonstration and it should be validated in relevant forms (such as communication within the project team, internal and external expert review, and necessary simulation calculation). 1.6.2.3 Verification test in the design and implementation stage Flight control system, subsystem design, and implementation include a series of verification test activities. In the engineering design, processing, manufacturing, assembly, and commissioning of the flight control system and subsystem, various effective verification activities should be carried out to identify problems, to improve the design, and to optimize and iterate the system. Possible activities include computational checks, prototype manufacturing, assembly, and subsystem test. After the corresponding verification, commonly comes the unit scheme improvement, subsystem-level scheme optimization, and even system-level scheme change. The iterations

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of design optimization and improvement depend on the accuracy of the requirements proposal, the designers’ understanding of the users’ requirements, the engineering experience of the subsystem and unit designers, and even the assurance of the management and quality system. 1.6.2.4 Integration and test verification After the unit of the flight control system under development is designed, produced, and debugged and validated through independent verification tests, the system development proceeds to the stage of system integration and test verification. This process should be carried out in a simulated environment that is as real as possible to verify and validate the whole system, with the main purposes as follows: 1. To verify the interface correctness and compatibility between subsystems and units of the flight control system. 2. To verify that the function and performance of the whole flight control system meet the requirements of design specifications. 3. To expose and determine potential hardware and software failure modes and mechanisms and make corrections to increase the reliability of the flight control system. 4. To familiarize the pilot with the control characteristics of aircraft through man machine integrated tests, expose potential design defects as early as possible, formulate operational procedures and emergency treatment measures, and ensure flight safety. 5. To validate the accuracy of the design requirements of the whole flight control system, that is, to validate if the flight control system design meets the requirements of aircraft use. The corresponding acceptance test procedure (ATP) should be adopted to validate that the flight control system meets the requirements of the design specifications. The integrated test of the flight control system is generally divided into two stages: the integrated test on bench and the “iron bird” integrated test. 1.6.2.5 Onboard ground test and flight test After the “iron bird” integrated test, the flight control system should be installed on the aircraft according to the specified technical requirements for an onboard ground test. The onboard ground test generally includes three parts: flight control system performance check test, structural mode coupling test, and all-aircraft electromagnetic compatibility test. The purpose of the performance check test is to run the real hardware and software of the flight control system in a real aircraft, check the functions and performance of the flight control system and compare the results with those from the “iron bird” integrated test. The main purpose of the structural mode coupling test is to obtain the open-loop frequency

Introduction 37 response of the aircraft flight control system including the structural mode transfer function to determine if its surge margin meets the design requirements. The purpose of the all-aircraft electromagnetic compatibility test is to check if there are electromagnetic compatibility issues inside or outside the airborne systems. After the verification of the flight control system through the “iron bird” integrated test and onboard ground test, it enters the flight test stage, which is the final stage of the flight control system evaluation and also the test with the most authoritative results. As the essential performance of the system can only be shown through the test flight of various flight subjects in the flight envelope, this stage is crucial for the development of the flight control system. Flight tests are usually carried out on several aircraft, with a specific test target for each aircraft. The specifics of flight test are mainly based on relevant standards and regulations (such as the tactical and technical indexes and technical status of model setting proposed by the use department) and the test flight assignment book submitted by the development unit, mainly aiming to test if the functions and performance indexes of the flight control system meet the requirements of the design specifications. For a newly developed flight control system, the flight test generally has two stages: development (adjustment) stage and validation stage. In the development stage, problems of the flight control system, hardware, and software are fully exposed and solved, and then control law is optimized and finally determined. In the validation stage, the flight control system and flight quality are verified and validated to meet the overall design requirements.

1.6.3 Procedures of a verification test The most important lesson learned is that plan shall be formulated and management shall be carried out for the verification of the flight control system from the start of the project. At the early stage of the project, adequate verification plans and procedures shall be prepared, with suitable verification tools and adequate test verification facilities. Test verification facilities mainly include independent laboratory, integrated laboratory, aircraft available for test flight, and equipment used for onboard ground test and flight test. After requirements are approved, requirements verification methods and criteria can be determined and documented. An appointed engineering representative needs to ensure the full coverage of the verification. At this stage, a senior engineer should be appointed as the person in charge of each requirement who will have abundant experience to participate in the verification plan and authority department so as to ensure strict and overall requirements verification.

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The requirement itself is deemed wrong if there is a need to ask whether there is a problem with verification methods or criteria, and a description is therefore required. If the problem is not addressed, designers may not be clear about the desired requirement and may fail to ensure requirement conformity. If one requirement does not have responsible personnel in the early stage of the project or there is any similar problem, it means this requirement may have been ignored in the design stage, so that the system cannot meet this design requirement. Especially in the late stage of the project, there may be no one responsible for the requirement due to a lack of specific responsibility regulations and this requirement may not be fully verified. Another important lesson learnt is that all departments of the flight control system should use the same verification process and tracing tools, the lack of which may lead to the failure to receive requirements or notification in a timely manner. The use of different databases or platforms to track and manage verification activities may lead to disordered engineering and state management. On the basis of domestic and foreign civil aircraft development experience and design concepts, the design and verification process of the flight control system is summarized and analyzed. The whole double “V” model is divided into two parts: requirements definition and validation, and system integration and verification, as shown in Fig. 1.17. To gradually achieve the design objectives and clearly identify the development milestones and responsibilities, the flight control system verification is divided into five levels according to the traceability breakdown relationship. The main levels are as follows: 1. Unit level: in the detailed design stage, complete documents of detailed design requirements of unit and software design requirements, according to the design requirements of the last level, validate the design requirements through review, analysis, or test and formulate test plans and procedures. In the flight control system integration and verification stage, there is complete software testing and unit function and performance testing through simulation and an independent test bed, so as to verify if the unit meets unit-level requirements, including the quality qualification test and acceptance test. 2. Equipment level: in the detailed design stage, complete documents of hardware and software design requirements, according to the design requirements of the last level, validate the design requirements through review, analysis, inspection, or testing and formulate test plans and procedures. In the flight control system integration and verification stage, there is complete software code testing and equipment function and performance testing through simulation and an independent test bed, so as to verify if the hardware and software

Introduction 39

Figure 1.17 Model of flight control system design and verification process.

meet design requirements, including the quality qualification test, acceptance test, and software test. 3. Subsystem level: in the detailed design stage, complete documents of subsystem design specifications, according to subsystem-level design requirements, validate the design requirements through review, analysis, inspection, or testing and formulate test plans and procedures. In the system integration and verification stage, there is complete subsystem integration through simulation and the subsystem test bed, so as to verify if the subsystem meets design requirements, including the quality qualification test and the hardware and software integrated test.

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4. Flight control system level: in the preliminary design stage, define the architecture and functions of the flight control system according to aircraft-level requirements, form specifications of the flight control system, validate the flight control system level design through review, simulation, and analysis and ensure the accuracy and integrity of flight control system level requirements. In the flight control system integration and verification stage, through analysis, review, and inspection, validate the flight control system level design specifications through the development test and flight control system integration test and verify if the functions, performance, and software and hardware of the flight control system meet the requirements of the flight control system design specifications, including the flight control system integration test, software and hardware verification test, “iron bird” integrated test, engineering simulator, and formal airworthiness verification test. 5. Aircraft level: define the design requirements of the flight control system according to aircraft design requirements and objectives, validate the design requirements through simulation and analysis, and ensure their accuracy and integrity. In the flight control system integration and verification stage, validate flight control system design requirements through analysis, review, inspection, and the development test, “iron bird” integrated test, and onboard ground test and show if the design of the flight control system meets the requirements of design specifications. In the flight test and certification stage, validate the flight control system design requirements and aircraft-level requirements through analysis and review and flight test and show the design of the flight control system satisfies the flight control system requirements and aircraft-level requirements, including the verification of the aircraft interface, onboard integration and airworthiness certification test, and prove the flight control system conforms to the airworthiness requirements to the airworthiness authority.

CHAPTER 2

Unit test of the flight control system The airborne unit is an integral part of an aircraft and its system. It is the key to the realization of aircraft functions, completion of missions, and flight safety, as well as one of the decisive factors for the smooth and successful development of a type of aircraft. Airborne unit development is the most important work throughout the development of aircraft type. Design department, customers, main machine, and quality supervision departments pay close attention to the whole process from the determination of unit design requirements, selection of suppliers, development of F type (principle scheme), development of C type (prototype sample), development of S type (installed test) to verification and finalizing, thus ensuring the development quality and progress of the airborne unit is particularly important. As large aircraft undertake long flight duration, high security requirements, complex missions, and many system cross-links, airborne units contain a large number of LRUs (several thousand and even tens of thousands of LRUs). In addition, as hundreds of units or departments and tens of thousands of personnel participate in the development, airborne unit development is a huge project of aircraft type development. Unit testing is the most basic technical means to validate and continuously optimize and improve the functions, performance, and environmental adaptability of the unit. The test covers all test items necessary for each stage of the airborne unit development by the airborne unit development unit. Generally, the development unit organizes or entrusts a third party to undertake the test. Among them, the development stage transition test, test of key items, airworthiness verification test, first flight safety test, and qualification test should be attended and witnessed by representatives of the main machine unit, customers, and the quality supervision department. This chapter mainly introduces the functional (performance) test, environmental adaptability test, and durability test related to the qualification test of the airborne unit of the flight control system. The content and methods of other tests in the development process are similar to that of the qualification test and can be taken as reference. Aircraft system tests and flight tests involving the airborne unit will be described in Chapters 5
8 and tests related to airworthiness requirements will be described in Chapter 9, Airworthiness Verification Test of Flight Control System.

Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00002-4 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

41

42

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2.1 Overview The physical properties of the airborne units of the flight control system are generally divided into electronic, electromechanical, and mechanical products and they include fields of automatic control, electric power, microelectronics, electromagnetism, machinery, hydraulics, and software engineering. According to the requirements of the aircraft airborne unit qualification test (functional and performance test, including the environmental adaptability test covering the natural environment, mechanical environment, and electromagnetic environment and the durability test) and the national military standards (military aircraft) or civil aircraft airworthiness standards (civil aircraft) as well as special specifications for airborne units, test requirements, test items, test methods, and judging criteria of units are determined. Different physical properties, fields, and unit installation locations lead to different test content and the requirements of different systems and units. For example, the impedance characteristic test and oil immersion test of electrohydraulic servo actuators commonly used in flight control systems are special test items of airborne units of the flight control system.

2.1.1 Basic process of unit development The development of airborne units starts from the signing of the technical agreement or development contract of the airborne units. After the agreement takes effect, the development of airborne units can be carried out. Generally, the technical agreement is attached with the design requirements and interface control documents of the units. During the development process, the design and testing of units shall be carried out in accordance with the provisions on the management of units and type development, and with the requirements of relevant specifications. Airborne units are developed in the sequence of scheme stage (F), prototype sample stage (C), sample stage (S), type finalizing stage (D), and batch production stage (P). Each stage completes its design, trial production, and verification. F stage: The development of new airborne units generally starts from the F stage. The principle test of such airborne units should be carried out in this stage to verify the design scheme. After the scheme passes the review, the units will proceed to the C stage. C stage: The development of improved and retrofitted airborne units generally starts from the C stage. Relevant tests are carried out according to the specifications or technical conditions of airborne units to confirm they meet the requirements of the product specification. Only after they pass the integrated test of the system or subsystem can they enter the S stage. S stage: In addition to the shelf airborne unit directly selected, other newly developed and improved and retrofitted airborne units shall go through the S stage and complete the

Unit test of the flight control system 43 qualification test in accordance with the qualification test outline. The units can only enter the D stage after they pass the qualification test and receive the test flight conclusions. D stage: The stage of process stability assessment of airborne units, that is, product type finalizing. P stage: The finished product is placed into the batch production stage.

2.1.2 Categories of unit tests According to the development process, testing of the airborne unit can be divided into the principle test in F stage, the prototype sample acceptance test in C stage, the ex-factory check and acceptance test in S stage, the special test of sample in S stage, the first flight safety test in S stage, and the qualification test in S stage. According to the nature of the test, the qualification test of the airborne unit can be divided into the functional and performance test, power supply characteristic test, dielectric strength test, stiffness and strength test, reliability test, life test (durability test), mechanical environment test, natural environment test, electromagnetic environment protection test, and special test (software testing and evaluation, testability test).

2.1.3 Basic principles for the selection of test items For the electronic, electromechanical, and mechanical airborne units of the flight control system, the basic principles for the selection of test items are as follows: 1. Test items can be selected according to product specifications and the technical agreement of airborne units or type development specifications or the airworthiness verification plan. 2. Test items can be selected according to the characteristics of electronic, electromechanical, and mechanical airborne units. For example, mechanical units do not need electromagnetic compatibility, lightning protection, and high-intensity radiated fields (HIRF) tests. 3. Tests can be divided according to the development stages. For example, the three-proof, lightning protection and HIRF tests may not be conducted in C stage. 4. The specific functional and performance test items can be determined according to the functional items and performance indexes of airborne units. 5. Life tests such as durability test shall be carried out for airborne units with life index requirements, and reliability related tests shall be carried out for airborne units and electronic and electrical units with requirements for mean time between failure index. 6. For airborne units with airworthiness requirements, the applicable airworthiness provisions shall be specified when conducting the verification test and the test shall be completed in accordance with the airworthiness management documents.

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2.1.4 Classification of the flight control system unit The flight control system will cause great differences in airborne unit configuration items and names due to different aircraft, system configurations, cross-links, and airborne unit configurations. The airborne unit configuration of the background aircraft flight control system described in Section 1.6 is an example of the basic categories of airborne units. According to the physical properties, airborne units can be divided into electronic, electromechanical, and mechanical units, as shown in Table 2.1. According to their functions, airborne units can be divided into control and display units, computational control units, control units, mechanical drive units, mechanical transmission units, mechanical actuator units, and sensor units, as shown in Table 2.2.

2.1.5 Basis and requirements for the preparation of the unit test outline Tests of airborne units are based on technical agreement, design requirements, and technical specifications of airborne units. Test outlines shall be independently prepared for the first flight safety test and type finalizing/qualification test and they shall be separately compiled into books and jointly bound into a volume. The outline of the first flight safety test shall include the outline of the functional and performance test, the outline of the environmental test (natural environment and mechanical Table 2.1: Categories of flight control system airborne units by physical properties. No.

Category

1

Electronic

2

3

Applied physics knowledge

Microelectronics, electromagnetism, computer science, software engineering Electromechanical Microelectronics, electromagnetism, solid mechanics, fluid mechanics, structural mechanics

Mechanical

Solid mechanics, fluid mechanics, structural mechanics, mechanical principle

Typical airborne units Flight control computer, automatic flight control computer, flaps and slats controller Flight control panel of mechanical principle, trim control panel, automatic flight control device, flaps and slats control handle, air brake control handle, horizontal stabilizer control handle, aileron actuator, rudder actuator, elevator actuator, spoiler actuator, control command displacement sensor, angular rate gyroscope assembly, acceleration sensor, angle of attack/angle of sideslip sensor, static pressure sensor, total pressure sensor Stick back, pulley, cable, guide part, spring load mechanism, ball screw mechanism, variable-angle reducer, rotary actuator, cable tension compensator

Unit test of the flight control system 45 Table 2.2: Categories of flight control system airborne units by functions. No.

Category

Main functions

Typical airborne units

1

Control and display

2

Computational control

Status selection button and status, control surface deflection angle and warning information display Information communication, control logic and control law resolution

3

Control unit

4

Mechanical drive

5

Mechanical transmission

Directly accepts the pilot’s control force and control displacement, transmit or converts to electrical signals to external mechanism, and output as operational command Not directly mechanically connected to the driven object, but exerts acting force on driven object through many drive links Transmits mechanical displacement and force (or torque)

Flight control panel, horizontal stabilizer trim and cutoff control panel, trim control panel, automatic flight control unit Flight control computer, actuator controller, automatic flight control computer, flaps and slats controller Flaps and slats control handle, air brake control handle, horizontal stabilizer control handle, steering wheel (column), pedal control mechanism

6

Mechanical actuator

Connects with the actuated object mechanically and exert acting force to it

7

Sensor

Converts pilot’s control command, aircraft attitude, control surface position, speed, overload, atmospheric dynamic and static pressure and other physical signals to electrical signals according to certain law

Flaps and slats drive unit

Stick back, pulley, cable, guide part, spring load mechanism, variable-angle reducer, rotary actuator, cable tension compensator Deflection limiter, back actuator, aileron trim mechanism, rudder trim mechanism, elevator trim mechanism, aileron actuator, rudder actuator, elevator actuator, horizontal stabilizer actuator, spoiler actuator, air brake actuator, wingtip braking device, rotary actuator, ball screw actuator Controls command force sensor, control command displacement sensor, control surface position sensor, angular rate gyroscope, acceleration sensor, angle of attack/angle of sideslip sensor, static pressure sensor, total pressure sensor

environment), the outline of the electromagnetic compatibility test, and the outline of the durability test. For the type finalizing/qualification test of airborne units, it is necessary to write a preparation description of the test outline. The preparation description shall describe the number, composition, state, function (performance) of test pieces and the sequence of environmental tests (the test sequence shall meet the requirements of relevant standards), the information about the test undertaker, test items, test environment, and preparation of the test outline, the test items that have been completed in early stage and the test results, as well as the applicability of test items, special requirements of tests, and the principles of

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reduction and supplement. The preparation description shall include the overview, test pieces, division of labor, functional (performance) test, environmental qualification test, electromagnetic environment effect test (electromagnetic compatibility, lightning effect, HIRF), durability test, testability test, software testing and evaluation, airworthiness review items, and other descriptions. The outline of the type finalizing/qualification test shall include the outline of the functional and performance test, environmental qualification test, reliability qualification test, durability test, testability test, software testing and evaluation, electromagnetic compatibility test (including electrostatic protection test), direct lightning effect protection, indirect lightning effect protection, and HIRF protection test. Software testing and evaluation is introduced in Chapter 3, Verification and Validation of Flight Control System Airborne Software. What should be noted is that without particularly significant changes, the conclusions of the first flight test can be used to support the design qualification test.

2.2 Basic functional/performance test The basic functional and performance test is a compulsory test for all airborne units to assess if the airborne unit meets the technical agreement and design requirements. The test basis is the technical agreement, design requirements, and type reference specifications of airborne units.

2.2.1 General requirements of test The functional and performance testing of airborne units is carried out under normal temperature and the atmospheric conditions at the test site and only the functions and performance of the test pieces under normal temperature are assessed. According to the specific requirements for the tested pieces, special test conditions, such as test bed, exciter, test equipment, load environment, and other supporting facilities, shall be prepared. 1. Environmental conditions Test temperature: 15 C
35 C Relative humidity: 20%
80% Air pressure: air pressure at test site Magnetic field intensity: ,1 Gs 2. Testing equipment The testing equipment used in functional and performance testing are mostly special equipment, such as the special test bed for actuators and the cabin control mechanism and the turntable of exciter, gyro, and accelerometer that are used to simulate the

Unit test of the flight control system 47

3.

4.

5.

6.

mechanical and electrical input of tested airborne units. They should be all within the validity period of calibration and the technical indexes of them should meet the test requirements of the unit under test. Test equipment All technical indexes of test equipment shall be inspected as acceptable in accordance with relevant national standards or metrological specifications and within the validity period of calibration, with accuracy not lower than one third of the accuracy of the performance indexes of the unit under test. Energy configuration The energy system used in the functional and performance test covers the power source and hydraulic source and the energy configuration shall take into account the energy demands of the airborne units in the test. Power source: generally DC 117 V, 28 V, 6 15 V, AC 380 V 400 Hz, 115 V 400 Hz, 36 V 400 Hz, 7 V 1800 Hz. The power source should be configured according to the power supply system, power, precision, and number of channels of airborne units. Hydraulic source: the hydraulic source of corresponding redundancy, pressure system, and flow shall be configured according to the operating requirements of the aircraft hydraulic system and airborne units. For example, a hydraulic source with flows of 150, 160, and 80 L/m should be configured for three independent sets of 21 MPa pressure system. Interface requirements The mechanical and electrical interfaces of the testing equipment and special test bed shall match with the tested airborne units. During the development of the testing equipment, the interface requirements for the test equipment should be put forward. In cases where the interface does not match with the unit, a fixture or adapter cable can be used to solve the problem. Regulations on quality assurance The functional and performance test of airborne units is part of the quality conformance check and ex-factory acceptance and qualification test. Unless otherwise specified, quality conformance check and ex-factory check and acceptance shall be conducted for all units and at least two test pieces shall be selected to have the qualification test. Test pieces submitted for inspection must be verified products. If any test item fails to meet the criteria, the test piece is deemed unacceptable and the test piece shall be prepared for inspection again until it passes the inspection.

2.2.2 Control and display unit According to functions, the control and display units of the flight control system can mainly be divided into flight control panel, trim control panel, horizontal stabilizer position

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indicator, trim and cutoff control panel, and automatic flight control device. The flight control panel is taken as an example below to introduce the functional and performance testing of control and display units. Other control and display units have similar tests. 2.2.2.1 Test objectives and test requirements The test aims to check and test the basic functions and performance of the flight control panel to validate that it meets the design requirements and works within the specified range. The flight control panel shall be tested in accordance with the requirements of acceptance specifications and product specifications one by one and each test result shall conform to the criteria. The tester for the flight control panel must be constructed and the inspection and test of the flight control panel shall be carried out on the tester. The general test requirements in Section 2.2.1 shall be followed unless otherwise specified. 2.2.2.2 Test items and test methods The main test items of the flight control panel include power supply voltage, power supply impedance, indication by indicator lamp, brightness adjustment, switch quantity output, and switch quantity input. 1. Appearance structure Inspection items of appearance structure inspection include appearance, weight, overall dimensions, installation form, interchangeability, error prevention, and interface and the inspection methods include visual inspection, electronic balance, and ruler. 2. Power supply impedance Digital multimeter is used to measure the impedance value of the power supply from the high terminal to the low terminal of the airborne unit electrical connector. 3. Indication by indicator lamp Use a test cable to connect the tester with the tested control panel, connect power to the tester and tested control panel, toggle the lamp check switch on the tester panel to conduct the lamp check test, and then visually check if the status indicator lamp on the control panel is on. 4. Brightness adjustment Use a test cable to connect the tester with the tested control panel, connect power to the tester and tested control panel, toggle the lamp check switch on the tester panel, light all status indicator lamps, rotate the “power for light guide plate” and “power for signal lamp” knobs on tester to have brightness adjustment of the light guide plate and signal lamp, and then visually check if the brightness of the light guide plate and signal lamp changes with the operation of knob. 5. Switch quantity output Use a test cable to connect the tester with the tested control panel, connect power to the tester and tested control panel, toggle the tested switches on the control panel in sequence, and visually check and record the switch status and signal output characteristic.

Unit test of the flight control system 49 Table 2.3: Expected value of power supply impedance, lamp check test, and brightness adjustment test. No. 1 2 3 4 5 6

Test items

Test items

28 V power-to-ground impedance

28 V input 1 28 V input 2 Power supply for light guide plate Power supply for signal lamp

Light guide plate power-to-ground impedance Signal lamp power-to-ground impedance Lamp check test Brightness adjustment test

Expected value

6. Switch quantity input Taking the test of the four-redundancy switch quantity input as an example, use a test cable to connect the tester with the tested control panel, connect power to the tester and tested control panel, toggle the corresponding switch on the tester according to the switch status in Table 2.3, provide the four-redundancy tested signal, and observe if the tested indicator lamp is on. 7. Power supply voltage Use a test cable to connect the tester with the tested control panel, connect power to the tester and tested control panel, adjust the voltage of the DC stabilized power supply to the limit value with specified deviation, such as adjusting 28 V power to 18
32 V, check if the function (performance) of the flight control panel meets the requirements of the product specifications, and record inspection and measurements in the test results. 2.2.2.3 Judging criteria and results handling If the measurement results of the appearance structure inspection of airborne units are consistent with the expected value, the structure of the airborne unit is deemed acceptable. For example, the appearance shall be intact and not damaged, the connector shall not be bent or retracted, the weight and overall dimensions shall meet requirements, the locking mechanism of the mounting bracket shall be well locked and not release off, the electrical interface shall be error free and the plate layout shall meet design requirements. If the power supply impedance inspection result is consistent with the expected value in Table 2.3, the test item is deemed acceptable; otherwise, the test item is deemed unacceptable. When the lamp check switch is turned on, the status indicator lamp of the control panel shall be all on. When the lamp check switch is turned off, the status indicator lamp of the control panel restores to its initial state. If the indicator lamp display inspection result is consistent with the expected value in Table 2.3, the test item is deemed acceptable; otherwise, the test item is deemed unacceptable.

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If the status indicator lamp can be well controlled by the lamp check switch and the brightness of the status indicator lamp and the light guide plate can be adjusted through the “power for light guide plate” and “power for signal lamp” knobs, the test item is deemed acceptable; otherwise, the test item is deemed unacceptable. The expected values of the test are shown in Table 2.4. If the switch quantity output inspection result is consistent with the expected value in Table 2.4, the test item is deemed acceptable; otherwise, the test item is deemed unacceptable. The principle of signal voting for the four-redundancy switch quantity input check is that the voting output shall adopt three or more than three consistent input signals. The corresponding relationship between the tester switch state and the control panel status indicator lamp and the expected test results are shown in Table 2.5. If the switch quantity Table 2.4: Expected value of switch quantity output test. Control panel switch name

Gear

System power Off switch On Flight control Automatic computer switch Off Fly-by-wire Elevator Semimechanical transform band Mechanical switch Aileron transform Fly-by-wire band switch Semimechanical Mechanical Rudder transform Fly-by-wire band switch Mechanical

Tester indicator lamp

Output characteristic 28 V On On 15 V to ground Four-redundancy 15 V ground, one circuit of 28 V, and ground motor control

Expected result

L1, L2 L7
L10 L11
L16

Four-redundancy 15 V ground, one circuit of 28 V, and ground motor control

L17
L22

On 15 V to ground

L23
L26

Table 2.5: Expected results of “OFF” indication function test of primary flight control panel. Tester switch and status Control panel status indicator lamp “OFF” indicator lamp of primary flight control computer

Channel 2

Channel 3

Channel 4

Channel 5

Up Up Up Up Down Others

Up Up Up Down Up

Up Up Down Up Up

Up Down Up Up Up

Expected results

Unit test of the flight control system 51 inspection results are consistent with the expected results, the test item is deemed acceptable, otherwise, the test item is deemed unacceptable. If the power supply voltage test result is consistent with the product specification, the test item is deemed acceptable; otherwise, the test item is deemed unacceptable.

2.2.3 Computer unit The flight control computer mainly includes the primary flight control computer, actuator controller, automatic flight control computer, and flaps and slats controller. The primary flight control computer (hereinafter referred to as the flight control computer) is taken as an example below to introduce the content of the functional and performance testing of computer units. The content of the functional and performance tests of other computer units are similar. The flight control computer has functions including the control/monitoring of branch computation, multichannel analog input port, multichannel discrete input/output, multichannel bus, synchronization, and channel voting. Performance mainly covers accuracy, degree of asynchronization, cross data transmission, main processor, input/output processor, analog input/output, discrete input/output, bus, in-channel self-monitor threshold, power consumption characteristics, secondary power supply characteristics, and functional board cards. 2.2.3.1 Test objectives and test requirements Through graded testing, the functions and performances of each functional board card, complete machine, and software and hardware of the flight control computer can be fully tested to ensure that the computer has reached the design requirements before the system test and works properly in the validation scope. The test of the flight control computer shall be carried out in accordance with the requirements of acceptance specifications and product specifications item by item, and each test result shall conform to the criteria. The functional and performance test of the flight control computer should be carried out on a special test device. The test device is divided into manual and automatic test devices. Automatic test device are generally adopted for a flight control computer, including an automatic test device for the board card, integrated test facilities for the flight control computer, and development integrated facilities for software. The test device shall generally have the following functions: 1. supplying all necessary power for the tested board; 2. generating different input/excitation signals necessary for the board card;

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3. having the analog load ability; 4. being able to switch between input signal and load; 5. being able to measure different signals accurately, such as analog quantity, discrete and bus signal; 6. having the ability to process data and verify test results; and 7. having the functions of displaying, printing, and reading the stored data. 2.2.3.2 Test items and test methods The functional and performance test of the flight control computer is mainly divided into the board card test, complete machine test, and software and hardware integrated test. Specific test items include case structure inspection, power supply characteristic test, performance test of functional board card, and computer subsystem test. The test items and methods are as follows: 1. Case structure inspection The inspection items of the case structure inspection include appearance, weight inspection, installation form, interchangeability, error prevention, cooling mode, and electrical interfaces. The appearance, weight, interchangeability, error prevention, cooling mode, and electrical interfaces are inspected through visual inspection and with electronic scale, measuring scale, and special test equipment. 2. Power supply characteristic test Use a board card automatic test device to run a macroprogram and collect the external input power of the flight control computer channel by channel, the voltage in the MBIT test chamber and all power supply values for external equipment, and then compare them with expected values. 3. Performance test of functional board card The performance test items of the functional board card shall be determined according to the types of board card inserted in the flight control computer. They include the functional (performance) test of the PS board, the functional (performance) test of the CPU board, the functional (performance) test of the DIO board, the functional (performance) test of the AIO board, and the functional (performance) test of the MBI board. 1. Functional (performance) test of the PS board. Use a board card automatic test device to conduct the PS test and to record test data. Test items include the power supply effectiveness test and power supply acquisition value test. 2. Functional (performance) test of the CPU board. Use an automatic test device to conduct the test and record test data. Test items include RAM test, ROM test, timer interruption test, processor instruction test, watchdog test, DFTI test, IRQ test, and IDCE test.

Unit test of the flight control system 53 3. Functional (performance) test of the DIO board. Use an automatic test device to conduct the auxiliary test and record test data. Test items include discrete input/output test and discrete quantity input/output. 4. Functional (performance) test of the AIO board. Use an integrated test device to conduct the auxiliary test and record the test data. Test items include analog input test, analog input BIT test, sensor excitation test, and analog input monitoring port test. 5. Functional (performance) test of the MBI board. Simulate the bus communication between devices, conduct the auxiliary test through an integrated test device, and record the test data. Test items include BC-RT test of MIL-STD1553B bus, RT-BC test of MIL-STD-1553B, and ARINC429 bus receiving/sending test. 6. Computer subsystem test Use a computer automatic test device to conduct the test and record the test data. Test items mainly include case ID number test, configuration table version number test, digital quantity synchronization test, discrete quantity synchronization test, and channel fault logic test. 2.2.3.3 Judging criteria and results handling 1. If the measured results of case structure inspection are consistent with the expected values, the case structure will be deemed acceptable. That is, to say, it has an intact appearance, the plug-in is not retracted, the weight meets the requirements, the locking mechanism of the mounting bracket locks securely and the electrical interfaces, error prevention, and vent port meet requirements of the drawings. 2. If the macroprogram does not report faults during the power supply characteristic test and the test values recorded by the automatic test device are consistent with the expected values, the power supply is deemed acceptable. The expected results of the power supply characteristic test are shown in Table 2.6. 3. An integrated test device is used for the performance test of the functional board card. If no faults are reported by the test program in the test process and the results meet the Table 2.6: Expected results of power supply characteristic test. No. 1 2 3 4 5 6 7

Test content

Test items

External input power 5 V output test 15 V output test Excitation power supply voltage Excitation power supply frequency Sensor power supply Sensor power supply

28 V 5V 15 V 7V 1800 Hz 15 V 215V

Expected results

Remarks

54

Chapter 2 Table 2.7: Expected values of power collection data.

No. 1 2 3 4 5 6 7 8

Test content

Test items

Power supply effectiveness test

PSV_ 1 28 V PSV_ 1 5 V PSV_ANALOG 7 V 1800 Hz 28 V 15 V 215V 5V

Power collection data test

Expected results

Remarks

requirements on expected values, it is deemed acceptable, or it is deemed unacceptable. The expected values of power collection data are shown in Table 2.7. 4. An integrated test device is used for the computer subsystem test and if the test results conform to the requirements on expected values or no faults are reported by the test program, the test is deemed passed; otherwise it is deemed failed.

2.2.4 Manipulator unit The manipulator unit is used to directly receive the pilot’s control force and displacement and transmit or convert them to electrical signals to the external mechanism as the pilot’s control instructions. Typical manipulator units include flaps and slats control handle, air brake control handle, horizontal stabilizer control handle, steering wheel (column), control mechanism, and pedal control mechanism. The main performance indexes include operating stroke, operating force, polarity, stop, return performance, operating force
displacement curve, linearity, hysteresis, frequency characteristic, and step characteristic. 2.2.4.1 Test objectives and test requirements It aims to test the basic functions and performance of the manipulator unit, to ensure it meets the requirements of the special technical conditions of the product and it works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to criteria. 2.2.4.2 Test items and test methods The main items of the functional/performance test of the manipulator unit are resistance, insulation resistance, dielectric strength, function, operating stroke, operating force, polarity, stop, return performance, operating force
displacement curve, linearity, hysteresis, frequency characteristic, and step characteristic.

Unit test of the flight control system 55 Pretest preparation: install the unit to be tested on the test bed, the installation and commissioning of the sensor and loading equipment shall meet requirements and the whole test system shall be debugged to conform to specifications. The installation, connection, mechanical load, electrical load, and energy of the unit under test shall be consistent with the actual installed condition or follow the requirements of the product specification. 1. Resistance measurement Conduct resistance or on
off measurement for each coil and electrical component of the unit under test and convert the resistance to that under 20 C. 2. Insulation resistance Use a 500 V tramegger (250 V tramegger for high-altitude test) with precision level no lower than level 2.5 to measure the insulation resistance under relative humidity of 45%
75%. 3. Dielectric strength Increase the voltage between circuits and between the circuit and case to the specified voltage, keep for 1 minute and ensure there is no breakdown. During the test, uniformly increase the voltage to the specified value and the increase speed should ensure the voltage reading can be read. One unit under test can only have three dielectric strength tests and they shall be carried out in a sequence of high-temperature test, damp heat test, and endurance test. 4. Function test Under the regulated working mode, installation, connection, and load conditions, inspect the flexibility and stability within the full-stroke movement range. During the operation, there shall be no loss of control and obvious jam, the mechanical output and electrical output functions shall meet the requirements, and the products with return or stop function shall meet requirements. 5. Operating stroke Measure the working stroke of the operating end (at the grip point of the handle) under regulated working mode, installation, connection, and load conditions. 6. Operating force Under the regulated working mode, installation, connection, and load conditions, measure the starting force of the operating end or the operating position and complete the required operating force for a specified action. 7. Polarity Under the regulated working mode, installation, connection, and load conditions, measure the movement direction of the operating end, the movement direction of the mechanical output end, and the relationship between positive and negative electrical outputs.

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8. Gear For operational units with gear requirements, measure the stroke, mechanical output, and electrical output of the operating end under the regulated working mode, installation, connection, and load conditions when it is under different gear conditions. 9. Return performance For operational units with return requirements, put the operating end at the limit position or the position specified in the product specifications under the regulated working mode, installation, connection, and load conditions, and then slowly return and stop the operating end and measure the return position of the operating end. Then, operate in the reverse direction and measure, and the return position difference is the return performance. 10. Operating force
displacement curve For operational units (such as side stick control device) with relatively accurate operating force
displacement characteristic curves, make the operating end have full-stroke movement under the regulated working mode, installation, connection, and load conditions. The operating speed shall not be greater than 5 mm/s or it shall be subject to product specification. The operating force and displacement shall be recorded in real time and the operating force
displacement curve shall be drawn. 11. Linearity Under the regulated working mode, installation, connection, and load conditions and within the full-stroke (slightly less than the limit stroke) range, measure the corresponding relationship between the operating end input displacement (or input force) and electrical output (or mechanical output) and it shall not be less than 10 points within the whole measuring range. The ratio of the maximum difference between the actual output and theoretical output of the operational unit to the output when the maximum input instruction is given is the linearity. 12. Hysteresis Under the regulated working mode, installation, connection, and load conditions, have full-stroke cycling of the operating end, the cycling frequency is 0.001
0.05 Hz, and the waveform is a triangular wave or sine wave. Record the output
input curve and it shall show a complete cycle. Then, find the maximum difference of the output positions corresponding to the same input signal on the curve, which is the hysteresis. 13. Frequency characteristic Under the regulated working mode, installation, connection, and load conditions, apply the operating force of specified frequency and amplitude to the operating end and record the amplitude
frequency characteristic and phase
frequency characteristic curves of the output relative operating force.

Unit test of the flight control system 57 14. Step characteristic Under the regulated working mode, installation, connection, and load conditions, apply the step operating force of the specified amplitude to the operating end and record the time course curve. 2.2.4.3 Judging criteria and results handling 1. Resistance It shall comply with the design requirements specified in the product specification. The formula is: R20 5

254 URt 234 1 t

where R20 is the resistance under 20 C, with Ω as the unit; t is the ambient temperature when the measurement is carried out, with  C as the unit; and Rt is the measured resistance when the ambient temperature is t, with Ω as the unit. 2. Insulation resistance It is generally required that the insulation resistance between electric circuits and between the electric circuit and the main case should not be less than 50 MΩ under room temperature, not be less than 20 MΩ after salt spray, mold, endurance test, and under the condition of high temperature and high altitude, and not be less than 10 MΩ under damp heat condition. It can also be estimated according to the product specification. 3. Dielectric strength It is generally required that the dielectric strength between electric circuits and between each circuit and the main case shall not have breakdown when it lasts for 1 minute under the power condition regulated in Table 2.8. It can also be estimated according to the product specification. 4. Function test It is generally required that the starting and conversion under different working modes and the movement within the full-stroke range shall be flexible and smooth without the loss of control or obvious jam, and the gear position shall be clear and accurate. Table 2.8: Condition of dielectric strength test. Test conditions Voltage (effective value) (V) Supply power (minimum) (kV A) Power frequency (Hz)

High temperature

Damp heat

After durability test

500 0.5 50

375

250

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Chapter 2

5. Operating stroke, operating force, polarity, gear They shall comply with the requirements of the product specification. 6. Return performance It shall not be greater than 2 mm generally, or comply with the requirements of the product specification. 7. Operating force
displacement curve It shall comply with the requirements of the product specification. 8. Linearity It is lower than 2% or 3% generally, or subject to the requirements of the product specification. 9. Hysteresis It is lower than 0.2% generally, or subject to the requirements of the product specification. 10. Frequency characteristic, step characteristic It shall comply with the requirements of the product specification.

2.2.5 Mechanical drive unit The mechanical drive unit is not directly connected with the driven object mechanically, but exerts force on the driven object through many drive links, so as to realize the actuation function. The typical mechanical drive unit is the centralized coaxial flaps and slats power drive unit (PDU) widely used for large aircraft. By the types of energy used, the flaps and slats PDU can be divided into dual-hydraulic drive type, dual-electronic drive type, and hydraulic-electronic drive type (hydraulic takes the main part and electronic plays the supplementary role). The dual-electronic drive type flaps and slats PDU is taken as an example below to introduce the functional and performance testing of a mechanical drive unit. Test items of other mechanical drive units are similar. 2.2.5.1 Test objectives and test requirements It aims to test the basic functions and performance of the flaps and slats PDU to ensure that it meets the design requirements and works accurately within the validation scope. The test shall be carried out in accordance with the special technical conditions of the product clause by clause and each test result shall conform to the criteria. 2.2.5.2 Test items and test methods Test items of the flaps and slats PDU usually include resistance, insulation resistance, dielectric strength, function, maximum output torque, maximum output angle, maximum

Unit test of the flight control system 59 output speed, protective threshold, emergency braking characteristics, efficiency, polarity, conversion performance, stability, clearance, and radial runout of output shaft. Pretest preparation: install the flaps and slats PDU on the test bed; the installation and commissioning of the sensor and loading equipment shall meet requirements. The air pressure, temperature, and humidity in the test environment shall meet the requirements of the product specification. The installation, connection, mechanical load, electrical load, and energy of the flaps and slats PDU shall be consistent with the actual installed condition or follow the requirements of the product specification. 1. Resistance measurement Conduct resistance or on
off measurement for each coil and electrical components of the flaps and slats PDU and convert the resistance to that under 20 C. 2. Insulation resistance Use a 500 V tramegger (250 V tramegger for high-altitude test) with precision level no lower than level 2.5 to measure the insulation resistance under relative humidity of 45%
75%. 3. Dielectric strength Increase the voltage between circuits and between the circuit and case to the specified voltage, keep for 1 minute and ensure that there is no breakdown. During the test, uniformly increase the voltage to the specified value and the increased speed should ensure the voltage reading can be read. One flaps and slats PDU can only have three dielectric strength tests and they shall be carried out in the sequence of hightemperature test, damp heat test, and endurance test. 4. Function test Inspect the starting, stop, and conversion, as well as the movement flexibility and stability of the flaps and slats PDU within the full-stroke range under different working modes. During the operation, there shall be no loss of control and obvious jam. Check if the flaps and slats PDU transits steadily and smoothly from one working mode to another. 5. Maximum output angle Measure the maximum output angle of the output end of the flaps and slats PDU under regulated load and working mode. 6. Maximum output torque Make the flaps and slats PDU under regulated working mode, apply a signal, that is, enough to generate the maximum output force and use a force measuring device to measure the output force when the output speed is 0. The maximum output force in two directions shall be measured. 7. Maximum output speed Under the regulated working mode, installation, connection, and load conditions, make the flaps and slats PDU reach the maximum working speed and use a speed

60

8.

9.

10.

11.

12.

13.

14.

Chapter 2 measuring device to measure the maximum speed of the flaps and slats PDU. The maximum output speed in two directions shall be measured. Protective threshold Protective threshold generally includes loading moment protection and speed protection. Supply rated voltage to the flaps and slats PDU under the regulated working mode, and then gradually increase the loading moment or speed at the mechanical output end until the protection function is enabled. Then, measure the loading moment or speed, that is, the protective threshold. Emergency braking characteristics Place the flaps and slats PDU under the regulated working mode, installation, connection, and load conditions and then operate it at the maximum working speed to enable the emergency braking function until it stops completely. Record the time course of output displacement or output speed of the flaps and slats PDU in this process. Efficiency Make the flaps and slats PDU operate under the regulated working mode, installation, connection, load conditions, and speed and then measure the output torque N, output speed n, voltage U at power end, and current I at power end. Then, as per the formula η 5 NUn=UUI, calculate the efficiency η of the flaps and slats PDU. Polarity Measure the static direction relation between the output and input of the flaps and slats PDU, which usually includes the relation between the direction of directive current and the rotation direction of the mechanical output end and the relation between the positive and negative output voltage of the position sensor of the flaps and slats PDU and the movement direction of the mechanical output end. The measurement shall be carried out according to the polarity requirements specified in the product specification. Conversion performance Under the regulated working mode, installation, connection, and load conditions and initial state, make the flaps and slats PDU convert to regulated working mode and measure the conversion threshold value and the transient state of mechanical output. Conversion performance includes the conversion performance when the flaps and slats PDU has faults and the conversion performance between normal working modes. Stability Simulate the installation, connection, and inertial load conditions of the flaps and slats PDU on the aircraft. The input signal and working mode shall conform to the requirements of the product specifications and the product shall not have harmful oscillation or instability. Clearance For the flaps and slats PDU, the radial clearance and axial clearance of the output shaft should be measured. Fix the motor housing and make the measuring head of the micrometer

Unit test of the flight control system 61 approach to the bearing position. Then, apply a force vertical to the spindle at a position within the cross section of the shaft extension according to the product specifications in one direction and then in the reverse direction. The difference between the two readings shown by the micrometer is the radial clearance. Fix the motor housing and put the micrometer at the top of the shaft extension. Then, apply a regulated force along the spindle according to the product specifications in one direction and then in the reverse direction. The difference between the two readings shown by the micrometer is the axial clearance. 15. Radial runout of output shaft For the flaps and slats PDU, the radial runout of the output shaft should be measured. Fix the motor housing and put the measuring head of the micrometer on the spindle and approach to the shaft extension end as far as possible, and then rotate the spindle slowly. The difference between the maximum and minimum readings shown by the micrometer is the radial runout of the output shaft. 2.2.5.3 Judging criteria and results handling 1. Resistance measurement It shall comply with the requirements of the product specification. The formula is: R20 5

254 URt 234 1 t

where R20 is the resistance under 20 C, with Ω as the unit; t is the ambient temperature when the measurement is carried out, with  C as the unit; and Rt is the measured resistance when the ambient temperature is t, with Ω as the unit. 2. Insulation resistance It is generally required that the insulation resistance between electric circuits and between the electric circuit and the main case of the electromechanical actuator should not be less than 50 MΩ under room temperature, not be less than 20 MΩ after salt spray, mold, and endurance test, and under the condition of high temperature and high altitude, and not be less than 10 MΩ under damp heat condition. It can also be estimated according to the product specification. 3. Dielectric strength It is generally required that the dielectric strength between electric circuits and between each circuit and the main case of the actuator shall not have breakdown when it lasts for 1 minute under the power condition regulated in Table 2.9. It can also be estimated according to the product specification. Table 2.9: Condition of dielectric strength test. Test conditions Voltage (effective value) (V) Supply power (minimum) (kV A) Power frequency (Hz)

High temperature

Damp heat

After durability test

500 0.5 50

375

250

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Chapter 2

4. Function test It is generally required that the starting and conversion as well as the movement of the flaps and slats PDU within the full-stroke range should be flexible and stable under different working modes, there shall be no loss of control and obvious jam during operation, and the flaps and slats PDU should transit steadily and smoothly from one working mode to another. 5. Maximum output torque, maximum output angle, maximum output speed, protective threshold, emergency braking characteristics They shall comply with the requirements of the product specification. 6. Efficiency It is greater than 80% generally, or subject to the requirements of the product specification. 7. Polarity, conversion performance They shall comply with the requirements of the product specification. 8. Stability There shall be no harmful oscillation or instability. 9. Clearance, radial runout of output shaft They shall comply with the requirements of the product specification.

2.2.6 Mechanical transmission unit The mechanical transmission unit is used to transmit mechanical displacement and force (or torque) and achieve specific mechanical functions, such as direction change, rotary motion to linear motion, deceleration, force sensing, and guiding. The mechanical transmission unit does not consume energy but transmits mechanical energy. Typical units include pull rod, pulley, steel cable, guide parts, spring load mechanism, ball screw actuator, variable-angle reducer, rotary actuator, and cable tension compensator. The spring load mechanism, cable tension compensator, and ball screw actuator are taken as examples below to introduce the functional/performance testing of the mechanical transmission unit. Test items of other mechanical transmission units are similar. 2.2.6.1 Spring load mechanism 1. Test objectives and test requirements It aims to test the basic functions and performance of the spring load mechanism, to ensure it meets design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to the criteria. 2. Test items and test methods The main test item of the spring load mechanism is force
displacement curve test.

Unit test of the flight control system 63 3. Pretest preparation Install the spring load mechanism on a special test bed, the installation and commissioning of the sensor and loading equipment shall meet requirements. 4. Force
displacement curve test a. Control the loading equipment to make the spring load mechanism operate slowly in the full stroke. In order to eliminate the influence of inertia force and damping force on the test data, the motion speed is usually # 1 mm/s and the motion acceleration is # 1 mm/s2. b. In the full-stroke motion of the spring load mechanism, record the force and displacement of the spring load mechanism at each sampling point. c. Judging criteria and results handling. Draw the force
displacement curve of the spring load mechanism according to recorded force and displacement at each sampling point, and then compare it with the force
displacement envelope curve. If the curve is within the envelope curve, the test item will be deemed acceptable; otherwise it is deemed unacceptable. A typical force
displacement envelope curve of the spring load mechanism is shown in Fig. 2.1. 2.2.6.2 Cable tension compensator 1. Test objectives and test requirements It aims to test the basic functions and performance of the cable tension compensator, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to criteria.

Figure 2.1 Typical force
displacement envelope curve of spring load mechanism.

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2. Test items and test methods Test items of the cable tension compensator mainly include tension test, transmission error test, and transmission friction test. Pretest preparation: install the cable tension compensator on a special test bed and adjust the initial tension of the cable tension compensator according to the product specification. The installation and commissioning of the sensor and loading equipment shall meet the requirements. 1. Tension test Stabilize the ambient temperature to a regulated temperature and change the relative distance between the cable tension compensator and the loading end to permissible maximum and minimum values. After they become stable, use a tensiometer to measure the tension of steel cable connected with the cable tension compensator. 2. Transmission error test a. Stabilize the ambient temperature to a regulated temperature; b. Apply regulated load on loading end and control the input end to have full-stroke reciprocating motion at a regulated speed. In order to eliminate the influence of inertia force and damping force on the test data, the motion speed is usually set as # 1 mm/s and the motion acceleration is set as # 1 mm/s2. c. In the full-stroke reciprocating motion of the input end, record the input end displacement X1 and loading end output displacement X2 of the cable tension compensator at each sampling point. 3. Transmission friction test a. Stabilize the ambient temperature to a regulated temperature; b. Apply regulated load on loading end and control the input end to have full-stroke reciprocating motion at a regulated speed. In order to eliminate the influence of inertia force and damping force on the test data, the motion speed is usually set as # 1 mm/s and the motion acceleration is set as # 1 mm/s2. c. In the full-stroke reciprocating motion of the input end, record the input end transmission force F1 and loading end loading force F2 of the cable tension compensator at each sampling point. 2.2.6.3 Judging criteria and results handling 1. Tension test If the measured tension is within the regulated range, it is deemed acceptable. If the measured tension is not within the regulated range, it is deemed unacceptable. 2. Transmission error test According to the ideal transmission ratio n of the cable tension compensator and measured input end displacement X1 and loading end output displacement X2 of the cable tension compensator, calculate the transmission error Δ 5 jX1 n 2 X2 j. If the maximum transmission error Δ is within the regulated scope, the test item is deemed

Unit test of the flight control system 65 acceptable. If the maximum transmission error Δ is not within the regulated scope, the test item is deemed unacceptable. 3. Transmission friction test According to the ideal transmission ratio n of the cable tension compensator and measured input end transmission force F1 and loading end loading force F2 of the cable tension compensator, calculate the friction force f 5 jF1 2 F2 nj. If the maximum friction force f is within the regulated scope, the test item is deemed acceptable. If the maximum friction f is not within the regulated scope, the test item is deemed unacceptable. 2.2.6.4 Ball screw actuator 1. Test objectives and test requirements It aims to fully test the function and performance of the ball screw actuator, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to the criteria. 2. Test items and test methods The main items of the functional and performance test of the ball screw actuator are transmission accuracy and transmission efficiency. Pretest preparation: Install the ball screw actuator on a special test bed, the installation and commissioning of the sensor and loading equipment shall meet requirements. 1. Transmission accuracy test a. Stabilize the ambient temperature to a regulated temperature, fix the lead screw in the axial direction, fix the nut of the ball screw actuator in the spindle direction and apply the regulated axial load to the nut of the ball screw actuator. The load direction is opposite to the movement direction of the nut. b. Control the lead screw to rotate at a regulated speed, make the nut have full-stroke to-and-fro motion and record the lead screw’s rotation angle θ and the nut’s motion displacement X at each sampling point. 2. Transmission efficiency test a. Stabilize the ambient temperature to a regulated temperature, fix the lead screw in the axial direction, fix the nut of ball screw actuator in the spindle direction, and apply regulated axial load to the nut of the ball screw actuator. The load direction is opposite to the movement direction of the nut. b. Control the lead screw to rotate at a regulated speed, make the nut have full-stroke to-and-fro motion and record the lead screw’s rotational angular speed ω and transmission torque T and the nut’s motion speed V and loading force F at each sampling point.

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2.2.6.5 Judging criteria and results handling 1. Transmission accuracy test Firstly, calculate the target value Xtarget of nut displacement as per the formula Xtarget 5 ðθ=360ÞS on the basis of the measured rotational angle θ at each sampling point. In the formula, S is the lead of the lead screw. Then, calculate
the transmission
accuracy ΔX at each sampling moment as per the formula ΔX 5
Xtarget 2 X
. If the maximum transmission accuracy ΔX is less than or equal to the required value, the test item is deemed acceptable. If the maximum transmission accuracy ΔX is greater than the required value, the test item is deemed unacceptable. 2. Transmission efficiency test Calculate the transmission efficiency at each sampling moment as per the formula η 5 jFV j=Tω on the basis of the lead screw’s rotational angular speed ω and transmission torque T and the nut’s motion speed V and loading force F at each sampling point. If the minimum transmission efficiency η is greater than or equal to the required value, the test item is deemed acceptable. If the minimum transmission efficiency η is lower than the required value, the test item is deemed unacceptable.

2.2.7 Mechanical actuating unit The mechanical actuating unit is directly connected with the actuated object mechanically and exerts acting force on the actuated object. Typical mechanical actuating units include the back actuator, aileron trim mechanism, rudder trim mechanism, elevator trim mechanism, aileron actuator, rudder actuator, elevator actuator, horizontal stabilizer actuator, spoiler actuator, air brake actuator, wingtip braking device, and slats rotary actuator. By the types of power and energy used, the mechanical actuating unit can be divided into hydraulic actuator, electromechanical actuator, and electrohydraulic actuator. The contents of the functional and performance testing of the hydraulic actuator and electromechanical actuator are described below. 2.2.7.1 Hydraulic actuator 1. Test objectives and test requirements It aims to test the basic functions and performance of the, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to the criteria. 2. Test items and test methods Test items of the hydraulic actuator generally include resistance, insulation resistance, dialectic strength, function, working stroke, external sealing, internal oil

Unit test of the flight control system 67

3.

4.

5.

6.

7.

8.

leakage, oil spillover, insensitive and zero-bias current, maximum output force, maximum speed and speed difference, zero drift, polarity, return performance, conversion performance, linearity, hysteresis, output synchronization, frequency characteristic, step characteristic, stability, impedance characteristic, output consistency, and interference level. Pretest preparation: the hydraulic actuator is installed on a special test bed, the installation and commissioning of the sensor and loading equipment meet requirements and the commissioning of the complete system is finished. The air pressure, temperature, and humidity in the test environment shall meet the requirements of the product specification. The installation, connection, mechanical load, electrical load, and energy of the hydraulic actuator shall be consistent with the actual installed condition or follow the requirements of the product specification. Resistance measurement Conduct resistance or on
off measurement for each coil and electrical components of the actuator and convert the resistance to that under 20 C. Insulation resistance Use a 500 V tramegger (250 V tramegger for high-altitude test) with precision level no lower than level 2.5 to measure the insulation resistance under relative humidity of 45%
75%. During the test, it is not necessary to supply oil to the actuator, but the wet coil should be immersed in the working fluid. Dialectic strength Increase the voltage between circuits and between the circuit and case to the specified voltage, keep for 1 minute and ensure there is no breakdown. During the test, uniformly increase the voltage to the specified value and the increase speed should ensure the voltage reading can be read. One hydraulic actuator can only have three dielectric strength tests and they shall be carried out in the sequence of hightemperature test, damp heat test, and endurance test. Function test Inspect the starting and conversion and also the movement flexibility and stability of the hydraulic actuator within the full-stroke range under different working modes. During the operation, there shall be no loss of control and obvious jam. Check if the hydraulic actuator transits steadily and smoothly from one working mode to another. Working stroke Measure the working stroke of the mechanical output end of the hydraulic actuator under regulated load and working mode. External sealing a. Sealing performance under low pressure: apply specified low pressure to each oil inlet connector of the hydraulic actuator and maintain for a specific time period, and then check the external leakage condition.

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b. Sealing performance under oil inlet pressure: apply specified high pressure to each oil inlet of hydraulic actuator and maintain for a specific time period. Pressure shall be posed at each limit position of the element of distributing slide valve and electrohydraulic servo valve and the piston of actuator cylinder to check external leakage condition. c. Sealing performance under oil return: pose rated oil inlet pressure at the oil inlet port and oil return port of the hydraulic actuator, maintain for a specific time period and then check the external leakage condition. d. Sealing performance in operation: after the hydraulic actuator has full-stroke reciprocating operation for no less than 100 cycles under regulated working mode, installation, connection, and load conditions, check the external leakage condition. e. Internal leakage amount: pose rated oil inlet pressure at oil inlet and measure the oil leakage amount at the oil inlet. The internal leakage amount of actuator cylinder, slide valve, and electrohydraulic servo valve shall be measured at different positions and when the solenoid valve is energized or not energized and it shall be measured for 3 minutes after the state maintains for 2 minutes. f. Oil spillover amount: For a hydraulic actuator with two hydraulic systems (A and B), it shall operate under dual system mode for a certain period of time before the test to make the cavity inside the hydraulic actuator be filled with working fluid. And then, rated oil inlet pressure shall be posed at the oil inlet of system A (or B) and the oil return pressure specified in product specifications shall be posed at the oil return port. Measure the outlet flow at the oil inlet and the return pipe nozzle of system B (or A). For a hydraulic actuator with a triple hydraulic system, the measurement method can be regulated in product specifications according to the principles above. 9. Insensitive and zero-bias current Measure the control current in two directions of the electrohydraulic servo valve when the actuator cylinder generates continuous output motion with a speed of 0.1 mm/s in two directions. The absolute value of the algebraic difference of the two control currents is insensitive current and the algebraic sum of the two control currents is zero-bias current. 10. Maximum output force Supply rated oil inlet and oil return pressure to the hydraulic actuator, make the hydraulic actuator under regulated working mode, and pose the signal that can generate the maximum output force and use a force measuring device to measure the output force when the output speed is 0. The measurement shall be conducted in two directions.

Unit test of the flight control system 69 11. Maximum speed and speed difference Supply rated oil inlet and oil return pressure to the hydraulic actuator and make the hydraulic actuator receive the maximum input signal when it is under regulated working mode and no load. Measure the maximum output speed of the hydraulic actuator in two directions and the maximum difference between the speeds in the two directions is the maximum speed difference. 12. Zero drift Due to the change of service conditions, the required input change to keep the original output unchanged is called zero drift, which is expressed by the percentage of the change of zero-bias current to rated current. During the measurement, the hydraulic actuator is under regulated working mode, installation, connection, and load conditions. a. Zero drift of oil inlet pressure: when the oil inlet pressure is within 85%
102% of the rated oil inlet pressure, the zero drift is not more than 2% or the required value of the product specification. b. Zero drift of oil return pressure: when the oil return pressure is within 0%
20% of the rated oil return pressure, the zero drift is not more than 4.5% or the required value of the product specification. c. Zero drift of temperature: when the temperature is 230 C to 135 C, the zero drift should not be greater than 6 1% or should be the required value of the product specifications per change of 28 C. d. Zero drift of acceleration: it shall be measured when the acceleration is within the range of change specified in the product specification. 13. Polarity Measure the static direction relation between the output and input of the hydraulic actuator, which usually includes the relation between the servo valve current direction and the movement direction of the mechanical output end of the hydraulic actuator and the relation between the positive and negative output voltage of the position sensor of the hydraulic actuator and the movement direction of the mechanical output end of the actuator. The measurement shall be carried out according to the polarity requirements specified in the product specification. 14. Return performance For a hydraulic actuator with a mechanical return function, under regulated working mode, installation, connection, and load conditions, make it have mechanical return, and then measure the return time, return speed, and return accuracy. 15. Conversion performance Under regulated working mode, installation, connection, and load conditions and initial state, make the actuator convert to regulated working mode and measure the conversion threshold value and the transient state of the mechanical

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19.

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Chapter 2 output. Conversion performance includes the conversion performance when the hydraulic actuator has faults and the conversion performance among normal working modes. Linearity Place the hydraulic actuator under regulated working mode, installation, connection, and load conditions and use a point-by-point method to measure the corresponding relationship between the input signal and the output position within the range of full stroke (slightly lower than the limit range). There should be no less than 10 points in the whole range. The ratio of the maximum difference between the actual output displacement and theoretical output displacement of the hydraulic actuator to the output displacement under the instruction of maximum input is the linearity. Hysteresis Place the hydraulic actuator under regulated working mode, installation, connection, and load conditions and make its output have a full-stroke reciprocating cycle, with cycle frequency of 0.001
0.05 Hz and waveform in a triangular wave or sine wave. Record the output
input curve and it shall show a complete cycle. Then, find the maximum difference of the output position corresponding to the same input signal on the curve and it is the hysteresis. Output consistency Place the hydraulic actuator under regulated working mode, installation, connection, and load conditions and make it input the maximum instruction signal, with frequency as 0.05 Hz generally, and then record the input
output characteristic curve when each channel works independently. The ratio of the maximum difference of the output positions of channels at the same input value to the output displacement under maximum input instruction is the output consistency. Frequency characteristic Place the hydraulic actuator under regulated working mode, installation, connection, and load conditions, input regulated frequency signal to the hydraulic actuator and record the amplitude
frequency characteristic and phase
frequency characteristic curve (Bode diagram) of the output to input. Step characteristic Make the hydraulic actuator under regulated working mode, installation, connection, and load conditions, input the regulated step signal to the hydraulic actuator and then record the time course curve of the output of the hydraulic actuator. Stability Simulate the installation, connection, and inertial load conditions of the hydraulic actuator-control surface system on the aircraft. The input signal and working mode shall conform to the requirements of the product specifications and the hydraulic actuator shall not have harmful oscillation or instability.

Unit test of the flight control system 71 22. Impedance characteristic Place the hydraulic actuator under regulated working mode, installation, connection, and inertia load conditions, apply a regulated external sine load to the output end of the hydraulic actuator and then record the amplitude
frequency characteristic and phase
frequency characteristic of the output to the external load. 23. Interference level For an actuation system that has two or more than two hydraulic actuators driving the same control surface, the output force consistency of two or more than two hydraulic actuators should be assessed. Place two or more than two hydraulic actuators under regulated working mode, installation, connection, inertia load, and external load conditions, input the regulated amplitude and frequency instructions to the hydraulic actuator and then measure the output force of the two hydraulic actuators. 2.2.7.2 Judging criteria and results handling 1. Resistance measurement It shall comply with the requirements of the product specification. The formula is: R20 5

254 URt 234 1 t

wherein R20 is the resistance under 20 C, with Ω as the unit; t is the ambient temperature when the measurement is carried out, with  C as the unit; and Rt is the measured resistance when the ambient temperature is t, with Ω as the unit. 2. Insulation resistance It is generally required that the insulation resistance between electric circuits and between the electric circuit and main case should not be less than 20 MΩ under room temperature, not be less than 2 MΩ after salt spray, mold, and endurance test and under the condition of high temperature and high altitude, and not be less than 1 MΩ under damp heat condition. 3. Dialectic strength It is generally required that the dielectric strength between electric circuits and between each circuit and the main case shall not break down when it lasts for 1 minute under the power condition regulated in Table 2.10. Table 2.10: Condition of dielectric strength test. Test conditions Voltage (effective value) (V) Supply power (minimum) (kV A) Power frequency (Hz)

High temperature

Damp heat

After durability test

500 0.5 50

375

250

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4. Function test It is generally required that the starting, stop, and conversion, as well as the movement of the hydraulic actuator within the full-stroke range should be flexible and stable under different working modes, there should be no loss of control and obvious jam during operation, and the hydraulic actuator should transit steadily and smoothly from one working mode to another. 5. Working stroke It should comply with the requirements of the product specification. 6. External sealing a. Sealing performance under low voltage: it is generally required to apply a pressure that can maintain a 2-meter-high liquid column at the connector of each oil inlet pipe of the hydraulic actuator for 2 hours and there shall be no visible external leakage. If there are disputes on the conclusion made with the test method, extend the maintaining time to 12 hours and there shall be no obvious external leakage (the sealing part may be wet but there are no water drops). b. Sealing performance under oil inlet pressure: it is generally required to apply 1.5 times the rated oil inlet pressure to each oil inlet of the hydraulic actuator and pose the pressure at each limit position of the element of distributing slide valve and electrohydraulic servo valve and the piston of actuator cylinder. The pressure shall be posed for 3 minutes each time and shall be no obvious external leakage (the sealing part may be wet but there are no water drops). c. Sealing performance under oil return: it is generally required to pose the rated oil inlet pressure at the oil inlet port and the oil return port of the hydraulic actuator and maintain for 3 minutes. There shall no obvious external leakage (the sealing part may be wet but there are no water drops). d. Sealing performance in operation: under regulated load and working mode, the hydraulic actuator is required to have a full-stroke reciprocating operation for no less than 100 cycles and there shall be no obvious external leakage (the sealing part may be wet but there are no water drops). During the high-temperature test, low-temperature test, endurance test, and life test, the external leakage of the hydraulic actuator shall not exceed 2 drops per minute or 4 mL/h in operation. e. Internal leakage amount: the internal leakage amount in the qualification test, periodic test, and service period shall not be greater than two times of that under acceptance conditions generally. 7. Oil spillover amount It shall comply with the requirements of the product specification. 8. Insensitive and zero-bias current The insensitive current is not greater than 0.1% generally, or subject to the requirements of the product specification.

Unit test of the flight control system 73 9. Maximum output force It shall comply with the requirements of the product specification. 10. Maximum speed and speed difference They shall comply with the requirements of the product specifications and the speed difference is generally 6 15% of the maximum speed. 11. Zero drift a. Zero drift of oil inlet pressure: not more than 2% or the required value of the product specification. b. Zero drift of oil return pressure: not more than 4.5% or the required value of the product specification. c. Zero drift of temperature: when the temperature is 230 C to 135 C, the zero drift should not be greater than 6 1% or should be the required value of the product specifications per change of 28 C. d. Zero drift of acceleration: It shall be measured when the acceleration is within the range of change specified in the product specification. 12. Polarity, return performance, conversion performance They shall comply with the requirements of the product specification. 13. Linearity It is lower than 2% or 3% generally, or subject to the requirements of the product specification. 14. Hysteresis It is lower than 0.2% generally, or subject to the requirements of the product specification. 15. Output consistency, frequency characteristic, step frequency They shall comply with the requirements of the product specification. 16. Stability The system shall have no harmful oscillation or instability. 17. Impedance characteristic It is generally required that the dynamic stiffness should be greater than the static stiffness. When approaching to the natural frequency, the phase
frequency response curve positively penetrates the frequency axis (from negative value to positive value) and remains positive within the range greater than the natural frequency of the system. 18. Interference level Compare the output force difference of two or more than two hydraulic actuators and it should comply with the requirements of the product specification. 2.2.7.3 Electromechanical actuator 1. Test objectives and test requirements It aims to test the basic functions and performance of electromechanical actuator, to ensure it meets the design requirements and works accurately within the validation scope.

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The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to the criteria. 2. Test items and test methods Test items of the electromechanical actuator generally include resistance, insulation resistance, dialectic strength, function, maximum output force (torque), maximum output displacement (angle), protective threshold, emergency braking characteristic, step response characteristic, efficiency, zero drift, polarity, conversion performance, linearity, hysteresis, frequency characteristic, stability, impedance characteristic, clearance, and radial runout of output shaft. Pretest preparation: install the electromechanical actuator on a special test bed, the installation and commissioning of the sensor and loading equipment shall meet requirements. The air pressure, temperature, and humidity in the test environment shall meet the requirements of the product specification. The installation, connection, mechanical load, electrical load, and energy of the electromechanical actuator shall be consistent with the actual installed condition or follow the requirements of the product specification. 1. Resistance measurement Conduct resistance or on
off measurement for each coil and electrical components of the electromechanical actuator and convert the resistance to that under 20 C. 2. Insulation resistance Use a 500 V tramegger (250 V tramegger for high-altitude test) with precision level no lower than level 2.5 to measure the insulation resistance under relative humidity of 45%
75%. 3. Dialectic strength Increase the voltage between circuits and between the circuit and case to specified voltage, keep for 1 minute and ensure there is no breakdown. During the test, uniformly increase the voltage to the specified value and the increased speed should ensure the voltage reading can be read. One actuator can only undergo three dielectric strength tests and they shall be carried out in the sequence of high-temperature test, damp heat test, and endurance test. 4. Function test Inspect the starting, stop, and conversion and also the movement flexibility and stability of the electromechanical actuator within the full-stroke (speed) range under different working modes. During the operation, there should be no loss of control and obvious jam. Check if the electromechanical actuator transits steadily and smoothly from one working mode to another. 5. Maximum output displacement (or angle) Measure the maximum output displacement (or rotation angle) of the output end of the electromechanical actuator under the regulated load and working mode.

Unit test of the flight control system 75 6. Maximum output force (or torque) Place the electromechanical actuator under regulated working mode, apply a signal, that is, enough to generate maximum output force (torque) and use a force measuring device to measure the output force when the output speed is 0. The maximum output force in two directions should be measured. 7. Maximum output speed Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions, adjust it to the maximum working speed and measure the maximum speed of the electromechanical actuator. The maximum output speed should be measured in two directions. 8. Protective threshold Protective threshold generally includes loading moment protection and speed protection. Supply rated voltage to the electromechanical actuator, place the unit under regulated working mode, and then gradually increase the loading moment (torque) or speed at the mechanical output end. 9. Emergency braking characteristics Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions and then operate it at the maximum working speed to enable the emergency braking function until it stops completely. Record the time course of output displacement or output speed of the electromechanical actuator in this process. 10. Step response characteristic Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions and input regulated step command (it is the position command for the position electromechanical actuator and it is the speed command for the speed electromechanical actuator) to the electromechanical actuator. Record the time course of output displacement or output speed of the electromechanical actuator in this process. 11. Efficiency Make the electromechanical actuator operate under regulated working mode, installation, connection, load conditions, and speed and then measure the output force F, output speed V, voltage U at power end, and current I at power end of the electromechanical actuator. Then, as per the formula η 5 ðFV=UIÞ, calculate the efficiency η of the electromechanical actuator. 12. Zero drift Due to the change of service conditions, the required input change to keep the original output unchanged is called zero drift, which is expressed by the percentage of the change of zero-bias current to rated current. During the measurement, the electromechanical actuator is under regulated working mode, installation, connection, and load conditions. a. Zero drift of power voltage: when the power voltage changes within the specified range, keep the output of the electromechanical actuator unchanged and measure the change of the input instruction of the electromechanical actuator.

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Chapter 2 b. Zero drift of temperature: when the ambient and product temperatures change within the specified range, keep the output of the electromechanical actuator unchanged and measure the change of the input instruction of the electromechanical actuator. c. Zero drift of acceleration: when the acceleration changes within the specified range, keep the output of the electromechanical actuator unchanged and measure the change of the input instruction of the electromechanical actuator. Polarity Measure the static direction relation between the output and input of the electromechanical actuator, which usually includes the relation between the direction of directive current and the movement direction of the mechanical output end of the electromechanical actuator and the relation between the positive and negative output voltage of the position sensor of the electromechanical actuator and the movement direction of the mechanical output end. The measurement shall be carried out according to the polarity requirements specified in the product specification. Conversion performance Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions and initial state, convert it to regulated working mode, and measure the conversion threshold value and the transient state of mechanical output. Conversion performance includes the conversion performance when the electromechanical actuator has faults and the conversion performance among normal working modes. Linearity Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions and use a point-by-point method to measure the corresponding relationship between the input instruction and the output (the output is speed for speed electromechanical actuator and the output is displacement for position electromechanical actuator) within the range of full stroke (slightly lower than the limit range). There should be no less than 10 points in the whole range. The ratio of the maximum difference between the actual output and theoretical output of the electromechanical actuator to the output under the instruction of maximum input is the linearity. Hysteresis Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions and make its output have a full-stroke reciprocating cycle, with a cycle frequency of 0.001
0.05 Hz and a waveform in a triangular wave or sine wave. Record the output
input curve and it should show a complete cycle. Then, find the maximum difference of the output position corresponding to the same input signal on the curve and it is the hysteresis. Frequency characteristic Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions, input regulated frequency signal to the

Unit test of the flight control system 77

18.

19.

20.

21.

electromechanical actuator, and record the amplitude
frequency characteristic and phase
frequency characteristic curve of the output to input instruction. Stability Simulate the installation, connection, and inertial load conditions of the electromechanical actuator on the aircraft. The input signal and working mode should conform to requirements of the product specifications and the system should not have harmful oscillation or instability. Impedance characteristic Place the electromechanical actuator under regulated working mode, installation, connection, and load conditions, apply a regulated external sine load to the output end of the electromechanical actuator and then measure the amplitude
frequency characteristic and phase
frequency characteristic of the output to the external load. Clearance For electromechanical actuator having rotary output, the radial clearance and axial clearance of the output shaft should be measured. Fix the motor housing and make the measuring head of micrometer approach to the bearing position. Then, apply a force vertical to the spindle at a position within the end face of the shaft extension according to the product specifications in one direction and then in reverse direction. The difference between the two readings shown by the micrometer is the radial clearance. Fix the motor housing and put the micrometer at top of shaft extension. Then, apply a regulated force along the spindle according to the product specifications in one direction and then in reverse direction. The difference between the two readings shown by the micrometer is the axial clearance. Radial runout of output shaft For electromechanical actuator having rotary output, the radial runout of the output shaft should be measured. Fix the motor housing and put the measuring head of micrometer on the spindle and approach to the shaft extension end as far as possible, and then rotate the spindle slowly. The difference between the maximum and minimum readings shown by the micrometer is the radial runout of the output shaft.

2.2.7.4 Judging criteria and results handling 1. Resistance measurement It shall comply with the requirements of the product specification. The formula is: R20 5

254 URt 234 1 t

where R20 is the resistance under 20 C, with Ω as the unit; t is the ambient temperature when the measurement is carried out, with  C as the unit; and Rt is the measured resistance when the ambient temperature is t, with Ω as the unit.

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2. Insulation resistance It is generally required that the insulation resistance between electric circuits and between the electric circuit and main case of the electromechanical actuator should not be less than 50 MΩ under room temperature, not be less than 20 MΩ after salt spray, mold, and endurance test and under the condition of high temperature and high altitude, and not be less than 10 MΩ under damp heat condition. It can also be estimated according to the product specification. 3. Dialectic strength It is generally required that the dielectric strength between electric circuits and between each circuit and the main case of the electromechanical actuator should not have a breakdown when it lasts for 1 minute under the power condition regulated in Table 2.11. It can also be estimated according to the product specification. 4. Function test It is generally required that the starting and conversion as well as the movement of the electromechanical actuator within the full-stroke range should be flexible and stable under different working modes without loss of control and obvious jam during operation, and the actuator should transit steadily and smoothly from one working mode to another. 5. The maximum output force (torque), maximum output displacement (angle), maximum output speed, protective threshold, emergency braking characteristic, and step response characteristic should comply with the requirements of the product specification. 6. Efficiency. It is greater than 80% generally, or subject to the requirements of the product specification. 7. Zero drift a. Zero drift of power supply voltage: it should not be greater than 2% or should be the required value of the product specification. b. Zero drift of temperature: it should not be greater than 6 1% or should be the required value of the product specifications per change of 28 C. c. Zero drift of acceleration: it shall be measured when the acceleration is within the range of change specified in product specification. It should comply with the requirements of the product specification. 8. Polarity and conversion performance. They should comply with the requirements of the product specification. Table 2.11: Condition of dielectric strength test. Test conditions Voltage (effective value) (V) Supply power (minimum) (kV A) Power frequency (Hz)

High temperature

Damp heat

After durability test

500 0.5 50

375

250

Unit test of the flight control system 79 9. Linearity It is 2%
3% generally, or subject to the requirements of the product specification. 10. Hysteresis It is 0.2% generally, or subject to the requirements of the product specification. 11. Frequency characteristic It should comply with the requirements of the product specification. 12. Stability There should be no harmful oscillation or instability. 13. Impedance characteristic It is generally required that the dynamic stiffness should be greater than the static stiffness. When approaching to the natural frequency, the phase
frequency response curve positively penetrates the frequency axis (from negative value to positive value) and remains positive within the range greater than the natural frequency of the system. 14. Radial runout of clearance output shaft It should comply with the requirements of the product specification.

2.2.8 Sensor unit By functions, the flight control system sensors are mainly classified into pilot command sensor, aircraft motion sensor, control surface motion sensor, atmospheric data sensor and sensor for other purposes (for the realization of other special functions other than the basic three-axis control function). 1. Pilot command sensor It mainly includes pilot instruction displacement sensor and pilot directive power sensor. 2. Aircraft motion sensor It mainly includes angular rate gyroscope assembly, accelerometer, attitude system, and inertial navigation device. 3. Control surface motion sensor It mainly includes control surface position sensor, flaps tilt sensor, and slats tilt sensor. 4. Atmospheric data sensor It mainly includes dynamic pressure, static pressure, total pressure, airspeed, pressure-altitude and altitude difference, and angle of attack/angle of sideslip sensors. 5. Sensor for other purposes It mainly includes overload sensor, throttle lever position sensor, landing gear control sensor, engine speed sensor, and wheel load sensor. 2.2.8.1 Pilot instruction displacement sensor Generally, pilot instruction displacement sensor adopts a differential transformer sensor, such as RVDT and LVDT. Its main function is to detect the pilot control displacement.

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RVDT is taken as an example below to introduce content about the functional and performance test of this kind of sensor. Test items of LVDT are similar. 1. Test objectives and test requirements It aims to test the function and performance of RVDT, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result shall conform to the criteria. Unless otherwise specified, test requirements for RVDT shall be subject to the general requirements in Section 2.2.1. 2. Test items and test methods The main test items of RVDT include appearance structure, power supply voltage, linearity, null voltage, power consumption, impedance, measuring range, gradient, precision, phase shift, sum voltage, and channel consistency. 1. Appearance structure Items for appearance structure inspection include appearance, weight, overall dimensions, installation form, interchangeability, error prevention, and electrical interface and the inspection methods include visual inspection, electronic balance, and ruler. Relevant test data shall be recorded. 2. Power supply voltage According to the requirements of the product specification, adjust the power supply voltage of RVDT to the limit value of the specified deviation 6 1% and test the performance indexes of RVDT. 3. Null voltage Connect RVDT and special tester according to the definition of electrical interface, fix the RVDT on the optical dividing head, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), and test the minimum output voltage of the spindle near the zero mark. 4. Power consumption Connect RVDT and tester according to the definition of electrical interface, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), and measure the input power of the exciting winding of RVDT when there is no load. 5. Impedance a. Input impedance: supply 7 V 1800 Hz AC power to RVDT according to the power supply voltage requirements in article (2) and use a milliammeter to measure its exciting current. b. Output impedance: supply 7 V 1800 Hz AC power to RVDT according to the power supply voltage requirements in article (2) and measure the output voltage under no load and the output voltage and output current under rated load.

Unit test of the flight control system 81 6. Measuring range Connect RVDT and special tester according to the definition of the electrical interface, fix the RVDT on the optical dividing head, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), set the RVDT at the position of maximum working stroke, and use a digital multimeter to measure the output voltage of the RVDT. 7. Gradient Connect RVDT and special tester according to the definition of the electrical interface, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the RVDT operates within the range of maximum working stroke, and calculate the ratio of output voltage to stroke. 8. Linearity Connect RVDT and special tester according to the definition of the electrical interface, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the RVDT operates within the range of maximum working stroke, conduct linear processing with the least square method or endpoint method, and calculate maximum linearity error. 9. Precision Connect RVDT according to the definition of the electrical interface, supply 7 V 1800 Hz AC current to RVDT according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the RVDT operates within the range of maximum working stroke, and compare them with theoretical output voltage. The ratio of the maximum voltage difference to the voltage in the whole measuring range is the maximum precision error. 10. Phase shift Install RVDT on a special test fixture and set it at the position of maximum working stroke, and then measure the phase shift of the output voltage of secondary winding to the voltage of primary winding. 11. Sum voltage Install RVDT on a special test fixture, use an instrument with high precision to measure high-medium and low-medium voltage at the output end of RVDT, and then calculate the sum of the two voltages. 12. Channel consistency Install redundant RVDT on a special test fixture, use a multichannel test instrument to measure the maximum output voltage of each channel and then calculate the maximum output voltage difference of channels.

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2.2.8.2 Judging criteria and results handling 1. If the results of the appearance structure inspection of airborne units are consistent with the expected value, the structure of the airborne unit is deemed acceptable. That is to say, the appearance is intact and not damaged, the connector is not bent or retracted, the weight and overall dimensions meet requirements, the structure is well installed without fall off and the electrical interface is error free. 2. Power voltage It should meet the requirements of the product specification. 3. Null voltage It should meet the requirements of the product specification, or subject to the regulations in Section 3.9.5 of Ref. [1]. 4. Power consumption It should comply with the regulations in Section 3.5.2 of Ref. [1] or requirements of the product specification. 5. Impedance The input impedance is calculated as per the formula below and the result should comply with the regulations in Section 3.9.2 of Ref. [1]. Rz 5

U I

where Rz is the input impedance, with Ω as the unit; U is the exciting voltage, with V as the unit; and I is the exciting current, with A as the unit. The output impedance shall be calculated as per the formula below and the result shall comply with the regulations in Section 3.9.2 of Ref. [1]. Z5

ðE 2 UÞ I

where Z is the output impedance, with Ω as the unit; E is the no-load output voltage, with V as the unit; U is the output voltage under rated load, with V as the unit; and I is the output current under rated load, with A as the unit. 6. Measuring range It should meet the requirements of the product specification. 7. Gradient Calculate the ratio of the output voltage to stroke and it should comply with the regulations in Section 3.9.3 of Ref. [1]. 8. If the results of linearity, precision, phase shift, sum voltage, and channel consistency comply with the regulations in Ref. [1] or requirements of the product specification, the test items are deemed acceptable; otherwise the test items are deemed unacceptable.

Unit test of the flight control system 83 2.2.8.3 Pilot directive power sensor Generally, the pilot directive power sensor adopts a strain pressure sensor. Its main function is to detect the pilot control force. The control wheel/column force sensor is taken as an example below to introduce the functional and performance testing of this kind of sensor. Tests of other force sensors are similar. 1. Test objectives and test requirements It aims to test the function and performance of control wheel/column force sensor, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be carried out in accordance with the special technical conditions of the product clause by clause and each test result shall conform to the criteria. Unless otherwise specified, test requirements for control wheel/column force sensor shall be subject to the general requirements in Section 2.2.1. 2. Test items and test methods Main test items of the control wheel/column force sensor include appearance structure, power voltage, null voltage, measuring range, impedance, gradient, linearity, precision, repeatability, and channel consistency. 3. Appearance structure Items for appearance structure inspection include appearance, weight, overall dimensions, installation form, and electrical interface and the inspection methods include visual inspection, electronic balance, and ruler. 4. Power supply voltage According to the requirements of the product specification, adjust the power supply voltage of control wheel/column force sensor to the limit value of the specified deviation and check the performance of the control wheel/column force sensor, which should meet the requirements of the product specification. 5. Null voltage Connect the control wheel/column force sensor through an electrical connection and supply power to the control wheel/column force sensor according to the power supply voltage requirements in article (2). Test and record the output value when the input voltage is 0. 6. Measuring range Connect the control wheel/column force sensor through an electrical connection and install it on a special test fixture, supply power to the control wheel/column force sensor according to the power supply voltage requirements in article (2), apply the maximum power for the sensor, and test the output voltage. 7. Impedance a. Input impedance: when the output end of the control wheel/column force sensor has an open circuit, use a digital ohmmeter or other relevant instrument to measure the impedance at the input end.

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Chapter 2 b. Output impedance: when the input end of the control wheel/column force sensor has a short circuit, use a digital ohmmeter or other relevant instrument to measure the impedance at the output end. Gradient Apply a rated power supply voltage to the control wheel/column force sensor according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the control wheel/column force sensor operates within the range of maximum working stroke, and calculate the ratio of the output voltage to the stroke. Linearity Connect the control wheel/column force sensor through an electrical connection, apply a rated power supply voltage to the control wheel/column force sensor according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the control wheel/column force sensor operates within the range of maximum working stroke, conduct linear processing with the least square method or endpoint method, and calculate maximum linearity error. Precision Connect the control wheel/column force sensor through an electrical connection and install it on a special test fixture, apply rated power supply voltage to the control wheel/column force sensor according to the power supply voltage requirements in article (2), use an instrument with high precision to measure the output voltage at each measuring point when the control wheel/column force sensor operates within the range of maximum working stroke, and compare them with theoretical output voltage. The ratio of the maximum voltage difference to the voltage in the whole measuring range is the maximum precision error. Repeatability Install the control wheel/column force sensor on a special test fixture, use an instrument with high precision to make the control wheel/column force sensor operate within the range of maximum working stroke, measure the output voltage when the given inputs change continuously and repeatedly in the full measuring range in the same direction (direct or reverse) and calculate the percentage of the ratio of the unrepeated maximum value for maximum difference and the mean value in the full measuring range in the direct and reverse stroke. The result should comply with the requirements of the product specification. Channel consistency Install the control wheel/column force sensor on a special test fixture, use a multichannel test instrument to measure the maximum output voltage of each channel and then calculate the maximum output voltage difference of the channels.

Unit test of the flight control system 85 2.2.8.4 Judging criteria and results handling If the results of the appearance structure inspection of airborne units are consistent with the expected value, the structure of the airborne unit is deemed acceptable. In other words, the appearance is intact and not damaged, the connector is not bent or retracted, the weight and overall dimensions meet requirements, the structure is well installed and the electrical interface is error free. If the results of the power supply voltage, null voltage, measuring range, impedance, gradient, linearity, precision, repeatability, and channel consistency comply with the requirements of the product specification, the test items are deemed acceptable, or the test items are deemed unacceptable. 2.2.8.5 Aircraft motion sensor The main function of the aircraft motion sensor is to detect the motion parameters of the aircraft. The angular rate gyroscope assembly (hereinafter referred to as the gyro assembly) is taken as an example below to introduce the functional and performance testing of this kind of sensor. Test items of other similar sensors are similar. 1. Test objectives and test requirements It aims to test the basic functions and performance of the gyro assembly, to ensure it meets design requirements and works accurately within the validation scope. The test should be carried out in accordance with the special technical conditions of the product clause by clause and each test result should conform to the criteria. Unless otherwise specified, test requirements for the gyro assembly should be subject to the general requirements in Section 2.2.1. 2. Test items and test methods The main test items of the gyro assembly include appearance structure, power voltage, insulation resistance, preparation time, starting current and working current, natural frequency and damping ratio, calibration factor, threshold, resolution, zero-bias, polarity, measuring range, and channel consistency. 1. Appearance structure Items for appearance structure inspection include appearance, weight, overall dimensions, installation form, and electrical interface and the inspection methods include visual inspection, electronic balance, and ruler. 2. Power supply voltage According to special technical conditions of the product, adjust the power supply voltage of gyro assembly to the limit value of specified deviation and check its performance. 3. Insulation resistance Place the gyro assembly in a nonworking state and measure the insulation resistance with a megohmmeter of corresponding voltage according to the special technical conditions of the product.

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4. Preparation time It generally refers to the starting time of the gyro assembly motor. For a gyro assembly that requires temperature control, the time required to reach the required working temperature should also be included. Generally, the motor characteristic method or rate response method is used for inspection. a. Motor characteristic method: measure with a second chronograph with an error of less than 0.1 second or other timers of corresponding precision, start the timer when power is connected to the gyro assembly motor, turn off the timer when the gyro assembly rotor reaches the rated speed (the speed can be measured with a working current monitor, rotor rotation detector or other instruments that can measure the rotor speed), and then read the time shown by the timer. b. Rate response method: install the gyro assembly on a rate turntable, make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane, and then connect the output of the gyro assembly with the recorder of timer of corresponding precision. Then, make the rate turntable rotate at a selected constant rate and start the recorder to record the time from the transient energization moment of the gyro assembly motor to the time when the output of the gyro assembly reaches a stable state or a specified value. 5. Heating time Use a second chronograph with an error of no more than 0.1 second or other timers of corresponding precision to measure the time from the connection to temperature control circuit to the time when the gyro assembly reaches the specified performance requirements. The specific method shall follow the special technical conditions of the product. 6. Starting current and working current According to the requirements of special technical conditions of the product, an ammeter of specified precision or other instruments should be used to measure the current values of the gyro assembly. When the power to the gyro assembly is connected, the current measured at the transient moment when the starting voltage is applied to the gyro motor is the starting current. When the gyro rotor reaches the rated speed and the working voltage stabilizes, the current measured is the working current. 7. Natural frequency and damping ratio The test methods of gyro assembly natural frequency and damping ratio are frequency response method and step input method. For a gyro assembly with a torquer, when the frequency response method is adopted, sinusoidal current can be input to the torquer to produce an electroexcitation dynamic test method of equivalent sinusoidal angular speed to substitute the angular exciter. When the step input method is adopted, after the gyro assembly works normally, a certain step current (positive or negative) can be input to the self-test torquer to substitute the hang-up station.

Unit test of the flight control system 87 a. Frequency response method: Install the gyro assembly on an angular exciter and make its input reference axis parallel to the angular exciter turntable axis and vertical to local horizontal plane. After the gyro assembly works normally, apply half measuring range to the angular exciter, but the input of the sinusoidal angular speed shall not be greater than 60 degrees/s to change the excitation frequency. Measure the input frequency fn of the gyro assembly whose output lags behind the input speed 90 degrees and the input frequency fn is the natural frequency of the gyro assembly. Adjust the input angular speed, make the output of imaginary frequency (phase shift 90 degrees) component be 0.5 V, reduce the input frequency as 1/10 of the natural frequency or lower and keep the angular speed constant. The reading of the real frequency (in-phase) component under this condition is the damping ratio. For a gyro assembly with a torquer, sinusoidal current can be input to the torquer to produce an electroexcitation dynamic test method of equivalent sinusoidal angular speed to substitute the angular exciter to have the test above. b. Step input method: Install the gyro assembly on a hang-up station and make its input reference axis parallel to the hang-up station turntable axis and vertical to local horizontal plane and then connect the output of the gyro assembly with recorder. After the gyro assembly works normally, make the hang-up station operate at a specified speed and then reduce the speed to 0 suddenly, and then record the output change characteristic curve of the gyro assembly. For a gyro assembly with a self-test torquer, hang-up station may not be used. After the gyro assembly works normally, input a certain step current (positive or negative) to the self-test torquer and then use a recorder to record its output change characteristic curve. Through the comparison between the recorded output change characteristic curve of the gyro assembly and the step characteristic curve of the second-order damping system or calculation, the gyro assembly’s natural frequency and damping ratio can be determined. Refer to Ref. [2], Appendix A. 8. Calibration factor Install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane. It is required that the speed of the rate turntable can change steadily and continuously according to the selected rate point and there is no overshoot when reaching each rate point. According to the threshold and measuring range of the gyro assembly specified in the technical conditions of the product, select the point after conducting R5 series common ratio progressive increase (progressive decrease) according to Ref. [3] and

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appropriate rounding and uniform deletion is allowed. However, there should be no less than 11 points in the measuring range. On the premise of ensuring no overshoot, the point can be selected in the principle of uniform interval change and appropriate increase at low rate for the rate turntable. After the gyro assembly works normally, make the rate turntable increase from 0 rate gradually to positive maximum measuring range of the tested gyro assembly by selected rate point and then decrease to 0 by the selected rate point. After that, gradually increase to negative maximum measuring range in reverse direction and then decrease to 0 by the selected rate point. After the rate stabilizes at each measuring point, read the reading (sampling) for no less than 10 times continuously (sampling interval is 1 second each time and the sampling integral time should meet the requirements of special technical conditions of the product, same below) and take the arithmetic mean value as the output signal value of the measuring point. 9. Threshold Install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane. After the gyro assembly works normally, measure the output of the gyro assembly when the input rate is 0, and then apply a rate input of regulated threshold to the gyro assembly steadily without overshoot and then record the output of the gyro assembly. The change of the output shall be greater than 50% of the output corresponding to the calibrator factor. 10. Resolution Install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane. After the gyro assembly works normally, apply a consistent input rate 20
50 times that of the threshold to the gyro assembly and then apply a rate increment of specified resolution to the gyro assembly steadily without overshoot. The output increment of the gyro assembly shall be greater than 50% of the output increment corresponding to the calibrator factor. After that, return to the original constant input rate and apply a rate decrement of specified resolution steadily without overshoot. The measured output increment of the gyro assembly shall be greater than 50% of the output increment corresponding to the calibrator factor. Repeat the procedure in the reverse direction of the constant input rate. 11. Zero-bias For an ordinary gyro assembly: If a nontilting rate turntable is used, install the gyro assembly on a rate turntable to make its input reference axis parallel to the rate turntable axis (if a tilting rate turntable is used, the rate turntable axis should be made parallel to the local horizontal plane and the output axis of the gyro assembly should be made

Unit test of the flight control system 89 vertical to the local horizontal plane). After the gyro assembly works normally, apply a rate of 20
50 times that of the threshold and a frequency of 0.5
1 Hz to make the turntable have clockwise and anticlockwise back and forth swing. Cut off the power supply for the gyro assembly and continue to swing the rate turntable until the gyro assembly rotor stops rotation. Make the gyro assembly output axis vertical to local horizontal plane carefully, reconnect the power supply for the gyro assembly and measure its output signal after it works normally. According to the precision requirement, correct the influence of earth rotation. For gyroassemblies with a torquer: Make the output axis of the gyro assembly vertical to local horizontal plane and connect power to the torque through a low frequency power supply. The frequency is less than the natural frequency of the gyro assembly and the amplitude is not larger than 1/10 of the maximum torque current, making the gyro assembly frame swing. Slowly reduce the torque current to zero. Then, make the gyro assembly work normally and measure its output signal. According to the precision requirement, correct the influence of the Earth’s rotation. 12. Polarity a. Gyro assembly output polarity: install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane. After the gyro assembly works normally, make the rate turntable operate at an appropriate speed to apply a positive rate input to the gyro assembly and record the polarity of the gyro assembly output signal. b. Self-test torque polarity: for a gyro assembly with a self-test torque, this item should be inspected. After the gyro assembly works normally, input an appropriate positive input current to the torque and record the polarity of the gyro assembly output signal. 13. Measuring range Install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to the local horizontal plane. After the gyro assembly works normally, apply a positive maximum input to the gyro assembly and record its output signal. When the input rate decreases, the output signal decreases in proportion accordingly. Repeat the procedure in the reverse direction of the input rate. 14. Channel consistency Install the gyro assembly on a rate turntable and make its input reference axis parallel to the rate turntable axis and vertical to local horizontal plane. After the gyro assembly works normally, apply positive maximum input to the gyro assembly and record the output signal in each channel and calculate the maximum output voltage difference of channels.

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15. Judging criteria and results handling If the results of the appearance structure inspection of airborne units are consistent with the expected value, the structure of the airborne unit is deemed acceptable. In other words, the appearance is intact and not damaged, the connector is not bent or retracted, the weight and overall dimensions meet requirements, the structure is well installed, and the electrical interface is error free. If the results of power voltage, insulation resistance, preparation time, starting current and working current, natural frequency and damping ratio, calibration factor, threshold, resolution, zero-bias, polarity, measuring range, and channel consistency comply with the special technical conditions of the product, the test items are deemed acceptable; otherwise the test items are deemed unacceptable. The calibrator factor is calculated with the data processing formula below: P P P n ni51 Xiyi 2 ni51 Xi ni51 yi K5 Pn 2 P n ni51 Xi2 2 i51 Xi where K is the calibration factor of the gyro assembly, V (degrees/s); is the input rate at the ith measuring point, (degrees/s); yi is the measured output of the gyro assembly at the ith measuring point, V; and i 5 1, 2, 3, . . . , n, n is the number of measuring points. 2.2.8.6 Control surface motion sensor The control surface motion sensor generally adopts the differential transformer sensor and proximity sensor. A differential transformer sensor, such as LVDT and RVDT, has the main function of detecting the control surface deflection angle. Content on the functional and performance test of this kind of sensor has been described in Section 2.2.8.1, thus it will not be elaborated here. The proximity sensor is usually in the form of a Hall unit and its main function is to detect the control surface motion synchronization. A flaps tilt sensor is taken as an example below to introduce the functional and performance test of this kind of sensor. 1. Test objectives and test requirements It aims to test the function and performance of flaps tilt sensor, to ensure it meets the design requirements and works accurately within the validation scope. The test shall be conducted in accordance with the special technical conditions of the product clause by clause, and each test result should conform to the criteria. Unless otherwise specified, test requirements for the flaps tilt sensor shall be subject to the general requirements in Section 2.2.1.

Unit test of the flight control system 91 2. Test items and test methods Main test items of the flaps tilt sensor include appearance structure, power voltage, power supply current, measuring range, precision, and channel consistency. 3. Appearance structure Inspection items of appearance structure inspection include appearance, weight, overall dimensions, installation form and electrical interface and the inspection methods include visual inspection, electronic balance and ruler. 4. Power supply voltage Adjust the power supply voltage of flaps tilt sensor to the limit value of specified deviation and check its performance. 5. Power supply current Put the flaps tilt sensor magnetic target on a special tooling, fix the probe on a special fixture, apply exciting power at the input end, and connect the output end with the digital multimeter. The digital multimeter is set at DC gear. Adjust the position of the probe to make it be situated in a position with and without inductivity and then measure the current with a digital multimeter when there is inductivity and when there is no inductivity. 6. Measuring range Put the flaps tilt sensor magnetic target on a special tooling, fix the probe on a special fixture, control the automatic turntable to operate, and then measure the angle from one end to the other end of the flaps tilt sensor. 7. Precision Put the flaps tilt sensor magnetic target on a special test bed, fix the probe on a special fixture, apply exciting power at the input end and connect the output end with the digital multimeter. The digital multimeter is set at DC gear. Connect the power supply, control the automatic turntable to operate and monitor the output voltage. Take the first jump of the output voltage (from high level to low level) as the starting zero point of the flaps tilt sensor. In the rotation process of the turntable, record the angle value of each jump and the jump angle difference of two adjacent high levels and low levels is the output voltage accuracy of the flaps tilt sensor. 8. Channel consistency Put the flaps tilt sensor magnetic target on a special test bed, fix the probe on a special fixture, apply exciting power at the input end, and connect the output end with the digital multimeter. The digital multimeter is set at DC gear. Connect the power supply, control the automatic turntable to operate, monitor the output voltage and calculate the maximum difference of output voltage of channels.

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9. Judging criteria and results handling If the results of the appearance structure inspection of airborne units are consistent with the expected value, the structure of the airborne unit is deemed acceptable. In other words, the appearance is intact and not damaged, the connector is not bent or retracted, the weight and overall dimensions meet requirements, the structure is well installed, and the electrical interface is error free. If the results of power voltage, power supply current, measuring range, precision, and channel consistency comply with the requirements of the product specification, the test items are deemed acceptable; otherwise the test items are deemed unacceptable. 2.2.8.7 Atmospheric data sensor An atmospheric data sensor generally adopts a semiconductor piezoresistive sensor and it is mainly used to detect signals including dynamic pressure, static pressure, airspeed, pressure-altitude, and angle of attack/angle of sideslip. This kind of sensor is generally classified as an aircraft avionics system and it will not be elaborated here. 2.2.8.8 Sensor for other purposes Sensors for other purposes, except for an overload sensor, are used for other systems of aircraft and they will not be elaborated here. The main function of the overload sensor is to detect the overload at wingtip and vertical tail. The functional and performance testing of this kind of sensor has been described in Section 2.2.8.3, thus it will not be elaborated further here.

2.3 Strength and stiffness test Strength and stiffness testing shall be carried out for mechanical and electromechanical airborne units.

2.3.1 Test objectives and test requirements The purpose of this test is to evaluate the strength and stiffness of an airborne unit under a specified stress state and confirm whether it meets the design requirements. Testing equipment: they generally include test bed, loading equipment, test equipment, exciter, and energy equipment. Test bed: it should have sufficient stiffness and strength to ensure the accuracy of the test data. The mechanical interfaces connecting with the unit under test and the connection stiffness should be as same as that under the normal working state of the unit under test (or they will not affect the test conclusion after analysis).

Unit test of the flight control system 93 Loading equipment: they should have enough loading capacity and loading accuracy as well as reliable protective measures to prevent product damage caused by overload. Test equipment: sufficient sensors should be provided, the installation form of sensors shall be reasonable and reliable, antiinterference measures shall be taken for the test system and the overall test accuracy and sampling rate shall be evaluated and inspected. Exciter: it should match with the electrical interface connecting with the unit under test and sufficient functions should be equipped to ensure the unit under test works normally in the test process. Energy equipment: the power and hydraulic source provided by the energy equipment should have sufficient power and high quality and the oil brand of the hydraulic source connected with the unit under test should meet requirements.

2.3.2 Test items and test methods The main items of the strength (stiffness) test include working load test, design load test, and stiffness test. 1. Working load test It assesses the capability of the unit to bear the working load. The test should cover the main or the most serious working condition, mode, and position. In the test, increase the load to 100% working load first and then under the loaded state, test the function and performance of the unit, and finally measure the residual deformation of the unit after load removal. 2. Design load test It assesses the capability of unit to bear design load. The test shall cover the main or the most serious working condition, mode, and position. In the test, increase the load to 100% design load first, and then under the loaded state test the function and performance of the unit. 3. Stiffness test It assesses the capability of the unit to resist the deformation caused by external load. The test should cover the main or the most serious working condition, mode, and position. In the test, the load is increased to 40% design load generally.

2.3.3 Judging criteria and results handling Working load test: under 100% working load, compare the measured functional and performance values and the residual deformation value of the unit after load removal with the design values. The measured values shall meet design requirements. Design load test: under 100% design load, compare the measured functional and performance values of the unit lasting for no shorter than 3 seconds with the design values. The measured values shall meet design requirements.

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Stiffness test: compare the measured deformation value of the unit under external load with the design values. The measured value shall meet design requirements.

2.4 Power characteristic test Aircraft power supply system generally provides six types of power source for an airborne electrical unit. According to the categories of direct power supply sources for an aircraft power supply system, the airborne electrical unit is divided into the following six categories: 1. 2. 3. 4. 5. 6.

28 V DC electrical unit. 270 V DC electrical unit. Single-phase 115 V 400 Hz AC electrical unit. Three-phase 115 V/200 V 400 Hz AC electrical unit. Single-phase 15 V variable-frequency AC electrical unit. Three-phase 115 V/200 V variable-frequency AC electrical unit.

For power supply sources for the six types of electrical unit above and the electrical unit (such as gyro assembly) that pass other units on aircraft (such as flight control computer) but do not have power source transformation and processing, power characteristic tests shall be carried out, that is, power supply compatibility test of airborne electrical unit.

2.4.1 Test items and reduction principles Electrical units with different power supply sources have different test items of the power characteristic test. Test items of the six types of airborne electrical unit are shown below and they should be deleted and selected according to the relevant system design specifications. 1. Power characteristic test of 28 V DC electrical unit There are 12 test items of the power characteristic test for a 28 V DC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.12. 2. Power characteristic test of 270 V DC electrical unit There are 12 test items of the power characteristic test for the 270 V DC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.13. 3. Power characteristic test of 115 V 400 Hz AC electrical unit There are 17 test items of the power characteristic test for the 115 V 400 Hz AC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.14.

Unit test of the flight control system 95 Table 2.12: Test items of power characteristic test for 28 V DC electrical unit. Power supply state Normal working

Conversion Abnormal working Emergency Power failure

Test items

Name

LDC101 LDC102 LDC103 LDC104 LDC105 LDC106 LDC201 LDC301 LDC302 LDC401 LDC601 LDC602

Load characteristic Voltage limit under normal steady state Voltage distortion frequency spectrum Voltage pulsation Normal voltage transient Voltage spike Power supply conversion interruption Voltage limit under abnormal steady state Abnormal voltage transient Voltage limit under emergency steady state Power failure due to fault Polarity reversal

LDC, DC low voltage.

Table 2.13: Test items of power characteristic test for 270 V DC electrical unit. Power supply state

Test items

Name

Normal working

HDC101 HDC 102 HDC 103 HDC 104 HDC 105 HDC 106 HDC 201 HDC 301 HDC 302 HDC 401 HDC 601 HDC 602

Load characteristic Voltage limit under normal steady state Voltage distortion frequency spectrum Voltage pulsation Normal voltage transient Voltage spike Power supply conversion interruption Voltage limit under abnormal steady state Abnormal voltage transient Same as HDC 302, not requiring another test Power failure due to fault Polarity reversal

Conversion Abnormal working Emergency Power failure HDC, DC high voltage.

4. Power characteristic test of the three-phase 115 V/200 V 400 Hz AC electrical unit There are 19 test items of the power characteristic test for the three-phase 115 V/ 200 V 400 Hz AC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.15. 5. Power characteristic test of 15 V variable-frequency AC electrical unit There are 16 test items of the power characteristic test for the 15 V variablefrequency AC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.16. 6. Power characteristic test of 115 V/200 V variable-frequency AC electrical unit There are 18 test items of the power characteristic test for the 115 V/200 V variablefrequency AC electrical unit under five power supply states: normal working, conversion, abnormal working, emergency, and power failure, as shown in Table 2.17.

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Table 2.14: Test items of power characteristic test for single-phase 115 V 400 Hz AC electrical unit. Power supply state Normal working

Conversion Abnormal working Emergency Power failure

Test items

Name

SAC101 SAC102 SAC104 SAC105 SAC106 SAC107 SAC108 SAC109 SAC110 SAC111 SAC201 SAC301 SAC302 SAC303 SAC401 SAC601 SAC602

Load characteristic Voltage and frequency limit under normal steady state Voltage modulation Frequency modulation Voltage distortion frequency spectrum Total voltage distortion DC component Normal voltage transient Normal frequency transient Voltage spike Power supply conversion interruption Voltage and frequency limit under abnormal steady state Abnormal voltage transient Abnormal frequency transient Voltage and frequency limit under emergency steady state Power failure due to fault Phase reversal

SAC, Single-phase AC.

Table 2.15: Test items of power characteristic test for three-phase 115 V/200 V 400 Hz AC electrical unit. Power supply state Normal working

Conversion Abnormal working Emergency Power failure

TAC, Three-phase AC.

Test items

Name

TAC101 TAC102 TAC103 TAC104 TAC105 TAC106 TAC107 TAC108 TAC109 TAC110 TAC111 TAC201 TAC301 TAC302 TAC303 TAC401 TAC601 TAC602 TAC603

Load characteristic Voltage and frequency limit under normal steady state Voltage phase difference Voltage modulation Frequency modulation Voltage distortion frequency spectrum Total voltage distortion DC component Normal voltage transient Normal frequency transient Voltage spike Power supply conversion interruption Voltage and frequency limit under abnormal steady state Abnormal voltage transient Abnormal frequency transient Voltage and frequency limit under emergency steady state Three-phase power failure Single-phase and two-phase power failure Phase sequence reversal

Unit test of the flight control system 97 Table 2.16: Test items of power characteristic test for single-phase 15 V variable-frequency AC electrical unit. Power supply state

Test items

Name

SVF101 SVF102 SVF104 SVF106 SVF107 SVF108 SVF109 SVF110 SVF111 SVF201 SVF301 SVF302 SVF303 SVF601 SVF603

Load characteristic Voltage and frequency limit under normal steady state Voltage modulation Voltage distortion frequency spectrum Total voltage distortion DC component Normal voltage transient Normal frequency transient Voltage spike Power supply conversion interruption Voltage and frequency limit under abnormal steady state Abnormal voltage transient Abnormal frequency transient Power failure due to fault Phase reversal

Normal working

Conversion Abnormal working Power failure SVF, Single-phase variable-frequency.

Table 2.17: Test items of power characteristic test for three-phase 115 V/200 V variablefrequency AC electrical unit. Power supply state

Test items

Name

TVF101 TVF102 TVF103 TVF104 TVF106 TVF107 TVF108 TVF109 TVF110 TVF111 TVF201 TVF301 TVF302 TVF303 TVF401 TVF601 TVF602 TVF603

Load characteristic Voltage and frequency limit under normal steady state Voltage phase difference Voltage modulation Voltage distortion frequency spectrum Total voltage distortion DC component Normal voltage transient Normal frequency transient Voltage spike Power supply conversion interruption Voltage and frequency limit under abnormal steady state Abnormal voltage transient Abnormal frequency transient Voltage and frequency limit under emergency steady state Power failure due to fault Single-phase and two-phase power failure Phase sequence reversal

Normal working

Conversion Abnormal working Emergency Power failure

TVF, Three-phase variable-frequency.

As units of the flight control system have different working principles and characteristics and the aircraft power supply system has a special power supply and distribution configuration for some electrical units, the test items of the power characteristic test of the

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electrical unit should be clipped on this basis to avoid the occurrence of over or under requirements. The basic principles of clipping are as follows: 1. Nonsecondary power supply unit of the flight control system shall have a power characteristic test. Capacitive, motor, and electronic units shall have an inrush starting current and duration test. The inductive unit shall have a voltage spike test under power failure. An AC nonlinear unit shall have current harmonics test. A high-power unit shall have a current modulation test. An AC electrical unit shall have a power test. 2. Test items of power characteristic test under steady state. Under normal and abnormal steady-state voltage and frequency limit, if the airborne electrical unit is required to work normally, it shall be verified at the abnormal steady-state voltage and frequency limit. Under abnormal steady-state voltage and frequency limit, performance requirements for airborne electrical unit can be reduced or the unit may stop working and the performance of the unit should be verified under normal and abnormal steady-state voltage and frequency limit. The performance of the three-phase AC airborne electrical unit should be checked under emergency steady-state voltage limit and frequency limit. DC voltage distortion and voltage pulsation tests may not be carried out for a DC electrical unit (such as solenoid valve and motor) insensitive to voltage modulation, frequency modulation, voltage waveform distortion, and DC component. 3. Test items of voltage and frequency transient test. If the flight control system has consistent requirements on the working performance of the airborne electrical unit under normal and abnormal power supply conditions, it is only necessary to conduct an abnormal voltage transient test and frequency transient test. If the requirements on the working performance under the two states are different (the unit may stop working under abnormal power supply), normal voltage transient and frequency transient test and abnormal voltage transient and frequency transient test shall be carried out. 4. Test items of voltage spike test. Voltage spike test may not be carried out for an airborne electrical unit insensitive to voltage spike (such as heating resistance wire, solenoid valve, and motor) or high-power airborne electrical unit (such as the unit whose rated current exceeds 10 A or which can prove the great attenuation of voltage spike generated by voltage spike signal generator). 5. Test items of power supply conversion interruption test. For a DC airborne electrical unit adopting uninterrupted power supply measures (such as uninterrupted busbar and nonsimilar redundancy), a power supply interruption test may not be carried out. For an AC airborne electrical unit, except for those whose performance can be clearly predicted during the power supply interruption such as resistance wire and motor unit, a power supply interruption test shall be carried out (especially the AC electrical unit with digital circuits). For an airborne electrical unit which is sensitive to power supply interruption, the test shall be carried out for different power supply interruption periods within the limit power supply interruption time.

Unit test of the flight control system 99 6. Test items of power failure test. Except for a resistance wire unit whose performance during power failure can be clearly predicted, all other airborne electrical units shall have a power failure test (such as the phase loss or negative phase sequence of three-phase AC unit and the polarity reversal of DC airborne electrical unit or single-phase AC airborne electrical unit) to verify that the unit has necessary protection function under power failure and will not damage or cause other unsafe states of the system and also verify the unit can restore to its normal working performance automatically after a normal power supply is recovered after a power failure.

2.4.2 Test objectives and test requirements The purpose of a power characteristic test is to test whether the airborne electrical unit can achieve relevant functions or meet performance indexes when the aircraft power supply system is under the five working states described in Tables 2.12
2.17. As the power supply system under each working state provides different power supply characteristics for an airborne electrical unit, they have different requirements for the performance of the airborne electrical unit. Therefore relevant technical documents such as the special technical conditions and outline of power characteristic test of the electrical unit shall clearly specify the performance requirements to the unit under different working states of the aircraft power supply system and a power characteristic test shall be carried out. Test equipment and the environment of the power characteristic test mainly include excitation equipment, testing equipment, test system, and test cable. Among them, the test system is the collective name for all other equipment in the test except for the excitation equipment and monitoring equipment, mainly including the test power source, sensor, analog load, and power characteristic parameter test equipment. 1. Test cable Power characteristic refers to the characteristic within 10 cm of the power input interface of the unit under test. If the electrical unit does not have its own cable, the power input end is at the power input interface of the unit. If the electrical unit is delivered with its own cable, the power input end is at the connection part between its own cable and the electrical unit. If the electrical unit has its own power conversion device, the power input end is the power input end of its power conversion device and the internal power distribution shall be realized by the unit itself. To sum up, the cable of the unit under test shall be exactly the same as the installed state, including the wire specification and length. 2. Excitation equipment An airborne electrical unit of the flight control system has signal cross-linking with other units or systems on the aircraft. When carrying out a power characteristic test, the

100 Chapter 2 cross-linking signals shall be simulated through excitation equipment. The excitation equipment in the test should meet the following requirements: a. The excitation equipment shall be able to provide various input and output signals cross-linked with the unit under test, including analog signal, digital signal, and hydraulic signal. b. The signal acquisition or transmission cycle and logical time sequence related to the cross-linking between the excitation equipment and the unit under test shall be consistent with the installed state. c. To ensure the effectiveness of the test, the excitation equipment shall be able to accurately simulate the real environmental conditions of the aircraft during the transient test which deviates from the normal power characteristic for a few milliseconds. d. The working power supply of the excitation equipment must be an independent power source and the excitation equipment shall not share a test power source with the unit under test to avoid the change of power supply voltage in the test affecting the normal operation of the excitation equipment or even damaging the excitation equipment. 3. Monitoring equipment Monitoring equipment provided by the research unit of the unit under test is necessary to judge if the test results are acceptable. Monitoring equipment shall meet the following requirements: a. Monitoring equipment shall be able to monitor all the performances of the unit under test and have the functions including monitoring, measuring, recording, and communication. b. Monitoring equipment shall be able to accurately and completely sample and record all output signals of the unit under test, rather than only sample or detect part of them. c. The communication mode and data update rate between the monitoring equipment and the unit under test shall be consistent with the installed state. d. The test data output by the unit under test shall be recorded by a device that can record the voltage and current curve in real time. An ordinary digital multimeter, indicator light, and pointer instrument should not be used as the monitoring equipment of the unit under test as these instruments cannot accurately capture the transient characteristics of the output of the unit. e. The sampling rate of the transient characteristic monitoring equipment shall be at least three times higher than that of the collected signal to ensure that the performance of the unit can be accurately measured in the transient test. f. The monitoring equipment shall have data storage and playback functions for easy observation and retrieval. g. The monitoring equipment should have the function of judging various performance parameters as far as possible, so that a corresponding fault code or alarm information can be provided for the test personnel when one or more performance parameters deviates from the normal range.

Unit test of the flight control system 101 4. Test system The test system shall be provided by the party undertaking the test. In order to ensure that the test results are accurate and reliable and meet the requirements of relevant standards, the test system should meet following requirements: a. The test power supply shall be able to provide various power supply characteristics to the unit under test, including various steady-state, transient-state, and fault power supply characteristics. b. The test power supply shall have sufficient power capacity, the output power characteristic parameters shall not have obvious on-load deviation due to the work of the unit under test, and it shall be ensured that relevant parameters can be adjusted. In a voltage transient test, it is not allowed that the transient voltage output by the test system cannot reach the specified value because the load current of the unit under test increases during the transient overvoltage test. c. The test power supply shall meet the transient characteristics required by various test conditions. d. The test instruments used for testing the power characteristic parameters and load characteristic of the unit under test shall meet the requirements of relevant standards. See Refs. [4,5]. Sufficient sampling rate should be ensured for the measurement of voltage spike, impulse current, and voltage distortion. e. All equipment of the test system shall be calibrated to be acceptable and within the validity period.

2.4.3 Test items and test methods The content and items of the power characteristic test are determined according to the technical agreement of the unit and relevant specifications and standards and the test methods are subject to the aircraft power characteristic parameter test methods in Ref. [5]. In Section 2.4.5, a flight control computer is taken as an example to introduce the power characteristic test.

2.4.4 Judging criteria and results handling The result handling and judgment of the power characteristic test shall be subject to the aircraft power characteristic parameter test methods in Ref. [5].

2.4.5 Test example: Power characteristic test of flight control computer 2.4.5.1 Test requirements The test shall be carried out according to the aircraft power supply characteristics specified in the aircraft power characteristic parameters test method in Ref. [5] or aircraft type specification.

102 Chapter 2 Taking a flight control computer powered by a three-circuit 28 V DC power source as an example, it shall work normally after setting one power supply circuit as normal and setting other power supply circuits under power supply interruption, abnormal steadystate low voltage limit, abnormal transient-state undervoltage, abnormal steady-state high voltage, or abnormal transient-state overvoltage states. When all power supply circuits are in an abnormal working state, the flight control computer can work by reducing the workload or stop working but not be damaged or cause an unsafe state. During the test, any power supply circuit can be selected for testing and the other two circuits will be disconnected. 2.4.5.2 Test requirements The testing shall be carried out according to the aircraft power supply characteristics specified in the aircraft power characteristic parameters test method in Ref. [5] or the aircraft type specification. Taking a flight control computer powered by a three-circuit 28 V DC power source as an example, it should work normally after setting one power supply circuit as normal and setting the other power supply circuits under power supply interruption, abnormal steady-state low voltage limit, abnormal transient-state undervoltage, abnormal steady-state high voltage, or abnormal transient-state overvoltage states. When all power supply circuits are in abnormal working state, the flight control computer can work by reducing the workload or stop working but not be damaged or cause an unsafe state. During the test, any power supply circuit can be selected for testing and the other two circuits will be disconnected. 2.4.5.3 Test items and test methods The test items include load characteristic test, normal working test of power supply system, abnormal working test of power supply system, and emergency working test of power supply system. 1. Load characteristic test a. Test objective This test aims to verify if the limit values of load power, impulse current, current distortion, and current spectrum of a 28 V DC electrical unit are within the specified range in the flight control computer specification. b. Test methods Install the UUT (unit under test, it refers to flight control computer here, same as below), excitation equipment and monitoring equipment as shown in Fig. 2.2. Except for the current, other parameters shall be measured within 10 cm of the UUT power input end. The current shall be measured on the 28 V DC positive line.

Unit test of the flight control system 103

Figure 2.2 Cross-link in the power characteristic test of flight control computer. Table 2.18: Limit values of load and current distortion. Parameters Impulse current Spike voltage Load power

Requirements on limit values 5 times of rated current (0.01 s # t # 0.2 s), as shown in Fig. 2.3 As shown in Fig. 2.4 Not exceeding 90 W in single channel

Notes: X, time, with second (s) as the unit; Y, the ratio of transient current to rated current.

Turn on the power supply and adjust the voltage to the steady-state rated DC voltage 28 V. Connect power output and supply power to UUT. Record the impulse current and duration, steady-state current, and load power when the UUT starts working and compare them with the limit values in the special technical conditions of the flight control computer. Then, disconnect the power output and record the spike voltage when the UUT is disconnected from the power. The limit values of the load and current distortion of the flight control computer are shown in Table 2.18. c. Judging criteria and results handling Record the results of the load characteristic test. If the limit values of the load power, impulse current, and the spike voltage and current distortion caused by power failure are within the range specified in Table 2.18 or meet the requirements of the product specification, the test is deemed passed.

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Figure 2.3 Limit values of maximum impulse current of electrical unit.

Figure 2.4 Limit values of DC input power voltage spike.

2. Normal working test of power supply system a. Steady-state voltage limit test i. Test requirements: the flight control computer adopts a three-circuit power supply mechanism, that is, two circuits of a 28 V DC power supply and one circuit of an emergency power supply by the flight control battery. The steady-state characteristic limits of an electrical unit of a 28 V power supply system are shown in Table 2.19. ii. Test methods: disconnect the power supply and install the flight control computer, excitation equipment, and monitoring equipment as shown in Fig. 2.5.

Unit test of the flight control system 105 Table 2.19: Steady-state characteristic limits of electrical unit of 28 V power supply system. DC voltage

Normal

Abnormal

Emergency

22
30 V

20.5
32 V

8
32 V

Notes: All indexes are measured at the input end of electrical unit.

Figure 2.5 Connection of flight control computer under steady-state voltage limits.

Conduct the functional and performance test under the rated normal steady-state voltage (28 V), maximum steady-state voltage (30 V), and minimum steady-state voltage (22 V) according to the test method in Section 2.2.3.2, record the test results, and verify the compliance of the function and performance of the flight control computer during the steady-state power supply period. During the normal working test of the power supply system, the flight control computer shall be safe and meet the functional and performance requirements. As the emergency steady-state limit voltage index has covered the normal steady-state voltage index, this test may not be needed if the emergency steadystate voltage limit test is passed. Judging criteria and results handling: if the function and performance of the flight control computer meet the requirements of the special technical conditions of product, the test is passed.

106 Chapter 2 b. Normal voltage transient test i. Test requirements: as the test conditions of the abnormal voltage transient test are stricter than those of the normal voltage transient test, this test may not be carried out if the abnormal voltage transient test is passed. ii. Test methods: disconnect the power supply and install the flight control computer, excitation equipment, and monitoring equipment according to Fig. 2.5. Set the output voltage of the DC power supply according to the limits specified in Fig. 2.6 and set the test conditions of the normal voltage transient test according to Table 2.20. In Table 2.20, the voltage transient requirements to the flight control computer from AA to HH are specified. Within 1 ms, the voltage should rise or decrease to the transient-state voltage shown in Table 2.20 from the steady-state voltage and maintain for the transient voltage duration time shown in Table 2.20. The voltage beyond the duration should recover to the steady-state voltage. For test condition EE, the transient overvoltage duration of 40 V is 50 ms, with an interval of 1 minute, and the test is carried out three times. For test condition HH, the transient undervoltage duration of 16 V is 50 ms, with an interval of 1 minute, and the test is carried out three times. For each test condition, the performance of the flight control computer is monitored during the voltage transient period and the flight control computer is verified according to the performance test procedure to see if it reaches the regulated performance when the aircraft power supply system works normally.

Figure 2.6 Limits in normal voltage transient test of 28 V DC power supply system.

Unit test of the flight control system 107 Table 2.20: Test conditions of normal voltage transient test. Duration of TransientDuration from steady transient-state state Test Steady-state state to transient state of voltage (V) voltage (ms) voltage (ms) conditions voltage (V)

Duration from transient state to steady state of voltage (ms)

Transient overvoltage AA BB CC DD EE

30 30 30 30 30

,1 ,1 ,1 ,1 ,1

40 40 36 36 40 (3 times)

50 50 110 110 50 (per 1 min)

,1

50 50 50 (per 1 min)

,1

,1 ,1

Transient undervoltage FF GG HH

22 22 22

,1 ,1 ,1

16 16 16 (3 times)

,1

After the power supply is restored to the normal steady-state limit, a performance test is carried out according to Section 2.2.3.2 to verify that the flight control computer can achieve the regulated performance when the aircraft power supply system works normally. After all the tests, adjust the voltage to the rated steady-state voltage 28 V and conduct the functional and performance test according to Section 2.2.3.2 to verify the compliance of the function and performance of the tested object during the steady-state power supply period. Judging criteria and results handling: if the function and performance of the flight control computer meet the requirements of special technical conditions of product under normal voltage transient condition, the test is deemed passed. c. Voltage spike test i. Test requirements: DC input voltage spike limit is shown in Fig. 2.7, peak pulse rise time shall be less than 5 μs, and the total pulse duration shall be greater than 10 μs. Spike generator open-circuit internal resistance: 50 6 5 Ω. Spike generator open-circuit voltage waveform: as shown in Fig. 2.7. Times of spike voltage: the flight control computer should be able to bear the positive-polarity and negative-polarity spike voltage for 50 times in 1 minute. ii. Test methods: disconnect the power supply and install the flight control computer, excitation equipment, and monitoring equipment according to Fig. 2.8. Supply power at the normal steady-state voltage and make the flight control computer enter the normal working state, and then apply a positive and

108 Chapter 2

Figure 2.7 Spike generator open-circuit voltage waveform.

Figure 2.8 Connection of spike test equipment.

negative spike voltage for 50 times respectively to the product within 1 minute. The tested object should be able to withstand the regulated spike voltage without fault. After passing the spike voltage, the performance test is conducted according to Section 2.2.3.2 and the test results are recorded. iii. Judging criteria and results handling: record the test result when the power supply system works normally, and give the conclusion of the test according to the functional and performance test results before and after the test. If the function and performance of the flight control computer meet the requirements of the special technical conditions of the product, the test item is deemed passed.

Unit test of the flight control system 109 3. Abnormal working test of power supply system a. Abnormal steady-state voltage limit test According to the steady-state characteristic limit voltage given in Table 2.19, the test is conducted under rated normal steady-state voltage (28 V), maximum steadystate voltage (30 V), and minimum steady-state voltage (22 V), respectively, according to the requirements in Section 2.2.3.2 and the test results are recorded to verify the compliance of the function and performance of the flight control computer during the steady-state power supply period. During the normal working test of the power supply system, the tested object should be safe and reliable. Except for the performances regulated in the special technical conditions whose standards can be relaxed, other regulated technical performance shall be provided. After the power supply system recovers to its normal working range, it should completely meet all functional and performance requirements. As the emergency steady-state limit voltage index has covered the abnormal steady-state voltage index, this test may not be carried out if the emergency steadystate voltage limit test is passed. b. Abnormal voltage transient test. i. Test requirements: the abnormal voltage transient curve is shown in Fig. 2.9. ii. Test method: supply power at normal steady-state voltage and make the flight control computer enter the normal working state, and then conduct an overvoltage and undervoltage surge test according to conditions in Table 2.21.

Figure 2.9 Abnormal voltage transient curve.

110 Chapter 2 Table 2.21: Test conditions of abnormal voltage transient test.

Test Steady-state conditions voltage (V)

Duration from steady state to transient state of voltage (ms)

Duration from transient state to Transient-state Duration of transient- steady state of voltage (ms) voltage (V) state voltage (ms)

Transient overvoltage A B

C D E

F

30 30

30 22 22

22

,1 ,1 Then

45 45 43.4

Then

40

Then

36.6

Then

32.7

Then ,1 ,1 ,1 Then

30 45 (3 times) 45 45 43.4

Then

40

Then

36.6

Then

32.7

Then ,1

22 45 (3 times)

100 100 Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually
100 (per 0.5 s) 100 100 Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually
100(per 0.5 s)

,1 ms 100 ms 200 ms

50 50 Voltage increases gradually Voltage increases gradually Voltage increases gradually Voltage increases gradually
50 (per 0.5 s) 50

,1 ms 15 ms 30 ms

600 ms 3s 6s
,1 ms ,1 ms 100 ms 200 ms 600 ms 3s 6s
,1 ms

Transient undervoltage G H

I J

30 30

30 22

,1 ,1 Then

8 8 12

Then

17

Then

22

Then

28

Then ,1 ,1

30 8 (3 times) 8

60 ms 4.85 s 1s
,1 ms ,1 ms (Continued)

Unit test of the flight control system 111 Table 2.21: (Continued)

Test Steady-state conditions voltage (V) K

L

22

22

Duration from steady state to transient state of voltage (ms)

Duration from transient state to Transient-state Duration of transient- steady state of voltage (ms) voltage (V) state voltage (ms)

,1 Then

8 12

Then

17

Then ,1

22 8 (3 times)

50 Voltage increases gradually Voltage increases gradually
50 (per 0.5 s)

15 ms 30 ms 60 ms
,1 ms

Hybrid transient M

N

30

22

,1 ,1 Then

8 later 45 43.4

Then

40

Then

36.6

Then

32.7

Then ,1 ,1 Then

30 8 later 45 43.4

Then

40

Then

36.6

Then

32.7

Then

22

50
100 Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually
10
100 Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually Voltage decreases gradually


,1 ms 100 ms 200 ms 600 ms 3s 6s
,1ms 100ms 200ms 600 ms 3s 6s


The voltage transient requirements to the flight control computer are specified from test condition A to N. Within 1 ms, the power supply voltage should rise or decrease to regulated transient-state voltage from the steady-state voltage and maintain for the time shown in the table. After passing the time, the voltage should recover to the steady-state voltage. For test condition C and F, the transient overvoltage interval is 0.5 second and the test is carried out three times. For test condition I and L, the transient undervoltage interval is 0.5 second and the test is carried out three times. For test condition M and N, transient overvoltage test is carried out immediately after the transient undervoltage test and the voltage restores to the steady-state voltage finally. For each test condition, the

112 Chapter 2 performance of the flight control computer is monitored during the voltage transient period and the flight control computer may stop working in this period but not be damaged or cause an unsafe state. After the tests, adjust the voltage to the rated steady-state voltage 28 V, conduct a power-on test, and record the test results. Judging criteria and results handling: analyze the test results of the normal working test of the power supply system, and give the conclusion of the test according to the functional and performance test results before and after the test. If the function and performance of the flight control computer meets the requirements of the special technical conditions of product, the test is deemed passed. 4. Emergency working test of power supply system a. Emergency steady-state voltage limit test. Disconnect the power supply and install the flight control computer, excitation equipment, and monitoring equipment according to Fig. 2.8. According to the requirements in Table 2.12, verify the compliance of the function and performance of the flight control computer during the emergency steady-state power supply period under rated normal steady-state voltage (28 V), maximum steadystate voltage (32 V), and minimum steady-state voltage (18 V) respectively. During the emergency working test of the power supply system, the flight control computer shall be safe and meet functional and performance requirements. b. Emergency transient-state voltage limit test. i. Test methods: under the emergency working condition of the power supply system, there are no requirements for a transient limit test generally. As the limit under emergency working condition is equivalent to the limit under the abnormal working condition, this test can directly use the conclusion of the abnormal transient-state voltage limit test. ii. Handling of test interruption: handle according to Section 3.6 in Ref. [6] in case of test interruption due to other reasons. iii. Judging criteria and results handling: a comprehensive conclusion of a power characteristic test is given on the basis of the test results of the flight control computer under normal, abnormal, and emergency working conditions of the power supply system.

2.5 Dielectric strength test Dielectric strength test is carried out between mutually insulated units or between an insulated unit and their ground to judge if the insulating materials or insulation clearance of the element is appropriate by applying regulated voltage in regulated time to determine if

Unit test of the flight control system 113 the element works safely under rated voltage and if it can withstand the overpotential caused by switching, surge, and other similar phenomena. Electrical units of the flight control system mainly include electronic type and electromechanical type equipment. Electromechanical type equipment RVDT is taken as an example below to describe the voltage withstand test of an insulating medium of this kind of unit.

2.5.1 Test objectives and test requirements The purpose of this test is to investigate the dielectric strength of RVDT, so as to determine whether RVDT can work safely under rated voltage and assess whether the insulating material or insulation clearance of RVDT is appropriate. Test equipment required for this test mainly includes high-voltage power source, voltage measuring instrument, current leakage detector, and fault indicator. Specific requirements on the test and equipment are as follows. 1. Property and value of test voltage: the test voltage is 500 V (rms), 50 Hz AC power, and the power waveform shall be as close to sine wave as possible. The power and output impedance shall ensure there is no significant waveform distortion or voltage change under various test loads. 2. The duration of application of voltage is 60 seconds. 3. The speed of application of voltage is 500 V/s (rms) (the voltage shall be uniformly increased from 0 to regulated value). 4. The error of the voltage measuring instrument should not be greater than 5%. 5. If there are requirements on electric current leakage, a current leakage detector should be used for measurement and its error should not be greater than 5% of the regulated value.

2.5.2 Test items and test methods 1. Preparation for test Use a test cable to connect the RVDT with a high-voltage power source and then connect the voltage measuring instrument and current leakage detector. 2. Test voltage and application point Test voltage shall be applied between windings of RVDT and between the winding and housing. 3. Speed and duration of applying voltage They are subject to regulations in Section 2.5.1(2) and (3). 4. Testing of test piece The fault indicator shall be monitored during the test to judge whether the test piece of the unit under test has disruptive discharge and current leakage. After the test, the

114 Chapter 2 function and performance of the test piece shall be tested to determine the influence of the dielectric strength test to the test piece. 5. Attentions When conducting quality conformity test, if applying a higher test voltage, the duration can be shortened, the specific value is given in the special technical conditions of the product, and repeated electrical stress should not be posed to the same medium. When it comes to the end of the test, the voltage should be gradually decreased to prevent surge.

2.5.3 Judging criteria and results handling If there is no breakdown, sparkover, or corona between windings of RVDT and between the winding and housing and the performance test result satisfies the requirements of the product specification, the dielectric strength of the RVDT is deemed acceptable.

2.6 Mechanical environment test Mechanical environment test is a general term for a series of tests, including the acceleration test, vibration test, noise test, and impact test.

2.6.1 Acceleration test 2.6.1.1 Test objectives and test requirements An acceleration test is a test that must be carried out for all units of the flight control system to verify that they can work normally under the expected acceleration and ensure they will not be separated from the mounting bracket or cause other dangers in this environment. Different positions on the aircraft have different acceleration environments, so the units should determine the values for the acceleration test according to the installation position. Usually, a large aircraft is divided into five areas: (1) front and middle fuselage; (2) rear fuselage; (3) horizontal tail and vertical tail; (4) inner side of wings; and (5) outer side of wings. The product specifications will give the acceleration test values of units in each area. 2.6.1.2 Test items and test methods An acceleration test includes the acceleration performance test and structure test, which respectively assess the working performance and structural integrity of units under an acceleration environment. Refer to Ref. [7] for test items and methods. The main steps are as follows: 1. Install the unit on the test bed. 2. Conduct appearance inspection and functional and performance test for the unit before the test.

Unit test of the flight control system 115 3. Load according to the acceleration performance test values, maintain the test value for 1 minute, and carry out in-test functional and performance test at the same time. 4. Unload and conduct appearance inspection and functional and performance test of the unit. 5. Load according to the acceleration structure test values, and maintain the test value for 1 minute (in-test functional and performance test are not required unless there are special requirements). 6. Unload and conduct posttest appearance inspection and functional and performance test of the unit. 7. Change the loading direction of the unit under test and repeat (3)
(6). To simplify the test, all pretest and posttest test items can be uniformly carried out before and after all loading tests, but failure to pass the test may cause some difficulties in the analysis of failure reasons. 2.6.1.3 Judging criteria and results handling The appearance, function, and performance test of the unit during and after the test shall meet requirements and the unit shall have no residual deformation of structure, cracks, looseness, fall-off, and other mechanical damages, or it is deemed to have failed the test.

2.6.2 Vibration test 2.6.2.1 Test objectives and test requirements A vibration test is a necessary test for all units of the flight control system to determine the capability of the units to withstand a vibration environment. Under an expected vibration environment, the performance of units should not degrade and the structure should not be damaged. As the aircraft has different vibration environments at different positions, the units shall select test values for the vibration test according to the installation position. Generally, a large aircraft is divided into 16 vibration areas: (1) nose; (2) middle fuselage; (3) rear fuselage; (4) fuselage tail cone; (5) inner side of horizontal tail; (6) outer side of horizontal tail; (7) central wing; (8) inner side of wing; (9) outer side of wing and aileron; (10) engine blower casing; (11) main landing gear compartment; (12) landing gear; (13) side wall of fuselage; (14) flaps; (15) engine nacelle and pylon; and (16) engine and accessories (excluding blower casing). The vibration test values, duration, and weight attenuation factor of the units in each area are given in the aircraft type specification. 2.6.2.2 Test items and test methods The vibration test includes the functional vibration test and endurance vibration test, which respectively assess the functional integrity and structural integrity of the units in a vibration environment.

116 Chapter 2 The test items and methods are referred to in Ref. [8]. The main steps are as follows: 1. Install the unit on the test bed. 2. Conduct appearance inspection and functional and performance test for the unit before the test. 3. Load according to the functional vibration test values and carry out in-test functional and performance test at the same time. The test shall last for 0.5 hours. 4. Unload and conduct appearance inspection and functional and performance test of the unit. 5. Load according to the endurance vibration test values and the test duration shall follow the requirements of aircraft type specification. Unless there are special requirements, intest functional and performance tests are not required. 6. Unload and conduct appearance inspection and functional and performance test of the unit. 7. Load according to the functional vibration test values and carry out in-test functional and performance test at the same time. The test shall last for 0.5 hours. 8. Unload and conduct appearance inspection and functional and performance test of the unit. 9. Change the loading direction of the unit under test and repeat (3)
(8). To simplify the test, all pretest and posttest test items can be uniformly carried out before and after all loading tests, but failure to pass the test may cause some difficulties in the analysis of failure reasons. 2.6.2.3 Judging criteria and results handling The appearance, function, and performance test of the unit during and after the test shall meet requirements and the unit shall have no residual deformation of structure, cracks, looseness, fall-off, and other mechanical damages, or it is deemed to have failed the test.

2.6.3 Noise test The noise test of units has two purposes: one is to assess the working performance and strong noise resisting capability of the units in an environment with strong noise and to test the response of units to strong noise; and the other is to assess the sound pressure level of noise emitted from units installed in the cockpit. For units installed near the engine, the response of them to strong noise should be tested. For units with motors or fans, the sound pressure level of the noise emitted by them should be assessed. The installation positions of units on the flight control system are mostly far away from this area and are in a good installation environment. Therefore it is unnecessary to assess their noise resisting capability generally. However, for units with motors such as the angular rate gyroscope assembly, the sound pressure level of the noise emitted by them should be measured.

Unit test of the flight control system 117 An angular rate gyroscope assembly is taken as an example below to introduce the noise testing of this kind of unit. 2.6.3.1 Test objectives and test requirements The test aims to measure the noise caused by an angular rate gyroscope assembly and assess if the sound pressure level of the noise meets the requirements of noise radiation limits. When carrying out outdoor measurement, the wind speed should be lower than 6 m/s (equivalent to grade 4 wind) and wind cover should be used to avoid the effects brought by strong electric and magnetic field, strong wind, high and low temperature, and air exhaust impact of units under test. The test equipment mainly includes a microphone, test cable, installation fixture, and sound calibrator. Test equipment mainly includes integrating-averaging sound level meter and timer. The precision of the instruments, meters, and test devices used to control or monitor test parameters should be checked before the test and conform to the relevant national standards or calibration specifications of metrological departments. The precision should not be lower than one third of the performance error and the instruments should be within an effective calibration period and in normal operation. 2.6.3.2 Test items and test methods The noise test of units shall include the direct test on the sound pressure level of the noise and also the correction considering background noise. The test frequency range is generally 100 Hz
10 kHz and the measurement results are given at one thirds of an octave. If the SNR does not meet the test requirements, the effective test frequency band should be determined. Refer to Ref. [9] for specific test methods and record the test results. 2.6.3.3 Judging criteria and results handling If the noise radiation value of an angular rate gyroscope assembly meets the expected value specified in the product specification, the angular rate gyroscope assembly is judged to meet the requirements for noise radiation; otherwise it is judged to not meet the requirements for noise radiation.

2.6.4 Shock test 2.6.4.1 Test objectives and test requirements A shock test is a necessary test for all units of the flight control system to determine the shock resisting capability of the units. Under an expected shock environment, the performance of the units shall not degrade and the structure shall not be damaged.

118 Chapter 2 A shock test of units of large aircraft generally includes a functional shock test and crash safety test. The crash safety test includes a crash shock test and sustained acceleration test. A crash safety test is not mandatory for every unit but necessary for the following units. 1. The units that may fall off and endanger the passengers after being damaged in an aircraft crash, such as the steering wheel, control column, and pedal in the cockpit. 2. The units that may penetrate the fuel tank, piping, or damage adjacent systems and lead to fire or injurious explosions after being damaged in an aircraft crash. 3. The units that may affect the normal operation of the lifesaving system, fire extinguishing system, and fault recording system after being damaged in an aircraft crash. The shock test values of all units are generally the same and the functional shock, crash shock, and sustained acceleration test values and duration are given in the aircraft type specification. 2.6.4.2 Test items and test methods The test items and methods can be referred to in Ref. [8]. The main steps are as follows: 1. Install the unit on the shock test bed. 2. Calibrate the functional shock waveform. 3. Conduct an appearance inspection and functional and performance test for the unit before the test. 4. Apply a functional shock load and have a shock three times in each shock direction. The time interval between two adjacent shocks shall ensure the shocks will not be mutually affected when they have response on the unit, generally being greater than five times that of the shock duration. Conduct the functional and performance test for the unit in test process. 5. End the functional shock test in the direction and conduct appearance inspection and functional and performance test for the unit. 6. Change the shock load direction of the unit under test and repeat (4)
(5). 7. Start the crash shock test if necessary. 8. Calibrate the crash shock waveform. 9. Conduct appearance inspection and functional and performance test for the unit before the test. 10. Apply the crash shock load and have a shock twice in each shock direction. The time interval between two adjacent shocks shall ensure the shocks will not be mutually affected when they have response on the unit, generally being greater than five times that of the shock duration. The unit is generally in an abnormal working state in the test process. 11. End the crash shock test in the direction and conduct appearance inspection for the unit. 12. Change the shock load direction of the unit under test and repeat (10)
(11).

Unit test of the flight control system 119 13. Install the unit on acceleration test bed and start the sustained acceleration test. 14. Apply sustained acceleration load to the unit under test and maintain it for 3 seconds after it reaches the test value, and then unload. The unit is generally in an abnormal working state in the test process. 15. End the test in the direction and conduct appearance inspection for the unit. 16. Change the acceleration direction of the unit under test and repeat (14)
(15). 2.6.4.3 Judging criteria and results handling The appearance, function, and performance test of the unit during and after the functional shock test should meet requirements and the unit should have no residual deformation of structure, cracks, looseness, fall-off, and other mechanical damage, or else it is deemed to have failed the test. If the unit is found to have separated from its mounting bracket in the appearance inspection after the crash safety test, it is deemed to have failed the test.

2.7 Natural environment test Natural environment test is a general term for a series of tests, including the low-pressure (altitude) test, high-temperature test, low-temperature test, temperature shock test, temperature
altitude test, single event effect test, solar radiation test, rain test, icing test, damp heat test, mold test, salt spray test, sand and dust test, and immersion test. It mainly aims to assess the adaptability of units to changes in environmental factors such as temperature and air pressure in the process of storage, transportation, and use, as well as the adaptability to harsh natural environments such as high temperature, intense coldness, damp heat, mold, salt fog, dust, rain, and snow in the global scope. The applicability of test items of the natural environment test is closely related to the installation position of the units on the aircraft. A conventional large aircraft is divided into five areas, as shown in Table 2.22. Table 2.23 lists the test items of the natural environment test applicable to units installed in different areas.

2.7.1 Low pressure (altitude) test 2.7.1.1 Test objectives and test requirements The low-pressure test aims to determine the adaptability of units to a low-pressure environment during storage, transportation, and operation. Mechanical products mainly assess the external tightness of hydraulic and pneumatic units. Electronic products mainly assess whether the components with a sealed chamber are damaged and whether

120 Chapter 2 Table 2.22: Division of areas for national environment test on conventional large aircraft. Areas

Location

Area Area Area Area

In cockpit In cargo compartment and lower equipment compartment In engine nacelle Wholly or partially exposed to the outside environment of the aircraft skin and landing gear compartment wall Other positions, such as wing, empennage, equipment compartment at tail of fuselage, radar compartment, equipment compartment at the left and right of front landing gear, equipment compartment in front of and behind main landing gear, and the equipment compartment in front of and behind wing body rectifying section

a b c d

Area e

Table 2.23: Test items of natural environment test in different areas. Areas Test items

a

b

Low air pressure

Storage O Working O Rapid decompression O* High temperature Storage O Working O Low temperature Strorage O Working O Temperature shock O* O Temperature
altitude O O Temperature
humidity
altitude O* O* Single event effect O O Solar radiation Cyclic heat effect O Steady-stage long-term photochemical effect O* Rain test Rain test with wind source N/A Water drop N/A Waterproof Damp heat Mold Salt fog Sand and dust test Icing test

Infusing

O O O Sand blowing Dust blowing Icing Freezing Ice accretion Units for emergency use on water

* means that more details on applicability can be found in relevant chapters.

c

d

e

O O O* O O O O O O O* O N/A N/A N/A N/A

O O N/A O O O O O O O* O N/A N/A N/A O*

O O N/A O O O O O O O* O N/A O O* O*

O O N/A O O O O

N/A N/A N/A O*

Aircraft, engine nacelle, and some large units O O O N/A N/A N/A N/A N/A

O O O O* O* N/A N/A N/A

O O O N/A N/A N/A N/A N/A

O O O O O O O O

N/A N/A N/A O* N/A

Unit test of the flight control system 121 their function (performance) is affected by the fatigue load caused by the change of air pressure. The low-pressure test includes three test procedures: storage test, working test, and rapid decompression test. The storage test is applicable to all nonpure mechanical units. The working test is applicable to electromechanical and electronic units. The rapid decompression test is applicable to key units (grade A) and important units (grade B) installed in area a and b (see Table 2.22 for definition of area division), such as primary flight control computer, actuator controller, flaps and slats controller, automatic flight control computer, three-axis angular rate gyroscope assembly, threeaxis acceleration sensor assembly, command sensor, cockpit control device, and control handle. The rapid decompression test can be carried out together with the working test or separately. The test procedures and requirements for the three tests are shown in Table 2.24. The allowable pressure error is less than 6 5%. 2.7.1.2 Test equipment and environmental requirements The test chamber shall meet the requirements of test conditions specified in Table 2.24 and there shall be auxiliary instruments monitoring various test conditions. When the pressure restores, the air filled into the test chamber should be dry and clean, without contaminating the unit under test. 2.7.1.3 Test items and test methods 1. Low-pressure storage test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Table 2.24: Requirements for test procedures of low-pressure test. No. Test procedure 1

Storage test

2

Working test

3

Rapid decompression test

Test pressure

Duration

57 kPa (corresponding to 4550 m) 57 kPa (corresponding to 4550 m)

No shorter than 1 h

Drop to 18.8 kPa (corresponding to 12,200 m) from 57 kPa (corresponding to 4550 m)

Pressure change rate No greater than 10 Pa/min No greater than 10 kPa/min

Time necessary for completing intermediate data measurement No shorter than 10 min, the Pressure restoration decompression time shall speed: No greater not be greater than 15 s than 10 kPa/min

122 Chapter 2 Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data as the initial test data. Place the unit in a structure for storage or transportation, put it in a low-pressure chamber, keep the mechanical and electrical connection part to be exposed in the storage or transportation process exposed in its original state, and provide the mechanical and electrical connection part by covering them in the storage or transportation process. An appropriate spacing shall be kept between the unit and test chamber wall, bottom and top, so that the air can circulate freely. Drop the pressure in chamber to 57 kPa at a rate not greater than 10 kPa/min and then maintain the pressure for no shorter than 1 hour. Then, restore the pressure in chamber to normal test atmospheric pressure at a rate of no greater than 10 kPa/min. Take the unit out from the chamber and make it reach a stable temperature under the normal test atmospheric condition. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications, take the test data as final test data and compare them with initial test data. 2. Low-pressure working test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, record the test data, and take them as the initial test data. Put the unit in a low-pressure chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. When installing the unit, its working state required in the test process should be taken into account to reserve sufficient space and freedom for it to function. After the unit is installed, it should start working and be checked. No faults should be caused by inappropriate installation. Drop the pressure in the chamber to 57 kPa at a rate not greater than 10 kPa/min, start the unit to work according to relevant standards or product specifications, check its performance, and take the data as intermediate test data. Then, restore the pressure in the chamber to normal test atmospheric pressure at a rate of no greater than 10 kPa/min. Take the unit out from the chamber and make it reach a stable temperature under the normal test atmospheric conditions.

Unit test of the flight control system 123 Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications, take the test data as the final test data and compare them with the initial test data. 3. Rapid decompression test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a low-pressure chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in the actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. Drop the pressure in the chamber to 57 kPa at a rate not greater than 10 kPa/min. Then drop the pressure in the chamber to 18.8 kPa from 57 kPa within 15 seconds and maintain the pressure within 10 minutes. After that, restore the pressure in the chamber to normal test atmospheric pressure at a rate of no greater than 10 kPa/min. Take the unit out from the chamber and make it reach a stable temperature under the normal test atmospheric conditions. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications, take the test data as final test data and compare them with the initial test data. 2.7.1.4 Judging criteria and results handling In the test process, the appearance of the unit under test should not have obvious deformation or damage and the test sample should not lose its function or work abnormally. The initial test data, intermediate test data, and final test data measured during the test should conform to relevant standards or product specification. 2.7.1.5 Test cases A type of rubber trim mechanism is taken as an example to introduce the low-pressure test of units. The rudder trim mechanism is an electromechanical product installed in the cockpit. According to the regulations, the test procedures of the low-pressure test that shall be carried out are storage test, working test, and rapid decompression test. The test items are shown in Table 2.25.

124 Chapter 2 Table 2.25: Test data requirements of low-pressure test. No.

Test procedures

1

Storage test

2

Working test

3

Rapid decompression test

Initial test data Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance

Intermediate test data N/A

Pull-in voltage and release voltage of brake, making current of brake, noload current of motor N/A

Final test data Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance Pull-in voltage and release voltage of brake, making current of brake, no-load current of motor, rated current of motor, the maximum working current of motor, RVDT characteristic, working speed, inertial slippage, axial clearance

2.7.2 High temperature test 2.7.2.1 Test objectives and test requirements The high-temperature test aims to determine the adaptability of units to the hightemperature environment during storage, transportation, and operation. Mechanical products mainly assess motion stability and the external tightness of the hydraulic and pneumatic units. Electronic products mainly assess operation stability and whether there is damage due to thermal stress. The high-temperature test includes two test procedures: storage test and working test. The storage test and working test are applicable to all mechanical, electromechanical, and electronic units. Requirements for the high-temperature test are shown in Table 2.26, the definition of area division is shown in Table 2.22, the allowable temperature error is 6 2 C and the temperature change rate is generally not greater than 10 C/min.

Unit test of the flight control system 125 Table 2.26: Requirements for high-temperature test. Test temperature No. 1 2

Test procedures

a

b

Storage test









Working test

c 

d

e

Duration





48 h





Time required for completing intermediate test data measurement 30 min

70 C 70 C 150 C 70 C 70 C 

Continuous 55 C 70 C 90 C 70 C 70 C working Short-time 70 C N/A N/A N/A N/A working

2.7.2.2 Test equipment and environmental requirements The test chamber shall meet the test conditions specified in Table 2.26 and have sensors or auxiliary instruments to monitor various test conditions. To maintain the uniformity of test conditions, forced air circulation can be adopted, but the air speed around the unit under test shall not exceed 1.7 m/s to prevent impractical heat conduction in the unit under test. The absolute humidity in the test chamber shall not exceed 20 g/m3. 2.7.2.3 Test items and test methods 1. High-temperature storage test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance and take the test data as initial test data. Place the unit in a structure for storage or transportation and put it in a temperature chamber. Keep the mechanical and electrical connection part to be exposed in the storage or transportation process exposed in its original state, and apply appropriate covering measures at the mechanical and electrical connection part to be protected in the storage or transportation process. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. Increase the temperature in the test chamber to the temperature specified in Table 2.26 at a rate not exceeding 10 C/min and preserve the temperature for 48 hours. During the temperature preservation period, the relative humidity in the test chamber shall not be greater than 15%. Then, restore the temperature in the test chamber to the normal test atmospheric condition at a speed of no more than 10 C/min. After the temperature of the unit stabilizes, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications, take the test data as final test data and compare them with the initial test data.

126 Chapter 2 2. High-temperature working test Units in an airtight compartment should have tests at 55 C and 70 C while units in a nonairtight compartment need to be tested at 70 C only. Short-time working units only need several working cycle tests within a regulated time while long-time working units should always be in a working state. The test profile is shown in Fig. 2.10. Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. When installing the unit, its working state required in the test process should be taken into account to reserve sufficient space and freedom for it to work. After the unit is installed, it should start to work and be checked. Any faults caused by inappropriate installation are not allowed. Increase the temperature in the test chamber to the temperature specified in Table 2.26 at a rate not exceeding 10 C/min and preserve the temperature for the time specified in Table 2.27. Then, start the unit to work according to relevant standards or product specifications and check its performance, and take the data as intermediate test

Figure 2.10 Temperature profile of high-temperature working test.

Unit test of the flight control system 127 Table 2.27: Requirements for temperature preservation time. Weight of unit under test G (kg) G # 1.5 1.5 , G # 15 15 , G # 150 G . 150

Temperature preservation time (h) 1 2 4 8

data. For units that should have a high-temperature short-time working test, the temperature in the test chamber should be increased to the temperature specified in Table 2.26 and then the temperature should be stabilized for the time specified in Table 2.27 after completing the steps above. After that, start the test sample to work for 30 minutes according to relevant standards or the product specifications and check its performance and take the data as intermediate test data. Restore the temperature in the test chamber to the normal test atmospheric condition at a speed of no more than 10 C/min. After the temperature of the unit stabilizes, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit as per relevant standards or product specifications, take the test data as final test data and compare them with initial test data. The high-temperature storage test and high-temperature working test can be carried out in succession. That is, to say, after the temperature is preserved for a specified time in the high-temperature storage test, the unit will start working to have an intermediate test, and then the temperature will be restored to the normal test atmospheric condition for the final test. The test conclusion obtained in this way is the test conclusion of both the high-temperature storage test and high-temperature working test. 2.7.2.4 Judging criteria and results handling In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally. The initial test data, intermediate test data, and final test data measured during the test shall conform to relevant standards or the product specification. 2.7.2.5 Test example An aileron trim mechanism is taken as an example below to introduce the high-temperature test. As it is a mechanical product installed in the cockpit, the high-temperature test procedures that shall be carried out according to regulations include storage test and working test (including continuous working test and short-time working test). The test methods shall follow the requirements in Section 2.7.2.3 and the test items are shown in Table 2.28.

128 Chapter 2 Table 2.28: Examples of required test data for high-temperature test. No.

Test procedures

Initial test data

1

Storage test

2

Working Continuous test working test

Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage

Short-time working test

Intermediate test data N/A

Working current, position feedback voltage range

Working current, position feedback voltage range

Final test data Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage Working stroke, full-stroke working time, working current, position feedback voltage range, axial clearance, no-load inertial slippage

2.7.3 Low temperature test 2.7.3.1 Test objectives and test requirements The low-temperature test aims to determine the adaptability of units to a low-temperature environment during storage, transportation, and operation. Mechanical products mainly assess motion stability and the external tightness of hydraulic and pneumatic units. Electronic products mainly assess operation stability and whether there is damage due to thermal stress. The low-temperature test includes two test procedures: storage test and working test. The storage test and working test are applicable to all mechanical, electromechanical, and electronic units. The environmental temperature for the low-temperature storage test and working test is required to be 255 C, the allowable temperature error is 6 2 C, and the temperature change rate is generally not greater than 10 C/min. 2.7.3.2 Test equipment and environmental requirements The test chamber shall meet the test conditions specified in Section 2.7.3.1 and have sensors or auxiliary instruments to monitor various test conditions. To maintain the uniformity of test conditions, forced air circulation can be adopted, but the air speed around

Unit test of the flight control system 129 the unit under test shall not exceed 1.7 m/s to prevent impractical heat conduction in the test sample. 2.7.3.3 Test items and test methods 1. Low-temperature storage test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit, check its appearance, and take the test data as initial test data. Place the unit in a structure for storage or transportation and put it in a temperature chamber. Keep the mechanical and electrical connection part to be exposed in the storage or transportation process exposed in its original state, and apply appropriate covering measures at the mechanical and electrical connection part to be protected in the storage or transportation process. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. Drop the temperature in the test chamber to 255 C at a rate not exceeding 10 C/min and preserve the temperature for 24 hours after it stabilizes (follow the time specified in Table 2.27). Then, restore the temperature in the test chamber to normal test atmospheric condition at a speed of no more than 10 C/min. After the temperature of the unit stabilizes, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications. Take the test data as final test data and compare them with the initial test data. 2. Low-temperature working test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. When installing the unit, its working state required in the test process should be taken into account to reserve sufficient space and freedom for it to work. After the unit

130 Chapter 2 is installed, it should start to work and be checked. Any faults caused by inappropriate installation are not allowed. Drop the temperature in the test chamber to 255 C at a rate not exceeding 10 C/min and preserve the temperature for the time specified in Table 2.27. Then, start the unit to work according to relevant standards or the product specifications and check its performance, record the result, and take the data as intermediate test data. After that, restore the temperature in the test chamber to normal test atmospheric condition at a speed of no more than 10 C/min. After the temperature of the unit stabilizes, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specifications. Take the test data as final test data and compare them with initial test data. The low-temperature storage test and low-temperature working test can be carried out in succession. That is, to say, after the temperature is preserved for the specified time in the low-temperature storage test, the unit will start to work to have the intermediate test, and then the temperature will be restored to normal test atmospheric condition for the final test. The test conclusion obtained in this way is the test conclusion of both the low-temperature storage test and low-temperature working test. 2.7.3.4 Judging criteria and results handling In the test process, the appearance of the unit under test should not have obvious deformation or damage and the test sample should not lose its function or work abnormally. The initial test data, intermediate test data, and final test data measured during the test should conform to relevant standards or the product specification. 2.7.3.5 Test example A type of rudder mechanical backup actuator is taken as an example to introduce the lowtemperature test of units. As an electromechanical product, the rudder mechanical backup actuator is installed on the rear beam of the horizontal stabilizer. According to regulations, the low-temperature test procedures that shall be carried out include the storage test and low-temperature working test. Test items are shown in Table 2.29.

2.7.4 Temperature shock test 2.7.4.1 Test objectives and test requirements The temperature shock test aims to determine the adaptability of units to the sudden change of atmospheric temperature around during storage, transportation, and operation. Mechanical products mainly assess motion stability and the external tightness of hydraulic and pneumatic units. Electronic products mainly assess operation stability and whether there is fatigue damage due to thermal stress.

Unit test of the flight control system 131 Table 2.29: Examples of required test data for low-temperature test. Test No. procedure 1

Storage test

2

Working test

Initial test data Working sealing, insensitive current, zero-bias current, resistance, insulation resistance, maximum speed, working threshold, limit cycle, actuator cylinder sensor output voltage, maximum working stroke, limit stroke, position accuracy, hysteresis, frequency characteristic, internal leakage Working sealing, insensitive current, zero-bias current, resistance, insulation resistance, maximum speed, working threshold, limit cycle, actuator cylinder sensor output voltage, maximum working stroke, limit stroke, position accuracy, hysteresis, frequency characteristic, internal leakage

Intermediate test data N/A

Working sealing, insensitive current, zero-bias current

Final test data Working sealing, insensitive current, zero-bias current, resistance, insulation resistance, maximum speed, working threshold, limit cycle, actuator cylinder sensor output voltage, maximum working stroke, limit stroke, position accuracy, hysteresis, frequency characteristic, internal leakage Working sealing, insensitive current, zero-bias current, resistance, insulation resistance, maximum speed, working threshold, limit cycle, actuator cylinder sensor output voltage, maximum working stroke, limit stroke, position accuracy, hysteresis, frequency characteristic, internal leakage

The temperature shock test is applicable to electrical units, displays, instruments, and units with glass products installed in area a and all units installed in area b, c, d, and e (the definition of area division is shown in Table 2.22). In the test, the high temperature is 70 C, the low temperature is 255 C, and the allowable temperature error is 6 2 C. The time from a high-temperature test chamber and a low-temperature test chamber shall not be longer than 5 minutes and the cycle times shall be three times. 2.7.4.2 Test equipment and environmental requirements The high-temperature test chamber shall meet the requirements of Section 2.7.2.2 and the low-temperature test chamber shall meet the requirements of Section 2.7.3.2. The volume of the test chamber shall ensure that the test chamber temperature can reach the 6 2 C error range after the test sample is put in the chamber for no more than 10% of the test temperature preservation time (see Table 2.27). 2.7.4.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1.

132 Chapter 2 Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. Increase the temperature in the test chamber to 70 C and hold the temperature for the time specified in Table 2.27. Then, transfer the unit under test to a test chamber with temperature adjusted to 255 C in 5 minutes and hold the temperature for the time specified in Table 2.27. And then transfer the unit under test to a test chamber with temperature adjusted to 70 C in 5 minutes and hold the temperature for the time specified in Table 2.27. The test steps above are defined as one cycle, and a total of three cycles are completed. For large or heavy units under test, the time from one test chamber to another test chamber may be appropriately relaxed according to the minimum time actually required. After completing three cycles, take the unit out from the chamber and make it reach a stable temperature under the normal test atmospheric condition. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or product specifications. Take the test data as final test data and compare them with initial test data. 2.7.4.4 Judging criteria and results handling In the test process, the appearance of the unit under test should not have obvious deformation or damage and the test sample should not lose its function or work abnormally. The initial test data, intermediate test data and final test data measured during the test should conform to relevant standards or the product specification. 2.7.4.5 Test example An aileron mechanical backup actuator is taken as an example to introduce the temperature shock test of units. As an electromechanical product, the aileron mechanical backup actuator is installed on the wing rear beam near the wingtip. According to regulations, the temperature shock test shall be carried out and the test items are shown in Table 2.30.

Unit test of the flight control system 133 Table 2.30: Examples of required test data for temperature shock test. Test procedure Temperature shock test

Initial test data Working sealing, internal leakage, maximum working stroke, maximum speed, hysteresis, linearity, output consistency, frequency characteristic, step characteristic, resistance, insulation resistance, null voltage, fault threshold

Intermediate test data N/A

Final test data Working sealing, internal leakage, maximum working stroke, maximum speed, hysteresis, linearity, output consistency, frequency characteristic, step characteristic, resistance, insulation resistance, null voltage, fault threshold

2.7.5 Temperature
altitude test 2.7.5.1 Test objectives and test requirements The temperature
altitude test aims to determine the adaptability of units to the individual or combined effects of high-temperature, low-temperature, and low-pressure environments during storage, transportation, and operation. The temperature
altitude test has in total 10 test procedures. The test procedures and requirements are shown in Table 2.31. For electromechanical and electronic units that work at an altitude of more than 4550 m, all 10 test procedures should be completed. For other units, only three test procedures, that is, ground low-temperature storage test, ground high-temperature storage test, and atmospheric pressure low-temperature frosting test are required to be completed. The allowable temperature error is 6 2 C and the allowable pressure error is 6 5%. 2.7.5.2 Test equipment and environmental requirements The test chamber shall meet the requirements of test conditions specified in Section 2.7.5.1 and there shall be sensors and auxiliary instruments monitoring various test conditions. The temperature change rate in the test chamber shall not be greater than 10 C/min and the pressure change rate in the test chamber shall not be greater than 1.7 kPa/s. 2.7.5.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in

134 Chapter 2 Table 2.31: Requirements for temperature
altitude test. No. 1 2

Ground low-temperature storage test Grounding low-temperature working test

Test temperature

Test altitude (atmospheric pressure)

Duration

262 C

Ground

2h

255 C

Ground

When the temperature stabilizesa, the time required to complete the test item When the temperature stabilizesa, the time required to complete the test item The time required for frosting and defrosting The time required to complete the test item



3

Low-temperature lowpressure working test

255 C

12,000 m

4

Atmospheric pressure lowtemperature frosting test Normal-temperature atmospheric pressure working test Ground high-temperature storage test Ground high-temperature continuous working test Ground high-temperature discontinuous working test Air high-temperature continuous working test Air high-temperature discontinuous working test

210 C

Ground

Room temperature

Ground

85 C

Ground

16 h

70 C

Ground

4h

70 C

Ground

30 min

30 C

12,000 m

4h

47 C

12,000 m

30 min

5 6 7 8 9 10 a

Test procedures

For the unit under test, the time required for the temperature stabilizes shall follow regulations in Table 2.27.

the test and protect the mechanical and electrical connection part in actual work with a covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely. When installing the unit, its working state required in the test process should be taken into account to reserve sufficient space and freedom for it to work. After the unit is installed, it should start to work and be checked. Any faults caused by inappropriate installation are not allowed. 1. Ground low-temperature storage test Drop the temperature in test chamber to 262 C when the unit under test is not connected to a power supply or other power sources, and hold the temperature for 24 hours after the temperature stabilizes (follow the time specified in Table 2.27). Conduct appearance inspection and external sealing inspection of the unit under test without changing the temperature in the chamber and take the data as intermediate test data. 2. Ground low-temperature working test Drop the temperature in the test chamber to 255 C and preserve the temperature for the time specified in Table 2.27. Then, start the unit to work according to relevant

Unit test of the flight control system 135

3.

4.

5.

6.

7.

8.

standards or the product specifications and check its performance. Take the data as intermediate test data. Repeat the test procedure for three times. Low-temperature low-pressure working test Drop the temperature in the test chamber to 255 C and preserve the temperature for the time specified in Table 2.27. Then, maintain the temperature in the test chamber and adjust the pressure in the chamber to that at an altitude of 12,000 m. Start the unit to work according to relevant standards or the product specification and check its performance, Take the data as intermediate test data. Atmospheric pressure low-temperature frosting test Drop the temperature in the test chamber to 210 C when the unit under test is not connected to a power supply or other power source and wait for the temperature in the unit become stable (follow the time specified in Table 2.27). Open the door of the chamber and make it frost on the surface of the unit under test. The door should be open long enough to melt the frost without evaporating the moisture. Then, close the door and start the unit to work according to relevant standards or the product specification and check its performance. Take the data as intermediate test data. Repeat the test procedure for three times. Normal-temperature atmospheric pressure working test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Start the unit to work according to relevant standards or the product specification and check its performance. Take the data as intermediate test data. Ground high-temperature storage test Increase the temperature in the test chamber to 85 C when the unit under test is not connected to a power supply or other power source, and hold the temperature for 16 hours after the temperature stabilizes (follow the time specified in Table 2.27). Conduct appearance inspection and external sealing inspection of the unit under test without changing the temperature in the chamber and take the data as intermediate test data. Ground high-temperature continuous working test Increase the temperature in the test chamber to 70 C. After the temperature of the unit under test stabilizes (follow the time specified in Table 2.27), make the unit work continuously for 4 hours according to relevant standards and product specifications and record the temperature every 30 minutes in this process. After completing this step, check the performance of the unit according to relevant standards or product specifications under the same test conditions and take the data as intermediate test data. Ground high-temperature discontinuous working test Increase the temperature in the test chamber to 70 C. After the temperature of the unit under test stabilizes (follow the time specified in Table 2.27), make the unit

136 Chapter 2 complete four work cycles according to relevant standards and the product specification. Each work cycle lasts for 30 minutes and the temperature shall be recorded every 10 minutes in this process. After completing the first three work cycles, stop the supply of a power source for 15 minutes and then complete the fourth work cycle. After the fourth work cycle is completed, check the performance of the unit according to relevant standards or the product specification and take the data as intermediate test data. 9. Air high-temperature continuous working test Increase the temperature in the test chamber to 30 C. After the temperature of the unit under test stabilizes (follow the time specified in Table 2.27), adjust the pressure in the test chamber to that at an altitude of 12,000 m and then make the unit work continuously for 4 hours according to relevant standards and the product specification, recording the temperature every 30 minutes during this process. After completing this step, check the performance of the unit according to relevant standards or the product specification under the same test conditions and take the data as intermediate test data. 10. Air high-temperature discontinuous working test Increase the temperature in the test chamber to 47 C. After the temperature of the unit under test stabilizes (follow the time specified in Table 2.27), adjust the pressure in the test chamber to that at an altitude of 12,000 m and make the unit complete four work cycles according to relevant standards and the product specification. Each work cycle lasts for 30 minutes and the temperature shall be recorded every 10 minutes during this process. After completing the first three work cycles, stop the supply of the power source for 15 minutes and then complete the fourth work cycle. After the fourth work cycle is completed, check the performance of the unit according to relevant standards or the product specification and take the data as intermediate test data. 2.7.5.4 Judging criteria and results handling After all the 10 test procedures are completed according to the regulations above, take the unit under test out of the test chamber and make it reach a stable temperature under normal test atmospheric condition. Then, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally. The initial test data, intermediate test data, and final test data measured during the test shall conform to relevant standards or the product specification.

Unit test of the flight control system 137 2.7.5.5 Test example A horizontal stabilizer actuator is taken as an example to describe the temperature
altitude test. As an electromechanical product installed in the fairing at the top of the vertical tail, the horizontal stabilizer actuator should complete all of the 10 test procedures of the temperature
altitude test according to the regulations. The test items are shown in Table 2.32.

2.7.6 Temperature
humidity
altitude test 2.7.6.1 Test objectives and test requirements The temperature
humidity
altitude test aims to determine the adaptability of units to the environments of low temperature and low pressure and high temperature and high humidity during their operation. This test is applicable to electronic units and other units installed in aircraft that are not in a pressurized compartment and have no temperature control. It is mainly used to design mature and nonairtight units and other units with a shell cover. The temperature
humidity
altitude test includes two environments, that is, low temperature and low pressure, and high temperature and high humidity. The specific requirements are shown in Table 2.33. The pressure reduction rate shall be kept within the range of 3.5
5 kPa/min, the allowable temperature error is 6 2 C, the allowable pressure error is 6 5%, and the allowable error of relative humidity is 6 5% of measured value. 2.7.6.2 Test equipment and environmental requirements A temperature
humidity
altitude combined test chamber is used for the test. The test chamber (laboratory) should meet the requirements of the test conditions specified in Section 2.7.6.1 and have sensors or auxiliary instruments to monitor various test conditions. 2.7.6.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance, and then record the test data and take them as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. An appropriate spacing shall be kept between installed units and between the unit and test chamber wall, bottom and top, so that the air can circulate freely.

138 Chapter 2 Table 2.32: Examples of required test data for temperature
altitude test. No. 1 2

Test procedures

Initial test data

Intermediate test data

Final test data

Ground low-temperature storage test Ground low-temperature working test

Resistance, insulation resistance, working sealing, internal leakage, actuation speed, starting characteristic, position sensor output characteristic, pressure switch state

Working sealing

Resistance, insulation resistance, working sealing, internal leakage, actuation speed, starting characteristic, position sensor output characteristic, pressure switch state

3

Low-temperature lowpressure working test

4

Atmospheric pressure low-temperature frosting test

5

Normal-temperature atmospheric pressure working test

6

Ground hightemperature storage test Ground hightemperature continuous working test

7

8

Ground hightemperature discontinuous working test

9

Air high-temperature continuous working test

10

Air high-temperature discontinuous working test

Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state Working sealing, starting characteristic, position sensor output characteristic, pressure switch state

Unit test of the flight control system 139 Table 2.33: Requirements for temperature
humidity
altitude test. No. 1 2

Test conditions Low temperature and low pressure High temperature and high humidity

Temperature 

Altitude (atmospheric pressure)

255 C

12,200 m (18.8 kPa)

60 C

Ground

Humidity Normal test atmospheric conditions 95%

Start the test and the steps are as follows: 1. Drop the temperature in test chamber to 255 C in 2 hours. 2. Stabilize the temperature in the test chamber, reduce the pressure to 18.8 kPa at a rate of 3.5
5 kPa/min and stabilize it. This step takes 2.5 hours. 3. Restore the temperature and pressure in the test chamber to normal test atmospheric conditions within 30 minutes. 4. Keep the same temperature and pressure, increase the relative humidity in the test chamber to 95% and hold it for 2.5 hours. 5. Keep the relative humidity at 95%, and increase the temperature in the test chamber to 60 C within 30 minutes. 6. Keep the temperature (60 C) and relative humidity (95%) unchanged in the test chamber for 6 hours. 7. Keep the relative humidity (95%) unchanged, and reduce the temperature in the test chamber uniformly to that under normal test atmospheric condition within 8 hours. 8. Keep the relative humidity (95%) and the temperature under normal test atmospheric condition for 2 hours in the test chamber. 9. Repeat (1)
(8) for at least three times. 10. Restore the environment in the test chamber to normal test atmospheric conditions. 11. Drop the temperature in test chamber to 255 C within 2 hours. 12. Keep the same temperature in the test chamber, reduce the pressure in the test chamber to 18.8 kPa at a rate of 3.5
5 kPa/min and hold it. This step takes 2.5 hours. 13. Restore the environment in the test chamber to the normal test atmospheric condition in 30 minutes. After completion of the steps above, take the unit out from the chamber and make it reach a stable temperature under the normal test atmospheric condition. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as final test data and compare them with initial test data.

140 Chapter 2 2.7.6.4 Judging criteria and results handling In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally. The initial test data and final test data measured during the test shall conform to relevant standards or the product specification. The determination of the initial test data and final test data of the temperature
humidity
altitude test can refer to the temperature
altitude test.

2.7.7 Single event test A single event test shall be carried out for electronic units of the flight control system, such as the flight control computer and control panel. The flight control computer is taken as an example below to introduce the single event test of this kind of unit. 2.7.7.1 Test objectives and test requirements The test aims to get the relationship between single event upset cross section, latchup cross section of a device in the flight control computer, and incident ion LET (linear energy transfer) and measure the sensitivity of a device single event upset and latchup. A unit that has not undergone a total dose test should be selected with priority. If the unit has undergone a total dose test, the effect of the total dose effect on the single event effect shall be evaluated. When the number of incident particles reaches 107 ion/cm2, the effect of the total dose effect shall be taken into consideration. The operation of test personnel in a radiation source area should comply with requirements of Ref. [10]. Where there are provisions, measurement uncertainty analysis should be carried out according to Ref. [11]. 2.7.7.2 Test equipment and environmental requirements Equipment used in a single event effect test mainly includes a heavy ion accelerator, vacuum target chamber, temperature test unit, beam measurement system, single event effect test system, test circuit board, bias circuit, cable and switch system. The test environment shall be in accordance with the unified test requirements in Section 2.2.1, and the electrostatic protection shall meet the requirements in Ref. [12]. 2.7.7.3 Test items and test methods 1. Formulation of test scheme Before the radiation test, a test scheme covering the following content should be prepared: a. name, category, packaging type, manufacturer, and quantity of devices;

Unit test of the flight control system 141

2.

3.

4.

5.

b. determination of the category of the single event effect test, the type of radiation source, radiation source, and the unit to which the irradiation source belongs; c. determination of the fluence monitoring method; d. determination of the single event effect test scheme; e. design and layout of test circuit board; f. estimation of device LET threshold value and value range, determination of the type and energy of ions, ion selection sequence, and radiation time; g. estimated potential problems in test and measures taken for them; h. schedule of the radiation test; and i. specified judging criteria of the test (if applicable). Preparation of devices Unless otherwise specified, the units under test shall be randomly selected from the parent body and shall have the same packaging type. Each device shall be individually numbered so that it can be compared before and after radiation. For electrostatic sensitive devices, necessary antistatic protection measures shall be taken. If necessary, the device shall be unsealed before the test to remove the protective layer outside the chip passivation layer, the chip size shall be measured and recorded, and the characteristics of the chip shall be recorded by taking photos. After unsealing, the device shall be tested, and only those devices that pass the test can be sent for a followup test. Storage of test circuit board Place the connected test circuit board in the radiation field and the following requirements shall be followed: a. In the process of fixing the test circuit board in a mobile test board support, the test circuit board should be prevented from short circuit and shaking to ensure moving consistency between the test circuit board and the mobile support. b. The test circuit board and detector shall be placed on a mobile test board support in the vacuum target chamber. The device and the beam shall be aligned to ensure the beam can be incident upon the sensitive part(s) on the device surface. Debugging of test system Connect the test circuit board, single event effect test system, and related recording equipment. After applying and adjusting the voltage applied to the socket, checking and verifying the bias circuit is normal, insert the device and apply the specified bias voltage, run the single event effect test system, perform the regulated functional test, display the working state of the device, and check whether the device and test system work properly. Radiation test a. General rules The beam measurement system shall be calibrated according to regulations or on a regular basis. The beam uniformity data shall be obtained before the radiation test. In the radiation process, the test system monitors the single event effect of the tested sample

142 Chapter 2 and the beam measurement system monitors the fluence rate of incident particles in real time, and relevant test phenomena, test data, and radiation time are recorded. The single event latchup test should be conducted under rated working voltage. b. Selection of ion species Unless otherwise specified, the ion species specified in the test scheme and the selected effective LET data points shall not be less than 5. c. Test process The radiation shall be carried out according to the ion species, effective LET value, ion fluence rate, and radiation time determined in the test scheme. The beam measurement system shall monitor the fluence rate of incident ions in real time. SEU (single event upset) and SEL (single event latchup) test processes are shown in Figs. 2.11 and 2.12. Notes: In the test process, if the fluence of device reaches 107 ion/cm2 and there is still no single event effect, increase the effective LET value or change the test method. 2.7.7.4 Judging criteria and results handling The test results and recorded test results should meet the requirements of the product specifications or special technical conditions.

2.7.8 Solar radiation test The solar radiation test is used to evaluate the ability of units directly exposed to solar radiation within their life cycle to withstand the heat effect or photochemical effect caused by solar radiation during a hot season. Given the installation position of units on aircraft, units in a closed area that are not directly exposed to solar radiation do not need a solar radiation test, such as the angular rate gyroscope assembly and accelerometer arranged in the middle of the cargo hold. Whereas the units installed in the cockpit and close to the cockpit glass should have a solar radiation test, such as the primary flight control panel, automatic flight control device, steering wheel, and control handle. The primary flight control panel is taken as an example below to introduce the solar radiation test. 2.7.8.1 Test objectives and test requirements The solar radiation test is used to evaluate the ability of the primary flight control panel directly exposed to solar radiation to withstand the heat effect or photochemical effect caused by solar radiation. The solar radiation test requirements of the primary flight control panel shall follow Chapter 6 in Ref. [13].

Unit test of the flight control system 143

Figure 2.11 Single event upset test process.

In the cyclic heat effect test, the total radiation intensity shall follow the regulations in Section 2.1.1 of Ref. [14]; the maximum daily cycle temperature is 44 C, the temperature in the test chamber is 30 C, and the test cycle is 3
7 days. In a steady-state long-term photochemical effect test, the total radiation intensity shall follow the regulations in Section 2.1.1 of Ref. [14]; the test temperature is 49 C and the test cycle is 56 days.

144 Chapter 2

Figure 2.12 single event latchup test process.

2.7.8.2 Test equipment and environmental requirements Test equipment includes a test chamber, a base on which the test piece is installed, an auxiliary measuring instrument, and a solar radiation lamp. Unless otherwise specified, the environmental requirements of the test shall be subject to the general requirements in Section 2.2.1. 2.7.8.3 Test items and test methods The test contains two procedures: procedure I—cyclic test (heat effect); and procedure II— steady-state test (photochemical effect). The solar radiation test method of the primary flight control panel shall follow the requirements in Chapter 7 of Ref. [13].

Unit test of the flight control system 145 2.7.8.4 Judging criteria and results handling The test results and recorded test results shall meet the requirements of the product specifications or special technical conditions.

2.7.9 Rain test 2.7.9.1 Test objectives and test requirements The rain test aims to evaluate the ability of the shells of units to avoid rainwater infiltration and evaluate performance during or after the rain. Mechanical units mainly assess if there is lubrication failure due to rain. Mechanical and electrical and electronic units mainly assess the insulation performance of the electric circuit. The rain test has three procedures: rain test with air source, raindrop test, and waterproof test. The rain test with air source is applicable to units in area d without rain protection measures. The raindrop test is applicable to the unit body and units under the covering cap of the engine nacelle as well as units in area d with rain protection measures. The waterproof test is applicable to large units, units’ bodies, and the engine nacelle structure that cannot undergo rain test with air source. The area of the units is defined in Table 2.22. Requirements for the three test procedures are shown in Table 2.34. 2.7.9.2 Test equipment and environmental requirements The test chamber shall meet the test condition requirements specified in Section 2.7.7.1 and have sensors or auxiliary instruments to monitor various test conditions. 1. Rain test with air source The rain chamber shall have sufficient rainfall capacity, with adjustable rainfall speed throughout the test and a turntable used to install the tested unit. Raindrops in the rain chamber are produced by a sprinkler and the raindrops’ diameter is 0.5
4.5 mm. The sprinkler shall be designed to form water drops. Table 2.34: Requirements for rain test procedures. No.

Test procedure

Rainfall intensity

2

Rain test with air source Raindrop test

3

Waterproof test

No less than 15 cm/h No less than 10 cm/h No less than 15 cm/h

1

Raindrop diameter 0.5
5 mm

Test duration

0.5
4.5 mm

No shorter than 30 min for each rain exposure surface No shorter than 15 min

2
4.5 mm

No shorter than 40 min

146 Chapter 2 The air source shall ensure that it can blow the rain from 0 degree (horizontal) to 45 degrees uniformly to a side of the unit under test. 2. Raindrop test The water tank used for the test shall be able to provide water drops of 2 280130 0 L=m h. Water drops out from a distributor. The distributer shall be designed to ensure that water drops evenly and its design can refer to the schematic diagram in Ref. [15]. The arrangement of the test equipment shall ensure that all upward surfaces of the units under test can be simultaneously exposed to water drops in the test. 3. Waterproof test The test nozzle shall be made into a square spray lattice or other staggered lattice. At least one nozzle shall be installed 450
500 mm away from the surface of the unit under test every 0.55 m2. The minimum pressure of the nozzle is 375 kPa, and the raindrop diameter is 2
4.5 mm. 2.7.9.3 Test items and test methods 1. Rain test with air source Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance. Take the test data as initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. For equipment with a sealed cavity, it shall be heated to at least 10 C higher than the rainwater temperature at the beginning of every 30 minutes rain period to produce a negative pressure difference inside the unit under test to prevent leakage. During the test, all surfaces of the units under test which may be exposed to wind and rain in work shall be subject to the rain test with air source, and each surface shall be exposed to the rain for 30 minutes. Take the unit out from the chamber, use a sponge to remove the water on the outer surface of the unit slightly and then put it under normal test atmospheric conditions until its temperature stabilizes. Then, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as final test data and compare them with initial test data. 2. Raindrop test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1.

Unit test of the flight control system 147 Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance and take the test data as initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. When installing the unit, its working state required in the test process should be taken into account to reserve sufficient space and freedom for it to work. After the unit is installed, it should start to work and be checked. Any faults caused by inappropriate installation are not allowed. Make the unit under test work and be exposed to the rainwater dropping from a height of 1 6 0.1 m at a constant speed for at least 15 minutes. The water level of the distributor producing rainwater shall be 75 mm. Take the unit out from the chamber, use a sponge to remove the water on outer surface of the unit slightly and then put it under normal test atmospheric condition until its temperature stabilizes. Then, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as final test data and compare them with initial test data. 3. Waterproof test Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit. Check its appearance and take the test data as initial test data. Put the unit in a test chamber, simulate its actual use state for installation and connection and close all its entrances and vent holes. Spray water on all exposed surfaces of the unit under test for at least 40 minutes. Take the unit out from the chamber, use a sponge to remove the water on the outer surface of the unit slightly and then put it under normal test atmospheric condition until its temperature stabilizes. Then, test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as final test data and compare them with initial test data. 2.7.9.4 Judging criteria and results handling In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally.

148 Chapter 2 Table 2.35: Examples of required test data for rain test. Test procedure Raindrop test

Initial test data

Intermediate test data

Working voltage, working current, N/A transmission efficiency, static friction torque, driving moment, holding torque, braking time, brake releasing time, transmission shaft stiffness, manual brake releasing torque, insulation resistance, power supply characteristics

Final test data Working voltage, working current, transmission efficiency, static friction torque, driving moment, holding torque, braking time, brake releasing time, transmission shaft stiffness, manual brake releasing torque, insulation resistance, power supply characteristics

The initial test data and final test data measured during the test shall conform to relevant standards or the product specification. 2.7.9.5 Test example The flaps and slats WTB unit is taken as an example to introduce the rain test. The flaps and slats WTB unit is a mechanical and electrical product and it is installed near the wingtip on the rear beam. The raindrop test procedure in the rain test must be performed according to regulations and the test data required for the test are shown in Table 2.35.

2.7.10 Icing test 2.7.10.1 Test objectives and test requirements The icing test aims to determine units’ adaptability to icing environment resulting from rapid changes in temperature, altitude, and humidity. Mechanical and electrical units mainly investigate whether the movement of moving units is impeded. The icing test has three procedures: icing test, freezing test, and glaciation test. The icing test is applicable to units in area d whose surface is freezing under damp air whose temperature is higher than the freezing point after extremely low-temperature soakage. The freezing test is applicable to units in area d and e whose movement will be impeded or restricted by icing (such as primary control plane actuator, horizontal stabilizer actuator, spoiler actuator, torsion bar of high lift control system, variableangle brake, rotary actuator, roller screw mechanism). The glaciation test is applicable to the equipment in area d whose performance will be affected by the thickness of ice caused by freezing of free water and ponding on surface. This test procedure is usually used to verify the impact of the typical ice thickness on the performance of the unit and determine the maximum allowable ice thickness before enabling a deicing scheme. The area of the units is defined in Table 2.2.

Unit test of the flight control system 149 Table 2.36: Requirements for icing test procedures. Icing test Low temperature High temperature Altitude (atmospheric pressure) Relative humidity Cycle index



255 C 30 C Ground 95% 3

Freezing test 

220 C 30 C 12,200 m (18.8 kPa) 95% 25

Glaciation test 220 C N/A Ground N/A N/A

The three test procedures and requirements are shown in Table 2.36. The allowable temperature error is 6 2 C, the pressure error is 6 5%, and the relative humidity error is 6 5% of measured value. Pressure change rate should be kept in the range of 3.5
5 kPa/min and the speed of temperature rise should not be more than 3 C/min. 2.7.10.2 Test equipment and environmental requirements The test chamber shall meet the requirements of the test conditions specified in Table 2.36 and have sensors or auxiliary instruments to monitor various test conditions. For the icing test procedure, as it needs to quickly switch between high-temperature humid environment and low-temperature dry environment, it is recommended to use two independent test chambers to create two different environments. 2.7.10.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance. Take the test data as the initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. Remove the atypical contaminants such as oil, grease, and dust on the surface of the unit. 1. Icing test a. Drop the temperature in the test chamber to 255 C, keep the atmospheric pressure and relative humidity in the chamber under normal atmospheric conditions and hold the temperature for the time specified in Table 2.26. b. After the heat preservation, put the unit under test in a test chamber at a temperature of 30 C and relative humidity of 95% as soon as possible and monitor the surface temperature of the unit under test.

150 Chapter 2 c. Keep the temperature and relative humidity in the test chamber unchanged until the surface temperature of the unit under test reaches 3 C
7 C. Then, put the unit under test in a test chamber with a temperature of 255 C and atmospheric pressure and relative humidity under normal atmospheric conditions as soon as possible and hold the temperature for the time specified in Table 2.26. Repeat steps (a) to (c) for three times. After the heat preservation in the third step (c) is completed, increase and hold the temperature in the test chamber to 210 C. When the surface temperature of the unit under test reaches 215 C to 5 C, make the unit under test work and test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. 2. Freezing test a. Drop the temperature in the test chamber to 220 C, keep the atmospheric pressure in the chamber under normal atmospheric conditions, and hold the temperature for the time specified in Table 2.26. Hold the temperature in the test chamber and drop the atmospheric pressure in the test chamber to 18.8 kPa at a rate of 3.5
5 kPa/min. Hold the temperature and pressure for at least 10 minutes. b. Increase the temperature in the test chamber at a rate not exceeding 3 C/min and increase the relative humidity in the test chamber to no less than 95% at the same time, and keep it. Maintain the environment until the ice and frost on the surface of the unit under test fully melts or the surface temperature of the unit under test reaches 0 C
5 C. In this process, the temperature in the test chamber shall not exceed 30 C. c. Increase the atmospheric pressure in the test chamber to normal atmospheric conditions at a constant speed and control the time for the rise pressure to within 15
30 minutes. After the rise in pressure, reduce the relative humidity in the test chamber to normal atmospheric conditions. d. Drop the temperature in the test chamber to 220 C, keep the atmospheric pressure in the chamber under normal atmospheric conditions, and hold the temperature for the time specified in Table 2.26. Repeat steps (a) to (d) for 25 times. After the last step (d) is completed, make the test sample work and test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. 3. Glaciation test Put the unit under test in a low-temperature environment and spray water mist on it to make it ice up. The temperature of the water mist should be close to freezing point. The ice must be clear and solid and it should contain no impurities or bubbles. The recommended freezing temperature is 210 C to 21 C and it should be determined better through the test.

Unit test of the flight control system 151 The thickness of the ice shall be uniform and subject to the special specification of the unit. When the thickness of the ice reaches the required level, stop spraying water, make the unit under test work, test its electrical performance, mechanical performance, and other performances and check its appearance. Take the test data as the final test data and compare them with the initial test data. 2.7.10.4 Judging criteria and results handling In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally. The initial test data and final test data measured during the test shall conform to relevant standards or product specification. 2.7.10.5 Test example The flaps roller screw mechanism is taken as an example below to introduce the icing test. As a mechanical product, the flaps roller screw mechanism is installed in the flaps slide rail fairing. According to regulations, the freezing test procedure is required in the icing test and content that should be monitored in the test is shown in Table 2.37.

2.7.11 Damp heat test 2.7.11.1 Test objectives and test requirements The damp heat test aims to evaluate the adaptability of units to a high-temperature and high-humidity environment and it mainly assesses the corrosion resistance of units. The damp heat test has one procedure: the damp heat test. The damp heat test is applicable to all units. Test procedures and requirements are shown in Table 2.38. The allowable temperature error is 6 2 C and the relative humidity error is 6 5% of measured value. Table 2.37: Examples of required test data for icing test. Test procedures Freezing test

Initial test data Axial load, transmission efficiency, no-load accuracy, operation stability, static friction torque

Intermediate test data N/A

Final test data Axial load, transmission efficiency, no-load accuracy, operation stability, static friction torque

152 Chapter 2 Table 2.38: Requirements for damp heat test procedure. Test stage High-temperature stage

Temperature

Relative humidity

Test duration



95%

10 3 24 h

60 C

2.7.11.2 Test equipment and environmental requirements The test chamber shall meet the requirements of the test conditions specified in Table 2.38 and have sensors or auxiliary instruments to monitor various test conditions. The structure of the test chamber and the way of placing accessories shall prevent the condensed water from dropping onto the unit under test. The condensed water in the test chamber shall be continuously discharged. The test chamber shall have good air ventilation and the wind speed in the working space shall be 0.5
2 m/s. The test chamber shall be provided with a well-insulated binding post or a cable access device for testing the unit under test. The test chamber shall be provided with a lighting device and an observation hole (window) for the purpose of observing the unit under test and the temperature and humidity in the chamber. The water for humidification should be distilled water or deionized water, with resistivity being not less than 500 Ω m. In addition to water, pollutants that may cause rusting or corrosion or other substances should not be introduced into the test chamber. 2.7.11.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance. Take the test data as initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. The test takes 24 hours as a cycle and each cycle has four stages: temperature rise, high temperature and high humidity, temperature drop, low temperature and high humidity. The specific steps are as follows: 1. In 2 hours, increase the temperature to 60 C and the relatively humidity to 95% in the test chamber. The temperature and humidity shall be controlled to ensure condensation on surface of the unit under test. 2. Maintain the temperature at 60 C and relative humidity at 95% for at least 6 hours.

Unit test of the flight control system 153 3. After 8 hours, drop the temperature in the test chamber to 30 C. In this process, the relative humidity in the test chamber shall be maintained above 85%. 4. When the temperature in the test chamber reaches 30 C, the relative humidity should be 95% and maintain this condition for 8 hours. Repeat steps (1) to (4) for 10 times. Before the end of step (4) in the fifth cycle and 10th cycle and when the unit under test is under the temperature of 30 C and relative humidity of 95%, make the unit under test work and test its performance according to relevant standards or product specifications. Record the results and take them as intermediate test data. After the 10 cycles are completed, restore the temperature and humidity in the test chamber to normal test atmospheric conditions and make the temperature of the unit under test become stable in the test chamber. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as final test data and compare them with initial test data. 2.7.11.4 Judging criteria and results handling In the test process, the appearance of the unit under test should be free from bubbles, wrinkling, pitting, unsoldering, loosening, peeling, permanent deformation, and structural damage. Loss of certain surface gloss is allowed. The initial test data, intermediate test data, and final test data measured during the test shall conform to relevant standards or the product specification. 2.7.11.5 Test example The flaps and slats WTB unit is taken as an example below to introduce the damp heat test. As a mechanical and electrical product, the flaps and slats WTB unit is installed on the rear beam near the wingtip. According to regulations, the damp heat test must be conducted. Items that should be tested are shown in Table 2.39.

2.7.12 Mold test 2.7.12.1 Test objectives and test requirements The mold test aims to evaluate the antimold ability of nonmetal materials used in units. The mold test has one test procedure: mold test. The mold test is applicable to all units. The mold test is undertaken for electrical, electronic, optical, and communication units of critical importance (such as primary flight control computer, actuator controller, flaps and slats controller, automatic flight control computer), and appearance inspection, and a

154 Chapter 2 Table 2.39: Examples of required test data for damp heat test. Test procedures

Initial test data

Damp heat Working voltage, working test current, transmission efficiency, static friction torque, driving torque, control torque, braking time, brake release time, transmission shaft stiffness, manual brake release torque, insulation resistance, power supply characteristic

Intermediate test data

Final test data

Working voltage, working current, insulation resistance, power supply characteristic

Working voltage, working current, transmission efficiency, static friction torque, driving torque, control torque, braking time, brake release time, transmission shaft stiffness, manual brake release torque, insulation resistance, power supply characteristic

Table 2.40: Strains used in mold test. No. 1 2 3 4 5

Strain name

Strain no.*

Aspergillus niger Aspergillus flavus Aspergillus versicolor Penicillium funicalosum Chaetomium globosum

3.3928 3.3950 3.3885 3.3872 3.4254

Notes: The *strain no. is the strain no. preserved by the Beijing Institute of Microbiology, Chinese Academy of Sciences.

performance test shall be carried out for them. The test cycle is 84 days. For the mold test for other devices, only the appearance inspection is required and the test cycle is 28 days. The test is conducted under alternate changes of temperature and humidity and it cycles every 24 hours. In the first 20 hours, keep the temperature at 30 C 6 1 C and relative humidity at 95% 6 5%. In later 4 hours, keep the temperature at 25 C 6 1 C and relative humidity at 95% 1 10% for at least 2 hours and the longest time for temperature and humidity change is 2 hours. In the change period, keep the temperature at 24 C
31 C and relative humidity at no less than 90%. The strains used in the test are shown in Table 2.40. Before the test, the strains for the test shall be tested one by one, and strains that are not pure, variants, or beyond the culture or preservation time shall not be used. According to the needs of the unit under test, in addition to the five bacterial strains listed in Table 2.40, strains that have been proven to cause corrosion to the product can be added, but they must be specified in the test record. 2.7.12.2 Test equipment and environmental requirements The test chamber shall meet the requirements of Section 2.7.12.1 and it shall have auxiliary temperature and humidity detection devices and automatic continuous recording devices inside.

Unit test of the flight control system 155 The test chamber shall be set with air holes and ventilation devices to prevent an air pressure increase in the chamber. During the ventilation, the temperature in the chamber shall not be lower than 24 C and the relative humidity shall not be lower than 80%. The air speed in the working space of the test chamber shall be controlled within the range of 0.5
2 m/s. The water used in the test chamber should be distilled water and its resistivity should not be less than 500 Ω m. Steam should not be directly introduced into the chamber for use and rust or other pollutants of the test unit should not be brought to the surface of the unit under test. Units passing the salt spray test and sand and dust test shall not be used in this test. Generally, the unit under test will not be cleaned. If relevant standards require cleaning, units under test should be added for comparison. The cleaning of the unit under test shall be completed 72 hours before the spraying of spore suspension. When cleaning the unit under test, avoid contaminating the surface of the unit. 2.7.12.3 Test items and test methods The inorganic salt solution, spore suspension, and reference sample shall be prepared according to the requirements of Ref. [16] and spore activity shall be tested. Before being tested, the unit shall have appearance inspection and surface contamination. Defects and other conditions that may contribute to mold growth shall be paid with special attention. Functional and performance testing of the unit under test should be carried out according to technical conditions of the unit before the test (if required). Put the unit under test on a sample rack of the mold test chamber according to the storage, transportation, or working condition. The contact area between the unit under test and the sample rack shall be as small as possible. The air around the sample should be freely circulating. Put the reference sample close to the unit under test vertically, but it should be in contact with the unit. After the unit and reference sample are pretreated for 4 hours according to the conditions specified in Section 2.7.12.1, strains can be applied. The unit under test and the reference sample shall be applied with strains at the same time. Use a sprayer or other similar atomizing device to spray prepared mixed spore suspension on possibly exposed inner and outer surfaces of the unit under test. After the unit under test and the reference sample are applied with strains, keep them under the conditions specified in Section 2.7.12.2 in the test chamber for 7 days and check the mold growing condition of the reference sample. The mold coverage area on the surface should not be less than 90%, otherwise the test is invalid. If the mold growing condition on the surface of the reference sample meets the requirements, the test time shall be calculated from the date of the strains’ application.

156 Chapter 2 During the test, the mold test chamber shall be ventilated every 7 days, preferably at a time when the temperature and humidity have alternated. During the ventilation, the temperature and humidity shall conform to Section 2.7.12.2 and the total amount of air change shall be one fifth of the chamber’s volume. After the test, immediately check the mold growing condition on the surface of the unit under test, mainly by visual inspection. If necessary, observe with a magnifying glass. Record the growth site, covering area, color, growth form, growth density, and growth thickness of mold. If necessary, take photos of them. The growth of nontest bacteria is allowed on the surface of the unit under test and it should also be evaluated for test results. After the appearance inspection of the unit under test, conduct a performance test according to relevant standards or the specification. Record the final test data and compare them with the initial test data. 2.7.12.4 Judging criteria and results handling The mold growing condition on the surface of the unit under test shall be evaluated according to Table 2.41. After the mold test is completed, the mold growing condition shall be superior to class 2 (classes 0, 1, 2 are acceptable).

2.7.13 Salt spray test 2.7.13.1 Test objectives and test requirements The salt spray test aims to evaluate the ability of units to resist the influence of a salt spray atmosphere. It mainly assesses the corrosion resistance of the coating and surface treatment layer of units in a salt spray atmosphere. Table 2.41: Classes of mold growing condition. Class

Mold growing degree

0 1

No mold Micro growth

2 3 4

Mold growing condition

No mold growing Rare or limited mold growth and reproduction. The mold growth area is less than 10% of the total area of the unit under test and the matrix is rarely used or not destroyed. Chemical, physical, and structural changes are hardly observed Slight growth Mold colonies spread intermittently or loosely on the surface of the matrix. The mold growth area accounts for less than 30% of the total area and the mold reproduces moderately Medium growth The mold grows and reproduces in large quantities and accounts for less than 70% of the total area. Chemical, physical and structural changes are found on the surface of matrix Serious growth The mold grows and reproduces in large quantities and accounts for over 70% of the total area. The matrix is decomposed or rapidly deteriorated

Unit test of the flight control system 157 The salt spray test has one procedure: the salt spray test. It is applicable to all units. Salt solution is prepared with chemically pure sodium chloride and distilled water or deionized water with resistivity not lower than 50,000 Ω cm. Salt solution with sodium chloride content at 4%
6% is prepared with 5 units of sodium chloride and 95 units of water. At least two collectors should be used generally, with one arranged near a nozzle and one away from all nozzles. The collector shall not be sheltered by the unit under test and the collected liquid of the unit and other objects shall not drip into the collector. All collected liquid can be collected together to measure their sedimentation rate and pH value under 35 C. In the effective space, test a clean collector at any position. The time for continuous spray collection shall be at least 16 hours and the average salt spray sedimentation in every 80 cm2 horizontal collection area (diameter 10 cm) shall be 1
2 mL/h. The collected liquid after spraying should have a pH value of 6.5
7.2 under 35 C. Before the test, the test chamber shall have a no-load test run with a continuous spray time of 16
24 hours. When it is confirmed that stable test conditions can be maintained, the unit can be sent for test. A test chamber used within 5 days may not have a no-load test run. The temperature in the effective space of the test shall be 35 C. The time that the unit under test bears continuous spray shall be 48 hours or subject to relevant standards or technical documents. The longer time should be used. 2.7.13.2 Test equipment and environmental requirements The materials of the test chamber and its accessories shall be resistant to salt spray corrosion and shall not affect the test results. The test chamber shall have reasonable and proper exhaust holes to prevent a pressure difference affecting the uniform distribution of salt mist. The structure of the test chamber shall be strong and durable with a large volume. Salt mist shall not be directly sprayed on the unit under test. The liquid on the test chamber and its accessories shall not drip on the test sample. The salt mist should be settled evenly on the unit under test and the salt solution that has been in contact with the unit shall not be recycled for use. The atomizer shall be corrosion and wear resistant without deformation and with good interchangeability. It should be able to produce fine, moist, and evenly dispersed salt mist. The compressed air used for spraying shall be free of impurities, oil dirt, and can be heated and humidified stably. Spray pressure should be as low as possible to meet the requirement of spraying at a desired rate. 2.7.13.3 Test items and test methods Use solvent that will not cause corrosion or a protective film to remove the dirt or temporary protective layer on the surface of the unit under test until there is no moisture on the surface. Organic coating shall not be cleaned with organic solvent and end faces and

158 Chapter 2 contact surfaces that do not need coating shall be coated with a wax layer or other substances for protection. Under normal atmospheric conditions, the appearance and performance of the unit under test shall be checked and the results shall be taken as initial test data. Appearance inspection items of the unit include metal surface, coating, surface treatment layer for corrosion prevention, multimetal contact area, uncoated electronic unit and circuit, mechanical system prone to failures, electrical insulation, thermal insulation, high stress area, concentrated salt solution area, and gap. Put the units under test in a test chamber and the average angle between the units under test with the vertical plane of the test chamber shall be 15
30 degrees. The units under test shall not touch or shelter each other, nor contact other metals or water-absorbing materials. Concentrated liquid drops shall not drip from one unit to another. The spacing between the units shall ensure that the salt mist is settled freely on the tested surface of each unit. Adjust the temperature of the test chamber to 35 C. After the temperature of the unit under test stabilizes for at least 2 hours, start spraying. The salt mist shall be sprayed for continuously 48 hours (or subject to the time in relevant standards or technical documents, it shall not be less than 48 hours) and the salt mist sedimentation rate and pH value should be checked every 12 hours. After the test, put the unit under test under normal atmospheric conditions for 48 hours for recovery and drying. If required by appearance inspection, the unit shall be flushed with flowing water with temperature not exceeding 38 C. Clean compressed air can be used to blow off water drops. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. 2.7.13.4 Judging criteria and results handling Loss of certain surface gloss and local corrosion of the protective layer on the outer surface of the unit under test are allowed. The paint layer shall be free from bubbles or fall-off and the base metal shall not have corrosion. Nonmetal shall not have obvious bubbles or expansion and serious corrosion is not allowed at the contact between metals. Measured initial test data in the test process and final test data shall conform to relevant standards or product specifications. 2.7.13.5 Test example The flaps and slats WTB unit is taken as an example to introduce the salt spray test. As a mechanical and electrical product, the flaps and slats WTB unit is installed on the rear beam near the wingtip. According to regulations, the salt spray test is required. Items to be tested are shown in Table 2.42.

Unit test of the flight control system 159 Table 2.42: Examples of required test data for salt spray test. Test procedure Salt spray test

Initial test data

Intermediate test data

Final test data

Working voltage, working current, transmission efficiency, static friction torque, driving torque, control torque, braking time, brake release time, transmission shaft stiffness, manual brake release torque, insulation resistance, power supply characteristic

Working voltage, working current, insulation resistance, power supply characteristic

Working voltage, working current, transmission efficiency, static friction torque, driving torque, control torque, braking time, brake release time, transmission shaft stiffness, manual brake release torque, insulation resistance, power supply characteristic

2.7.14 Sand and dust test 2.7.14.1 Test objectives and test requirements The sand and dust test aims to evaluate adaptability of the unit to a sandy and dusty environment. The sand and dust test has two procedures: blowing dust test and blowing sand test. The blowing dust test mainly assesses the ability of the unit’s shell to resist dust penetration. The blowing sand test mainly assesses the ability of units to resist particle abrasion, ensure motion stability, and resist pipeline blocking. The blowing dust test and blowing sand test are applicable to units in area d, exposed units in area b after the cargo hold door is opened, and also the exposed units in area e after control plane deflection or laying-off. The area of the units is defined in Table 2.22. The dust particles used for the blowing dust test are angular silica powder. Its silica content is 97%
99% by mass and it is composed of the following sizes: 1. 2. 3. 4.

100% passes 150 μm sieve. 96%
100% passes 106 μm sieve. 88%
92% passes 75 μm sieve. 73%
77% passes 45 μm sieve.

The sand particles used for the blowing sand test are subangular quartz sand. The silica content is above 95% by mass, average roundness factor is 0.2, and the Moh’s hardness is 7. The sand is composed of the following sizes: 1. 2. 3. 4.

0.5%
1.5% fails to pass 850 μm sieve. 1.2%
2.2% fails to pass 600 μm sieve. 13.8%
15.8% fails to pass 425 μm sieve. 36%
38% fails to pass 300 μm sieve.

160 Chapter 2 5. 27.6%
29.6% fails to pass 212 μm sieve. 6. 11.7%
13.7% fails to pass 150 μm sieve. 7. 4.2%
6.2% fails to pass 150 μm sieve. The two test procedures and requirements are shown in Table 2.43 and the allowable temperature error is 6 2 C. 2.7.14.2 Test equipment and environmental requirements The test chamber shall have good sealing performance, the cross-sectional area of the working space shall be two times greater than the cross-sectional area of the unit under test, and the effective volume shall be 3.3 times greater than the volume of the unit under test. The test chamber shall have instruments that can measure and control the sand and dust concentration, air speed, temperature, relative humidity, and other auxiliary devices. Before the dusty air in the test chamber is applied to the unit under test, approximate laminar flow is allowed. A sand separator shall be set in the test chamber to ensure that the fan can circulate the air repeatedly when there is no sand. 2.7.14.3 Test items and test methods Make the unit under test reach a stable temperature under the normal test atmospheric conditions specified in Section 2.2.1. Under normal test atmospheric conditions, test the electrical performance, mechanical performance, and other performances of the unit and check its appearance. Take the test data as initial test data. Put the unit in a test chamber and simulate its actual use state for installation and connection. Keep intact the plugs, enclosures, and test plates used in actual work but not in the test and protect the mechanical and electrical connection part in actual work with a covering. Make the most critical and weakest surface of the unit face the dust airflow direction. 1. Blowing dust test Control the temperature in the test chamber to 23 C, relative humidity to less than 30%, adjust the air speed to 8.9 6 1.2 m/s, adjust the dust particle input volume, make the blowing dust concentration be 10.6 6 7 g/m3, and maintain the conditions above for 6 hours. Table 2.43: Requirements for sand and dust test procedures. Test procedures Blowing dust test Blowing sand test

No. Temperature 1 2 3

23 C 60 C 60 C 60 C

Relative humidity

Air speed

Sand and dust concentration

Test duration

,30% N/A N/A ,30%

8.9 6 1.2 m/s 1.5 6 1 m/s 8.9 6 1.2 m/s 18
29 m/s

10.6 6 7 g/m3 N/A 10.6 6 7 g/m3 2.2 6 0.5 g/m3

6h 16 h 6h 1.5 h

Unit test of the flight control system 161 Stop inputting dust particles and adjust the air speed to 1.5 6 1 m/s, increase the temperature in the test chamber to 60 C, keep the heating rate to less than 10 C/min, and maintain these conditions for 16 hours. Keep the temperature in the test chamber at 60 C, adjust the air speed to 8.9 6 1.2 m/s, continue to input dust particles, control blowing dust concentration to 10.6 6 7 g/m3, and maintain the conditions above for 6 hours. Stop the test chamber and restore the temperature in the test chamber to normal test atmospheric conditions. Take out the unit under test, shake off or brush off the dust particles on the unit, pay attention to avoid other sand and dust falling on the unit, do not use a hair dryer or vacuum cleaner to remove the dust particles on the unit. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. In the appearance inspection, check if there is dust accumulation at the bearing, grease seal, and lubrication parts of the unit and also record the position of the dust accumulation and accumulation volume. 2. Blowing sand test Control the temperature in the test chamber at 60 C, heating rate at less than  10 C/min, and relative humidity less than 30%, and make the unit under test reach a stable temperature. Adjust the air speed to 18
29 m/s, adjust the sand feeder to ensure the blowing sand concentration is 2.2 6 0.5 g/m3, and maintain the conditions above for 1.5 hours. Change the installation direction of the test sample to make the next key (or weak) surface of the unit under test face the sand blowing airflow. Repeat the steps above until all key (or weak) surfaces of the test sample are tested. If it is required to start the unit under test in the blowing sand test, the unit shall start in the last hour of the test. Check the working performance of the test sample according to relevant standards or product specifications and record the data as intermediate test data. Stop the test chamber and restore the temperature in the test chamber to normal test atmospheric conditions. Take out the unit under test, shake off or brush off the dust particles on the unit, avoid other sand and dust falling on the unit, do not use a hair dryer or vacuum cleaner to remove the dust particles on the unit. Test the electrical performance, mechanical performance, and other performances of the unit and check the appearance of the unit according to relevant standards or the product specification. Take the test data as the final test data and compare them with the initial test data. In the appearance inspection, attention should be paid to the abrasion and blocking caused by sand particles as well as the penetration of sand and dust.

162 Chapter 2 2.7.14.4 Judging criteria and results handling In the test process, the appearance of the unit under test shall not have obvious deformation or damage and the test sample shall not lose its function or work abnormally. The initial test data and final test data measured during the test shall conform to relevant standards or product specifications. 2.7.14.5 Test example An aileron mechanical backup actuator is taken as an example to introduce the sand and dust test. As a mechanical and electrical product, the aileron mechanical backup actuator is installed on the rear beam near the wingtip. According to regulations, the blowing sand test is required for the sand and dust test and the items to be tested are shown in Table 2.44.

2.8 Electromagnetic environment protection test The units of the flight control system include mechanical and electrical and electronic units, and adopt the computer and bus transmission method, which will make the system or units more sensitive to electromagnetic interference, and require an electromagnetic environment protection test. The electromagnetic environment protection test of the flight control system is an effective measure to test the electromagnetic interference resistance of units of the flight control system and an effective means to ensure aircraft safety. The electromagnetic environment protection test includes the electromagnetic emission and susceptibility test, lightning direct effect test, lightning-induced transient susceptibility test, and high-intensity radiated field protection test.

2.8.1 Electromagnetic emission and susceptibility test All electronic, mechanical, and electrical units of the flight control system shall undergo the electromagnetic emission and susceptibility test and only their test items are different. The selection of test items should follow the working principle, power supply mode, and Table 2.44: Examples of required test data for sand and dust test. Test procedures Blowing sand test

Initial test data Working tightness, resistance, insulation resistance, null voltage, insensitive current, zero-bias current, motion stability

Intermediate test data N/A

Final test data Working tightness, resistance, insulation resistance, null voltage, insensitive current, zero-bias current, motion stability

Unit test of the flight control system 163 sensitive source of the units. The specific deletion principles are shown in Table 2.47. The test items shall comply with the provisions of Ref. [17] and the technical agreements of the finished products. Unless otherwise specified in the technical documents, the test items generally selected for the units of the flight control system are shown in Table 2.45. 2.8.1.1 Test objectives and test requirements The electromagnetic emission and susceptibility test aims to evaluate if the flight control units or mechanical and electrical units meet the requirements of Ref. [17]. The test can quantitatively check the electromagnetic compatibility of units and help find sensitive components to eliminate incompatibility. The configurations, precautions, and general requirements for the electromagnetic emission and susceptibility test shall follow Sections 3.1.1, 4.6, and 4 of Ref. [18]. 2.8.1.2 Test equipment and environmental requirements Test equipment and environmental requirements shall follow the requirements of Ref. [18]. The electromagnetic emission and electromagnetic sensitivity test should be completed in the electromagnetic compatibility laboratory and the electromagnetic compatibility laboratory shall pass the quality certification of “China National Accreditation Service for Conformity Assessment (CNAS)” and “Defense Science and Technology Industry Laboratory Accreditation Committee (DILAC)” to have the ability to test electromagnetic emission, electromagnetic sensitivity, and other items. Test equipment such as the tester used in the test to monitor and record the working state and indicators of the unit shall be provided by the unit development unit. Moreover, the tester itself shall have electromagnetic protection measures to withstand the test environment. Table 2.45: Electromagnetic effect test items of flight control system units. No. 1 2 3 4 5 6 7 8

Finished products

Electromagnetic effect test items

Computer Actuator Angular rate gyroscope assembly Acceleration sensor LVDT/RVDT Handle and switch units Control panel units Trim units

CE102, CS101, CS106, CS114, CS115, CS116, RE102, RS103 CS114, CS115, CS116, RE102, RS103 CE102, CS101, CS106, CS114, CS115, CS116, RE102, RS103 CE102, CS101, CS106, CS114, CS115, CS116, RE102, RS103 CS114, CS115, CS116, RE102, RS103 CE102, RE102 (made with the system) CE102, CS101, CS106, CS114, CS115, CS116, RE102, RS103, CE107 CE102, CS101, CS106, CS114, CS115, CS116, RE102, RS103

164 Chapter 2 The test cable shall be consistent with the installed cable in model, protection mode, and ending mode. Cables longer than 10 m can be cut into 10 m cables, and for cables shorter than 10 m, the test cable shall be made according to the length of installed cable. 2.8.1.3 Test items and test methods The content of the electromagnetic emission and susceptibility test includes conducted emission, conducted susceptibility, radiated emission and radiated susceptibility. The test items are shown in Table 2.46. The test items shall be deleted according to the technical agreement, technical requirements, or the model specification. The application range of the test items is shown in Table 2.47. Test methods of electromagnetic emission and susceptibility test shall follow the requirements of Ref. [18]. 2.8.1.4 Judging criteria and results handling The test data and results are directly output by the unit under test. The requirements for curve processing are as follows: 1. The curve chart of the amplitude and frequency shall be automatically drawn by means of X
Y axis output mode. Except for verifying the curve chart, manually collected data will not be accepted. 2. Each curve chart shall show applicable limit values. Table 2.46: Items of electromagnetic emission and susceptibility test. No. Items 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19

Name

CE101 25 Hz
10 kHz power cable conducted emission CE102 10 kHz
10 MHz power cable conducted emission CE106 10 kHz
40 GHz antenna terminal conducted emission CE107 Power cable spike signal (time domain) conducted emission CS101 25 Hz
50 kHz power cable conducted susceptibility CS103 15 kHz
10 GHz antenna terminal intermodulation conducted susceptibility CS104 25 Hz
20 GHz antenna terminal undesired signal curbing conducted susceptibility CS105 25 Hz
20 GHz antenna terminal intermodulation conducted susceptibility CS106 Power cable spike signal conducted susceptibility CS109 50 Hz
100 kHz shell current conducted susceptibility CS114 10 kHz
400 MHz cable bundle injection conducted susceptibility CS115 Cable bundle injection pulsed excitation conducted susceptibility CS116 10 kHz
100 MHz cable and power cable damped sinusoid transient conducted susceptibility RE101 25 Hz
100 kHz magnetic field radiated emission RE102 10 kHz
18 GHz electric field radiated emission RE103 10 kHz
18 GHz antenna harmonic wave and spurious output radiated emission RS101 25 Hz
100 kHz magnetic field radiated susceptibility RS103 10 kHz
18 GHz electric field radiated susceptibility RS105 Transient electromagnetic field radiated susceptibility

Unit test of the flight control system 165 Table 2.47: Application range of electromagnetic emission and susceptibility test items. No. Items

Application range

1

CE101

2

CE102

3

CE106

4 5

CE107 CS101

6

CS103

7

CS104

8

CS105

9 10 11

CS106 CS109 CS114 Within the frequency range of 10 kHz
400 MHz, this item is applicable to all interconnected cables of units and systems, including power cables CS115 This item is applicable to all interconnected cables of units or systems, including power cables CS116 This item is applicable to all interconnected cables including power cables and single power conductors, but it is unnecessary to conduct return cable test independently RE101 Applicable to radiated emission of all unit or system shells and all interconnected cables, not applicable to the base frequency of communication transmitter RE102 This items is applicable to radiated emission of unit or system shells and all interconnected cables, not applicable to the base frequency of transmitter or radiation of antenna, with a frequency range of 2 MHz
18 GHz (test frequency upper limit up to 1 GHz or 10 times of the maximum working frequency of EUT, the greater one will be used) RE103 If the test has a transmitter with fixed antenna, this item can be used to substitute CE106. CE106 requirement is preferred unless unit and system design characteristics interfere with its use. This requirement is not applicable to EUT required bandwidth or 6 5% base frequency (take the wider one). According to the working frequency range of EUT, the starting frequency of the test is as follows: 10 kHz
3 MHz, 10 kHz, 3
300 MHz, 100 kHz, 300 MHz
3 GHz, 1 MHz, 3 GHz
18 GHz, 10 MHz. The frequency upper limit of the test is 18 GHz or 20 times the maximum working frequency of EUT (the lower one is used). For units applying a waveguide, this item is not applicable when the frequency is lower than 0.8 times the waveguide cutoff frequency RS101 This item is applicable to unit and system shells and all interconnected cables, not applicable to EUT antennas. For units installed on naval aircraft, it is only applicable to ASW aircraft

12 13 14 15

16

17

It is applicable to AC and DC power cables of units (including return cables, but excluding conductors at output end of EUT power supply) It is applicable to all power conductors (including return cables) except for the conductors at output end of the power supply of the equipment under test (EUT), with frequency range of 10 kHz
10 MHz It is applicable to EUT transmitters and receiver antenna terminals, not applicable to units with nonremovable fixed antennas It is applicable to AC and DC power cables for equipment and systems on aircraft It is applicable to AC and DC input power cables for units and systems, excluding return cables. If EUT works under DC power supply, it is suitable for 25 Hz
50 kHz. If EUT works under AC power supply, it is suitable to start from the second harmonic of EUT power supply frequency to 50 kHz This receiver front-end susceptibility requirement is applicable to receiver units and systems, such as communication receivers, radio-frequency amplifiers, transceivers, radar receivers, and electronic countermeasures receivers This receiver front-end susceptibility requirement is applicable to receiver units and systems, communication receivers, radio-frequency amplifiers, transceivers, radar receivers, and electronic countermeasures receivers This receiver front-end susceptibility requirement is applicable to receiver units and subsystems, communication receivers, radio-frequency amplifiers, transceivers, radar receivers, and electronic countermeasures receivers This item is applicable to all EUT power cables

(Continued)

166 Chapter 2 Table 2.47: (Continued) No. Items 18

RS103

19

RS105

Application range This requirement is applicable to unit and system shells and all interconnected cables, with an applicable frequency range of 10KHz
18 GHz When the unit or system is located externally on a reinforced (shielded) platform or facility, this item is applicable to unit and subsystem shells. It is also applicable to key units and subsystems externally installed on army aircraft for safety purposes

3. Each curve should have a frequency resolution of at least 1% or twice the width of the measuring receiver and an amplitude resolution of at least 1 dB. 4. Two sets of curves: EUT test curve and test system calibration and inspection curve, shall also be provided. The products that meet the specified limit value requirements after the test are conforming products. The electromagnetic compatibility test task description of units shall specify the acceptance/ failure criteria for units. In the procedure of the electromagnetic compatibility test, detailed working parameters and methods for monitoring these parameters shall be defined. The basic judging criteria for acceptance are as follows: 1. About disaster When and after the unit is exposed in an electromagnetic environment or when and after it is applied with single strike, multiple strikes, or burst transient voltage, its function shall be kept valid. In the event of any system interruption, the continued performance and effectiveness of the functions of the primary or standby units shall be ensured, which shall be assessed and approved by the chief engineer system and the user. The affected units shall not provide misleading information and shall be able to recover automatically after the test signal is removed. 2. About danger and large effects When the unit is in a specified electromagnetic compatibility test, it must not be adversely affected. When the unit is exposed to a level higher than the specified electromagnetic compatibility test level, it is not required to complete the normal function, but is required to have the system function recover after the signal is removed. Unit intervention is allowed for recovering its functions. The unit is not required to provide the expected function in a one-strike, multistrike, or burst test, but shall be able to perform its expected function after the flash point pulse is applied. The test shall not cause permanent invalidity or damage of the unit. When the unit is exposed in a lightning environment, it is allowed to recover its function through unit intervention.

Unit test of the flight control system 167

2.8.2 Lightning direct effect test The lightning direct effect test is applicable to units externally mounted on aircraft skin, including external objects (such as wind vane of angle of attack sensor and landing light). Most of the electrical units of the flight control system are installed within the aircraft’s skin and this test is only applicable to the units of the flight control system installed outside the skin. 2.8.2.1 Test equipment and test requirements The effectiveness and reliability of lightning protection measures for units should be verified through analysis and testing. The test shall be carried out by a lightning laboratory with relevant test conditions. The test conditions include safety, experienced lightning simulation test personnel, and eligible test equipment and measuring instruments. The test equipment shall ensure correct and reliable grounding and insulation and the discharge circuit shall avoid unnecessary flashover and short circuit. The units shall be grounded and installed through the verified circuit according to the current flow direction in a lightning strike. The control and measuring system required by the test should adopt reliable high-voltage isolation measures to avoid high voltage. The accuracy of measuring instruments shall meet the requirements of measurement and metering and be within the valid period of the national legal measurement department. The accuracy of the measurement system shall be guaranteed by an effective electromagnetic compatibility design to avoid distortion of test results caused by the interference of a strong electromagnetic pulse in a lightning simulation test. Refer to Ref. [19] for requirements on the test waveform. 2.8.2.2 Test items and test methods Refer to Ref. [19] for test items and methods of the lightning protection qualification test for military aircraft. 2.8.2.3 Judging criteria and results handling Refer to Ref. [19] for judging criteria and results handling of the lightning protection qualification test for military aircraft.

2.8.3 Lightning-induced transient susceptibility test The lightning-induced transient susceptibility test shall have two procedures: damage tolerance test and functional disturbance test.

168 Chapter 2 All class I, II, and III units located in level 3
5 level areas of aircraft shall undergo the lightning indirect effect test. The relationship between the level area division of aircraft and the installation position is shown in Table 2.47. As the flight control system is the key to the safety of an aircraft and its units are required to complete this test generally (including units below class III), the units shall select the test waveform and level according to their importance and installation position. The importance degree of the units and installation position shall follow the agreement of the units and the signal interconnection relationship shall be used as a basis to determine whether to conduct an independent test or if there is an interconnection to test each subsystem. According to the influence of system failure modes on aircraft, the security level of the electronic/electrical systems can be divided into the following five categories: 1. Class I system: electronic/electrical systems whose failure will cause disaster of aircraft. 2. Class II system: electronic/electrical systems whose failure will cause danger of aircraft. 3. Class III system: electronic/electrical systems whose failure will cause large failure of aircraft. 4. Class IV system: electronic/electrical systems whose failure will cause small failure of aircraft. 5. Class V system: electronic/electrical systems whose failure will not affect the operation ability of aircraft or increase the workload of units. The test level of units is divided into 3
5 levels according to the installation position of units and interconnected cables in the aircraft, and the division of the test level is shown in Table 2.48. 2.8.3.1 Test objectives and test requirements The lightning-induced transient susceptibility test aims to evaluate the susceptibility of units to lightning discharge. Through the test, we can determine whether the lightning protection measures taken for the units are appropriate, so as to minimize the damages of units caused by lightning. The lightning environment, the type of lightning, and the lightning area of the unit under test shall be confirmed according to Ref. [20]. Test equipment and environmental requirements for the lightning-induced transient susceptibility test shall meet the requirements of Ref. [21]. 2.8.3.2 Test items and test methods The lightning-induced transient susceptibility test has two procedures: damage tolerance test and functional disturbance test. For the damage tolerance test, a plug and pin

Unit test of the flight control system 169 Table 2.48: Division of test level of units. Test level Cable bundle Position of unit and interconnected cable

Pin

Cockpit, nose equipment compartment, cargo hold, lower equipment compartment, environmental control equipment compartment, hydraulic equipment compartment, rear equipment compartment, wing box Nose radome area, front landing gear compartment, main landing gear compartment and fairing, landing gear, vertical stabilizer box section, leading edge of vertical stabilizer, leading edge of wing, trailing edge of wing, engine suspension area, engine nacelle, horizontal stabilizer, elevator, outer skin of aircraft, fuselage-wing fairing area Rudder, wingtip

One-strike and Burst Test multistrike test item options

3

3

3

4

4

4

5

5

5

Table 2.49: Waveform and level for unit signal and power pin damage tolerance test. Waveform group A Level of test 3 4 5

3 (Voc/Isc)

4 (Voc/Isc)

5 (Voc/Isc)

600 V/24 A 1500 V/60 A 3200 V/128 A

300 V/60 A 750 V/150 A 1600V/320 A

300 V/300 A 750 V/750 A 1600V/1600A

injection test or cable bundle induction test of the tested piece are conducted to verify the damage tolerance of the test piece. A functional disturbance test shall be conducted on a fully configured and fully running unit with interconnected cables and interface loaded on the basis of the interconnection of subsystems. The test level is divided into 3
5 levels. 1. Damage tolerance test The damage tolerance test shall be carried out according to Ref. [21]. Table 2.49 specifies different levels of waveform and level. The test waveform is shown in Figs. 2.13
2.15. 2. Functional disturbance test The functional disturbance test shall be carried out in accordance with Ref. [21]. The test waveform and test level are shown in Table 2.50. The single impulse test of the cable bundle in the table can be integrated in multiple impulse tests. In this case, the first transient test level of multiple impulse tests is replaced by the test level in a single impulse test. When conducting a power cable test, it shall be ensured that the current of

170 Chapter 2

Figure 2.13 Voltage/current waveform 3. Notes 1: Voltage and current do not have to be in the same phase. The waveform can be sine wave or cosine wave.

Figure 2.14 Voltage waveform 4.

a single conductor does not exceed the corresponding pin test current specified in Table 2.49. The test waveform is shown in Figs. 2.13
2.19. 2.8.3.3 Judging criteria and results handling The handling of test results and judging criteria shall be subject to Ref. [21].

Unit test of the flight control system 171

Figure 2.15 Current/voltage waveform 5.

2.8.4 High-intensity radiated field protection test The HIRF protection test complements the electromagnetic compatibility test. HIRF test mainly has three procedures: class I radio susceptibility test, class II radio susceptibility test, and aircraft level low level swept frequency coupling test. The test items to be done for the units shall be classified according to the safety level of the electronic/electrical system. HIRF radio sensitivity test divides class I system into two categories: the first category is class I control function, for which the pilot is not part of working circuit; and the second category is class I display function, for which the pilot is within the working circuit through pilot/system information exchange. The first category requires the class I radio susceptibility test and aircraft level low level swept frequency coupling test. The second category requires the class I radio susceptibility test. Class II and III systems are required to complete the class II radio susceptibility test. Class IV and V systems are not required to complete the high-intensity radiated field protection test. According to the principles above, class III (or above) units of the flight control system installed in open or half open area should conduct the HIRF test, such as actuators and control plane position sensors. 2.8.4.1 Test objectives and test requirements Units shall have sufficient HIRF protection ability to prevent electronic or mechanical and electrical units that execute key functions causing disastrous accidents due to direct or indirect HIRF function. HIRF test is an effective means to test the HIRF protection ability of units. 2.8.4.2 Test equipment and environmental requirements Test equipment and environmental requirements shall be subject to Ref. [21]. HIRF test should be completed in the electromagnetic compatibility laboratory and the electromagnetic compatibility laboratory shall pass the quality certification of “China National Accreditation Service for Conformity Assessment (CNAS)” and “Defense Science

172 Chapter 2 Table 2.50: Test waveform and test level of functional disturbance test. Test level

Cable bundle type

Waveform group

Test waveform number

3

Shield

J

1 3

Unshielded

G

2 3 4

4

Shield

J

1 3

Unshielded

G

2 3

Multiple pulse Waveform instructions

Single pulse

70 μs pulse (current) Pulse sine wave 1 and 10 MHz 6.4 μs pulse (voltage) Pulse sine wave 1 and 10 MHz 70 μs pulse (voltage) 70 μs pulse (current) Pulse sine wave 1 and 10 MHz 6.4 μs pulse (voltage) Pulse sine wave 1 and 10 MHz

300 V/ 600 A 600 V/ 120 A 300 V/ 300 A 600 V/ 120 A 300 V/ 600 A 750 V/ 1500 A 1500 V/ 300 A 750 V/ 1500 A 1500 V/ 300 A 750 V/ 1500 A 1600V/ 3200 A 3200 V/ 640 A 1600V/ 3200 A 3200 V/ 640 A 1600V/ 3200 A

4 5

Shield

J

1 3

Unshielded

G

2 3 4

70 μs pulse (current) Pulse sine wave 1 and 10 MHz 6.4 μs pulse (voltage) Pulse sine wave 1 and 10 MHz 70 μs pulse

First

After 13

Multiple pulse group

300 V/ 150 V/ (N/A) 300 A 150 A 600 V/ 300 V/ 360 V/6 A 120 A 60 A 300 V/ 150 V/ (N/A) 300 A 150 A 600 V/ 300 V/ 360 V/6 A 120 A 60 A 150 V/ 75 V/ (N/A) 300 A 150 A 750 V/ 375 V/ (N/A) 750 A 375 A 1500 V/ 750 V/ 900 V/15 A 300 A 150 A 750 V/ 375 V/ (N/A) 750 A 375 A 1500 V/ 750 V/ 900 V/15 A 300 A 150 A 3750 V/ 187.5 V/ (N/A) 750 A 375 A 1600V/ 800 V/ (N/A) 1600A 800 A 3200 V/ 1600V/ 1920V/ 640 A 320 A 32 A 1600V/ 800 V/ (N/A) 1600A 1600A 3200 V/ 1600V/ 1920V/ 640 A 320 A 32 A 800 V/ 400 V/ (N/A) 1600A 800 A

and Technology Industry Laboratory Accreditation Committee (DILAC)” in order to be able to test the items in the HIRF test. Test equipment such as the tester used in the test to monitor and record the working status and indicators of the unit shall be provided by the unit development unit. Moreover, the tester itself shall receive electromagnetic protection measures to withstand the test environment. The test cable shall be consistent with the installed cable in model, protection mode and ending mode. For cables longer than 10 m, they can be cut into 10 m cables. For cables shorter than 10 m, the test cable shall be made according to the length of installed cable.

Unit test of the flight control system 173

Figure 2.16 Current waveform 1.

Figure 2.17 Voltage waveform 2.

Figure 2.18 Multistrike.

174 Chapter 2

Figure 2.19 Burst.

1. Requirements for class I radio susceptibility test a. Unit under test i. The test environment in the laboratory shall reflect the installation condition of the unit under test on the aircraft, such as the composition and laying of cable harness, cable shielding terminal and installation, the lap method and lap bar impedance of units, and the state change of units. ii. The system shall be tested when devices work normally, such as the input sensor. The input sensor can use a simulator as long as the simulator can accurately represent the terminal impedance of the sensor and it has been evaluated to meet the HIRF requirements with regard to its installation location. iii. For units and systems that should complete class I radio susceptibility test, if the unit and system are not part of the pilot working circuit, they are categorized as a control function and the aircraft level low level coupling test should be completed. b. Test level and modulation requirements For a class I radio susceptibility test, the test level shall be subject to the values given in Table 2.51. The modulation mode is as follows: i. From 100 to 400 MHz, 1 kHz square wave modulation (SW), modulation depth 90%. ii. From 400 MHz to 4 GHz, pulse modulation (PM), pulse width 4 μs (or wider), pulse repetition frequency 1 kHz. iii. From 4 to 18 GHz, PM, pulse width 4 μs (or wider), pulse repetition frequency 1 kHz.

Unit test of the flight control system 175 Table 2.51: Level requirements for class I radio susceptibility test (unit: V/m). Test frequency range 100
200 MHz 200
400 MHz 400
700 MHz 700 MHz
1 GHz 1
2 GHz 2
4 GHz 4
6 GHz 6
8 GHz 8
12 GHz 12
18 GHz

Peak value

Average value

100 100 700 700 2000 3000 3000 1000 3000 2000

100 100 50 100 200 200 200 200 300 200

iv. From 400 MHz to 18 GHz, for systems/units with low frequency response such as flight control system, 50% duty ratio at rate of 1 Hz is used for SW and PM signal on/off. v. Modulation of other signals related to units and systems, such as clock, data, intermediate frequency, internal processing, and modulation frequency. The test level shall be determined according to the worst case of the harness position on the aircraft. The interface of the units connected to these harnesses shall be tested at the worst level of the interface. 2. Requirements for class II radio susceptibility test a. Unit under test For the class II radio susceptibility test, only the unit level test is required and this test shall include the sensor or sensor simulation. b. Test level and modulation requirements For the class II radio susceptibility test, the test level shall be subject to the values given in Table 2.52 and the modulation mode is the same as the modulation mode in the class I radio susceptibility test. 2.8.4.3 Test items and test methods The content of the high-intensity radiated field test includes the class I radio susceptibility test, class II radio susceptibility test, and aircraft level low level swept frequency coupling test. The high-intensity radiated field test method shall follow the requirements of Ref. [21]. 2.8.4.4 Judging criteria and results handling The products that meet the specified limit value requirements after the test are conforming products.

176 Chapter 2 Table 2.52: Level requirements for class II radio susceptibility test (unit: V/m). Test frequency range 100
200 MHz 200
400 MHz 400
700 MHz 700 MHz
1 GHz 1
2 GHz

Peak value

Average value

Test frequency range

Peak value

Average value

100 100 600 600 600

100 100 50 100 100

2
4 GHz 4
6 GHz 6
8 GHz 8
12 GHz 12
18 GHz

750 750 600 750 600

100 100 100 150 100

The HIRF protection test task description of units shall specify the acceptance/failure criteria for units. In the procedure of HIRF protection test, detailed working parameters and methods for monitoring these parameters shall be defined. The basic judging criteria for acceptance are as follows: 1. About disaster When and after the unit is exposed in HIRF environment or when and after it is applied with single strike, multiple strikes, or burst transient voltage, its function shall be kept valid. In the event of any system interruption, the continued performance and effectiveness of the functions of the primary or standby units shall be ensured, which shall be assessed and approved by the chief engineer system and the user. The affected units shall not provide misleading information and shall be able to recover automatically after the test signal is removed. 2. About danger and larger effects When the unit is in a specified HIRF test, it must not be adversely affected. When the unit is exposed to a specified HIRF level, it is not required to complete the normal functions, but is required to have the system function recover after the signal is removed. Unit intervention is allowed for recovering its functions. The unit is not required to provide expected function in a one-strike, multistrike, or burst test, but shall be able to perform its expected function after the flash point pulse is applied. The test shall not cause permanent invalidity or damage of the unit. When the unit is exposed to a lightning environment, its functions can be recovered through intervention.

2.8.5 Electrostatic discharge protection test Electrostatic discharge is the transfer of electrostatic charge between two objects with different electrostatic potential due to direct contact or electrostatic field induction. As electrostatic discharge may cause personal injury or unit damage, units that may cause electrostatic discharge due to personal contact shall be taken with electrostatic discharge protection measures in their design, including mechanical, electronic, and avionic electronic units.

Unit test of the flight control system 177 2.8.5.1 Test objectives and test requirements The electrostatic discharge test aims to evaluate the immunity to interference of units to ensure that they will not have permanent performance degradation and can still execute expected functions even under electrostatic pulse air discharge. The onboard units shall be capable of withstanding a series of electrostatic pulses. The severe level selected for electrostatic discharge is 15,000 V and points to a specific human contact position on the electrostatic discharge unit under test (EUT). There shall be 10 pulses of positive and negative voltage polarities at each selected position. The EUT shall be in an energized state and operate in a required manner. Test points shall include the following applicable locations: any points within the control or keyboard area and any other human contact points, such as knobs, buttons, indicator LEDs, gaps, barriers, connector housings, and other areas accessible to operators. 2.8.5.2 Test items and test methods The electrostatic discharge protection test method shall be subject to Ref. [21]. 2.8.5.3 Judging criteria and results handling The handling of test results and judging criteria shall be made by referring to those of the electromagnetic compatibility test.

2.9 Reliability test The reliability test associated to unit development includes the environmental stress screening (ESS) test, reliability preexposure test, reliability growth test, reliability qualification test, and reliability acceptance test.

2.9.1 Environmental stress screening test 2.9.1.1 Test objectives and test requirements The ESS test is a test for developed and produced units to implement ESS in order to detect and eliminate early failures caused by defective components, manufacturing processes, or defects introduced due to other reasons. ESS is mainly applicable to electronic units as well as electrical, mechanical, electrical, and photoelectric units. The manufacturer shall conduct ESS on as many electronic circuit boards, components, and units as possible. Generally, the manufacturer shall screen incoming components twice according to the regulations and relevant requirements.

178 Chapter 2 2.9.1.2 Test equipment and environmental requirements The test equipment and environmental requirements to have ESS of units shall be subject to Ref. [22]. 2.9.1.3 Test items and test methods For circuit boards and electronic, electrical, and mechanical units of assemblies, test items and test methods shall be subject to Ref. [22] or [23], if possible. All nonelectrical units except for pure mechanical units shall have ESS according to Ref. [22]. 2.9.1.4 Test results handling and judgment Test results handling and judgment shall be subject to Ref. [22].

2.9.2 Reliability preexposure test 2.9.2.1 Test objectives and test requirements The reliability level of products is improved by applying appropriate environmental stress and workload to the units and improving their design after finding design defects of them. The reliability preexposure test shall be carried out for newly developed units and key units for reliability, especially the units with high technical capacity. When necessary, the reliability preexposure test scheme shall be approved by the ordering party. 2.9.2.2 Test equipment and environmental requirements The test equipment and environmental requirements shall be subject to Ref. [24]. 2.9.2.3 Test items and test methods The reliability preexposure test is a part of the unit development test and it should be combined with the unit development test as far as possible. The environmental stress spectrum and workload of the test shall be provided or approved by the ordering party. The reliability preexposure test can adopt an accelerated stress or simulation test to identify weak links and induce failure or verify the design allowance. The test items and test methods shall be subject to Ref. [24]. 2.9.2.4 Judging criteria and results handling Handling of test results and judgment shall be subject to Ref. [24].

Unit test of the flight control system 179

2.9.3 Reliability growth test 2.9.3.1 Test objectives and test requirements By applying the comprehensive environmental stress in a simulated actual service environment, the test aims to expose potential defects of units and take corrective measures for them to ensure the reliability of the unit reaches the specified requirements. The reliability growth test shall have a clear growth goal and a growth model, and the life profile and task profile of the test, the number of test pieces, and the technical status shall be put forward by the ordering party. The evaluation shall be carried out before and after the reliability growth test to determine the certainty and effectiveness of the test. A successful reliability growth test can substitute the reliability qualification test, but it should be approved by the ordering party. 2.9.3.2 Test requirements, methods, and results judgment The test equipment and environmental requirements, test methods, and results judgment shall follow the requirements of Ref. [25] or other relevant standards.

2.9.4 Reliability qualification test 2.9.4.1 Test objectives and test requirements The reliability qualification test aims to verify if the reliability design of units has reached the specified reliability requirements. The reliability qualification test shall be conducted for units with reliability index requirements, especially the units with critical missions or high technical content. The reliability qualification test is generally conducted by a third party. It should be carried out for a system product at the highest possible level, so as to fully assess the interface condition and improve the authenticity and reliability of the test. The qualification test can be carried out in combination with the type finalizing test or life test of units. The test piece of the qualification test shall be under the type finalizing state and approved by the ordering party. The quantity, test scheme, life profile, mission profile, and failure judging criteria of the test piece shall be given by the ordering party. The qualification test shall be conducted after the environmental qualification test and ESS. Evaluation must be conducted before and after the test. 2.9.4.2 Test requirements, methods, and results judgment The test equipment and environmental requirements, test methods and results judgment shall follow the requirements of Ref. [24] or other relevant standards.

180 Chapter 2

2.9.5 Reliability acceptance test 2.9.5.1 Test objectives and test requirements The reliability acceptance test aims to verify if the reliability of batch produced units reaches the specified reliability level. The batch acceptance test shall be conducted for units required by the ordering party. The test pieces of the acceptance test shall be randomly selected from batch produced products and the quantity and batch of test pieces shall be determined by the ordering party. The test scheme, life profile, mission profile, and failure qualification criteria shall be given by the ordering party. The acceptance test shall be conducted after ESS and the evaluation shall be conducted before and after the test. 2.9.5.2 Test requirements, methods, and results judgment The test equipment and environmental requirements, test methods, and results judgment shall follow the requirements of Ref. [24] or other relevant standards.

2.10 Endurance test 2.10.1 Test objectives and test requirements The endurance test must be carried out for both mechanical and electrical units of the flight control system to verify the ability of units to withstand the expected working stress and environmental stress and to maintain normal working in the specified working time. The conditions of the endurance test include the environmental conditions, working conditions, and service and maintenance conditions of units. For each unit, the environmental conditions, working conditions, and service and maintenance conditions of the endurance test shall be specified in the product specifications of units in light of actual conditions. Environmental conditions include the mechanical environment, natural environment, and electromagnetic environment. When formulating environmental conditions, the economic and technical feasibility of the test shall be paid with attention and only the environmental conditions and reasonable values that are most sensitive to the durability of units should be selected as far as possible. For example, a hydraulic actuator should select hydraulic oil temperature as the environmental condition and a pure mechanical unit should select atmospheric temperature as the environmental condition. Working conditions generally include input command amplitude/frequency, working time/ times, and workload. They shall be given according to the specific working conditions and service characteristics of units. The extra coefficient of working time/times is generally 1.5.

Unit test of the flight control system 181 Service and maintenance conditions generally include regular lubrication, regular replacement of vulnerable parts, and regular cleaning. The items and frequency shall not exceed actual service and maintenance work.

2.10.2 Test items and test methods The test items and methods shall be given in the product specifications and the main test steps are as follows: 1. Prepare for testing the onboard units. 2. Install the unit on the test bed and conduct joint commissioning with other units. 3. Conduct appearance inspection and the functional and performance test for the unit before the test. 4. Start the test in accordance with specified environmental conditions, working conditions, and maintenance conditions and conduct the functional and performance test in the test process. 5. Stop the test after the specified time (generally the cycle of the first restoration), conduct appearance inspection and the functional and performance test, and then repair the specified items of the unit. 6. Repeat (2)
(5) until the cumulative working time of the unit under test reaches the specified requirements (generally the total life), and then stop the test.

2.10.3 Judging criteria and results handling The endurance test generally adopts one to two units to have test. If a failure (the failure mode shall be determined by the special specification of the unit) occurs in the endurance test, the test is stopped. The formula used for the life cycle calculation of the unit is: T0 5

αβT K

where T0 is the life value of the unit obtained through the test; α is the test stress coefficient (in the test, the environmental stress and working stress are usually equivalent to the actual use condition of the unit, α 5 1 is taken); β is the sample coefficient (β 5 0.7 is taken when one unit is used for test and β 5 1 is taken when two units are used for the test); T is the normal working time of the unit from the start of the test to the emergence of the failure (the average value is taken when two units are used for the test); and K is the empirical coefficient and generally 1.5 is taken.

2.11 Testability test Testability is a design characteristic that can in a timely and accurate manner determine the state of the system or unit, such as working, nonworking, or performance degradation, and

182 Chapter 2 isolate its internal failures. In other words, the unit has the ability of self-diagnosis and failure isolation. The testability test is applicable to electronic units with testability requirements, such as flight control computers and controllers.

2.11.1 Test objectives and test requirements Testability test is a technical means to determine whether the system or unit can meet the specified testability requirements and evaluate the expected effectiveness of testability. The testability requirements generally include failure detection rate, failure isolation rate, failure detection time, failure isolation time, and false alarm rate. Testability test requirements are as follows: 1. There should be clear requirements on the quantity and technical status of the units used for the testability test. 2. All the failure injection operations in the testability test shall be carried out under indoor natural environmental conditions and the natural environmental conditions shall be subject to Section 2.2.1. 3. The test units mainly include failure injection equipment, signal acquisition equipment, excitation equipment, test power supply, common test instruments and tools, and relevant tools software. The test equipment shall meet the requirements of test execution, parameters, and safety requirements. The test equipment that can be used for measurement shall have passed metrological verification and their measuring accuracy shall be at least one third of the tolerance of the measured parameters, such as signal acquisition equipment, excitation equipment, and other test equipment. The test equipment that cannot be used for measurement shall be discussed by the expert group to confirm if it can be used for the testability test. 4. The test scheme and test procedures shall be made and the failure mode shall be screened according to the technical agreement and technical status of the units. In the test scheme, the principle of sample allocation and selection, preliminary sample size, sample size supplement, alternative fault sample base, parameter evaluation, and other information shall be specified.

2.11.2 Test items and test methods According to the model requirements or technical agreement, the test verification content generally includes the following aspects: 1. BIT detection and failure isolation capability of UUT. 2. UUT’s compatibility with selected test equipment. 3. The degree of conformance between BIT detection and failure isolation indication and offline test results. 4. The effectiveness of the model used to predict the testability index.

Unit test of the flight control system 183 The test methods shall be subject to Ref. [26]. If necessary, appropriate methods and criteria in Ref. [27] can be used and additional verification can be implemented to obtain sufficient testability data used for the evaluation and as part of the testability verification results for compiling documents.

2.11.3 Judging criteria and results handling Conduct the test according to test procedures and record the noninjectable failure review data, injectable failure data, and natural failure data. The judging criteria of the test include the judging criteria for the intact condition of the unit, judging criteria for failure detection, and judging criteria for failure isolation. The criteria above are all given according to relevant technical indexes specified in product specifications or technical agreement of the units.

2.12 Test piece selection and test sequence of the unit qualification test The principles for the selection of test pieces for the unit qualification test are generally as follows: 1. For the functional test, performance test, natural environment test, mechanical environment test, and electromagnetic protection test of mature units, one piece (set) of unit is generally selected as the test piece. 2. For newly developed units, a piece (set) of test pieces will be generally added for tests (acceleration test, lightning protection test) that causes damage to units or relatively long tests (mold test). 3. For durability test, one piece (set) or two pieces (sets) of test pieces are generally selected and the number of test pieces depends on the value of dispersion coefficient and the test cycle. 4. The test pieces of the reliability test can be combined with those of the environmental test for use. The test sequence of the unit qualification test shall be subject to Ref. [6] and the general principles are as follows: 1. For developmental tests, as part of the performance study of the model machine, the test sequence begins with the most severe test items to obtain the trend of failure of the unit under test at the early stage of the test. 2. In case of limited test pieces, the test sequence of the developmental test starts from the least severe test items, so as to get as much information as possible before the unit is destroyed in the test.

184 Chapter 2 3. In a standard qualification test of units, the test sequence is generally arranged to ensure that the results of the previous tests will be exposed or strengthened by the latter test. This test sequence has the greatest impact on the unit under test. 4. For units with known service conditions, the tests are generally arranged in the emerging sequence of environmental factors that have the greatest impact on the unit. The qualification test of units is generally recommended to be carried out in the sequence specified in (3).

2.13 Organization and implementation of the unit qualification test The organization and implementation of the qualification test of the unit of the flight control system shall follow General Specification for Installation and Test of Electronic Units (GJB1403—1992). The unit test task shall be assigned by the ordering party in the form of technical agreement or order contract. The test outline or acceptance specification shall be prepared by the manufacturer or the testing unit and the test scheme shall be executed in accordance with the technical agreement or relevant model standards and approved by the ordering party. The test outline shall be reviewed according to the requirements and the review of the test outline for class III (or above) units shall be organized by the ordering party. The electromagnetic environment protection test and reliability test shall be completed by a third party. Tests of all units shall be attended by military representatives stationed in the factory and the responsibilities of the test personnel shall be clearly specified.

References [1] Commission of Science, Technology and Industry for National Defense. General specification for aircraft pilot controllable sensor (GJB2349—1995), 1995. [2] Commission of Science, Technology and Industry for National Defense. Test method for rate gyro assembly (GJB669—1989), 1989. [3] Standardization Administration of the People’s Republic of China, Preferred Numbers—Series of Preferred Numbers (GB/T321-2005), Standards Press of China, Beijing, 2005. [4] Aerospace Industry of the People’s Republic of China. Digital test of performance parameters of aircraft power supply system (HB6448—1990), 1990. [5] Commission of Science, Technology and Industry for National Defense. Test methods for aircraft power supply characteristic parameters (GJB5189—2003), 2003. [6] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment—General rules (GJB150.1—1986), 1986. [7] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment’s—Acceleration test (GJB150.15—1986), 1986. [8] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment—Vibration test (GJB150.16—1986), 1986.

Unit test of the flight control system 185 [9] Standardization Administration of the People’s Republic of China, Acoustics--Noise Emitted by Machinery and Equipment--Measurement of Emission Sound Pressure Levels at a Work Station and at Other Specified Positions—Site Simple Method (GB/T17248.3—1999), Standards Press of China, Beijing, 1999. [10] Ministry of Health. Basic health standards for radiological protection (GB4792—84), 1984. [11] Commission of Science, Technology and Industry for National Defense. Expression and evaluation of uncertainty in measurement (GJB3756—1999), 1999. [12] Commission of Science, Technology and Industry for National Defense. Electrostatic discharge control program for protection of electronic products (GJB1649—1993), 1993. [13] Commission of Science, Technology and Industry for National Defense. Military equipment laboratory environment test method—part 7: solar radiation test (GJB150.7A—2009), 2009. [14] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment—Solar radiation (GJB150.7—1986), 1986. [15] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment—rain test (GJB150.8—1986), 1986. [16] Commission of Science, Technology and Industry for National Defense. Environmental test methods for military equipment—mold test (GJB150.10—1986), 1986. [17] Commission of Science, Technology and Industry for National Defense. Electromagnetic emission and susceptibility requirements for military equipment and subsystems (GJB151A—1997), 1997. [18] Commission of Science, Technology and Industry for National Defense. Measurement of electromagnetic emission and susceptibility for military equipment and subsystems (GJB152A—1997), 1997. [19] Commission of Science, Technology and Industry for National Defense. Lightning protection qualification test techniques for military aircraft (GJB3567—1999), 1999. [20] Commission of Science, Technology and Industry for National Defense. Lightning protection for military aircraft (GJB2639—1996), 1996. [21] RTCA SEBS. RTCA DO-160E Environmental Conditions and Test Procedures for Unit, 2004. [22] Commission of Science, Technology and Industry for National Defense. Environmental stress screening process for electronic products (GJB1032—1990), 1990. [23] Commission of Science, Technology and Industry for National Defense. The guidance of quantitative environmental stress screening on electronic products (GJB/Z34—1993), 1993. [24] Commission of Science, Technology and Industry for National Defense. Reliability qualification and acceptance test (GJB899A—2009), 2009. [25] Commission of Science, Technology and Industry for National Defense. Reliability growth test (GJB1407—1992), 1992. [26] Commission of Science, Technology and Industry for National Defense. Testability program for material (GJB2547—95), 1995. [27] Commission of Science, Technology and Industry for National Defense. Maintenance test and evaluation (GJB2072—1994), 1994.

CHAPTER 3

Verification and validation of flight control system airborne software 3.1 Overview 3.1.1 Purpose and significance of verification and validation Flight control system airborne software of large aircraft are generally divided into airborne software of the fly-by-wire flight control system airborne software, high lift control system airborne software, and automatic flight control system. The functions of flight control system airborne software generally include hardware resource management, operating system, task management, interface management, sensor data management, bus data management, redundancy management, system BIT, control mode management, control law calculation, and fault warning integration. The configuration items of airborne software of the flight control system of large aircraft are shown in Table 3.1. In view of the high safety and reliability requirements of the flight control system of large aircraft, the airborne software of large aircraft is divided into class A, B, C, D, and E according to Software Considerations in Airborne Systems and Equipment Certification (RTCA DO-178B). Class A: software whose failure may cause or lead to system failure and disaster of aircraft. Class B: software whose failure may cause or lead to system failure and danger of aircraft. Class C: software whose failure may cause or lead to system failure and large failure of aircraft. Class D: software whose failure may cause or lead to system failure and small failure of aircraft. Class E: software whose failure may cause or lead to system failure but will not affect the performance of aircraft or the workload of pilot. The class of airborne software shall be determined after comprehensive analysis according to following factors. Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00003-6 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

187

188 Chapter 3 Table 3.1: Configuration items of airborne software of flight control system. No. 1

2

Name

Subsystem

Fly-by-wire flight control system airborne software

Fly-by-wire flight control system airborne software

Rudder trim Fly-by-wire flight mechanism control control system software airborne software

3

Automatic flight control system airborne software

Automatic flight control system

4

Automatic flight control postback actuator control software

Automatic flight control system

5

Automatic flight control device software

Automatic flight control system

6

High lift system software

High lift system

LRU

Software description

System scheduling and management; cross transmission, voting, and monitoring of input redundancy signals; control law calculation; thrust asymmetric compensation command calculation; cross transmission and monitoring of output signals; fault synthesis and reporting; signal acquisition and transmission at control plane position; communication control of system internal data bus; Information exchange and sharing among units; self-test and system maintenance control Rudder Trim mechanism bus data transmission; bus data transmission; control law calculation; loop monitoring; steering speed control System scheduling; computer Automatic management; cross flight communication; input/output control management; bus data signal computer processing; BIT; fault management; modal conversion logic; control law calculation Automatic Receiving, analyzing and processing the data from AFCC and feeding flight back the actuator state; command control computer reception; position and speed servo closed-loop calculation; drive command sending; clutch control; double branch synchronization; BIT monitoring and fault reporting; position signal acquisition and transmission of fuel regulating valve Receiving and analyzing data from Automatic AFCC and sending them to display flight module after processing; sending the control control data of AFCU panel unit to device AFCC and display module; BIT and fault recording and reporting Flaps and Flaps and slats control; computer slats control hardware management; task computer management; interface management; channel monitoring; system BIT; fault integration Fly-by-wire flight control computer

Software grade A

D

B

B

B

B

(Continued)

Verification and validation of flight control system airborne software

189

Table 3.1: (Continued) No. 7

Name

Subsystem

LRU

Flaps and slats power drive unit software

High lift system

Flaps and slats power drive unit (PDU)

Software description

Software grade

B Receiving the stroke difference command or the motion direction command of override switch from flaps and slats control computer; control plane closed-loop control and brake device on off control; feeding the working status of PDU and brake device back to flaps and slats control computer through the ARINC429 bus for system monitoring

1. Influence of software failure on aircraft safety, mission completion, aircraft flight accident analysis, fault judgment, and maintenance analysis. 2. System architecture of the software (such as parallel, nonsimilar, independent architecture). 3. Structure of the computer on which the software runs (e.g., whether it has fault detection and channel monitoring function). 4. Whether the software uses multiversion and nonsimilar design methods. According to the principles above, the safety level of airborne software of the flight control system of large aircraft is mostly class A or class B (see Table 3.1). Therefore it is extremely important to ensure the quality and correctness of software. In addition to the strict control of the software development process, the verification and validation (V&V) of software is also of great significance. “Verification” refers to a process to verify if specified requirements have been satisfied, if the product in each stage of the software life cycle follows the requirements of last stage of the life cycle (such as correctness, completeness, consistency, and accuracy) and if standards, measures, and agreement in the stage are satisfied through inspection and provision of objective evidences, so as to establish the relevant basis for starting the next stage of activities in the life cycle. “Validation” refers to a process to verify if the requirements for a specific intended use have been met and if the completed final products meet the requirements of established software through the inspection and provision of objective evidences. Software V&V supports all life cycle processes and it evaluates the results of both the software development process and the software validation process. The software validation process aims to test and report the errors that may have been formed in the software development process to ensure the conformity and traceability of system requirements,

190 Chapter 3 software requirements, software architecture, software design, and source code allocated to software and to ensure the final executable object code can satisfy system requirements.

3.1.2 Basic requirements of verification and validation The organizational structure of software V&V is related to V&V and there are different requirements for the software review, analysis, and testing. Software review aims to evaluate software elements according to a specified process to identify potential problems of software elements. Software review is a process of software product inspection organized by relevant personnel, usually in the form of a meeting. For software verification, especially in the early stage of software development, review is a frequently used verification method with the best effect. 3.1.2.1 Software review Software review is a process of examining the intermediate and final software products in the software development process using review forms or similar auxiliary means. The object of the software review “software product” refers to various technical files and source codes generated in software development activities that should be delivered, generally including contracts, project plans, requirement specification, design files, source codes, user files, support and maintenance files, test plan, testing specifications, standards, or other work products. The software review mainly aims to: 1. review and evaluate the product or process at each stage to assess its compliance with specified requirements; 2. identify existing defects; 3. take remedial actions; and 4. identify possible improvements in performance, safety, and economic efficiency. Software review generally includes peer review, management review, and auditing review: 1. Peer review: the evaluation of technical content and product quality organized by the software product developer or one (or more) member of the developer team. 2. Management review: the assessment of work progress by a management representative in order to make decisions on follow-up work. 3. Auditing review: the evaluation of the compliance of software products with relevant specifications, standards, and agreements conducted by personnel other than the software project team. Software review methods are generally divided into the following five types: 1. Code review: a systematic examination of program source code, generally peer review is adopted.

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2. Pair programming: two people “pair up” for software development and review each other’s code. 3. Review: a standard method of peer review, in which auditors discover software defects by performing well-defined processes. 4. Walkthrough: a method of peer review, in which the product developer leads members of the development team and other interested participants in inspecting the software product, and participants can ask questions and raise doubts in this process. 5. Technical review: a method of peer review, in which a group of qualified personnel check the realization of expected use of software products and the deviations of the software products from relevant specifications or standards. In the development process of the flight control system of large aircraft, software review is generally conducted through the combination of internal review and external review. The external review is also called the formal review, in which the setting of the software review points is determined according to the class of software. The basic requirements for the review of different classes of software are shown in Table 3.2. 3.1.2.2 Software analysis Software analysis includes traceability analysis, interface analysis, hazard analysis, and risk analysis. Traceability analysis: the traceability of system requirements for all development sets, user document sets, and test document sets are analyzed for consistency, completeness, and correctness to verify the software requirement is realized in the software related to the correct design, code, and testing information. Interface analysis: the interface analysis aims to ensure the integrity, accuracy, and consistency of hardware-to-software, software-to-software, and software-to-user interfaces. Review is a process of examining the intermediate and final software products in the software development process using review forms or similar auxiliary means. 3.1.2.3 Software testing Software testing is the process of verifying if a system meets specified requirements or identifying the differences between actual and expected results. As defined in Software Engineering Terminology 1983, software testing is a process of running or testing a system by manual or automatic means to verify if it meets specified requirements or to identify the differences between expected and actual results. The airborne software of the flight control system of large aircraft is a kind of embedded real-time multiredundancy control airborne software. Combined with the hardware of flight control computer, it realizes all functions and performance of the flight control system, such

Table 3.2: Review requirements for different classes of software. Class A Review content Software development task description Software development plan Quality assurance plan Configuration management plan Software verification plan Software conformity review plan Software requirement standard Software design standard Software code standard Software module importance analysis report Software requirement specification Interface requirement specification Software preliminary design specification Software interface design specification Database design specification (if necessary) Software detailed design specification Software test plan Software testing specification Software test record Software test report Software development summary report

B

C

D

E

Internal

External

Internal

External

Internal

External

Internal

O O O O O O O O O O O O O O O O O O O O O

O O

O O O O O O O O O O O O O O O O O O O O O

O O

O O O O O O O O O

O O

O O

O O

O O O O

O O O

O O O O O O O O O O O

O

O

O

O

O

O O O

O O O

O O O O O O O O O

O O O O O O O O

O O O

O O

External

O

O

Internal

External

O

O O

O

Notes: In the table, “internal” means internal review; “external” means external review; and “O” means review must be conducted. The blank does not have special requirements but it shall be taken as an review support document. For consolidated documents, review requirements are applicable to documents after consolidation.

1. Internal review: it is generally organized by the software project leader. The review process, content, and requirements, as well as the members and responsible persons of the review group shall also be determined by the project leader. In the internal review, peer experts shall be employed to undertake an in-depth review of technical aspects. Review records and comments shall be documented and managed and controlled or included for configuration management after the review. 2. External review: it is generally organized by the software development unit according to the task description or contract. A software review group composed of professional design personnel, who are not directly involved in the software development, software and hardware system design personnel, user representatives, and representatives of the quality department of the chief engineer unit, external interface equipment unit, and development unit, shall be established.

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as the operating system, control law calculation, redundancy management, and built-in test. To sum up, the airborne software of the flight control system has the following characteristics: 1. The flight control system is a real-time multitask scheduling management software. Many tasks such as control law calculation and redundancy management work in parallel. 2. The flight control system has complex external interfaces and has many interfaces to connect with external sensors, computers, and actuating equipment (such as ARINC429 bus interface, GJB2894 bus interface, AFDX interface, RS422 bus interface, RS232 interface, A/D conversion interface, D/A conversion interface, and discrete quantity interface). 3. Strong real-time performance of the software: the airborne software of the flight control system has many algorithms related to time. The main task cycle of the software shall be strictly controlled within specified time and all tasks such as data acquisition, redundancy management, and control law calculation are required to be completed within a certain time. 4. Simultaneous development and verification, various test environments: software development and testing follow the V&V model and it is required to conduct different levels of testing in different development stages so as to find problems early. Therefore different test environments should be established. Some tests need professional test tools. Besides, special simulation test target machines should be configured as the processors are different. 5. High reliability and safety: the core airborne software of the flight control system is a key software ensuring high safety and its unreliable performance will bring disastrous consequences. Therefore some advanced technologies to improve reliability and safety such as fault-tolerant technology, N-version technology, safety monitoring, and safety isolation technology shall be adopted in the design of airborne software. The requirement of high safety and reliability greatly increases the workload of testing. To improve the safety and reliability of software, the logic of software products becomes more complex and testing becomes more difficult after multiple technologies are adopted. Software testing is mainly based on software requirements. From the life cycle of software, the software testing can be divided into unit test, integration test, and system test (including the Component testing based on software integration and the configuration item test based on software and hardware integration). From the tested program, the software testing can be divided into static testing and dynamic testing. From the internal structure of the tested program, the software testing can be divided into black box testing and white box testing. These testing methods divide software testing from different perspectives. In actual testing, these testing methods are used together.

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3.1.3 Basic process of verification and validation The verification and validation process of the software life cycle is mainly divided according to the stages of the software life cycle model, and the “V” type software development model is usually adopted. On this basis, the verification and validation activities of the airborne software of the flight control system can be divided into the system analysis and design stage, software planning stage, software requirements analysis stage, software design stage, software implementation stage, and software testing stage. In the verification and validation activities of the airborne software of the flight control system, review, analysis, and test are mainly used to determine whether the software meets the system requirements and meets the system design requirements. The verification and validation activities of the airborne software during its life cycle are shown in Fig. 3.1. 3.1.3.1 Verification and validation in system analysis and design stage In the system analysis and design stage, requirements of the flight control system, hardware and software environment, and other requirements are analyzed and determined, the

Figure 3.1 Airborne software life cycle validation activities for the flight control system.

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software/hardware interface is designed, and finally the airborne software development task description of the flight control system is formed. Critical analysis, traceability analysis, requirements evaluation, hazard analysis, and risk analysis are also conducted for software verification and validation. 1. Critical analysis: according to the functions of the flight control system and its influence on aircraft safety, the software level and the accuracy and completeness of the software requirements assigned by the system are verified. 2. Traceability analysis: through the analysis of the input documents in this stage, such as system design requirements and technical agreement/contract, system requirements fully or partially realized by flight control system airborne software are identified and software requirements traceability analysis is conducted starting from the system requirements. 3. Requirements evaluation: it is conducted to evaluate if the system function, performance, input/output, and interface, and other technical requirements meet system requirements when the flight control system testing is carried out. 4. Hazard analysis: it is conducted to analyze if the system requirements described in the airborne software development task description of the flight control system are potentially hazardous, for example, if the system requirements are inconsistent with user requirements. 5. Risk analysis: technical and management risks are identified to provide suggestions for elimination, reduction, or mitigation of risks. Upon the completion of the system analysis and design stage, review work in this stage shall be organized to verify if the software development task description is consistent with the system requirements as required in the technical agreement/contract and the feasibility and testability of the system’s functional requirements, design, operation, and maintenance requirements are validated. In the system testing stage, the system is tested according to the flight control system testing plan and test instructions to generate a corresponding system test report and confirm if the flight control system meets the functional and performance requirements in the system development task description. 3.1.3.2 Verification and validation in software planning stage In the software planning stage, the software project team plans the software development, quality assurance, configuration management, qualification examination, and verification work according to model development progress, flight control system development progress, computer hardware equipment development progress, and other constraint conditions, specifies software development standards, and generates a plan guiding the development process of software.

196 Chapter 3 The software development plan, quality assurance plan, and configuration management plan in the software planning stage adopt hazard analysis and risk analysis to conduct verification and validation activities. 1. Hazard analysis: potential hazards in the airborne software plan and development standards of the flight control system are analyzed. Inconsistency between plans and failure to properly guide the software development process may lead to failure of the whole software development. 2. Risk analysis: management risks are identified to give suggestions for the elimination, reduction, or mitigation of risks. Upon the completion of the software planning stage, the software development plan, software quality assurance plan, software configuration management plan, software Certification Compliance Plan, software verification plan, and other documents are formed and review work in this stage is organized to review if the structure of the software plan is clear, if the activities in software development process and combined process are defined, if software plans are mutually coordinated, and if the plans are specific, reasonable, and feasible. 3.1.3.3 Verification and validation in software requirements analysis stage In the software requirements analysis stage, the software configuration item functions, performance, external interface, internal interface, software and hardware requirements, and safety-related requirements are detailed item by item according to the airborne software development task description of the flight control system and finally the software requirement specification is formed. Traceability analysis, interface analysis, requirements evaluation, hazard analysis, risk analysis, and configuration management evaluation are conducted to have verification and validation activities. 1. Traceability analysis: the traceability from software requirements to system requirements, from system requirements to software requirements, and between software requirements and system requirements are analyzed for correctness, consistency, and completeness. 2. Interface analysis: it is conducted to verify and validate if the interface between the software and the hardware and other systems is correct, consistent, accurate, and testable. 3. Requirements evaluation: it is conducted to evaluate if the function, performance, and interface of software configurations item meet requirements when the configuration items of the flight control system airborne software are tested. 4. Hazard analysis: it is conducted to determine the system hazards introduced by the software, identify the software requirements that lead to each system hazard and confirm the treatment, control, or mitigation to each hazard by the software.

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5. Risk analysis: on the basis of the review work conducted before and risk analysis results, risk analysis is conducted for potential risks in software requirements to provide suggestions on the elimination, reduction, or mitigation of risks. 6. Configuration management evaluation: it is conducted to verify that the configuration management process of the airborne software of the flight control system is complete and sufficient. Upon the completion of the software requirements analysis stage, the software requirement specification and software interface requirement specification are formed. A review is organized at this stage to evaluate the correctness, consistency, completeness, accuracy, readability, and testability of the software/interface specification and the software requirement specification is verified and validated to see if it conforms to software requirement standard. In the configuration item testing stage, software configuration items are tested according to the configuration item testing plan of the airborne software of the flight control system and test instructions to generate the corresponding test report and confirm if the airborne software configuration items of the flight control system meet the functional and performance requirements in the software requirement specification. 3.1.3.4 Verification and validation in software design stage In the software design stage, structural diagrams and software flow charts are used to establish the architecture of the airborne software of the flight control system and the relationship between software units, to define the data interface and control interface of software units, to design global data structure, to specify design constraints, and finally to form the software preliminary design specification and detailed design specification. In addition, traceability analysis, interface analysis, design evaluation, hazard analysis, and risk analysis are also conducted to provide verification and validation activities. 1. Traceability analysis: the traceability from software design to requirements, from requirements to design, and between design and requirements are analyzed in terms of correctness, consistency, and completeness. 2. Interface analysis: it is conducted to verify and validate the correctness, consistency, completeness, accuracy, readability, and testability of interfaces between software units. 3. Design evaluation: it is conducted to evaluate if the functions of units and the interfaces between units meet software design requirements when component testing of airborne software of the flight control system is conducted. 4. Hazard analysis: it is conducted to verify that the logical algorithm design and associated data fulfill key requirements without introducing new hazards. 5. Risk analysis: on the basis of the review work conducted before and risk analysis results, risk analysis is conducted for potential risks in software design to provide suggestions on the elimination, reduction, or mitigation of risks.

198 Chapter 3 Upon the completion of the software design stage, documents such as software preliminary design specification and software detailed design specification are formed and a review at this stage is organized to confirm the correctness, consistency, and accuracy of design elements and confirm if the software design conforms to the design requirements of the software design standard. In the software component testing stage, the unit is tested according to the testing plan of the flight control system and test instructions, and a corresponding test report is generated to confirm if the airborne unit of the flight control system meets the software design requirements.

3.1.3.5 Verification and validation in software implementation stage In the software implementation stage, source codes of the airborne software of the flight control system should be written to ensure the codes conform to programming standards and the functional requirements of the completed codes are consistent with the content of the design document, so as to finally form source codes. In addition, traceability analysis, interface analysis, code assessment, hazard analysis, and risk analysis are also conducted to have verification and validation activities. 1. Traceability analysis: the traceability from the source codes of airborne software of the flight control system to design specification, from design specification to source codes, and between source codes and design are analyzed for correctness, consistency, and completeness. 2. Interface analysis: it is conducted to verify and validate if the detailed information of interfaces between software units meets the requirements of software design. 3. Code assessment: it is conducted to assess if the editing of source codes meets the code editing specification and if the logic design and data structure of the unit meet the detailed design requirements of the software when airborne software unit testing is conducted. 4. Hazard analysis: it is conducted to verify that the edited source codes and associated data fulfill key requirements without introducing new hazards. 5. Risk analysis: on the basis of the review work conducted before and risk analysis results, risk analysis is conducted for potential risks in the source codes written by programmers to provide suggestions on the elimination, reduction, or mitigation of risks. Upon the completion of the software implementation stage, software source codes and executable object codes are generated and the review work in this stage is organized to evaluate the correctness, consistency, and accuracy of the source codes and evaluate if the software source code conforms to the software code standard.

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In the software unit testing stage, the software unit is tested according to the test plan and test instructions to generate a corresponding test report and confirm if the airborne software unit of the flight control system meets the software design requirements. 3.1.3.6 Verification and validation in software testing stage In the software testing process, units of software products shall be evaluated and integrated (not only the software units but also the software and hardware integration) and all verification and validation activities of planned testing stages (units, airborne units, configuration items, and system testing) shall be completed to finally form test reports of corresponding stages. In addition, testing evaluation, hazard analysis, and risk analysis are also conducted to have verification and validation activities. 1. Testing evaluation: it is conducted to evaluate if the test results of each stage in the software testing process meet the corresponding test requirements. 2. Hazard analysis: it is conducted to verify if the activities of each test stage meet the test requirements of the corresponding stages without introducing new hazards, such as the test environment and test methods. 3. Risk analysis: on the basis of the review work conducted before and risk analysis results, risk analysis is conducted for potential risks in each testing stage to provide suggestions on the elimination, reduction, or mitigation of risks. Upon the completion of each software testing stage, the software unit test report, component testing report, configuration item test report, and system test report are formed and in each test stage a review is organized to determine the adequacy of software testing and to evaluate the effectiveness of test results, the structure of the software, and the coordination between interfaces. Explanations are demanded for situations in which the test requirements are not met and there are differences between test results and actual results.

3.2 Software testing As an important part of software verification, software testing has three purposes. The first purpose is to verify whether the software meets the requirements on software quality characteristics of the software development contract or specification, the system/subsystem design document, the software requirement specification, and the software design specification. The second purpose is to identify software errors through testing. The third purpose is to provide a basis for the evaluation of software product quality. The complexity of airborne software, the diversity of system requirements, the usual changes of software requirements, and other factors make software errors inevitable. As the development process of the flight control system airborne software is usually accompanied with constant revision and repeated optimization due to the aircraft development schedule

200 Chapter 3 and the development personnel’s perception of the flight control system is a gradual process, human errors made by software developers in a series of activities are inevitable, no matter how skilled they are. For these reasons, all participants in the FCS airborne software project shall attach importance to software testing and shall never ignore the importance of software testing. The software project team shall organize about 40% human source to the testing and the software testing shall run through the entire life cycle of the software definition and development. The standard referenced for the airborne software testing of military aircraft in China is generally Military Software Testing Guide (GJB/Z141—2004). The standard referenced for the airworthiness review of airborne software of civil aircraft and software testing is RTCA DO-178B (the new version is DO-178C). With the increasingly growing requirements for the flight safety of military aircraft, standard DO-178B is starting to be applied in airborne software testing of military aircraft. In order to ensure the quality of the aircraft airborne software, third party testing is generally required for the airborne software of military aircraft. Airborne software of civil aircraft must pass RTCA DO-178B/C to show its compliance with civil aviation regulations.

3.2.1 Unit testing The airborne software of the flight control system is composed of different software modules. Each software module has a clearly defined subfunction and this subfunction is not interdependent on the functions of other software modules at the same level. The unit testing of airborne software of a flight control system tests these basic modules and main control paths to verify if each software unit can realize its functions and meet the requirements of performance and interface. Software unit testing is an indispensable link in the process of software testing. Through this link, low-level errors hidden in the software can be found, so that the difficulty in locating and finding errors in subsequent tests of airborne units and configuration items can be avoided. 3.2.1.1 Testing plan Before the dynamic testing of the software unit, it is necessary to make a general plan of the unit testing process of the airborne software of the flight control system, which shall cover the composition of the test environment used in the software unit testing, the division of labor of the testing personnel, the testing content, the testing methods, and the testing

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progress, etc., so that the software testing efficiency can be improved and the testing cost can be lowered. The software unit testing plan shall conform to and coordinate with the software development plan. The testing methods and requirements adopted in the software unit testing process must be consistent with the content described in the testing plan and the testing personnel shall carry out the software unit testing activities in strict accordance with the plan. The software unit testing plan guides software testing personnel to avoid risks in the process of software testing and avoid frequent problems causing delay of the testing process, promotes the communication between testing personnel, and assists quality management so as to make the testing management easier. 3.2.1.2 Testing methods The airborne software unit testing of the flight control system includes software unit static testing and software unit dynamic testing. The software unit testing mostly adopts the white box testing method and the parallel testing of several units is allowed. Just as the name suggests, white box testing means that the program is regarded as in a transparent white box and the structure and control process of the program are fully visible, and the testing personnel test the software according to its internal logical structure. As white box testing requires an in-depth understanding of the software, including its structure, components, and their relations, as well as the internal operation principle and logic, the software unit testing can be undertaken by a programmer. The airborne software unit testing method of the flight control system is shown in Fig. 3.2. Before the software unit dynamic testing, the static analysis of software source codes, code review, and document review should be conducted to ensure the source codes conform to the code editing specification. The static analysis can use TestBed to have code analysis for the airborne software of the flight control system. Code review verifies data transmission and data call between software interfaces and the consistency between them and the software detailed design. Document review reviews the consistency between the content in the software detailed design specification and the preliminary design specification. For software dynamic testing, the white box testing method is adopted to focus on the concrete implementation of the unit, internal logic structure, data flow, and control flow and to complete the interface testing of units, local data structure testing of units, testing of all independent execution paths in units, testing of various error handling, and boundary condition testing of units. The running environment of software unit testing is a host. Since the software unit itself is not an independent program, a driver module and stub module shall be developed for each

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Figure 3.2 Airborne software unit testing method of the flight control system.

unit testing to build a software unit testing environment. The driver module is simply a unit that receives test data and transmits the data to the unit under test, and then outputs the relevant results. A stub module is a replacement for those modules that belong to the unit under test. The environment of airborne software unit testing of the flight control system is shown in Fig. 3.3. 3.2.1.3 Testing specification The content of the airborne software unit testing of the flight control system mainly comes from the software detailed design specification. Software testing personnel review the documents and extract the test case required by the software unit testing. The design of test examples shall be based on the principles of test requirements and test methods to ensure the correctness and traceability of test items. Software testing personnel design test examples according to the function points in the software detailed design specification and give the input data and corresponding output data of the software unit. Among them, the designed test examples shall meet the functional requirements of the software unit to achieve functional coverage. Meanwhile, equivalence partitioning and boundary value analysis methods shall be used for source codes to design

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Figure 3.3 Environment of airborne software unit testing of the flight control system.

test examples and achieve the logic coverage of the software unit, such as statement coverage, decision coverage, and condition coverage. Equivalence partitioning divides the data collection in the software unit. In each class, the function of a typical value in the testing is same as the function of other values in the class and the testing personnel select a set of groups in the equivalence class as test examples for testing and identify the errors in the program. Boundary value analysis focuses on boundary conditions of input and output space to identify test examples. Practices show that the program is correct when dealing with a large number of intermediate values, but error is most likely to appear at the boundary. For example, in a control law operation module of the flight control system, there are many flight parameter values and most of them are floating point numbers. The error of one decimal point may greatly affect the function of the unit. Each software test case shall have a unique identification number and make a detailed description of requirements tracking, test input, expected test results, criteria for evaluating test results, and test steps, so as to facilitate software testing personnel to carry out the unit testing activities and design examples used for software regression testing. 3.2.1.4 Testing results According to the content of the software unit testing plan, the airborne software testing personnel set up a software unit testing environment, use software testing tools, import the data of software test case, and conduct dynamic testing for the software unit. The process of airborne software unit testing of the flight control system is shown in Fig. 3.4.

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Figure 3.4 Process of airborne software unit testing of the flight control system.

Through the analysis of the software unit testing results, it can be judged whether the test case meet the software coverage requirements in the software unit testing plan. For the software units not reaching the coverage rate, test examples shall be supplemented for dynamic testing until the requirements are met. Coverage is used to measure the test integrity. According to software grade and model requirements, the coverage of the airborne software unit testing of the flight control system includes statement coverage, decision coverage, and condition coverage. Statement coverage aims to execute every executable statement in the program at least once, decision coverage aims to make every decision in the program adopt the true branch and fault branch at least once, that is, the true/fault value of the decision should all be satisfied, and the condition coverage aims to satisfy the possible value of every condition in every decision at least once. In the airborne software unit testing process, if it is found that the output results of the program are inconsistent with the expected results, it is necessary to analyze the potential errors in the program. After confirmation, the errors shall be fed back to the programmer for modification and then regression testing shall be conducted for the software unit until the software unit passes the test.

3.2.2 Component testing The component testing of airborne software of the flight control system is based on software preliminary design specification, with a purpose of checking the interface relations between software units and modules and the correctness of the realization of module functions. Component Testing is a transition of testing from unit to module. When making a testing plan, the architecture of the airborne software of the flight control system, including the hierarchical relationship and dependency relationship between modules, shall

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be considered, while the division of test items and the integration strategy of the software unit are particularly important for testing design. Reasonable test item division and unit integration strategy can shorten the testing time, reduce technical risks, and improve testing efficiency. 3.2.2.1 Testing plan The component testing plan includes the division of component testing items, the definition of testing types, the division of labor, and the testing schedule. It is used to guide the whole component testing process. The test items are divided according to the results of the software preliminary design and the integration granularity and integration strategy of the software airborne unit are determined based on the functional complexity of modules and the coupling degree between modules. The test items 100% cover all the functions of the software airborne unit defined by the software preliminary design. The component testing strategy of the airborne software of the flight control system is a hybrid integration testing mode, as shown in Fig. 3.5. 1. A big bang integration testing strategy is adopted for BIT software functional modules that are closely coupled with external hardware. 2. An integration testing strategy combining bottom-up with big bang is adopted for control law operation modules with complex internal functions, call relationship, and data coupling. 3. A bottom-up integration testing strategy is adopted for redundancy management modules with independent functions, and simple interface relations.

Figure 3.5 Component testing strategy of airborne software of the flight control system.

206 Chapter 3 3.2.2.2 Testing methods The methods adopted for the component testing of airborne software of the flight control system include static analysis, code review, and dynamic testing. Dynamic testing combines the white box testing method and the black box testing method. Static testing is undertaken before the dynamic testing. Static analysis and dynamic testing can be completed with software testing tools. The static analysis and code review of component testing focus on the function interface and call check. Dynamic testing focuses on the correctness of module functions, through the method of combining stubbing and source code calling generally. The environment of the component testing is the same as the environment of unit testing, and component testing can be carried out in the environment of unit testing. When conducting the control law software component testing of the flight control system, its characteristics of multimode, multirate set, and continuity shall be considered and the testing personnel shall carry out functional testing of the control law software through the data of continuity testing. In practical engineering application, the control law simulation model generates continuity test data through Simulink simulation software. The data include input data and output data. Meanwhile, the driver program of software testing is written to preprocess test data and call the control law software under test. Finally, the driver program is tested with the TestBed software testing tool (i.e., the control law software is tested) and the software testing results are analyzed. The test data uncovered will be analyzed and the data range not covered will be generated again through the control law model for supplement to the test data until all the test data in the cycle is covered. Then, the control law software will be verified. The process of the component testing of control law software of the flight control system is shown in Fig. 3.6. 3.2.2.3 Testing specification The content of the airborne software component testing of the flight control system mainly comes from the software preliminary design specification. The design methods of test examples are the same as those of unit testing. For the design of examples for component testing, emphasis shall be given to: 1. the correctness of parameter passing and result returning between airborne equipment (or units). 2. the correctness of the functions of airborne units after assembly of airborne equipment (or units). 3. the correctness of the global data structure.

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Figure 3.6 Method of component testing of control law software of the flight control system.

3.2.2.4 Testing results The execution environment, testing process, and testing adequacy judgment of component testing are same as that of unit testing. With regard to their differences, the first is that component testing substitutes the stub modules, driver programs, and global data constructed in the unit testing process with actual airborne equipment (units) and actual global data and gradually assembles airborne equipment (units) into large airborne equipment. The second is that the component testing coverage shall cover all the functions

208 Chapter 3 of the software airborne unit defined in the software preliminary design specification generally and reach 100% coverage based on function call.

3.2.3 Configuration item testing The configuration item testing of flight control system airborne software aims to check the consistency between software configuration items and software requirement specification (including interface requirement specification). The configuration item testing tests the functions, performance, interfaces, and safety under various states defined in the requirements specification and tests the correctness and integrity of the interfaces between software configuration items and configuration items, between configuration items and hardware, and between configuration items and users. Configuration item testing is a test combining software and hardware. When making a testing plan, all external interfaces of the airborne software of the flight control system shall be considered, including interface type, interface data type, and format. 3.2.3.1 Testing plan Content of the configuration item testing plan covers the items division of configuration item testing, the definition of testing types, division of labor of personnel, and the test schedule to guide the whole configuration item testing process. The division of test items is mainly based on the software requirements specification (including the interface requirement specification) and test items 100% cover all the functions, performance, interfaces, and safety requirements defined in the software requirement specification (including the interface requirement specification). 3.2.3.2 Testing methods The configuration item testing of airborne software of the flight control system mainly adopts the black box testing method, without considering the implementation details inside the software. The test environment is generally the target machine 1 exciter mode, including host, flight control computer, bus simulator, and aircraft model simulator. The test environment is shown in Fig. 3.7.

Figure 3.7 Configuration item testing environment of airborne software of the flight control system.

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1. The host completes software loading and test data processing. 2. The flight control computer/simulator runs the airborne software of the flight control system under test. 3. The external environment simulator completes the bus communication and hard-wired interface simulation between the flight control computer and the external system. 4. The aircraft model simulator completes the data operation of the aircraft motion equation. The airborne software configuration items of the flight control system after integration are running in the flight control computer and an external environment simulator is used to provide input data and complete output data collection, so as to verify the airborne software functions and the conformity of the performance and other indexes of the whole flight control system. 3.2.3.3 Testing specification The content of the airborne software configuration item testing of the flight control system shall be subject to the software requirement specification. The design of test examples shall cover the design of functions, interfaces, boundary, allowance, performance, and safety. 1. The design of the examples for functional testing has equivalence partitioning of software functions, including valid equivalence class and invalid equivalence class. The adoption of equivalence partitioning for the design of test examples not only ensures the completeness of software airborne unit functional testing, but also avoids the redundancy design of the testing. 2. The design of the examples for interface testing mainly verifies the data transmission of hard-wired and bus interfaces of the airborne software of the flight control system. 3. The design of the examples for boundary testing is based on the boundary conditions of the flight control system airborne software in input and output spaces, including the minimum value, slightly higher than the minimum value and normal value, slightly below the maximum value and the maximum value, slightly higher than the maximum value, and slightly lower than the maximum value. The value determination is associated with the type and resolution of corresponding data. If it is slightly higher than the maximum value, the maximum value plus a resolution is generally taken. If it is slightly lower than the maximum value, the maximum value minus a resolution is generally taken. In case of several input variables, the test case shall be designed based on “single fault” assumption in reliability theory. In other words, the software failure is basically caused by a single fault. The design of test examples takes a boundary value for a variable and takes a normal value for other variables. 4. The design of the examples for the allowance test is based on the system FLASH storage allowance requirements, processing time requirements, and allowance

210 Chapter 3 requirements in software requirements to determine the system FLASH and processing time allowance size. 5. The design of the examples for performance testing is based on the performance requirements in software requirements to determine relevant testing methods. The performance test of bus transmission rate adopts an oscilloscope for calculation. 6. The design of the examples for safety testing is based on the safety requirements in software requirements to determine relevant testing methods. The watchdog functional testing observes the system running results by adding a delay statement after the watchdog clearing statement. 3.2.3.4 Testing results The test items and requirements listed in the configuration item testing plan have been completed. All the abnormalities in the testing have been properly explained and handled. The design and implementation of the software configuration items have been evaluated. In other words, the configuration item testing of the airborne software of the flight control system has been completed. Configuration item testing requirements have 100% covered all functional, performance, interface, and safety requirements defined in software requirement specification (including the interface requirement specification).

3.2.4 System testing When the airborne software of the flight control system has completely passed the unit testing, component testing, and configuration item testing, then it progresses on to system testing. The purpose of the testing is to check if the complete software configuration items can be well-connected with other equipment and systems under the semiphysical simulation environment and satisfy the functional and performance requirements in the software development task description. 3.2.4.1 Testing plan The system testing plan of the airborne software of the flight control system is based on the software development task description. It mainly focuses on the cross-link between modules and divides test items according to the software requirements proposed to improved software reliability in the software development task description. The division of test items shall fully cover the functional and performance requirements specified in the software development task description. 3.2.4.2 Testing methods The system testing of the airborne software of the flight control system adopts the black box testing method and analyzes and determines if the software meets functional and performance requirements according to the software development task description.

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Figure 3.8 “V-shaped” development process.

The environment used for the software system testing is the real target aircraft environment, that is, the “iron bird” integration test bed of the flight control system. The test equipment includes ground debugging equipment, flight control system testing equipment, avionics system testing equipment, and an oscilloscope. The ground debugging equipment runs software testing tools and ground maintenance equipment software. Together with avionics system testing equipment and flight control system testing equipment, it provides partial simulation data input for software testing and echoes output data to complete the whole system testing function. The test environment refers to the airborne software development and support environment of the flight control system shown in Fig. 3.8. 3.2.4.3 Testing specification The design of examples for system testing of the airborne software of the flight control system is based on the software development task description. The types of test examples are the same as those of the configuration item testing, mainly covering function, performance, allowance, boundary, and safety testing. 3.2.4.4 Testing results All test items and test types defined in the system testing plan and all test examples designed in the testing specification of the airborne software of the flight control system have been completed. The correction or effective explanation of the problems found in the testing marks the completion of the system testing of the flight control system airborne software. The system testing validates the traceability from software code implementation to the development task description as well as the compliance with the functional, interface, and performance requirements specified in the code and development task description.

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3.3 Model-based flight control system airborne software development and testing methods 3.3.1 Overview of model-based development methods With the increasingly mature and extensive use of the fly-by-wire flight control technology, the flight control computer adopts redundancy technology to form multichannel control to improve aircraft safety, and the control law configuration also becomes increasingly complex to pursue excellent flying quality. All these factors promote the growing scale and complexity of the airborne software of the flight control system. The actual model development experience shows that in the development process of the airborne software of the flight control system, the software requirement changes go throughout the whole development process and result in a sharp increase in the workload of software verification, validation, tracking, and analysis. The airborne software of the flight control system is faced with the dual challenges of improving the development efficiency and ensuring high safety. The traditional airborne software development of the flight control system follows the typical “V-shaped” development process. The core process in the whole software life cycle is software code editing. The software requirements and software design made in the design stage are all regarding code writing. The unit testing, integration testing, and system testing in the verification process are all to verify the codes. Therefore the traditional airborne software development method of the flight control system can be called the “code-based development process,” which has the disadvantages including hard verification of the early development process, multiple descriptions (ambiguity), difficult maintenance changes in the later stages, long development cycle, and high cost. For flight control system airborne software of more complex large aircraft of larger scale and higher safety, the traditional software development method can hardly meet the requirements. With the development and improvement of software engineering and the emergence and application of new design methods, software development methods are constantly changing. Model-based software development methods provide a new solution for the airborne software development of flight control system. For the use of model-based airborne software development method, the traditional “V-shaped” process is shown in Fig. 3.9 and the “Y-shaped” development process is shown in Fig. 3.9. The main difference is that the focus shifts from code writing to model establishment. The code can be generated automatically through the model. If a certified code generator is used, testing time and cost can be saved. Model-based development tools have developed rapidly in recent years. In the field of embedded software, they are mainly SCADE Suite, MATLAB Simulink, Rhapsody, TAU,

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Figure 3.9 “Y-shaped” development process.

RoseRT, and STOOD. Among them, SCADE Suite has been widely used in European aerospace, high-speed rail, energy, and other high safety fields for many years due to its compliance with the high safety requirements in key application fields and its simple and easy use. In recent years, it has been widely used in many high-safety fields in China, such as aerospace, with the following characteristics: 1. Adopting the formal design method, it ensures design integrity and unambiguity with strict mathematical theory. 2. Adopting the formal modeling method, it is simply to learn and use. It uses “modelbased” development method rather than the traditional “code-based” development method, reducing the workload of developers. 3. It generates high-quality C code that meets A-level certification of DO-178B and does not need unit testing automatically, ensuring strict consistency between the codes and model. 4. Adopting efficient and reliable simulation and testing methods, it can have quantitative verification of each development process. The SCADE model is implemented through the hybrid of data flow diagram and state machine diagram, which is similar to the Simulink and Stateflow implementation methods commonly used in control law algorithm design. Therefore SCADE is widely used in the control law software design. The basic development process is shown in Fig. 3.10.

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Figure 3.10 Basic SCADE model development process of flight control system control law software.

Figure 3.11 Traditional software testing.

3.3.2 SCADE model testing and verification features The software verification process includes three parts: review, analysis, and testing. In terms of the software development process, the development process that we are familiar with includes requirements analysis, preliminary design, detailed design, and code editing. Although the review and analysis can ensure the consistency between the preliminary design and requirements analysis and between the detailed design and the preliminary design, they cannot quantitatively measure the completeness. Therefore unit testing is used to ensure the consistency between code and detailed design, while code coverage analysis is conducted to measure the completeness of unit testing (Fig. 3.11). When SCADE modeling is used to make preliminary design and detailed design based on the model development method, the software development risks including unclear requirements, different meanings, or ambiguity caused by the natural language used in the document description of the traditional development method can be effectively avoided. Then, the analog simulation (black box testing) provided by SCADE is conducted to verify if the functions are realized and MTC model coverage (white box testing) is used to analyze and measure if the analog simulation is complete, so as to achieve the quantitative measurement of the verification. In addition, since the KCG code generator can ensure the

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Figure 3.12 SCADE model-based software testing.

consistency between the generated code and the model, the unit testing that is required by the code-based software development method can be completely omitted and code static testing can be changed to model static testing accordingly (Fig. 3.12).

3.3.3 SCADE software testing process 3.3.3.1 Testing strategy The best strategy of SCADE-based model testing is to have simulation testing of each functional node on the host or test bed as early as possible. In other words, once the SCADE model is established, the simulation verification of the model shall be started as soon as possible. Based on previous model engineering experience, the following testing strategies are proposed: 1. Static check is the first step of the whole SCADE testing, including checking the syntax and semantics (completed by SCADE editor), having walkthrough to check the rationality of the naming rules of variables and parameters in the model, having walkthrough to check if each functional node/module has detailed and specific description, having walkthrough to check if the model can trace to last layer of specification, and having walkthrough to check if the SCADE model structure and algorithm are reasonable. 2. Check if the examples for simulation testing are designed based on requirements (detailed design documents of flight control system or control law design documents). 3. If new functions or modifications are added to the model, regression testing shall be conducted for verification.

216 Chapter 3 3.3.3.2 Testing process One of the biggest differences between model-based software testing and traditional software testing is analog simulation. For SCADE model testing, analog simulation is completed with the help of the SCADE simulator. The SCADE simulator is a visual debugging environment whose essence is to simulate and debug the model by executing the C code generated by KCG. In SCADE software testing, simple analog simulation cannot achieve all the testing objectives and it still needs to be combined with other testing methods. The testing methods include “open loop” and “closed loop.” These two methods complement each other and are both dynamic testing methods. The former is equivalent to unit testing and component testing, and the latter is equivalent to configuration item testing and system integration testing (Fig. 3.13). 3.3.3.2.1 Open-loop simulation

1. With the use of the SCADE simulator and SCADE MTC (SCADE model coverage analysis), open-loop simulation can be conducted to verify the interpolation, first-order or second-order transfer function, trigonometric function, saturation, and the public

Figure 3.13 Principle of SCADE software testing process.

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functions constructed to reduce the complexity of model during the construction of control law module nodes before a closed-loop simulation. By inputting several sets of test examples, analog simulation (functional test, interface test) and coverage analysis of SCADE nodes can be completed. 2. Integrate the generated codes into the flight control system airborne software, compile and generate executable object codes, and load them to the flight control computer or target CPU board card. The external test equipment only simulates the interfaces of excitation signals necessary for system software (such as bus data, hard-wired signal, analog and discrete quantity) rather than the dynamic flight simulation environment. At this time, the internal and external interfaces, control law polarity and transmission ratio, control logic, partial BIT, and redundancy management functions of the flight control system can be tested. 3.3.3.2.2 Closed-loop simulation

The flight control system is a strong real-time closed-loop control system and the closed-loop simulation test is the best way to verify the software. The test case for closed-loop simulation should be based on requirements and they should better be provided by the flight control system designer or the control law designer, or imported from the Simulink model established in the control law design simulation. In addition, in order to ensure the high reliability of tests, these test examples shall be reused in the following test processes: 1. Model-in-the-loop. The document of test examples can be loaded to the SCADE model and then the SCADE simulator and SCADE MTC are used to complete the analog simulation (functional testing and interface testing) and coverage analysis. Or the SCADE model is encapsulated as S-Function and is used to substitute the corresponding control law module in Simulink as the simulation testing in the Simulink environment can only complete the functional test but cannot achieve coverage analysis. 2. Software-in-the-loop. On the host, the source code of control law software (generated by SCADE KCG) is compiled and executed and the component testing of control law software is conducted by using a desktop development environment (such as VC1 1) to indirectly verify the SCADE model. 3. Hardware-in-the-loop. Development tools (such as Tornado) are integrated through embedded software, and executable object codes of the flight control system airborne software (including control law software) are cross-compiled, generated, and loaded or filled into the flight control computer. The excitation signals required by the flight control computer are dynamically provided by external simulation or real equipment.

218 Chapter 3 3.3.3.3 SCADE model testing To conduct a complete testing, there shall be at least three elements, that is, test environment, test examples, and coverage analysis. 3.3.3.3.1 Test environment

Available test environments include the SCADE simulator (provided by SCADE), host (PC), and target machine (flight control computer). 1. The SACDE simulator can realize functional simulations at any level and use MTC (model coverage analysis) in the same environment to verify the completeness of the SCADE model simulation. 2. The control law software codes generated by SCADE KCG (SCADE code generator) can be integrated on a PC machine. The control law algorithm is composed of some mathematical models and logical models and the software is independent from the platform. One method is to conduct integration testing with a desktop development environment, such as Visual C1 1. By loading test examples and running control law software, the SCADE model is indirectly verified. The other method is to convert the SCADE model into S-Function and run this S-Function in Simulink to verify the functions of the SCADE model. 3. Software and hardware integration testing is carried out on the flight control computer, which is a necessary test. The conformity of the software can only be confirmed after it runs in a real environment. 3.3.3.3.2 Test examples

After the test environment is established, test examples shall be written according to the requirements. The reason why the testing shall be conducted based on requirements is that this kind of testing is the most efficient in finding errors. There are two types of test examples developed according to requirements, that is, normal range test examples and robustness (abnormal range) test examples. The test case are applicable to unit, integration, system, and regression testing. Traditional unit testing is omitted in the SCADE model testing and it is replaced with analog simulation. However, the testing process and the basic principles followed by test examples are the same as traditional software testing. It is suggested that the design and implementation of the examples for control law software testing should focus on the following matters: 1. Test examples shall be designed for the normal operation of the control law software and also the state transition that is not allowed by the software requirements. 2. For the large number of time-related functional modules used in control law, such as filter, integrator, and delay, the test case shall iterate over the code many times, the

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characteristics of the module in the context shall be checked, and test examples for the arithmetic overflow protection mechanism shall be designed. For the conditions that the calculated value is often used for the cycle count in the calculation of the control law, test examples that may exceed the cycle count range should be designed. As the description of state machine is widely used in the calculation of control law (especially the automatic flight control law), test examples shall be designed for every valid trigger and invalid trigger under each state. For real and integral variables, test examples shall be designed based on equivalence classes and boundary values of legal and illegal values. Besides, every valid boundary and invalid boundary shall have at least one test example and every valid equivalence class and invalid equivalence class shall have one test example. In the process of testing, testing personnel shall strictly follow the test case, test items, and test steps, record the testing results, and finally form the test report. For different testing stages, test examples used for integration testing, system testing, and regression testing shall be clearly specified when designing the test case and testing personnel shall not change them at will. As the testing load is huge, in order to improve work efficiency, automatic testing tools or plug-ins should better be used and it will be good if some test examples can be automatically generated. After the testing, the testing results should be evaluated and some quantitative results with regard to the completion of software testing and testing quality should be obtained, such as testing coverage, grade of errors found, and testing pass rate. It should be analyzed to determine if the defects are caused by missing testing or recurrence. Missing testing indicates that the test case are not complete and relevant test examples shall be immediately supplemented. Recurrence indicates the system or software has problems and careful troubleshooting is required. If the defect is caused by recurrence, the troubleshooting workload is generally huge.

3.3.3.3.3 Model coverage analysis

As a software verification technique, coverage analysis aims to evaluate if the software testing process is complete, to ensure the quality of test examples and thus indirectly improve the quality of software products. The role of model coverage for the model-based development method and traditional development method is similar and it is a major means for software quality assurance. The purpose of SCADE model coverage analysis is to evaluate the results of SCADE model simulation verification. The SCADE model is different from codes in that the model describes functions while codes describe the implementation of the functions. When defining model coverage criteria,

220 Chapter 3 in addition to the model’s structure and the actions to capture the model (if it is activated during operation), the model’s functional coverage shall also be considered. SCADE conducts model coverage analysis through MTC, mainly covering the following three aspects: 1. To verify that all system requirements have been realized through the SCADE model (coverage of system requirements by control law software). 2. To verify that the SCADE model has no unexpected functions (structure coverage of control law software SCADE model). 3. To complete model coverage analysis through the test case based on high-level requirements, and test examples shall be supplemented for derived requirements. The SCADE MTC process is shown in Fig. 3.14 (please refer to SCADE user manual for specific use methods): 1. Test examples Test examples are developed according to the control law requirements and the source of inherent errors in the software development process, including normal range test examples and robustness (abnormal range) test examples. 2. Coverage criteria a. The requirements coverage and model structure coverage (model branch coverage, decision coverage, MC/DC coverage) of control law software SCADE model must reach 100%. b. SCADE brings coverage criteria of DC, unique method MC/DC, and shielding method MC/DC. For the SCADE model of control law software, it is recommended to adopt the shielding method MC/DC model instrumentation. c. Basic coverage criteria for state machine: i. State coverage criteria. The generated test examples can test each state, which is similar to the statement coverage in structural testing. This is the weakest test coverage criterion that can be met the most easily and requires the test case the least.

Figure 3.14 SCADE MTC process.

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ii. Transformation coverage criteria. Every transformation in the state machine shall be tested at least once, which is similar to decision coverage in a structural testing. iii. Transformation coverage criteria. For every entry/exit transformation pair in the state machine, both of them shall be tested. iv. When executing the state machine coverage criteria, it should be noted that the data flow coverage within the state only needs to consider the situation when the state is activated, and the coverage in the migration only needs to consider the situation when the migration is activated. SCADE RM (SCADE requirements management tool) or DOORS (requirements management tool) is used to validate the corresponding relation between the flight control system airborne software design document or control law design document and test examples and ensure that all requirements have corresponding test examples, so as to get coverage analysis data 100% based on requirements finally and obtain analysis reports. The coverage can be obtained by executing the simulation program. After the simulation, if the coverage rate does not reach 100% when analyzing the coverage data, the reasons may be as follows: a. The test case are not complete. b. The system requirements are inadequate. c. One branch of the requirement is never reached (e.g., the flight control computer has a Reset function, but the Reset is never needed during flight). d. Dead requirements (e.g., y 5 x 1 2.0; output 5 (x . 0.0) and (y , 1.0)). According to the reasons found in the coverage analysis, add test examples and description to a branch that is not needed according to the system requirements, so as to achieve 100% coverage. Finally, reports are generated according to the coverage activities.

3.3.3.3.4 Code integration testing

After the model is validated, the KCG code generator automatically converts the control law SCADE model into engineering-oriented source code and ensures that the model is consistent with the source code. As the software developed based on the SCADE model still needs to be integrated with the software developed with the traditional mode, the verification of the software integration and input/output is mainly conducted through analysis and testing. The input and output data of Simulink can be edited into a script file and loaded to testing engineering established in the desktop development environment (such as VC1 1). Through the script, the input data can be read conveniently and the software under test can

222 Chapter 3 be executed. Then, the analysis results will be compared with the output results provided by Simulink and the consistency between the software and Simulink model can be verified. 3.3.3.3.5 Software and hardware integrated testing

The software and hardware integrated testing should be carried out in a target machine environment. The testing process is the same as the traditional system testing process. For specific methods, please refer to relevant content in section 3.2.

3.4 Software whole life cycle support environment 3.4.1 Basic requirements of environment The software whole life cycle support environment is a software and hardware integrated platform that supports the design, verification, and maintenance of software products. The tool set includes the software tools that support the software development-related processes, activities, and tasks. The environment integration mechanism provides unified support for tool integration and software development, maintenance, and management. Through the software whole life cycle support environment, various activities of software development can be supported, thus greatly improving the efficiency and quality of software development. According to the construction and use experience of software whole life cycle support environment in previous model development, the following requirements are put forward: 1. The environment shall be able to cover all relevant work in the whole life cycle of the software, such as design, testing, verification, maintenance, and configuration management. It shall be ensured that the software requirement analysis management, software design, code editing, integration, compiling, downloading, and debugging can be completed in the software design stage; the unit testing, component testing, configuration items (integration) testing, and system testing can be completed in the software testing and verification stage; and configuration management can be conducted for the whole life cycle of the software. 2. The environment shall integrate the differences of the testing and verification activities in process, function, and platform from requirements to model design, source code generation, and target code loading, so as to maintain the connection and fluency of the testing and verification activities in this process, improve the degree of automation of testing and verification activities, and make continuous iterative testing and verification possible. With features including function customization, high integration, simplified operation, and unified data, it can achieve the integration effect through the integration of process, function, and platform. 3. It shall support the characteristics of airborne software of the flight control system. Due to the particularity of the airborne software running environment, the environment must support some special characteristics of aircraft equipment, such as aviation bus

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protocol, memory usage specification, programming specification, and compilation specification. In addition, the high safety and high reliability requirements of the airborne software of the flight control system require the software safety level mostly at class A or B. Therefore in addition to the static testing and dynamic testing of the code, coverage analysis shall also be conducted for the target code of class A software. 4. In addition to the testing conducted directly on the target machine in the software testing stage, simulation may also be conducted for the model in the design stage. The requirements for different software simulation tools are different. As the testing may be conducted on the target machine again because the test results are affected by the inconsistency between the tester and target machine, the capability of the simulator and its difference with the real airborne unit should be considered when establishing the test environment, and the effects on the test results shall be analyzed. 5. With regard to the development progress, the platform shall be planned synchronously at the beginning of the life cycle of the software project and the environmental construction shall be implemented step by step according to different stages of software development. Then, the environmental construction planning shall be written in the Software Development Plan. The software environment construction shall cover all the work in the whole life cycle of the software.

3.4.2 Environment architecture The main objectives in the whole life cycle of the software shall include software design, software testing, software delivery, maintenance and upgrading, and the corresponding process management. The whole life cycle environment of the software enables software designers, software testing personnel, software upgrade and maintenance personnel, and process management personnel to work in a unified way and supports the work of different personnel in different stages of software development, effectively reducing the cost and time of software testing. The principle of architecture of the software whole life cycle support environment is shown in Fig. 3.15, with constitutional parts as follows: 1. Test object: target system used for software development and testing. It refers to airborne units that integrate software, such as the flight control computer. The airborne computer can have software downloading, interactive debugging, and software operation through mode selection. 2. Simulation excitation system. All external devices that are cross-linked with the test object are simulated and software and hardware interfaces consistent with the test object are provided. 3. Data acquisition/storage system. The output data of the test object is acquired and stored as the data basis for software debugging and testing. The data source is mainly test objects and the simulation excitation system.

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Figure 3.15 Principle of architecture of the software whole life cycle support environment.

4. Software development/debugging system. Software editing and compiling functions and software and hardware interfaces consistent with ground maintenance and upgrading of airborne units are provided to complete software downloading and interactive debugging. 5. Software testing/analysis system. The ability of software unit testing, integration testing, system testing, and test data analysis is ensured. The functions including generation of test examples, preparation of test report, and analysis of test data and management of test process in each stage can be completed. Test examples can be applied in test objects through the simulation excitation system and test results can be acquired through the data acquisition/storage system. 6. Software process management system. Engineering management in the whole software life cycle can be completed, such as development process management, quality process management, and configuration management.

3.4.3 Environment composition and functions The support environment can simulate the working process of the fly-by-wire flight control system airborne software, automatic flight control system, and high lift system to realize the testing and verification of the flight control system computer and software. By establishing a peripheral excitation environment required by the test system, providing a hardware working platform for the test system and combining it with the test system software, a complete testing and verification platform can be established to finish the design, debugging, solidification, testing. and verification of airborne software of the fly-by-wire

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flight control system, automatic flight control system. and high lift system. The environment shall have following functions: 1. 2. 3. 4. 5. 6. 7. 8.

Software development and testing function. Software configuration management function. Software requirements management function. Excitation simulation and signal injection function. Test data and test examples input and testing process control functions. Test data acquisition and recording function. Test data display and analysis function. Virtual control and display function.

The support environment is composed of the flight control test system and software development test environment. The flight control test system mainly includes the public test system, the fly-by-wire flight control system airborne software, the automatic flight control system, and the high lift system. The software development test environment mainly includes the signal test board of software development computer, requirements management computer, configuration management computer, test data management computer, PFC DIF equipment, AFCC DIF equipment, FSECU DIF equipment, fly-by-wire flight control computer, automatic flight control computer, and flaps and slats control computer. The composition of the support environment is shown in Fig. 3.16.

Figure 3.16 Flight control system airborne software development and support environment.

226 Chapter 3 The support environment can be used for the independent software development test of PFC, AFCC, and FSECU, as well as for their system level cross-linking test. While meeting the test requirements above, some equipment will be shared to achieve the optimal design of the test system. The backbone communication network of the test system adopts a reflective memory network to ensure the real-time data interaction of the system. 3.4.3.1 Composition and functions of the public test system The public test system provides functions including the avionics simulation for PFC, AFCC, and FSECU, aircraft body and dynamics simulation, aircraft control system simulation, public test system control, virtual display of avionics instrument, and test system data acquisition. Specifically, it covers the public test control unit, avionics instrument unit, data acquisition server, avionics and fly-by-wire flight control system airborne software simulation unit, and signal conditioning unit. 1. Public test control computer. Composed of a high-performance computer and test control software, it can download and manage the running of the avionics system simulation excitation model and the aircraft body and dynamics simulation excitation model, complete the editing and execution of automatic test examples, and inject bus and nonbus faults through the model. 2. Data acquisition service computer. Composed of the data acquisition lower computer software, data acquisition server, and optical fiber reflective memory network, it can acquire and store the analog quantity, discrete quantity, and bus data in the test process of PFC, AFCC, and FSECU through the optical fiber reflective memory network. 3. Flight simulation unit. Through the aircraft body and dynamics simulation excitation model, it can realize the aircraft body and dynamics simulation. 4. Avionics simulation unit. It simulates the avionics system interface and function and drives the running display of avionics virtual instruments. 5. Virtual instrument. Composed of a high-performance computer and avionics instrument simulation software, it simulates the PFD virtual interface and EICAS virtual interface of aircraft avionics system. 6. Control simulation unit. It simulates control devices such as control column and related switches, flaps and slats control handle, flaps and slats override control panel, horizontal stabilizer control handle, primary flight control panel, trim control panel, horizontal stabilizer cutoff control panel, and brake control handle and takes the data as the input to the flight control system. 3.4.3.2 Composition and functions of flight control test system The flight control test system consists of a fly-by-wire flight control test system, an automatic flight control test system, and a high lift control test system. It cross-links with the public test system to complete the simulation test of the whole flight control system.

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Meanwhile, the fly-by-wire flight control system airborne software, the automatic flight control system, and the high lift system can all complete the independent test of the subsystem. 3.4.3.2.1 Test control computer of flight control system

The test control computer of the flight control system includes the test control computer of the fly-by-wire flight control system airborne software, the test control computer of automatic flight control system, and the test control computer of the high lift system. Mainly consisting of a high-performance computer and test control software, it can complete the simulation excitation of the test of the three subsystems of the flight control system, the model downloading, running management, editing and execution of automatic test examples, and can also inject bus and nonbus faults through the model. 3.4.3.2.2 Simulation excitation unit

The simulation excitation unit consists of the PFC simulation excitation unit, AFCC simulation excitation unit, and FSECU simulation excitation unit. By running the simulation excitation model, the software simulates the functions and interface characteristics of internal and external cross-linking devices of PFC, AFCC, and FSECU computers to provide a simulation excitation source for airborne computers. 3.4.3.2.3 Airborne computer and simulated target machine of the flight control system

The flight control test system provides a real airborne computer and simulated target machine. The real airborne computer provides a real airborne environment for software development and debugging and software testing. The simulated target machine provides a simulation interface and software running environment consistent with the real airborne computer. The test control computer of subsystems of the flight control system can switch between the real airborne computer and the simulated target machine so as to constitute a test environment of the real airborne computer or the simulated target machine. The simulated target machine can be developed and delivered for use ahead of the real target machine, so that it can provide an environment for software development, debugging, and testing at the early stage of software development and support the rapid design and verification of software. 3.4.3.2.4 Signal adapter unit

Consisting of a signal conditioning board card, a signal conditioning box, and connecting cables, the signal adapter unit provides the PFC signal adapter unit, AFCC signal adapter unit, and FSECU signal adapter unit for the conditioning of discrete quantity and analog signal, so that the test system signal can be matching with the signal of the automatic flight control computer and flaps and slats control computer.

228 Chapter 3 3.4.3.3 Functions and composition of software development and testing environment The software development and testing environment consists of the signal test board of the software development computer, requirements management computer, configuration management computer, test data management computer, PFC DIF equipment, AFCC DIF equipment, FSECU DIF equipment, fly-by-wire flight control computer, automatic flight control computer, and flaps and slats control computer. Computers with different functions have data communication through the Internet to realize the resource sharing of the whole software development and testing environment and complete the collaborative work of software design, testing, and software process management. 3.4.3.3.1 Software development computer

The software development computer completes all the software design work and supports the software, documents, and code editing in the stages of system requirements analysis, software preliminary design, software detailed design, software code editing, unit testing, software integration, and component testing. 3.4.3.3.2 Software process management system

The software process management system includes the software requirements management computer and configuration management computer. The software requirements management computer is installed with DOORS requirements management tools and the configuration management computer is installed with APTECH JBCM7 software configuration management tools to complete the configuration management work in the whole life cycle of the software. 3.4.3.3.3 Test data management computer

The test data management computer has unified the management of the test example data and test results data in each stage of the software and it can prepare test reports based on these data. 3.4.3.3.4 DIF equipment

According to the requirements of each subsystem of the flight control system, DIF equipment is installed and configured with the relevant software compiling and downloading tools to debug and download the software. 3.4.3.3.5 Signal test board

The signal test board is directly connected with the signal adapter unit. In the process of software debugging and testing, the testing of and signal injection to the nonbus interface of the flight control computer can be directly conducted through the signal test board.

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3.4.4 Construction process of software whole life cycle support environment The software whole life cycle environment covers the whole process of software development, testing, and maintenance. The basic principles for the environment construction are as follows: 1. The environment construction shall be planned at the software development planning stage. 2. The environment construction shall follow closely the software development process so as to meet the environmental requirements in each stage of software development. 3. Incremental development mode shall be adopted, which meet the different requirements in each stage of software development, and ensure the development progresses to the maximum extent. The whole life cycle environment construction of software can be divided into the planning stage, implementation stage, environment evaluation stage, and use and maintenance stage. 3.4.4.1 Planning stage The software development environment shall be planned according to the software/ hardware requirements specified in the software development task description. It mainly includes the adopted requirements development methods and tools, design methods and tools, coding languages, compilers, linking and compiling programs, and loading programs, and the hardware platform of the tools used. According to the software development process, a clear environment construction plan shall be made in stages, reasonable cost and workload estimation shall be carried out, risks shall be assessed, and focus shall be given to the development and delivery schedule of the airborne computer used for the test. 3.4.4.2 Implementation stage The implementation of software whole life cycle environment is a gradual iterative process and the incremental development shall be executed step by step. For example, in the stage of software coding, at least the construction of a pure software development environment shall be completed. In the stage of unit testing, the configuration of testing tools and the construction of a pure software testing environment shall be completed at least. In the stage of configuration item testing, the construction of a target machine simulator environment shall be completed at least. And in the stage of system testing, the construction of the target machine and system cross-linking environment shall be completed. 3.4.4.3 Environment evaluation stage The software verification environment (including tools) shall be evaluated to ensure adequate, correct, and effective software verification. The software verification environment includes the review environment, analysis environment, and test environment and different

230 Chapter 3 evaluation methods are adopted for different verification environments. The software review environment refers to the various forms used in the software review process. The content in the forms shall be verified. Software analysis and test environment generally adopt verification tools software, test equipment, target computer, target computer simulator, or host computer simulator to analyze, test, and verify the software. The evaluation methods for the analysis environment and test environment are as follows: 1. The selection of software verification tools shall meet the requirements of relevant national standards, military standards, and DO-178B. If the verification activities specified for the software are canceled, reduced, or automatically carried out due to the use of a certain verification tool, the verification tool shall be certified. 2. The certification of software verification tools shall be subject to the tools’ compliance with relevant operating requirements under normal working conditions. The operating requirements shall include: a. Description of the tool’s functions and technical features. b. Information for users, such as installation guide and user manual. c. Description of the tool’s operating environment. d. A target computer, a target computer simulator or a host simulator can be used to complete the software testing process. The following two points shall be considered: i. The difference between the target machine and emulator or simulator and the influence of these differences on error detection and function verification ability and the errors not detected due to the difference between the target machine and the emulator or simulator shall be detected in other software verification activities and explained in the software plan. ii. For purchased software verification tools, the certification certificate of the verification tool issued by an authoritative institute and certificate of service history of the verification tool shall be provided and approved by the user. 3.4.4.4 Use and maintenance stage In the use and maintenance stage, the most important work is to have the environment change and upgraded according to the maintenance and upgrading of the software.

3.5 Software safety and reliability test 3.5.1 Safety and reliability The flight control system is a typical complex embedded type real-time control system. The airborne software of the flight control system has the characteristics of complexity,

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timeliness, high safety, and high reliability. As the airborne software quality of the flight control system directly affects the flight safety of the aircraft, the safety analysis, design, and verification of the airborne software of the flight control system shall be carried out to effectively improve the quality and safety of the software. Military Software Quality Metrics (GJB5236—2004) defines software reliability and safety as follows: 1. Reliability: it is an internal and external quality attribute of software related to software defects, faults, and failures and will not cause system failures. 2. Safety: it is a feature of software use quality that particularly emphasizes that the safety is not only related to software, but also to the whole system, as well as to system risks, hazards, and accidents, and it will not cause a system accident. Software errors are the source of software safety and reliability problems. Reliability is only analyzed down to failure, but safety goes down to accident, the final link. Whether a failure would cause a hazard and ultimately an accident is what distinguishes software reliability from safety. If software error identification, error traps, fault tolerance, or error elimination is not directly related to a system hazard, it is usually not related to safety, but to reliability or availability. The causes of reliability and safety problems are shown in Fig. 3.17. There is no definite subordination relationship between safety and reliability. Safety does not ensure absolute reliability and reliability does not ensure safety either. Reliability emphasizes that normal functions must work while safety emphasizes unexpected hazards will not happen. Software in fields like aerospace should pay more attention to safety.

Figure 3.17 Causes of reliability and safety problems.

232 Chapter 3 As the causes of software reliability and safety problems are the same, there is no clear boundary between software reliability and safety technologies for software developers. Therefore the internal quality of software shall be paid high attention, including the design and programming criteria and specifications.

3.5.2 Safety analysis and testing According to the requirements of DO-178B, software safety analysis must be carried out in the whole software life cycle. The specific work requirements in each stage of software development are shown in Fig. 3.18 and the process of software safety analysis is shown in Fig. 3.19. Software failure modes are mainly identified in the following categories: 1. Functional logic failure: functional coverage analysis, functional integrity, rationality and adequacy analysis, hierarchical traceability analysis. 2. Relationships failure: control/data migration failure, functional combination failure analysis, functional constraint analysis. 3. Interface failure: external interface, internal interface, information transfer. Examples of common software failure modes are shown in Table 3.3.

3.5.3 Reliability analysis and testing The software reliability test is a random test conducted to ensure and verify the reliability of software. The results of the software reliability test are used to evaluate the reliability level of software and to verify whether software products meet the reliability requirements. For most airborne software, DO-178B/C series standard is adopted. In the process of software development, conducting the development in strict accordance with the requirements on development objectives corresponding to the software safety level can

Figure 3.18 Safety analysis work in life cycle.

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Figure 3.19 Software safety analysis process.

obtain the development confidence of corresponding reliability and safety. However, it is also proposed in the airworthiness regulations that for some software alternative methods can be used to illustrate the compliance of airborne software with the confidence of corresponding safety level, including exhaustive testing, service history, reliability testing, and multiversion nonsimilarity testing. For many newly developed airborne software of flight control systems, it is difficult to have exhaustive testing and the service history cannot be described as it is hard to test all states. However, airborne software of multiple versions of the flight control system is limited by resources and development cycle. The relationship between software reliability testing and software testing is shown in Fig. 3.20. It is a kind of random black box testing method, belonging to software dynamic testing. Compared with traditional software testing activities, the reliability test can use the same test environment and test result analysis method as software testing and has its own unique software testing data generation method and software reliability evaluation technology.

234 Chapter 3 Table 3.3: Common software failure modes. Category

Failure modes

Input processing category

Failure in time sequence Failure in content

Output processing category

Failure in time sequence Failure in content Failure in processing process

Software fails to identify or mask the data sent in advance Software fails to identify or mask the data sent late Software receives data too slowly, resulting in data loss The software with failure does not mask data not within the normal range Software does not mask the data with out-of-tolerance accuracy Software does not mask deficient data Software does not mask redundant data Software does not mask reversed/misplaced data Software outputs data ahead of time Software outputs data late the software with failure outputs data with accuracy out of tolerance Software outputs data not within the normal range The software with does not complete relevant function Software completion function error

Figure 3.20 Relationship between software reliability testing and software testing.

Software requirements and users’ use of the software are reflected in the test data and the quantitative measurement in software reliability testing is reflected in the evaluation. The software reliability test is generally divided into the software reliability placement test, software reliability verification test, and software reliability growth test. The methods

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adopted in the three tests are basically the same and the only difference is that in the specific test operation, the software reliability growth test needs to modify the problems in the software while testing. The reliability test steps of the airborne software of the flight control system are the same as those of the general software reliability tests, as shown in Fig. 3.21. To modify the failure synchronously in the process of the software reliability test, the defects need to be corrected and modified so as to realize the growth of software reliability.

Figure 3.21 Reliability test steps of airborne software of the flight control system.

236 Chapter 3 In view of the characteristics of the airborne software of the flight control system, the following factors must be considered for the reliability test with the steps above for the airborne software of the flight control system: 1. Because of the special nature of the airborne software reliability test of the flight control system, it is necessary to carry out the reliability test for the software and build a special software environment for the software reliability test. Generally speaking, since the purpose of the software reliability test is to evaluate the characteristics of the software in the final use process, the operating environment of the software reliability test shall be close to that of software system testing. In other words, the input and output of the airborne software of the flight control system shall be tested when it is simulated that the pilot operates the aircraft. The test environment of the software reliability test can be built by relying on the engineering simulator and the “iron bird” integrated test bed of the flight control system. In the test of the engineering simulator, the typical software operation profile can be designed to simulate the operation input of the pilot. The software failure set can be obtained through the qualitative or quantitative analysis of the test data of the engineering simulator. The “iron bird” integrated test bed of the flight control system can also verify the software. In the process of analyzing the data of the “iron bird” integration test of the flight control system, the faults caused by software failures can be identified. The object of the test may be an intermediate product or a final product of the airborne software of the flight control system, depending on the requirements. 2. From the perspective of constructing a software reliability test profile, the airborne software of the flight control system is used to execute aircraft control tasks and the running of the software is in a strict cycle. From the power supply connection for the computer to the control of the system’s closed-loop running, the software running results are continuous and circular. The test profile of the airborne software of the flight control system is a combination of several operations, which is more complex than the general test profile. Since the state of the airborne software of the flight control system in the running process is related to the state information input received by the aircraft and the randomness of the input is very large, the state input of the aircraft, such as altitude, speed, and attitude must be considered when constructing the test profile. 3. Airborne software reliability test results of the flight control system are ultimately used to evaluate the software reliability, so in the process of the reliability test, the data required for the software reliability evaluation shall be sampled and acquired in a targeted manner. In addition, for some intermediate process data, they shall be acquired according to the data requirements for the software reliability test. For example, statistics of defect density and fault density can be made to provide reference for software reliability evaluation.

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For embedded airborne software with high safety and reliability requirements for the flight control system, the reliability test and evaluation is very important and significant. However, the engineering practice is difficult and there are mainly the following problems: 1. As the airborne software of the flight control system has direct and critical impact on the flight safety of the aircraft and its failure may lead to system function failure and even danger to the aircraft and a catastrophic failure accident, the software safety level is usually defined as class A or class B. In the process of software development from planning, design, coding, testing to use, and maintenance, the software engineering methods must be strictly observed. After flight control system airborne software passes the unit testing and configuration item testing, it has a low defect rate and a high reliability. After it passes the system testing and is installed, it has a very low failure rate. According to the empirical value of reliability data, when the failure rate of software is one order of magnitude higher than that of hardware, the failure rate of the software is less than 1025/flight hour. The statistical characteristics of software failure are completely different from the failure data of general software. It is impossible to obtain complete failure information in a short time. Making an estimation with incomplete failure information may cause a large deviation and it requires long-term and massive data accumulation. 2. The idea of the software reliability test is to carry out random testing according to the statistical law of users’ actual use of the software. Different from traditional software testing, it introduces the concept of the test profile. Software usage is modeled in the form of a test profile and then use by users is simulated and taken as the test input. The purpose of constructing the operation profile is to generate software reliability test data and use the data to support the execution of the software reliability test. The operation profile is defined according to the input value of the system based on the time distribution or the probability distribution within the input range to describe the actual situation of the software. However, the development of the operation profile depends on the functional description of the software and software reliability engineers’ mastery and understanding of the use of the software. The operation profile developed by different software reliability engineers may be different, but it must reflect the actual use of the software objectively and effectively. To construct a test profile meeting requirements, for flight control system airborne software, the work of the flight crew, flight control system designers, software development, and testing personnel should be closely combined to ensure that the flight crew transmits the software functional requirements, performance requirements, and operation requirements to the software reliability engineering personnel accurately and completely. However, this requirement is hardly to be strictly implemented in engineering practice. 3. The airborne software of the flight control system of large aircraft will be installed after the system testing generally in the aircraft development cycle. After the delivery of the

238 Chapter 3 integrated product, it is required that preliminary software reliability indexes should be obtained. If the software reliability test is to be conducted, it shall be simultaneously conducted with the software testing and verification. Therefore adequate model development time and resources must be guaranteed.

CHAPTER 4

Flight control system control law and the flying quality evaluation test The aircraft control system has developed from simple to complex and from mechanical to digital. Before 1940 the aircraft control system was basically composed of some simple mechanical components (cable, pulley, pull rod, rocker arm, etc.), with the advantages of portability, reliability, and low cost. During the World War II period, as the size and weight of aircraft and the flight speed and altitude increased, the control plane hinge moment also increased. As it is difficult to handle the control surface merely with the pilot’s physical strength, the hydraulic booster was introduced to the control system. It can be said that the adoption of the hydraulic booster in the control system was the first major change in the development of the aircraft control system. The boost control cuts off the direct connection between the control column and the aerodynamic load on the control plane, reduces the pilot’s control force, and gives full play to the maneuverability of the aircraft. However, the boost control also increases the complexity of the system. In order to make the pilot feel the force while operating the control plane, an artificial sensor device must be used. The original artificial sensor device was a simple single-gradient spring, which gradually developed into a multigradient spring with preload, a hydraulic load mechanism, and a velocity pressure controlled load mechanism to make the change of rod force and displacement gradient appropriate under different flight states or different displacements of the input rod. In the nonboost control system, it is customary to use a tab to remove the force acting on the control column in a balanced flight. In the boost control system, an electric-driven tab effect mechanism is used to substitute for the tab as it is unable to exert its efficiency in the boost control system. With the expansion of the range of the flight envelope, the aerodynamic characteristics (such as stability and control plane efficiency) of the aircraft when flying at different altitudes and speed varied drastically. To provide the pilot with relatively consistent control characteristics, many complex mechanisms (such as nonlinear mechanism and variable transmission ratio mechanism) further complicate the control system. After 1950 the design of aircraft developed toward high altitude and high speed; the flight envelope expanded sharply and the dynamic quality of aircraft flight deteriorated (the most obvious phenomenon is that the damping of aircraft pitch control and the stability decreased). Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00004-8 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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240 Chapter 4 As a result, the flying quality under certain flight conditions was hardly accepted for pilots and the precision control ability of the aircraft was affected. These problems caused by the aerodynamic layout could no longer be solved by simply improving the design of the machinery control system. Therefore the stability augmentation system (SAS) emerged to increase the stability of aircraft based on aircraft motion information feedback. Although the SAS increases the stability of the aircraft, it reduces the maneuverability of the aircraft to some extent. To solve this issue, the pilot’s command signal is introduced and subtracted from the aircraft motion information feedback signal, so that the SAS will not affect the pilot’s control. When there is no pilot command, the feedback control can effectively suppress the aircraft’s disturbed motion, so that the stability and maneuverability of the aircraft can be both taken into account, that is, the so-called control augmentation system (CAS). At the end of the 1960s, with the emergence of high maneuverability requirements and advanced aerodynamic layout, computer technology and control theory developed rapidly and the aircraft control system experienced the second major technological change—the emergence of fly-bywire control system. With the development of the fly-by-wire control system, the artificial sensor device returns to the fixed rod force-rod displacement index and other indexes (such as rod force-overload, rod displacement-angular speed) were realized by the control computer. Fly-by-wire control is different from the conventional machinery control system. In terms of form, it is a control system adopting electrical signal transmission, and regarding content it is a full-authority feedback control system with the aircraft (motion parameter) as the control object. Fly-by-wire control can be defined as a flight control system adopting redundant electrical signal transmission with aircraft motion parameters as control parameters. The important difference between it and the machinery control system is not only that it replaces the mechanical chain with cables, but more importantly, the pilot no longer directly controls the deflection of the aircraft control plane but controls the motion of the aircraft. Fly-by-wire control facilitates the synthesis, conversion, transmission, and control of various signals and boosts the upgrading and optimization of the system. The core of the design and verification of the fly-by-wire control system is the safety of the system and the applicability of the control law. The flight control system control law and flying quality evaluation test discussed in this chapter is the control law design and verification for the flying quality of the engineering simulator and flight control system “iron bird” integration test bed.

4.1 Overview Despite the development of the aircraft control system, its fundamental purpose is to enable pilots to safely and comfortably complete various established flight missions as easily as possible. Flying quality is a comprehensive ability to measure the maneuverability and stability of an aircraft and also an assessment index of the burden made by pilots to

Flight control system control law and the flying quality evaluation test 241 complete flight missions. The flying quality specification, which describes the quantitative design criteria for the quality of the aircraft control system, improved gradually through many scientific experiments and production practices. The earliest flying quality specification can be traced back to 1908. Now, the commonly used flying quality specifications include US military standards MIL-F-8785 C and MIL-STD—1797. The corresponding flying quality specifications in China include Flying Qualities of Piloted Airplanes (Fixed Wing) (GJB185—86) and Flying Qualities Standard for Airplane with Fly-by-Wire Control System (GJB2874—97). Based on descriptions of flying qualities and handling qualities, the basic requirements for the flying qualities of an aircraft are as follows: 1. Under the condition of man machine closed loop, the aircraft shall be continuously controllable. 2. The aircraft control response shall be stable, fast, and accurate enough. 3. Workload of pilots shall be minimized. Generally, the control law can be understood as a kind of digital (or analog) special control system. For a modern fly-by-wire flight control system, the control law is a series of control plane command algorithms under existing technical conditions (control device, actuation system, and sensor system) to make the aircraft meet the requirements of flying quality specifications within the range of a given flight envelope. From the perspective of flying quality, modern aircraft designers are not only designing an aircraft, but more importantly, designing a man aircraft system. An aircraft with a flight control system must be suitable for users and cater to their needs. In this sense, the starting point and final verification of flying quality design should focus on pilots. The modern aircraft flight control system has a high degree of complexity. The control law structure and algorithm need to make the aircraft meet the flying quality requirements under various configurations, modes, and missions. The all-digital simulation, engineering simulator test, “iron bird” integration test, “iron bird” human machine combined test, onboard ground test, and flight test are all related to the control law and flying quality evaluation work. 1. All-digital simulation is mainly used for nonphysical objective evaluation of the whole flight envelope and all modes. It highlights the control law structure and parameter validation under each design state and the main users are designers. 2. The engineering simulator test comprehensively verifies the control law, flying quality and flight performance, highlights the subjective feeling and evaluation of pilots and is used to improve and optimize the design of control law. 3. The “iron bird” integration test is a combined test to verify the flight control system and its cross-linked system. It highlights the system interface, logic, functional, and

242 Chapter 4 dynamic (static) performance of system hardware equipment in the loop and the comprehensive verification of fault modes, including the comprehensive control law and flying quality verification, such as state switch, polarity transmission ratio, and stable reserve. 4. The “iron bird” human machine combined test is mainly conducted to verify the flying quality when the software and hardware are in a realistic environment. It highlights the human machine combined characteristics analysis and verification when the pilot is in a physical simulation environment. 5. The onboard ground test is mainly used to further assess the characteristics of the flight control law in the real environment of aircraft, such as the interface with other relevant airborne systems and the stability after closure through the aircraft structural mode, so as to further determine the filter parameters of the feedback signal and highlight the relevance between the system and the structure. 6. The flight test is the final method to evaluate the flight control law, including comprehensive inspection of control law and flying quality. It highlights performance evaluation and comprehensive verification in real environment.

4.1.1 Design requirements for flying quality of large transport aircraft There are mainly two flying quality requirements for large transport aircraft: military standard and airworthiness standard. The two standards are closely related in terms of performance and safety and also have many differences. Military standard mainly applies to military aircraft and airworthiness standard is mainly for civil aircraft. As safety is the largest concern for civil aircraft, the airworthiness standard is selected to highlight safety for civil aircraft. As performance is most important for military aircraft, the military standard is selected to highlight the quality rating for military aircraft. Regarding its development, large transport aircraft are widely used in peacetime, and it is inevitable that their design requirements follow the airworthiness standards for civil aircraft. Therefore the selection of flying quality requirements for large transport aircraft shall be subject to military standards and also introduce relevant airworthiness safety provisions. Military standards related to the flying quality requirements of large transport aircraft mainly include Flying Qualities of Piloted Airplanes (Fixed Wing) (GJB185—1986) and Flying Qualities Standard for Airplane with Fly-by-Wire Control System (GJB2874—1997). GJB185—1986 was prepared according to the lessons learned from China’s experience in aircraft design and flight practices regarding flying quality (US military standard MIL-F-8785C was also referenced at the same time). This standard is applicable to aircraft with a machinery control system. After introducing the concept of the “equivalent system,” it can also be used for aircraft with a SAS, aircraft with a CAS, and aircraft with a fly-by-wire control system. It mainly covers the scope of application, aircraft classifications, flying quality standard

Flight control system control law and the flying quality evaluation test 243 regulations, definition, terminology and symbols, longitudinal flying quality, lateral (directional) flying quality, stall and screw, other flight qualities, primary flight control system characteristics, secondary flight control system characteristics, atmospheric disturbance, supplementary requirements for the flying quality of seaplanes, and quality assurance. GJB2874—1997 was prepared based on the research and analysis of US military standard Recommended Military Standard and Manual—Flying Quality of Atmospheric Vehicle (AFWAL_TR-82 3081) and China’s aircraft development and flying quality standard preparation experience. It mainly stipulates the requirements in the air and on the ground for military aircraft with a fly-by-wire control system and is applicable to various aircraft with a fly-by-wire control system. It mainly details the scope, reference documents, definitions, general requirements, detailed requirements, items needing explanation, and appendixes. The differences between GJB2874—1997 and GJB185—1986 are similar to those between US military standards MIL-F-8785C and MIL-HDBK—1797. Compared with GJB185—1986, GJB2874—1997 highlights the customization based on the customers’ needs, adds many new provisions to the design and validation of the aircraft with fly-by-wire control system, and makes major changes to the standard structure.

4.1.2 Design requirements for flight control system control law of large transport aircraft The design requirements for the control law of the flight control system of large transport aircraft mainly cover two aspects: flying quality requirements and functional requirements. The flying quality requirements have been described in section 4.4.1. As large transport aircraft are generally class III aircraft, specifications and requirements for class III aircraft are generally adopted. During aircraft development, some content of GJB2874—1997 can be deleted according to the actual situation of the aircraft, relevant new requirements can be supplemented, and design experience of domestic and foreign advanced military and civil aircraft can be introduced as part of the flying quality design requirements (such as FAR-25). As flying quality design requirements are wide-ranging, corresponding evaluation criteria are usually adopted according to the development needs of the flight control law in different stages. The functional requirements for the control law of the flight control system mainly relate to precision and correlation and are from airworthiness standards and national military standards General Specification for Flight Control Systems of Piloted Aircraft (GJB2191— 1994), General Specification for Automatic Flight Control Systems and SAS and CAS of Piloted Aircraft (GJB3819—1999), and General Specification for Piloted Aircraft Autopilot (GJB1690—1993). GJB2191—1994 is a relatively comprehensive specification of the flight control system, which is not only applicable to the conventional machinery control system, but also applicable to the fly-by-wire control system. As a guiding document for the design,

244 Chapter 4 installation, and test of the flight control system, it has regulated key and major events in the development of the flight control system. 1. GJB2191—1994 is prepared based on the US military standard General Specification for the Design, Installation and Test of Flight Control Systems of Piloted Aircraft (MIL-F-9490D) and follows the principle of equivalence. The US military standard MIL-F-9490D is not only a standard commonly used for flight control system design in the United States, but also around the world. It is not only applicable to military transport aircraft and civil aircraft, but also to bomber aircraft and fighter aircraft. 2. GJB3819—1999 mainly specifies the stability and control augmentation of the flight control system and the automatic flight control system. As a guideline for the stability augmentation, control augmentation, and the design of automatic flight control system, it is applicable to the design of the automatic flight control system and manual flight control system and is a refinement of GJB2191—1994. 3. GJB1690—1993 specifies the autopilot of the flight control system and the automatic flight control system in line with that of GJB3819—1999 and GJB2191—1994. In actual aircraft development, the functions of the control law of the flight control system are generally determined according to the mission requirements of the aircraft. For a modern large transport aircraft, its manual control system generally has functions including three-axis control augmentation, stall protection, overspeed protection, overload protection, and automatic trim and its automatic flight control system generally has functions including altitude holding/selection, course holding/selection, speed holding/selection, horizontal navigation, vertical navigation, glide slope, and localizer. The design requirements for these functions are generally made based on national military standards as well as the characteristics of aircraft by modifying and adjusting the scope of indicators and provisions appropriately. With regard to the verification of these functions, they will be evaluated in an engineering simulator test according to design requirements generally and then the pilot will give suggestions for functional and logical improvement. If there are failures in the flight control system, sensor signal, control surface, engine, and energy, the flight control law will be affected in performance and function. Therefore the design requirements for the control law of the flight control system shall include the requirements for flying quality and function under various failures.

4.2 Stage division and objectives of the evaluation test According to the change of the design input state of the flight control system in each development stage, the control law and flying quality evaluation test can be divided into four stages.

Flight control system control law and the flying quality evaluation test 245 1. Stage 1: Structural demonstration of control law This stage starts from aircraft preresearch or aircraft project approval with the absence of a large-scale wind tunnel test. The aircraft data are mainly from CFD calculation and theoretical estimation and the engineering simulator is also in the initial construction stage (or borrowing an engineering simulator of other models). The main design work in this stage includes the preliminary selection of the parameter range of force and displacement of the control device, theoretical analysis of the main aerodynamic characteristics of aircraft, preliminary determination of stability augmentation control law structure and key relevant parameters, preliminary design of the parameter adjusting law of the control law, and the determination of the perturbation range of aerodynamic and control law parameters through theoretical analysis. After the completion of the theoretical design work, the pilot will evaluate the flying quality on an engineering simulator and select the best control law structure and control device parameters. 2. Stage 2: Detailed design of control law In this stage, major wind tunnel tests and more detailed CFD calculations are carried out, and power systems, sensors, and actuators obtain partial test data. With regard to the design of the control law, the control augmentation, boundary protection, high lift control, and detailed design of the automatic flight control system are mainly conducted. The engineering simulator test is mainly conducted to evaluate the flying quality by the pilot after all the work above is completed, so that the control law parameters can be determined. After control law synthesis, an overall flying quality evaluation will be conducted again. 3. Stage 3: Overall verification before maiden flight Before the maiden flight of the aircraft, the wind tunnel test, power system test, sensor test, and actuator test have been completed and the control law itself has gone through several rounds of engineering simulator testing. After confirming the final design input data and the maiden flight status, a new round of control law design checks and engineering simulator test verification will be conducted. Besides, after the completion of the ground integration test (“iron bird” integration test) of the flight control system, the “iron bird” semiphysical environment can be used to conduct the pilot-in-loop test to comprehensively evaluate the control law and flying quality. In this stage, the engineering simulator and the “iron bird” integration test environment can be used for the training of the pilot and the pilot also becomes familiar with the maiden flight to ensure the success of the flight. 4. Stage 4: Test flight and verification after maiden flight In this stage, the flight performance data packet is optimized according to the test flight data to provide a more realistic flight simulation environment for the test flight, so as to create conditions for the subsequent training on the units. The engineering simulator and “iron bird” flight packages and control law model can also be used for the research and development of the training simulator. Generally, the engineering

246 Chapter 4 simulator can be used for ground emergency response training before the test flight of risk subjects. In the process of the test flight, in case of a fault in the air, ground troubleshooting with the engineering simulator in the “iron bird” integration test environment is the best approach. The engineering simulator test mainly aims to: a. verify the robustness of the control law of flight control system. b. verify the parameter validity of the control law of flight control system. c. verify the acceptability of mode conversion to transient state of the control law of flight control system. d. verify whether the rod force/rod displacement of the control system meets the fullenvelope flight control requirements. e. verify the matching degree of aircraft maneuverability and stability. f. verify if the energy and actuator power requirements are met. The human machine combined test in the “iron bird” integration test environment mainly aims to: 1. further verify the internal interface (mechanical, electrical, bus), function and performance of each subsystem and confirm whether the internal interface relationship of each subsystem is correct and whether the function and performance (dynamic and static) meet the design requirements of the subsystem. 2. verify the interface, signal transmission and logic relationship between each subsystem in the flight control system and confirm that the interface (mechanical, electrical, and bus) relationship in the flight control system is correct. 3. verify the interface (mechanical, electrical, bus) between the flight control system and other aircraft systems, including the hydraulic system, power supply system, avionics system, take-off and landing control system, and aircraft actuation surface, and confirm the interface relationship is correct. 4. verify the control channels (longitudinal, directional, and transverse) of the aircraft and the function, performance (time domain and frequency domain), and control logic between channels during manual control and automatic control, and confirm whether they meet the requirements of flight control system design specification. 5. have a comparative analysis of the results of the human machine combined test in the “iron bird” integration test environment and theoretical calculation results and quality simulator test results before the maiden flight, and comprehensively verify the consistency between the flight control system and the design state. 6. establish the confidence of the test pilot through simulated test running, fault simulation, and maiden flight profile drill. 7. conduct simulated flight for subjects with high risk in the test flight stage and improve the handling plan. 8. reproduce and remove the faults during the test flight.

Flight control system control law and the flying quality evaluation test 247

4.3 Design requirements for the engineering simulator 4.3.1 Composition Engineering simulator (sometimes called quality simulator) is an important device for designing flight control law and verifying the flying quality of aircraft, an essential device for the design of a flight control system, especially the design of a fly-by-wire flight control system, as well as a main environment and platform for pilots to intervene in aircraft design during the whole life cycle. The engineering simulator and flight training simulator are basically the same in composition and they highlight the analog simulation of flight performance-related content, such as package, manual and automatic flight control system and control law, engine, landing gear, as well as flight management. The difference is that the engineering simulator is widely in use. Apart from the equipment layout, the function and performance are closely related to the quality evaluation that needs to be consistent with that of real aircraft; the cockpit instrument and others can adopt the general system. The engineering simulator of a large transport aircraft is mainly composed of the aircraft performance simulation system, cockpit simulation system, integrated control management system, avionics simulation system, dynamic simulation system, vision system, sound simulation system, computer interface and network system, and the engineer analysis and appraisal system. The basic composition is shown in Fig. 4.1 and the network topology structure is shown in Fig. 4.2. As shown in Fig. 4.2, network data transmission is divided into two levels of management. The data communication between the aircraft performance simulation system, avionics simulation system, integrated control and management system, control load system, dynamic simulation system, vision system, and computer interface system has strict reliability and real-time requirements. It is realized through a reflective memory real-time network, which is managed as a primary network. The internal communication of the aircraft performance simulation system and avionics simulation system is realized through the Internet, which is managed as a secondary network.

4.3.2 Main functions The engineering simulator of a large transport aircraft shall have following main functions: 1. Design and evaluation of control law of manual flight control system. 2. Design and evaluation of control law of automatic flight control system (including autothrottle). 3. Design and evaluation of flight control systems such as automatic navigation system. 4. Evaluation of flying qualities such as aircraft maneuverability and stability.

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Figure 4.1 Basic composition of engineering simulator.

5. 6. 7. 8. 9.

Evaluation of aircraft flight performance. Evaluation of interconnected system or human machine work efficiency. Fault recurrence, troubleshooting analysis, and study of ground test and flight test. Support for preliminary training for maiden flight. Evaluation of test flight methods.

4.3.3 Design requirements The requirements for parts of the engineering simulator related to flight performance are even higher than those for the flight training simulator, but the rest can be lower than those of the flight training simulator, such as vision system, instrument simulation system, and sound simulation system. The main performance indicators are as follows:

Flight control system control law and the flying quality evaluation test 249

Figure 4.2 Network topology block diagram of engineering simulator.

1. Simulation frame period: 16.7 ms. 2. Transmission delay time: ≯120 ms. 3. The aircraft flight performance, power device performance, aerodynamic effect, control characteristics, and flying quality shall be simulated and the effects of various climate conditions, runway conditions, and icing on aircraft performance shall be simulated. 4. The dynamic force to the pilot in the takeoff, landing, and maneuvering flight process of the aircraft shall be simulated. The aircraft and aircraft system simulation, engineer analysis, and appraisal system are the key systems of the engineering simulator, which will be described in detail below. Other systems are relatively generic and are not described here. 4.3.3.1 Simulation requirements for aircraft and aircraft system The fidelity of the aircraft and aircraft system simulation models in the engineering simulator directly affects the evaluation results of pilots. According to the Flight Simulation Device Certification and Use Rules (CCAR60) and Flight Simulation Training Device Design & Performance Data Requirements issued by the Civil Aviation Administration of China, as well as the design requirements and test results of aircraft, the main performance requirements are as follows:

250 Chapter 4 1. Taxiing performance The ground taxiing performance of aircraft shall be simulated and the landing gear support compression change, tire friction, lateral force and other ground reaction effects shall be considered in the modeling process. 2. Takeoff performance The ground takeoff performance of aircraft shall be simulated and the landing gear support compression change, tire friction, lateral force, and other ground reaction effects shall be considered in the modeling process. 3. Climb performance The climb performance of aircraft shall be simulated and the effects of environment and ground effect to flight and control shall be considered in the modeling process. 4. Cruise performance The cruise performance of aircraft shall be simulated and the effects of highaltitude Mach number effect to flight and control shall be considered in the modeling process. 5. Ground deceleration performance The ground deceleration performance of aircraft shall be simulated and the braking efficiency reduction effect caused by the large humidity of the brake disc and too high braking temperature shall be considered in the modeling for aircraft ground taxiing. 6. Engine acceleration and deceleration performance The engine acceleration and deceleration performance of aircraft shall be simulated. 7. Static handling quality High-fidelity simulation shall be conducted for the static characteristics of control devices such as steering wheel, rudder pedal, front wheel steering control device, pitch trim device, throttle lever, and brake pedal. 8. Dynamic control response High-fidelity simulation shall be conducted for the dynamic characteristics of aircraft under pitching, rolling, and yaw control. 9. Longitudinal dynamic characteristic High-fidelity simulation shall be conducted for the engine power and configuration changes (including flaps/slats change, spoiler or air brake change, and landing gear change), the longitudinal trim performance of aircraft, longitudinal maneuvering stability and static stability of aircraft, response of the aircraft to stall state, and the dynamic response characteristics of the aircraft under longitudinal control. 10. Directional dynamic characteristics High-fidelity simulation shall be conducted for the directional dynamic characteristics of the aircraft, including the minimum control speed in air, the rolling response, spiral stability, engine failure trim, rudder response, stable sideslip, and Dutch roll.

Flight control system control law and the flying quality evaluation test 251 11. Dynamic characteristics in landing process High-fidelity simulation shall be conducted for the dynamic characteristics of the aircraft in the normal landing process, landing with smallest flaps or no flaps, crosswind landing, landing with one faulted engine, and missed approach. 12. Flight envelope protection simulation The flight envelope protection function and performance of aircraft shall be simulated. 4.3.3.2 Requirements for engineer analysis and appraisal system As a main platform based on the analysis of the control law and flying quality evaluation test results of an engineering simulator, the engineer analysis and appraisal system shall be able to realize data display, data recording, data analysis, off-line flight simulation, model identification, and evaluation functions. 1. Data display The data display function of the engineer analysis and appraisal system is mainly exerted to monitor the process of the control law and flying quality evaluation test. It shall support the display of two types of test data: data in dynamic text form and data in dynamic curve form. The data in dynamic text form includes the landing gear control status signal, aircraft wheel load signal, autopilot on/off signal, and flight control system work mode signal. The data in dynamic curve form includes the aircraft motion status signal, pilot control signal, and aircraft attitude information. By combining the numerical form and the dynamic curve form, the flight condition of the aircraft during the test is comprehensively displayed. Besides, the data display interface should be able to be adjustable, such as the range of interface curve curvilinear coordinates, selection of curve signal, selection of text signal, number of curves, and texts. 2. Data recording The data recording function of the engineer analysis and appraisal system is mainly exerted to prepare data for the appraisal and problem analysis of the control law and flying quality evaluation test. It shall have the function of selecting to record data or not and the function of recording the data. There are many situations in the test process. For example, there may be a situation that does not need recording data or a situation that requires stopping data recording, such as during flight preparation, state adjustment, and test subject transition. Under this situation; redundant data can be reduced and data space can be saved. Different test subjects need to record different data. For example, manual control characteristic inspection does not need to record data of automatic flight. The stall protection function inspection not only needs to record the control information and aircraft motion state, but also needs to record the intermediate quantity alarm angle of attack and the maximum angle of attack calculated with the control law. During off-line flight simulation, all input commands and aircraft motion states shall be recorded. Therefore the engineer analysis and appraisal system should be able to set up several files and record data at the same time, and each file can choose different data.

252 Chapter 4 3. Data analysis The most critical function of the engineer analysis and appraisal system is data analysis, which is exerted for the appraisal of the control law and flying quality evaluation test. It shall be able to read, display, and analyze the recorded data. Due to the different time and quantity of data recording and different data file size and data type, data loading shall support various data types and also support the reading of big data (above 1 G). Different test subjects have different requirements for data analysis. For example, the result analysis of the control stability test generally requires a curve display and characteristic calculation function; the forward and reverse rolling test requires drawing of the response curve of the transverse control command, roll angle, and roll angle rate and to calculate the time required for rolling change to evaluate the rolling characteristics; the result analysis of the functional inspection test generally requires drawing of a correlation curve of functional parameters, functional targets, and limit values to accurately analyze the completion of functions; the stall protection function inspection requires drawing of the correlation curve of aircraft angle of attack, alarm angle of attack, and maximum angle of attack; and the result analysis of the human machine combined test generally requires drawing a curve family according to different types. For normal items, the changes of three-axis control quantity, three-axis control plane deflection, three-axis angular rate, and altitude, as well as the trim features of the aircraft should be analyzed. At the same time, the contrast analysis of different test results of the same subject shall also be considered. To better evaluate test results, the pilot needs to repeat the test several times and comprehensive analysis of several test results is required. For example, for the push-and-pull attitude angle, the control quantity of several times and corresponding altitude responses shall be drawn together to have comprehensive analysis of the intercepted attitude angle predictive ability. 4. Off-line flight simulation The off-line flight simulation function of the engineer analysis and appraisal system is mainly exerted for control playback, problem analysis, and problem positioning of the control law and flying quality evaluation test. The data recorded is always limited and cannot cover all flight-related quantities of the aircraft, especially the intermediate variables of the solution of the control law and the aircraft model. As the purpose of the test is to develop and evaluate the control law, some problems in the test process sometimes cannot be analyzed thoroughly with the recorded data and off-line simulation and test process recurrence are required to view any variables concerned. Based on these considerations, the off-line flight simulation shall be able to read the data files required for the model operation and support the comparison between the offline simulation results and data records to ensure the effectiveness of the off-line flight simulation. 5. Model identification and evaluation The model identification and evaluation function of the engineer analysis and appraisal system is mainly exerted to provide technical support for the prediction of pilot-induced

Flight control system control law and the flying quality evaluation test 253 oscillations (PIO) in control law and the flying quality appraisal test. With the development of the PIO prediction technology, PIO prediction is no longer limited to the subjective evaluation of the physical person (pilot) in the loop, but develops by the objective evaluation of the physical person (pilot) in the loop. It not only avoids the inaccurate driver mode in traditional PIO theoretical analysis, but also addresses the issue of large randomness in human machine combined testing and effectively improves the rationality of the PIO prediction. The key to this method lies in identifying the physical person (pilot) and the system model, including the nonlinear link. The engineer analysis and appraisal system shall have the model identification function based on data, so as to draw the PIO prediction evaluation chart based on the PIO prediction criteria. The main performance indicators and requirements of the engineer analysis and appraisal system are as follows: a. Interface refresh rate: not less than 80 Hz. b. Network communication rate: not less than 10 Mb/s. c. Data recording performance: not less than 2 h record, record packet loss probability less than 1023. d. Degraded for use, CPU availability factor and memory usage are both not less than 30%.

4.4 Test items and methods Whether it is the flight control system control law and flying quality evaluation test in the engineering simulator environment or the human machine combined test in the “iron bird” integration test environment, they are both simulation tests for pilots in the loop. They are conducted to constantly optimize and improve the design by exposing the defects of control law and flying quality design as early as possible and familiarize the pilot with the basic control modes of the aircraft and aircraft control stability characteristics, in order to build confidence for the maiden test flight. The procedures and methods of the evaluation tests conducted under the two test environments participated in by the pilot are basically the same. Firstly, the control law engineer or system engineer forms the test task description according to the task planning and development progress, then the test flight engineer prepares the test outline and test task list, and then the evaluation test is carried out according to the task list. Before the test, the test flight engineer shall explain in details the content and requirements of the task list and the key points of the control action to the pilot and all participants. After the test, the pilot shall comment on the test and discuss with the system engineer and the control law engineer. If necessary, the pilot shall review the subject execution parameters and video and finally give the evaluation or suggestions for improvements.

254 Chapter 4 After completing a stage or a series of correlated test tasks, the control law engineer will comprehensively analyze the correlation of a certain parameter at different states and compare it with theoretical calculation results to obtain the final modification opinions, and then give feedback to the pilot. After the flight control system or control law design is improved, the pilot will reevaluate until a satisfactory result is reached.

4.4.1 Planning for test tasks The control law and flying quality evaluation test is mainly divided into four stages and the test task planning follows the principle of “from top level to details, from architecture to parameters, and from normal to failure.” Pilots with experience flying aircraft of a similar background are invited for the evaluation. With the progress of the test task, the number of pilots invited and the units they work for must follow the principle of “from divergence to concentration.” The control law and flying quality evaluation test mainly includes: 1. 2. 3. 4. 5. 6. 7. 8. 9.

Evaluation of rod force-rod displacement, damping and rocker arm characteristics. Evaluation of mode and logical functions. Evaluation of control augmentation control law and quality. Evaluation of high lift system control law function and performance. Evaluation of automatic flight function and performance. Evaluation of boundary protection control law. Evaluation of failure mode function and performance. Evaluation of maiden flight status. Evaluation of risky subject performance.

In different stages of the flight control system development, relevant functional and performance evaluation is carried out according to the design and development of the flight control system function, control law, and artificial feel system. Generally speaking, the structure of the control law of the artificial feel system and CAS will be evaluated first, then comes the normal modal control function, the control law of the high lift control system and the manual flight control system, and the failure mode and the automatic flight control system, before evaluating finally the maiden flight state and risky subjects. If there are problems in the state evaluation of a certain round, a supplementary evaluation after the improvement will be conducted in subsequent evaluation work.

4.4.2 Selection of test state points In the control law design process of modern aircraft flight control systems, to ensure design quality (covering various envelopes) and to improve the design efficiency (merging states with similar performance), it is relatively reasonable to select 1000 2000 state points for design, about 50 200 state points for engineering simulator tests, and about 10 20 state

Flight control system control law and the flying quality evaluation test 255 points for the human machine combined test in the “iron bird” integration test environment. The elements that constitute the state point generally include flight stage, aircraft weight, aircraft center of gravity, flaps and slats configuration, flight altitude, flight speed, landing gear state, and power system state. In the absence of any constraints, the state points can be selected with an almost exhaustive method based on the design experience of the control law engineers. The development of computer technology makes it possible to select a large number of design state points, but it is still unrealistic and unnecessary for tests. A multiconstraint test state point selection principle and method is proposed below: 1. The selection of test state points shall show obvious differences, including both the state points with good quality and state points with poor quality deemed by designers. 2. The selection of test state points shall be wide-ranging. In other words, the selection shall cover the boundary points of various envelopes. 3. The selection of test state points shall take into account the aircraft service conditions and pilot control habits. 4. According to a certain design requirement, the natural aircraft characteristics at the design state point shall be calculated, and all the state points are described in the image with a specific index quantity and the dense point in the image shall be taken as a test state point. 5. The aerodynamic characteristics data of aircraft shall be analyzed and they shall be taken as test state point candidates according to the use limit conditions and control efficiency limit conditions of the aerodynamic control plane and the flight state conditions. 6. The design results of the control law parameters shall be analyzed to select the midpoint of the segment on the curve of parameter adjusting law. Then, the parameter in the typical state shall be calculated and taken as the candidate for the test state point. 7. The test state points necessary for logical selection of the flight control system shall be combined, such as the space conversion condition. 8. The state points for the human machine combined test in the “iron bird” integration test environment can be screened according to the concerns of the pilot in several rounds of quality simulator testing.

4.4.3 Test items The flight control system control law and flying quality evaluation test includes the test in the engineering simulator environment, the human machine combined test in the “iron bird” integration test environment and the onboard ground test. Major test items are shown below:

256 Chapter 4 1. Control law design and evaluation of manual flight control system. It includes the type selection and evaluation of the pilot control mechanism, configuration evaluation and parameter optimization of manual flight control law, boundary protection logic evaluation and boundary validation, evaluation of high lift control function and modal conversion to transient state, as well as the evaluation of flight control system degradation and fault protection logic and performance. 2. Control law design and evaluation of automatic flight control system. It includes the evaluation of functions of the automatic flight control system, such as automatic pilot, automatic navigation, automatic approach, automatic throttle, and flight direction, and the evaluation of the control law configuration of various modes, such as holding of angle of pitch, vertical speed, flight path angle, altitude holding, gradient holding, course selection, horizontal straight navigation line, horizontal arc navigation line, level change, vertical navigation trail control, autothrottle thrust control, autothrottle speed control, as well as parameter optimization. 3. Design and evaluation of flight control systems such as automatic navigation system. Taking the flight control system, automatic flight control system, and manual flight control system as a three-layer control loop, it includes the evaluation of control law configuration under various modes such as automatic navigation and the automatic approach, as well as parameter optimization. 4. Flying quality evaluation of aircraft. According to the design requirements for aircraft flying quality, it makes a comprehensive analysis and verification of the flying quality of aircraft under different aircraft configurations and different flight states. 5. Flight performance evaluation of aircraft. It makes comprehensive evaluation of the basic flight, endurance, maneuverability, takeoff/landing, aerodynamic effect, etc. of aircraft. 6. Evaluation of interconnected system and human machine work efficiency. It makes a comprehensive evaluation of other aircraft systems associated with aircraft performance and flight control systems, such as the logic and power of the power supply system, the wheel load (wheel speed) signal of the landing gear system, the flow and power of the hydraulic system, the layout of control and display devices, and the man machine work efficiency. 7. Fault recurrence and analysis. It solves different problems exposed in the ground test and flight test, checks failure modes, finds out the causes of failures, and proposes solutions, so as to improve the design of the system and control law. 8. Preliminary training for maiden flight screw. Assists the chief test pilot to familiarize themselves with aircraft failures and special handling measures, understand the comprehensive performance of the aircraft, and support the maiden flight to the maximum.

Flight control system control law and the flying quality evaluation test 257 9. Evaluation of test flight methods. Conduct simulation training on key and risky subjects, such as the handling measures for the failure caused because the flaps are not put down according to requirements, and develop safer and more reasonable test flight plans.

4.4.4 Test control action As the flight control system control law and flying quality evaluation test is completed through the pilot’s control action, the control action shall be designed in a scientific way according to the test items. In fact, after long-term practice of the flight test, many standard actions have been formed for selection. The test control actions used for the engineering simulator test and the human machine combined test conducted in the “iron bird” integration test environment and the test items are shown below. 1. Single-pulse control: it can be divided into longitudinal, transverse, and directional single-pulse control. The control amplitude is 1/4, 1/2, and 1 times that of the full control stroke. During the test, the stroke is gradually increased and the free response is 10 s after the control action is completed. This control action is suitable for checking the aircraft’s initial response and equivalent time delay, checking the aircraft’s short-period modal frequency and damping, as well as for parameter identification. 2. Time-pulse control: it can be divided into longitudinal, transverse, and directional time-pulse control. The control requirements and application range of this control action are the same as that of single-pulse control and the only difference is that the one-way control is changed to two-way continuous control, as shown in Fig. 4.3. 3. Step control: it can be divided into longitudinal, transverse, and directional step control. The control is increased to the specified amplitude and maintained for 3 7 s. The amplitude is 1/4, 1/2, and 1 times that of the full control stroke respectively. To ensure that the flight state does not change too much, the longitudinal control can perform an appropriate reverse control first, and then perform the reverse step control. This control action is suitable for checking the initial response and equivalent time delay of the aircraft, checking the modal characteristics of the aircraft, identifying key parameters, and confirming the control gradient.

Figure 4.3 Time-pulse control.

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Figure 4.4 “3211” Control.

4. “3211” control: this control action is only applicable to longitudinal control. Conduct a set of step control of longitudinal time series, pull rod for 3 s, push rod for 2 s, pull rod for 1 s, push rod for 1 s, and finally return to center. The amplitude of pulling and pushing the rod shall be the same so as to minimize the longitudinal response generated by the aircraft. The amplitude shall be determined by the pilot and it should be kept for 10 s after the completion of the control. This control action is suitable for parameter identification and can be used to check the initial response, following the quality and stability of the aircraft, as shown in Fig. 4.4. 5. Sinusoidal frequency sweep: equal-frequency or variable-frequency control is conducted for the control column. The frequency range is 0.2 5 Hz. This sweep control can be manually completed by the pilot or generated by a specific signal generator. This control action is suitable for measuring the dynamic quality of the aircraft system, including the stability margin of the flight control system, and for parameter identification, especially the parameter identification in the frequency domain. 6. Wind-up turn: this control action is only applicable to transverse control. Keep the speed constant, gradually press the rod to increase the roll angle and fully pull the rod to gradually increase the normal overload until it reaches the maximum overload and angle of attack. This control action is used to verify maximum overload and angle of attack, check the gradient of rod force, and check control harmony. 7. Decelerated turn: keep the altitude constant, turn under different engine status and at different deceleration level, gradually increase the roll angle, pull rod to the maximum stroke. This control action is to check the performance of the angle of attack and overload limiter and check the control force varying with the angle of attack and overload. 8. Quick turn: at a given altitude, keep the speed constant through the throttle and quickly turn and pull the rod fully to reach the maximum angle of attack and normal overload. This control action is used to check horizontal maneuverability and control the coordination and control force. 9. Maximum roll and pitch combined control: rapidly pull and press the rod and gradually increase the longitudinal and transverse control amplitude until the rod is pulled and pressed fully. This control action is suitable for checking the performance

Flight control system control law and the flying quality evaluation test 259

10.

11.

12.

13.

14.

of the angle of attack limiter under coupled longitudinal and transverse actions, the longitudinal and transverse control coordination, the maximum control force and maximum speed, as well as for checking the structural strength and load. Accelerated and decelerated level flight: Keep the 1 g flight at a constant altitude. Accelerate and decelerate the flight through throttle and measure the changes of rod force and rod displacement. The purpose of this test is to check the speed stability of the aircraft. BTB (back-to-back): keep the altitude and speed constant, press rod and roll to an angle, quickly press the rod in opposite direction after it stabilizes to make the aircraft roll in the opposite direction to the same angle, and then restore the horizontal flight after stabilization. Keep the pitch angle constant during the roll process. This control action is suitable for checking aircraft coordination, lateral control response, and integrated attitude control capabilities. Coordinated sideslip: keep the flight altitude and speed constant, press the rod and pedal the rudder, keep the flight course constant. This control action is used to check the longitudinal, transverse, and directional control coordination of the aircraft and check the steady-state sideslip quality and crosswind landing correction capability. Rod push and pull to the given pitch angle: quickly pull the rod to the given pitch angle and keep for 3 5 s. This control action is suitable for checking the predictability, stability, and accuracy of pitch control. Precise approach/deviation-correction landing: keep on the glide path accurately according to the guided approach route first and then land according to the landing mark made on the runway. Sometimes, it is set deviated from the center line of the runway before landing. Thus it shall be quickly corrected to the center line of the runway or designated marker point for landing. This control action is used to check the man machine closed-loop flying quality of approach and landing.

These control actions can be roughly divided into three categories: open-loop control, openloop and closed-loop combined control, and closed-loop control. Open-loop control is dominated by quantitative evaluation. Open-loop and closed-loop combined control is evaluated quantitatively and qualitatively. Closed-loop control is dominated by the subjective evaluation of test pilots.

4.4.5 Test task list As mentioned above, the flight control system control law and flying quality evaluation test involves many states that cannot be fully reflected in one report and the descriptions made by the pilot in the test process shall be recorded in time; it is suggested to further specify the test tasks in the form of a task list, clarifying the test process and standardizing the test operation, which is conducive to the test organization and management, test result recording, test schedule speeding up, and test report preparation.

260 Chapter 4 For the flight control system of large transport aircraft, the test task list shall determine the test sequence according to the combination of the aircraft weight, center-of-gravity position, and the mode of the flight control system and a task list under a specific combination shall determine the action based on the aircraft flight altitude and speed. In this way, the setting of each state point can complete all the test actions under this state point and effectively improve the test schedule. Table 4.1 shows the test task list in a certain state. For other test items, the state points and standard actions of the aircraft should be replaced with corresponding content. Table 4.1: Example of test task list. Task listno: 001

Weight: 90,000kg

Flight state

Speed (km/h, Ma)

Center of gravity: intermediate of aircraft Operation mode: normal Standard action

Evaluation

“3211” control for longitudinal, transverse, and directional control 381

Push/Pull the stick to make the pitch angle to the target value Turn the wheel to make the roll angle to the target value

In according with “Cooper_Harper” rating

Coordinated sideslip

Cruise configuration

“3211” control for longitudinal, transverse, and directional control 500

Push/Pull the stick to make the pitch angle to the target value Turn the wheel to make the roll angle to the target value

In according with “Cooper_Harper” rating

Coordinated sideslip “3211” control for longitudinal, transverse, and directional control 179

Push/Pull the stick to make the pitch angle to the target value Turn the wheel to make the roll angle to the target value

In according with “Cooper_Harper” rating

Coordinated sideslip

Take off configuration

“3211” control for longitudinal, transverse, and directional control 244

Push/Pull the stick to make the pitch angle to the target value Turn the wheel to make the roll angle to the target value Coordinated sideslip

In according with “Cooper_Harper” rating

Flight control system control law and the flying quality evaluation test 261

4.4.6 Preparation of test report The control law and flying quality evaluation test may go through multiple rounds of evaluation. It mainly depends on the experience of the designer and the aircraft development process supporting the test, the pilot evaluation level, and the construction of test conditions. It is impossible to pass the evaluation test at one time even when everything goes well. After the completion of each round of the evaluation test, the test report must be written in time to truthfully reflect the test process and test results. The content of the test report mainly includes test objective, test principle, test process, and test results. Test objectives and test principle can be briefly described, while the test process and test results shall be described in detail. The description of the test process mainly covers the preparation before the test, personnel arrangement, implementation time, appraisal method, and the implementation of the test task list. The sorting and analysis of test results is the core of the test report. According to the test task list, the pilot’s comments shall be summarized and analyzed, preliminary analysis conclusions of test results shall be given, and the problems and suggestions put forward for the test shall be classified and summarized to provide clear and real basis for the test analysis report. The comment form filled in by the pilot and the test task sheet filled in by the test flight engineer shall be archived together as annexes to the test report.

4.4.7 Test analysis report The analysis report of the flight control system control law and flying quality evaluation test is the main basis for the optimization and improvement of the flight control system and control law and aircraft aerodynamic characteristics. It is generally completed by a background aircraft flight control system designer or control law designer. For each round of the control law and flying quality evaluation test, the pilot will come up with a lot of opinions and suggestions, which can be harsh criticism sometimes. The designer must take them seriously. However, affected by the flight environment and natural physiology and psychology, the pilot may give comments that differ minimally or largely, or even give conclusions contrary to the mathematical simulation. In such cases, the designer should analyze relevant conclusions in the test analysis report and reach a consensus. The test analysis report focuses on summarizing and analyzing the pilot’s comments, classifying the comments according to relevant evaluation provisions, drawing the state points and improvement trend of the pilot’s comments of the different categories on the altitude-speed envelope and parameter adjustment gain curve chart, respectively, confirming the optimization and improvement of the state envelope, and adjusting the parameters to a reasonable range according to the calculation results. If the pilot’s comments lead

262 Chapter 4 to a large slope change of the parameter adjustment gain curve, it is necessary to communicate and confirm with the pilot proposing the comments, and conduct the test again if necessary. Meanwhile, the pilot may put forward many man machine work efficiency-related cockpit layout and display control problems. These problems should be notified in a timely manner to the personnel of relevant disciplines. If they have real impact on the evaluation test, the engineering simulator shall be improved according to the actual conditions of the aircraft.

4.5 Data collection, processing, and evaluation methods 4.5.1 Overview The results of the flight control system control law and flying quality evaluation test cover two aspects, the evaluation opinions (subjective evaluation and suggestions for improvement) of the pilot and the flight data and curves (objective analysis) recorded in the test process. The correct understanding of pilot’s subjective evaluation and suggestions and overall recording and analysis of test data are crucial for the optimization and improvement of the control law.

4.5.2 Requirements for data collection The requirements for data collection mainly include the requirements for the type of collected parameters (physical), data sampling rate, and precision. The parameters shall include: 1. Control actions of pilot, such as action method, action time, and action amplitude. 2. Operating modes and internal parameters of the flight control system, such as channel information and voting values. 3. Aircraft control surface information, such as deflection angle and speed. 4. Flight parameters, such as altitude, speed, and dynamic pressure. 5. Aircraft attitude parameters, such as angle of attack, angle of sideslip, angle of pitch, roll angle, and yaw angle. 6. Engine status parameters, such as speed, fuel consumption, and exhaust temperature. 7. Landing gear system parameters, such as wheel load and wheel speed. 8. External injection information, such as initial status information. 9. Fault information, such as flight control system fault and engine fault. The parameters above may be switch quantity, digital quantity, analog quantity, and data bus signal. For the convenience of analysis, the consistency of signals in time shall be ensured and the reflective memory network with good real-time performance should be

Flight control system control law and the flying quality evaluation test 263 used as a unified source for data reception. Generally, the sampling rate should be no less than 25 Hz, and 50 Hz would be better. It is unnecessary to have a higher sampling rate.

4.5.3 Requirements for data processing The purpose of collecting test data is to realize data review, video/audio playback, control motion capture and positioning, and fast offline data identification. Data review, video/ audio playback, and control motion capture and positioning are mainly used for test appraisal and evaluation. When test results are inconsistent with the pilot’s evaluation results, data review and video/audio playback analysis can be conducted to reach consistency. Fast offline data identification is mainly used for flying quality analysis and objective evaluation according to relevant standards. Through data review, the specific action point can be confirmed based on video/audio playback by extracting a set of parameters of interest, such as longitudinal control and response parameters. Therefore the timer for the test data and the video/audio data must be synchronized. The control motion capture and positioning can help extract typical aircraft responses, analyze system delays and determine control law gains, and improve the efficiency of test data analysis. It is especially important when the data volume is large and the processing cycle is very long. The control motion capture and positioning shall be able to capture the 14 control motions described in section 4.4.4. By adopting the standard control motion model and data normalization and the method of minimum variance, all the desired control actions can be found after two cycles, and the optimal control motion can be found after six loop iterations.

4.5.4 Objective evaluation methods The control law engineer shall attach importance to the objective evaluation of test results. In other words, according to the flying quality specifications of military aircraft and the design requirements for the control law of the flight control system, the flying quality can be evaluated through test data identification and the evaluation results can be compared with the results of mathematical simulation so as to provide a basis for subsequent optimization and improvement. The evaluation method described in Flying Qualities of Piloted Airplanes (Fixed Wing) (GJB185—86) and Flying Qualities Standard for Airplane with Fly-by-Wire Control System (GJB2874—97) is generally used for the flying quality evaluation of military aircraft. The evaluation shall cover at least the following content: 1. Longitudinal flying qualities a. Longitudinal static stability b. Long period stability

264 Chapter 4 c. Flight path stability d. Aircraft short period stability e. Control feeling in maneuvering flight 2. Heading flying qualities a. Heading mode characteristics b. Heading dynamic response characteristics c. Roll control performance d. Heading control performance e. Heading characteristics under steady sideslip f. Heading control in go-around g. Heading control in diving h. Heading control under asymmetric thrust It should be noted that the items above are mainly put forward for requirements on flying qualities of military land-based aircraft. For civil aircraft, seaplanes, carrier-based aircraft, and unmanned aerial vehicles, Airworthiness Standard for Transport Aircraft (CCAR25), Specification of Carrier-based Aircraft—Flight Performance (GJB3718—1999), Specification of Carrier-based Aircraft—Flying Qualities (GJB3719—1999), General Specification of Unmanned Aerial Vehicles (GJB2347—1995), and General Specification for Unmanned Aerial Vehicle Flight Control and Management System (GJB5201—2003) shall be taken as reference. By the functions of the flight control system, the objective evaluation of the control law is divided into control law evaluation of the manual flight control system, control law evaluation of the high lift control system, and control law evaluation of the automatic flight control system. 1. Objective control law evaluation of manual flight control system: parameter identification is performed for test data to get corresponding flying quality indexes. Mathematical simulation results are compared with pilot’s subjective evaluation to verify mathematical simulation results, laying the basis for providing the frequently used environment for the “iron bird” integration test, such as the human machine combined test environment and onboard ground test environment. 2. Objective control law evaluation of high lift control system: the evaluation content are mainly transition characteristics and modal conversion to transient state. During the design and mathematical simulation of the flight control system, the steady-state process can only be evaluated and the transition characteristics can be preliminarily predicted based on the wind tunnel test data, and then the characteristics will be put into the flight envelope to give the pilot basic feelings, which will be further improved after the flight test obtains the real relevant aerodynamic data of the aircraft. Before the maiden flight, the steady-state performance of the high lift control system should be

Flight control system control law and the flying quality evaluation test 265 confirmed through high-speed taxiing and a front wheel lifting test to ensure the safety of the maiden flight and subsequent test flights. 3. Objective control law evaluation of automatic flight control system: the preliminary functional and performance verification is carried out by experienced designers and the pilot completes the final operational evaluation. It should be pointed out that if the evaluation is carried out on an engineering simulator without a motion sensing system, the aircraft transition and modal conversion to a transient state can only be objectively evaluated through theoretical calculation and the control logic can be evaluated by comparing the test results with the expected results.

4.5.5 Subjective evaluation methods The common subjective evaluation methods in the world are Cooper Harper rating scale and PIO trend assessment scale, which are also regulated in Flying Qualities of Piloted Airplanes (Fixed Wing) (GJB185—86) and Flying Qualities Standard for Airplane with Fly-by-Wire Control System (GJB2874—97). 1. Cooper Harper rating scale The Cooper Harper rating scale is shown in Fig. 4.5. It conducts a subjective evaluation mainly from two aspects: the maneuverability of aircraft and the burden of the pilot to complete the flight mission. The characteristics of the aircraft are described and different PR values (from 1 to 10) of the rating scale on requirements for the pilot in a mission or operation are provided. Their specific meaning is described below: a. PR 5 1, it means good and satisfactory flying quality. All characteristics are ideal and well-coordinated. b. PR 5 2 3, it means the mission can be completed well but there is a little room for improvement. It may be caused by difference in the pilot’s habits or other insignificant issues. c. PR 5 4 6, it means the flight control system or aircraft must be improved. The remarkable characteristic is that the mission can be basically completed through the pilot’s endeavor, but the workload is huge and a lot of energy is required. PR 5 7 9, it means quite poor flying quality and the limit state has been reached. The pilot has tried his best but still fails to complete the mission well. d. PR 5 10, it means maneuvering failure, a very abnormal phenomenon. It only occurs in a certain fault state or under certain conditions. The result is that the mode or flight state must be changed to avoid flight accidents. 2. PIO trend assessment scale The use of the PIO trend assessment scale may cause PIO of the aircraft. The rating criteria and procedures are shown in Fig. 4.6. “Undesired motion” refers to a single

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Figure 4.5 Cooper Harper rating scale.

overshoot or small, fast, and damped periodic motion, and “oscillation” refers to a motion more than half a cycle or one overshoot.

4.6 Management of control law and the flying quality evaluation test The flight control system control law and flying quality evaluation test involves a wide range of expertise, many participants, long duration, complex data records, and large workload of data analysis. In order to better obtain the test conclusion and optimize it and ensure the safety of the test (such as the safety interlocking of the motion system), scientific

Flight control system control law and the flying quality evaluation test 267

Figure 4.6 Pilot-induced oscillations trend assessment scale.

test management should be carried out, which covers test organization, division of labor, and safety guarantee measures.

4.6.1 Planning of test Following the basic idea of staged implementation of the overall planning, emphasis shall be placed on the mathematical simulation and engineering simulator test, “iron bird” integration test, onboard ground test, and flight test as well as on the major problems that should be solved regarding the flight control system control law and flying quality evaluation. The test task description, test outline, test task list, test report, and test analysis report shall be carefully reviewed and validated to estimate the differences between potential evaluation opinions and design targets and make a plan of

268 Chapter 4 an additional test as soon as possible. When necessary, an industry expert review meeting shall be held to review the test documents in different stages.

4.6.2 Preparation for test Before the flight control system control law and flying quality evaluation test, all preparations for the test must be made, covering the state of test participants and test equipment (engineering simulator) or the state of the “iron bird” integrated test bed. The specific work is as follows. 1. Test participants They must pass strict posttraining and undertake corresponding postwork, be familiar with relevant test technical documents, and master test objectives, test principles, test processes, and the test environment. They are mainly the pilot, test flight engineer, system engineer, control law engineer, and equipment engineer. a. Pilot The pilot participating in the test must have a solid level of flight mechanics and control theory, as well as long-term flight experience, preferably as a member of the chief pilot team. During the test, the pilot is responsible for completing the task list, evaluating the control law and flying quality as required, and filling in the pilot opinions form. b. Test flight engineer The flight test engineer must have some knowledge of flight mechanics, control theory and avionics, understand aircraft and flight control systems, and possess some flight knowledge. They are responsible for the training of test pilots, explaining the task list, filling in the task list according to the opinions of the test pilots, and cooperating with designers to complete the flight test summary. c. Control law engineer The control law engineer must be a control law development engineer and a flying quality design engineer of the background aircraft, who is mainly responsible for setting flight status, modifying control law parameters, filling in the parameter change form, and recording and analyzing test data. d. System engineer The system engineer must be a design engineer of the flight control system and other related aircraft systems of the background aircraft, who is responsible for fully describing the aircraft system to pilots and other test participants, providing mathematical models and control logic of related aircraft systems and airborne equipment, solving problems of the system in the test process, and carrying out the evaluation results of the pilot on the aircraft system in the system design in a timely manner.

Flight control system control law and the flying quality evaluation test 269 e. Equipment engineer The equipment engineer must be a developer of the engineering simulator used for the test, who shall be able to ensure the consistency between the equipment status and the test items, be responsible for the opening, closing, and safe and reliable operation of the equipment, operating the integrated control management system, and controlling the test process. f. Safety officer The safety officer is responsible for patrol inspection at the test site and also cooperate with the commander for timely organization and dealing with any special situations such as fire, mechanical abnormality, and personnel entering a dangerous area, so as to ensure the safety of personnel and equipment. g. Commander The commander is the authoritative person in charge of the test, who is responsible for the coordination of test participants, test schedule, task arrangement, command of the test process, and supervision of the test’s quality and safety. 2. Test equipment a. Artificial feel simulation system: it shall ensure the load mechanism is at the required technical status for the test. b. System model: it shall ensure the requirements for airborne equipment, such as sensor, actuator, engine, and landing gear, and the system model meet the requirements of the test. c. Control law state: the control law state has completed the full mathematical simulation evaluation and reached the corresponding quality design requirements. d. Flight simulation system: the flight envelope state run by the flight simulation system is consistent with the test requirements, including aircraft state, engine state, control law state, and related system state. e. Instrument and vision system: the instrument display and vision system shall be validated to ensure correct cockpit display. f. Network transmission: the network shall be tested to ensure correct data transmission and transmission delay within the specified range.

4.6.3 Control of test process Test process control mainly covers pretest inspection, test implementation, and posttest evaluation. 1. Pretest inspection: check cockpit equipment to ensure instrument display, visual display, and communication/recording system work normally; check the console and flight simulation model to ensure the effective control of the console on the test and confirm the version of the control law is the version for the test; fully communicate with the pilot on test items, test methods, and evaluation criteria.

270 Chapter 4 2. Test implementation: before each operation, the console selects and confirms the test status. The test pilot confirms that the landing gear control handle, flaps (slats) control handle, and the switch on the flight control panel are at normal position. The pilot completes the regulated control actions in the test task list and fills in the list with the evaluation results according to requirements. The console maintains communication with the cockpit to observe and record test data synchronously. 3. Posttest evaluation: after the pilot steps out of the cockpit, communicate with the pilot in time, answer the questions raised by the pilot and analyze the flight phenomena together, record questions and suggestions, gather comments of the pilot, keep the flight data, close the cockpit, and turn off the console.

4.6.4 Summary of test After completion of each test, the pilot evaluation opinion form and recorded data shall be well managed, the test data shall be analyzed and processed, and the test report and test analysis report shall be prepared to draw lessons and prepare for the next test.

CHAPTER 5

Combined test of the flight control subsystem 5.1 Overview The flight control system is one of the key systems with the highest requirement for aircraft safety, the most complex cross-linking relationship, the most airborne equipment, and the most stringent requirements for verification. To ensure the quality of development, resolve technical risk, and speed up the development progress, the flight control system shall generally adopt the thought of “hierarchical design and validation level by level,” emphasize the key point for development in each stage and make comprehensive verification, and give full play to the initiative and advantages of the main machine manufacturer and the developer of the subsystem and airborne equipment, so as to ensure the smooth development of the aircraft. “Hierarchical design” is embodied in three aspects, that is, main machine manufacturer, developer of subsystem, and developer of airborne equipment. First level: as the main machine manufacturer, based on the in-depth study of all top-level requirements such as aircraft conceptual design requirements and flying quality design requirements, it should confirm the design requirements of the flight control system, form design specifications of the flight control system and the overall technical scheme of the flight control system, further decompose the design requirements of the flight control system, divide the system module or subsystem scientifically, and put forward the design requirements of each subsystem, such as fly-by-wire flight control system, automatic flight control system (AFCS), high-lift system, cockpit control device system, mechanical backup control system, and control law, as well as airborne software. Second level: as a subsystem developer, on the basis of in-depth analysis of top-level design requirements such as subsystem design requirements and specifications in related disciplines, it should confirm the design specifications of the constituent subsystem and form the subsystem design scheme and airborne equipment design requirements. Third level: as an airborne equipment developer, on the basis of in-depth analysis of top-level design requirements such as airborne equipment design requirements and specifications in related disciplines, it should confirm the design specifications of constituent subsystem airborne equipment and form the scheme of airborne equipment.

Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00005-X © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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272 Chapter 5 The work above is actually the validation process of the flight control system, as part of the system design process. The conformity of the design results can be verified by the digital simulation method, which is not described in detail in this book. “Validation level by level” is also embodied in three aspects, that is, developer of airborne equipment, developer of subsystem, and main machine manufacturer. First level: as an airborne equipment developer, it should complete the manufacturing of parts and synthesis of airborne equipment, to verify and validate that airborne equipment meets the design requirements. Second level: as a subsystem developer, it should complete the synthesis from airborne equipment to subsystem, to verify and validate the subsystem meets the design requirements. Third level: as the main machine manufacturer, it should complete the synthesis from subsystem to system, to verify that the system meets the requirements of the system design specification, and then meet the requirements on the flight control system of the overall and flying quality design of the aircraft. The work above is actually the verification process of the flight control system. The verification technology at the first level has been introduced in chapter 2 of this book, the verification technology at the second level is to be introduced in this chapter, and the verification technology at the third level will be introduced in chapters 6 and 7. The main purpose of the comprehensive verification of the flight control system subsystem is to verify whether the function, performance, and environmental adaptability of the subsystem meet the design requirements of the subsystem under the subsystem test environment. Through the verification test, the design defects and problems of the subsystem are exposed in advance and the potential failure modes and mechanisms of the subsystem are identified. Corrective measures are taken to increase the reliability of the subsystem. Static verification is carried out to verify the correctness and compatibility of interfaces between the subsystem and other subsystems and other aircraft systems. The subsystem developer is responsible for the combined test of the flight control subsystem, which is carried out according to the internal development process and quality management system of the developer and is generally implemented at the location of subsystem development. The flight control system of a large aircraft mainly requires the following subsystem tests: 1. 2. 3. 4. 5.

Combined test of pilot control units (PCU). Combined test of fly-by-wire flight control system. Combined test of high-lift system. Combined test of AFCS. Combined test of machinery control system (if any).

Combined test of the flight control subsystem 273

5.2 Combined test of the pilot control units 5.2.1 System introduction The cockpit control system of large aircraft generally includes the longitudinal control device, lateral control device, and heading control device. Longitudinal and lateral control devices are generally central disc (column) type, central long or short rod type, and side rod type (active side rod and passive side rod). The heading control device is generally designed with a front wheel steering control device and wheel brake control device. The central disc (column) type includes distributed type and integrated type. The cockpit control system of each aircraft is basically similar. The basic functions of the cockpit control system include: 1. converting the mechanical control command to electric signal under the fly-by-wire control mode. 2. providing the pilot with appropriate control feel (rod force displacement). 3. completing the trim and autopilot on/off and synchronization on horizontal stabilizer through the switches and press buttons set on the steering wheel. 4. realizing the trim of aileron and rudder through the aileron and rudder trim button on trim control panel. 5. for aircraft with machinery control system, controlling the deflection of the corresponding control plane through the connection between the cockpit control system and mechanical system with the pull rod. Distributed central disc (column) type longitudinal and lateral control devices and heading control devices comprehensively designed with a front wheel steering control device and wheel brake control device are taken as examples to introduce technical content about the testing of the cockpit control system. Longitudinal, lateral, and heading three-axis main control devices are mainly introduced. Other auxiliary control devices have simple functions and principles and the test method and judging criteria are basically similar; thus they are not described in this book. 5.2.1.1 Cockpit lateral control channel The left and right pilots realize the lateral control of the aircraft by controlling the two conventional steering wheels. In normal operation, the left and right steering wheels are connected with a set of link mechanisms to realize the synchronous movement of the left and right steering wheels. The link mechanism is provided with an arming mechanism or a spring pull rod to realize the separation between the left and right lateral control devices in case of an emergency. The trim control panel on the central console controls the motion of the aileron trim mechanism to realize the lateral trim of the aircraft.

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Figure 5.1 Cockpit lateral control channel.

Under the fly-by-wire control mode, the cockpit lateral control channel converts the pilot’s mechanical control commands into electric signals through the steering wheel displacement sensor and force sensor, and the electric signals are transmitted to PFC through ACE for the control law calculation of the flight control system, thus controlling the deflection of the corresponding control plane. The cockpit lateral control channel includes the steering wheel, lateral load mechanism, aileron trim mechanism, lateral arming mechanism, hydraulic damper, steering wheel angular displacement sensor, steering wheel force sensor, pull rod, and base. The diagram of the cockpit lateral control channel is shown in Fig. 5.1. 5.2.1.2 Cockpit heading control channel Left and right pilots realize the heading control of the aircraft through two conventional vertical pedals. The pedal position can be adjusted back and forth through the pedal adjustment motor

Combined test of the flight control subsystem 275

Figure 5.2 Cockpit heading control channel.

(it can be manually adjusted sometimes) to meet the needs of pilots with different body shapes. Through the trim control panel on the central console, the motion of the rudder trim mechanism can be controlled to realize the heading trim of the aircraft. In normal operation, the left and right pedals are connected through a set of link mechanisms to realize the synchronous movement of the left and right pedals. The link mechanism is generally provided with a spring pull rod to realize the separation between the left and right pedals in case of an emergency. Under the fly-by-wire control mode, the cockpit heading control channel converts the pilot’s mechanical control commands into electric signals through the pedal displacement sensor and force sensor, and the electric signals are transmitted to PFC through ACE for the control law calculation of the flight control system, thus controlling the deflection of corresponding control plane. The cockpit heading control channel includes the pedal assembly, heading load mechanism, rudder trim mechanism, hydraulic damper, pedal command displacement sensor, pedal force sensor, spring pull rod, pull rod, and base. The diagram of the cockpit heading control channel is shown in Fig. 5.2. 5.2.1.3 Cockpit longitudinal control channel Left and right pilots control the elevator of the aircraft through two conventional central control columns. In normal operation, the left and right control columns are connected

276 Chapter 5 through a set of link mechanisms to realize the synchronous movement of left and right control columns. The link mechanism is generally provided with an arming mechanism and or a spring pull rod to separate the left and right control columns in case of emergency. Under the fly-by-wire control mode, the cockpit longitudinal control channel converts the pilot’s mechanical control commands into electric signals through the control column displacement sensor and force sensor, and the electric signals are transmitted to PFC through ACE for the control law calculation of the flight control system, thus controlling the deflection of the corresponding control plane. The cockpit longitudinal control channel includes the control column assembly, longitudinal load mechanism, longitudinal arming mechanism, hydraulic damper, control column displacement sensor, pull rod, and base. The diagram of the cockpit longitudinal control channel is shown in Fig. 5.3. Most large aircraft adopt a horizontal stabilizer to achieve longitudinal trim. They do not set a trim mechanism similar to the lateral or heading channel, but realize longitudinal automatic trim through the driving of the actuator of the horizontal stabilizer.

5.2.2 Test objectives The purpose of the combined test of the cockpit control device subsystem is to obtain a preliminary synthesis of the system under the subsystem test environment and to check and verify the subsystem function, performance, interface, and index conformity. The strength test is carried out according to the service load, and the durability test of the subsystem is carried out to test the durability index of the subsystem.

5.2.3 Test requirements 5.2.3.1 Requirements for the tested object All tested objects are “S” type parts, whose technical statuses and indexes shall be inspected by the manufacturer and ensured that they meet relevant technical requirements, and relevant certificates of them shall be submitted. After the tested objects are installed and commissioned according to relevant drawings and technical conditions, it shall be ensured that the control channel moves steadily without an obvious idle stroke and all objects work normally through visual inspection when the steering wheel, control column, and pedal are operated within the full stroke. The null output voltage of each command sensor shall be measured and shall meet the requirements. If it does not meet the requirements, the null position of the command sensor shall be adjusted.

Combined test of the flight control subsystem 277

Figure 5.3 Cockpit longitudinal control channel.

5.2.3.2 Environmental requirements for the combined test The supporting stiffness of the test bed shall be greater than 1 3 107 N/m and the installation of test pieces and test equipment shall be considered. 5.2.3.3 General requirements for test equipment 1. Requirements for accuracy and measuring range of test system Test system accuracy and measuring range shall meet the test requirements of the combined test of cockpit control system. Table 5.1 shows the requirements for accuracy and measuring range of the test system for the cockpit control system test of a type of aircraft.

278 Chapter 5 Table 5.1: Requirements for accuracy and measuring range of test system. Items Angular displacement Tension force Torque Linear displacement sensor Acceleration Signal generator

Accuracy (%F.S.)

Measuring range

1 1 1 1 1 1

6 90 6 200 kg 6 100 N m 6 75 mm 6 200 g G

2. Requirements for test times In the case of good repeatability of test data (the difference is less than 5%), repeat the test for three times and take the average value. In the case of poor repeatability of test data (the difference is greater than 5%), repeat the test for no less than five times and take the average value. 3. Sampling frequency The sampling interval of the test system shall not be greater than 5 ms.

5.2.4 Test items and methods 5.2.4.1 Performance test of the lateral control channel 1. Motion clearance inspection Rotate the steering wheel clockwise and anticlockwise respectively to the maximum and then return the steering wheel back to the neutral position at a constant and slow speed. At the same time, measure and record the rotating angles of the left and right steering wheels and the output of the steering wheel displacement sensor. Conduct the operation above for left and right steering wheels, respectively. Disconnect the hydraulic damper and repeat the above operations. 2. Force displacement test of steering wheel Rotate the left steering wheel from the neutral position clockwise to the limit position at a constant and slow speed, and then rotate anticlockwise back to the null position and then to the limit position. The total operation time within the full stroke is about 80 100 s. At the same time, measure and record the rotating angle of the left and right steering wheels, the rotating angle of the mechanical input rocker arm, the output of the steering wheel displacement and force sensor, the steering wheel command displacement and force sensor, and the force at the measuring point of the left steering wheel. Repeat the operation above for the right steering wheel. Disconnect the hydraulic damper and repeat the operation above.

Combined test of the flight control subsystem 279 3. Inspection of aileron trim lamp area Press the “left wing down” button on the trim control panel and release the button when the “aileron neutral position” lamp is off. Then press the “right wing down” button on the trim control panel and release the button when the “aileron neutral position” lamp is off. 4. Inspection of trim range Take the neutral position of the steering wheel as the start point and press the switch of trim mechanism until the steering wheel stops rotating. At the same time, record the maximum rotating angle of the left and right steering wheels clockwise and anticlockwise, the time needed, the trim indication position, and the output of the steering wheel force and displacement sensor. 5. Release force and release angle test Fix the left steering wheel at the neutral position and then rotate the right steering wheel clockwise until the two steering wheels release. At the same time, measure and record the angle when the steering wheels release, the force at the measuring point of the steering wheel, and the output data of the steering wheel force sensor, and record whether the release signal is normally transmitted. Then, restore the arming mechanism to connect the two steering wheels, rotate the right steering wheel anticlockwise, and repeat the operation above. Fix the right steering wheel at the neutral position and repeat the test above. 6. Damping characteristic test of steering wheel Rotate the steering wheel clockwise until it reaches the limit position and hold it at the limit position. After releasing, record the response curves of the rotating angle of the left and right steering wheels. Rotate the steering wheel anticlockwise until it reaches the limit position and repeat the operation above. Calculate the relative damping coefficient with the test curve. Conduct the operation above for both left and right steering wheels. Disconnect the hydraulic damper and repeat the operation above. 5.2.4.2 Performance test of the heading control channel 1. Motion clearance inspection Push the pedal back and forth, respectively, to the limit position and then return the pedal back to the neutral position slowly. At the same time, measure and record the rotating angle of the left and right pedals and the output of the pedal command displacement sensor. Conduct the operation above for both left and right pedals. 2. Pedal force and pedal displacement characteristic test Set the pedal and rudder trim mechanism at the neutral position. Push the left pedal of the left pilot to the limit position from the neutral position, and then return to the neutral position and push the right pedal of the left pilot to the limit position. The total operation time within the full stroke is about 80 100 s. At the same

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3.

4.

5.

6.

time, measure and record the displacement of the pedal of the left and right pilots, the force at the measuring point of the pedal of the left pilot, the rotating angle of the mechanical input rocker arm, and the output of the pedal command displacement sensor and directive force sensor. Analyze the starting force, dead zone, friction force, and the static characteristics of the control input to the command sensor with the test curve. Repeat the operation above for the pedal of the right pilot. Disconnect the hydraulic damper and repeat the operation above. Pedal adjusting range test Set the pedal and rudder trim mechanism at the neutral position. Press the switch of the adjusting mechanism of the pedal of the left pilot to make the pedal move to the front and back limit position, respectively, and then record the pedal motion stroke. Repeat the operation above for the pedal of the right pilot. Pedal emergency control test Set the pedal and rudder trim mechanism at the neutral position. Fix the pedal of the left pilot at the neutral position, and then push the left and right pedals of the right pilot to the limit position respectively. At the same time, measure and record the displacement of the pedal of the right pilot, the force at the measuring point of the pedal of the right pilot, and the output of the pedal command displacement sensor and pedal force sensor. Fix the right pedal at the neutral position and repeat the operation above. Trim range test Take the neutral position of the pedal as the start point and press the switch of trim mechanism until the pedal stops moving. At the same time, record the maximum angle of the pedals of the left and right pilots to move back and forth, the time needed, and the trim indication position. Pedal damping characteristic test Push the left pedal of the left pilot forwards until it reaches the limit position and hold it at the limit position. After release, record the response curves of the displacement of pedal of the left and right pilots. Push the right pedal of the left pilot forward until it reaches the limit position and repeat the operation above. Calculate the relative damping coefficient with the test curve. Repeat the operation above for pedal of the right pilot. Disconnect the damper and repeat the operation above.

5.2.4.3 Performance test of the longitudinal control channel 1. Motion clearance inspection Push the control column back and forth to the limit position, respectively, and then return the control column back to the neutral position slowly. At the same time, measure and record the rotating angle of the left and right control columns and the output of the control column command displacement sensor.

Combined test of the flight control subsystem 281 Conduct the operation above for both left and right control columns, respectively. 2. Force and displacement characteristic test of control column Push the left control column from the neutral position to the limit position slowly, and then return to the neutral position and pull back to the limit position. At the same time, measure and record the displacement of the left and right control columns and the force at the measuring point of the left steering wheel. Analyze the starting force, dead zone, friction force, and the static characteristics of the control input to the command sensor with the test curve. Repeat the operation above for the right control column. Disconnect the hydraulic damper and repeat the operation above. 3. Release force and release angle test Fix the left control column at the neutral position and push the right steering wheel forwards until the two control columns release. At the same time, measure and record the angle when the control columns release and the force at the measuring point of the steering wheel, and record if the arming signal is normally sent out. Restore the connection between the two control columns, pull the right control column back and repeat the operation above. Fix the right control column at the neutral position and repeat the operation above. 4. Damping characteristic test of control column Push the control column until it reaches the limit position and hold it at the limit position. After release, record the response curves of the displacement of the left and right control columns. Pull the control column back until it reaches the limit position and repeat the operation above. Calculate the relative damping coefficient with the test curve. Conduct the operation above for both the left and right control columns. Disconnect the hydraulic damper and repeat the operation above. 5.2.4.4 Durability test Based on the fatigue load displacement spectrum block of the cockpit control system provided by the aircraft strength discipline, a durability test is carried out for lateral, heading, and longitudinal control channels within a specified time. The fatigue load displacement spectrum block of the cockpit control system of a type of aircraft is shown in Table 5.2. “Neutral position - specified positively biased position - neutral position - specified negatively biased position - neutral position” is a fatigue load displacement cycle. After the completion of each spectrum block test, a force and displacement characteristic test will be conducted to draw force displacement characteristic curve and transmission accuracy curve, so as to check the force displacement characteristic and transmission accuracy changes of the control channel. After each spectrum block test of the lateral and heading control channels is completed, 10 times of full-stroke trim control will be performed for the aileron and rudder trim mechanism,

282 Chapter 5 Table 5.2: Fatigue load displacement spectrum block of cockpit control system of a type of aircraft. Control load /maximum Control displacement /maximum control load % control displacement % 10 50 100

10 50 100

Total control times per 100 flight hours

Control times per 100 flight hours /time

Proportion in total control times /%

3889 1611 156 5656

68.76 28.48 2.76 100

respectively. A full stroke refers to the process of “Neutral position - specified positively biased position - neutral position - specified negatively biased position - neutral position.” After each spectrum block test of the lateral control channel is completed, fix the control device at the side of the captain and rotate the steering wheel at the side of copilot clockwise and anticlockwise to perform an arming test. Then, fix the control device at the side of the copilot and rotate the steering wheel at the side of the captain clockwise and anticlockwise to perform an arming test. After each spectrum block test of the longitudinal control channel is completed, fix the control device at the side of the captain, and push and pull the control column at the side of copilot to have an arming test. Then, fix the control device at the side of the copilot and push and pull the steering wheel at the side of the captain to perform an arming test.

5.2.5 Criteria for the assessment of test results 5.2.5.1 Performance test of the lateral control channel 1. Motion clearance inspection The motion clearance of the steering wheel is just the angle difference between the steering wheel after it is returned to and stopped at the neutral position after being rotated clockwise and anticlockwise. The motion clearance of the steering wheel and the steering wheel command displacement sensor shall meet the design requirements of the cockpit control system. 2. Force displacement test of steering wheel Force displacement curve of left and right steering wheels: take the position with zero rod force of the steering wheel as the origin of coordinates to draw the force displacement curve of the steering wheel at the input end and determine the steering wheel’s clearance, starting force, starting stroke, maximum force, and frictional force. The test results include the maximum rotating angle of steering wheel, starting force, frictional force, maximum force, clearance, starting stroke, and the rotating angle

Combined test of the flight control subsystem 283 of the mechanical control output rocker arm, all of which shall meet the design requirements of the cockpit control system. Motion independence of steering wheel: when the steering wheel is rotating in full stroke, record the output curve of steering wheel directive force, control column directive force, and command displacement at the free end. The test results include the control column directive force, the control column command displacement, and the steering wheel directive force at the free end, all of which shall meet the design requirements of the cockpit control system. Directive force displacement characteristic curve of steering wheel: take the position with zero rod force of the steering wheel as the origin of coordinates to draw the directive force displacement curve of the steering wheel at the input end and determine the clearance from the steering wheel to the command displacement sensor, starting force, starting stroke, maximum force, and frictional force. The test results include the maximum output voltage of sensor, starting force, frictional force, maximum force, clearance, and starting stroke, all of which shall meet the design requirements of the cockpit control system. Transmission accuracy curve: take the rotating angle of the steering wheel at the input end as abscissa to draw the curve of the steering wheel relative to the rotating angle of the command sensor and the steering wheel at the free end and determine the clearance. The test results include the rotating angle of the steering wheel at the input end, the rotating angle of the steering wheel at the free end, maximum difference between left command displacement sensor channels, maximum difference between right command displacement sensor channels, maximum difference between left and right command displacement sensors, clearance between steering wheels, clearance between left command displacement sensors, and clearance between right command displacement sensors, all of which shall meet the design requirements of cockpit control system. 3. Inspection of aileron trim lamp area Record the motion stroke of the steering wheel and steering wheel command displacement sensor in two operations. The test results include the motion angle of steering wheel and the motion voltage range of the steering wheel command displacement sensor, all of which shall meet the design requirements of the cockpit control system. 4. Trim range test Data obtained from the analysis of test results include trim range, rotating angle of steering wheel, and the static characteristics of the steering wheel displacement sensor and steering wheel force sensor relative to aileron trim mechanism displacement sensor. Test results include full-stroke trim time, rotating angle of left and right steering wheels, rotating angle of left and right steering wheel command sensors, and the maximum force of the left and right steering wheel directive force sensors, all of which shall meet the design requirements of the cockpit control system.

284 Chapter 5 5. Release force and release angle test Data or curves that can be obtained from the analysis of test results include the release force and release angle when the two steering wheels release, the output of the steering wheel force displacement sensor, the steering wheel force and (test) steering wheel displacement curve, and the steering wheel force and (command) displacement sensor curve. Test results include the release force of the left and right steering wheels when they release clockwise and anticlockwise, respectively, the output voltage and the release angle of the directive force sensor, the release angle of the command displacement sensor, and the indication of the microswitch, all of which shall meet the design requirements of the cockpit control system. 6. Damping characteristic test of steering wheel According to the recorded test curve, the relative damping coefficient of the steering wheel lateral control channel can be calculated and it shall meet the design requirements of the cockpit control system. 5.2.5.2 Performance test of the heading control channel 1. Motion clearance inspection The motion clearance of the pedal is just the angle difference between the pedal after it is returned to and stopped at the neutral position after being operated back and forth. The motion clearance of the pedal and the pedal command displacement sensor shall meet the design requirements of the cockpit control system. 2. Pedal force and pedal displacement characteristic test Force displacement curve of left and right pedals: take the position with zero pedal force of the pedal as the origin of coordinates to draw the force displacement curve of the pedal at the input end and determine the pedal’s clearance, starting force, starting stroke, maximum force, and frictional force. The maximum stroke, starting force, frictional force, maximum force, clearance, starting stroke of the pedal, and the rotating angle of the rudder mechanical backup output rocker arm obtained from the test shall meet the design requirements of the cockpit control system. Directive force displacement characteristic curve of pedal: take the position with zero pedal force of the pedal as the origin of coordinates to draw the directive force displacement curve of the pedal at the input end and determine the clearance from the pedal to command displacement sensor, starting force, starting stroke, maximum force, and frictional force. The maximum output voltage of the sensor, starting force, frictional force, maximum force, clearance, and starting stroke obtained from the test shall meet the design requirements of the cockpit control system. Transmission accuracy curve: take the stroke of pedal at the input end as abscissa to draw the curve of the pedal stroke relative to the rotating angle of command sensor and pedal at the free end and determine the clearance. The pedal stroke at the input end, the

Combined test of the flight control subsystem 285

3.

4.

5.

6.

pedal stroke at the free end, maximum difference between left command displacement sensor channels, maximum difference between right command displacement sensor channels, maximum difference between left and right command displacement sensors, clearance between pedals, clearance between left command displacement sensors, and clearance between right command displacement sensors obtained from the test shall meet the design requirements of cockpit control system. Pedal adjusting range test Record the motion stroke of left and right pedals and the output of pedal force sensor and pedal displacement sensor. The angle of the arm of force of the pedal at the neutral position, the front limit angle of petal, and the back limit angle of pedal obtained from the test shall meet the design requirements of cockpit control system. Pedal emergency control test Analyze the test curve to obtain the pedal force and pedal displacement curve and pedal force sensor and pedal displacement sensor curve. The emergency control force of the pedal obtained from the test shall meet the design requirements of the cockpit control system. Trim range test Data obtained from the analysis of test results include trim speed and trim range, the motion stroke of pedal, and the static characteristics of the pedal displacement sensor and pedal force sensor relative to the rudder trim mechanism displacement sensor. The high and low speed full-stroke trim time, full-stroke trim centering time, the forward rotating angle of left and right pedals, the stretching stroke of the command sensor, the back rotating angle of the left and right pedals, and the withdrawing stroke of the command sensor obtained from the test shall meet the design requirements of the cockpit control system. Pedal damping characteristic test According to the recorded test curve, the relative damping coefficient of the pedal heading control channel can be calculated and it shall meet the design requirements of the cockpit control system.

5.2.5.3 Performance test of the longitudinal control channel 1. Motion clearance inspection The motion clearance of the control column is just the angle difference between the control column after it is returned to and stopped at the neutral position after being operated back and forth. The motion clearance of the control column and the control column command displacement sensor shall meet the design requirements of the cockpit control system. 2. Force and displacement characteristic test of control column Force displacement curve of left and right control columns: take the position with zero rod force of the control column as the origin of coordinates to draw the

286 Chapter 5 force displacement curve of the control column at the input end and determine the control column’s clearance, starting force, starting stroke, maximum force, and frictional force. The maximum rotating angle of the control column, starting force, frictional force, maximum force, clearance, starting stroke, and the rotating angle of the elevator mechanical backup sector gear obtained from the test shall meet the design requirements of the cockpit control system. Motion independence of steering wheel: when the control column is rotating in full stroke, record the output curve of control column directive force, steering wheel directive force, and command displacement at the free end. The steering wheel directive force, the steering wheel command displacement, and the control column directive force at the free end obtained from the test shall meet the design requirements of the cockpit control system. Directive force displacement characteristic curve of control column: take the position with zero rod force of the control column as the origin of coordinates to draw the directive force displacement curve of the control column at the input end and determine the clearance from the control column to command displacement sensor, starting force, starting stroke, maximum force, and frictional force. The maximum stroke of the sensor, starting force, frictional force, maximum force, clearance, and starting stroke obtained from the test shall meet the design requirements of the cockpit control system. Transmission accuracy curve: Take the rotating angle of the control column at the input end as abscissa to draw the curve of the control column relative to the rotating angle of the command sensor and the control column at the free end and determine the clearance. The rotating angle of the control column at the input end, the rotating angle of the control column at the free end, maximum difference between left command displacement sensor channels, maximum difference between right command displacement sensor channels, maximum difference between left and right command displacement sensors, clearance between steering wheels, clearance between left command displacement sensors, and clearance between right command displacement sensors obtained from the test shall meet the design requirements of cockpit control system. 3. Release force and release angle test Data or curves that can be obtained from the test include the release force and release angle when the two control columns release, the output of the control column force displacement sensor, the control column force and displacement curve, and the control column force and displacement sensor curve. The release force of the left and right control columns when they release in back and forth operation, the output voltage and the release angle of the directive force sensor, the release angle of the command displacement sensor, and the indication of the microswitch obtained from the test shall meet the design requirements of the cockpit control system.

Combined test of the flight control subsystem 287 4. Damping characteristic test of control column According to the recorded test curve, the relative damping coefficient of the control column longitudinal control channel can be calculated and it shall meet the design requirements of the cockpit control system. 5.2.5.4 Durability test During the test, the mean time between failures of the cockpit control system (airborne equipment) and the replacement condition of the airborne equipment shall be recorded. If any faults are found and confirmed to be associated with the airborne equipment, the product can be replaced and the working time of the airborne equipment shall be recorded, and then the durability test of the cockpit control system can be continued. For airborne equipment with faults, the developer of the cockpit control system should coordinate with the airborne equipment manufacturer to effectively analyze the fault mechanism and propose improvement measures. After the improvement measures are taken, the durability test can be carried out for the airborne equipment independently. If the cockpit control system has great changes in state, the durability test shall be carried out again. For test results and any problems or faults in the test, dynamic and static performance analysis and reliability analysis shall be conducted and a conclusion of the durability test shall be given, and an implementation scheme for the continuous optimization of the airborne equipment shall be provided as well. The durability test of the cockpit control system shall meet the design time requirements for the cockpit control system.

5.3 Combined test of the fly-by-wire flight control system 5.3.1 System introduction The fly-by-wire flight control system of large aircraft generally includes a computer subsystem, sensor subsystem, actuator subsystem, cockpit display, and control subsystem. The following control functions of the aircraft can be realized through the elevator, aileron, rudder, horizontal stabilizer, and spoiler: 1. Three-axis control augmentation. 2. Flight envelope protection (stall warning and protection, normal overload protection, overspeed warning and protection, roll angle rate limit, slant angle protection, and pitch angle limit). 3. Slant/pitch attitude holding.

288 Chapter 5 4. 5. 6. 7. 8.

Three-axis manual trim. Automatic pitch trim. Assisted roll and deceleration in air of multifunctional spoiler. Manual and automatic lift destruction and drag increase of spoiler. Deflection limit of rudder.

The fly-by-wire flight control system of the background aircraft adopts the architecture of “digital fourfold-redundancy computer (PFC) 1 actuator controller (ACE),” as shown in Fig. 5.4, with the ability of working with faults for three times. The fly-by-wire flight control system has three working modes, that is, normal working mode, degraded working mode, and simulated backup working mode. 1. Normal working mode: when the PFC is normal and all cross-linking signals are effective, the fly-by-wire flight control system is under normal working mode. Under this mode, all functions of the fly-by-wire flight control system can be realized by controlling the aircraft’s main control plane through PFC and ACE. 2. Degraded working mode: when received signals such as atmospheric or inertial navigation data become invalid, it will transfer to degraded working mode automatically and the aircraft’s main control plane can be controlled through PFC and ACE. Under this mode, the functions of the fly-by-wire flight control system related to atmospheric or inertial navigation data will become invalid and the flying quality of the aircraft will be affected. 3. Simulated backup working mode: when three or more PFCs lose functions or the simulated backup working mode is manually selected, the fly-by-wire flight control system will work under simulated backup mode. Under this mode, the aircraft’s main control plane is directly controlled through ACE, the gains are adjusted according to the state of flaps, and the functions including three-axis stability augmentation control, three-axis manual trim, manual ground lift destruction and drag increase, assisted roll and deceleration in air of the multifunctional spoiler can be realized. Also, the flying quality of the aircraft is degraded. When the fly-by-wire flight control system is under normal working mode, the flying quality of the aircraft meets the first level standard requirements within the service flight envelope and meets the second level standard requirements outside the service flight envelope and within the available flight envelope. The flying quality of aircraft in special stage shall meet the design requirements on flying quality. Under the simulated backup mode, the flying quality meets the second level standard requirements within the flight envelope.

5.3.2 Test objective Combined test of fly-by-wire flight control system aims to verify and validate that the flyby-wire flight control system meets functional, performance, and safety requirements on a

Combined test of the flight control subsystem 289

Figure 5.4 Architecture of fly-by-wire flight control system for a large transport aircraft.

290 Chapter 5 simulation test bed that tries to best reflect the working environment of the fly-by-wire flight control system. The test objectives include: 1. 2. 3. 4. 5. 6. 7. 8.

verifying the correctness and compatibility of control channel interfaces;. verifying if the control logic meets design requirements. verifying the accuracy and comprehensiveness of BIT design. verifying the redundancy strategy. verifying if actuator system meets design requirements. verifying if the function and performance meet design requirements;. verifying the fault logic and alarm display. verifying and checking if the fly-by-wire flight control system and its airborne equipment meet the airworthiness conformity requirements. 9. completing the durability test of the fly-by-wire flight control system.

5.3.3 Test requirements 5.3.3.1 Requirements for the tested object All tested objects are “S” type parts, whose technical status and indexes shall be inspected by the manufacturer and assured to meet relevant technical requirements. It shall be ensured that they have passed the acceptance test before delivery and are provided with a conformity certificate or relevant record. All airborne equipment shall be installed on the corresponding test bed according to the drawings, the connecting cables between them shall be made as per the interface control documents of each system, and the length of cables may be increased or decreased according to the test environment. However, the electrical characteristics of cables shall be consistent with the airborne cables as far as possible. The installation of mechanical and electrical airborne equipment (cockpit airborne equipment, actuators, etc.) shall be as consistent as possible with that of other equipment on the aircraft. All performance indexes of the test equipment and the accuracy of the test sensor and instrument shall meet the test requirements. They shall have passed relevant metering certification and be within the validity period. 5.3.3.2 Environmental requirements for the combined test Unless otherwise specified, the combined test of the fly-by-wire flight control system shall be carried out in a normal laboratory, which shall cover the ambient temperature, humidity, atmospheric pressure, etc. It shall be ensured that the combined test environment provides relevant airborne equipment of the fly-by-wire flight control system and relevant aircraft system equipment with analog power to ensure their normal operation, sets an independent power distribution cabinet, and

Combined test of the flight control subsystem 291 reserves a power conversion interface for power conversion with the aircraft. It is required that the 28 V DC power is adjustable within 18 32 V and every circuit of the power supply is monitored and can have an on off setting. Among the circuits, the bus bar of the flight control system shall be available for the analog of instantaneous on off of 10 ms. The AC power shall adopt 220 V and the frequency shall be 50 Hz. The combined test environment shall provide simulated hydraulic energy to the airborne equipment of the fly-by-wire flight control system and ensure required flow, pressure, and contamination level.

5.3.3.3 General requirements for test equipment The main test equipment and requirements for the combined test of the fly-by-wire flight control system are described below: 1. Fly-by-wire flight control system tester and requirements To facilitate the combined test of fly-by-wire flight control system, a tester with relatively complete functions shall be configured. Its basic functions include: 1. Releasing and truncating various input and output signals of fly-by-wire flight control computer. 2. Releasing and truncating various input and output signals of ACE (actuator controller). 3. Detecting various signals through the make and break points set on the tester. 4. Injecting test signals such as orders or faults at the make and break points manually and completing the test under various modes. The tester of the fly-by-wire flight control system shall be set according to the redundancy level and it shall be ensured that the redundancy signal in the tester is not cross-linked. The test point design of the tester shall meet the requirements of the system static testing. If possible, there should be as many set testing points as is beneficial. The injection point design is mainly conducted for the redundancy management test. In addition to injecting signal, the grounding and suspension states of the signal shall also be given generally. Parameters monitored by the fly-by-wire flight control system tester shall at least include: 1. power characteristic value (primary and secondary power supply). 2. segment detection points of signal chains. 3. Analog quantity, discrete quantity, and digital quantity of front interface and rear interface of fly-by-wire flight control computer and the detection quantity on each SRU board. 4. value of the software node characteristic quantity of fly-by-wire flight control computer. 5. BIT detection values of sensor, computer, and servo actuation system.

292 Chapter 5 2. Actuator test bed and requirements The actuator of the fly-by-wire flight control system is installed on the actuator test bed. The basic requirements for the test bed are as follows: 1. The test bed shall be able to simulate the stiffness, including the installation stiffness of the actuator and the connection stiffness between the actuator and the control plane. 2. The test bed shall be able to simulate the quality and rotational inertia of the control plane. 3. It shall be ensured that the test bed can be equipped with a loading system and can simulate the aerodynamic load of the aircraft. Requirements for the loading system are described in detail in chapter 8 of this book. 3. Cockpit control device test bed and requirements The basic requirements of the combined test of the fly-by-wire flight control system on the cockpit control device test bed are as follows: 1. It shall be ensured that the test bed can be installed with all real airborne equipment of the cockpit control system and their relative position shall be consistent with that on the aircraft. 2. The test bed shall consider the installation of test sensors such as rod force sensor and rod displacement sensor. 4. Mechanical displacement signal generator The mechanical displacement signal generator is used to simulate the pilot’s operation. It gives the control column (steering wheel) and pedal an input order to complete the signal injection for pilot’s operation. The mechanical displacement signal generator is mostly the electrohydraulic servo system and its control accuracy, bandwidth, and stroke shall meet test requirements. The specific requirements are as follows: 1. The mechanical displacement signal generator shall include actuator cylinders of different strokes to meet different test requirements. 2. The mechanical displacement signal generator shall include two types of actuators: large-stroke low-bandwidth actuator and small-stroke high-bandwidth actuator. The large-stroke actuator with low frequency response has a bandwidth generally not less than 15 Hz and the small-stroke actuator with a high frequency response has a bandwidth generally not less than 30 Hz. 3. The mechanical displacement signal generator shall have a safety protection function under various faults. 5. Flight simulation system and requirements In the combined test of the fly-by-wire flight control system, the flight simulation system is used for the real-time resolving of the aircraft motion equation and engine thrust equation and to simulate the flight process of the aircraft. The use of a digital computer as a simulation computer may inevitably introduce a time delay, which should be minimized. The period for a digital computer to

Combined test of the flight control subsystem 293 resolve the 6-DOF aircraft motion equation and engine thrust equation is generally controlled to be within 6 ms. The simulation computer shall have enough storage capacity to store all kinds of flight parameters online. It must have a high-speed A/D interface for the input of analog signals from external sensors and a high-speed D/A interface for driving simulation equipment such as rate turntable, angle of attack turntable, acceleration turntable, and loading system. 6. External cross-linking system simulator and requirements The fly-by-wire flight control system is cross-linked with the avionics system, power system, engine throttle actuator mechanism, landing gear control system, and antiicing and deicing system. During the test, the cross-linking signal can be provided by the corresponding simulator. The basic requirements for the external cross-linking system simulator are as follows: 1. The interface of the simulator shall be designed and configured according to the corresponding interface control document. 2. The simulator shall be able to simulate the cross-linking system to provide signals to the flight control system. 3. The simulator shall be able to respond to the flight control system signals that should be responded, so that the fly-by-wire flight control system can make judgment according to the response signals. 7. Test system and requirements The test system includes test equipment including test sensor, data acquisition and processing system, signal isolation amplifier, and modem. 1. Test sensor The test sensor shall meet the test requirements of the combined test of the system. Table 5.3 shows the requirements for the accuracy and measuring range of the test sensor for a combined test of the fly-by-wire flight control system. After the test sensor is installed on the test bed, its test accuracy and measuring range shall be calibrated to eliminate the installation error and confirm they meet the test requirements of the combined test of the system. Table 5.3: Requirements for accuracy and measuring range of test sensor for test of fly-by-wire flight control system. Items Angular displacement Tension force Torque Linear displacement sensor Acceleration Pressure sensor

Accuracy (%F.S.)

Measuring range

0.25 0.25 0.25 0.25 1 0.25

6 60 0 300 kgf 6 100 N m 6 100 mm 6 50 g 0 35 MPa G

294 Chapter 5 2. Data acquisition and processing system Data acquisition and processing system shall be able to record the signals of the test sensor, the working status signals of the fly-by-wire flight control system and the signals of the fly-by-wire flight control computer in real time. The sampling frequency of the data acquisition and processing system shall be two to three times that of the natural frequency of measured signals. It shall have enough storage capacity to record at least all the data collected during a flight takeoff and landing period for test analysis. The data acquisition and processing system shall have high antiinterference performance and a digital/analog converter for voltage output and current output. With strong software functions and a good human/machine interface, it shall also be able to provide functions including data sorting and the provision of required parameters and signal sources dedicated for a cross-linking system for the excitation test, as well as data analysis, processing, and playback. 3. Signal isolation amplifier The signal isolation amplifier completes the isolation and amplification of signals between the fly-by-wire flight control system, other systems of the aircraft and the test system. The signal isolation amplifier shall be able to avoid the mutual interference caused by the impedance mismatch between cross-linked systems and be able to amplify the weak current signals to some extent. The specific requirements are as follows: a. The number of channels shall be determined according to actual requirements of the test. b. The input and output shall be numerically displayed with a display accuracy not less than three places after the decimal point. c. The bandwidth of each channel shall not be less than 2 kHz. d. Each channel shall have a zero setting knob. e. Each channel shall be effectively isolated in input and output. 4. Modem It is used to modulate and demodulate equipment to complete the generation and output of carrier signals from modulation and demodulation, respectively. In the combined test of the fly-by-wire flight control system, it is necessary to conduct signal acquisition and analysis for airborne equipment of the fly-by-wire flight control system with some of its output signals as carrier signals. As such, the signals should be demodulated. When the carrier signal generated by some airborne equipment externally simulated needs to be input to other equipment, modulation is required. Therefore the modem equipment is indispensable. The requirements are as follows: a. The number of modulation and demodulation channels shall be determined according to specific test requirements, generally no less than four channels. b. The reference power supply for modulation and demodulation shall be 7 V 1800 Hz AC power.

Combined test of the flight control subsystem 295 c. The demodulated and output signal shall be a voltage signal with a voltage range of 10 to 110 V. d. The voltage range of the modulated and input signal shall be 10 to 110 V. 5. Loading system and requirements The loading system is also called the control plane aerodynamic hinge moment simulation system. It simulates the aerodynamic force on the control surface of the elevator, aileron, rudder, multifunctional spoiler, ground spoiler, and flaps (slats) during actual flight of the aircraft so as to test the static and dynamic performance of actuators under the action of aerodynamic force and the effects on the performance of fly-by-wire flight control system. As the control plane aerodynamic load changes with the flight state, such as flight speed and flight altitude, the loading system shall adopt electrohydraulic servo control. An electrohydraulic servo system with high dynamic response is commonly used. The output force of the actuator cylinder of the loading system shall be determined according to the maximum aerodynamic load on different control surfaces, the output stroke of the actuator cylinder shall be determined according to the stroke of actuators on each control surface and the output force of the loading system changes with the flight state. When the input signal of the loading system is 10% of the maximum output force, the bandwidth shall not be lower than 8 Hz. The loading system is a force system with strong position motion interference. In order to ensure the control accuracy of the force system, the redundant force of the loading system shall be less than 5% of the output force. When designing the loading system, safety protection measures under the state of power loss and voltage loss shall be considered and protection measures in the electrical, mechanical, and hydraulic aspects shall be provided. 6. Software development and debugging equipment (DIF) and requirements In the test process of the fly-by-wire flight control system, software development and debugging equipment (DIF) can be used to debug, monitor, and test the airborne software in the fly-by-wire flight control computer. The main functions of software development and debugging equipment shall include: 1. software loading and unloading. 2. providing man machine interface, input of debug command, register operation (modification), storage operation (read, write, transfer, compare, fill), break point operation, fault injection, I/O operation, program execution (continuous, single step, trace), and time test. 3. displaying the result of the execution of the command by the fly-by-wire flight control computer.

296 Chapter 5 4. displaying the command processing content and the real-time process of information change. 5. recording information and forming document. 7. Acceleration turntable and requirements The acceleration turntable is used to excite the overload sensor of the aircraft. The control command comes from the flight simulation system and the output signal of the sensor is provided to the fly-by-wire flight control computer, forming a closed-loop control of the fly-by-wire flight control system and flight motion. The performance of the acceleration turntable shall meet the following requirements: 1. Overload simulation range: 0.1 12 g. 2. Load capacity: The load capacity of a single servo turntable shall not be less than 5 kg. 3. Linear acceleration resolution: 0.02 g. 4. Linear acceleration accuracy: 1 3 10 4. 5. Bandwidth of servo turntable: It is not less than 8 Hz when the amplitude A 5 0.5 , amplitude error |ΔA/A| , 10%, phase shift |Δφ| , 10 . 8. Rate turntable and requirements The rate turntable is used to excite the rate gyro of aircraft, the control command comes from the flight simulation system and the rate gyro output signal is provided to the fly-by-wire flight control computer, forming a closed loop of the fly-by-wire flight control system and the aircraft motion, so that the static characteristics such as the transmission ratio of the rate feedback channel and the polarity can be inspected. In order to connect the lead-in/lead-out wires of rate gyro with nonrotating equipment, the turntable shall have a conductive slip ring with good performance. The performance of the rate turntable shall meet following requirements: 1. Load capacity: no less than 15 kg. 2. Rate range: 6 0.001 to 6 150 /s. 3. Accuracy: under the position mode, the positioning accuracy of the angular position is 6 3v; under the rate mode, when angular spacing is 10  , the accuracy is 2 3 10 24. 4. Bandwidth: not less than 10 Hz when the input signal A 5 1 , amplitude error |ΔA/A| , 10%, phase shift |Δφ| , 10 . 9. Angle of attack turntable and requirements The angle of attack turntable is used to excite the angle of attack sensor of the aircraft. The control command comes from the flight simulation system and the output signal of the angle of attack sensor is provided to the fly-by-wire flight control computer, forming a closed loop of the fly-by-wire flight control system and the aircraft motion, so that the static characteristics such as the transmission ratio of angle of attack feedback channel and the polarity can be inspected.

Combined test of the flight control subsystem 297 The performance of the angle of attack turntable shall meet the following requirements: Load capacity: no less than 10 kg. Angle range: 30 to 160 . Positioning accuracy of angle position: 6 3v. Frequency characteristic: when the amplitude is 1 , the dynamic response frequency is 6 Hz, 10% amplitude error is allowed, and 10 phase error is allowed. 10. Preparation for Test and Precautions Before the combined test of fly-by-wire flight control system, the installation, connection, and debugging of the system shall be completed, mainly covering the following items: 1. Completion of the conductive performance, insulation, and impedance inspection of cables. 2. Completion of the neutral adjustment of the actuator and the null adjustment of the control plane position sensor. 3. Completion of power supply (including primary power supply and secondary power supply). 4. Completion of system null inspection. 5. Completion of system interface inspection. 6. Completion of system self-test. 1. 2. 3. 4.

5.3.4 Test items and test methods The combined test of fly-by-wire flight control system mainly tests the following: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

interface inspection. actuator system test. BIT. redundancy management test. control logic check. modal conversion test. polarity and transmission ratio inspection. open-loop characteristic test. time-domain response test. stability test. closed-loop frequency response test. fault simulation and alarm display test. durability test.

5.3.4.1 Interface inspection The interface inspection of the fly-by-wire flight control system aims to verify whether the interface between the various airborne equipment of the system meets the requirements of

298 Chapter 5 the interface control document, whether the airborne equipment can communicate normally, and whether the system works normally. As interface inspection is the key to the smooth implementation of the combined test of flyby-wire flight control system, it shall be carried out carefully and thoroughly. If the system fails to pass the interface inspection, the causes shall be identified and the next test can be carried out after the faults are resolved. The interface inspection method is shown below: 1. Measure the input and output interfaces of each airborne equipment of the fly-by-wire flight control system according to the interface control document. 2. Connect all airborne equipment after passing the interface test. 3. Supply liquid and connect power. 4. Input control command to the system and check whether the system works normally. 5.3.4.2 Actuator system test The actuator system is the actuator mechanism of the fly-by-wire flight control system and its performance directly affects the system functions and performance. In the combined test of the system, the dynamic (static) state performance of the actuator system is verified to determine the static transmission ratio of the input command point of the actuator servo actuator control surface and confirm the polarity. The following feature and stability of the servo actuator system can be checked through the time-domain response and frequency response tests. For servo actuators with various control modes, such as fly-by-wire control and mechanical control modes, as well as working modes such as primary/backup, backup/primary, and primary/primary working modes, the correctness of the switching between control modes and working modes as well as the static and dynamic performance of the actuator working under different control modes and working modes shall be checked. The actuator system test includes static parameter inspection, transmission ratio measurement, time-domain response test, and frequency-domain response test. When the servo actuator system test is performed, the influence of rotational inertia and aerodynamic load on the control plane should be considered. The control plane rotational inertia can be replaced with a simulation element. The actuator system test shall be performed under unloaded and loaded conditions and the test results shall be compared to indirectly evaluate the dynamic stiffness of the servo actuator. In the flying process of the aircraft, as the control plane motion is affected by aerodynamic load, inertial load, friction force, and damping load, different load forms have different effects on the performance of the servo actuator system.

Combined test of the flight control subsystem 299 According to the working principle of the servo actuator, the load on the actuator will affect the change of the actuator motion speed, thus affecting the dynamic characteristics of the actuator system. Therefore the frequency response capability of the loading system shall be improved as much as possible and the effects of redundant force shall be minimized to achieve highly realistic aerodynamic load simulation. 1. Static parameter inspection The static parameter inspection mainly inspects the maximum output force, no-load output speed, output stroke, and static stiffness of the actuator. For actuators with mechanical control mode, the mechanical input dead zone and mechanical input minimum starting force shall also be inspected. 1. Measurement of maximum output force: keep the actuator at the neutral position, make the output end in series connection with the tension pressure sensor, fix the other end of the tension pressure sensor, input the maximum command in positive and negative directions to the actuator. After the piston rod of the actuator becomes stationary, measure the maximum output force in the stretching direction of the actuator. 2. Measurement of maximum output speed: keep the actuator at the neutral position, free the output end, input the maximum command in positive and negative directions to the actuator, measure the output displacement and time course curve of the actuator, and calculate the maximum output speed of the actuator. 3. Measurement of maximum output stroke: keep the actuator at the neutral position, free the output end, input the maximum command in positive and negative directions to the actuator, and measure the maximum output displacement of the actuator. 4. Measurement of static stiffness of actuator: connect the output end of the actuator with the loading actuator cylinder, apply pressure to the actuator, and input a fixed command to the actuator. The actuator shall stay at a middle position and use the loading actuator cylinder to apply load of different sizes to the actuator. Record the load applied by the loading actuator cylinder and the displacement of relevant actuator, and calculate the static stiffness of the actuator. 5. Measurement of mechanical input dead zone and minimum starting force: apply a control command not greater than 0.1 mm/s to the direction of the actuator at the mechanical input rocker arm of the actuator, record the displacement change curve of the mechanical input rocker arm, the input force change curve of the input rocker arm, and the motion curve of the actuator, and read the mechanical input dead zone value and minimum starting force from the curve. 2. Transmission ratio measurement Apply command input of certain amplitude at the command input point of the actuator and measure the actuator displacement and control surface deflection.

300 Chapter 5 Change the amplitude of input command gradually and the change of command amplitude shall consider two directions: positive direction and negative direction. To improve the measurement accuracy, it is suggested to increase the density of measurement points in the saturated area of the actuator. Draw the actuator input command point actuator output and actuator input command point control surface deflection transmission curves according to measured results, analyze the polarity, and calculate the transmission ratio from the curves. 3. Time-domain response test By applying a step command of different amplitudes to the actuator system, the time-domain response test can get a time course curve of the actuator output displacement and control surface deflection angle. Through calculation, time-domain response indexes such as rise time, overshoot, stabilization time, and oscillation times can be obtained. The results of the time-domain response test directly show the following features and stability of the actuator system. For the input of the step command of small amplitude, the time-domain response of the system includes nonlinear effects such as the dead zone and hysteresis, and the effects of the nonlinear effects to the performance of the actuator system can be obtained through the analysis of the response curve. For the input of step command of large amplitude, the effect of saturation phenomena on the performance of the actuator system can be studied. 4. Frequency-domain response test By applying a sinusoidal frequency sweeping signal, the frequency response test can obtain the frequency response of actuator input displacement and the control plane deflection angle output relative to the input command signal. Through the analysis of the frequency response test and calculation results, the bandwidth of the servo actuator system can be determined. When the frequency-domain response test is carried out, large, medium, and small amplitudes of the sinusoidal frequency sweeping signal shall be taken for the test. If the amplitude is too small, the frequency response characteristics cannot be judged due to nonlinear effects. If the amplitude is too large, frequency-domain characteristic analysis also cannot be carried out due to flow saturation (speed saturation). 5.3.4.3 BIT detection test The BIT of the fly-by-wire flight control system aims to identify hidden faults through the built-in test to ensure flight safety. It aims to verify whether the logical functions of the fly-bywire flight control system BIT are correct, whether the fault threshold setting is reasonable, and whether the fault detection rate, fault coverage rate, false alarm rate, and detection time meet the design requirements of the fly-by-wire flight control system. BIT can be carried out under two modes, that is, the mode when the fly-by-wire flight control system works normally and the mode under which the fault mode is set manually

Combined test of the flight control subsystem 301 according to the analysis results, so as to confirm if the BIT can accurately detect the fault and locate it. For the first mode, the correct result should be that the system passes the BIT without faults. If it fails to pass the BIT, the monitoring mechanism and threshold setting used shall be analyzed. For the second mode, it depends on fault setting which shall have the maximum coverage rate. After it passes the BIT, the state of the fly-by-wire flight control system shall be confirmed before testing other items. 5.3.4.4 Redundancy management test The redundancy management test aims to check the monitoring coverage ability of the flyby-wire flight control system to faults, including the rationality of monitoring threshold (amplitude threshold and delay threshold), the fault synthesis ability, and the correctness of the fault warning function. The redundancy management test is generally carried out under the open-loop mode of the fly-by-wire flight control system. With the connecting, breaking, grounding, and preset signals input ability provided by the tester of the fly-by-wire flight control system, faults are input manually to observe and record the fault display on the display device in the cockpit and software development equipment (DIF) is used to check the fault display of the fly-by-wire flight control computer. Parameters such as the type, quantity, and time sequence of input faults shall be given after analyzing the specific redundancy configuration and management scheme and alarm logic of the fly-by-wire flight control system. Generally, the following fault modes shall be considered: 1. Sensor faults, including sensor output, excitation input, and power supply faults. 2. Discrete quantity input/discrete quantity output fault. 3. Out-of-step, cross-transmission error, output command fault, channel fault, and logic fault of the fly-by-wire flight control computer. 4. Electrical and hydraulic mechanical faults of the actuator system. 5. Energy faults provided by aircraft to the fly-by-wire flight control system, such as a power fault and hydraulic source fault. 5.3.4.5 Control logic check The control logic check aims to verify if the realization of control logic and functions of the fly-by-wire flight control system is as expected. The control logic check covers control logic, control law function logic, and actuator work logic.

302 Chapter 5 1. Control logic open-loop check Control logic open-loop check generally includes system start logic, control command generation logic, trim logic, and operation panel use logic checks. 1. Start logic function check Start logic function check mainly checks the starting conditions and time of the fly-by-wire flight control system and if the start of fly-by-wire flight control system meets the design requirements under different power states, hydraulic source states, actuator states, computer states, and sensor (such as gyro and wheel load) states. 2. Control command generation logic check Control command generation logic check mainly checks if the control displacement of the multiredundancy steering wheel/column and pedal and the command sensor can generate effective control command logic under different working modes of the fly-by-wire flight control system and checks the combination of several faults of multiredundancy command sensor. 3. Trim logic check Trim logic check mainly checks the display correctness of trim command, trim function, and trim position under different working modes of the fly-by-wire flight control system. Generally, the signal transmission link from the trim command of the fly-by-wire flight control system to the control plane and the trim command generation logic are different under different working modes and the test items must cover the trim logic under various modes of the fly-by-wire flight control system. 4. Operation panel use logic check As part of pilot control, the operation panel of the fly-by-wire flight control system must have functions that can be simply, smoothly, and accurately realized. In the operation panel check test, the system function and logic of the control device, as well as the man machine efficiency of the panel shall be evaluated. For functions involving flying quality, such as the modal conversion function of the flyby-wire flight control system, the transient response of an aircraft in a typical flight state shall be evaluated when the fly-by-wire flight control system is under the closed-loop state. 2. Control logic closed-loop check The control logic closed-loop check mainly includes start logic, trim logic, control command generation logic, modal conversion logic, and control plane use logic checks. 1. Start logic check Start logic check includes the check of the start and start time of the fly-by-wire flight control system. The air start and ground start modes can be switched by setting wheel load signal. 2. Trim logic check The trim logic check includes the check of horizontal stabilizer automatic trim logic, manual electrical trim logic, manual mechanical trim logic, manual cutoff

Combined test of the flight control subsystem 303 logic, automatic cutoff logic and check of mutual priority, aileron trim check, and rudder trim check. The trim function logic can be verified by setting the conditions for accessing automatic trim logic and automatic cutoff logic through the flight control system tester. 3. Control command generation logic check The control command generation logic check includes the check of pilot (copilot) command displacement and directive force logic and the pilot (copilot) control priority logic of manually set faults. The control priority logic can be verified by setting command displacement and directive force faults on the fly-bywire flight control system tester. 4. Modal conversion logic check Modal conversion logic check includes the check of normal manual switch logic and system mode check in the case of signal fault, including the manual mode conversion on the main flight control panel to degraded system mode in the case of atmospheric data failure and inertial navigation data failure, the conversion of simulated backup mode in case of PFC failure, the conversion of mechanical mode and actuator primary/backup mode in the case of ACE failure. In the case of atmospheric and inertial navigation data failures, the mode conversion is realized through fault simulation of the flight control system tester. In the case of PFC and ACE failures, the mode conversion is realized through fault simulation of the power supply. 5. Control plane use logic check The control plane use logic check includes the check of control plane bias limit function, aileron lift increase function, ground automatic lift destruction and drag increase function, automatic air deceleration function, and ground and air manual deceleration function. The control plane use logic can be verified by setting control plane use logic through the flight control system tester. 3. Control law function logic check Many functions (more than 80%) of the fly-by-wire flight control system are realized in the control law, mainly including boundary protection, attitude holding, signal fault reconstruction, direct chain work logic, spoiler use logic, control plane bias limit logic, and aileron-assisted lift logic. The control law function logic check checks the function start and exit conditions and if the function of the fly-by-wire flight control system is normal after the function is started on one hand, and checks if the fly-by-wire flight control system meets the design requirements under the closed loop state after the function is started and if the transient response of the aircraft meets the design requirements when the function is started, exited, or loses efficiency on the other hand. 4. Actuator system work logic check Actuator system fault logic check mainly inspects the actuator system’s work logic from the requirements of fly-by-wire flight control system, rather than checks its function and

304 Chapter 5 performance only from the internal loop of the actuator system described in section 5.3.4.2. The actuator system performs to record and report the conversion logic of primary and backup backup and primary working mode, the fault switch logic of the primary and backup backup and primary working mode, the entry/exit logic of the primary and primary working mode, the conversion of the working mode and working way of the actuator when the fly-by-wire flight control system has mode conversion, the lock logic of the spoiler of the wing at the other side when the spoiler of a wing fails, the automatic cutoff logic of horizontal stabilizer, and the fault status of the actuator system. When it comes to the logic of the performance of the fly-by-wire flight control system, the transient response of the system when the logic is converted shall be evaluated. 5.3.4.6 Modal conversion test Modal conversion test includes the modal conversion function test and modal conversion to transient state test, aiming to verify whether the fly-by-wire flight control system can be automatically or manually converted from one working mode to another working mode, and whether the response transient state of the aircraft in the modal conversion process can meet the design requirements on flying quality. When the modal conversion test is performed, the modal conversion logic and conditions must be clarified and verified one by one, and the transient response of aircraft during the modal conversion must be measured at the same time. The modal conversion test checks the modal conversion function first. According to the modal conversion conditions, the test state is set and the modal conversion is completed automatically or manually to confirm whether the mode can be converted normally under the condition that the conversion condition is met and whether the mode will be converted incorrectly under the condition that the conversion condition is not met. After the modal conversion function is checked as correct, the transient response of the aircraft in the modal conversion process should be measured. In the modal conversion to transient state test, the aircraft motion equation must be connected to the fly-by-wire flight control system to form a closed loop. Based on the modal conversion conditions, choose the corresponding flight state, adjust the fly-by-wire flight control system to the modal conversion state, automatically or manually convert the mode, and test and report the aircraft’s attitude angle signal, overload signal, and angular rate signal in the modal conversion process at the same time, and judge if the transient response of the aircraft in modal conversion process meet the design requirements on flying quality through analysis of measured data. 5.3.4.7 Polarity and transmission ratio inspection Polarity and transmission ratio inspection aims to check if the polarity and transmission ratio of each channel of the fly-by-wire flight control system are consistent with the design.

Combined test of the flight control subsystem 305 For the polarity and transmission ratio inspection, it is unnecessary to select all flight states, but select typical and representative flight states for check. Polarity and transmission ratio inspection shall be performed for the fly-by-wire flight control system under different working modes. For the control link of the control device control plane and feedback sensor control plane, not only the overall polarity and transmission ratio of the control link shall be checked, but also the polarity and transmission ratio of each link shall be checked. If possible, the polarity and transmission ratio inspection should better directly excite the sensor so that the characteristics of the sensor can be included in the test. 5.3.4.8 Time-domain response test The time-domain characteristic of the fly-by-wire flight control system is actually the common characteristic of the fly-by-wire flight control system and a natural aircraft. In other words, after step input is applied at the control command sensor or pilot control point (input according to the elevator control channel, horizontal stabilizer control channel, aileron control channel, rudder control channel and spoiler control channel, and finally input according to the longitudinal control channel, lateral control channel, and heading control channel), the aircraft motion output and time course curve is tested and recorded and it is just the time response characteristic. The time-domain response test of the fly-by-wire flight control system aims to analyze the time-domain response characteristics of the fly-by-wire flight control system by measuring the time response curve of the fly-by-wire control aircraft under the step control input according to the aircraft time-domain response judging criteria. The time-domain response can visually characterize the characteristics of the fly-by-wire flight control system, which is consistent with what is observed and experienced during actual flight. The time-domain response test of the fly-by-wire flight control system, by controlling the control column and pedal and applying standard input to the system when the aircraft is in a closed loop, measures, records, and analyzes the change course of aircraft motion response with time. Typical standard input includes step, pulse, or bidirectional pulse and input amplitude includes small amplitude, medium amplitude, and large amplitude, which can directly analyze the performance of the system under the input of different amplitudes. Attention should be paid to the influence of clearance, dead zone, saturation, and other factors brought by the control input on the test results. The time-domain response test of the fly-by-wire flight control system shall cover all flight envelope and flight states and be performed under all working modes of the fly-by-wire flight control system. It is better to use the 6-DOF equation for the flight motion equation,

306 Chapter 5 or use the simplified small-perturbation linear equation if restrained by conditions, but the influence of the simplified aircraft motion equation on the test results must be considered. The verification of flight boundary limit functions of the fly-by-wire flight control system, such as angle of attack limit, overload limit and roll rate limit shall be carried out in the time-domain response test of the system and large amplitude shall be input for the verification. In this way, the limit functions can be analyzed and the performance of the system and aircraft dynamics under the condition of deep saturation can be examined. The time-domain response test of the fly-by-wire flight control system can also verify the transient conversion of different working modes of the system and the transient changes caused by the aircraft configuration transformation [such as landing gear control and flaps (slats) control] under different flight states. 5.3.4.9 Stability test Designing aircraft to be stable and maneuverable has always been the goal of aircraft designers. However, without stability, it is impossible to talk about maneuverability and the comprehensive design of aircraft aerodynamic layout, center of gravity design, and engine thrust design. The aircraft can be statically unstable but not dynamically unstable. Large aircraft are generally designed as statically stable. Is the statically stable aircraft definitely stable with the flight control system? The answer is no. Will the statically unstable aircraft be stable with the flight control system? The answer is yes. Therefore, it is particularly important to study and verify the stability of an aircraft with the fly-by-wire flight control system, which covers two aspects: is the aircraft stable? And is the stability margin enough? Those are the problems to be confirmed in the stability test. The stability test of the fly-by-wire flight control system aims to verify the stability of the aircraft with the fly-by-wire flight control system in a relatively real system test environment and test and analyze its stability margin. 1. Closed-loop stability test The basic principle of the closed-loop stability test is that standard excitation control such as step, pulse, bidirectional pulse, and sine are applied to the aircraft with the flyby-wire flight control system (fly-by-wire flight control system 1 natural aircraft), motion outputs of the aircraft such as attitude angle, angular speed, and overload are measured, and the stability of the aircraft is analyzed according to the principle of control theory. The aircraft is stable if its motion does not diverge or oscillates in constant amplitude. 1. Manually operate the control column (wheel) or pedal of the aircraft randomly and simulate the standard control above as far as possible to observe if the aircraft motion is stable from the real control plane and display and check if there is buffeting and limit cycle oscillation.

Combined test of the flight control subsystem 307 2. Make the aircraft fly in a certain stable flight state and apply different amplitude of wind, angle of attack, pitch angle, roll angle, sideslip angle through the flight simulation system to verify the stability of the fly-by-wire flight control system including to the interference. 3. Operate the aircraft control column (wheel) or pedal in the standard excitation way above through a mechanical position signal generator and observe if the aircraft motion is stable from the real control plane and display and check if there is buffeting and limit cycle oscillation. Step signals or sinusoidal signals are generally selected as standard signals. Due to the inaccuracy of aircraft aerodynamic data and system modeling and the difficulty of mathematical description of the flight environment, the aircraft designed to be stable may be unstable in an actual flight environment. Although this uncertainty is taken into account in the design of the fly-by-wire flight control system and the Monte Carlo method is used for the parameter uncertainty analysis, it is of special significance to use the interference test to find an unstable aircraft configuration in a real system test. It can determine the relative stability level and the absolute stability, and can study the effects of actuator deflection and rate limit. The closed-loop stability test shall cover the entire flight envelope and various working modes of the fly-by-wire flight control system and give emphasis to the flight states sensitive to stability. Great attention shall be paid to areas with low altitude high speed and high altitude small indicated airspeed. 2. Stability margin test The fly-by-wire flight control system is a system based on model design and the determination of control law parameters is based on the aircraft and the aircraft system model. Although these uncertainties are taken into account in the design, the unestimable internal parameters of the system and the changes of external factors will still pose a threat to the stability of the aircraft. Practical experience shows that as long as the fly-by-wire flight control system including the aircraft has the stability margin meeting the requirements of the relevant specifications, the stability of aircraft in actual flight can be ensured. In other words, an appropriate stability margin can ensure the aircraft flies stably when external factors and internal parameters change within a certain range. The stability margin test should cover the entire flight envelope of the aircraft and be carried out under all working modes of the fly-by-wire flight control system. When the stability margin test is carried out, the aircraft may adopt longitudinally and directionally separated linear small-interference motion equations. The aircraft flies steadily and horizontally at the selected state point and the control device is generally set at the neutral position. A typical stability margin test method is used in the closed fly-by-wire flight control system test with aircraft motion equation, and the test principle is shown in Fig. 5.5. Use a

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Figure 5.5 Principle of stability margin test of the fly-by-wire flight control system.

dynamic frequency response analyzer to add a sinusoidal frequency sweeping excitation signal from the command synthesis point (the addition of the internal control command and the external test command) and record the frequency response curve of the internal control command before the command synthesis relative to the input command (input command 5 internal control command 1 frequency sweeping command). The key point of the method is the input at the same point and the response under that input. Typical stability margin test method has following features: 1. As the closed-loop system is continuously opened when the stability margin test is carried out, the test will not affect the normal operation of the closed-loop system. When the closed-loop system is working, the sinusoidal frequency sweeping command is added to test the input and output and the stability margin of the closed-loop system. This method is suitable for the combined test of the flight control subsystems, “iron bird” integration test, onboard ground test, and flight test of the flight control system. 2. This test method requires designing interfaces for the system. In other words, the test command input interface shall be set at a command synthesis point and control command output interface shall be set in system. Besides, special test equipment (i.e., the open-loop

Combined test of the flight control subsystem 309 frequency sweeping phase inverter in Fig. 5.5) shall be designed and the system input command shall be calculated according to the command synthesis algorithm at the command synthesis point of the system. Under the open loop of the fly-by-wire flight control system, the stability margin test method generally has the following two ideas: 1. After inverting the internal control command of the system, superpose with the frequency sweeping command and then return to the system. Under this case, the system input command 5 (frequency sweeping command system internal control command) 1 system internal control command, and the system output 5 system internal control command. This method actually constructs the open-loop system. 2. Disconnect the closed-loop system in the flight simulation system and add a frequency sweeping command representing control plane deflection. At this time, the output of the open-loop system is the control plane deflection of the system. What should be noted is that the fly-by-wire flight control system generally adopts multifeedback control and when constructing a closed-loop system for the stability margin test, the break point must be able to completely disconnect the system control loop. In the stage of the combined test of the subsystem, various testing methods can be used for comparative verification. As the stability margin is a concept of a linear system and the tested system is a nonlinear system with many nonlinear factors such as dead zone and integrator, the measured stability margin is closely related to the sinusoidal excitation amplitude. Therefore the input amplitude shall be carefully selected to make the system as close as possible to a linear system, which can weaken the influence of nonlinear factors such as dead zone, and also make any link of the system not be in a saturated state. The frequency range for stability margin measurement shall be subject to ensuring that the point of intersection between the amplitude frequency curve and 0 dB line can be found and the phase frequency curve reaches 180 . The heading control system of the aircraft is a multiinput and multioutput system and the stability margin can be calculated based on the lateral and heading channels in the test. 5.3.4.10 Closed-loop frequency response test The closed-loop frequency response test aims to obtain the frequency characteristics of high-order fly-by-wire flight control system (with aircraft dynamic motion equation), thus obtaining the characteristic parameters of equivalent low-order short-period motion with equivalent matching method. The control law in the stage of the combined test of the fly-by-wire flight control system is not very perfect. The defects of the control law should be fully exposed through test and then analysis and test shall be carried out for the quality of the fly-by-wire flight control

310 Chapter 5 system through a closed-loop frequency response test, providing a test basis and improvement suggestions for further optimization of the control law. The actual fly-by-wire flight control system is a system with many nonlinear factors, so the closed-loop frequency response characteristics obtained under different amplitude inputs are different. For this reason, the input amplitude shall be carefully determined in the test. The input amplitude determination principle is the same as the amplitude selection principle in the stability margin test. The input frequency range of the closed-loop frequency response test is generally 0.1 10 rad/s and the frequency interval is divided by 10 parts of logarithm for every 10 frequency range. If the a mplitude curve does not attenuate after ω $ 10 rad/s in the test, the upper limit frequency can be extended by one time. The closed-loop frequency response test shall be conducted on all design points of the flight envelope and the aircraft motion can be described with a longitudinally and directionally separated small-interference motion equation and linear aerodynamic derivative. The principle diagram of the closed-loop frequency response test is shown in Fig. 5.6. 5.3.4.11 Fault simulation and alarm display test 1. Fault simulation test In the design process of the fly-by-wire flight control system, a fault tree is established and FHA (functional hazard analysis) is conducted according to the system composition, working principle, and its relationship with aircraft structure, power, and other aircraft systems to determine the hazards of fault modes of the system on the

Figure 5.6 Principle of closed-loop frequency response test of the fly-by-wire flight control system.

Combined test of the flight control subsystem 311 aircraft. According to the hazards of the fault modes on the aircraft, FAR25 Department divides the faults into five levels, which are defined as follows. Level I: the faults that may cause or lead to functional failure of the fly-by-wire flight control system and hence result in catastrophic aircraft accident. Level II: the faults that may cause or lead to system functional failure and hence result in a dangerous aircraft accident. Level III: software which failure may cause or lead to system functional failure and hence result in a failure state that has large effects on the aircraft. Level IV: software which failure may cause or lead to system functional failure and hence result in a slight failure state of the aircraft. Level V: software which failure may cause or lead to system functional failure and it will not affect the working performance of the aircraft or the workload of the pilot. Fault mode and hazard analysis is important work in aircraft design. Through analysis, the fly-by-wire flight control system of large aircraft has many fault modes and its single faults and fault combinations may reach several hundreds and level I and II faults account for about 20% of all similar faults of the aircraft. That is why people always pay close attention to the safety of the fly-by-wire flight control system. The fault simulation test of the fly-by-wire flight control system shall cover all fault modes that may be analyzed to confirm whether these fault modes are consistent with their hazard analysis, and verify the rationality of system redundancy and reconfiguration for fault avoidance and disposal measures at the same time. As the fly-by-wire flight control system has many fault modes which are not described here, this book describes the test methods of typical level I and II fault modes of the fly-by-wire flight control system below. These fault modes mainly include: 1. Power fault. 2. Hydraulic source fault. 3. Control plane clamping stagnation (including at neutral position and limit position). 4. Control plane uncommanded motion. 5. Control plane failure (control free). 6. Trim function fault (uncommanded motion and clamping stagnation). 7. Unilateral control stagnation (including at neutral position and limit position). 8. Airborne equipment fault (including PFC, ACE, command sensor, feedback sensor, and actuator). The difficulty of the fault mode test is the fault injection method. In other words, how to inject faults into a complete system that works normally without affecting its faults. The other difficulty is the assessment on the impact of the fault on the aircraft. 1. Power fault Simulate single faults and fault combinations of all bus bars, single or redundancy power supply faults in the redundancy power supply of PFC and ACE, check the working state of the fly-by-wire flight control system under the fault

312 Chapter 5 modes above, and also check the working state of the system under the emergency power supply. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the fly-by-wire flight control system, as well as the correctness of fault alarm and display. 2. Hydraulic source fault Simulate the pressure loss and low pressure (the pressure before and after the low pressure warning of the hydraulic system) faults of the hydraulic system, check the working mode of the fly-by-wire flight control system under the fault modes above, including the single faults and fault combinations of the hydraulic system, and also check the working state of the fly-by-wire flight control system under the emergency status of the hydraulic system. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the flyby-wire flight control system, as well as the correctness of fault alarm and display. 3. Control plane clamping stagnation The control plane clamping stagnation faults may include single clamping stagnation faults or several clamping stagnation fault combinations involving the aileron, elevator, rudder, ground spoiler, multifunctional spoiler, and horizontal stabilizer. Check the working state of the fly-by-wire flight control system under the fault modes above. The simulation test of the control plane clamping stagnation fault shall be carried out under the closed-loop state of the fly-by-wire flight control system and the fault mode shall be realized by setting the control plane deflection in the flight simulation system. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the fly-bywire flight control system, as well as the correctness of the fault alarm and display. 4. Control plane uncommanded motion The control plane uncommanded motion faults may include single uncommanded motion faults or several uncommanded motion fault combinations involving the aileron, elevator, rudder, ground spoiler, multifunctional spoiler, and horizontal stabilizer. Check the working state of the fly-by-wire flight control system under the fault modes above. The simulation test of the control plane uncommanded motion fault shall be carried out under the closed-loop state of the fly-by-wire flight control system and the fault mode shall be realized by setting the control plane deflection in the flight simulation system. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the fly-by-wire flight control system, as well as the correctness of the fault alarm and display.

Combined test of the flight control subsystem 313 5. Control plane failure (control free) The control plane failure (control free) may include single failures (control free) or several failure combinations (control free) involving the aileron, elevator, rudder, ground spoiler, multifunctional spoiler, and horizontal stabilizer. The simulation test of the control plane failure (control free) shall be carried out under the closed-loop state of the fly-by-wire flight control system and the fault mode shall be realized by setting the control plane deflection in the flight simulation system. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the fly-by-wire flight control system, as well as the correctness of the fault alarm and display. 6. Trim function fault The trim function fault may include trim channel clamping stagnation or uncommanded motion of the aileron, elevator and horizontal stabilizer. The simulation test of the trim function fault shall be carried out under the closed-loop state of the fly-by-wire flight control system and the fault mode shall be realized by setting the trim command (or superposing control command) in the flight simulation system. Confirm the consistency of the fault mode and its hazards with the FHA analysis, the validity of the redundancy design, fault reconstruction, fault conversion logic, and fault tolerance measures of the fly-by-wire flight control system, as well as the correctness of the fault alarm and display. 7. Unilateral control stagnation The unilateral control stagnation fault may include the stagnation faults of the control column, steering wheel and pedal of the left and right pilots. The aiming/override functions of the cockpit control system under unilateral control stagnation shall be verified to confirm the consistency of the fault mode and its hazards with the FHA analysis. 8. Airborne equipment fault The airborne equipment fault may include faults of PFC, ACE, command sensor, feedback sensor, and actuator. The fault isolation and handing ability of the fly-bywire flight control system and its consistency with the FHA analysis shall be verified. The simulation test of the airborne equipment fault shall be carried out under the closed-loop state of the fly-by-wire flight control system and the fault mode can be realized through power cutoff and redundancy signal out of tolerance. 2. Alarm display test The alarm display function of the fly-by-wire flight control system mainly includes unified voice alarm and stick shaker, as well as CAS information and display of flight control diagram. The information displayed on the diagram mainly includes control plane state, control plane deflection, PFC/ACE state, system mode and stage of hydraulic system. The fault alarm function test mainly checks whether the logic of the alarm functions above designed for the fly-by-wire flight control system are correct, whether the alarm

314 Chapter 5 effect is obvious, for example, whether the voice form conforms to the habits of pilots, whether the man machine efficiency is good, and whether the sound frequency and size are appropriate, as well as whether the amplitude and frequency of the stick shaker are appropriate and whether the pilot can accept that. The fault display function test mainly checks whether the CAS information and display of flight control diagram are correct and whether the modal conversion display is appropriate. When the alarm display test of the fly-by-wire flight control system is conducted, set the alarm trigger conditions and system state on the tester of the fly-by-wire flight control system to visually observe the correctness of the alarm logic and whether the display information and voice meet the design requirements. After the system exits the alarm, check whether the alarm display and voice prompt meet the requirements. When the alarm display test of the fly-by-wire flight control system is conducted, set the working state and actuator fault on the tester of the fly-by-wire flight control system to verify the control plane state display under corresponding state. With regard to control plane deflection check, control devices such as control column, steering wheel, pedal, brake control handle, and pitch trim switch can be operated to verify the indication of control plane deflection polarity and stroke after the operation. By setting faults of PFC/ ACE, the display of the PFC/ACE status can be verified. The display correctness of system mode can be checked by manually switching and setting faults to make the system degrade. The status of the hydraulic system can be verified by setting the working state (low pressure or normal pressure) of the hydraulic system. The alarm display test shall be conducted in combination with the fault simulation test to evaluate whether the alarm display design of the system is reasonable and whether it could clearly and accurately reflect the state change of the system. 5.3.4.12 Durability test As one of the major combined tests of the fly-by-wire flight control system, the durability test is used to assess the long-term working reliability of the fly-by-wire flight control system under the closed-loop state. There are no strict regulations on the test method and it is suggested to adopt the man machine combined simulation method. The durability test of the fly-by-wire flight control system is carried out under laboratory temperature and finishes taxiing, takeoff, climb, cruise, approach, and landing motions in the full flight envelope. It can also use the mechanical displacement signal generator to excite the control column, steering wheel, and pedal for simulation control to assess the durability of single control channel. At the stages of takeoff, cruise, and landing, adjust the power voltage (power fault). That is to say, increase and decrease the primary power supply voltage of the system and hold it at each for 30 s within the power voltage range required for normal working of the closed-loop system.

Combined test of the flight control subsystem 315

Figure 5.7 Profile of durability test.

At the stages of takeoff, cruise, and landing, set the hydraulic source low pressure fault. In other words, lower the pressure of the hydraulic system and hold it for 30 s within the hydraulic pressure range required for normal working of the closed-loop system. In the durability test of the fly-by-wire flight control system, the time of each test flight profile can be determined according to the requirements of the designer. The test flight profile shall cover different weights and configurations of the aircraft and shall be a typical mission profile used for aircraft. Fig. 5.7 shows a typical test flight profile for the durability test of the fly-by-wire flight control system of the background large aircraft. The weight of the aircraft is large, medium, and small weight, the center of gravity is the normal center of gravity, the cruise altitude of the aircraft is 9000 11,000 m, and the flight time of the aircraft in the cruise stage is at least 10 min. After the completion of a profile test, a PBIT test of the system shall be conducted to confirm the system is in normal state. During the flight stage of the durability test, 0.1 1 Hz frequency sweeping signals shall be added to each control channel at a certain interval (20 30 s) and the amplitude shall be 10% maximum stroke. Manual free control is also allowed during the flight.

5.3.5 Test results and judging criteria 5.3.5.1 Results of the open-loop test and judgment In the open-loop test of the fly-by-wire flight control system, the test results can be evaluated according to certain error requirements by comparing the test results with simulation results. The input signal of the open-loop test is imported to the simulation model to calculate the output of the simulation model. The test results are compared with the simulation results and their curves shall be within the required error range. Fig. 5.8A shows the test curve of a test of the elevator control channel, Fig. 5.8B shows the simulation output curve relative to the test input, and Fig. 5.8C shows the comparison result of the two curves after they are put in the same coordinate system.

316 Chapter 5

Figure 5.8 Comparison of open-loop test curve and simulation curve of elevator control channel.

Combined test of the flight control subsystem 317 5.3.5.2 Results of the time-domain and closed-loop characteristic test and judgment The evaluation of the test results of the time-domain and closed-loop characteristic test are conducted in two steps. First step: have comparative analysis of the test results and simulation results. If they are basically consistent, start the second step. Otherwise, check the test status and test environment and analyze the reasons for the differences. If necessary, conduct the test again. Second step: conduct flying quality analysis and evaluation for the test results. The results of the closed-loop characteristic test should have equivalent matching and then the matching results shall be analyzed and evaluated. 5.3.5.3 Results of the stability margin test and judgment The evaluation of the test results of the stability margin test are conducted in two steps. First step: have comparative analysis of the test results and simulation results. If they are basically consistent, start the second step. Otherwise, check the test status and test environment and analyze the reasons for the differences. If necessary, conduct the test again. Second step: read the amplitude margin and phase margin in the test results and check whether they meet the requirements. In other words, the amplitude margin shall be greater than 6 dB and the phase margin shall be greater than 60 . 5.3.5.4 Results of the durability test and judgment In the durability test, the mean time between failures of the fly-by-wire flight control system (airborne equipment) and the replacement condition of the airborne equipment shall be recorded in detail. It will be judged as a fault in case of any of the following conditions: 1. Failure to pass PBIT. 2. Fault message appears. 3. The airborne equipment, structural parts, or components are fractured, broken, or damaged and they affect the function, performance, or structural integrity of the fly-bywire flight control system. In the durability test, if a fault appears and the fault is confirmed as an associated fault of the airborne equipment, the airborne equipment can be replaced and its normal working hours shall be recorded, and then the test can be continued. Then, cooperate with the manufacturer to analyze the fault of the airborne equipment. The durability test can be conducted for the airborne equipment independently again after improvement measures are taken.

318 Chapter 5 If the flight control system has software upgrading in the durability test and the software version does not affect system functions, the test can be continued. If the fly-by-wire flight control system has great changes in state, the system durability test shall be conducted again. For the test results and the problems or faults found in the test, the dynamic (static) performance analysis and reliability analysis shall be carried out to give a durability conclusion and put forward an implementation plan for the continuous optimization of airborne equipment. The durability test of the fly-by-wire flight control system shall meet the design requirements of the fly-by-wire flight control system in the time course.

5.4 Combined test of the high-lift system 5.4.1 System introduction The high-lift system of large aircraft is used for the lift augmentation control in the takeoff, climb, approach, and landing stages of the aircraft, so as to improve the low-speed performance of the aircraft, reduce the landing speed, and shorten the distance of run. The high-lift subsystem is generally divided into two types: central driving type and distributed driving type. This book takes the most commonly used central driving high-lift system of large aircraft as an example to introduce the combined test. The test methods for distributed driving high-lift system are similar. The high-lift system is generally composed of flaps and slats control handle, flaps and slats command sensor, flaps and slats override module, flaps and slats controller (FSECU), flaps and slats power drive unit (PDU), flaps and slats mechanical drive device, flaps and slats actuator, flaps and slats position sensor, flaps tilt sensor, slats tilt sensor, and flaps and slats WTB unit. The energy, controller, and drive device of the high-lift system are all configured based on dual-redundancy and the software is developed in an engineering way according to class B requirements to meet the safety requirements of the aircraft for the high-lift system. The architecture of the high-lift system of the background aircraft is shown in Fig. 5.9. The flaps and slats control handle has six gears, as shown in Table 5.4. Each gear of the handle is designed with a slot to prevent the pilot from misoperation, and “1” and “4” are the gears that will be passed no matter which gear is to be selected.

5.4.2 Test objective The combined test of the high-lift system aims to verify and validate that the function, performance, and safety of the high-lift system meet design requirements on a simulation test bed in the most realistic test environment.

Combined test of the flight control subsystem 319

Figure 5.9 Architecture of high-lift system.

The test is carried out to: 1. verify the interface, function, and performance of control channel. 2. verify the functional transformation logic. 3. verify the self-test function and the accuracy of fault synthesis, fault reporting, and information recording. 4. verify the redundancy strategy. 5. verify the function and dynamic (static) performance of the airborne equipment composing of the system. 6. verify the fault logic and alarm display information.

320 Chapter 5 Table 5.4: Corresponding relationship between gears of flaps and slats control handle and control plane deflection angle. Gear of control handle 0 1 2 3 4 5

Slats deflection

Flaps deflection

Corresponding configuration

0 18 18 18 26 26

0 0 15 27 27 41

Cruise Initial arrival Plateau takeoff Takeoff Go-around and arrival Landing

5.4.3 Test requirements 5.4.3.1 Requirements for the tested object All airborne equipment of the high-lift system are “S” type parts and have passed the acceptance test before delivery and have conformity certificate or relevant record. The combined test of the high-lift system shall be carried out on a special test bed and all test pieces shall be arranged and installed to be the same or as similar as possible with those on actual aircraft. The connecting cables between them shall be connected according to the relevant interface control document and their electrical characteristics shall be consistent with the airborne cables as far as possible, while the length of cables can be adjusted according to the specific test environment. Since the high-lift system is symmetric left and right around the structure center plane of the aircraft, it is suggested to build a half test bed and install a half system only and load simulation can be adopted at the other side. The test bed shall pass the acceptance inspection and metrological examination and be provided with a certificate. The test cables shall be made according to the requirements of the relevant interface control documents. All performance indexes of the test equipment and the accuracy of the test sensor and instrument shall meet test requirements and be within the validity period. 5.4.3.2 Environmental requirements for the combined test 1. Requirements for power supply The power supply in the test environment shall meet the following requirements: 1. The type, voltage, power, and quality of the power supply in the test environment shall meet the power requirements of airborne equipment and other equipment in the test. 2. Effective grounding measures shall be taken for the test environment to meet system grounding requirements. 3. Effective antistatic measures shall be taken for the test environment to meet the antistatic requirements of the system.

Combined test of the flight control subsystem 321 2. Requirements for special test bed The special test bed shall have the following functions: 1. It shall be able to simulate the wing front beam, rear beam, and middle fuselage structure and provide a support structure for the installation of flaps and slats PDU, flaps and slats mechanical drive device, flaps and slats actuator, flaps and slats position sensor, flats and slats tilt sensor, flaps and slats WTB unit, etc. The support stiffness shall not be lower than the installation stiffness on the aircraft. 2. It shall be able to simulate the deformation of the wing along the altitude direction and assess the adaptability of the flaps and slats mechanical drive device to wing deformation (bending and twisting). 3. It shall be able to apply a simulated aerodynamic load at the control points of the flaps and slats actuator. 5.4.3.3 Requirements for the tester The tester of the high-lift system shall meet the following requirements: 1. It can be connected with the system airborne equipment, and an aviation connector matching with the system airborne equipment shall be used as the electrical connector. 2. It can power up flaps and slats controller, flaps and slats PDU, flaps and slats position sensor, flaps and slats tilt sensor, flaps and slats WTB unit, etc. for independent operation. 3. It has signal excitation function to provide all external cross-linking signals for the system. 4. It has a signal detection function to detect the internal signals of the system. 5. It has a signal monitoring function to monitor the bus data and discrete quantity of the system. 6. It can set cross-linking system faults. 7. It can modify the cross-link bus data manually and set the cross-link signal/data fault. 5.4.3.4 Preparation for the test and precautions The following debugging work shall be completed before the combined test of the high-lift system: 1. 2. 3. 4. 5. 6. 7.

Cable conduction, insulation, and impedance inspection. Power supply inspection (including primary power supply and secondary power supply). Neutral position check. Interface inspection. BIT test. Control logic and brake logic inspection. Polarity and transmission ratio inspection.

322 Chapter 5 5.4.3.5 Test requirements All instruments used in the high-lift system test shall be strictly calibrated and within the validity period. The requirements for the accuracy of the test equipment are as follows: 1. The measuring range of the sensor shall be suitable for the measured equipment and the accuracy of the sensor is 0.1% F.S. 0.3% F.S. 2. The resolution of time measuring equipment shall be # 0.1 s. 3. The measuring range of the torque measuring equipment shall be suitable for the measured equipment, with a resolution # 1 N•m. 4. The measuring range of the linear displacement measuring equipment of the ball screw nut shall be suitable for the maximum stroke of the ball screw, with a resolution # 0.1 mm. The test point layout for the combined test of the high-lift system of the background aircraft is shown in Fig. 5.10, mainly including the torque measuring point, speed measuring point, and the electrical signal measuring point.

5.4.4 Test items and methods The combined test of the high-lift system mainly tests: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

polarity inspection. control logic check. brake logic check. control test under normal mode. control test under degraded mode. Backup mode retraction test. flaps (slats) tilt function test. flaps (slats) asymmetry test. unconventional control test of flaps and slats control handle. fault simulation test. durability test.

5.4.4.1 Polarity inspection The polarity inspection of the high-lift system is mainly conducted to inspect the polarity of flaps and slats position sensor, the response polarity of the system in handle control and the response polarity of the system in override control. The methods and steps of the polarity inspection are as follows: 1. Connect power to the system normally. 2. Operate the flaps and slats control handle from gear “0-5,” and all moving airborne equipment of the system shall move normally.

Combined test of the flight control subsystem 323

Figure 5.10 Test point layout for combined test of high-lift system.

3. Check if the motion direction of all moving airborne equipment of the system is consistent with the flaps and slats “release” direction. 4. In the case of inconsistent direction, check and adjust the polarity of modules of the system. 5. Repeat steps (1) (3) until the motion direction is consistent with command direction. 6. Stop after the system runs to position after the directions become consistent. 7. Operate the flaps and slats control handle from gear “5-0,” and all moving airborne equipment of the system shall move normally. 8. Check if the motion direction of all moving airborne equipment of the system is consistent with the flaps (slats) “withdraw” direction. 9. In case of inconsistent direction, check and adjust the polarity of modules of the system.

324 Chapter 5 10. Repeat steps (6) (8) until the motion direction is consistent with the command direction; 11. Stop and wait for about 1 min after the system runs to position. 12. Press the arm button on flaps and slats override control panel and switch to standby mode. 13. Switch the rotary switch to the “release” position. 14. Check if the motion direction of all moving airborne equipment of the system is consistent with the flaps and slats “release” direction; 15. In the case of inconsistent direction, release the rotary switch and adjust the brake system. 16. Repeat steps (13) (14) until the motion direction is consistent with the command direction. 17. Release the rotary switch and wait for about 1 min. 18. Switch the rotary switch to the “withdraw” position. 19. Check if the motion direction of all moving airborne equipment of the system is consistent with the flaps and slats “withdraw” direction. 20. In the case of inconsistent direction, release the rotary switch and check and adjust modules of the system. 21. Repeat steps (18) (19) until the motion direction is consistent with command direction. 22. The system restores to the initial zero position. In the test, measure and record the output voltage of the flaps and slats position sensor and the deflection angle of the flaps (slats) with the test system and visually observe the motion direction of the moving airborne equipment of the system. 5.4.4.2 Control logic check The control logic check of the high-lift system is mainly conducted to check whether the conversion between normal mode and backup mode is normal by pressing the arm button on the flaps and slats override control panel in the flaps and slats motion process. The methods and steps of the control logic check are as follows: 1. Connect power to the system normally. 2. Operate the flaps and slats control handle from gear “0-5,” and all moving airborne equipment of the system shall move normally. 3. Press the arm button on flaps and slats override control panel at random in the normal motion process of the system. 4. Switch the rotary switch on the flaps and slats override control panel to “withdraw” position. 5. Release the rotary switch. 6. Repeat steps (2) (4) twice.

Combined test of the flight control subsystem 325 7. Make the arm button on flaps and slats override control panel pop up and connect power to the system again. 8. Operate the flaps and slats control handle from gear “5-0,” and all moving airborne equipment of the system shall move normally. 9. Press the arm button on flaps and slats override control panel at random in the normal motion process of the system. 10. Switch the rotary switch on flaps and slats override control panel to the “release” position. 11. Release the rotary switch. 12. Repeat steps (9) (10) twice. In the test, measure and record the system working mode and flaps (slats) deflection angle and observe the motion direction of the moving airborne equipment of the system. 5.4.4.3 Brake logic check The brake logic check of the high-lift system is conducted to check whether the brake logic of the flaps and slats drive device and flaps and slats WTB unit is normal under various working conditions. The methods and steps of the brake logic check are as follows: 1. Connect power to the system. 2. Operate the flaps and slats control handle from gear “0-5,” and all moving airborne equipment of the system shall move normally. 3. Check if the brake logic of flaps and slats drive device and flaps and slats WTB unit meets design requirements after the flaps (slats) deflect to position. 4. Operate the flaps and slats control handle from gear “5-0” to control the motion of the system. 5. Input emergency brake command at random and send it to FSECU. 6. Check if the brake of flaps and slats drive device and logic of antiwithdrawing brake device meet design requirements of the system when the system has emergency brake. 7. Cancel the emergency brake command and the system restores initial zero position. 8. Press the arm button on flaps and slats override control panel. 9. Switch the rotary switch on flaps and slats override control panel to the “release” position. 10. Release the rotary switch. 11. Check if the brake of flaps and slats drive device and logic of antiwithdrawing brake device meet design requirements. 12. Switch the rotary switch on flaps and slats override control panel to the “withdraw” position. 13. Release the rotary switch after the system reaches the mechanical limit position and stops.

326 Chapter 5 14. Check the brake of flaps and slats drive device and logic of antiwithdrawing brake device when the system reaches the limit position and brakes under override state. In the test, measure and record the brake time of flaps and slats drive device and antiwithdrawing brake device. 5.4.4.4 Control test under normal mode The flaps (slats) control test under normal mode is divided into no-load test and loaded test. The no-load test can check the control time and control sequence of flaps and slats. In the loaded test, based on the check on control time and control sequence of flaps and slats, the holding function of flaps (slats) mechanical drive device, the synchronicity of the motion of the ball screw nut, the synchronicity of the motion of the gear-rack mechanism, as well as the torque distribution and power distribution on the flaps (slats) mechanical drive system can be checked when the flaps (slats) bear aerodynamic load within the flight envelope. The steps of the flaps (slats) control test under normal mode of the high-lift system are as follows: 1. Follow following steps and methods for no-load test: 1. Connect power to the system normally. 2. Operate the flaps and slats control handle from gear “0” to “3.” 3. The system corresponds to the command of the handle, flaps and slats are “released” in sequence, and wait for about 2 min after the system stops. 4. Operate the flaps and slats control handle from gear “3” to “0.” 5. The system corresponds to the command of the handle, flaps and slats are “withdrawn” in sequence, and wait for about 2 min after the system stops. 6. Operate the flaps and slats control handle from gear “0” to “1.” 7. The system corresponds to the command of the handle, only the slats are “released,” and wait for about 2 min after the system stops. 8. Operate the flaps and slats control handle from gear “1” to “0.” 9. The system corresponds to the command of the handle, only the slats are “withdrawn,” and wait for about 2 min after the system stops. 10. Operate the flaps and slats control handle from gear “0” to “2.” 11. The system corresponds to the command of the handle, flaps and slats are “released” in sequence, and wait for about 2 min after the system stops. 12. Operate the flaps and slats control handle from gear “2” to “0.” 13. The system corresponds to the command of the handle, flaps and slats are “withdrawn” in sequence, and wait for about 2 min after the system stops. 14. Operate the flaps and slats control handle from gear “0” to “5.” 15. The system corresponds to the command of the handle, flaps and slats are “released” in sequence, and wait for about 2 min after the system stops.

Combined test of the flight control subsystem 327 16. Operate the flaps and slats control handle from gear “5” to “0.” 17. The system corresponds to the command of the handle, flaps and slats are “withdrawn” in sequence, and wait for about 2 min after the system stops. 18. Operate the flaps and slats control handle from gear “0” to “4.” 19. The system corresponds to the command of the handle, flaps and slats are “released” in sequence, and wait for about 2 min after the system stops. 20. Operate the flaps and slats control handle from gear “4” to “0.” 21. The system corresponds to the command of the handle and flaps and slats are “withdrawn” in sequence. 2. Follow following steps and methods for loaded test: 1. Apply the minimum working load of flaps (slats) within the flight envelope of aircraft to the flaps and slats driving system and repeat the test steps of no-load test in sequence. 2. Apply the maximum working load of flaps (slats) within the flight envelope of aircraft to the flaps and slats driving system and repeat the test steps of no-load test in sequence. 3. Apply the minimum holding load of flaps (slats) within the flight envelope of aircraft to the flaps and slats driving system and repeat the test steps of no-load test in sequence. 4. Apply the maximum working load of flaps (slats) within the flight envelope of aircraft to the flaps and slats driving system and slats line and repeat the test steps of no-load test in sequence. 5. Apply the minimum working load of flaps (slats) at different angles within the flight envelope of aircraft to the flaps and slats driving system and repeat the test steps of no-load test in sequence. 6. Apply the maximum working load of flaps (slats) at different angles within the flight envelope of aircraft to the flaps and slats driving system and repeat the test steps of no-load test in sequence. In the control test under normal mode of the high-lift system, the distribution of test points for the flaps (slats) driving system is shown in Figs. 5.11 and 5.12. In the test, following data shall be recorded: 1. Flaps (slats) position and accuracy. 2. The angular displacement and speed of flaps (slats) PDU when it actuates the output shaft of airborne equipment. 3. The linear displacement of ball screw nut at test point 12 15 in Fig. 5.11. 4. The torque at test point 1 11 in Fig. 5.11. 5. Loading state and loading process. 6. The angular displacement at test point 11 18 in Fig. 5.12. 7. The torque at test point 1 10 in Fig. 5.12. 8. Flaps (slats) control plane motion condition.

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Figure 5.11 Distribution of test points for flaps driving system.

Figure 5.12 Distribution of test points for slats driving system.

5.4.4.5 Control test under degraded mode Under the degraded fault mode, the flaps and slats drive device has a single motor fault and the output speed is halved. Thus flaps (slats) control plane control time is doubled. However, the control logic is the same as that under normal fault mode. Under this working mode, the flaps (slats) control time shall be checked when the flaps and slats PDU has single motor fault. The steps and recorded data of the control test under degraded mode are the same as that of control test under normal mode described in section 5.4.4.4. 5.4.4.6 Backup mode retraction test The backup mode retraction test mainly checks the motion of flaps (slats) control plane when the system is in an override working mode and there is no load. The steps and methods of the backup mode retraction test are as follows: 1. Connect power to the system normally. 2. Press the “arm” button on flaps and slats override control panel. 3. Switch the rotary switch on flaps and slats override control panel to the “release” position.

Combined test of the flight control subsystem 329 4. Switch the rotary switch on flaps and slats override control panel to the “off” position about 15 s later. 5. Switch the rotary switch on flaps and slats override control panel to the “off” position after the flaps and slats actuator moves to the mechanical limit position and stops. In the test, measure and record the motion condition of flaps (slats) control plane with the test system. 5.4.4.7 Slats tilt test The slats tilt test is carried out under no load. In the slats release process, the slats tilt fault state is simulated and the output of slats tilt sensor and the monitoring and cutoff function of the flaps and slats controller are tested when it is detected that the slats tilt exceeds the threshold. The methods and steps of the slats tilt test are as follows: 1. Connect power to the system normally. 2. Set slats tilt fault through slats tilt sensor about 5 s after the slats start to move. 3. After slats stop moving, disconnect system power supply and adjust the slats tilt sensor to restore it to normal working state. 4. Connect power to the system again and operate the flaps and slats control handle from gear “0” to “4” manually. 5. Set slats tilt fault through slats tilt sensor again about 2 s after the slats start moving for the second time. 6. After slats stop moving, disconnect power supply to the system and adjust the slats tilt sensor to restore it to normal working state. 7. Operate the flaps and slats control handle to gear “0” manually. In the test, measure and record the motion condition of flaps (slats) control plane and measure the output of slats tilt monitoring device with the test system. 5.4.4.8 Flaps tilt test Flaps tilt test is carried out under no load. In the flaps release process, the flaps tilt fault state is simulated and the output of flaps tilt sensor and the monitoring and cutoff function of flaps and slats controller are tested when it is detected that the flaps tilt exceeds the threshold. The methods and steps of flaps tilt test as well as recorded data are the same as that of the slats tilt test described in section 5.4.4.7. 5.4.4.9 Flaps (slats) asymmetry test The monitoring and cutoff functions of the flaps and slats controller are tested when it is detected that the flaps (slats) asymmetry exceeds the threshold. In the flaps

330 Chapter 5 and slats control process, the flaps and slats asymmetry out of threshold is simulated. In the test, measure and record the motion condition of flaps (slats) control plane and the working state of the system with the test system. 5.4.4.10 Unconventional control test of flaps and slats control handle The unconventional control test of the flaps and slats control handle is carried out under no load and it aims to check the system command and confirm its effectiveness. Due to differences in aerodynamic characteristics and overall design requirements of large aircraft and the design of flaps (slats) control plane control gears and logic, the control handle gears and logic design will also be different. The unconventional operation of the flaps and slats control handle of the background aircraft mainly has the following five types of unconventional control modes, and their test methods and steps are shown below: 1. Mode 1 1. Connect power to the system, operate the control handle from gear “0” to “3” manually, and operate the handle quickly to gear “1” when the slats correspond to the command of the handle to release. 2. The slats can be set to the angle indicated by gear “1” and the flaps do not respond to the command. 3. After slats stop moving, operate the control handle from gear “1” to “0” manually. 4. Slats withdraw normally. 2. Mode 2 1. Connect power to the system, operate the control handle from gear “0” to “3” manually, and operate the handle quickly to gear “2” when the slats correspond to the command of the handle to release. 2. The flaps and slats move to the angle indicated by gear “2” only. 3. After flaps stop moving, operate the control handle from gear “2” to “0” manually. 4. Flaps and slats withdraw normally. 3. Mode 3 1. Connect power to the system, operate the control handle from gear “0” to “5” manually, and operate the handle quickly to gear “4” when the slats correspond to the command of the handle to release. 2. The flaps and slats move to the angle indicated by gear “4” only. 3. After flaps stop moving, operate the control handle from gear “4” to “0” manually. 4. Flaps and slats withdraw normally.

Combined test of the flight control subsystem 331 4. Mode 4 1. Connect power to the system, operate the control handle from gear “0” to “5” manually, and operate the handle quickly to gear “3” when the slats correspond to the command of the handle to release. 2. The flaps and slats move to the angle indicated by gear “3” only. 3. After flaps stop moving, operate the control handle from gear “3” to “0” manually. 4. Flaps and slats withdraw normally. 5. Mode 5 1. Press the arm switch on override control panel before connecting power to the system and operate the handle after connecting power to the system. 2. The system does not correspond to the command of the handle. In the test, measure and record the motion condition of the flaps (slats) control plane and the working state of the system with the test system. 5.4.4.11 Fault test The fault test shall include the test of the flaps and slats control handle signal fault, flaps and slats override control panel fault, flaps and slats controller fault, flaps and slats PDU fault, flaps and slats tilt sensor fault, flaps and slats position sensor fault, and flaps (slats) stagnation. Set flaps and slats control handle signal fault, flaps and slats override control panel fault, flaps and slats controller fault, flaps and slats PDU fault, flaps and slats tilt sensor fault, flaps and slats position sensor fault, and flaps and slats stagnation through the tester of the high-lift control system. Measure and record the motion condition of flaps (slats) control plane, the working state of the system, and the fault alarm information with the test system. 5.4.4.12 Durability test The durability test is conducted on a special test bed and the system is installed and configured under all states. The durability test is conducted with the system being loaded. In other words, the maximum working load of deflection angles of flaps (slats) within the flight envelope of the aircraft is taken for load simulation to complete the control test of a certain flight profile of the high-lift system. During the test, data of the following items are recorded: 1. 2. 3. 4.

Displacement of slats rack. Displacement of flaps screw nut. Flaps (slats) control time. Load simulation system state and load.

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5.4.5 Test results and judging criteria 5.4.5.1 Polarity inspection Measure the output voltage of the flaps and slats position sensor and corresponding flaps (slats) deflection. When the flaps and slats control handle is operated to high gear from low gear, flaps (slats) are released and the polarity of the output voltage of the flaps and slats position sensor shall meet the system design requirements. When the flaps and slats control handle is operated to low gear from high gear, flaps (slats) are withdrawn and the polarity of the output voltage of the flaps and slats position sensor shall meet the system design requirements. When the arm button on the override control panel is pressed and the rotary switch is switched to “release” position, flaps (slats) are released and the output voltage of the flaps and slats position sensor shall meet system design requirements. When the rotary switch is switched to the “withdraw” position, flaps (slats) are withdrawn and the output voltage of the flaps and slats position sensor shall meet the system design requirements. If the arm button on the flaps and slats override control panel is pressed in the flaps (slats) control process, the flaps (slats) shall stop moving. 5.4.5.2 Control logic check If the arm button on the flaps and slats override control panel is pressed in the flaps (slats) control process, the flaps (slats) shall stop moving. 5.4.5.3 Brake logic check During system braking, the flaps and slats PDU shall brake before the antiwithdrawing brake device. During brake releasing, the antiwithdrawing brake device shall release the brake before the flaps and slats PDU. 5.4.5.4 Control test under normal mode 1. The flaps (slats) shall operate normally and the control time and sequence shall conform to the design requirements. 2. Slats and flaps position and accuracy shall conform to design requirements. 3. Flaps holding function: after the required load is applied according to the design, the mechanical drive device shall have no modules damaged and the linear displacement measured by the linear displacement sensor shall not change. 4. Flaps synchronicity: under the same command angle, the difference of linear displacement at test point 12 15 in Fig. 5.6 shall not be greater than 5 mm.

Combined test of the flight control subsystem 333 5. Slats holding function: after the required load is applied according to the design, the mechanical drive device shall have no modules damaged and the angular displacement measured by the angular displacement sensor shall not change. 6. Slats synchronicity: under the same command angle, the difference of angular displacement at test point 11 18 in Fig. 5.7 shall not be greater than 0.25 . 5.4.5.5 Control test under degraded mode The judging criteria for the control test under the degraded mode are the same as for the control test under normal mode, as described in section 5.4.5.4. What should be noted is that under the degraded mode, the flaps (slats) control time is doubled. 5.4.5.6 Backup mode retraction test Under the backup mode, the motion, control time, and accuracy of the flaps (slats) control plane shall meet the design requirements. 5.4.5.7 Slats tilt test The output of slats tilt sensor shall meet the design requirements. The slats shall operate normally if there is no slats tilt fault. The slats shall stop moving and keep still at the current position in case of any slats tilt fault. 5.4.5.8 Flaps tilt test The output of the flaps tilt sensor shall meet the design requirements. The flaps shall operate normally if there is no flaps tilt fault. The flaps shall stop moving and keep still at the current position in case of any flaps tilt fault. 5.4.5.9 Flaps and slats asymmetry test The control plane shall move normally if there is no flaps (slats) asymmetry fault. The flaps (slats) shall stop moving and keep still at the current position in case of any asymmetry fault. 5.4.5.10 Unconventional control test of flaps and slats control handle In the unconventional control test of flaps (slats) control handle, the flaps (slats) control plane shall move normally without joggle. They will correspond to the latest command input only. 5.4.5.11 Fault test The motion state, working state, and fault alarm information of the flaps and slats shall meet design requirements.

334 Chapter 5 5.4.5.12 Durability test The flaps and slats drive device and actuator shall have no modules damaged and the displacement of slats rack and flaps screw nut and flaps and slats control time shall meet design requirements.

5.5 Combined test of the automatic flight control system 5.5.1 System introduction As an external loop of the flight control system, the AFCS is cross-linked with the flight management system, autothrottle actuator, flight parameter recording equipment, central warning system, and central maintenance system, etc. to realize autopilot, flight director, automatic navigation, automatic approach/landing, and fault warning. The AFCS is composed of units including the automatic flight control computer (AFCC), automatic flight control panel (AFCP), back drive actuator, autothrottle actuator (ATM), and control switch. Aircraft of some types, such as C-17, are not provided with an independent AFCC, but the control logic and control law resolving functions are completed by the primary flight control computer. The architecture of the AFCS of the background aircraft is shown in Fig. 5.13. The core airborne equipment of the AFCS is AFCC, which receives commands from AFCP and information about PFCS and FMS, and forms three axes (pitch, roll, yaw) and autothrottle actuator control quantity after signal processing, cross transmission, input voting, control law resolving, and output voting. Under the autopilot working mode, AFCS outputs control commands to PFCS to control the aircraft control plane to achieve autopilot. Under the autothrottle working mode, AFCS outputs control commands to the engine control system to control the aircraft engine to achieve speed control and thrust control. Under the flight director working mode, AFCS outputs the director’s commands to the integrated avionics system and generates an operation director on the PFD for the pilot’s reference to realize the flight director function. Meanwhile, the AFCS status display area of PFD displays the current working mode, longitudinal/lateral/throttle working mode, alarm, and other information.

5.5.2 Test objectives After completing the supporting mechanical equipment level acceptance test, the AFCS shall have a system level combined test, which aims to: 1. check the correctness and compatibility of the interfaces of airborne equipment of the system;

Combined test of the flight control subsystem 335

Figure 5.13 Architecture of automatic flight control system of the background aircraft.

2. verify if the interface between the system and external systems meets design requirements; 3. verify if the system control logic meets design requirements; 4. verify if the system alarm and display meet design requirements; 5. verify if system function meets design requirements; 6. verify if system performance meets design requirements;

336 Chapter 5 7. check the correctness and comprehensiveness of system BIT design; 8. check the correctness and rationality of system redundancy management strategy; 9. check if the system and its airborne equipment meet the airworthiness compliance requirements; 10. Check if the system design meets system design requirements, and expose design defects and errors so as to improve the system design.

5.5.3 Test requirements 5.5.3.1 Requirements for tested objects All tested airborne equipment used in the combined test of the AFCS are “S” type ground parts, whose function and performance meet the requirements of special technical conditions and whose technical status conforms to the technical agreement. They are provided with an inspection certificate for delivery and relevant record. The test system used for the combined test of the AFCS includes ground airborne equipment of the system used for the test, simulation element, simulator or exciter, and test devices. 5.5.3.2 Environmental requirements for the combined test It shall be ensured that the environment for the combined test of AFCS can produce dynamic flight parameters, simulate the functions and data of AFCS cross-linked system, and provide an integrated verification environment for AFCS system mechanical equipment. The environment used for the combined test of the AFCS mainly includes the cockpit system, flight simulation system, avionics simulation system, visual simulation system, data acquisition, monitoring and analysis system, and adapter and signal detection panel. 1. Cockpit system The cockpit system is used to provide an operation environment in the cockpit, with airborne equipment including the test control column (wheel), pedal, three-axis loading mechanism and its mounting rack, test throttle lever and console, back drive actuator and corresponding rack, test sensor, and seat. The installation position and mode of airborne equipment shall be consistent with that on the actual aircraft as far as possible. 2. Flight simulation system The flight simulation system can provide the aircraft dynamic model and engine model, etc. and calculate the flight parameters of the aircraft under each flight state in real time, which requires accurate calculation and good real-time performance. 3. Avionics simulation system

Combined test of the flight control subsystem 337

4.

5.

6.

7.

The avionics simulation system simulates the cross-linking of equipment between the AFCS and avionics system and also simulates communication with the bus data crosslinked with the AFCS. The main airborne equipment and systems cross-linked between the AFCS and avionics system include the atmospheric data computer, inertial navigation equipment, flight management computer, display and control system, instrument/microwave landing system, radio altimeter, accident recording equipment, central warning computer, and central maintenance computer. The redundancy configuration of each airborne equipment and system and the data form of the bus interface shall be consistent with that on actual aircraft to realize crosslinked equipment algorithm simulation and bus real-time simulation. Instrument and visual simulation system The instrument simulation system shall be able to simulate the real-time flight data and AFCS status information displayed by the instrument system, which shall be consistent with the actual display of the primary flight control display on actual aircraft. It shall match with the scene to ensure sufficient accuracy and real-time performance. Through the visual simulation system, external information related to the position and attitude of the simulated aircraft can be displayed, which requires accurate display, good real-time performance and no image lag. Data acquisition, monitoring, and analysis system The data acquisition, monitoring, and analysis system collects and records bus and nonbus signals in the test process, monitors the data in real time, and then processes and analyzes them. Adapter and signal detection panel The adapter can provide the interface connecting the dynamic test simulation system and system equipment. The signal detection panel is used by test personnel to detect the signal in the system. Tester The tester of the fly-by-wire flight control system shall be able to connect or disconnect the power supply to the system, test main variables and characteristic parameters of the system, inject excitation signals, and complete the connection of the AFCC/controller and actuating airborne equipment, fault simulation, and modulation/ demodulation of input/output signals.

5.5.3.3 General requirements for test equipment The test equipment shall meet the parameter test requirements of the combined test. All instruments used for the combined tests must be strictly calibrated and within the validity period.

338 Chapter 5 5.5.3.4 Preparation for the test and precautions The following preparations shall be made before the test: 1. The airborne equipment is inspected according to the product acceptance specification. 2. The system tester is checked and tested according to requirements. 3. The static and dynamic test simulation environment has completed equipment integration and debugging. 4. All interfaces connected with the airborne equipment are inspected or tested.

5.5.4 Test items and methods 5.5.4.1 Functional test The functional test of the AFCS includes interface inspection, control logic check, alarm and display inspection, redundancy management test, and BIT test. 1. Interface inspection Interface inspection shall be carried out after the system static impedance inspection is completed. The main items of the interface inspection include: 1. the inspection of bus, discrete quantity, and analog quantity input and output signal interfaces between airborne equipment in the system (AFCC or flight control computer, AFCP, column/wheel /pedal back drive actuator, and autothrottle controller). 2. the inspection of bus, discrete quantity, and input and output signal interfaces between cross-linked airborne equipment simulators outside the system (avionics simulation system, fly-by-wire flight control simulation system and engine simulation system). 3. the inspection of switch quantity interfaces between the system and steering wheel and throttle. 4. the inspection of bus monitoring input data of the system. Inspection method for system interfaces: input and record data to interfaces through equipment exciter and tester according to the AFCS interface control document and observe whether the input value is consistent with the output value. 2. Control logic check The main items of the control and display logic inspection include: 1. the on/off logic inspection of working way (autopilot, flight director, autothrottle). 2. the on/off logic inspection of working mode (attitude hold, altitude hold, and heading hold, etc.) and working mode conversion logic inspection. 3. the control logic inspection when the navigation mode is degraded to default mode. 4. the inspection on indication of AFCP signal lamp and primary flight control display during logic conversion.

Combined test of the flight control subsystem 339 Control logic check method: connect power to the system, load open-loop simulation model on AFCS tester, set initial working conditions through tester or atmospheric data and inertial navigation exciters, select different working modes on control panel, and observe signal lamp indication on control panel and the AFCS working mode display on PFD. 3. Redundancy management test The redundancy management test mainly checks various redundancy management functions, including the correctness of monitoring, voting, fault handling, and reporting. The test items of redundancy management test include: 1. input data voting. 2. computer monitoring (including status monitoring, unrolling monitoring, channel monitoring and branch monitoring). 3. computer degraded logic test. Input data redundancy voting check method: input data of avionics system, fly-bywire flight control system and engine system to ATE, the software of the AFCS has dual-source or multisource data voting to the data according to redundancy configuration, check if the decision table voting algorithm, signal monitoring threshold, and monitoring time delay setting are reasonable and consistent with the design. Computer monitoring logic check method: simulate fault information (including AFCC hardware fault, cable fault, and power supply fault), observe if the system works normally and if the reported fault information is correct through ATE. Computer degraded logic check method: according to the conversion relationship between working ways (primary and primary, primary and backup, single machine) of the computer system, set corresponding conversion conditions through tester or control power supply to verify the correctness and rationality of the logic conversion from primary and primary working way to primary and backup working way or single machine working way, or from primary and backup working way to single machine working way. 4. Alarm and display inspection The alarm and display inspection mainly inspects if the alarm level and display of autopilot disconnection and autothrottle disconnection are correct. The alarm and display inspection method is described below: 1. Connect power to the tester, simulator, and airborne equipment and set system open-loop state and initial state on the tester. 2. Switch to the automatic flight control mode on the AFCP, observe whether the automatic flight control mode is selected on the monitoring page of the tester, and observe whether the corresponding signal lamp is effective on the control panel. 3. Disconnect the working mode through the disconnecting switch on the control panel, steering wheel, or throttle lever and observe on the monitoring page of the tester whether the working mode is disconnected, whether the alarm information is

340 Chapter 5 displayed, whether the alarm level is correct, and whether the corresponding signal lamp on the control panel is effective. 4. By setting fault data through the tester, observe whether the working mode is disconnected, whether the alarm information is displayed and whether the alarm level is correct through the monitoring page of the tester. The alarm information can be added or deleted according to system functions. Different alarm information can be triggered through switches or settings on the tester. 5. BIT test BIT test items: check the completeness of the test items of system BIT (PUBIT, PBIT, IFBIT, MBIT) and the accuracy of fault reporting. Verify the PBIT and MBIT interlocking conditions. The BIT test method is described as follows: 1. Set the system state according to normal and fault conditions, connect power to the system, and observe system PUBIT test results on the tester. 2. Set different combinations of system PBIT interlocking conditions on the tester, start BIT and observe system PBIT test results. The MBIT interlocking condition inspection is similar to that of PBIT. 3. Connect power to the system, set different fault modes, start PBIT/MBIT, and observe whether the system can detect and report faults on the tester. 4. Set different fault modes under the automatic flight control working mode, observe and record IFBIT test results, and analyze the detection coverage and reliability of IFBIT. 5.5.4.2 Performance test The performance test of the AFCS mainly includes: 1. 2. 3. 4.

polarity inspection. transmission ratio inspection. dynamic (static) performance test. stability margin inspection.

1. Polarity inspection The polarity inspection of AFCS includes the polarity inspection under three modes, that is, autopilot working mode, flight director working mode, and autothrottle working mode. 1. Autopilot working mode In the polarity inspection under autopilot working mode, the polarity change of controlled quantity will cause deflection motion of the corresponding control plane and the working polarity and following performance of back drive actuator will be inspected.

Combined test of the flight control subsystem 341 Polarity inspection method under autopilot working mode: when the system works in an open-loop state, switch to autopilot and enter the autopilot working mode; set the target value of the controlled quantity on the tester or AFCP greater than current value, and then observe the motion direction of corresponding control plane through the instrument system to see if it is consistent with the design value; set the target value lower than the current value on the tester or AFCP and observe again if the motion direction of corresponding control plane is consistent with the design value. 2. Flight director working mode In the polarity inspection under flight director working mode, the polarity change of controlled quantity will cause a motion change of the flight director lever. Polarity inspection method under flight director working mode: when the system works in an open-loop state, switch to flight director and enter the flight director working mode; set the target value of the controlled quantity on the tester or AFCP greater than current value and then observe the change of relative position between the flight director lever and fixed aircraft through the instrument system; set the target value lower than the current value on the tester or AFCP and observe again the change of relative position between the flight director lever and fixed aircraft. 3. Autothrottle working mode In the polarity inspection under autothrottle working mode, the polarity change of controlled quantity will cause a motion change of the throttle lever. Polarity inspection method under autothrottle working mode: when the system works in an open-loop state, switch to autopilot and connect the autothrottle to enter the autothrottle working mode; set the target value of the controlled quantity on the tester or AFCP greater than current value and then observe the motion of the throttle lever; set the target value lower than the current value on the tester or AFCP and observe again if the motion direction of throttle lever is consistent with the design value. 2. Transmission ratio inspection The transmission ratio inspection mainly checks whether the transmission ratio of control branches from the control quantity to fly-by-wire flight control command under various working modes is consistent with the design and whether the transmission ratio of autothrottle actuator is consistent with the design. Transmission ratio inspection method: set test state points in the flight simulation system, ensure the system is in a closed-loop state; connect autopilot, flight director, and autothrottle; connect the working mode to be inspected; set different input controlled quantity of polarity through ATE, and observe if the curve is consistent with the simulation curve through the test curve. 3. Dynamic (static) performance test Dynamic (static) performance test aims to check whether the performance of each working mode meets the design requirements at different state points. In the flight

342 Chapter 5 envelope, sufficient flight states (different aircraft weight and center of gravity, altitude and speed) are selected as test state points. Dynamic (static) performance test method: set test state points in the flight simulation system, ensure that the system is in a closed-loop state, connect autopilot, flight director, and autothrottle; connect the working mode to be inspected; observe through the instrument system to see if it meets functional requirements, that is, altitude hold and heading hold etc.; and observe if it meets design requirements through the test curve. Inspect the working mode one by one and set different state points for test again after completion. 4. Stability margin inspection Stability margin inspection aims to check whether the amplitude margin and phase margin of each working mode meets the design requirements at different state points. In the flight envelope, sufficient flight states (different aircraft weight and center of gravity, altitude and speed) are selected as test state points. Stability margin inspection method: set test state points in the flight simulation system, ensure the system is in a closed-loop state; connect autopilot, flight director, and autothrottle; connect the working mode to be inspected; use a dynamic frequency response analyzer to input sinusoidal frequency sweeping signal of certain amplitude and frequency from the ACE test chamber of the fly-by-wire flight control system to actuator of corresponding control plane; observe the motion of corresponding control plane and observe whether the stability margin meets design requirements through the test curve. Inspect the working modes one by one and set different state points for the test again after completion. 5.5.4.3 Fault simulation test The fault simulation test of the AFCS can be performed in a redundancy management test or separately. It shall include the following content: 1. 2. 3. 4. 5.

Power fault Internal equipment fault of system External equipment fault of system Signal fault Cable fault

The fault simulation test can usually be conducted synchronously with the system alarm and display test.

5.5.5 Test results and judging criteria The test results of the functional test, performance test, and fault simulation test, etc. of the combined test of AFCS shall meet the expected value in the test outline. If the measured value is not consistent with the expected value, the test shall be repeated and the test

Combined test of the flight control subsystem 343 method and corresponding operation shall be inspected. The judging criteria of the system combined test shall be subject to top-level documents of the system, such as system interface document, design requirements, and design specifications.

5.6 Combined test of the machinery control system 5.6.1 System introduction Large aircraft are generally equipped with machinery control system for different purposes. Although many large overseas aircraft adopted the fly-by-wire flight control system in early stage, they all retained machinery control system to some degree. For example, aircraft C-17, A320 and B777 all retained the least machinery control system. Aircraft C-17, under the failure of the fly-by-wire flight control system, can improve the aircraft failure protection ability and ensure that the aircraft can safely complete the flight mission through the mechanical control of elevator, lower rudder, aileron and horizontal stabilizer. Aircraft B777 retains mechanical control on a pair of spoiler and horizontal stabilizer. Aircraft A320/330/340 retains mechanical control on the rudder and stabilizer. With the technological development of aircraft and further improvement of the reliability of the fly-by-wire flight control system, many new large aircraft do not retain machinery control system while adopting the fly-by-wire flight control system, such as B787, A380 and A400M. Therefore, when developing the flight control system for one type of aircraft, the technical level of the developer and the reliability level of the fly-by-wire flight control system and its airborne equipment must be considered for determining if the machinery control system is needed. The machinery control system of the background aircraft is taken as an example to introduce the combined test of the machinery control system. The machinery control system of the background aircraft is composed of aileron machinery control system and horizontal stabilizer machinery control system. Its design idea is the same as that of aircraft B777. In other words, when the electrical system is completely cut off, the pilot can still control the aircraft to fly in a horizontal and straight line until the electrical system is restarted. The machinery control system is composed of cockpit control device (lateral control device, horizontal stabilizer trim control handle unit), cable gearing and mechanical backup actuator. The cockpit control device is mechanically connected with the cable gearing input end through the sector gear, the cable gearing output end is mechanically connected to the mechanical backup actuator through sector gear, and the mechanical backup actuator is mechanically connected with corresponding control plane through rocker arm. When the pilots control the steering wheel and the horizontal stabilizer trim

344 Chapter 5

Figure 5.14 Layout of aileron machinery control system.

control handle unit separately or simultaneously, corresponding control plane will deflect to change the aircraft’s motion attitude. The layout of aileron machinery control system is shown in Fig. 5.14 and the layout of horizontal stabilizer machinery control system is shown in Fig. 5.15.

5.6.2 Test objectives The combined test of machinery control system mainly aims to: 1. verify the mechanical interface of the system. 2. verify the function, performance and working stability of the system. 3. verify the flying quality of system aircraft.

5.6.3 Test requirements 5.6.3.1 Requirements for tested objects and tested system 1. All tested airborne equipment are “S” type ground parts and the installation and layout of tested objects shall be subject to relevant requirements of technical conditions for system installation and debugging.

Combined test of the flight control subsystem 345

Figure 5.15 Layout of horizontal stabilizer machinery control system.

2. Before the test, the tested object shall be inspected to see whether there are visible cracks and obvious defects on pull rod, rocker arm, pulley and support and whether the steel cables have fractured wires. 3. The aileron machinery control switch on the flight control panel shall be set in allmechanical state for debugging and test. 4. After the system is well adjusted according to the content in (2) above, press the steering wheel to the left and right manually and push and pull the horizontal stabilizer trim handle manually to check whether the mechanical system has phenomena of clamping stagnation or unstable motion. If there are no such phenomena, test can be carried out, or the fault shall be removed first and then the test can be conducted. 5.6.3.2 Environmental requirements for the combined test The combined test of machinery control system shall be conducted on a special test bed. The specific requirements are shown below: 1. The test bed supports the tested system, airborne equipment, test equipment and test sensor connectors. The test bed shall be designed in 1:1 in principle. If limited by the test site, the wing can be folded parallel to fuselage, the cable at vertical tail can be

346 Chapter 5 shortened, the relative position of the support can be reduced, the wrap angle remains unchanged, and the bed stiffness shall be large enough. 2. If the wing is folded, the process support and process pulley shall be connected with the tested system and the design requirements of the process support and process pulley shall meet the design requirements of the tested system. 5.6.3.3 General requirements for test equipment 1. The measuring range of the sensor shall be generally about 150% of the measured value and the accuracy of the sensor shall be 0.1% F.S. 0.3% F.S. The test signals shall include steering wheel force signal, steering wheel displacement signal, horizontal stabilizer handle force signal, horizontal stabilizer handle displacement signal, left and right aileron control plane rocker arm displacement signal and horizontal stabilizer control plane rocker arm displacement signal. 2. All instruments used for the test must be strictly calibrated and within the validity period.

5.6.4 Test items and methods 5.6.4.1 Aileron machinery control system 1. Polarity inspection Supply pressure to aileron mechanical control actuator and set trim mechanism at neutral position. Rotate the steering wheel in full stroke anticlockwise and clockwise respectively (recommended for 10 times) and check if the motion direction of the left and right wing control plane rocker arm is consistent with that of the steering wheel. 2. Return performance test Supply pressure to aileron mechanical backup actuator and set trim mechanism at neutral position. Rotate the steering wheel anticlockwise slowly to the maximum stroke, and then release it. The steering wheel will automatically return and measure the deflection angle of the steering wheel after the return, that is, the difference between the position of the steering wheel after the anticlockwise return and the neutral position of the steering wheel. In the same way, rotate the steering wheel clockwise slowly to the maximum stroke, and then release it. The steering wheel will automatically return and measure the deflection angle of the steering wheel after the return, that is, the difference between the position of the steering wheel after the clockwise return and the neutral position of the steering wheel. 3. Functional test of transmission ratio adjustment device Set the transmission ratio of the transmission ratio adjustment device at 1, supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, rotate the steering wheel anticlockwise and clockwise to full stroke

Combined test of the flight control subsystem 347 respectively and then measure the rotation angle of the steering wheel and deflection angle of the control plane rocker arm. Set the transmission ratio of the adjustment device at 0.5, supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, rotate the steering wheel anticlockwise and clockwise to full stroke respectively and then measure the rotation angle of the steering wheel and deflection angle of the control plane rocker arm. Convert the transmission ratio of the transmission ratio adjustment device from 1 to 0.5 and record the conversion time; convert the transmission ratio of the adjustment device from 0.5 to 1 and record the conversion time. 4. Clearance measurement Disconnect between the rocker arm unit at the tail end of the left and right wings and the left and right aileron mechanical backup actuators, disconnect the spring in aileron load mechanism and fix the rocker arm unit at the tail end of the left and right wings. Apply a 30 50 N control force to the left and right steering wheels slowly (clockwise and anticlockwise) until they reach the limit position and do not move, and then measure the rotation angle of the steering wheel, that is, the system clearance. Rotate the steering wheel in the reverse direction with the same method and measure the rotation angle of the steering wheel. The sum of the clearance in both directions is just the clearance of the aileron machinery control system. 5. System stiffness measurement Disconnect between the rocker arm unit at the tail end of the left and right wings and the left and right aileron mechanical backup actuators, connect the load mechanism and fix the rocker arm unit at the tail end of left and right wings. Rotate the steering wheel clockwise and anticlockwise respectively (maximum control force up to 220 N) and measure the wheel force and wheel displacement at the corresponding steering wheel position in real time. Through the analysis of measured data, the system stiffness can be obtained through fitting. Disconnect between the rocker arm unit at the tail end of the left and right wings and the left and right aileron mechanical backup actuators, disconnect the spring in load mechanism and fix the rocker arm unit at the tail end of left and right wings. Rotate the steering wheel clockwise and anticlockwise respectively (maximum control force up to 220 N) and measure the wheel force and wheel displacement at the corresponding steering wheel position in real time. Through the analysis of measured data, the system stiffness can be obtained through fitting. Fix the front-end sector gear in cable gearing, disconnect the spring in load mechanism, rotate the steering wheel clockwise and anticlockwise respectively (maximum control force up to 220 N) and measure the wheel force and wheel displacement at the corresponding steering wheel position in real time. Through the analysis of measured data, the stiffness of the system without a load mechanism can be obtained through fitting.

348 Chapter 5 Fix the front-end sector gear in cable gearing, connect the load mechanism, rotate the steering wheel clockwise and anticlockwise respectively (maximum control force up to 220 N) and measure the wheel force and wheel displacement at the corresponding steering wheel position in real time. Through the analysis of measured data, the stiffness of the system with a load mechanism can be obtained through fitting. 6. Plotting of wheel force-wheel displacement-control plane rocker arm displacement relation curve Supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, rotate the steering wheel anticlockwise and clockwise to the limit position respectively, measure the steering wheel force and displacement and the deflection displacement of control plane rocker arm, plot the wheel force-wheel displacement-control plane rocker arm displacement relation curve, analyze and record data, and curve and calculate the displacement at the limit position of steering wheel and control plane rocker arm as well as the starting force of the system. 7. Frictional force and unbalanced force measurement Supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, disconnect the spring in load mechanism, rotate the steering wheel anticlockwise and clockwise to the limit position respectively, and calculate the frictional force and unbalanced force of the system with the recorded curve. Disconnect the aileron mechanical backup actuator, set trim mechanism at neutral position, disconnect the spring in load mechanism, rotate the steering wheel anticlockwise and clockwise to the limit position respectively, measure the steering wheel force and steering wheel displacement and calculate the frictional force and unbalanced force of the system with the recorded curve and data. The formula used for calculation of frictional force and unbalanced force is shown below: f5

pdirect 2 preverse 2

(5.1)

pdirect 1 preverse (5.2) 2 wherein, f refers to frictional force; F refers to unbalanced force; pdirect refers to control force of direct stroke; and preverse refers to control force of reverse stroke. 8. Mass and relative damping coefficient measurement Disconnect the aileron mechanical backup actuator and the spring in load mechanism and then connect a spring of known stiffness and mass (e.g., stiffness 23 N/mm, mass 0.25 kg) at the steering wheel. Fix the other end of the spring, rotate the steering wheel anticlockwise and clockwise respectively until it reaches the limit position and does not move, release the wheel suddenly and the spring will return instantly, and then record the output displacement of the steering wheel. F5

Combined test of the flight control subsystem 349

Figure 5.16 Diagram of oscillation curve.

The basic form of the displacement curve is shown in Fig. 5.16. The damping coefficient of the system can be calculated according to the output displacement curve and data. The formulas used for the calculation of equivalent mass m and relative damping coefficient ζ of the system are respectively shown below: m5

K G 2 5 0:02533 KT 2 2 0:034G ω2 3g
 ln aa12 =2π ζ 5 rffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

 2 1 1 ln aa12 =2π

(5.3)

(5.4)

where in, T 5 t2 2 t1 period, s; G refers to spring mass, kg; K refers to spring stiffness, kgf/cm; g 5 980 cm/s2. If reasonable test results are obtained, more sets of springs should be prepared for progressive increase or decrease by referring to the stiffness and mass data above to obtain an ideal oscillation curve, if possible. 9. Trim stroke and trim time measurement Press the aileron trim switch and rotate the steering wheel clockwise and anticlockwise respectively until it reaches the limit and does not move. At the same time, record the steering wheel’s motion amplitude and time course curve and read the trim stroke and trim time. 10. Step characteristic test Supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, apply 10 and 15 step signal at the steering wheel respectively with a displacement signal generator, measure the displacement signal of steering wheel and the output displacement signal of control plane rocker arm and process them into a time course curve of the tested signals. Through the analysis of test data and curve,

350 Chapter 5 the system delay time, rise time (time to reach 95% steady-state value), overshoot, oscillation times and steady-state time (time to reach 6 5% steady-state error zone) can be calculated. 11. Frequency characteristic test Supply pressure to aileron mechanical backup actuator, set trim mechanism at neutral position, apply sinusoidal signals with amplitude of 10 and 15 and frequency of 0.1 2 Hz at the steering wheel respectively with a displacement signal generator, measure the force and displacement signal of steering wheel and the output displacement signal of control plane rocker arm and process them into a Bode diagram. Through the analysis of the Bode diagram, the phase lag of the system can be calculated. 5.6.4.2 Mechanical backup system of the horizontal stabilizer 1. Polarity inspection Supply pressure to the mechanical backup actuator of the horizontal stabilizer, push and pull the horizontal stabilizer trim handle unit in full stroke for over 10 times to make the horizontal stabilizer control system operates flexibly. Polarity inspection can be conducted according to the motion direction. 2. Return performance test Supply pressure to the mechanical backup actuator of the horizontal stabilizer, push the trim handle slowly forward to the maximum stroke and then release it. The trim handle will return automatically. Measure the deflection angle of the handle after the return, that is, the difference between the position of the handle after the return and the neutral position. Pull the trim handle slowly backward to the maximum stroke and then release it. The trim handle will return automatically. Measure the deflection angle of the handle after the return, that is, the difference between the position of the handle after the return and the neutral position. 3. Clearance measurement Disconnect between the mechanical backup actuator of the horizontal stabilizer and the rocker arm at tail end of the cable gearing, fix the rocker arm at tail end of the cable gearing, apply control force (forward and backward) slowly at the control handle position of the horizontal stabilizer until the handle reaches the limit position and does not move, measure the displacement of the handle, that is, system clearance. Control the handle in the reverse direction with the same method and measure the displacement of the handle. The sum of the clearance in both directions is just the clearance of the entire horizontal stabilizer control system. 4. System stiffness measurement Disconnect between the mechanical backup actuator of the horizontal stabilizer and the rocker ram at the tail end of cable gearing, fix the rocker arm unit at the tail end of cable gearing, push and pull the trim handle forwards and backwards, measure the force

Combined test of the flight control subsystem 351

5.

6.

7.

8.

and displacement at the control handle of the horizontal stabilizer (up to 200 N), and analyze the test data to fit the system stiffness. Measurement of handle control force-handle control displacement-control plane rocker arm deflection speed and measurement of frictional force and unbalanced force Supply pressure to the mechanical backup actuator of the horizontal stabilizer, push and pull the trim handle forwards and backwards, measure the force and displacement at the control handle of the horizontal stabilizer, draw the force displacement relation curve, measure the deflection speed of the rocker arm at the control plane rocker arm, analyze test data and curve and calculate frictional force and unbalanced force with formulate (5.1) and (5.2). Disconnect the mechanical backup actuator of the horizontal stabilizer, measure the force and displacement at the control handle of the horizontal stabilizer, draw the force displacement relation curve, analyze test data and curve and calculate frictional force and unbalanced force of the system with formulate (5.1) and (5.2). Starting force measurement Supply pressure to the mechanical backup actuator of the horizontal stabilizer, measure the force at the control handle of the horizontal stabilizer and the motion displacement of the mechanical input rocker arm, draw the force displacement relation curve, analyze test curve and data to determine the starting force of the system. Mass and relative damping coefficient measurement Disconnect the mechanical backup actuator of the horizontal stabilizer and connect a spring of known stiffness and mass (e.g., stiffness 18 N/mm, mass 0.22 kg) at the control handle of the horizontal stabilizer. Fix the other end of the spring, apply artificial interference signal at the handle and the spring can release suddenly after it is released, record the output signal of the handle, analyze the test data and curve and calculate the damping coefficient of the system with formula (5.3) and (5.4). Step characteristic test Disconnect the mechanical backup actuator of the horizontal stabilizer, apply step signal of 5 amplitude at the control handle of the horizontal stabilizer through a displacement signal generator, measure the displacement signal of the handle and the output displacement signal of the rocker arm at the tail end and process them into time course curve of tested signals, and then calculate the delay time of the tested system, rise time (time to reach 95% steady-state value), overshoot, oscillation times and the steady-state time (time to reach 6 5% steady-state error zone) according to the curve. Disconnect the mechanical backup actuator of the horizontal stabilizer, apply step signal of 10 amplitude at the control handle of the horizontal stabilizer through a displacement signal generator, measure the displacement signal of the handle and the output displacement signal of the rocker arm at the tail end and process them into time course curve of tested signals, and then calculate the delay time of the tested system, rise time

352 Chapter 5 (time needed to reach 95% steady-state value), overshoot, oscillation times and the steadystate time (time needed to reach 6 5% steady-state error zone) according to the curve. 9. Frequency characteristic test Disconnect the mechanical backup actuator of the horizontal stabilizer, apply sinusoidal signals with amplitude of 5 and frequency of 0.1 2 Hz at the handle with a displacement signal generator, measure the handle displacement and handle force and the output displacement of the rocker arm at the tail end and analyze the test data and curve to obtain the frequency characteristic of the output displacement of rocker arm at tail end relative to the handle displacement and handle force. Disconnect the mechanical backup actuator of the horizontal stabilizer, apply sinusoidal signals with amplitude of 10 and frequency of 0.1 2 Hz at the handle with a displacement signal generator, measure the handle displacement and handle force and the output displacement of the rocker arm at the tail end and analyze the test data and curve to obtain the frequency characteristic of the output displacement of rocker arm at tail end relative to the handle displacement and handle force.

5.6.5 Test results and judging criteria 1. Polarity inspection Operate the steering wheel and the control handle of the horizontal stabilizer and record the deflection direction of the control plane rocker arm. The test results shall include the deflection direction of the control plane rocker arm in each channel, and the deflection direction shall meet the design requirements of the system. 2. Return performance test Operate the steering wheel and the control handle of the horizontal stabilizer and record the difference between the return position and neutral position of the control mechanism. The test results shall include the difference between the return position and neutral position of control mechanism in each channel and the test results shall meet design requirements. 3. Functional test of transmission ratio adjustment device Under different states of the transmission ratio adjustment device, record the conversion time of the aileron and transmission ratio adjustment device. The test results shall include the rotation angle of the steering wheel, the deflection angle of the aileron control plane rocker arm and the conversion time of transmission ratio adjustment device under different states of transmission ratio adjustment device, which shall meet the design requirements of the system. 4. Clearance measurement Measure the clearance between the aileron and horizontal stabilizer machinery control system and it shall meet the design requirements of the system.

Combined test of the flight control subsystem 353 5. System stiffness measurement According to the test data, obtain the stiffness of aileron and horizontal stabilizer machinery control system through fitting. The results shall include the stiffness of aileron channel with load mechanism, the stiffness of system without load mechanism, the stiffness of cockpit control device channel with load mechanism, the stiffness of cockpit control device channel without load mechanism and the stiffness of horizontal stabilizer control channel, and they shall meet the requirements of relevant specifications. 6. Force displacement-control plane rocker arm displacement relation curve Through analysis of the test data and curve, the force displacement-control plane rocker arm displacement relation curve of the aileron and horizontal stabilizer control systems and the starting force of the systems can be obtained. The results shall include the control force and control displacement of each channel, displacement of control plane rocker arm and starting force, which shall all meet design requirements of the system. 7. Frictional force and unbalanced force measurement Through analysis of the test data and curve, the frictional force and unbalanced force of the aileron control system can be obtained. The results shall include the frictional force and unbalanced force of each channel and they shall meet the design requirements of the system. 8. Mass and relative damping coefficient measurement Through analysis of the test data and curve, the equivalent mass and relative damping coefficient of the aileron and horizontal stabilizer control system can be obtained. The results shall include the equivalent mass and relative damping coefficient of each channel and they shall meet the design requirements of the system. 9. Trim stroke and trim time measurement Through analysis of the test data and curve, the trim stroke and trim time of the aileron can be obtained. The results shall include the trim stroke and trim time and they shall meet the design requirements of the system. 10. Step characteristic test Through analysis of the test data and curve, the delay time, rise time, overshoot and steady-state time of each control system can be calculated and they shall meet the design requirements of the system. 11. Frequency characteristic test The stroke amplitude of the steering wheel is 10 and the frequency is 0.50 Hz. The phase lag of the control plane rocker arm relative to the force of the steering wheel shall not be greater than 60 , or shall meet the design requirements of the system. The stroke amplitude of the handle of the horizontal stabilizer is 5 and 10 and the frequency is 0.33 Hz. The phase lag of the rocker arm relative to the force of the handle shall not be greater than 60 , or shall meet the design requirements of the system.

CHAPTER 6

“Iron bird” integration test of the flight control system 6.1 Overview Whether it is for the civil aircraft developed as per airworthiness requirements or military aircraft developed as per national military standards, the “iron bird” integration test of the flight control system is the most important, time-consuming, and costly system-level verification test with the most comprehensive verification content. Its importance is mainly reflected in the following aspects: 1. It is the first cross-linking test of the flight control system and the virtual controlled object “aircraft,” that is, the hardware-in-loop test. The virtual aircraft is realized through the flight simulation system calculating aircraft 6DOF nonlinear dynamic characteristics. 2. It is the first integration test of the flight control system and relevant aircraft systems (hydraulic system, power supply system, avionics system, and landing gear control system, etc.), which is also called the integrated test of aircraft system. The physical test environment of aircraft system with the flight control system as the core is established to realize interface design verification, energy demand verification, and logic function verification. 3. It is the first cross-linking test conducted for all subsystems of the flight control system, that is, the integrated verification test. Through the real bus interface, energy interface, electrical interface, and mechanical interface, the airborne equipment and systems of the flight control system are integrated for internal interface design verification and logic design verification of the flight control system. 4. It is the first time for the pilots of the first flight crew to evaluate aircraft performance and experience the flight in a real aircraft system and realistic flight environment, that is, the pilot-in-loop (PIL) test (also called man
machine combined test). The vision system and sound system are adopted to provide pilots with a realistic visual and auditory experience. 5. It is the last overall laboratory verification of the flight control system. All indexes of the system are confirmed to meet the requirements of system design specification and aircraft use and the test conclusion is the basic condition for the first flight of the aircraft. Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00006-1 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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356 Chapter 6 6. The effects of the flight control system faults and other aircraft system faults on flight safety will be verified in this test, such as the hydraulic system fault, power supply system fault, and sensor fault. Meanwhile, the pilot’s fault handling ability will also be evaluated and trained in this test. The “iron bird” integration test of the flight control system is under construction and testing from the aircraft requirements definition to the flight, which is time-consuming and costly. It is always conducted from the test flight for adjustment, test flight for scientific research, and test flight for identification after the first flight till the aircraft identification, aiming to continuously develop and verify the flight control system so as to improve and optimize the aircraft design. The “iron bird” integration test of the flight control system is a verification test and a system installed state integration test carried out after the unit qualification test of the aircraft system, combined test of system, and control law state freezing and software qualification test. Sometimes the unit qualification test is conducted synchronously with the “iron bird” integration test of the flight control system due to the development cycle. However, before the first flight, the airborne equipment must complete the tests required for safety except for the three-proof tests and durability test. Therefore, the “iron bird” integration test of the flight control system has strict requirements for the tested system, airborne equipment, and other aircraft systems participating in the test. The test shall be completely consistent with the installed state, including the software version. The “iron bird” integration test of the flight control system mainly aims to: 1. Further verify the internal interface (mechanical, electrical, bus), function, and performance of each system, and confirm that the internal interface relationship of each system is correct, and that the function and performance (dynamic and static) meet the design requirements of the system. 2. Verify the interface, signal transmission, and logic relationship between each subsystem of the flight control system, and confirm that the internal interface (mechanical, electrical, bus) of the flight control system meets the interface definition requirements. 3. Verify the interface (mechanical, electrical, bus) between the flight control system and other aircraft systems, such as the hydraulic system, power supply system, avionics system, takeoff and landing control system, aircraft control plane, and other aircraft structures, and confirm that it meets the interface definition requirements. 4. Verify the function, performance (time domain and frequency domain), and control logic of each control channel (longitudinal, heading and lateral) and between channels of the aircraft during manual and automatic control, and confirm that they meet the requirements of the flight control system design specification. 5. Verify the hydraulic source pressure and flow, power supply voltage and current, sensor signal quality, and safety redundancy required for the normal and backup

“Iron bird” integration test of the flight control system 357

6.

7.

8.

9.

10.

11.

operation of the flight control system, and confirm that they meet the system working requirements. Verify the control law and software of the flight control system, and confirm that its static characteristics (transmission ratio) can ensure the realization of control law and software and are consistent with the design values, and the dynamic characteristics meet the design requirements of the aircraft flying quality. Verify the BIT (built-in-test) function, redundancy strategy, and modal conversion logic of the flight control system, and confirm that they meet the requirements of the flight control system design specification and control law design requirements. Verify the use and maintenance functions of the flight control system, such as fault synthesis, display alarm, fault reporting, and information storage, and confirm that they meet the requirements of the design specification of the flight control system. Verify the man
machine ergonomics characteristics of the display control of the flight control system, especially the coordination on the control of the three-axis control channel, so as to meet the pilot’s control requirements. Verify the aircraft control characteristics (normal and backup) of pilots of the first flight crew and the fault handling measures, conduct subjective evaluation according to the Cooper
Harper scale, and ensure the pilot-induced oscillations (PIO) assessment results meet design requirements for flying quality and fault handling measures are effective. Make pilots of the first flight crew experience the flight and support the training on pilot skills for pilots of the first flight crew, including the training on use of the equipment in the cockpit.

To achieve the test purpose and engineering practice, the flight control system test of large transport aircraft shall adopt the method of graded synthesis and integration by levels. The “iron bird” integration test of the flight control system is divided into four levels: system “iron bird” integration and verification, flight control system integrated verification, flight control system “iron bird” integration test verification, and man
machine combined test. Level 1: System “iron bird” synthesis and verification. The cockpit control system, machinery control system, fly-by-wire flight control system, automatic flight control system, and high-lift system installed on the “iron bird” test bed are further combined and verified. Level 2: Comprehensive verification of the flight control system. The five systems of the flight control system are integrated on the “iron bird” test bed. The aircraft hydraulic system and power supply system may be the ground simulated energy system, the avionics system may be the simulation system, the control plane load may not be applied, and some sensors may use the simulator. Level 3: “Iron bird” integration test of flight control system. On the basis of level 2 comprehensive verification, the aircraft hydraulic system, power supply system, avionics

358 Chapter 6 system, and other systems conforming to the installed state are integrated with the flight control system. Load must be applied on the control plane and the sensor must be a real part conforming to the installed state, but it may run on an aircraft motion sensor drive device (turntable for short). Level 4: Man
machine combined test. After the above objectives and quantitative evaluation tests, the flight control system shall meet the requirements of the design specification and the design requirements for flying quality. On this basis, the pilots of the first flight crew make subjective and qualitative evaluation of the aircraft and flight control system and the visual and audio system should be connected to the test.

6.2 Test environment and test support equipment The most prominent features of the “iron bird” integration test of the flight control system are as follows. First, the airborne equipment of the flight control system (“S” type parts) and their installation, arrangement, and connection on the test bed shall be consistent with the state on the aircraft for the maiden flight. Second, the aircraft support system required by the flight control system test shall be consistent with the state of the aircraft for the maiden flight. Third, it shall have the most realistic aircraft environment in addition to the aircraft dynamic characteristics. The above features are the overall requirements for the “iron bird” integration test environment and equipment for the flight control system. To build an “iron bird” integration test environment of the flight control system having the above features or meeting the above requirements, the environment should be constructed from overall design, arrangement of test pieces, configuration of test systems, configuration of support systems, configuration of analog simulation systems, and configuration of exciter. 1. Overall design: the overall architecture of the test bed shall be planned according to the arrangement of aircraft control plane, landing gear, engine [including auxiliary power unit (APU)] and aircraft cockpit [including pilots control unit (PCU)] on the aircraft. 2. Arrangement and installation of test pieces: the arrangement and installation principle of the airborne equipment of the flight control system on the test bed shall be consistent with the state on the aircraft for the maiden flight. For the flight control system without a machinery control system, except that the installation of PCU and actuator shall be consistent with the state of the aircraft, other equipment’s installation can be appropriately and flexibly simplified. 3. Configuration of test systems: the “iron bird” integration test of the flight control system shall be configured with all aircraft systems related to the completion of the flight mission and flight safety. In addition to the aircraft hydraulic system, landing gear control system, and the aircraft cockpit layout, which shall be consistent with the

“Iron bird” integration test of the flight control system 359 state on the aircraft for the maiden flight, other airborne equipment for the test shall also be consistent with the state of the aircraft but can be arranged according to the specific position in the laboratory. 4. Configuration of support systems: test support systems such as ground hydraulic system, ground power supply, special tester, data test and analysis system (including test sensor and modulator), and special signal processing and analysis systems, such as dynamic signal frequency response analyzer and data bus signal analysis system, must be configured. 5. Configuration of analog simulation systems: analog simulation systems such as flight simulation system, avionics simulation system, pneumatic load simulation system, atmospheric data simulation system, visual simulation system, and audio simulation system must be configured. 6. Configuration of exciter: exciters such as displacement signal generator, three-axis rate turntable, and linear acceleration turntable must be configured. Generally, they mainly include equipment of the following types: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

“Iron bird” integrated test bed. Aircraft simulator cockpit. Vision system. Sound system. Sensor and test analysis system. Flight test interface (FTI). Flight simulation system. Flight control system tester (including tester of fly-by-wire flight control system, automatic flight control system, and high-lift control system). Simulation exciter of airborne equipment. Mechanical displacement signal generator. Control plane aerodynamic load simulation system. Ground hydraulic energy and ground power supply. Integrated control and management system of integrated test. Drive device of aircraft motion sensor (including three-axis rate turntable, linear acceleration turntable, rate turntable, angle of attack turntable, and dynamic and static pressure simulation system).

6.2.1 “Iron bird” integrated test bed “Iron bird” integrated test bed is a special equipment used for the “iron bird” integration test of a type of flight control system and also the core of the integrated test of the aircraft system. The overall layout of the test bed shall be designed according to the structural arrangement of the aircraft control plane, landing gear, engine (including APU), hydraulic

360 Chapter 6 system, and aircraft cockpit (including PCU) on the aircraft. The arrangement principle of the airborne equipment of the flight control system on the test bed shall be consistent with the state on the aircraft for the maiden flight. For the flight control system without machinery control system (cable drive, rod drive, or torsion bar drive), except that the installation of PCU, actuator, control plane, and hydraulic system shall be consistent with the state of the aircraft, other equipment’s installation can be appropriately and flexibly simplified. The basic functions of the “iron bird” integrated test bed and requirements involve the following aspects: 1. The dimensional design of the “iron bird” integrated test bed shall ensure that the control plane, landing gear, engine drive pump, electric pump, hydraulic system, and aircraft cockpit are arranged and installed based on the state of the aircraft for the maiden flight. 2. The “iron bird” integrated test bed is used to install airborne equipment of the flight control system and other airborne equipment supporting the test, such as the control plane, hydraulic system, aircraft cockpit and control devices, or simulation elements, such as aircraft simulator cockpit, engine drive simulator, and aerodynamic load simulation system drive device. 3. Considering the close relationship between the flight control system, hydraulic system, and landing gear control system, the three systems are generally built on an “iron bird” integrated test bed for testing. 4. Internal electronic airborne equipment of the flight control system or the airborne equipment supporting the system can be appropriately arranged at relevant positions, but the cable interface, length, and characteristics shall be consistent with the real aircraft. 5. The “iron bird” integrated test bed shall be open for the installation and commissioning of tested airborne equipment and test pieces and the site observation and maintenance required for the completion of the test, the installation and parameter testing of test sensors and installation, and maintenance required for the completion of the test, as well as ensuring the safety of operating around the test bed. Meanwhile, auxiliary facilities such as work platform, fence, and access shall be set. 6. The stiffness, strength, and natural frequency of the “iron bird” integrated test bed must be large enough and must not have undesired effects on test results. Therefore the natural frequency of the test bed shall be at least more than three times the working frequency of the tested system, generally designed as more than six times. The maximum working frequency selected for the dynamic test of elevator and rudder channels of large transport aircraft is generally 3 Hz and that selected for the dynamic test of aileron is generally 5 Hz. The maximum frequency selected for the aerodynamic load test of elevator, aileron, and rudder channels is no more than 3 Hz. Therefore the natural frequency of the rear

“Iron bird” integration test of the flight control system 361 section for the “iron bird” integrated test bed is designed to be greater than 9 Hz and the natural frequency of the wing section is designed to be 15 Hz. 7. The “iron bird” integrated test bed shall be of a building block type modular frame structure to ensure a closed local stress bearing and natural frequency. 8. The body of the “iron bird” integrated test bed is separated from a local structure installing test pieces and a transition connector is designed between the test bed and test piece, so that when the structure size of the test piece changes, the installation of the test piece can be ensured by changing the transition connector while not affecting the main structure of the test bed.

6.2.2 Aircraft simulator cockpit The aircraft cockpit is designed based on the mission and tactical and technical performance of the aircraft and fully considers the physiological and psychological factors of humans, as well as the feasibility of cockpit equipment and structure design. The man
machine interface is optimized and attention is paid to man
machine ergonomics. Aircraft cockpit design mainly considers cockpit aerodynamic configuration, vision, spatial size, functional configuration, information distribution, display control layout, and environment. It not only ensures that the pilot can operate all flight control devices and observe all kinds of instrument conveniently, but also provides vision, meeting the requirements of the specifications and allows pilots of different heights to operate the aircraft. The arrangement in the cockpit shall be flexible to some degree. The aircraft simulator cockpit is mainly used to provide pilots with a real cockpit control and display environment, so that pilots may feel that they are in a real aircraft cockpit to meet control needs. The main design content includes: 1. 2. 3. 4. 5. 6.

Overall layout of cockpit. Installation and arrangement of cockpit control device. Implementation of the aircraft cockpit vision design. Arrangement of instrument board. Installation and arrangement of main airborne equipment. Arrangement of pilot seats.

The aircraft simulator cockpit generally adopts a semimonocoque thin-walled structure. In order to facilitate machining and not affect the overall effect of the cockpit, the aircraft simulator cockpit frame and skin are designed according to a real aircraft cockpit configuration and interior decorative surface. The frame adopts a steel plate riveted structure rather than a “Z” profile and the skin adopts aluminum plate. In order to increase the overall stiffness of the aircraft simulator cockpit, the aluminum plate is riveted at the two ends according to its configuration curve, and liner plate and “L” profile are used at the

362 Chapter 6

Figure 6.1 Configuration of aircraft simulator cockpit.

bottom to connect with the frame and the skin along the aircraft simulator cockpit bottom surface. To ensure the installation of the test and testing equipment, the cockpit adopts an open structure. The configuration of the cockpit is shown in Fig. 6.1. In the aircraft simulator cockpit, the cockpit control devices (including left (right) control columns (wheels), pedal unit and command sensor, installation base and transmission connectors) of the flight control system, control display device of the flight control system, multifunctional display device of the avionics system, operation handle and press buttons of the hydraulic system, and landing gear control system are installed to complete the operation and control of the flight control system, hydraulic system, and landing gear control system, as well as the system state adjustment and setting. During the vision design of the aircraft simulator cockpit, the nose and windshield structures shall be designed and machined according to the aircraft structure and the vision in the cockpit shall be the same as that of a real aircraft. Meanwhile, the layout in the cockpit shall ensure that the pilot has a good inner vision and the main flight instruments are not shielded by the control column (wheel) under the normal flight state. The distance between the reference eye position of the pilot and the windshield shall refer to the actual distance of the aircraft, the left (right) position shall be the same as the seat center line and symmetrically arranged at two sides of the aircraft, and the reference eye position and aircraft horizontal line shall refer to the actual distance of the aircraft. The layout of the instrument board shall ensure that the operator can have normal and emergency operation at the preset operation position when it does not make a large body movement. For the instruments and control devices required for takeoff and landing, relevant operators shall be able to see and operate them when sitting at normal positions with shoulder straps fastened. A display control panel is set for the pilot and copilot in the cockpit. According to the position of the control panel, it can be divided into following areas: main instrument board, front control panel, top control panel, central console, left side console, and right side console.

“Iron bird” integration test of the flight control system 363 The main instrument board is machined, produced, and installed according to the aircraft state. The front control panel is located above the main instrument board and slightly below the right front sight line of the pilot and copilot. Light alarm devices, display control panel, and flight director display control panel are arranged and the shading device of the central instrument board is machined and installed according to the aircraft design state. On the top control panel, control units related to the flight control system are installed, such as the control panel of the flight control system, control panel of the hydraulic system, and control panel of the power fuel system. The central console is machined, produced, and installed according to the aircraft state. The installation positions of airborne equipment such as multifunctional display, stop and brake handle, brake control handle, flaps (slats) control handle, flaps (slats) override switch, throttle control board, throttle stop board, trim control board and horizontal stabilizer control handle, and the control panel of the flight control system, control panel of the hydraulic system, and control panel of power fuel system originally installed on top control panel shall be reserved. Other positions of the control panel are covered by blind plate. The left (right) console is machined, produced, and installed according to the aircraft drawing. Only the interfaces for installing the front wheel turning handle and verbal system are designed, the holes used for the installation of the front wheel turning handle and headphone microphone jack unit are reserved and other holes are covered by blind plate. The internal layout of the aircraft simulator cockpit is shown in Fig. 6.2.

Figure 6.2 Internal layout of aircraft simulator cockpit.

364 Chapter 6 The reference point of the seat is the locating point of the seat and it is determined according to the reference eye position. After the seat reference point is determined, the position of the cockpit instrument board, control device, and display that have a close relationship with the seat can be determined. For aircraft with a third pilot, the design of the aircraft simulator cockpit on the “iron bird” integrated test bed of the flight control system configures the seat for the pilot and copilot only and the seat for the third pilot is canceled. The position of the left (right) seat is on the same vertical plane as the design eye position. The pilot seat (including the slide rail) adopts there test piece, the seat of the pilot and copilot are installed in the slide rail on the cockpit floor and can slide forwards and backwards on the slide rail to help the pilot to sit on and get up from the seat easily, and they can also be fixed at a normal position.

6.2.3 Vision system The vision system on the “iron bird” integrated test bed is a compulsory test equipment for the PIL test. Relevant vision is generated according to the flight attitude information of the aircraft and displayed on the dome screen (may also be flat screen display, column screen display, or virtual image display, etc.) view field to provide a vivid outboard scene for the pilot, making the pilot feel in the scene. It is of great importance for the flying quality appraisal in flight stages including takeoff, landing, go-around, accurate tracking, minimumaltitude air drop, and maneuvering flight. Fig. 6.3 shows the vision system adopting the column screen display on the “iron bird” integrated test bed of the flight control system.

Figure 6.3 Vision system on “iron bird” integrated test bed of the flight control system.

“Iron bird” integration test of the flight control system 365 The vision system on the “iron bird” integrated test bed of the flight control system generally has the following functions: 1. Working mode: day, dusk, night visual image with texture. 2. Continuous landform, rivers, fields, runways, and iconic ground objects. 3. The illumination model can simulate the illumination characteristics in different time periods. 4. Atmospheric model can simulate different weather conditions such as rain, snow, clouds, and visibility. 5. Special effects: flame, smoke, dust, spray, wake, etc. 6. Infinite horizon effect. 7. During takeoff and landing, the pilot can judge the height-to-distance ratio between the flying speed and ground object through visual display. 8. Dynamic ocean effects. The vision system on the “iron bird” integrated test bed of the flight control system generally has the following performance indexes. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Field angle: horizontal 180 , vertical 50 . Number of channels: three channels, real image display. Graphics resolution: 1400 3 1050/channel. Graphic complexity: not lower than level 3. Display brightness: no lower than 6 ft  lala (lambert), brightness nonstatutory unit, 1la 5 3.18310 3 103 cd/m2. Graphic contrast ratio: 6:1. Graphic update rate: 60 Hz. Graphic refresh rate: 60 Hz. Transmission delay: 58 ms (when image update rate is 60 Hz). Imaging distance: not shorter than 8 m. Display geometric distortion: ,5% image height.

The hardware of the vision system mainly includes a graphics generation computer and projection display system. The software mainly includes database modeling software, realtime scene management driver software, and image distortion correction software. In the design implementation of the vision system, factors such as resolution, field angle, brightness, refresh rate, update rate, color, texture, shade, shadow, transparency, visibility, detail level, and region of interest shall be considered. The graphics generation computer can produce a realistic view outside the cockpit, including airport runway, lights, buildings, fields, rivers, roads, terrain, and landforms. The vision system shall be able to simulate a variety of meteorological conditions, such as visibility, fog, rain, and snow, as well as daytime, dusk, and night.

366 Chapter 6 The projection display system has an optical design according to the pilot’s eye point. It may be a real image display, such as flat screen, column screen, and dome screen display, or a virtual image display. The projector may be a conventional projector or a laser projector. The projection display device (the screen and the projector) is positioned outside the aircraft simulator cockpit. The composition principle of the vision system is shown in Fig. 6.4.

6.2.4 Sound system The sound system on the “iron bird” integrated test bed is a compulsory equipment for the PIL test. It is used to provide the pilot with a realistic sound environment and make them feel in the scene. Through sound synthesis and generation equipment as well as the sound generation equipment (loudspeaker box) arranged in the cockpit that produce the noise of the engine (piloting, acceleration, and deceleration), air flow, doors (open, close), landing gear (control), wheels touching the ground and brake friction, goods delivery, etc., it can enable the chief test pilot to experience the flight, master flying skills, and adapt to the flight environment.

Figure 6.4 Composition principle of vision system.

“Iron bird” integration test of the flight control system 367 Generally, the sound system on the “iron bird” integrated test bed shall be able to simulate the following noises: 1. 2. 3. 4. 5.

Noise of engine. Aerodynamic noise. Landing gear control noise. Flaps (slats) control noise. Landing and ground touching noise.

The sound system on the “iron bird” integrated test bed shall meet the following performance requirements while realizing various noise effects. 1. The frequency and amplitude of the simulated sound can reflect the state change of the aircraft. 2. When the aircraft simulation system freezes or simulates the crash, the simulated sound disappears automatically. The sound system can simulate noise in the aircraft environment in which the pilot can have a certain degree of intuitive understanding of the working state of each aircraft system according to the sound heard, and to know the change of current operation or aircraft state through the sound change.

6.2.5 Sensor and test analysis system The sensor and test analysis system consists of the test sensor, signal modulator, signal isolator, data acquisition and recording system, and dynamic signal frequency response analysis system. 1. Test sensor The main function of the test sensor is to convert the changes of physical quantities to electrical signals for the data acquisition and recording system and dynamic signal frequency response analysis system to collect, record, analyze, and process test signals. The main test physical quantities, types of sensors, and their technical requirements are as follows: a. Linear displacement sensor: it is used for linear displacement measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. b. Angular displacement sensor: it is used for angular displacement measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities.

368 Chapter 6 c. Tension pressure sensor: it is used for control force and load force measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. d. Tilt angle sensor: it is used for tilt angle measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. e. Torque sensor: it is used for torque measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. f. Wire pull sensor: it is used for flexible displacement test, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. g. Pedal force sensor: it is used for pedal force measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. h. Pressure sensor: It is used for actuator working pressure measurement, with accuracy greater than 0.25%, measuring range greater than the motion range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. i. Flow sensor: it is used for actuator working flow measurement, with accuracy greater than 0.25%, measuring range greater than the range limit of tested physical quantities, and bandwidth higher than the maximum working frequency of the tested physical quantities. 2. Signal conditioner The signal conditioner is a key device for the digital test analysis system to ensure signal accuracy, reliability, and safety and to avoid damage to the test analysis system caused by external abnormal signal or abnormal changes. The signal conditioner has functions including input/output protection, signal and power isolation, common mode rejection, series mode rejection, signal amplification, filter, signal excitation, cold junction compensation, and linear compensation. The signal conditioner provides power excitation to tension pressure sensor, pedal force sensor, torque sensor, linear displacement sensor, angular displacement sensor, tilt angle sensor, pressure sensor, and flow sensor and receives output signals from the sensors above. The front panel of the signal conditioner shall be equipped with a power switch, power indicator lamp, channel working indicator lamp, analog bus socket, and BNC socket. The

“Iron bird” integration test of the flight control system 369 BNC socket can be used to connect the conditioned signals to the data acquisition and recording system and dynamic signal frequency response analysis system. The signal conditioner shall be equipped with gain adjustment, zero point adjustment, sensor excitation adjustment, and filter selector switch, etc. Active filters with different cutoff frequencies can be selected according to the test requirements to realize various combinations of gains for selection. Meanwhile, overvoltage protection and the input filter circuit shall be provided to protect the signal conditioner module and restrain the interference signal on the transmission line. 3. Data acquisition and recording system and dynamic signal frequency response analysis system The data acquisition and recording system and dynamic signal frequency response analysis system shall, on the one hand, complete the test of static parameters, and on the other hand, complete the analysis and processing of dynamic signals in the test of system dynamic characteristics such as step characteristics and frequency response characteristics. This requires that the test analysis system not only can accurately collect and record the signals from the sensor, but also can quickly collect and store these signals in real time, and complete the processing and analysis of the test data. 4. Software of sensor and test analysis system The software of the sensor and test analysis system shall have functions including system configuration, testing and data acquisition, data processing and analysis, analysis display, and data output. a. System configuration i. Channel definition and configuration. ii. Channel calibration. iii. Setting of initial value and upper and lower limit values of test parameters. iv. Data file management (acquisition channel dimension, unit, file of data and test results). v. Test task management (test task name management, test log management, and test subject management). vi. Flexible setting of data acquisition rate, time (number of points), and channels, etc. b. Testing and data acquisition i. Record the sensor output signal at different test points in real time. ii. Have zero compensation and engineering unit conversion for sensor signals at different test points. iii. Conduct communication management for all equipment in the test to ensure coordinated and synchronous completion of test task. iv. Be able to complete the test of clearance, stiffness, control stick force and its displacement, frictional force
unbalanced force, and drive relationship of the flight control system.

370 Chapter 6 c. Data processing i. Process the test signal to obtain time course curve of the test signals. ii. Process the test data to obtain X
Y relation curve of relevant physical quantities. d. Analysis display i. Display and monitor the sensor output signals in real time. ii. Draw and output various graph curves conforming to test reports. iii. Graphic edit function: be able to display graphic curves of different test results on the same screen and have comparative analysis of them. iv. Be able to set any test channels as X-axis input or Y-axis input, be able to set the X-axis and Y-axis drawing proportion, display or hard copy output while drawing multiple curves, complete the acquisition and recording and curve display of signals of multiple channels and the annotation of graph curves. v. Be able to select channel corresponding to the test point flexibly, complete the test analysis of input signals, be able to set conversion coefficient corresponding to the test channel and the conversion formula shown through a mathematical expression, be able to automatically set X-axis and Y-axis drawing proportion, display or hard copy output while drawing multiple curves, complete the data storage of multiple channels and the annotation of graph curves, print and output test results meeting the requirements of test reports. e. Report generation and data output i. Generation of tabular data conforming to test report according to the test data. ii. Management of time course test data and record files. iii. Generation and output management of list files. iv. Generation and output management of graphic files. v. Time course data files, curves and graphic files. vi. Test plotting and report generation; vii. The format of all saved acquired data supports the call of software platforms such as Labwindows/CVI, Matlab, HP VEE, and Excel. In the “iron bird” integration test of the flight control system, as thousands of test items are required to be tested and some tests are to be cross-linked with the adjustment of the flight control system to be completed, it involves a complex test process and large data quantity. In the test, on the one hand, not only the test of static parameters shall be completed, and on the other hand, the analysis and processing of dynamic signals in the test of system dynamic characteristics such as step characteristics and frequency response characteristics shall also be completed. This requires that the test analysis system can not only accurately collect and record the signals from the sensor, but also can quickly collect and store these signals in real time, and complete the processing and analysis of the test data. The distributed test analysis system based on

“Iron bird” integration test of the flight control system 371 Operaonal control computer

LXI data switching exchange

Graph and data browsing computer

Nose data acquision point

Le wing data acquision data

Le wing data acquision data

T-tail empennage data Acquision point

Data acquision:2×48 channels +200 channels 2×EX 1000A+E 34980A

Data acquision:96 channels EX 1000A+EX 1629

Data acquision:96 channels EX 1000A+EX 1629

Data acquision:96 channels EX 1000A+EX 1629

Signal measurement: 89 channels Signal condioning: 16×6=96

Signal measurement: 89 channels Signal condioning: 16×6=96

Signal measurement: 96 channels Signal condioning: 16 ×6=96

Signal measurement: 42 channels Signal condioning: 4×16=48

Signal measurement: 187 channels Signal condioning: 40×5=200

Horizontal stabilizer actuator Horizontal stabilizer deflecon signal

Rudder actuator signal

Rudder deflecon signal

Elevator deflecon signal

Elevator actuator signal

Trailing edge flaps torque speed Leading edge slats torque speed Mul-funconal spoiler signal

Ground spoiler signal

Sensor and conditioner signal Transmission line

Actuator displacement force Right aileron deflecon angle

Ground spoiler signal

Trailing edge flaps torque speed Leading edge slats torque speed Mul-funconal spoiler signal

Actuator displacement force Le aileron deflecon angle

Flight simulaon signal

Flaps (slats) controller tester signal Automac flight control computer tester signal Actuator controller tester signal Fly-by-wire flight control computer tester signal

Transmission rao device displacement Postback actuator displacement

Handle control force

Pedal displacement ,foot Steering wheel displacement force Control column displacement force

LXI data bus

Conditioner and collector signal transmission line

Figure 6.5 Architecture of flight control system “Iron Bird” integration test and Measurement system.

the LXI bus becomes the optimal scheme. Fig. 6.5 shows the composition of the test analysis system for the “iron bird” integration test of the flight control system of a type of large transport aircraft.

6.2.6 Flight test interface The FTI is used to obtain and monitor the data information of the flight control computer. It generally realizes the communication with the flight control computer through the data bus and completes the acquisition and analysis of the internal data information of the flight control computer according to the FTI communication protocol of the flight control computer. It also displays the state parameters of the flight control system, so that engineers can observe and analyze the working state of the system. FTI generally receives the following types of data information from the flight control computer: 1. 2. 3. 4. 5. 6.

Working state information and its voting value. Alarm and fault state information and its voting value. Command sensor state information and its voting value. Feedback sensor state information and its voting value. Cross-linked airborne equipment state information and its voting value. Other relevant state information and its voting value.

372 Chapter 6

Figure 6.6 Composition principle of the flight test interface.

The information above can be discrete quantity and switch quantity signals or analog quantity signals or data bus signals. Possible buses include ARINC429, MIL-STD-1553B, AFDX, RS-422, CAN, and FC. The composition principle of FTI is shown in Fig. 6.6.

6.2.7 Flight simulation system The flight simulation system is one of the most important pieces of test equipment used in the “iron bird” integration test of the flight control system. In the “semiphysical” simulation test of the flight control system, the aircraft motion solved by the flight simulation system is the nonphysical part. The flight simulation system collects the control plane signal in real time and solves the nonlinear 6-DOF motion equation and the engine thrust equation of the aircraft to obtain the motion state parameters of the aircraft. The calculated aircraft motion state parameters are sent through the real-time communication interface to aircraft motion sensor excitation devices such as the three-axis turntable and line acceleration turntable as

“Iron bird” integration test of the flight control system 373 well as the avionics simulation system, vision system computer, and instrument system computer. The calculated flight altitude, flight speed, and other results are transmitted to the control plane load simulation system as the command signal of the aerodynamic load simulation system, to the avionics system simulation exciter as the command signal of the avionics system simulation exciter, and to simulation equipment such as inertial navigation, atmospheric data system, and radio altimeter. The flight simulation system can display the state of the aircraft, store the motion response curve and data of the aircraft, and print the curves. 1. Frame structure of flight simulation system The flight simulation system generally adopts a distributed computer structure and its composition principle is shown in Fig. 6.7. The flight simulation system is hierarchized according to the functions realized by the airborne equipment and then the functional units in each layer are encapsulated and organized by the object-oriented technology. The functional objects of the same layer have information and data exchange through the object interface, while objects of different layers have communication through the standard interface. The simulation scheduling management system on each node is operated based on the client/server mechanism, and the simulation control and simulation application are located between application layers and connect the

User

Node A

System control

System monitoring

Simulaon control

Standard interface

S1

Sn

Data transmission and communicaon middleware

Real-me simulaon service Data storage service Simulaon process management: engineering establishment, configuraon state monitoring, parameter modificaon, data storage

Simulaon control interface System monitoring

System control

Standard interface Simulaon scheduling management system

Service set

S1

Data transmission and communicaon middleware

Operang system Network and hardware system

Figure 6.7 Frame structure of flight simulation system.

Execuve layer

Operang system Network and hardware system

Sn

Support core layer

Simulaon scheduling management system

Service set

Simulaon applicaon

User Applicaon layer

Operang control interface

Node B

374 Chapter 6 client and the server. The distributed flight simulation system is divided into three levels: simulation application layer, distributed support core layer, and execution structure layer. a. Execution structure layer In the distributed flight simulation system, the operating system built on the computer hardware will be regarded as the execution center of specific simulation calculation tasks and the ultimate undertaker of distributed simulation. The execution structure layer provides the upper layer with a running platform and is the basis for the whole distributed flight simulation system. b. Distributed support core layer As the core of the distributed flight simulation system, the distributed support core layer is responsible for the operation scheduling, management, and information processing of the whole system, including data and communication middleware, simulation scheduling management, service set and standardized interface. The data and communication middleware provides a unified data processing place and cross-platform working mechanism for the kernel of the distributed flight simulation system and hides the differences between different network systems. A variety of communication protocols are used to achieve the transparent distribution of the simulation model and data between nodes, which are key to the distributed flight simulation system being able to achieve a distributed cross-platform simulation calculation. The simulation scheduling management system determines the best simulation strategy according to the specific requirements of the simulation task and controls the simulation execution process. Various information and data generated in the simulation process are collected by invoking the service as the basis for performance analysis of the flight control system and functions such as thread, communication, and data processing are controlled and managed. The standard interactive interface is a bridge realizing the communication between the simulation application layer and the distributed system support layer. The standardized design of the interface is conducive to the transparency of the system distribution and the reusability of the system model. From this, it can be seen that by providing distributed simulation technical support to the upper layer, the distributed core layer has specified separation of the simulation application description and distributed simulation, which is of great help to the functional expansion and reliability improvement of the whole system. In addition, the simulation application layer and distributed simulation support core layer can be independently and simultaneously developed, which is of great significance to shorten the development cycle of the distributed flight simulation system or to realize a new simulation system.

“Iron bird” integration test of the flight control system 375 c. Application layer The application layer is the top layer of the entire distributed flight simulation system and it is responsible for the concrete implementation and management of the simulation task object model, as well as the man
machine interaction for performance monitoring and simulation control. From a horizontal perspective, there are two categories of application layer, that is, simulation application category and monitoring and control category, and their relationship is similar to the client/ server model with middleware. The monitoring and control category, as the customer, submits performance data and other relevant information in the actual running of the system to the user (the operator). The simulation application category, as the concrete implementation of the simulation task, carries out the concrete simulation calculation. It uses object-oriented and modular technology to break down and distribute the simulation task and provides a series of functions related to the simulation model. 2. Flight simulation software package The flight simulation software package includes the aircraft system (including hydraulic system, fuel system, power system, and flight control system) performance calculation module, aircraft aerodynamic performance calculation module, aircraft mass performance calculation module, landing gear system performance calculation module, engine performance calculation module, and aircraft 6-DOF full amount motion equation calculation module. Its composition principle is shown in Fig. 6.8. a. Aircraft system performance calculation module. Dynamic models of aircraft systems including hydraulic system, fuel system, power system, and flight control system and their cross-linking relationship are established, and the models above are solved in real time and the control plane deflection angle, fuel consumption, and power extraction are output.

Figure 6.8 Composition principle of the flight simulation system.

376 Chapter 6 b. Aircraft aerodynamic performance calculation module The aircraft aerodynamic performance calculation module is used to describe the aerodynamic properties of aircraft. Its basic principle is to calculate the current six components of aircraft aerodynamic coefficient, aerodynamic force, and aerodynamic moment according to the given aircraft’s shape and flight conditions at each instant. Large aircraft generally include leading-edge slats, trailing-edge flaps, aileron, rudder, adjustable horizontal stabilizer, elevator, spoiler, and other control planes, as well as aircraft airborne equipment that have impacts on aircraft aerodynamic performance, such as engine, landing gear, and cargo compartment door, leading to multiple aircraft aerodynamic configurations, complex modeling, and large amounts of data processing. c. Aircraft mass performance calculation module The aircraft mass performance calculation module adopts a numerical calculation method for mass calculation according to the aircraft mass data and is used to describe the overall mass property of the aircraft. Its basic principle is to calculate the mass of the whole aircraft, center of mass, rotational inertia and product of inertia, rotational inertia matrix and its inverse matrix based on current residual fuel quantity and cargo loading condition. d. Landing gear system performance calculation module. The landing gear system performance calculation module consists of three parts, that is, landing gear control performance calculation module, front wheel turning performance calculation module, and aircraft wheel antiskid and brake performance calculation module. The landing gear control performance calculation module is used to calculate the control time and control angle of the aircraft landing gear and describe the influence of the control process on the aircraft’s aerodynamic performance. The front wheel turning performance calculation module, on the basis of establishing aircraft ground taxiing aerodynamic performance and front wheel turning system performance, calculates the dynamic relationship between the turning control command and turning angle and the directional motion and attitude changes of the aircraft in real time. The aircraft wheel antiskid and brake performance calculation module, on the basis of establishing aircraft ground taxiing aerodynamic performance and wheel antiskid and brake system performance, calculates the brake command and antiskid, brake, and bias correction performance during ground taxiing of the aircraft in real time. e. Engine thrust performance calculation module The engine thrust performance calculation module is used to describe the physical properties of the aircraft engine, that is, the dynamic response of thrust, fuel consumption rate, and exhaust temperature, etc. to speed and altitude, and the dynamic response of speed to the throttle.

“Iron bird” integration test of the flight control system 377 f. Aircraft 6-DOF full amount motion equation calculation module The basic principle of the aircraft 6-DOF nonlinear full amount motion equation calculation module is to calculate the linear acceleration and angular acceleration along the aircraft body axis by obtaining flight parameters such as aircraft attitude, speed and acceleration, aircraft position information and force and moment, and synthesizing various forces and moments of the aircraft such as gravity, aerodynamic force, engine thrust, and various torques, and then generate the resultant speed, angular speed, and Euler angle of the aircraft. 3. Communication interfaces of flight simulation system The flight simulation system should drive a variety of test equipment and have signal transmission and data communication with ground test equipment of the flight control system, display and control system, avionics system, aerodynamic load simulation system, vision system, and integrated test integrated management system. a. Interface with flight control system The flight simulation system receives the processed control plane deflection angle sensor signals from the data acquisition and processing system, as well as the engine throttle position signals. These signals are the inputs calculated with the aircraft 6-DOF nonlinear full amount motion equation. And the motion parameters of the aircraft are calculated based on the atmospheric signals provided by the atmospheric data system. b. Interface with turntable The turntable is a high-precision and wide-broadband ground simulation device. The flight simulation system provides aircraft motion parameters to the turntable in real time through the data interface. The three-axis turntable can accurately convert the aircraft attitude and angular rate motion parameters of the flight simulation system to three-axis mechanical motion. The linear acceleration turntable can accurately convert the aircraft acceleration motion parameters of the flight simulation system to acceleration mechanical motion. The aircraft motion sensor (angular rate gyro assembly, acceleration sensor) installed on the turntable can sense the aircraft attitude, angular rate, and normal overload acceleration. c. Interface with display and control system The display and control system displays various information and parameters of aircraft control and motion, among which the flight motion parameters of the aircraft come from the flight simulation system. In the “iron bird” integration test of the flight control system, no matter whether the simulated display system or the real display system of the aircraft is adopted, the flight-related information displayed by the display system includes the aircraft motion parameters from the flight simulation system and the state and fault display information of the aircraft system. As the flight simulation system and display and control system have a large amount of data transmission and have high requirements for real-time performance, gigabit Internet is generally adopted for communication.

378 Chapter 6 d. Interface with vision system The vision system receives motion parameters such as the position and attitude of the aircraft from the flight simulation system and generates the environment and scene outside the aircraft cabin in real time. As the flight simulation system and vision system have a large amount of data transmission and have high requirements for real-time performance, gigabit Internet is generally adopted for communication. e. Interface with avionics system Generally, the avionics system interface equipment adopts a reflective memory network card to receive signals such as air speed, altitude, and wheel load of the aircraft from the flight simulation system and then converts these signals to ARINC429 signals and sends them to the flight control system. Through the realtime network card, the flight control system receives signals such as the attitude and angular acceleration of the aircraft from the flight simulation system and converts these signals to ARINC429 signals and sends them to the avionics system. f. Interface with integrated test integrated management system As a top-layer management device of the “iron bird” integration test of the flight control system, the integrated test integrated management system manages all the equipment in the test process. The flight simulation system shall be able to receive the command information from the test management system through the real-time network, analyze the command according to the preset command spectrum, complete relevant operations according to the command, and finally send the feedback information to the test management system.

6.2.8 Flight control system tester The flight control system tester is a special test facility developed for “iron bird” integration test of the flight control system. It can be used for detecting and observing the main variables, characteristic parameters, and technical status of the system, injecting excitation, setting fault, detecting response, and completing partial data acquisition tasks. The parameters that can be monitored by the flight control system tester shall include at least the following content: 1. Power characteristic value (primary power supply, secondary power supply). 2. Segment detection point parameters of the signal chain. 3. The analog quantity, discrete quantity, and digital quantity signals of the front interface and rear interface of the flight control computer. 4. BIT detection values of sensor, flight control computer, and servo actuator system. 5. Observation and detection points required by the designer. The flight control system tester realizes the interconnection communication between airborne equipment of the flight control system, completes the signal transmission and data

“Iron bird” integration test of the flight control system 379 information exchange, realizes signal isolation and interconnection communication between the flight control system and the ground test equipment, and completes the signal transmission and data information exchange between the airborne equipment and the ground test equipment. The flight control system tester is composed of the tester chassis, signal disconnection module, signal isolation module, logic control and adder module, DC power module, interface adapter, cabinet, and adapter cable. Given the complex flight control system architecture of large transport aircraft and various airborne equipment, the flight control system tester is generally configured with three parts separately, that is, fly-by-wire flight control system tester, automatic flight control system tester, and high-lift control system tester. The composition of the flight control system tester is shown in Fig. 6.9. 1. Fly-by-wire flight control system tester The fly-by-wire flight control system tester completes the interconnection between the fly-by-wire flight control computer (PFC) and actuator control electronics (ACE) and other airborne equipment (display and control system, cockpit control command sensor, actuator surface position detection sensor, overload sensor) of the fly-by-wire flight control system and realizes the transmission of the switch quantity signal, DC analog quantity signal, and AC analog quantity signal during the interconnection of the airborne equipment above. Besides, during the signal transmission in the interconnection of the airborne equipment above, it disconnects the equipment and transmits the signals to the detection panel to detect the signals being transmitted or injects the excitation signals at the disconnection position to the airborne equipment. The fly-by-wire flight control system tester can isolate and drive the transmission of signals between the fly-by-wire flight control computer, the actuator controller, and the test support equipment and realize the communication between the fly-by-wire flight control system and the test support equipment. In order to detect the interconnection signals between the fly-by-wire flight control computer and the airborne equipment, after the signals above are isolated and driven, it can be connected to a test recording device for acquisition and recording of the above signals. 2. Automatic flight control system tester The automatic flight control system tester completes the interconnection between the automatic flight control computer (AFCC) and other airborne equipment and realizes the transmission of switch quantity signal, DC analog quantity signal, and AC analog quantity signal during the interconnection of the airborne equipment above. Besides, during the signal transmission in the interconnection of the airborne equipment above, it disconnects the equipment and transmits the signals to the detection panel to detect the signals being transmitted or injects the excitation signals at the disconnection position to the airborne equipment. It can isolate and drive the signals transmitted between the AFCC and the test support equipment.

380 Chapter 6 Mechanical and electrical system simulator Engine param meter recording system simulaon ARINC429 An icing/de-iicing system simulaon, 6 channeels, 2 redundancies Undercarrriage control system s simulaon, 4 channels, c 2 reedundancies

oller Flaps(slats)contro

Flaps(slats)co F ontrol handle,, 4 redun ndancies F Flaps(slats)ov verride switch h

High li system teester Analog, digital, discrete signal transmission disconn necon Analog, digital, discrete signal detecon Signal isolatiion drive, datta bus signal coupling

Flaps lt sensor,8,2 redundanccies 2 Slats tilt check device,2,2 redundancies Flaps control plane positio on sensor,2,2, redundancies r s Slats control plane positio on sensor,2,2, rredundanciess

Fly-by-wire flight f control computer Actu uator controller Command ffrom control mechanissm,3×2,2 dancies redund Command from brake handle,2, 4 redundancies Control display device

F Fly-by-wire fliight control system tester Analog, digital, discrete signal transmisssion disconn necon Analog, digital, discrete signal detecon Signal isolaon drive, data bus signal coupling Autom mac flight co ontrol computer

Automatic flight control device Engine throle control console

Automatiic flight contrrol system tester Analog, digital, discrete signal transmission disconnecon Analog, digital, discrete signal detecon Signal isolation drive, datta bus signal coupling

Elevator position,4,2 redundancies Aileron po osition,2,2 redundancies Rudder po osition,2,2 redundancies Stabilizer position,2,2 redundancies Three-axis angular rate,2, 4 redundancies Three-axis acceleration,2,4 redundancies Wing oveerload,2, 2 redundancies Vertical tail overload,2, 2 redundancies

Control colu umn postback actuator,4 redundancies r s Steering wh heel postback actuator,4 redundancies r s Petal postbaack actuator,4 redundancies

Avionics ssystem simulator Atmospheree data computer, 4 redundancies Inera/satellite integrated navigaon n, 4 redundancies Radio almeter, 2 redundancies

ARINC429data bus

Flight con ntrol system 2 redundancies Central maintenance system 2 redundancies

n Signal transmission line Airborne system sign nal transsmission line

Figure 6.9 Composition of flight control system tester.

AFDX data d bus

“Iron bird” integration test of the flight control system 381 3. High-lift control system tester The high-lift control system tester completes the interconnection between the highlift control system flaps and slats controller unit (FSECU) and other airborne equipment and realizes the transmission of the switch quantity signal, DC analog quantity signal, and AC analog quantity signal during the interconnection of the airborne equipment above. Besides, during the signal transmission in the interconnection of the airborne equipment above, it disconnects the equipment and transmits the signals to the detection panel to detect the signals being transmitted or injects the excitation signals at the disconnection position to the airborne equipment. It can isolate and drive the signals transmitted between the flaps (slats) controller and the test support equipment. The flight control system tester shall also be configured with a bus coupler to realize the transmission of data information during the interconnection communication between the fly-by-wire flight control computer PFC and other airborne equipment through the ARINC 429 data bus, MIL-STD-1553B data bus, and AFDX data bus. As an important equipment of the “iron bird” integration test of the flight control system, the flight control system tester shall have the following functions and performances: 1. General requirements a. All signals led to the front panel are classified and clearly marked. b. Signals are led from the disconnection terminal in a convenient, safe, and reliable way to complete detection. c. Signals are led from the disconnection terminal to the recording instrument in a convenient, safe, and reliable way to complete signal detection, recording, storage, and playback analysis. d. Signals are input to the disconnection terminal in a convenient, safe, and reliable way to complete signal input. e. Signals are input to the test object from the disconnection terminal through instruments such as signal generator in a convenient, safe, and reliable way to complete signal input. f. A coupler or a data bus connector is set in the connection path of the data bus signal so as to realize the interconnection communication between the data bus signal and other test devices in a convenient, safe, and reliable way. g. It shall be able to simulate the work of the redundant power supply and battery on the aircraft and provide power to the flight control computer. h. The insulation resistance between signal transmission cables shall be greater than 10 MΩ. i. The resistance of the signal transmission cable shall not be greater than 0.01 Ω. j. The signal transmission cables shall have same characteristics and performance of the airborne cables. k. Electrical connectors shall have the same characteristics and performance as the airborne electrical connectors.

382 Chapter 6 l. Grounding requirements. The design of the grounding system of the flight control computer shall prevent ground return and the return of the signal ground together with power ground and shield key signals effectively. m. Each circuit of the power supply input on the power connector of the flight control computer shall be returned to the ground wire correspondingly and the ground wire shall not be connected to the cabinet body from inside the cabinet. n. Shield ground. An independent plug-in shall be provided for signal shielding, the shielding pin shall be adjacent to the corresponding signal pin, and the shielding pin of the plug-in shall be connected to the chassis of the flight control computer. o. Excitation signal injection requirements. i. Programmable control excitation signal of random waveforms and functions shall be provided. ii. Standard waveforms such as sinusoidal wave, square wave, oblique wave, triangular wave, noise, and DC shall be generated and their frequency can be adjusted manually. iii. Random waveforms such as exponential increase, exponential fall, negative oblique wave, and sinx/x shall be set inside. iv. The frequency range shall be 1 μHz
80 MHz and the amplitude range shall be 10 mVpp
10 Vpp. v. In addition to AM, FM, PM, FSK, and PWM carrier wave modulation functions, a standard waveform signal (custom) can be loaded to form a carrier wave and then the carrier wave will be sent to the tested system. vi. It shall have linear and logarithmic scanning and pulse string mode. p. Signal detection requirements i. The test accuracy shall be better than 6 0.25%. ii. The test resolution shall be higher than 16 bit. iii. The acquisition rate shall be greater than 1 kSa/s. iv. Voltage signal (mV, V) shall be adopted for test mode. v. Differential mode shall be adopted as the signal input mode. vi. The input impedance shall be greater than 1 MΩ. vii. The measuring range shall be greater than 100 V. 2. Special requirements a. Different environmental conditions required by the test of the flight control system shall be provided to support the completion of all functional verification and performance tests of the flight control system. b. It shall have an interface adapter with complete functions that works reliably to realize the interconnection between the airborne equipment of the flight control system and the interconnection between the flight control system and the avionics system, landing gear control system, antiicing and deicing system, engine parameter recording system, and equipment.

“Iron bird” integration test of the flight control system 383 c. It shall realize the signal transmission and data information exchange between airborne equipment of the flight control system as well as between the flight control system and the avionics system, landing gear control system, antiicing and deicing system, engine parameter recording system, and equipment. d. A detection panel with complete functions that works reliably shall be configured. During the signal transmission between airborne equipment of the flight control system and between the flight control system and the avionics system, landing gear control system, antiicing and deicing system, engine parameter recording system and equipment, a disconnection block shall be used to lead the signal lines to the detection panel for leading-out detection of signals or excitation injection. e. The connection and signal transmission between the fly-by-wire flight control computer and airborne equipment such as actuator controller, command sensor, display and control system, and feedback sensor shall be realized. f. The connection and signal transmission between the fly-by-wire flight control computer and avionics system, landing gear control system, antiicing and deicing system, and equipment shall be realized. g. The connection and signal transmission between the fly-by-wire flight control computer and AFCC, flaps (slats), and equipment shall be realized. h. The connection with the airborne equipment cabinet shall be realized and the connection and signal transmission between the actuator controller and other airborne equipment shall be completed. i. The connection with the airborne equipment cabinet shall be realized and the connection and signal transmission between the AFCC and other airborne equipment shall be completed. j. The connection with the airborne equipment cabinet shall be realized and the connection and signal transmission between the flaps (slats) controller and other airborne equipment shall be completed. k. A disconnection terminal shall be set in the connection path of the analog quantity, switch quantity, and power supply signals and it shall lead to the front panel so as to disconnect or connect the signals conveniently and safely.

6.2.9 Avionics system exciter The avionics system exciter is used in the “iron bird” integration test of the flight control system to excite relevant equipment to generate the required system signals, mainly including the atmospheric data system, inertial navigation system, radio altimeter, remote data concentrator (RDC), flight data recording system, central warning system, central maintenance system, flight management system, display processing system, primary flight display, multifunctional display, multifunctional keyboard, trackball, display control panel, and engine indication and unit alarm system of the flight control system. Signals required

384 Chapter 6 for the working of the flight control system include atmospheric data, aircraft attitude, angular rate, aircraft weight, center of gravity position and the status information, alarm information, control plane position, and maintenance information and the aircraft trim position information of the flight control system shall be displayed on the primary flight display, multifunctional display, and engine indication and unit alarm system. Units such as menu, window, button, header, table, and graph shall be set on the operation interface of the avionics system exciter to receive the user’s operation input, realize the control or state adjustment of the system, and complete the input and sending or receiving and display of data information. The avionics system exciter shall be configured with a reliable simulation function to simulate its function and working process and realize convenient, safe, and reliable signal transmission with equipment such as the flight control computer. The composition of the avionics system exciter is shown in Fig. 6.10. The avionics system exciter shall be able to simulate the data exchange function and data storage function of the avionics system integrated processor and the functions of the RDC to realize the transmission and processing of the data information between the flight control computer, mechanical and electrical management computer, and avionics display processing system.

Figure 6.10 Architecture of airborne equipment simulation exciter.

“Iron bird” integration test of the flight control system 385 Requirements for the functions of the avionics system exciter are as follows: 1. It shall have complete interface simulation function of the avionics system and be able to simulate the data transmission of relevant system. The data bus signals shall be interconnected, can go through the bus coupler, and are interchangeable with the airborne equipment. 2. It shall realize the interconnection communication and data information transmission between airborne equipment, such as flight control computer and mechanical and electrical management computer, including various bus signals, and connect with external equipment through a bus coupler. 3. It shall have data bus signal simulation function, send and receive signals through the data bus, and control the complete operation interface to select the channel, set parameters of the simulation channel, and control the data information transmission of each simulation channel. 4. It shall have the display interface simulation function of avionics system to simulate the display interface of the corresponding system, so as to realize the processing, display, storage, and other functions of the transmitted data information. 5. A dual-redundancy real-time Internet environment shall be constructed to connect to the “iron bird” integration test of the flight control system, the cross-linking test of the flight control system and hydraulic system, power supply system, landing gear control system, the cross-linking test of the flight control system and avionics system, as well as the avionics devices required for debugging. 6. It shall be able to simulate the integrated processing computer and send control commands to the flight control computer and electrical and mechanical management computer and receive the status information sent from the flight control computer and electrical and mechanical management computer through the same data bus, and analyze, display, and record according to the corresponding ICD. 7. It shall be able to simulate the function of the RDC to ensure the automatic flight control system and high-lift control system can receive the control information sent from the integrated processing computer through the RDC. Accordingly, the status data returned from the automatic flight control system and high-lift control system shall be converted to AFDX bus data through the RDC and then be distributed to the integrated processing computer and display processing unit.

6.2.10 Mechanical displacement signal generator In the “iron bird” integration test of the flight control system, the mechanical displacement signal generator is used to produce mechanical displacement commands (including linear displacement and angular displacement) of various waveforms to simulate the pilot’s handling of the control column (wheel) and pedal and complete the mechanical input of the pilot’s control command.

386 Chapter 6 The main technical requirements for the mechanical displacement signal generator are as follows: 1. The frequency range and amplitude for frequency sweeping test can be set and changed through the control interface and there are outputs of pulse, sinusoidal, square wave, oblique wave, and triangle wave signals, etc. 2. The output and test frequency range is generally 0.1 mHz
10 Hz. 3. Signals may be generated in two forms: displacement and force. 4. The control channel is generally four-channel. 5. The control column (wheel) can be operated to have forward and backward straightline motion and clockwise and anticlockwise rotation and the pedal mechanism can be operated to have reciprocating straight-line motion simultaneously and the output stroke and force shall meet relevant control requirements. 6. The mechanical signal generator can output straight-line displacement and angular displacement, the bandwidth shall not be lower than 10 Hz when the amplitude fluctuates within 0.5 dB, the time for establishing step response shall not be greater than 10 ms, and the overshoot shall not be greater than 5%. 7. Displacement-control method. Static output displacement accuracy: # 6 0.25% F.S. 8. Force control method. Static output force accuracy: # 0.25% F.S. 9. Full-stroke accuracy of signal generator: # 0.3% F.S. 10. Dynamic load control accuracy: # 6 1.0% F.S. The mechanical displacement signal generator outputs the required displacement signal through setting the displacement (angular displacement, linear displacement) command or receiving an external control command after closed-loop control. The general form of the mechanical displacement signal generator is the hydraulic actuation servo system or electric actuation servo system. The composition principle of the mechanical displacement signal generator is shown in Fig. 6.11. As shown in Fig. 6.11, the multichannel servo controller is the control core of the mechanical displacement signal generator and shall have the following functions and performances: 1. The control of multiple control channels (generally four) shall be centralized in one control box as far as possible. To facilitate their use, the control box and external equipment shall be connected through quick release coupling and the control box shall adopt portable packaging. 2. Each control channel can achieve internal control and external control. The external control signal is input through A/D and function conversion can be conducted for external signals (function can be customized). 3. Displacement feedback control and force feedback control can be realized and the displacement feedback and force feedback control can achieve undisturbed switching.

“Iron bird” integration test of the flight control system 387

Figure 6.11 Composition principle of mechanical displacement signal generator.

4. In addition to realizing remote control, each control channel can also realize close control for the installation and debugging of the mechanical signal generator. 5. Each control channel can have null adjustment with software or manually. 6. It shall be able to provide waveforms such as sinusoidal wave, square wave, triangular wave, trapezoidal wave, arbitrary wave, frequency sweeping signal, and customized signal spectrum. 7. When the force control mode is adopted, the phase output of each hydraulic actuator cylinder between groups can be set according to the test requirements when the hydraulic actuator cylinders work in different groups, and the phase within a group is the same. 8. The feedback sensor shall provide the bridge voltage power supply to ensure control accuracy. 9. For hydraulic displacement signal generator, valve balance and valve flutter can be adjusted by software. 10. The controller shall have a fault self-diagnosis and warning function. 11. The controller shall set overload protection, amplitude protection, and difference protection functions. a. Overload protection: when the maximum value and minimum value of test load (or displacement) exceeds the preset value, the controller shall be able to stop for pressure relief treatment.

388 Chapter 6 b. Amplitude protection: when actual load and displacement amplitude exceeds the preset value, the controller shall be able to stop for pressure relief treatment. c. Difference protection: when the difference of synchronous loads exceeds the preset value, the controller shall be able to stop for pressure relief treatment. d. In the case of loss of control, the hydraulic actuator cylinder shall be able to safely return to the neutral position and lock. It shall be in the damping state during the return process to minimize the impact on the controlled object. e. The controller shall have accidental power failure protection function. In case of an accidental power failure, the equipment shall be in a damping state. 12. The controller software shall be able to modify the sensor coefficients and polarity and of nonlinear compensation. 13. The controller software shall be able to modify the servo valve control polarity. 14. In the normal working process of the equipment, the start of loading and unloading, stop, and emergency unloading can be interfered manually at any time. During the flight of the aircraft, the control plane bears the aerodynamic load, which is reflected in the hinge moment on the actuator of the flight control system. The actuator overcomes the hinge moment on the control plane and causes a deflection of the control plane, thus forming the aircraft control moment and causing a corresponding attitude change of the aircraft. The hinge moment directly affects the dynamic (static) performance (control accuracy, output speed, response time, frequency bandwidth) of the actuator and then affects the performance and flying quality of the flight control system. Therefore in the “iron bird” integration test of the flight control system, a set (probably multichannel) of the control plane aerodynamic load simulation system is required to realize the simulated application of the hinge moment of the actuator related to aircraft status (air speed, altitude, control plane deflection, and flight attitude, etc.). The control plane aerodynamic load simulation system shall have following main functions and performance: 1. It shall be able to simulate the hinge moment on all control planes of the aircraft on the “iron bird” integrated test bed of the flight control system. The form can be determined by the designer according to the sensitivity of actuator performance to the load, aerodynamic load spectrum, and cost, but the simulation of the hinge moment of the control plane must be achieved. 2. The two ways position control and force control shall be realized. It shall be ensured that the force control and position control can realize smooth transition and the modal switch can be realized. The position control aims to facilitate system installation and debugging. 3. The coordinated loading between any channels can be flexibly configured and independent single channel loading can also be realized.

“Iron bird” integration test of the flight control system 389 4. It shall have the electrical interface connecting test equipment such as the flight simulation system as well as the mechanical interface connecting the aircraft structure and test bed support structure. 5. When the input signal is 10% of the maximum output force, the bandwidth is no less than 8 Hz and the surplus force is less than 5% of the output force. 6. It shall have the function of automatic input and manual input of the load spectrum and the function can be set at will to calculate the load spectrum and realize the aerodynamic load simulation function. It is used for the debugging and maintenance of the load simulation system. 7. It can achieve table lookup function and define load in a simple and effective way. In other words, according to certain flight parameters and the given table, the simulation of the aerodynamic load is carried out according to the aircraft motion state table lookup results. 8. The equipment shall be grounded reliably, with high antiinterference ability and stable and reliable operation. No continuous oscillation of any amplitude is allowed. 9. It shall have a good man
machine interface to facilitate the operation and use by test personnel and the setting of initial conditions. It shall also have a real-time monitoring interface to accept remote control, receive the “Prepare,” “Start,” “Reset,” “Stop,” and “Emergency stop” control commands from the integrated test management system, and respond and give response information in a timely manner. 10. It shall have self-detection function with high coverage and multiple safety measures to ensure the safety of test personnel and test facilities. During the test, if there is any abnormality or overload, it shall be able to alarm automatically and turn off the control plane aerodynamic load simulation system in a timely manner. 11. The control plane aerodynamic load simulation system generally adopts a hydraulic servo loading method and connection structures such as hydraulic actuator cylinder, and the “iron bird” test bed and control surface shall have sufficient strength and stiffness to facilitate the installation, disassembly, and maintenance. The clearance between the hydraulic actuator cylinder and the connection structure shall be strictly controlled and well lubricated. It is recommended that the lugs at both ends of the actuator cylinder be connected with the connecting shaft through self-aligning bearings to ensure the clearance is within the acceptable range. The control plane aerodynamic load is distributed load. As the simulation of distributed load is difficult and costly and there is no difference between centralized load and distributed load for the performance of the actuator, centralized load simulation is generally adopted. The control plane aerodynamic load simulation system is a typical passive servo load simulation system and the commonly used schemes include aerodynamic load simulation, electrohydraulic servo loading simulation, and electric servo loading simulation. Most of them adopt the electrohydraulic servo loading simulation system.

390 Chapter 6 The output load command of the electrohydraulic servo load system formed according to the flight state (flight speed, flight altitude, aircraft attitude) of the aircraft and through the control plane deflection resolution is applied to the actuator of the flight control system through the controller of the electrohydraulic servo load system and the loading actuator, and the closed-loop control of the load is realized through the force sensor. The electrohydraulic servo aerodynamic loading system is mainly composed of controller (including hardware and software), electrohydraulic servo actuator, hydraulic oil source, control cabinet, and pipeline. The electrohydraulic servo actuator is composed of the electrohydraulic servo valve, hydraulic actuator cylinder, connector, hydraulic protection module, force sensor, displacement sensor, and damper, etc. The principle and composition of the electrohydraulic servo loading system are shown in Fig. 6.12. The control plane aerodynamic load simulation system generally adopts all-digital distributed mode and multimode integrated control mode. The control computer acquires aircraft motion parameters of the flight simulation system such as flight altitude, airspeed, and control plane deflection angle through a real-time network. The calculated control plane load force is taken as a force control command signal. Meanwhile, the displacement and force feedback signals

Figure 6.12 Composition and principle of aerodynamic load simulation system.

“Iron bird” integration test of the flight control system 391 are acquired in real time as control compensation parameters. After the synthesis by the control computer, control law calculation, and servo amplification, commands are sent to the electrohydraulic servo valve and the loading actuator is controlled to track the force in real time. Then, the system follows the motion of aircraft control plane on the “iron bird” test bed so as to complete the simulation of aerodynamic load on the control plane of the aircraft. The control plane aerodynamic load simulation system is mainly divided into the following two levels: The first level is the test management level with the monitoring computer as the core. It realizes test resource management, sets test parameters and servo controller parameters, and has the real-time man
machine interaction function. The second level is the real-time control level with VME slot 0 controller (embedded computer) as the core. It mainly includes the slot 0 controller module, analog signal processor module (including signal conditioning, servo valve driver amplification and other functions), real-time network card, signal adapter board, and model simulator module. This level completes system status monitoring and receives the loading commands transmitted by the simulator, obtains waveform data and feedback data, and transmits them to the monitoring computer for graphic display, provides valve current, provides excitation voltage to the load signal sensor, and amplifies the feedback voltage. The second level and the first level transmit data over the Internet.

6.2.11 Ground hydraulic energy and ground power supply In the “iron bird” integration test of the flight control system, the hydraulic system and power supply system of the aircraft shall provide the hydraulic energy source and power supply. In the preparation for test and control plane loop debugging, the ground power supply and ground hydraulic source are generally used to provide power and hydraulic energy for the flight control system. This is a very effective arrangement because the aircraft hydraulic system and power system have not completed relevant tests and their technical states have not been frozen yet at this stage. In addition, the mechanical displacement signal generator and the control plane aerodynamic load simulation system also need a ground hydraulic source or ground power supply to provide driving energy. 6.2.11.1 Ground hydraulic energy Ground hydraulic energy is also known as the ground hydraulic pump station and its working parameters shall be consistent with the airborne hydraulic energy parameters. The hydraulic energy demands of the ground test equipment shall also be taken into account. Specific design requirements and parameters of the ground hydraulic source shall be determined according to the requirements above.

392 Chapter 6 The ground hydraulic energy for the “iron bird” integration test of the flight control system is generally composed of the control system, hydraulic system, hydraulic pipeline, oil distribution cabinet, and auxiliary device. The main technical requirements and parameters of the typical ground hydraulic energy are as follows. 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20.

Rated working pressure: 28 MPa. Maximum working pressure: 31.5 MPa. Maximum pressure borne by the pipeline: 42 MPa. Oil return pressure: 0
1.0 MPa. Pressure pulsation: not exceeding 6 5% of rated working pressure. Maximum pressure full flow: 8 3 120 L/min (configured according to actual requirements). Working medium: YH-15 aviation hydraulic oil. Pollution degree of solid particles in oil: Level GJB420A-6. Working temperature of oil: # 80 C. Safety valve opening pressure: 31.5 MPa (adjustable). Pressure regulation way: 0.5
28 MPa proportional pressure regulation. Oil supply performance: constant pressure variable. External power supply: AC 380 V, 50 Hz, three-phase four-wire. Maximum power: 8 3 75 kW. Maximum speed of oil pump (motor): 1600 r/min (frequency control). Oil pump type: constant pressure variable plunger pump. Control mode: PLC, man
machine interface, frequency control. System heat dissipation mode: water cooling, air cooling. Noise: {70 dB. Continuous working time: $ 16 h.

6.2.11.2 Ground power supply system Ground power supply includes power supply for airborne equipment and power supply for test equipment. The airborne power supply shall be configured according to the technical requirements for power use and power consumption of the airborne equipment. The airborne power supply system may be 115 V, 400 Hz AC power, and 28 V DC power (see Fig. 6.13). The ground power supply shall be configured according to the actual power consumption of the test equipment and sufficient allowance shall be reserved. The power for test instruments and equipment shall be stabilized and the voltage fluctuation shall not exceed 6 2.5%. The power types for the test equipment generally include AC 220 V, 50 Hz; AC 115 V, 400 Hz; AC 7 V, 1800 Hz; DC 28 V, DC 6 15 V; and UPS voltage-stabilized power source.

“Iron bird” integration test of the flight control system 393

Figure 6.13 Typical ground AC power supply.

The quality of ground AC power supply shall meet the requirements of GJB 572—2005. The output voltage, frequency, and DC power supply voltage of medium-frequency power supply (AC 115 V, 400 Hz) shall be adjustable within a certain range. The power supply shall have convenient operation and digital display panels, have perfect fault warning and system protection functions, and have good electromagnetic compatibility. It shall be ensured that it can work normally in a laboratory environment and will not affect the normal operation of other equipment. The power supply system shall operate stably and reliably with low noise and meet the environmental requirements of the test site. The power supply shall be operated simply and used and maintained conveniently. According to GJB 572A—2005, the following requirements are put forward for the ground AC power supply: 1. Under rated working condition, the power input power factor shall not be less than 0.9 and the efficiency shall not be less than 85%. 2. With regard to the load/time characteristic of power overload, it is minimum 5 min for 125% rated capacity and it is minimum 10 s for 150% rated capacity. 3. For voltage waveforms with a distortion factor no more than 0.1, the voltage sensitive adjustment function shall be normal. 4. The steady-state voltage shall be within the range of 113
118 V. 5. The maximum difference between the three-phase voltages shall not be greater than 2 V. 6. The voltage phase difference shall be in the range of 118
122 . 7. Voltage waveform parameters shall meet the following requirements: peak coefficient 1.41 6 0.07, the distortion coefficient not greater than 0.04, and the DC component within
0.1 to 0.1 V. 8. The voltage modulation amplitude shall not be greater than 2.5 V (root-mean-square value).

394 Chapter 6 9. The steady-state frequency shall be kept within the range of 395
405 Hz. 10. The frequency modulation amplitude shall be within 4 Hz. 11. All switches and devices used to cut off the main loop in a fault state shall be reliably disconnected without human intervention. 12. The noise pressure level of the static variable power at 2 m away from the outer contour shall not be greater than 70 dB(A), and the noise pressure level of the unit power at 7 m away from the outer contour shall not be greater than 85 dB(A). Fig. 6.13 shows a typical ground AC power supply (CIF-4530M3P) and its main technical indexes are as follows: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19.

Input power supply: AC 3-phase 4-wire 1 GND, 220 V 6 10%, 50 Hz 6 10%. Output frequency: 350
650 Hz continuously adjustable. Frequency minimum resolution: 0.1 Hz. Frequency stability: 0.01%. Output voltage: 0
150 V(L-N) continuously adjustable. Voltage minimum resolution: 0.1 V. Voltage source stability: 0.1%. Voltage load stability: 0.5%(L-N), 1%(L-L). Transient characteristics of output voltage: 2 ms (resistance load 10%
90%). Rated output current: 125 A per phase, being able to withstand 2.5 times of surge current. Transient starting current for motor, addition of power amplifier and modulation of related parts. Load power factor: 0.5
1. Three-phase phase angle shift: # 1 (load balance), # 2 (100% load unbalance). Total harmonic distortion: # 2% (resistive load, THD-R measurement). Protection functions: Overload, short circuit, overtemperature protection. Operating environment: 0 C
40 C, 10%
90% RH (noncondensing). Cooling system: Air cooling. Insulation resistance: Input to housing 20 MΩ 500 V(DC), output to housing 20 MΩ 500 V(DC). Voltage withstand capacity: Input to housing 1000 VAC 1 min, input to output 1500 VAC 1 min.

6.2.12 Comprehensive test management system The test integrated management system is the core equipment to ensure the coordinated operation of all test devices in the “iron bird” integration test of the flight control system. The “iron bird” integration test of the flight control system of large transport aircraft involves almost all aircraft systems, more than 1000 pieces of airborne equipment, more

“Iron bird” integration test of the flight control system 395 than 3000 test pieces, and nearly 100 sets of test equipment. Aircraft airborne systems shall operate normally and test support equipment with different functions shall be managed. Therefore the operation of equipment shall be effectively managed and the testing shall be organized to ensure that the test is carried out safely and smoothly through the coordination and management of airborne systems for the “iron bird” integration test of the flight control system and test support equipment. In addition, it is necessary to realize state monitoring and fault warning of key and sensitive parts and provide a good supporting environment for test personnel to observe and monitor various test conditions, so as to organize and implement various tests flexibly, quickly, and conveniently. Therefore an integrated management system for the “iron bird” integration test of the flight control system should be established to ensure that the test is reasonably, safely, and efficiently organized with a standardized process. The test integrated management system shall have the following functions: 1. It is capable of the digital cooperative control between test support equipment and airborne equipment, be able to determine test status, formulate test process, control test operation, realize the sequential control over the power connection and power disconnection for the test system and support equipment, as well as manage the power supply status of the monitoring system and support equipment in real time. 2. It shall manage test files, test status, test parameters, and test results for the convenience of query, analysis, browsing, and output of test results. 3. It shall have a complete interface adapter for the mechanical and electrical connection with the flight control system to achieve the connection and communication with the airborne system. 4. It shall have a complete, stable, and reliable signal isolation function to ensure the safety and accuracy of analog signal transmission with the flight control system. 5. It shall have video information and audio information monitoring and transmission functions and be able to monitor and transmit the working condition of key airborne equipment such as the control plane in real time and save the image information, so as to realize the monitoring, scheduling, and management of the test system and test support equipment. 6. It shall have complete state detection and information transmission functions and be configured with fieldbus communication interface, network communication interface, and interface adapter with complete functions. The test support equipment shall comprise a complete network system to realize the detection, transmission, and saving of information about the voltage supply switch state of actuators and other important and key operating states in the test, so as to realize the monitoring of the operating state of the test system and the test support equipment. 7. The data exchange and display processing system shall be capable of the real-time data transmission and exchange between test support equipment and airborne equipment, the

396 Chapter 6 display of test status, test progress, and test results, as well as on-site query and comparison of historical data, simulation data, and test data, and the output of test results. 8. The data exchange and display processing system adopts the test data exchange of a real-time network and is configured with the test data management function of a disk array. Based on informationized and digital management, it plans and arranges test items and content, and releases test items and test status information to test support equipment and the test site. The operation of the test system and test support equipment is also scheduled according to test items and test status. The data in the test process is saved in real time and the functions of real-time query and browsing of test status information are provided. 9. It shall be able to read the information flow parameters exchanged by the airborne system, including the data information transmitted within the flight control system, the data information transmitted between the flight control system and the avionics system, as well as the data information transmitted between the flight control system and other airborne systems. The composition structure of the integrated test integrated management system is shown in Fig. 6.14. Test execuon and operaonal control

Secure storage of test informaon

Data exchange and visualizaon

Data acquision

Operaonal control layer

Video surveillance scheduling layer Data exchange and transmission layer

Cockpit simulaon

Visual system Avionic instrument system

Avionics system simulator

Fly-by-wire system tester

Airborne data exchange and transmissioin

Automac flight control tester

Command simulaon control system

High lift system tester

Mechanical displacement signal generator

Control plane loading system

Power distribution management#1 AC

Power distribution management#2

Angle of aack turntable

Site1#acquisition node

Dynamic signal analysis system

Three-axis turntable

Site2#acquisition node

Flight simulation system

Linear acceleration turntable

Site3#acquisition node

Ground hydraulic energy system

Site4#acquisition node

Flight control test system

Power distribution management#3

22220V/50Hz Site data bus

Control and power signal line

Video signal transmission line

Power distribution management#4

AC

380V/50Hz

Device operaon control layer

Power distribution management#5 Power distribuon control Internet signal line layer

Figure 6.14 Architecture of “Iron Bird” integration test control and management system.

“Iron bird” integration test of the flight control system 397

6.2.13 Aircraft motion sensor driver The flight control system generally uses an angular rate gyro assembly to test the angular rate of airframe motion, uses an overload sensor to test the normal overload and lateral overload of the aircraft, uses an angle of attack sensor to test the angle of attack of the aircraft, uses a vertical gyro or inertial navigation platform to test the attitude angle and yaw angle of the aircraft, and uses a dynamic (static) pressure sensor to test the total pressure and static pressure given by the airspeed head of the aircraft. In the “iron bird” integration test of the flight control system, as only the virtual flight of the aircraft can be performed and its motion is obtained through the calculation of the flight simulation system and physical motion of the aircraft fails to drive the aircraft motion sensor to output signal, the sensor must be installed on the corresponding driver (turntable) and the aircraft motion parameters calculated with the flight simulation system shall be taken as the driving command. The output shall conform to the sensor parameters defined by the interface of the flight control system. 6.2.13.1 Single-axis rate turntable Single-axis rate turntable is a necessary device for the “iron bird” integration test of the flight control system. In the open-loop test of the flight control system, a single-axis rate turntable is used to complete the polarity and transmission ratio inspection and angular rate gyro assembly characteristic inspection of the angular rate channel. In the closed-loop test of the flight control system, the angular rate gyro assembly is driven according to the angular rate command calculated with the flight simulation computer and the angular rate gyro assembly generates angular rate signals and sends them to the flight control computer to achieve the closed-loop control of the angular rate channel of the flight control system. The composition principle of the single-axis rate turntable is shown in Fig. 6.15. Requirements for the performance indexes of the single-axis rate turntable are as follows: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11.

Loading weight: . 20 kg. Table diameter: $ 450 mm. Table runout: {0.01 mm. Angle range: continuous infinite. Working mode: rate mode, position mode, position simulation, rate simulation, standard function (sine). Position resolution: , 0.36v. Angular position accuracy of position: { 6 5v. Angular position repeatability: { 6 2v. Rate range: 0.001
300 /s. Rate resolution: 0.0001 /s. Angular rate accuracy and rate stability: lower than 0.001% within 360 .

398 Chapter 6

DSP moon control module

CPU

Interface and buffer module

Bus extension module

Cycle measurement module

IPC control computer

High-speed serial interface DC 220V 380V

Power unit

AC

Analog interface

Signal interface adapter unit

Display device

Single-axis simulation turntable driving amplifier

Communicaon interface

Single-axis simulation turntable moon structure

User simulator

Figure 6.15 Composition principle of single-axis rate turntable.

12. Frequency response: 10 Hz, (with load, input signal amplitude 1 , 10% amplitude attenuation, 10 phase lag). 13. Number of slip rings: 32 rings, including four power cables. 6.2.13.2 Linear acceleration turntable In the “iron bird” integration test of the flight control system, a linear acceleration turntable is used to complete the polarity and transmission ratio inspection of the acceleration channel and the inspection of the overload sensor (also called the accelerometer sometimes). In the “iron bird” integration test of the flight control system, the overload sensor installed on the linear acceleration turntable will feel the linear acceleration value given by the linear acceleration turntable and the value will be sent to the flight control computer in the form of electrical signals to realize the closed loop of the overload signal in the flight control system. The basic composition of the linear acceleration turntable is shown in Fig. 6.16 and the electric control principle is shown in Fig. 6.17. Requirements for the performance indexes of the linear acceleration turntable are as follows: 1. Linear acceleration range: 0.1
12 g (stepless adjustable). 2. Standard g value uncertainty: 1 3 1024.

“Iron bird” integration test of the flight control system 399

Figure 6.16 Basic composition of linear acceleration turntable.

Figure 6.17 Electric control principle of linear acceleration turntable.

400 Chapter 6 3. 4. 5. 6. 7. 8. 9. 10.

Number of servo turntables: 3. Load weight of servo turntable: 5 kg. Diameter of servo turntable: 200 mm. Working angle of servo turntable: 360 continuous. Positioning uncertainty of angular position of servo turntable: 6 0.002 . Resolution of angular position of servo turntable: 0.0001 . Maximum angular rate of servo turntable: 500 /s. Bandwidth of servo turntable: Input sinusoidal signal, amplitude 1 , bandwidth 10 Hz, amplitude attenuation 10%, phase lag 10 . 11. Servo turntable user slip ring: 30 rings /2 A, shielded. 12. Steady-speed table user slip ring: 90 rings /2 A, shielded. Besides, the linear acceleration turntable shall also have the following functions: 1. Simulation interface: 6 10 V analog interface, high-speed serial port, high-speed reflective memory board. 2. Fault detection and position function. 3. Sound-light alarm function in case of fault and abnormality. 4. Overspeed, overcurrent, and outage protection functions. 5. It shall have good resistance to electrical and magnetic interference, be able to normally work in the laboratory environment, and not affect the normal operation of other equipment. 6. Reliable turntable, be able to work continuously for no less than 10 h. 7. It shall have a safety cover to prevent installed objects flying out and doing harm to personnel or equipment during the operation of the turntable. 6.2.13.3 Three-axis flight simulation turntable The three-axis flight simulation turntable is used to simulate the angular motion of the aircraft in air motion. In the “iron bird” integration test of the flight control system, the three-axis angular rate gyro assembly (or inertial navigation system) installed on the three-axis flight simulation turntable will feel the aircraft motion angular rate given by the three-axis flight simulation turntable and it will output angular motion signals to the flight control computer to consist of a closed-loop control for the “iron bird” integration test of the flight control system. The composition principle of the three-axis flight simulation turntable is shown in Fig. 6.18. Requirements for the performance indexes of the three-axis flight simulation turntable are as follows: 1. 2. 3. 4.

Structural type of turntable: U-O-O frame. Loading weight: 50 kg. Loading size: 500 3 450 3 450 (W 3 L 3 H). Three-axis definition: Rolling (inner axis), heading (middle axis), pitching (outer axis).

“Iron bird” integration test of the flight control system 401

Figure 6.18 Composition principle of three-axis flight simulation turntable.

5. Angle range: 6 120 . 6. Bandwidth of control system (amplitude difference lower than 6 1 dB, rolling 12 Hz, heading 8 Hz, pitching 8 Hz. 10 phase shift): 7. Axis perpendicularity: 6 10v. 8. Axis orthogonality: # 0.25 mm. 9. Axis perpendicularity: , 5v. 10. Axis cross: , 0.25 mm. 11. Position accuracy: 6 5v. 12. Position resolution: 0.0001 . 13. Speed range: Inner axis 400 /s, middle axis 200 /s, outer axis 150 /s. 14. Speed resolution: 0.001 /s. 15. Maximum acceleration: Inner axis 3000 /s2, middle axis 1250 /s2, outer axis 800 /s2. 16. Rotation accuracy: 5v. 17. Signal interface: digital, analog, network card. 6.2.13.4 Total (static) pressure simulator The total and static pressure simulator, also known as a dynamic (static) pressure simulator, is a kind of electrical/gas conversion device that converts voltage signals into pressure

402 Chapter 6 signals in proportion. It mainly simulates the total pressure and static pressure in aircraft flight. The dynamic (static) pressure output by the dynamic (static) pressure simulator drives the airborne atmospheric data sensor and the sensor output signal is input to the flight control system to complete the parameters adjustment of the control law. The normal operation of the total (static) pressure simulator requires an air source of corresponding pressure and flow. The air supplied shall be dehumidified, filtered, and dedusted to ensure the air is dry and clean. In addition, a vacuumizing system composed of vacuum pumps of sufficient displacement shall be configured. The continuous pipeline between the vacuumizing system and the total (static) pressure simulator shall be well sealed and its performance indexes shall be higher than those of the airborne total (static) pressure sensor. The total (static) pressure simulator mainly consists of a digital controller, digital filter, direct driving servo valve, air path module, quick response sensor, high-precision power supply, real-time communication interface, curve and data display interface, and cabinet. The composition principle of total (static) pressure simulator is shown in Fig. 6.19. Requirements for the functions of the total (static) pressure simulator are as follows: 1. 2. 3. 4.

Dynamic or static mode can be selected. Externally sent altitude and speed signals are received through the real-time network. Altitude and speed input signals of different amplitudes and frequencies are set. The dynamic change process of pressure, altitude, and speed signals is displayed with a curve in real time. 5. The physical units and dimensions of pressure, altitude, and speed are selected online. 6. The limit values of altitude and altitude change rate and speed and speed change rate are set to prevent damage to the tested products.

Figure 6.19 Composition principle of total (static) pressure simulator.

“Iron bird” integration test of the flight control system 403 Requirements for the main technical indicators of the total (static) pressure simulator are as follows: 1. 2. 3. 4. 5. 6.

7. 8.

9. 10. 11.

Static pressure (ps): 10
1200 hPa. Accuracy: 6 0.01% F.S. Total pressure (pt): 10
3500 hPa. Accuracy: 6 0.01% F.S. Altitude (H):
1000 to 3000 m. Accuracy: 6 1.0 m (0 m). 6 2.5 m (9000 m). 6 8 m (18,000 m). Mach number (Ma): 0
5, allowable error 6 0.005. Airspeed: (V) 50
1800 km/h. Allowable error 6 5 km/h (50 km/h). 6 2.1 km/h (100 km/h). 6 0.15 km/h (1000 km/h). Service environment: Working temperature 0 C
40 C. Relative humidity , 85% RH. Power supply: 220 V 6 10%, power 500 W. Communication mode: RS232, 485 standard serial port, IEE488 standard parallel port, 100 M Internet, VMIC card.

6.3 Debugging and preparation for the flight control system “iron bird” integration test The “iron bird” integration test of the flight control system of large aircraft is the most important ground laboratory test that is to be carried out after the completion of the qualification test and subsystem combined test of airborne equipment, when the airborne equipment, subsystem, and control law status is frozen and after the software passes the test. Therefore all airborne equipment of the flight control system and relevant support system airborne equipment installed on the “iron bird” integrated test bed of the flight control system as well as installation forms shall be consistent with the technical state of the aircraft for the maiden flight. In other words, they shall be “S” type ground test pieces. The flight control system of large aircraft is composed of hundreds of pieces of airborne equipment (LRU) that are replaceable at the site, such as various control components of the pilot, control panels, control computers, actuators, transmission mechanisms, and sensors. The flight control system of the background aircraft is composed of more than 1000 LRUs.

404 Chapter 6 For the “iron bird” integration test of the flight control system of large aircraft, it is a huge project to complete the installation and debugging of the “iron bird” integration test bed. The main work includes: 1. Cockpit control system. Left (right)control column module, left (right) pedal module, left (right) loading mechanism module. 2. Fly-by-wire flight control system. Display control equipment (trim handle module, indicator, trim and control panel), fly-by-wire flight control computer PFC, ACE, aileron, elevator, rudder, horizontal stabilizer, spoiler, and other control plane actuators, sensors, etc. 3. Machinery control system. Steel cable rod driving system, tension regulator, variable ratio device, deflection limiting mechanism, mechanical control actuator shared with the fly-by-wire flight control system, and the hand and foot operating module shared with the cockpit control system. 4. Automatic flight control system. Automatic flight control panel (auto flight control unit, AFCU), AFCC. 5. High-lift system. FSECU, flaps (slats) power drive unit, transmission devices, actuators, brakes, and sensors. After the installation of the airborne equipment of the flight control system, the cables connecting the airborne equipment shall be laid and connected, and then the static adjustment and inspection of the system shall be carried out. At the same time, the test support systems such as hydraulic system, power system, avionics system, and test guarantee equipment such as testers, simulators, ground hydraulic sources, ground power supply, test analysis system, and control plane aerodynamic load simulation system shall complete installation and debugging successively. The installation and debugging of the tested systems, test support systems, and test equipment can be conducted after the construction of the “iron bird” integrated test bed of the flight control system or conducted alternately according to the engineering design. However, due to the requirements of the engineering schedule, the design of the “iron bird” integrated test bed shall be carried out synchronously with aircraft structure design, the manufacturing of the “iron bird” integrated test bed shall be carried out synchronously with aircraft manufacturing, and the installation and debugging of the “iron bird” integrated test bed shall be carried out synchronously with the aircraft final assembly and system installation and debugging. During this period, the installation of the “iron bird” integrated test bed requires massive tooling and fixtures, such as aero vehicles, cranes, mobile vehicles, formwork frames, platforms, laser equipment, jigs and tools, and as many as hundreds of people may participate in the work. Therefore in order to ensure safety, quality, and progress, organizational management is particularly important.

“Iron bird” integration test of the flight control system 405

6.3.1 Static adjustment and inspection of the flight control system 1. Installation inspection of tested system The static adjustment inspection of the flight control system is conducted according to the technical conditions for installation and electrified inspection of five subsystems, respectively, cockpit control system, machinery control system, fly-by-wire flight control system, high-lift system, and automatic flight control system. These two kinds of documents are mandatory documents sent to the manufacturer to confirm the technical status of the flight control system. After the installation, debugging, and testing of the flight control system, it is confirmed that it meets the requirements of the technical documents above. a. Technical conditions for installation. The technical conditions for installation describe the installation position, method, and requirements of the airborne equipment of the flight control system and specify the initial position (zero position) and limit the stroke state of the airborne equipment. They generally cover the control display, actuator, and sensor of cockpit control system, machinery control system, high-lift system, and fly-bywire flight control system, as well as the installation position, method, and requirements of the airborne equipment of the automatic flight control system. According to the technical conditions for installation of the five subsystems of the flight control system, the mechanical characteristics of each airborne equipment LRU of the system such as the installation position, initial state (zero position), motion stroke, and motion accuracy shall be checked to see if they meet the requirements of the technical conditions for installation. For airborne equipment not meeting the requirements, they shall be reinstalled and debugged until they meet the requirements in principle. For the airborne equipment that cannot satisfy technical conditions for installation for some reason (such as the limit deflection angle of the control plane cannot reach the maximum value requirement), they shall be reevaluated and the difference shall not affect the test results and the correct judgment of the test results. b. Technical conditions for electrified inspection. The technical conditions for electrified inspection describe the methods and requirements for checking the electrical characteristics of various electronic and electrical airborne equipment of the flight control system and specify the requirements for the electrical characteristics, cable conduction characteristics (connected, short circuit, open circuit), insulation characteristics (correct grounding and bonding, inter-line), and load impedance of sensors of electronic and mechanical and electrical airborne equipment. According to the technical conditions for electrified inspection of the five subsystems of the flight control system, the electrical characteristics and cables of

406 Chapter 6 each electronic and mechanical and electrical airborne equipment of the system such as the cable conduction characteristics (connected, short circuit, open circuit), insulation characteristics (correct grounding and bonding, inter-line), and load impedance of sensors shall be checked to see if they meet the requirements of technical conditions for electrified inspection. For airborne equipment not meeting the requirements, the causes shall be identified until they meet the requirements in principle. For the airborne equipment that cannot satisfy the requirements for some reason (e.g., the impedance of signal transmission line is slightly lower than the requirement), they shall be reevaluated and the difference shall not affect the test results and the correct judgment of the test results. 2. Control-display function inspection of flight control system The inspection and validation of the control-display function of the flight control system refers to the inspection of the “man
machine” interface functions such as the setting, conversion, control logic, and display of the system working state under the condition that the open-loop technical state of the flight control system is established. The control-display function inspection starts from manual operation control (switch pull and button press), realizes relevant control and conversion response through the system (check the entry, exit and state conversion of working state or the change of truth values of relevant discrete quantity), and is completed through the display and reporting of the system (display of change of image, status, parameters, and watch/ listen by personnel). The inspection items also include the internal working status and logic transformation of the system which are not externally reported. For these changes, the changes of related variables in the system shall be checked to see if they meet the design requirements after external conditions are set. Besides, although the user (pilot and ground crew) has been consulted on the layout of the control-display airborne equipment, the arrangement of the control mechanism and the display mode etc. at the beginning of the design, relevant personnel shall also be invited to conduct on-site operation and observation in the system integration stage to further seek for their evaluation opinions. Please refer to the “iron bird” integration test of the flight control system for the content and method of the control-display function inspection of the flight control system. 3. Open-loop static characteristic test of flight control system The open-loop static characteristic test of the flight control system shall follow the principle of gradual development from the core to the periphery and from the local to the whole system and be completed by signal chains or subsystems and systems. The open-loop static characteristic test of the flight control system, on the basis of the performance test of airborne equipment, tests the power supply, power consumption, oil supply, signal polarity, transmission ratio, on/off logic, and other items of the whole

“Iron bird” integration test of the flight control system 407 system in the test environment. Tests of these items should be considered from the top level, and the test items, methods, and precautions are as follows: a. The power supply and power consumption test shall be carried out step by step according to the power supply sequence of each airborne equipment and the flight control computer with multiple plug-in boards shall also be connected in sequence to check the power supply. Protective measures such as fuses must be installed in the circuit to prevent fault expansion. Before the power supply and power consumption test, the voltage of the cable shall be tested first and then the power shall be cut off and connected to the airborne equipment under test. The grounding characteristic inspection is particularly important. It is required to check the connection of the ground wire, check whether the grounding point is correct, whether the isolated ground wire meets the requirements, and whether the ground wire of the test equipment is consistent with the design requirements. In particular, not only the ground wire shall be checked before the power supply but also the grounding condition shall be checked in a timely manner before new equipment is introduced. b. In the signal transmission ratio inspection, the corresponding relationship between mechanical zero position and electrical zero position shall be noted and attention shall be paid to see if the control plane deflection angle is calculated based on airflow direction or spindle direction. The transmission ratio of some signals changes with the flight state. For information chains with a high pass network or integral network, the test problem of the transmission ratio shall be solved. In the transmission ratio inspection, all branches of the chain shall be covered, including the branch of software. When one branch is tested, other branches shall be disconnected or grounded. c. In the signal polarity inspection, the polarity of each segment of the signal chain shall be strictly tested according to regulations and the polarity shall not be changed at will. This point shall be especially ensured for a system with software, or flight safety will be endangered. Meanwhile, attention should be paid to the polarity of the introduced equipment and the installation direction of the product on the equipment. After the static adjustment inspection of the flight control system, the static debugging of the flight control system shall be carried out. The static debugging is just the open-loop integration and testing of the system in essence. With the support of the test equipment, the airborne equipment and relevant software after passing design, processing, debugging, and performance test are well-connected into a complete openloop signal chain conforming to the design technical state according to the provisions of the interface control document. The integration of the “iron bird” test bed of the flight control system aims to constitute a flight control system that conforms to top-level design requirements and

408 Chapter 6 system design specifications. After necessary installation and debugging, the “iron bird” test bed for the flight control system creates the basic conditions for the closed-loop system test and verification of the “flight control system-aircraft” and makes preparation of the tested system. The first test bed integration of the flight control system mainly aims to identify problems and potential dangers and solutions to the problem.

6.3.2 Debugging and technical status of the cross-linking system The flight control system is cross-linked with the aircraft avionics system, hydraulic system, landing gear control system, power supply system, and engine control system, etc. through mechanical, hydraulic, electrical, and data buses to jointly support the completion of functions and performance of the aircraft’s flight control system. Therefore, before the “iron bird” integration test of the flight control system, it is necessary to check and confirm whether the interface between the flight control system and the system cross-linked with it on the “iron bird” test bed conforms to the interface control document, whether the cross-linked system completes installation and debugging, and whether the technical state of the flight control system and its subsystems meet the design requirements or design specifications. 1. Interface inspection The following three requirements shall be satisfied for the interface inspection. First, the cross-linked system participating in the “iron bird” integration test of the flight control system must be specific, that is, the avionics system, hydraulic system, power supply system, landing gear control system, and engine control system. The object of interface inspection must be airborne equipment eligible in a technical state of the aircraft for maiden flight. Second, cross-linked systems that do not participate in the “iron bird” integration test of the flight control system must have physical
physical interface inspection. These systems generally include the fuel system (including fuel measurement and management system) and environmental control system (including deicing and antiicing system). Third, for systems that cannot have physical
physical interface inspection, a cross-linked equipment simulator can be used to have interface inspection and the inspection shall cover the interface relationship and static and dynamic characteristic testing. a. Mechanical interface inspection Through on-site assembly and connection, the mechanical connection form, dimensions, and error prevention design between airborne equipment shall be checked to ensure the airborne equipment are connected securely and reliably and meet the requirements of the interface definition document. b. Hydraulic interface inspection Through on-site assembly of the connection of the pressure supply, the connection form, dimensions, pressure system, liquid flow direction, and error

“Iron bird” integration test of the flight control system 409 prevention design of the hydraulic interface between airborne equipment shall be checked to ensure the connection is reliable, the pressure system is consistent, there is no leakage, and the requirements of the interface definition document are met. Meanwhile, the effects of hydraulic characteristics on the flight control system shall be checked, such as the effects of hydraulic source fluctuation characteristics, hydraulic source switching characteristics, and hydraulic pulsation. c. Electrical interface inspection Through on-site assembly of the connection of the electricity supply, the connection form, dimensions, pressure system, liquid flow direction, and error prevention design of the electrical interface between airborne equipment shall be checked to ensure the connection is reliable, the pressure system is consistent and the requirements of interface definition document are met. Meanwhile, the effects of aircraft power characteristics on the flight control system shall be checked, such as the effects of power supply fluctuation characteristic and counter electromotive force, power supply switching characteristics, and AC component of DC power supply. d. Data bus interface inspection Through on-site assembly and connection, the data bus form, impedance matching, signal transmission ratio and signal polarity, control logic, and rate coordination, etc. shall be checked to ensure the form, polarity, and logic are correct and the requirements of the interface definition document are met. Meanwhile, if the attitude system participates in the “iron bird” integration test of the flight control system, the polarity and load characteristic of the turntable shall be checked. 2. Debugging of cross-linked system and requirements for its technical state a. Debugging of hydraulic system and requirements for its technical state The hydraulic system configuration that can ensure the safety of the flight control system of large transport aircraft generally requires at least three sets of independent hydraulic sources. Therefore the “iron bird” integrated test bed of the flight control system shall have three sets of hydraulic sources that meet the requirements of the flight control system and they shall meet the requirements of technical conditions for the installation and design specification of the hydraulic system after installation and debugging. To advance the preparation of the test, concurrent engineering is often adopted to ensure the simultaneous test and debugging. Therefore three sets of independent ground hydraulic sources corresponding to aircraft hydraulic sources shall be configured to provide hydraulic sources for flight control system debugging and preliminary combined test, and the interchangeability and maintenance convenience of the two hydraulic systems shall be ensured. b. Debugging of landing gear control system and requirements for its technical state As the aircraft bears the aerodynamic force, engine thrust, and gravity in air flight and bears the ground reaction force from the landing gear in addition to the

410 Chapter 6 forces above in high-speed taxiing, it requires that the flight control system must have both air and ground control modes to satisfy the aircraft control in two stress environments. The air configuration of the control law may be control augmentation configuration and the ground configuration may be direct chain configuration and there may also be a directional control mode that controls whether or not the cross-link is disconnected. It is a common method for the flight control system to judge if the aircraft is in the air state or ground state by extracting the landing gear control system wheel load and wheel speed and landing gear control signals as the conditions. As the landing gear control system is an important user of the hydraulic system, confirming the landing gear control system meets the design requirements is also an important part to assess the characteristics of the hydraulic system. Therefore the landing gear control system and hydraulic system are often established on the same test bed for the integration test. In other words, the “iron bird” integration test bed of the flight control system also contains a complete landing gear control system. Landing gear control system debugging aims to check whether the interface between the landing gear control system and other airborne systems on the test bed is correct, whether the landing gear control, turning, antiskid, and brake control functions and performance meet the requirements of design specifications, and whether the indication and alarm and signal transmission are correct. c. Debugging of avionics system and requirements for its technical state The flight control system completes aircraft flight according to the given task, needs to receive the task commands (such as horizontal navigation and vertical navigation) given from the avionics system, receives the aircraft and flight environment information (such as flight speed and flight Mach number) provided by the avionics systems, and sends flight control system state and fault information [such as flaps (slats) position and control plane deflection angle] to the avionics system. That is to say, the basic elements and commands, control and feedback related to the work of the flight control system are all related to the avionics systems. Therefore the “iron bird” integration test of the flight control system must include the aircraft avionics system. The cross-linking test of the two systems is also an important test item. The avionics system participating in the “iron bird” integration test of the flight control system includes the atmospheric data system, inertial navigation system, radio altimeter, integrated processing system, RDC, central warning system, central maintenance system, flight management system, display processing system, primary flight display, multifunctional display, multifunctional keyboard, trackball, display control panel, and engine indication and unit warning system. The atmospheric data system provides the flight control system with atmospheric data of the aircraft, the inertial navigation system provides aircraft attitude and angular rate signals, the

“Iron bird” integration test of the flight control system 411 integrated processing system provides aircraft weight and center of gravity position signals, the primary flight display, and the multifunctional display and engine indication and unit warning system displays the flight control system status information, warning information, control plane position and maintenance information, and aircraft trim position information. Therefore the debugging of the avionics system on the “iron bird” integration test bed of the flight control system is also an important part of test debugging and preparation, including interface inspection, transmission real-time performance inspection and performance test and it should be conducted to confirm if the system meets the requirements of the design specifications and interface definition document. Besides, to advance test preparation work quickly, concurrent engineering is often adopted to ensure test and debugging are carried out simultaneously. As a result, a simulation system corresponding to the aircraft avionics system should be configured to support the debugging and preliminary combined test of the flight control system and the interchangeability and maintenance convenience of the aircraft system and simulation system shall be ensured. d. Debugging and technical state of other systems Other aircraft systems participating in the “iron bird” integration test of the flight control system may also include the power supply system, hydraulic system, antiicing and deicing control system, and thrust control system. Power supply system: electronic and mechanical and electrical airborne equipment of the flight control system require the aircraft power supply system to provide sufficient redundant power and some electrical airborne equipment such as sensors also require the flight control system to convert the aircraft power into required secondary power supply. For example, the power supply system of an aircraft provides AC power to flaps (slats) drive unit, DC power to PFC, ACE, AFCC, and FSECU, aileron trim mechanism, rudder trim mechanism, etc., and ensures the three-redundancy power (including normal DC bus bar, uninterrupted bus bar, and flight control battery bus bar) provided to PFC and ACE. Before the “iron bird” integration test of the flight control system, the interface between the flight control system and the power supply system (possibly integrated in the mechanical and electrical management system) shall be verified and the performance of the flight control system when the power supply system is under normal and emergency conditions shall be checked. Considering the convenience of aircraft system integration, it is suggested to establish the power supply system test bed near the “iron bird” integrated test bed of the flight control system, so as to carry out aircraft level system integration and support the flight control system to complete the “iron bird” integration test. If restrained by test conditions or in the debugging stage of the flight control system test, ground power supply with the same redundancy and characteristics can be used to guarantee the test. However, the flight control system test items under power failure states and the cross-linking test between

412 Chapter 6 the flight control system and power supply system shall use the complete real power supply system on the aircraft as the working power supply of the flight control system. The hydraulic system, landing gear control system, avionics system, and other crosslinking airborne equipment are all test pieces (“S” type ground test pieces) with states consistent with the aircraft for the maiden flight and they shall have passed the delivery acceptance of the producer and manufacturer and be issued with the relevant conformity certificate or record and acceptance certificate. All simulation exciters shall meet the design requirements and have passed the delivery acceptance of the producer and manufacturer and be issued with the relevant conformity certificate and acceptance certificate. Antiicing and deicing control system: it shall be judged whether the aircraft is in an icing state and the state affects the stall characteristic of the aircraft. The flight control system with stall protection function needs to receive the state information of the antiicing and deicing control system and the icing state information of aircraft wing and empennage so as to complete the stall protection of the aircraft in the icing state. When the “iron bird” integration test of the flight control system is performed, the interface and communication between the flight control system and the antiicing and deicing control system (possibly integrated in the mechanical and electrical management system) shall be verified and the exciter of the antiicing and deicing control system shall be used to provide cross-linking signals for the flight control system to verify the stall protection function and performance of the control law of the flight control system. Thrust control system: the realization of functions including thrust/speed control, asymmetric compensation control of thrust, opening of ground taxiing spoiler, (horizontal and vertical) navigation control, and autothrottle control of the flight control system requires the thrust control system to provide signals including engine speed of sufficient redundancy, pressure ratio, reverse thrust, throttle position, and autothrottle cutoff. Additional engine signals may be provided for the thrust control system with FADEC to realize the optimized thrust extraction. When the “iron bird” integration test of the flight control system is performed, the interface and communication between the flight control system and the thrust control system (possibly integrated in the mechanical and electrical management system) shall be verified and the exciter of the thrust control system shall be used to provide cross-linking signals for the flight control system to verify relevant functions and performance of the control law of the flight control system.

6.3.3 Potential problems in the flight control system debugging process and cause analysis Although the flight control system has experienced the validation process of the system, subsystem, control law, software requirements analysis and design requirements, formed flight control system design specification, subsystem design requirements, control law design requirements, software design requirements, airborne equipment design requirements

“Iron bird” integration test of the flight control system 413 and technical agreement and undergone airborne equipment manufacturing and testing, control law design and engineering simulator verification, software coding and testing, and subsystem combined test, unexpected technical or engineering errors or inconsistencies are still inevitable in the debugging and preparation process for the “iron bird” integration test of the flight control system. The potential problems and causes are as follows: 1. Function loss or incorrect function of airborne equipment The function of airborne equipment is to support the system’s function and performance. One of the main achievements in the system requirements definition stage is to determine the function of airborne equipment. The incomplete functioning of airborne equipment may be caused by following reasons. a. The design requirements are incomplete, inaccurate, and not strict. Due to the lack of in-depth work in the requirements definition stage, poor consideration of influencing factors, and the fuzzy and even ambiguous design boundary, the design requirements of airborne equipment lack content or cause misunderstandings. b. The designers of airborne equipment misunderstand the design requirements. The designers of airborne equipment rely excessively on their previous experience, fail to read the airborne equipment design requirements and technical agreement carefully, misunderstand the design requirements, and carry out the airborne equipment design and manufacturing smugly. c. Technical coordination in the design process is incomplete. During the design and development stage of airborne equipment, the design requirements changes caused by various reasons are not implemented and the design, manufacturing, and process problems related to functions (performance) of airborne equipment exposed are not fed back to the host supervisor in time but are handled at will. d. The functions of airborne equipment exposed in the debugging are adjusted (cancelled, added, or adjusted). 2. Interface error or inconsistency The mechanical, electrical, hydraulic, and data bus interfaces between airborne equipment are the most common problems in the debugging process, such as the installation failure or failure of installation in position, polarity reversal (reversed connection of signal high and low ends, inverse phase of positive and negative terminals), and truth value error (level definition and logic circuit dislocation) of the airborne equipment caused by mechanical interface error. The interface error may be caused by the incorrect writing of the interface control document, the wrong understanding by the airborne equipment designers of the interface, or the designers’ inexperience. 3. Performance index of airborne equipment out of tolerance The airborne equipment or its signal processing performance does not meet design requirements; the index is not realized; the accuracy does not meet requirements; the

414 Chapter 6 performance is not stable or not compatible with actual load conditions; all this would cause abnormal performance and index out of tolerance. 4. Connection line error The insufficient understanding of the airborne equipment by system designers causes line form error, line redundancy error, or line parameter error. 5. High airborne equipment fault rate Due to transportation, installation, or airborne equipment reasons, the airborne equipment may break down after abnormal connection to power or work, or after liquid leakage or malfunction. The debugging for the “iron bird” integration test of the flight control system is an important and time-consuming work. Not only should the flight control system and other airborne systems be preliminarily confirmed to meet the requirements of relevant design specifications and design requirements and it should be confirmed whether the airborne equipment meets the design requirements and requirements of the technical agreement, but also the inconsistency of the system and defects of the airborne equipment shall be identified. Before this, all possible inconsistency shall be identified and solved further. This is the compulsory preparation and evaluation before the “iron bird” integration test of the flight control system.

6.4 “Iron bird” integration test of the cockpit control system 6.4.1 Overview Chapter 2 introduces the verification test of electronic, mechanical, and mechanical and electrical airborne equipment of the flight control system. Chapter 5 introduces the combined test of PCU. After the special test bed is established, the combined test of PCU is carried out to confirm that their function and performance meet the subsystem design requirements. Many configurations of the cockpit control system are mentioned in Chapter 5, but they still show generality. This chapter still takes the control column (wheel) type control mode as an example to introduce the “iron bird” integration test of the cockpit control system. Test items of systems of other configurations are similar. The “iron bird” integration test of the cockpit control system is the basis for the “iron bird” integration test of the flight control system and the premise for the subsequent integration test of the fly-by-wire flight control system, machinery control system (if any), and man
machine integration systems. After the cockpit control system is integrated, it shall meet the subsystem design requirements. The difference between the “iron bird” integration test and the combined test of the cockpit control system is that all installation and supporting structures of the airborne equipment of the system are consistent with the state of the aircraft for the maiden flight and the supporting

“Iron bird” integration test of the flight control system 415 base, beam, rib, and associated front console, middle console, left (right) control panel, top control panel, seat, and their adjustment devices are aircraft structures or airborne equipment. Under this condition, the man
machine ergonomics of the cockpit control system can be evaluated in an aircraft cockpit environment and the effect of the supporting stiffness of airborne equipment of the system on the control performance can also be evaluated. Broadly speaking, the cockpit control system shall also include the flaps (slats) control handle of flight control system, override switch, trim control handle of horizontal stabilizer, brake control handle and switch buttons of the automatic flight control panel and primary flight control panel, as well as the throttle control handle of the thrust control system, the front wheel turning handle of the landing gear control system, and the landing gear control handle. As these control devices are functional airborne devices and are relatively simple and their performance has been verified and validated in the integration test of relevant systems, they will not be described here. However, the man
machine ergonomics should still be validated in this test. The “iron bird” integration test of the cockpit control system mainly includes: 1. 2. 3. 4. 5.

Evaluation of man
machine ergonomics of pilot control system. Static performance testing of pilot control system. Dynamic performance testing of pilot control system. Fault mode verification of pilot control system. Study on effects of mechanism supporting stiffness on system dynamic (static) performance.

6.4.2 Test principle The test principle of the “iron bird” integration test of the cockpit control system is shown in Fig. 6.20. The mechanical displacement signal generator is used as a pilot input to test the output of the longitudinal command sensor, horizontal command sensor, heading command sensor, brake command sensor, and turning command sensor. These sensors may be of dualredundancy, four-redundancy, six-redundancy, and eight-redundancy. The consistency of airborne sensors in the same control channel and the difference between test sensors are the focus of test data analysis and processing. For aircraft with a machinery control system, the mechanical output of sector gear (reverse rocker arm) or rocker (bar system) in the corresponding channel should be tested.

6.4.3 Static evaluation of man
machine ergonomics of the cockpit control system The cockpit control system shall be adjusted according to technical conditions for installation. After the front console, central console, left (right) console, top control panel,

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Figure 6.20 Test principle of the “iron bird” integration test of cockpit control system.

seat, and its adjusting devices are adjusted according to the technical conditions for installation of relevant disciplines, experienced pilots (generally the pilots of the first flight crew) shall be hired to experience on-site if they conform to man
machine ergonomics. In other words, if they conform to the pilot’s posture and control habits, if the control functions are realized comfortably, if adjustment mechanisms are used conveniently, and if eye position is reasonable; finally they give conclusions of Good, Acceptable, or Unacceptable. Meanwhile, the body shape difference of pilots such as height, fat or thin, leg length, and arm length shall also be considered. The man
machine ergonomics may also be further evaluated in the flight experience in the man
machine combined test. The principle of the evaluation of man
machine ergonomics of cockpit is shown in Fig. 6.21. The evaluation of man
machine ergonomics of the cockpit control system conducts visibility analysis first to determine the visibility of the airborne equipment in the cockpit. If the visibility is poor, other control postures shall be selected for rejudgment. Then, the accessibility analysis should be carried out to determine the accessibility of airborne equipment in the cockpit. In the case of poor accessibility, other control postures shall be selected for rejudgment. According to the comments of pilots, improvement measures should be studied to finally reach a comfortable control posture. Finally, other man
machine ergonomic indicators shall be analyzed according to the specific situation and requirements.

“Iron bird” integration test of the flight control system 417

Figure 6.21 Evaluation of man
machine ergonomics of the cockpit control system.

Visibility has a significant influence on cockpit man
machine ergonomics and the airworthiness regulations have clear provisions on cockpit interior and exterior views. Generally, the important and frequently used instruments shall be located in the best visual areas, the less important and less frequently used instruments shall be located not exceeding the maximum visual areas when the pilots turn their head and the least used and unimportant instruments shall be located not exceeding the maximum visual areas when the pilots turn their head and eyes (the horizontal visual range of pilots’ two eyes is around 120 and the vertical visual range is about 35 up and down, respectively). Upper limb accessibility analysis is one of the main items for the evaluation of man
machine ergonomics in cockpit. The control devices in the cockpit include control column, control column (wheel), central console, side console, flight control panel unit, rudder panel, and top control panel and there are a wide variety. The control devices shall be reasonably arranged taking into account such factors as degree of importance, frequency of use, and control habit, so as to ensure that the arrangement of all control devices is within the accessible range of a human body of specified sizes and to ensure pilots of all body sizes can reach and control the aircraft.

6.4.4 Static performance testing of the cockpit control system The static performance testing of the cockpit control system aims to check the design conformity of the cockpit control system. Test items include the polarity inspection of

418 Chapter 6 lateral, heading, and longitudinal control channels, motion clearance inspection, control displacement
force characteristic test, trim speed and trim range test, trim mechanism inertial slip test, emergency control displacement
force characteristic test, damping characteristic test, control stiffness test, turning command output test, and brake command output test. The test items and test methods are described as follows. 1. Zero position inspection of airborne equipment and test sensors Before the static performance testing of the cockpit control system, the system shall be checked to see if it is stagnated or if it is at the neutral position under the natural state. If not, the system shall be adjusted to neutral position according to technical conditions for installation. Start the test to test the analysis system and aircraft airborne system and inspect whether the airborne sensor and test sensor are all at zero position (certainly not at the absolute zero position, but within the scope specified by system design and test requirements). If not, zero adjustment shall be carried out for the tested system. The neutral position and zero position inspection above shall be carried out before the testing of new items to ensure the system works normally. 2. Polarity inspection The polarity inspection aims to ensure the control polarity and command output polarity are consistent with the design and check whether the polarity of the test sensor is consistent with the control polarity. If not, the polarity shall be adjusted through software. This method is only applicable to test sensors and inspection and verification shall be carried out in the case of inconsistent polarity. The polarity inspection method is to manually control the cockpit control device and then check if the polarity of the cockpit control device is correct according to the recorded control polarity and airborne sensor output curve. Despite the simple nature of the control polarity inspection, it is of great importance. 3. Motion clearance inspection of channel The motion clearance inspection of the channel can also be considered as a return performance test of the channel. The inspection method is to slowly control the control column (wheel) or pedal manually in two directions from the neutral position, return slowly, and then record the return clearance of the channel. 4. Control displacement
force characteristic curve The control displacement
force characteristic curve is an important performance index of the cockpit control system, that is, the force gradient of the control column. The theoretical curve form is shown in Fig. 6.22. In the control displacement
force characteristic curve test, a mechanical displacement signal generator is usually adopted to drive the control column (wheel or pedal) at a constant speed in full stroke, that is, “Neutral—push forward—stop position—return—pull back—stop position—return.” In the process, the displacement and force of left (right) test sensors, the command

“Iron bird” integration test of the flight control system 419

Figure 6.22 Theoretical control displacement
force characteristic curve.

displacement and command force of airborne sensor, as well as the mechanical input point displacement
time course curve are recorded. If required by the analysis, corresponding data shall be selected to draw the displacement
force characteristic curve. 5. Trim speed and trim range test From the neutral position, control the trim mechanism in full stroke, record the change of corresponding trim displacement with time, check the trim position, and calculate whether the trim speed meets the design requirements. 6. Inertial slip test of trim mechanism Lead the switch quality test point of the trim mechanism, operate the trim mechanism from a neutral position, loosen the trim switch near the preset test point, record the displacement response curve at the test point of the trim switch and control column (wheel or pedal), and analyze the slip quality of the control mechanism at the disconnection point of the trim mechanism. The data analysis method is shown in Fig. 6.23. 7. Emergency control displacement
force characteristic test Emergency control is usually taken to prevent a cockpit control device from getting stuck and affecting flight safety. The emergency control devices usually include an arming mechanism and spring pull rod. If the arming mechanism is used, the left (right) control columns will release and they will not affect the other mutually. If the

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Figure 6.23 Inertial slip analysis method.

spring pull rod is used and the control device is stuck at one side, the pilot will operate the control device at the other side with greater force. The emergency control displacement
force characteristic test aims to test whether the control characteristics of the cockpit control device meet the design requirements when the cockpit control device is stuck. The arming mechanism usually checks if the control device at two sides can release normally and tests the control force and control angle during the release by setting the control device at one side as stuck. The spring pull rod also usually checks if the control force is acceptable and if the control stroke meets design requirements by setting the control device at one side as stuck. 8. Damping characteristic test The damping characteristic test method is to operate the control device to 50% of the full stroke (such as the half stroke after the control column is pushed forward),

“Iron bird” integration test of the flight control system 421

Figure 6.24 Damping characteristic curve.

release the control device quickly, return naturally, and then record the control device in a return displacement
time course curve. The curve shape is shown in Fig. 6.26. The damping coefficient ζ of the system is calculated as per the formula below. The period of oscillation TðT 5 t2 2 t1 Þ and number of oscillations N are read according to the curve in Fig. 6.24. 在此处键入公式
 ln aa12 =2π ζ 5 rffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

 2 1 1 ln aa12 =2π 9. Control stiffness test The control stiffness test method is to fix the output end of the line system and control the pilot control device first. The control force shall not be less than the maximum load force under normal control and not more than 80% of the limited load of the line system. And then record the control force
displacement at the control point, calculate F/L(N/mm), and obtain the control stiffness of the line system. 10. Turning and brake command output test The brake and turning control are part of the pedal control mechanism and they are usually designed jointly. In the test of the cockpit control system, it shall be tested if the turning displacement, force, and command signal outputs as well as the brake command outputs meet the design requirements. As a subsystem test of the “iron bird” integration test, it is impossible for the cockpit control system test to evaluate the performance of the entire aircraft according to the national military standard. The test results are judged as acceptable or not according to the theoretical design data of the cockpit control system. Different aircraft may have different theoretical design data.

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6.4.5 Dynamic performance testing of the cockpit control system The dynamic performance testing of the cockpit control system includes time-domain performance testing and frequency-domain performance testing. 1. Time-domain performance testing By applying a standard input to the hand and foot control channels (longitudinal, heading, and lateral) and recording the output of the command sensor at the same time, the relationship between the output voltage (converted to displacement) and the input displacement is drawn according to the time coordinate and the time-domain characteristics of the output and input are analyzed according to the principle of the control system. Typical input includes step, pulse, and bidirectional pulse and the input amplitude includes small amplitude, medium amplitude, and large amplitude. The lateral control channel is taken as an example below to introduce the method and steps of the time-domain characteristic test. The methods and steps of other control channel tests are similar. Step 1: Apply 10% step input clockwise at the left pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, the testing record data will be reviewed and analyzed and the curves of excitation signal and response signal will be drawn to obtain such parameters as rise time, steady-state time, oscillation times, and overshoot. Proceed to the next step after it is confirmed that they meet the design requirements. Step 2: Apply 10% step input anticlockwise at the left pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, data will be analyzed and processed according to Step 1. Step 3: Apply 10% step input clockwise at the right pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, data will be analyzed and processed according to Step 1. Step 4: Apply 10% step input anticlockwise at the right pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, data will be analyzed and processed according to Step 1. 2. Frequency-domain performance testing The test methods and steps of frequency-domain performance testing are basically the same as those of time-domain performance testing. The difference is that the standard input applied is sinusoidal frequency sweep input rather than step input. The test can be conducted from the left pilot position to the right pilot position. After the

“Iron bird” integration test of the flight control system 423 test is completed, play back and analyze the testing record, draw the curves of the excitation signal and response signal, and then obtain the amplitude margin and phase margin according to the control principle. The typical standard input is the sinusoidal frequency sweep signal and the input amplitude shall include small amplitude, medium amplitude, and large amplitude.

6.5 “Iron bird” integration test of the machinery control system 6.5.1 Overview The flight control system of large aircraft with a machinery control system shall be verified and confirmed to meet the system design requirements and flying quality requirements for the “iron bird” integrated test bed of the flight control system. Only the static characteristic, dynamic characteristic, and fault mode test of the machinery control system are introduced in this part and the man
machine combined test is described in Section 6.9. As mentioned above, the “iron bird” integration test of the cockpit control system is the basis for the “iron bird” integration test of the machinery control system. After it is confirmed that the cockpit control system satisfies the subsystem design requirements, the installation and debugging of the machinery transmission link meet the requirements of the technical conditions for installation and debugging of the machinery control system, and the mechanical backup actuator has technical indexes passing the test, other ground test equipment shall be installed and debugged, mainly including the displacement signal generator, ground hydraulic source, and test sensor. Up until now, the conditions for the test have been available and the test of the machinery control system can be conducted. The main test items of the “iron bird” integration test of the machinery control system are as follows: 1. 2. 3. 4. 5.

Evaluation of man
machine ergonomics of the machinery control system. Static performance testing of machinery control system. Dynamic performance testing of machinery control system. Fault mode verification of machinery control system. Study on effects of mechanism supporting stiffness on system dynamic (static) performance.

6.5.2 Test principle The test principle of the “iron bird” integration test of the machinery control system is shown in Fig. 6.25. The displacement signal generator is used as the pilot input to test the corresponding control plane deflection angle of the test sensor and the mechanical output of the leading

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Figure 6.25 Principle of mechanical backup control system “Iron Bird” integration test.

edge sector gear rocker arm (bar system) and trailing edge (sector gear) or rocker arm (bar system). The output of the airborne control plane deflection angel sensor is also recorded. These sensors may be of multiredundancy configuration. The consistency between test sensors and airborne sensors and the difference between test sensors are the focus of test data analysis and processing.

6.5.3 Evaluation of man
machine ergonomics of the machinery control system After the static evaluation of the man
machine ergonomics of the cockpit control system, the man
machine ergonomics of the cockpit control device with a machinery control system should be further evaluated. Meanwhile, experienced pilots (generally the pilots of the first flight crew) shall be hired to experience on-site if the cockpit control devices conform to man
machine ergonomics. In other words, if they conform to the pilot’s posture and control habits, if the control functions are realized comfortably, if adjustment mechanisms are used conveniently, and if eye position is reasonable. The rationality of the control force and control displacement shall be preliminarily confirmed and finally conclusions of Good, Acceptable, or Unacceptable shall be given. Meanwhile, the body shape difference of pilots such as height, fat or thin, leg length and arm length shall also be considered. The man
machine ergonomics may also be further evaluated in the flight experience in man
machine combined test.

“Iron bird” integration test of the flight control system 425

6.5.4 Static performance testing of the machinery control system The static performance testing of the machinery control system mainly aims to test the input stroke, starting force, and control force of hand and foot control mechanisms and the corresponding relationship between control plane deflection angle and input stroke and force. According to the test results, the idle stroke, stop stroke, and frictional force of each control channel are calculated and the force
displacement curve and displacement
displacement command output curve are drawn. Meanwhile, the consistency and difference between the test results of the test sensor and airborne sensor signals are compared. The starting force and control force of each control handle are tested with a test tool similar to a spring scale and the stall warning function and performance of the stick shaker are tested with a test device similar to an accelerometer. In the actual test, other static performance parameters are calculated from the test results of testing the control force
control displacement curve of each hand and foot control channel (longitudinal, heading, and lateral). The lateral control channel is taken as an example below to introduce the test methods and steps of the control force
control displacement test. Test methods and steps of other control channels are similar. Step 1: After the lateral control channel is confirmed to be in the correct state, the surrounding environment is cleared out without interference, the displacement signal generator and test sensor are firmly connected, and the test analysis system is in good condition, power and liquid should be supplied to the relevant airborne equipment and test equipment of the tested channel to ensure their normal operation. Step 2: Start the test analysis system and keep under a real-time recording state. Step 3: Operate at the left pilot position steadily and slowly for a cycle of 10 s and reciprocate for a complete cycle in the sequence of “Initial neutral position - deflect to the left to limit position - return to neutral position - deflect to the right to limit position return to neutral position.” Step 4: Operate at the right pilot position and control the steering wheel steadily and slowly for a cycle of 10 s, and reciprocate for a complete cycle in the sequence of “Initial neutral position - deflect to the left to limit position - return to neutral position - deflect to the right to limit position - return to neutral position.” Step 5: After the control, stop the reporting of the test and play back and analyze the steering wheel and other signals recorded by the test recording system, get the steering wheel force
wheel displacement relation curve through data processing and obtain control characteristic parameters. Typical steering wheel force
wheel displacement characteristics are shown in Fig. 6.26. According to the analysis and calculation results of the steering wheel force
wheel

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Figure 6.26 Theoretical control displacement
force characteristic curve.

displacement characteristics, the maximum control stroke, maximum control force, starting force, and starting stroke, as well as static lateral control characteristics, such as system clearance and frictional force, can be obtained.

6.5.5 Dynamic performance testing of the machinery control system The dynamic performance testing of the machinery control system includes time-domain performance testing and frequency-domain performance testing. 1. Time-domain performance testing By applying standard input to the hand and foot control channels (longitudinal, heading and lateral) and recording the output of the aileron control plane deflection angle sensor at the same time, the time course relationship between the output voltage (converted to displacement) and the input displacement is drawn according to the time coordinate, and the time-domain characteristics of the output and input

“Iron bird” integration test of the flight control system 427 are analyzed according to the principle of the control system. Typical standard input includes step, pulse, and bidirectional pulse and the input amplitude includes small amplitude, medium amplitude, and large amplitude. The lateral control channel is taken as an example below to introduce the method and steps of the frequencydomain characteristic test. The methods and steps of other control channel tests are similar. Step 1: Apply 10% step input clockwise at the left pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, the testing record data will be reviewed and analyzed and the curves of excitation signal and response signal will be drawn to obtain such parameters as rise time, steady-state time, oscillation times, and overshoot. Proceed to the next step after the design requirements are met. Step 2: Apply 10% step input anticlockwise at the left pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, data will be analyzed and processed according to Step 1. Step 3: Apply 10% step input clockwise at the right pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and it restores to stable, data will be analyzed and processed according to Step 1. Step 4: Apply 10% step input anticlockwise at the right pilot position, record and store the steering wheel displacement, lateral command sensor output, and test sensor output. After the system response ends and restores to stable, data will be analyzed and processed according to Step 1. 2. Frequency-domain performance testing The test methods and steps of frequency-domain performance testing are basically the same as those of time-domain performance testing. The difference is that the standard input applied is a sinusoidal frequency sweep input rather than step input. The test can be conducted from the left pilot position to the right pilot position. Play back and analyze the testing record, draw the curves of excitation signal and response signal, and then obtain the amplitude margin and phase margin according to the control principle. The typical standard input is the sinusoidal frequency sweep signal and the input amplitude shall include small amplitude, medium amplitude, and large amplitude. 3. Damping characteristic test Operate the hand and foot control device to 50% of the full stroke (such as the half stroke after the control column is pushed forward), release the hand and foot control device quickly, return naturally, and then record the return displacement
time response curve of the hand and foot control device. The theoretical curve is shown in Fig. 6.27.

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Figure 6.27 Damping characteristic curve.

The damping coefficient ζ of the system is calculated as per the formula below. Refer to the curve to read the period of oscillation T(T 5 t2
t1) and number of oscillation N.
 ln aa12 =2π ζ 5 rffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

 2 1 1 ln aa12 =2π

6.5.6 Fault mode verification of the machinery control system The fault mode of the machinery control system is stagnation. The test of the release function and mutually independent function of the control column (wheel) are completed under different stagnation states. 1. Stagnation at front end of mechanical link Use a test fixture to fix a pull rod of the control channel (longitudinal, heading and lateral) to simulate the stagnation. Manually operate the control column (wheel) until the control columns (wheels) on the two sides release. In the test, record the time course curve of control column (wheel) angular displacement sensor and force sensor. According to the time course curve, read the variable quantities of the steering wheel angular displacement sensor and force sensor at the control force sudden change point, that is, the release angle and release force at corresponding point. 2. Stagnation at terminal of mechanical link Use a test fixture to fix the sector gear (or rocker arm) at terminal of the cable gearing to simulate the stagnation. Manually operate the control column (wheel) until the control columns (wheels) on the two sides release. In the test, record the time course curve of control column (wheel) angular displacement sensor and force sensor. According to the time course curve, read the variable quantities of the steering wheel angular displacement sensor and force sensor at the control force sudden change point, that is, the release angle and release force at corresponding point. If the stiffness of the

“Iron bird” integration test of the flight control system 429 control channel link is weak, the control column (wheel) may not release when the mechanical link has stagnation at the terminal. If it does not release, the effects of this condition on system control performance should be evaluated.

6.5.7 Study on effects of mechanism support stiffness on system dynamic (static) performance The support stiffness of the airborne equipment and mechanism of the machinery control system has great effects on the performance and function of the machinery control system. If the supporting stiffness is insufficient, greater structural deformation will be caused in control, with the result that the transmission ratio of the system cannot meet the design requirements and the dynamic performance of the control channel will be affected. The support stiffness test is a key factor that distinguishes the “iron bird” integration test of the machinery control system from the system combined test. The reason is that the support stiffness of the airborne equipment and mechanism of the machinery control system on the “iron bird” test bed is basically consistent with that of the actual aircraft and the real aircraft structure is adopted. Support stiffness test method: use a test fixture to fix the output end of the cable gearing line system and operate pilot control devices. The control force shall not be less than the maximum load force of the normal control and not be greater than 80% of the restricted load of the line system. The control force
displacement at the control point shall be recorded and F/L(N/mm) shall be calculated. Thus the system stiffness can be obtained.

6.6 “Iron bird” integration test of the fly-by-wire flight control system 6.6.1 Overview The “iron bird” integration test of the fly-by-wire flight control system is a verification test that shall be carried out after the “iron bird” integration test of the cockpit control system. For aircraft with a machinery control system, the test can be carried out before the “iron bird” integration test of the fly-by-wire flight control system. As the high-lift control system is independent from the fly-by-wire flight control system and machinery control system, the “iron bird” integration test of the high-lift control system can be conducted simultaneously. The “iron bird” integration test of the fly-by-wire flight control system aims to verify and validate that the system meets the system design requirements in a realistic test environment. The test items include: 1. Basic status inspection. 2. Zero position inspection.

430 Chapter 6 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

Interface inspection. Actuator system test. Logic function inspection. Redundancy management test. BIT function inspection. Polarity and transmission ratio inspection. Time-domain characteristic test. Closed-loop frequency-domain characteristic test. Stability margin test. Display alarm function inspection. Failure effect test.

This section focuses on the basic status inspection, zero position inspection, interface inspection, actuator system test, logic function inspection, redundancy management test, and BIT function inspection, etc. of the fly-by-wire flight control system. Polarity and transmission ratio inspection, time-domain characteristic test, closed-loop frequency-domain characteristic test, stability margin test, display alarm function inspection, and failure effect test are introduced in the section of the “iron bird” integration test of the flight control system.

6.6.2 Basic status inspection and testing 1. Installation, adjustment, and debugging The basic status inspection of the fly-by-wire flight control system mainly covers the installation, adjustment, and debugging of airborne equipment, the laying inspection of cables, the installation inspection of hydraulic pipeline, and the installation, adjustment, and debugging of relevant airborne equipment. 2. Status inspection of control-display device, cockpit control device, and mechanical driving system a. Check the status of switches and buttons on the primary flight control panel, among which the “fault recovery” button shall be in normal status, the “fly-by-wire flight control computer” switch shall be in “automatic” status, the “elevator” knob shall be in “fly-by-wire” status, the “aileron” knob shall be in “fly-by-wire” status, and the “rudder” knob shall be in “fly-by-wire” status. b. Check and ensure the left and right buttons on the horizontal stabilizer trim cutoff control panel on the central console are both at the “normal popup” position. c. Check and ensure the brake control handle on the central console is at “0” position. d. Check and ensure the left and right steering wheels are at normal neutral position. e. Check and ensure the left and right control columns are at normal neutral position. f. Check and ensure the left and right control pedals are at normal neutral position.

“Iron bird” integration test of the flight control system 431 g. Pull out the neutral pin of the mechanical control mechanism and confirm the limit pin of cockpit control mechanism is at the normal limit position. 3. Impedance check, power check, and self-check a. Impedance check Impedance check is mainly conducted with ACE and PFC. The cable conduction check between the fly-by-wire flight control system tester and PFC and ACE shall be completed and then the grounding shall be confirmed as correct. Measure the impedance of the corresponding pins of PFC and ACE through the tester front panel and then measure the impedance of power for the test sensor and feedback sensor as well as the impedance of the corresponding PFC/ACE interface through the tester front panel. b. Power check The power check of the system can be carried out after the system adjustment, impedance check, and status check. Before the power connection, all PFC and ACE power supply sources and output power shall be disconnected through the tester front panel. Primary power supply check: start the analog power supply, supply power to each PFC and ACE one by one from the power analog control equipment of the flight control system, and check whether the power supply voltage of PFC and ACE is normal through the tester front panel. After ensuring the power supply of PFC and ACE is normal, connect the power supply for all PFC and ACE through the tester front panel. And then check the power supply voltage of PFC and ACE with load. Supply power to the primary flight control panel, trim control panel, aileron trim mechanism, rudder trim mechanism, horizontal stabilizer trim cutoff control panel, stall warning stick shaker 1, stall warning stick shaker 2, pedal adjustment switch module 1, and pedal adjustment switch module 2 from power analog control equipment of flight control system and observe whether the equipment is under normal status before and after the power connection. Secondary power supply check: check whether the power supply voltage supplied by PFC and ACE to the sensor is normal through the tester front panel. If it is normal, connect the power cables of PFC and ACE sensors from the tester front panel. c. Self-check Before the self-check, the fly-by-wire flight control system shall have been connected to power and pressure and the flight control system tester, airborne equipment exciter, and flight simulation system shall be in normal working state. After the fly-by-wire flight control system is connected to power and pressure, enter the CMS interface for the ground test of the tester, send MBIT (maintenance built-in-test) information, and start system MBIT to conduct system testing. The

432 Chapter 6 airborne equipment replacement test (zero adjustment) and functional test are carried out. Check the system status test results according to configuration information and confirm that all the equipment of the system is working normally. d. Test environment check The test equipment of control mechanism and control plane motion shall not be affected, and the mounting bracket, protective device, and positioning equipment shall be beyond the contact range of the control plane and other motion mechanisms.

6.6.3 Zero position and stroke inspection In the zero position and stroke inspection, the wheel load of the landing gear control system shall be set in effect in the fly-by-wire flight control system tester and the flaps of the highlift control system shall be lowered to landing position. In the zero position and stroke inspection, keep the aircraft on the ground, ensure the flyby-wire flight control system is in normal work condition, slowly operate the control column (wheel) and pedal in large amplitude and observe if the system is stagnated. After it is released, read the data of the FTI device, record the command displacement of the control column, steering wheel, and pedal and record the control plane position value of the aileron, rudder, and elevator. Operate the brake control handle, observe if there is stagnation in the control process and record the position value of the spoiler when the handle is at “0” position. Operate the trim control handle of the horizontal stabilizer and observe in case of stagnation. When the handle is at neutral position, observe if the trim indicator of the horizontal stabilizer is at the zero position. The elevator control channel is taken as an example below to introduce the zero position and stroke inspection. The inspection of methods for other control channels is similar. Zero position and stroke inspection of the elevator control channel shall include the following. 1. Push the left control column forward to check if the system is stagnated. 2. Pull the left control column back to check if the system is stagnated. 3. Release to make the control column in natural neutral position and record the control column displacement and command displacement. If it is not at the zero position, adjust the control column or control column command displacement sensor. 4. After the control column zero position is confirmed as correct, check the command voltage of the elevator actuator at ACE output end. 5. Disconnect the piston rod of the actuator cylinder at the outer side (the side far away from the rudder when the rudder is taken as a reference) of the left inner, left outer, right inner, and right outer elevator actuator, and the control plane.

“Iron bird” integration test of the flight control system 433 6. Through the sensor simulation interface of the tester, generate control column command displacement signal, pitch angle rate signal, angle of attack signal, and normal overload signal and set them as “0.” 7. Check the deflection angle of inner and outer elevators and adjust the piston rod of the actuator cylinder to make the control plane return to the center. 8. Adjust the length of the piston rod of the outer actuator cylinder so that it can be connected with the control plane. 9. Convert the control column command displacement signal to the physical signal through the tester and operate the control column to check whether the system is stagnated. 10. After the control column returns, test the null angle of the control plane. 11. Push the control column forward to the limit position and then test the deflection angle of the elevator. 12. Pull the control column back to the limit position and then test the deflection angle of the elevator. After completion of the zero position and stroke inspection above, the zero position and stroke of the control column, steering wheel, pedal, aileron, rudder, elevator, horizontal stabilizer, and the spoiler shall meet the requirements of technical documents such as Technical Conditions for Installation of Fly-by-wire Flight Control System and Technical Conditions for Installation of Flight Control System Sensor.

6.6.4 Testing of servo actuator system The dynamic (static) characteristics of the servo actuator system shall be inspected under normal power supply, voltage supply, and normal working conditions of the flight control system to verify the main/standby conversion of each control plane actuator and mechanical conversion of the fly-by-wire actuator, the deflection speed, and the dynamic characteristics of each actuator servo circuit under the main/main and main/standby working conditions. 6.6.4.1 Main/standby conversion function inspection of actuator The main/standby conversion function inspection of the actuator is to record the actuator solenoid valve working current (or voltage), electrohydraulic servo valve working current (or voltage), actuator controller state, fly-by-wire system disconnection state information when the actuator has main/standby conversion after the system passes the PBIT (preflight built-in-test) normally. The test steps and methods are as follows. 1. Connect the FTI equipment with the fly-by-wire flight control computer through the fly-by-wire flight control system tester to record the actuator solenoid valve working current (or voltage), electrohydraulic servo valve working current (or voltage), actuator

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2.

3.

4.

5.

controller state, fly-by-wire system disconnection state, and other signals that need to be recorded in the test. Check whether the tested system and environmental support equipment are normal and whether the test conditions can guarantee the normal implementation of the test. If the equipment is normal and environmental conditions are available, supply power and pressure to the tested system and environmental support equipment to ensure normal operation, and then start the FTI equipment to enter the recording state. After the system is supplied with pressure normally and powered on for the first time and PBIT is completed and regarded as normal, the control channel state of left (right) aileron actuator, rudder actuator, and elevator actuator shall be inspected through the FTI equipment. Then, the system is powered off and the control channel state of left (right) aileron actuator, upper (lower) rudder actuator, and inner (outer) elevator actuator is inspected through the FTI equipment. After the system is powered on for the second time and PBIT is completed and regarded as normal, the control channel state of left (right) aileron actuator, rudder actuator, and elevator actuator shall be inspected through the FTI equipment. Then, the system is powered off and the control channel state of the left (right) aileron actuator, rudder actuator, and elevator actuator is inspected through the FTI equipment. After the test is completed, stop the recording by the FTI equipment.

6.6.4.2 Fault return function and performance inspection of actuator The methods and steps of the fault return function and performance inspection of actuator are as follows: 1. Connect the FTI equipment with the fly-by-wire flight control computer through the fly-by-wire flight control system tester to record the aileron, elevator, rudder, multifunctional spoiler actuator, ground spoiler actuator, and other signals that need to be recorded in the test. 2. Check whether the tested system and environmental support equipment are normal and whether the test conditions can guarantee the normal implementation of the test. If the equipment is normal and environmental conditions are available, supply power and pressure to the tested system and environmental support equipment to ensure normal operation, and then start the FTI equipment to enter the recording state. 3. After the system is supplied with pressure normally and powered on and PBIT is completed and regarded as normal, the state of left (right) aileron actuator, rudder actuator, elevator actuator, multifunctional spoiler actuator, and ground spoiler actuator shall be inspected through the FTI equipment. 4. After the system operates normally, disconnect the hydraulic source of the left (right) aileron actuator, rudder actuator, elevator actuator, multifunctional spoiler actuator, and

“Iron bird” integration test of the flight control system 435

Figure 6.28 Servo actuator system test principle.

ground spoiler actuator, observe the working state of the actuators, and check the return state of the actuators. 5. After the test is completed, stop the recording of the FTI equipment. After the inspection on the main/standby conversion and return performance of the actuators above, the maximum speed, displacement, transmission ratio, time-domain characteristic, and frequency-domain characteristic tests of the servo actuator system shall be carried out. The method is to apply excitation signal through the D/A output end of ACE and record the output response of the servo actuator at the same time. The test principle is shown in Fig. 6.28. 6.6.4.3 Maximum output speed and displacement test methods and steps of actuator In the maximum output speed and displacement test of the actuator, apply the maximum input release/withdrawing step signals to the actuator, record the output displacement of the actuator, and analyze whether the maximum output speed and displacement of the actuator meet the design requirements. The test methods and steps are as follows: 1. Connect the FTI equipment with the fly-by-wire flight control computer through the fly-by-wire flight control system tester to record the aileron, elevator, rudder, multifunctional spoiler actuator, ground spoiler actuator, and other signals. 2. The aileron, elevator, rudder, multifunctional spoiler, and ground spoiler actuator working signals should be excited in test and the flight control system test analysis system is used to excite the signals (or for signal generator output). Signals are added to the TEST IN end of ACE from the disconnection part of the fly-by-wire flight control system tester front panel to finish excitation control.

436 Chapter 6 3. Check whether the tested system and environmental support equipment are normal and whether the test conditions can guarantee the normal implementation of the test. If the equipment is normal and environmental conditions are available, supply power and pressure to the tested system and environmental support equipment to ensure normal operation, and then start the FTI equipment to enter the recording state. 4. After the system operates normally, set the signal excitation (or signal generator output) amplitude of the flight control test analysis system corresponding to the maximum input value of aileron, elevator, rudder, multifunctional spoiler, and ground spoiler actuator, and then start the output to generate the step excitation signal. After the test is completed, stop the recording by the FTI equipment, play back the aileron, elevator, rudder, multifunctional spoiler, and ground spoiler actuator signals recorded by the FTI equipment and get the maximum output speed and displacement after data processing. 6.6.4.4 Polarity and transmission ratio inspection In the polarity and transmission ratio inspection, apply sequence voltage signals to the D/A output command point of ACE through the fly-by-wire flight control system tester, test the actuator displacement and control plane deflection angle, and use slow triangular wave signals as excitation sequence voltage signals. The sequence voltage signal shall be gradually increased and both the positive and negative directions shall be considered. The test points shall be densified in areas near the saturation region of the actuator. According to the data recorded by the test analysis system, the curves of the control plane deflection angle and actuator displacement corresponding to the input command are drawn and the polarity can be seen. The polarity and transmission ratio inspection is generally carried out under the normal mode of fly-by-wire flight control system, including cruise configuration polarity and transmission ratio inspection, flaps control process polarity and transmission ratio inspection, takeoff configuration polarity and transmission ratio inspection, plateau takeoff configuration polarity and transmission ratio inspection, landing configuration polarity and transmission ratio inspection, and signal fault polarity and transmission ratio inspection. Under the simulated backup mode, the longitudinal control channel transmission ratio inspection, lateral control channel transmission ratio inspection, heading control channel transmission ratio inspection, pitch rate transmission ratio inspection, rolling rate transmission ratio inspection, and yaw rate transmission ratio inspection shall also be carried out. 6.6.4.5 Time-domain characteristic test In the time-domain characteristic test, apply step signals of different amplitude to the D/A output command point of ACE through the fly-by-wire flight control system tester and record the time response of actuator displacement and control plane deflection angle.

“Iron bird” integration test of the flight control system 437 The time-domain characteristic curve can directly reflect the following performance and stability of actuator system. Time-domain indexes of the actuator system, such as rise time, overshoot, stabilization time, and oscillation times are calculated and then the performance of the actuator system is evaluated with the classical control theory. As for the determination of input signal amplitude, as the actuator system contains many inevitable nonlinear factors such as dead zone, hysteresis, and saturation, too small signal amplitude will affect the reflection of real characteristics and too large signal amplitude will cause damage to the tested system. According to the experience of previous models, the amplitude is generally 6 10% of the maximum stroke. 6.6.4.6 Frequency characteristic test In the frequency characteristic test, the sinusoidal frequency sweep signal is applied to the D/A command output point of ACE through the fly-by-wire flight control system tester and the frequency response of the actuator displacement and control plane deflection corresponding to the input signal is recorded. The input mode is U 5 U0 sin2πft The frequency response test of the actuator system aims to determine the system bandwidth under U input of different amplitude. Generally, the earliest frequency with
3 dB amplitude decrease and
90 phase shift is defined as the bandwidth. As for the determination of input signal amplitude, as the actuator system contains many inevitable nonlinear factors such as dead zone, hysteresis, and saturation and the hydraulic flow is limited, the bandwidth varies significantly with the change of input amplitude. Therefore only the system frequency response characteristics under small signal input can truly reflect the essence of the system, such as the input of 10% full amplitude or lower value.

6.6.5 Logic function inspection The fly-by-wire flight control system of large transport aircraft has characteristics including multiple functions, complex tasks and high safety requirements. As such, large transport aircraft shall not only have a complex control law structure, but also have a large number of control logic functions. The logic functions include power-on logic, fault recovery logic, control priority logic, modal conversion logic, trim function (horizontal stabilizer trim, aileron trim, rudder trim), control plane deflection limit function, automatic deceleration function, aileron auxiliary lift augmentation function, lift destruction and drag increase function, and flight boundary limit and protection function.

438 Chapter 6 The logic function inspection aims to inspect and verify the logic functions above and whether the aircraft transient state caused by logic conversion is within a proper range. Modal conversion function, automatic deceleration function, and flight boundary limit and protection function are taken as examples below to introduce the content and methods of the logic function inspection of the fly-by-wire flight control system. Content and methods of other logic function inspections are similar. 6.6.5.1 Modal conversion function inspection The fly-by-wire flight control system of large transport aircraft generally has normal working mode, degraded working mode, simulated backup working mode, and mechanical backup working mode (if any). The modal conversion function inspection mainly verifies the modal conversion logic of the four working modes above of the fly-by-wire flight control system, including the status display and control plane motion when the normal mode is converted to degraded mode as well as the conversion of normal mode to simulated backup mode and to mechanical backup mode under automatic and manual mode. The modal conversion function inspection of the fly-by-wire flight control system is conducted under normal power supply, pressure supply, normal working state, and given flight state. The flight state is usually set in the flight simulation system. 1. Normal working mode to degraded working mode Set the flight state to cruise state, control the control column, steering wheel, and pedal to have full-stroke motion, set the atmospheric data computer fault or inertial navigation device fault at the same time and continue to control the control column, steering wheel, and pedal to have full-stroke motion. In the test, record the state and command and control plane deflection angle of the fly-by-wire flight control system, analyze and record data and curves and verify the conversion function. 2. Normal working mode or degraded working mode to simulated backup working mode Set the flight state to cruise state, control the control column, steering wheel and pedal to have full-stroke motion, set the fly-by-wire flight control system under normal working mode or degraded working mode respectively, set fly-by-wire flight control computer PFC fault and continue to control the control column, steering wheel and pedal to have full-stroke motion. In the test, record the state and command and control plane deflection angle of the fly-by-wire flight control system, analyze and record data and curves and verify the conversion function. 3. Normal working mode, degraded working mode, and simulated backup working mode to mechanical backup working mode Set the flight state to cruise state, control the control column, steering wheel and pedal to have full-stroke motion, set the fly-by-wire flight control system under normal

“Iron bird” integration test of the flight control system 439 working mode or degraded working mode or simulated backup working mode respectively, set ACE fault and continue to control the control column, steering wheel and pedal to have full-stroke motion. In the test, record the state and command and control plane deflection angle of the fly-by-wire flight control system, analyze and record data and curves and verify the conversion function. 4. Normal working mode to mechanical backup working mode Manually operate the working mode conversion switch on the primary flight control panel to convert aileron control channel to “mechanical” working mode and check whether the conversion function is correct. In the modal conversion function inspection of the fly-by-wire flight control system, the flight state and aircraft attitude signal that need to be recorded can be directly recorded by the flight simulation system. The working state of the fly-by-wire flight control system, the AFCS working mode information, the flaps (slats) handle position and landing gear handle position, the control column (wheel) and pedal displacement, trim mechanism displacement, control plane deflection angle, flaps (slats) position signal, as well as state of atmospheric data system, inertial navigation device, data bus can be recorded by FTI equipment or the test recording system and data bus recording equipment of the flight control system. 6.6.5.2 Lift destruction and drag increase function inspection Large transport aircraft are generally equipped with ground spoiler and multifunctional spoiler, which destroy the lift and increase the drag in landing run stage so as to improve the aircraft landing performance. The lift destruction and drag increase function is realized in a way that the fly-by-wire flight control computer receives the wheel speed signal, wheel load signal and engine throttle state and then controls to open the spoiler through logical calculation. The lift destruction and drag increase function inspection aims to check whether the spoiler is opened normally or not by setting relevant state. The test items and methods are described as follows. 1. Set the fly-by-wire flight control system under normal state and the brake control handle at “1” (armed) position, and then select the flight state through the flight simulation system. 2. Through the control handle and switch in the cockpit, set the aircraft as grounded (wheel load effective or wheel speed lower than a given value), the flight speed greater than a given value, the engine in a slow state and can be opened through reverse pushing. Use a FTI equipment to record flight state, aircraft attitude, steering wheel, control column, and pedal displacement, and spoiler deflection angle signals.

440 Chapter 6 After lift destruction and drag increase function inspection is completed, play back and analyze the data recorded by the FTI and confirm whether the lift destruction and drag increase function meets the design requirements. 6.6.5.3 Flight boundary limit and protection function inspection The fly-by-wire flight control system of large transport aircraft is generally set with such active control functions as attitude hold (pitch angle hold and roll angle hold), stall protection, overspeed protection, overload protection, pitch angle protection, and slant angle protection under normal working mode, so as to realize convenient handling by pilots. The flight boundary limit and protection function inspection aims to verify the realization of the boundary limit and protection function above, as well as to assess the performance of the system and the dynamic characteristics of the aircraft under deep saturation. The detailed content is introduced in the time-domain characteristic test in Section 6.9 of this chapter.

6.6.6 Built-in-test functional inspection and testing Modern fly-by-wire flight control systems all have BIT functions, including power-up builtin-test (PUBIT), in-flight built-in-test (IFBIT), PBIT, and MBIT. BIT functional inspection and testing aims to check whether the conditions for BIT are correct. If the BIT condition does not meet requirements, the system shall exit BIT. The test items and steps are shown below: 1. After the system is connected with power, enter PUBIT. Exit automatically after the completion of PUBIT. 2. After the completion of PUBIT, enter IFBIT. Exit IFBIT after PFC receives the command of MBIT or PBIT. 3. Enter MBIT after PFC is inspected and the command of MBIT is received. Exit automatically after completion of the inspection. 4. Enter PBIT after PFC receives the command of PBIT. Exit automatically after completion of the inspection. In BIT functional inspection, use fly-by-wire FTI equipment to set BIT conditions (wheel load bearing, ground testing allowable, software voting WOW_SW $ 3, and at least software voting indicated airspeed of three channels # given indicated airspeed), set current PFC as PFC1, PFC2, PFC3, PFC4, send a BIT command, start BIT (PUBIT, IFBIT, PBIT, MBIT), display the BIT result from PFC, and check if BIT entry logic is correct. When PFC enters BIT, destroy BIT condition and check whether BIT exit logic is correct. PUBIT is carried out under the condition of normal hydraulic supply and power supply of the system, the PUBIT execution result or fault information is reported to airborne CMS, and the

“Iron bird” integration test of the flight control system 441 result can be read through FTI. After the PUBIT, enter IFBIT and its execution result or fault information is reported to the airborne CMS, and the result is read through FTI. In PBIT functional inspection, send PBIT through CMS or CMS module of FTI and start PBIT. After PBIT is executed, read the results through FTI. MBIT inspection is similar to PBIT inspection. MBIT is sent through CMS or CMS module of FTI to start MBIT. After MBIT is executed, read the results through FTI. In BIT functional inspection, the fault setting way is that to disconnect the power supply for sensor from the fly-by-wire flight control system tester front panel, simulate the pilot command sensor, feed the sensor fault back, cut off the ACE, PFC, primary flight control panel, trim control panel, and horizontal stabilizer trim cutoff control panel, simulate control fault and simulate the actuator fault by disconnecting the electrical interface of actuator feedback sensor. According to the results read by FTI such as the fault code detected in BIT functional inspection, fill the results in test record form and verify BIT inspection.

6.6.7 Redundancy management function inspection The flight control system of large transport aircraft is generally airborne system with high safety requirements. To meet the safety requirements, the fly-by-wire flight control system generally adopts redundancy design technology and redundancy configuration, which may be similar redundancy or dissimilar redundancy. The redundancy management function inspection of the fly-by-wire flight control system mainly verify the redundancy management monitoring and voting functions. The main inspection items include: 1. 2. 3. 4.

Sensor output and value monitoring and voting inspection. Analog signal monitoring and voting inspection. ARINC429 bus (possibly other buses) signal monitoring and voting inspection. Discrete quantity signal monitoring and voting inspection.

In the redundancy management function inspection of the fly-by-wire flight control system, set the command sensor signal and feedback sensor signal respectively, set analog signal and ARINC429 data bus signal and discrete quantity signal, and also record the working mode of the fly-by-wire flight control system, the redundancy signal input value of each channel, redundancy signal voting value of each channel and the channel fault state before/ after redundancy input. The stability margin inspection, display alarm test, and failure effect test, etc. of the fly-bywire flight control system are described in Section 6.9 in this chapter.

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6.7 “Iron bird” integration test of the high-lift system 6.7.1 Overview The “iron bird” integration test of the high-lift system of large aircraft is a laboratory test of system installed state in a realistic aircraft environment after the completion of qualification test (sometimes carried out synchronously) and subsystem combined test of system airborne equipment. The combined test of high-lift system is generally carried out by system supplier on a special test bed. Due to unrealistic test environment caused by development fund and development cycle, only the airborne equipment such as controller, drive device, actuator and sensor inside the system can be combined to solve the potential problems and defects in system internal interface, control logic and drive capability. In fact, for the high-lift system driven by central mechanism, torsion bar and mechanical mechanism, the combined test above is far from enough. Therefore, further combination is required on a more realistic “iron bird” integrated test bed of the flight control system, mainly covering the energy required by the system, that is, aircraft power supply (or hydraulic energy or both), information exchange required, that is, avionics system and fly-by-wire flight control system, mechanism support required, that is, aircraft leading edge and trailing edge and fuselage structure, control object required, that is, control plane and its aerodynamic load, as well as the inertial load and frictional load caused by them. Based on the analysis above, the establishment of a realistic and complete high-lift system test bed becomes an important part of the construction of the “iron bird” integrated test bed for flight control system. For a centrally driven high-lift system, the drive device, transmission device, actuator and motion mechanism are all connected through hardware mechanism and the high-lift control system and related airborne equipment and structures shall be installed according to the state of aircraft for maiden flight, and their support stiffness and strength shall be ensured. The “iron bird” integration test of high-lift system aims to verify and confirm that the system meets the system design requirements in a relatively realistic test environment. The test items and content mainly include: 1. 2. 3. 4. 5. 6.

Basic function and interface inspection; Control function and logic inspection; Coordination and performance inspection; Loaded and unloaded system function and performance inspection; Inspection on effects of wing deformation to flaps (slats) control; Failure effect test.

“Iron bird” integration test of the flight control system 443

Figure 6.29 The principle of “iron bird” integrated test for high lift system.

6.7.2 Test principle The high-lift system is generally composed of flaps and slats control handle, flaps and slats override control panel, FSECU, power drive unit (PDU), actuator, flaps tilt sensor, slats tilt sensor, flaps and slats position sensor and antiwithdrawing brake device. The “iron bird” integration test of high-lift system mainly includes interface inspection, control logic and function inspection, modal conversion function inspection, protection function inspection, fault warning function test, BIT and redundancy management function inspection and failure effect test. Its test principle is shown in Fig. 6.29. 1. Interface inspection Conduct inspection on the data bus signal and transmission line signal around FSECU and PDU. Verify the ARINC429 bus signal transmission between FSECU, PDU and avionics system, the MIL-STD-1553B bus signal transmission between FSECU and PFC, the analog signal transmission between FSECU and flaps position sensor, slats position sensor, flaps tilt sensor, slats tilt sensor, flaps and slats control handle and flaps and slats override control panel, the discrete signal transmission between FSECU and ACE, as well as the discrete signal transmission between PDU and override switch and antiwithdrawing brake device. 2. Control function and logic inspection Check if the flaps (slats) control logic and performance meet requirements under the normal, degraded and backup working modes of the high-lift system respectively.

444 Chapter 6 3. Modal conversion function inspection Check if the conversion function of high-lift system between different working modes meets requirements. 4. Protection function inspection Check the flaps asymmetry protection function, slats symmetry protection function, flaps tilt protection function, slats tilt protection function and flaps actuator system hold and protection function of the high-lift system. 5. Fault warning function test Check if the display results of high-lift system under different working modes and operating mode are correct and verify the fault state testing and alarm. 6. BIT and redundancy management function inspection Check the BIT function and redundancy management function of the high-lift system respectively. 7. Failure effect test Simulate the potential fault modes of the high-lift system and check the effects of the faults on the system. The effects of flaps and slats controller fault, command sensor fault, flaps and slats position sensor fault, flaps and slats tilt sensor fault, cross-linking system signal fault and power supply fault on the high-lift system are mainly checked. There are many signals to be tested in the “iron bird” integration test of high-lift system, including analog signal, discrete signal, data bus signal, angular displacement signal, linear displacement signal, torque signal, speed signal and time signal. The test data are acquired, recorded, displayed and their format is converted through the test sensor, signal conditioner and test data acquisition system of the flight control system test analysis system. For test items requiring load simulation, load simulator is used to simulate the aerodynamic load on flaps (slats) control plane respectively.

6.7.3 Interface inspection According to the relationship between the high-lift system and other aircraft systems introduced in section 1.6 of Chapter 1 of this book, system interfaces mainly include ARINC429 bus, MIL-STD-1553B bus, discrete and analog interfaces. 1. ARINC429 bus interface inspection The ARINC429 bus interface inspection includes the inspection of the ARINC429 bus interface between FSECU and avionics system and the ARINC429 bus interface between FSECU and PDU. Through the ARINC429 bus coupler, the ARINC429 communication data bus of the high-lift control system and the ARINC429 data bus of data bus detection equipment are cross-linked to read the information of the ARINC429 data bus. 2. MIL-STD-1553B bus interface inspection The MIL-STD-1553B bus interface inspection mainly includes the inspection of the MIL-STD-1553B bus interface between FSECU and PFC. Through the MIL-STD-1553B

“Iron bird” integration test of the flight control system 445 bus coupler, the MIL-STD-1553B communication data bus of the high-lift control system and the MIL-STD-1553B data bus of data bus detection equipment are cross-linked to read the information of the MIL-STD-1553B data bus. 3. Discrete interface inspection The discrete interface inspection includes the inspection of the discrete interface between FSECU and ACE and the discrete interface between PDU and override switch and antiwithdrawing brake device. a. Inspection on discrete interface between FSECU and ACE: Introduce the signal from the disconnection part of high-lift system tester front panel to the flight control system test analysis system (or connect to the test channel of digital multimeter, or connect to the acquisition channel of data storage system) to complete the recording of signal test. b. Inspection on discrete interface between PDU and override switch and antiwithdrawing brake device: Introduce the signal from the disconnection part of high-lift system tester front panel to the flight control system test analysis system (or connect to the test channel of digital multimeter, or connect to the acquisition channel of data storage system) to complete the recording of signal test. 4. Analog interface inspection The analog interface inspection includes the inspection of analog interface between FSECU and flaps position sensor and between slats position sensor and flaps and slats control handle. Introduce the signal from the disconnection part of high-lift system tester front panel to the flight control system test analysis system (or connect to the test channel of digital multimeter, or connect to the acquisition channel of data storage system) to complete the recording of signal test.

6.7.4 Control function and logic inspection The basic control logic of the high-lift system is as follows: When the pilot puts the flaps and slats handle in a corresponding gear, the flaps (slats) start to release or withdraw. When the flaps and slats handle has fault or the flaps and slats controller has fault, the flaps and slats can be controlled through the override switch on the flaps and slats override control panel. The control function and logic inspection shall be carried out under normal, degraded and backup working modes respectively to check whether the flaps (slats) control function and logic meet design requirements. 1. Control function and logic inspection under normal working mode Supply 28 V double-circuit DC power to FSECU, PDU controller and antiwithdrawing brake device of the high-lift system and supply 115 V 400 Hz doublecircuit AC power to PDU.

446 Chapter 6 Put the flaps and slats handle in corresponding gear and the flaps (slats) will start to release and withdraw to the target position and record the flaps control plane deflection angle (spiral screw displacement) and slats rack displacement and slats control plane deflection angle at the same time. For the control gear change of the flaps and slats control handle, gradual gear change operation and continuous gear change operation shall be conducted. The specific operation process is as follows. With regard to gradual gear change operation, start from “0” gear and change the gear gradually between “0-1,” “1-2,” “2-3,” “3-4,” “4-5” and “5-4,” “4-3,” “3-2,” “2-1,” “1-0,” and hold for about 1 min between each two gears and then operate at next gear. With regard to continuous gear change operation, start from “0” gear and change the gear between “0-1-3” “3-1-0” and “0-1-4-5” “5-4-1-0,” and hold for about 1 min between each two operations and then conduct next operation. For each operation, check whether the flaps and slats release sequence or withdrawing sequence and their logic meet the design requirements, and analyze whether the flaps and slats release time or withdrawing time meet the requirements according to the recorded time course curve of flaps and slats. 2. Control function and logic inspection under degraded working mode Supply 28 V double-circuit DC power to FSECU, PDU controller and antiwithdrawing brake device of the high-lift control system and supply 115 V 400 Hz single-circuit AC power to PDU. Put the flaps and slats control handle in required gear and the flaps (slats) will start to release and withdraw to the target position and record the flaps control plane deflection angle (spiral screw displacement) and slats rack displacement and slats control plane deflection angle at the same time. For the control gear change of the flaps and slats control handle, gradual gear change operation and continuous gear change operation shall be conducted. The specific operation process is as follows. With regard to gradual gear change operation, start from “0” gear and change the gear gradually between “0-1,” “1-2,” “2-3,” “3-4,” “4-5” and “5-4,” “4-3,” “3-2,” “2-1,” “1-0,” and hold for about 1 min between each two gears and then operate at next gear. With regard to continuous gear change operation, start from “0” gear and change the gear between “0-1-3,” “3-1-0” and “0-1-4-5,” “5-4-1-0,” and hold for about 1 min between each two operations and then conduct next operation. For each operation, check whether the flaps and slats release sequence or withdrawing sequence and their logic meet the design requirements, and analyze whether the flaps and slats release time or withdrawing time meet the requirements according to the recorded time course curve of flaps and slats.

“Iron bird” integration test of the flight control system 447 3. Control function and logic inspection under backup working mode Supply 28 V double-circuit DC power to FSECU, PDU controller and antiwithdrawing brake device of the high-lift control system and supply 115 V 400 Hz double-circuit AC power to PDU. Press the button on the flaps and slats override control panel, observe the flaps (slats) control plane motion condition, and verify the information transmitted to the fly-bywire flight control computer and avionics system. Press the arm button on flaps and slats override control panel. When it is necessary to release the flaps (slats), switch the knob on flaps and slats override control panel to “release” position, wait for about 15 s and switch the knob on flaps (slats) override control panel to “off” position. The flaps (slats) will stop after moving to the mechanical limit position and the release operation is completed. When it is necessary to withdraw the flaps and slats, switch the knob on flaps and slats override control panel to “withdraw” position, wait the flaps (slats) to withdraw, and then press to pop up the arm button on flaps and slats override control panel arm button to complete the withdrawing operation. Supply 115 V 400 Hz single-circuit AC power to PDU and check whether the control function, motion logic and motion time of the flaps (slats) meet design requirements according to the operation process above.

6.7.5 Modal conversion function inspection The high-lift system has three working modes, normal working mode, degraded working mode and backup working mode. Their conversion logic includes normal working mode converted to degraded working mode, normal working mode converted to backup working mode and degraded working mode converted to backup working mode. The backup working mode has the highest priority and the pilot can enter the mode directly according to the flight demands. The modal conversion logic of the high-lift control system and conditions are shown in Fig. 6.30. The modal conversion function inspection mainly checks whether the flaps (slats) working mode can be converted to override mode through the flaps and slats override control panel when flaps (slats) control plane is controlled under normal mode. The inspection methods and process are shown below: 1. Set the flaps and slats control handle in “0” gear, the arm button on flaps and slats override control panel in pop-up state (indicator lamp off) and the rotary knob at “off” position. 2. In the flaps (slats) release process under normal working mode, check the modal conversion function by converting flaps (slats) to backup working mode. The specific operation process is as follows.

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Figure 6.30 Modal conversion logic of the high-lift control system.

Operate the flaps and slats control handle from gear “1-5” and the flaps (slats) control plane deflects to the release direction. During the motion process of the control plane, press the arm button on flaps and slats override control panel at random and then switch the rotary switch on flaps and slats override control panel to “withdraw” position, and the flaps (slats) control plane shall deflect to withdrawing direction. Record the state of flaps and slats control handle and flaps and slats override control panel as well as the flaps (slats) control plane deflection angle time course. Analyze the recorded test data and confirm the flaps (slats) control plane release sequence and logic under normal working mode are correct and the flaps (slats) release time satisfies design requirements, the flaps (slats) control plane release and withdrawing and logic under backup working mode are correct and the flaps (slats) withdrawing time satisfies design requirements. 3. In the flaps (slats) withdrawing process under normal working mode, check the modal conversion function by converting flaps (slats) to backup working mode. The specific operation process is as follows. Operate the flaps and slats control handle from gear “5-1” and the flaps (slats) control plane deflects to the withdrawing direction. During the deflection process of the control plane, press the arm button on flaps and slats override control panel at random and then switch the rotary switch on flaps and slats override control panel to “release”

“Iron bird” integration test of the flight control system 449 position, and the flaps (slats) control plane shall deflect to release direction. Record the state of flaps and slats control handle and flaps and slats override control panel as well as the flaps (slats) control plane deflection angle time course. Analyze the recorded test data and confirm the flaps (slats) control plane withdrawing sequence and logic under normal working mode are correct and the flaps (slats) withdrawing time satisfies design requirements, the flaps (slats) control plane release sequence and logic under backup working mode are correct and the flaps (slats) release time satisfies design requirements.

6.7.6 Safety protection function inspection The high-lift system has safety protection functions including flaps asymmetry protection, slats asymmetry protection, flaps tilt protection, slats tilt protection, safety braking and flaps hold. The methods and process of the safety protection function inspection of the high-lift system are as follows. 1. Inspection of flaps asymmetry protection function a. Set flaps and slats control handle in “0” gear, the arm button on flaps and slats override control panel under pop-up state (indicator lamp off) and the rotary knob at “off” position. b. Release the torsion bar or U-joint at the driving system at the left or right side of the flaps control channel or set the output gradient of the sensor at left (right) of the flaps control channel as inconsistent. c. Operate the flaps and slats control handle from gear “0-1-5” and the flaps (slats) control plane shall deflect to the release direction. Record the state of the flaps and slats control handle and the deflection angle at two sides of the flaps (slats) control plane, observe the deflection angle difference at two sides of the flaps (slats) control plane, and confirm the flaps control plane stops moving when the deflection angle difference is greater than 3 (can be determined according to actual aircraft). 2. Inspection of slats asymmetry protection function a. Set flaps and slats control handle in “0” gear, the arm button on flaps and slats override control panel under pop-up state (indicator lamp off) and the rotary knob at “off” position. b. Release the torsion bar or U-joint at the driving system at the left or right side of slats control channel or set the output gradient of the sensor at left (right) of the slats control channel as inconsistent. c. Operate the flaps and slats control handle from gear “0-1-5” and the flaps (slats) control plane shall deflect to the release direction. Record the state of the flaps and

450 Chapter 6 slats control handle and the deflection angle at two sides of the flaps (slats) control plane, observe the deflection angle difference at two sides of the flaps (slats) control plane, and confirm the slats control plane stops moving when the deflection angle difference is greater than 3 (can be determined according to actual aircraft). 3. Inspection of flaps (slats) tilt protection function a. Set flaps and slats control handle in “0” gear, the arm button on flaps and slats override control panel under pop-up state (indicator lamp off) and the rotary knob at “off” position. b. Release the torsion bar or U-joint at the driving system at one side of flaps control channel or set the output gradient of the tilt sensor at one side of the flaps or slats as inconsistent. c. Operate the flaps and slats control handle from gear “0-1-5” and the flaps (slats) control plane shall deflect to the release direction. Record the state of the flaps and slats control handle, the deflection angle at two sides of the flaps (slats) control plane and data of flaps or slats tilt sensor, and confirm the flaps control plane stops moving when the flaps or slats deflection is greater than 4.5 (can be determined according to actual aircraft). 4. Inspection of safety braking function The inspection of safety braking function mainly includes the braking function inspection of PDU and antiwithdrawing brake device under normal and emergency braking, when the override rotary knob is switched from “release/withdraw” to “stop” under backup working mode and when it reaches the limit position for braking of the PDU under override and backup working mode. a. Set flaps and slats control handle in “0” gear, the arm button on flaps and slats override control panel under pop-up state (indicator lamp off) and the rotary knob at “off” position. b. Braking function inspection of PDU and antiwithdrawing brake device under normal and emergency braking Flaps or slats release process: Operate the flaps and slats control handle from gear “0-1-5” and the flaps (slats) control plane shall deflect to the release direction. Record the state of the flaps and slats control handle and the deflection angle at two sides of the flaps (slats) control plane, confirm whether the flaps or slats release angle meets design requirements when flaps or slats control plane stops moving and whether the brake logic of PDU and antiwithdrawing brake device meets design requirements when flaps or slats stop moving. Flaps or slats withdrawing process: Operate the flaps and slats control handle from gear “5-0” and the flaps (slats) control plane shall deflect to the release direction. Record the state of the flaps and slats control handle and the deflection angle at two sides of the flaps (slats) control plane, confirm whether the flaps or slats release angle meets design requirements when flaps or slats control plane stops

“Iron bird” integration test of the flight control system 451 moving and whether the brake logic of PDU and antiwithdrawing brake device meets design requirements when flaps or slats stop moving. c. Braking function inspection during flaps (slats) release process under backup working mode Flaps or slats release process: Switch the rotary switch on flaps and slats override control panel to “release” position and the flaps (slats) control plane shall deflect to the release direction. In this process, release the rotary switch and record whether the flaps (slats) release angle meets design requirements when the flaps (slats) control plane stop moving in the release process and whether the brake logic of PDU and antiwithdrawing brake device meets design requirements when flaps (slats) stop moving. Flaps or slats withdrawing process: Switch the rotary switch on flaps and slats override control panel to “withdrawing” position and the flaps (slats) control plane shall deflect to the release direction. In this process, release the rotary switch and record whether the flaps (slats) withdrawing angle meets design requirements when the flaps (slats) control plane stop moving in the withdrawing process and whether the brake logic of PDU and antiwithdrawing brake device meets design requirements when flaps (slats) stop moving.

6.7.7 Display and fault warning function test The high-lift system shall be set with complete display and fault warning function, including the display and fault warning function under normal working mode, degraded working mode and backup working mode. The display and fault warning function test is carried out with the control function and logic inspection, modal conversion function inspection and safety protection function inspection introduced in section 6.7.4
6.7.6. It mainly checks whether the display results of the highlift control system under different working modes are consistent with those under the working mode and operating mode and whether the fault is detected and warned. 1. Display and fault warning function test under normal working mode In the display and fault warning function test under normal working mode, pull the flaps and slats control handle from “withdraw” backward to gear “1,” “2,” “3,” “4” and “5” respectively and push from gear “5” forward to gear “4,” “3,” “2,” “1” and “withdraw” respectively, or conduct operation in other combinations, observe whether the flaps and slats position and motion state shown by the EICAS and flight control diagram are consistent with the operation control. 2. Display and fault warning function test under degraded working mode In the display and fault warning function test under degraded working mode, the flaps and slats control is completed when only one circuit of power supplies power and

452 Chapter 6 the other circuit of power is disconnected. It is checked whether the flaps position, slats position, flaps motion time, slats motion, EICAS and flight control diagram are consistent with the design. 3. Display and fault warning function test under backup working mode The display and fault warning function test under backup working mode completes the flaps and slats control by operating the press button on flaps and slats override control panel and check whether the flaps and slats motion and EICAS and flight control diagram are consistent with the design. If current slats position is greater than 18 and current flaps position is greater than 27 , the slats shall be set to override limit position 26 and the flaps shall be set to override limit position 41 . While if current slats position is lower than or equal to 18 and current flaps position is lower than or equal to 27 , the slats shall be set to override limit position 18 and the flaps shall be set to override limit position 27 .

6.7.8 Built-in-test and redundancy management function inspection BIT function inspection checks the starting process and display of PUBIT, PBIT and MBIT under normal working mode. The redundancy management function inspection also sets the cutoff fault of such airborne equipment as flaps position sensor, slats position sensor, flaps tilt sensor and slats tilt sensor one by one and monitors the input effectiveness of analog quantity and discrete quantity of relevant sensors and hard wires cross-linked with FSECU, including the effectiveness monitoring of flaps and slats controller under normal working mode and fault-free condition and the effectiveness monitoring when different deviation values or fault are manually set. Meanwhile, the flaps and slats controller and the voting algorithm of input data monitoring surface and output data monitoring surface in channel and the output data monitoring surface between channels are inspected, mainly including the voting of dual-redundancy analog quantity and voting of dual-redundancy discrete quantity.

6.7.9 Failure effect test In the failure effect test of high-lift system, possible failure mechanisms of the high-lift system are simulated and the system response when the failure occurs is checked. The failure mainly includes flaps and slats controller fault, command sensor fault, flaps and slats position sensor fault, flaps and slats tilt sensor fault, cross-linking system (fly-by-wire flight control system) signal fault and power supply system fault. 1. Signal failure effect test of flaps and slats control handle Set one, two, three or four circuits of RVDT signals of the control handle fail and operate the flaps and slats control handle from gear “0-5,” the flaps (slats) control plane shall

“Iron bird” integration test of the flight control system 453

2.

3.

4.

5.

deflect to the release direction. Or operate the flaps and slats control handle from gear “5-0” and the flaps (slats) control plane shall deflect to the withdrawing direction, and check if the flaps (slats) moves normally when signal of the flaps and slats control handle fails. Failure effect test of flaps and slats position sensor Set a single flaps and slats position sensor to have single-channel fault or doublechannel fault and two flaps and slats position sensor to have single-channel fault or double-channel fault, operate the flaps and slats control handle from gear “0-5,” the flaps (slats) control plane shall deflect to the release direction. Or operate the flaps and slats control handle from gear “5-0” and the flaps (slats) control plane shall deflect to the withdrawing direction, and check if the flaps (slats) control plane moves normally when the flaps and slats position sensor fails. Unconventional operation failure effect test of flaps and slats control handle Check whether the sudden withdrawing operation during the flaps (slats) control plane release process is normal and whether the sudden release operation during the flaps (slats) control plane withdrawing process is normal. Sudden withdrawing operation failure effect test in flaps (slats) release process: Operate the flaps and slats control handle from gear “0” to the preset gear (such as gear “3,” “4,” or “5”), the flaps (slats) control plane deflects to the release direction, operate the flaps and slats control handle back to any gears (such as gear “1,” “2” or “3”) in this process and observe if the flaps or slats move according to preset command and if there is shaking phenomena. Sudden release operation failure effect test in flaps (slats) withdrawing process: Operate the flaps and slats control handle from high gear (such as gear “5,” “4” or “3”) to low gear (such as gear “0” or “1”), the flaps (slats) control plane deflects to the withdrawing direction, operate the flaps and slats control handle back to any gears (such as gear “3,” “2” or “1”) in this process and observe if the flaps or slats move according to preset command and if there is shaking phenomena. Failure effect test of flaps and slats controller Operate the flaps and slats control handle from low gear (such as gear “0”) to high gear (such as gear “5”), the flaps (slats) control plane deflects to the release direction. Or operate the flaps and slats control handle from high gear (such as gear “5”) to low gear (such as gear “0”), the flaps (slats) control plane deflects to the withdrawing direction. In the flaps (slats) control plane release or withdrawing process, set relevant faults of FSECU and check if the flaps (slats) control plane is released or withdrawn according to preset command and if there is shaking phenomena. Failure effect test of PDU Operate the flaps and slats control handle from low gear (such as gear “0”) to high gear (such as gear “5”), the flaps (slats) control plane deflects to the release direction. Or operate the flaps and slats control handle from high gear (such as gear “5”) to low gear (such as gear “0”), the flaps (slats) control plane deflects to the withdrawing

454 Chapter 6 direction. In the flaps (slats) release or withdrawing process, set relevant faults of PDU and check if the flaps (slats) are released or withdrawn according to preset command and if there is shaking phenomena. 6. Failure effect test of flaps tilt sensor Set a single flaps tilt sensor to have single-channel fault or double-channel fault and several flaps tilt sensors to have single-channel fault or double-channel fault, operate the flaps and slats control handle from gear “0-5,” the flaps (slats) control plane shall deflect to the release direction. Or operate the flaps and slats control handle from gear “5-0” and the flaps (slats) control plane shall deflect to the withdrawing direction, and check if the flaps move normally when the flaps tilt sensor fails. 7. Failure effect test of slats tilt sensor Set a single slats tilt sensor to have single-channel fault or double-channel fault and several slats tilt sensors to have single-channel fault or double-channel fault, operate the flaps and slats control handle from gear “0-5,” the flaps (slats) control plane shall deflect to the release direction. Or operate the flaps and slats control handle from gear “5-0” and the flaps (slats) control plane shall deflect to the withdrawing direction, and check if the slats move normally when the slats tilt sensor fails.

6.8 “Iron bird” integration test of the automatic flight control system 6.8.1 Overview Before the “iron bird” integration test, the automatic flight control system of large aircraft has completed the qualification test of relevant airborne equipment, the subsystem combined test under subsystem test environment and the static interface cross-linking test of associated system and confirmed that the functions and performance of the airborne equipment and subsystem met the design requirements. As the combined test in the subsystem test environment takes airborne equipment such as automatic flight control panel, AFCC, autothrottle actuator (or FADEC), back drive actuator (generally available for aircraft) as the core, supplemented with the flight simulation system (including engine), avionics simulation system and fly-by-wire flight control simulation system, and the test system is composed of test support systems such as bus simulator and flight control system test analysis system, the fidelity is still not enough and more realistic aircraft airborne equipment are needed for the cross-linking test. The fly-by-wire flight control system in traditional concept is an internal control loop of the flight control system and the automatic flight control system is an external control loop of the flight control system. As the normal working of the automatic flight control system is based on the normal working of the fly-by-wire flight control system, only after the fly-bywire flight control system has passed full functional and performance verification can the “iron bird” integration test of the automatic flight control system carried out.

“Iron bird” integration test of the flight control system 455 The “iron bird” integration test of the automatic flight control system aims to further integrate the system in a more realistic test environment so as to prepare for the “iron bird” integration test and man
machine combined test of the flight control system. The specific test items include: 1. 2. 3. 4. 5. 6.

interface inspection; polarity and transmission ratio inspection; control logic, display and warning and redundancy function inspection; control function and performance test; stability margin test; BIT and fault test.

6.8.2 Test principles The “iron bird” integration test mainly consists of all or main airborne equipment of airborne systems such as automatic flight control system, fly-by-wire flight control system and avionics system as well as the ground test equipment of flight simulation system, automatic flight control system tester, hydraulic source, power source and flight control system test analysis system. The test principle is shown in Fig. 6.31.

Figure 6.31 Principle of automatic flight control system “Iron Bird” integration test.

456 Chapter 6 Generally, avionics simulation system is used in the early preparation and debugging stage of the “iron bird” integration test of the automatic flight control system. After the automatic flight control system is gradually improved, the real airborne avionics system is included in the test. The signal cross-linking for the “iron bird” integration test of the automatic flight control system is implemented through the automatic flight control system tester. All signals of cross-linking equipment of the automatic flight control system and inside the system are connected in series to realize signal disconnection through the disconnection block. The measuring and control simulation computer on the flight control system tester can realize the switch between signal simulation and physical simulation signals and can record the input and output of the AFCC and the data transmitted with the cross-linking system in real time. Through the flight control system test analysis system, the flight simulation system receives the aircraft control plane deflection angle and engine throttle lever position in real time, calculates the aircraft’s three-axis acceleration, angular acceleration, linear speed, angular rate and three-axis coordinate position in real time, calculates the flight altitude data necessary for the aircraft such as atmospheric inertial navigation data and sends them to the automatic flight control system in real time. The automatic flight control system is cross-linked with fly-by-wire flight control system through the automatic flight control system tester and flight control system test analysis system and flight simulation system, forming a semiphysical test environment for the automatic flight control system.

6.8.3 Interface inspection Before the airborne equipment of the automatic flight control system are cross-linked into a complete system, the airborne equipment shall be ensured to work normally first, which requires the conducting inspection of connecting cables and the impedance inspection of the pins of communication interface and electrical interface of the airborne equipment. After the airborne equipment of the automatic flight control system are cross-linked into a complete system, the cross-linking communication between the airborne equipment in the system shall be checked and the communication with external cross-linking system (such as fly-by-wire flight control system and avionics system) shall also be checked. The interface inspection mainly inspects bus interface, analog interface and discrete interface. The specific method of the interface inspection is that the parameters are set through the simulation measuring and control machine of the automatic flight control system or through real airborne equipment and then received through the receiving terminal of the automatic flight control system tester, and then the data sent by the automatic flight

“Iron bird” integration test of the flight control system 457 control system tester to the system are compared with the system feedback data to judge whether the interface is correct. 1. Bus interface inspection Take a bus signal ARINC429 as an example, for a 32-bit ARINC429 character, 1
8 bits are label bits, 9
10 bits are source/destination bits, 11
29 bits are data bits, 30
31 bits are symbol state bits and 32 bit is parity check bit. The validity of the data is checked through symbol state bit and the maximum data display is 524288. According to the variable measuring range, the resolution of data bits can be determined and the resolution 5 variable measuring range/524288. Thus the error of the ARINC429 data received by the automatic flight control system tester and the ARINC429 bus data sent to the system shall not exceed the size of two resolutions. 2. Discrete interface inspection Discrete signals are generally variables used to show the state of the system or airborne equipment. The data sent to the system by automatic flight control system tester shall be consistent with that received. 3. Analog interface inspection The error range of the analog signal is determined according to the use requirements of specific parameter variables. For example, a mechanical displacement signal generator is used to push the control column forward for 20 mm from zero position, receive the value of control column displacement fed back by the AFCC through the automatic flight control system and compare it with 20 mm. The error generally shall not exceed 1% of the variable’s full measuring range. Details of the interface inspection have been described in Section 6.3.2 State and debugging of cross-linking system.

6.8.4 Polarity and transmission ratio inspection After the interfaces are inspected as correct, the polarity and transmission ratio inspection of the system can be carried out. The interface inspection is carried out in real airborne flight control system environment on the “iron bird” integrated test bed of the flight control system. At this time, the control law has been validated by the simulator and realized through software code and all systems are in the same technical state as aircraft for maiden flight. The basic principle of the polarity inspection is to verify that the automatic flight control system and fly-by-wire flight control system operate normally. By setting current state value of the flight control system through the automatic flight control system tester or setting the expected target value through the automatic flight control panel, the output of the automatic flight control system changes current state command of the aircraft and determines if the polarity is correct according to the deflection direction of the control plane

458 Chapter 6 response command, that is, judge if the command output of automatic flight control system and the deflection polarity of the control plane are correct. For example, if automatic flight control system is under altitude hold mode and the target altitude is 1000 m, when actual state altitude is set to 1500 m, the fly-by-wire flight control system will output the longitudinal command to make the aircraft climb higher and the elevator will deflect to a certain direction. Similarly, the polarity inspection of the automatic flight control system under different modes can ensure the polarity correctness of the automatic flight control system. The basic principle of the transmission ratio inspection is that by setting aircraft state parameters such as aircraft atmospheric and inertial navigation data statically through the automatic flight control system tester or modifying the target parameters of aircraft state through the automatic flight control panel, the real airborne equipment command calculated by the AFCC is compared with the three-axis command of the simulation control law to verify if the polarity and transmission ratio of the control law are correct. The polarity and transmission ratio inspection shall cover the longitudinal, lateral and heading control channels, different flight states and different working modes of the aircraft. The inspection content and workload are massive due to combinations of them. Longitudinal polarity and transmission ratio inspection: It mainly covers the working modes of pitch angle hold (including current pitch angle and amplitude-limit pitch angle hold), altitude hold (including current altitude, vertical speed and amplitude-limit altitude hold) and vertical speed (including target vertical speed, current vertical speed and amplitudelimit vertical speed) under cruise, takeoff and landing configurations. Lateral polarity and transmission ratio inspection: It mainly covers the working modes of slant angle hold (current slant angle, roll angle rate, lateral overload, amplitude-limit roll angle command, amplitude-limit yaw command), heading hold (current heading, roll angle, and indicated airspeed, true airspeed, angle of attack, pitch angle, yaw angle rate), heading selection (target heading, current heading), track keeping (current track, roll angle), track selection (target track, current track) and altitude hold under cruise, takeoff and landing configurations. The pitch angle hold function of the automatic flight control system is taken as an example below to introduce the methods and process of the polarity and transmission ratio inspection of the automatic flight control system. The methods and process of inspection of other functions are similar. When the polarity and transmission ratio inspection is carried out under pitch angle hold mode, start the automatic flight control system first and then start the fly-by-wire flight control system and confirm the fly-by-wire flight control system is working under normal working mode. Set flight state points in flight simulation system, press the “autopilot”

“Iron bird” integration test of the flight control system 459 button and confirm the longitudinal operation under autopilot is conducted under “pitch attitude hold” mode and the heading operation is conducted under “heading hold” mode. The pitch angle hold mode is divided into current pitch angle hold and amplitude-limit pitch angle hold. Current pitch angle hold mode: Input current pitch angle value to 5 through the automatic flight control system tester and wait for 5 s; input current pitch angle value to 0 through the automatic flight control system tester again and wait for 5 s; and finally input current pitch angle value to
5 through the automatic flight control system tester and wait for 5 s. Press the “autopilot disconnection button” twice to disconnect autopilot, clear the warning level warning and record the whole-process test data. Amplitude-limit pitch angle hold mode: Set flight state points of the flight simulation system, rotate the “altitude adjustment knob” on AFCU panel to adjust the altitude to the altitude of set state point, set the target altitude as “state point altitude 500 m,” press the “autopilot” button, confirm the longitudinal operation under autopilot is carried out under “pitch angle hold” mode and the heading operation is carried out under “heading hold” mode, set current pitch angle
5 -0 -5 -10 -15 -20 through the automatic flight control system tester and the data changes at an interval of 1 s, and record the wholeprocess test data.

6.8.5 Control logic and display function inspection The control logic and display function inspections aim to verify the connection, exit and modal conversion logic of the automatic flight control system and the consistency between the display on and control logic of the flight control panel and the avionics display and control system. The control logic and display function inspections adhere to two basic principles. The first is that under the condition of meeting state conditions, connect the automatic flight control system and enable modal conversion, and observe the display of avionics display and control system and automatic flight control panel. The second is that under the condition of not meeting state conditions, enable conversion of modal downgrade and exit of the automatic flight control system and observe the avionics display and control system and the display on the automatic flight control panel. When the automatic flight control system works normally and the cross-linking signal is normal, press the mode button on the automatic flight control panel and the system will start working under the corresponding mode. The control logic and display function inspections mainly cover the logic and display functions including the priority of mode, entry/exit of autopilot, entry/exit of longitudinal

460 Chapter 6 disconnection, entry/exit of lateral disconnection, entry/exit of simultaneous operation, manual override under autopilot mode, entry/exit of longitudinal mode and entry/exit of lateral mode. The priority of mode and entry/exit of autopilot of the automatic flight control system are taken as examples to introduce the methods of control logic and display function inspections of the automatic flight control system. The methods and process of inspection of other functions are similar. 6.8.5.1 Mode priority logic and display function inspection The “approach/landing” mode has the highest level of priority and other modes have equivalent priority. That is to say, by entering a new mode the former mode will exit automatically. Only when the horizontal navigation mode is connected, the vertical navigation mode can be connected. Once the vertical navigation mode is connected, the horizontal navigation mode will exit and the longitudinal mode converts to atmospheric altitude hold mode. Under the autopilot working mode, if AFCS fails to complete current longitudinal mode or lateral mode control, current longitudinal or lateral mode will degrade to default mode and caution level warning will be output. If it fails to enter the default mode, disconnect AFCS and output warning level warning. Under the “approach/landing” mode, it will fail to exit the “approach/landing” mode by selecting longitudinal or lateral working mode. The automatic landing mode will only exit after the autopilot disconnection button is pressed. 6.8.5.2 Autopilot entry/exit logic and display function inspection 1. Autopilot entry logic and display function inspection a. Set the system to enter state A (BANK 5 0 , PITCH 5 0 , HC 5 5000 m, VIAS 5 561.6 km/h, cruise stage, other data are default values), press the “autopilot” button, confirm the flight director switch is at the off position and the “heading/track” reference is heading, observe AFCU “autopilot” “heading hold” signal lamps are on, and PFD shall display “autopilot” “pitch hold” “heading hold.” b. Set the system to enter other states, press the “autopilot” button, and the AFCU signal lamp and PFD do not display. 2. Autopilot exit logic and display function inspection a. Set the system to enter state A (BANK 5 0 , PITCH 5 0 , HC 5 5000 m, VIAS 5 561.6 km/h, cruise stage, other data are default values), connect AP and FD, press the “autopilot disconnection” button, connect altitude hold longitudinally

“Iron bird” integration test of the flight control system 461 and connect heading hold laterally, observe AFCU “autopilot” “heading hold” signal lamps are on, and PFD shall display “autopilot” “pitch hold” “heading hold.” b. Set the system to enter other states, press the “autopilot disconnection” button, and the AFCU signal lamp and PFD do not display.

6.8.6 Control function and performance test Automatic flight control system of large aircraft generally has basic control functions including pitch angle hold, altitude hold, vertical speed, slant angle hold, heading hold, heading selection, track keeping and track selection. The control function and performance test shall cover all the content above and it shall ensure that the above functions are realized and the performance meets the design requirements. The control function and performance test is also called static (dynamic) performance test of the automatic flight control system sometimes. It essentially checks the closed-loop function and static (dynamic) performance of the automatic flight control system. The basic idea of the function and performance test of the automatic flight control system is to test the realization accuracy of relevant functions while inspecting the functions, that is, performance. The assessment criteria are the function and performance indexes specified by system design requirements. The control function and performance test of the automatic flight control system is carried out in a closed-loop environment of the automatic flight control system. The “iron bird” semiphysical test environment composed of automatic flight control system, fly-by-wire flight control system, automatic flight control system tester and flight simulation system can make the aircraft “fly.” The initial state points (including the atmospheric data and inertial navigation data and characteristic parameters of the aircraft) of the automatic flight control system are set in the flight simulation system and the flight envelope adopts aircraft 6-DOF nonlinear motion equation. The test data recorded by the automatic flight control system tester are analyzed to confirm the functions and performance of the automatic flight control system meet design requirements. The pitch angle hold mode control function of the automatic flight control system is taken as an example below to introduce the methods of control function and performance test of the automatic flight control system. The inspection methods and process of other functions are similar. The methods and process of the pitch angle hold mode control function test of the automatic flight control system are as follows. 1. Start the automatic flight control system first and then start the fly-by-wire flight control system, and confirm the fly-by-wire flight control system is working under normal working mode.

462 Chapter 6 2. In the flight simulation system, set flight state points (select cruise, takeoff or landing configuration) and set the automatic flight control system to enter closed-loop state. 3. Press the “autopilot” button to enter autopilot mode, confirm the aircraft has longitudinal operation under “pitch attitude hold” mode and has heading operation under “heading hold” mode. 4. Press the “longitudinal disconnection” button to operate the control column to change the pitch angle to 8 . 5. Press the “longitudinal disconnection” button again to return to the “pitch attitude hold” mode, observe current pitch angle to change to 8 and wait for 3 s. 6. Press the “longitudinal disconnection” button to operate the control column to change the pitch angle to
5 . 7. Press the “longitudinal disconnection” button again to return to the “pitch attitude hold” mode, observe current pitch angle to change to
5 and wait for 3 s. 8. Press the “autopilot disconnection button” twice to disconnect autopilot mode, and clear warning level warning. Record the whole-process test data and analyze and confirm whether the functions and performance meet design requirements.

6.8.7 Built-in-test The automatic flight control system of large aircraft generally has BIT functions, including PUBIT, PBIT, IFBIT and MBIT. The basic test principle of BIT is described as follows. 1. PUBIT: Before connecting the AFCC, set the automatic flight control system in fault state and then power on the automatic flight control system. After the PUBIT is carried out, check if the results of the PUBIT meet design requirements through the automatic flight control system tester. 2. PBIT: After the AFCC is connected and the PUBIT is completed, set the automatic flight control system in fault state and connect PBIT. After the PBIT is carried out, check if the results of PBIT meet design requirements through the automatic flight control system tester. 3. IFBIT: Set the flight profile state of the aircraft in flight simulation system, operate the aircraft to enter test state, connect the autopilot working mode, ensure the automatic flight control system works normally and then set system faults through the automatic flight control system tester, and then read the IFBIT results through the automatic flight control system tester, check if the results of IFBIT meet design requirements. 4. MBIT. After the AFCC is connected and the PUBIT is completed, set the automatic flight control system in fault state and connect MBIT. After the MBIT is carried out, check if the results of MBIT meet design requirements through automatic flight control system tester.

“Iron bird” integration test of the flight control system 463 6.8.7.1 Power-up built-in-test The methods and process of the PUBIT of the automatic flight control system are as follows. 1. Under normal state, power the automatic flight control system up and read PUBIT time through the automatic flight control system tester. 2. Under normal working mode, read the PUBIT results through the automatic flight control system tester. 3. Power the AFCS off and set faults through the automatic flight control system tester, such as single AFCC fault, all AFCC fault, AFCCs inter-channel synchronization fault, AFCCs data cross transmission fault, AFCU fault, ATM fault, single-side BDA fault and double-side BDA fault. 4. Power the automatic flight control system up, read the PUBIT results and confirm if the PUBIT can detect the set faults. 6.8.7.2 Preflight built-in-test The methods and process of the PBIT of the automatic flight control system are as follows: 1. Under normal state, power the automatic flight control system up. 2. After 10 s, through the automatic flight control system tester, set dynamic pressure ,3.5 kPa, air/ground state as “ground,” ADC source effective, two engines closed due to reverse pushing, not started up and shut down, and start the PBIT. Read the PBIT time and results under normal working mode through the automatic flight control system. 3. Set faults through the automatic flight control system tester, such as single AFCC fault, all AFCC fault, AFCCs inter-channel synchronization fault, AFCCs data cross transmission fault, AFCU fault, ATM fault, single-side BDA fault and double-side BDA fault. 4. Start the PBIT, read the PBIT results and confirm if the PBIT can detect the set faults. 6.8.7.3 In-flight built-in-test The methods and process of the IFBIT of the automatic flight control system are as follows: 1. Under normal state, power the automatic flight control system up. 2. Set the aircraft to fly in cruise state and connect the autopilot mode. After the automatic flight control system enters IFBIT automatically, read the IFBIT results from the 429 bus of AFCC cross-linked to CMS through the flight control system tester. 3. Set faults through the automatic flight control system, such as single AFCC fault, all AFCC fault, AFCCs inter-channel synchronization fault, AFCCs data cross transmission fault, AFCU fault, ATM fault, single-side BDA fault and double-side BDA fault. 4. After the faults are set, read the IFBIT results and confirm if the IFBIT can detect the set faults.

464 Chapter 6 6.8.7.4 Maintenance built-in-test The methods and process of the MBIT of the automatic flight control system are as follows: 1. Under normal state, power the automatic flight control system up. 2. After 10 s, through the automatic flight control system tester, set dynamic pressure ,3.5 kPa, air/ground state as “ground,” ADC source effective, two engines in closed due to reverse pushing, not started up and shut down state, and start the PBIT. Read the MBIT time and results under normal working mode through the automatic flight control system tester. 3. Set faults through the automatic flight control system tester, such as single AFCC fault, all AFCC fault, AFCCs inter-channel synchronization fault, AFCCs data cross transmission fault, AFCU fault, ATM fault, single-side BDA fault and double-side BDA fault. 4. After faults are set, read the MBIT results and confirm if the MBIT can detect the set faults.

6.8.8 Failure effect test Large aircraft have high safety requirements for automatic flight control system (as the probability of catastrophic accident caused by automatic flight control system failure is less than 1025). Eliminating and reducing failures of automatic flight control system and minimizing the impact of failures are the keys to the safety design of automatic flight control system. Therefore, failure effect test has become a main content of the verification test of the automatic flight control system. In general, the automatic flight control system adopts measures such as redundancy design and redundancy management, fault reconstruction, modal conversion and degradation safety. The main fault modes of the automatic flight control system include power supply fault, AFCC fault, AFCU fault, ATM fault, BDA fault, AFCC inter-channel synchronization fault, AFCC cross data transmission fault, automatic flight control system connection not allowed by the fly-by-wire flight control system due to external fault and the fault of flight attitude sensor (atmospheric data computer, inertial navigation device and radio altimeter, etc.) externally cross-lined with the automatic flight control system. The failure effect test of the automatic flight control system aims to test whether the functions and performance of automatic flight control system meet the design requirements and flight safety requirements under the fault modes above of the system. The test methods and process are described as follows. 1. Effects of power supply fault Under normal power supply of the automatic flight control system, the flight simulation system and automatic flight control system tester are run to make the

“Iron bird” integration test of the flight control system 465 automatic flight control system work normally under a certain mode and then single and combined faults of each bus bar are simulated to make AFCC, AFCU, ATM and BDA work under single or redundancy power supply faults. The system is run under emergency power supply and the transient characteristics such as the overload at the power supply failure moment and angular rate are checked and set. After the power supply stabilizes, the automatic flight control system shall be checked to see whether it can ensure stable flight and whether the system power supply redundancy and safety design meet the requirements of flight safety. 2. Effects of airborne equipment fault Under normal power supply of the automatic flight control system, the flight simulation system and automatic flight control system tester are run to make the automatic flight control system work normally under a certain mode and AFCC fault, AFCU fault, ATM fault and BDA fault are simulated by powering off all airborne equipment or setting communication failures to check, after the airborne equipment fault of the automatic flight control system stabilizes, whether the fault reconstruction conforms to the fault working mode determined in design, the transient characteristics such as the overload at the fault setting moment and angular rate and whether the airborne equipment fault reconstruction of automatic flight control system and model conversion design meet the requirements of flight safety. 3. Effects of AFCC synchronization and cross communication fault Under normal power supply of the automatic flight control system, the flight simulation system and automatic flight control system tester are run to make the automatic flight control system work normally under a certain mode, a circuit or several circuits of AFCC synchronizing signals are disconnected and then the automatic flight control system is started and the AFCC synchronization and cross communication faults are determined according to the automatic flight control system fault reported by BIT. 4. Cross-linking system signal fault Cross-linking system signal faults mainly include two parts, that is, flight attitude signal faults of atmospheric data computer and inertial navigation system, etc. and faults of flight management system and fly-by-wire flight control system. By setting flight attitude signal faults, the transient characteristics of the automatic flight control system during modal conversion and fault setting are checked to see whether they meet flight safety requirements. For example, after the flight simulation system and automatic flight control system tester are started up, the automatic flight control system is set to enter the vertical climb mode longitudinally and then vertical speed signal faults are set. The automatic flight control system shall exit the vertical speed mode and convert to the default pitch angle hold mode and then set the pitch angle signal faults. The automatic flight control system shall also exit the control function, record the signal faults and the transient characteristics during modal conversion simultaneously and confirm if they meet flight safety requirements.

466 Chapter 6 After signal faults of flight management system are set, confirm if the automatic flight control system can exit automatic navigation safely and stably. After faults of fly-by-wire flight control system are set, confirm if the automatic flight control system can exit automatic navigation and autopilot functions safely and stably.

6.9 “Iron bird” integration test of the flight control system 6.9.1 Overview According to engineering development idea, in section 1.6 of this book, the flight control system of large aircraft is divided into cockpit control system, fly-by-wire flight control system, high-lift system and automatic flight control system. For aircraft with machinery control function, the flight control system also includes machinery control system. In the development process, the idea of hierarchical synthesis is adopted. In other words, the airborne equipment supplier shall solve the airborne equipment problems and the system supplier shall solve the subsystem problems. Airborne equipment problems shall not be left to subsystem and system and subsystem problems shall not be left to system as far as possible. If the synthesis result of last level requires changes of next level, that is an optimization problem which requires further improvement of the last level system design requirements. Although subsystems of the flight control system have their own independent functions and performance, they are also mutually interdependent and supported. Coordinating the relationship between subsystems and making it meet the requirements of flight control system design specification is an important part of the flight control system design, and the verification and validation of the flight control system is one of the main purposes and content of the “iron bird” integration test of the flight control system. According to the design idea of “hierarchical design and layer-by-layer synthesis” of the flight control system, Chapter 5 introduces the combined test of subsystems of flight control system of large aircraft in supplier test environment. It is fair to say that subsystem combined test can solve most design defects and problems of the subsystem. Of course, this is related to subsystem developers’ technical level, test conditions, economic strength and management level, etc. Therefore, the selection of supplier is crucial for the development level of airborne equipment and systems and the subsequent “iron bird” integration test cycle of flight control system. In section 6.4
6.8 of this book, we introduce the “iron bird” integration test of subsystems. In a more realistic “iron bird” integration test environment, the functions and performance of subsystems are further verified and their working stability is assessed. Meanwhile, their interdependence relations can be seen from previous sections. In other words, the “iron bird” integration test of the cockpit control system is the basis for subsequent “iron bird” integration tests of various systems. The “iron bird” integration test of the automatic flight control system shall be performed after the “iron bird” integration test of the fly-by-wire

“Iron bird” integration test of the flight control system 467 flight control system and the “iron bird” integration test of fly-by-wire flight control system requires the participation of high-lift control system in full states. The same is for the “iron bird” integration test of the machinery control system. In addition to the complex internal functional interdependence, performance support and interface relations, the normal working of the flight control system also requires the energy guarantee, information provision and command execution functions of other aircraft systems, such as hydraulic system, power supply system, avionics system, landing gear control system and power system. The matching verification of the flight control system with these aircraft systems is also one of the important purposes and content of the “iron bird” integration test of the flight control system. Specific test items include: 1. 2. 3. 4. 5. 6.

interface inspection; polarity and transmission ratio inspection; control logic, display and warning, redundancy function inspection; control function inspection and performance test; stability margin test; BIT and fault test.

6.9.2 Test principles The test principles of the “iron bird” integration test of flight control system of large aircraft are shown in Fig. 6.32. Test items such as interface inspection, polarity inspection, logic function inspection, trim function and trim logic check, modal conversion function inspection, system state and warning and display function inspection and system BIT shall be carried out in open-loop state of the flight control system. Test sensor and flight control system test analysis system are used for testing of test parameters. Data bus simulation equipment or FTI equipment is used for testing of internal information flow of the flight control system. Flight control system tester (including fly-by-wire flight control system tester, automatic flight control system tester and high-lift control system tester) is used for the testing of internally transmitted signals of the flight control system, or the signals to be tested are introduced from the flight control system tester to the test recording equipment. In the flying quality test, modal conversion to transit state test, logic function inspection, modal conversion function inspection, state and alarm display test and fault simulation test of the flight control system, the deflection angle signals of the control plane are introduced to the flight simulation system and then aircraft dynamics equation and engine thrust equation of the flight simulation system are used to calculate the aircraft motion attitude and motion state parameters. They are taken as the command signals of airborne equipment such as linear acceleration turntable, three-axis rate turntable, atmospheric data simulator

468 Chapter 6

Figure 6.32 Principle of flight control system “Iron Bird” integration test for large transport aircraft.

exciter, inertial navigation simulation exciter and intelligent probe simulation exciter to control the operation of each test equipment, drive three-axis overload sensor and angular rate gyro assembly and excite each simulator exciter, so as to provide feedback signals for fly-by-wire flight control computer, AFCC and flaps and slats controller.

6.9.3 Interface inspection The interface inspection covers the discrete interface between fly-by-wire flight control computer PFC, ACE, AFCC, FSECU and internal cross-linking interfaces such as ARINC429, MIL-STD-1553B data bus signal interfaces, and also the external cross-linking interfaces such as ARINC429, MIL-STD-1553B and AFDX data bus signal interfaces between internal equipment of the flight control system and airborne equipment including atmospheric data computer, inertial navigation, radio altimeter, central warning system, avionics display control (DPU) system, engine parameter recording system and central maintenance system. 1. Analog and discrete interface inspection Introduce the signals at the disconnection part of the front panel of flight control system tester to the flight control system test analysis system (connect to the test

“Iron bird” integration test of the flight control system 469 channel of the digital multimeter, or the acquisition channel of the data storage system) or use a FTI equipment to record system state data to complete the test and recording of the signals. 2. Data bus interface inspection If the data bus signal can be generated by controlling or adjusting the state of the airborne equipment of the tested system, the interface signal change can be detected by controlling or adjusting the state of the airborne equipment of the tested system at different positions. The state of the system is selected at the maximum position, the minimum position and the middle position and the detected interface signals are compared with current state of the system to determine whether they are consistent. If the data bus signal cannot be generated by controlling or adjusting the state of the airborne equipment of the tested system, the interface signal change shall be detected by using ground maintenance equipment or simulation exciter to set their interface signals in different states. The state of the tested system is selected at the maximum position, the minimum position and the middle position and the detected interface signals are compared with current state of the system to determine whether they are consistent. In the inspection of interface between the atmospheric data system and PFC, the dynamic (static) pressure simulator simulates different altitudes and speeds after the preparation for power-on of the atmospheric data equipment is completed. In the inspection of interface between the inertial navigation and PFC, the inertial navigation device is driven by turntable to give attitude signals after the preparation for power-on of the inertial navigation device is completed. In the inspection of interface between the radio altimeter and PFC, altitude signals are given through exciter and then radio altimeter data received by PFC and output after PFC voting are detected. Interface inspection of typical flight control system is taken as an example below to introduce the interface inspection methods and process. Other interface inspection methods and process are similar. 1. Interface inspection between fly-by-wire flight control computer PFC and FSECU Through the MIL-STD-1553B data bus coupler, the MIL-STD-1553B data bus signals of the flight control system are connected to data bus detection equipment. Through the ARINC429 data bus coupler, the data bus signals of FSECU connected to accident recording equipment (FDAU) and RDC are connected to data bus detection equipment. The information that PFC gets from FSECU is obtained through FTI equipment. a. Supply power and pressure to the flight control system and adjust the flight control system under normal working state. b. Supply power to test support equipment such as data bus detection equipment, FTI equipment and flight control system test analysis system.

470 Chapter 6 c. Run the application software, enter data bus testing and recording state as well as the test recording state. d. Operate the flaps (slats) control handle to different gears and hold at current gear for 10 s, change the flaps (slats) control handle to next gear until the operation of the flaps and slats control handle at all gears is completed. e. Use a data bus detection equipment to record the “flaps control plane deflection angle voting value” and “slats control plane deflection angle voting value” data information sent by FSECU to PFC through the MIL-STD-1553B data bus, use a data bus detection equipment to obtain the “flaps and slats control handle position,” “flaps control plane deflection angle” and “slats control plane deflection angle” data information sent by FSECU to accident recording equipment (FDAU) through ARINC429 data bus, and use a FTI equipment to obtain the “flaps and slats control handle position,” “flaps control plane deflection angle” and “slats control plane deflection angle” data information of PFC through RS-422 data bus. f. Play back and analyze the “flaps control plane deflection angle voting value” and “slats control plane deflection angle voting value” data information obtained by FTI equipment, data bus detection equipment and flight control system test analysis system, compare the flaps and slats control handle gear and flaps (slats) deflection angle signals, flaps (slats) control plane deflection angle data received by PFC, flaps position state signals received by ACE, flaps and slats control handle gear and flaps and slats control plane deflection angle data recorded by accident recording equipment and analyze whether the corresponding parameter results recorded by different equipment are consistent. 2. Interface inspection between flight control system and atmospheric data system The flight control computer (PFC, AFCC) and atmospheric data system adopt ARINC429 data bus communication and their interface inspection methods and process are described as below: a. Power on the flight control system and the system enters normal working state. b. Set multiredundancy true airspeed, indicated airspeed, true angle of attack, Mach number, corrected pressure altitude, static pressure, dynamic pressure and pressure altitude and other parameters through atmospheric data system simulation exciter (or data bus monitoring equipment) and send to PFC and AFCC. c. Use a FTI to record the bus data sent by atmospheric data system and received by PFC and AFCC through ARINC429 data bus and check and compare the consistency of the sent data and received data. 3. Interface inspection between flight control system and inertial navigation system The flight control computer (PFC, AFCC) and inertial navigation system adopt ARINC429 data bus communication and their interface inspection methods and process are described as below: a. Power on the flight control system and the system enters normal working state.

“Iron bird” integration test of the flight control system 471 b. Set multiredundancy pitch angle, roll angle, pitch rate, roll rate, yaw rate, ground speed, forward linear acceleration, lateral linear acceleration, normal linear acceleration and other parameters through inertial navigation system simulation exciter (or data bus monitoring equipment) and send to PFC and AFCC. c. Use a FTI to record the bus data sent by inertial navigation system and received by PFC and AFCC through ARINC429 data bus and check and compare the consistency of the sent data and received data.

6.9.4 Polarity and transmission ratio inspection The polarity and transmission ratio inspection is carried out under full states of the flight control system to verify the polarity and transmission ratio correctness of the fly-by-wire flight control system, including the polarity and transmission ratio inspection of steering wheel displacement, control column displacement, pedal displacement, pitch rate, slant rate, yaw rate, normal overload, lateral overload, angle of attack sensor, inertial navigation device, etc. to horizontal stabilizer, elevator, aileron, rudder and spoiler, etc. The polarity and transmission ratio inspection is conducted under normal working mode, degraded working mode and simulated backup working mode of the flight control system respectively. According to the aircraft configuration, the initial flight state points are set through flight simulation system, command signals are generated by controlling cockpit control mechanism, or simulation command or feedback signal is generated through simulation exciter and added to system inputs, command signal, feedback signal and control plane response signal are recorded, graphic curves are generated again and compared with control law simulation curves to judge their consistency. 6.9.4.1 Polarity and transmission ratio inspection under normal working mode The methods and process of the polarity and transmission ratio inspection under normal working mode of the fly-by-wire flight control system are as follows: 1. Start up the flight control system and enter normal working mode. 2. Set the flight state and initial values of parameters through the flight simulation system. 3. Through the flight simulation system, set different input command or feedback signal in corresponding sequence. The continuous changes of the parameters are shown in Table 6.1. The sequence is Mach number, dynamic pressure, true airspeed, indicated airspeed, pitch angle, slant angle and pressure altitude, showing the linear change respectively from current initial value to minimum value, minimum value to maximum value, maximum value to initial value. The signal change cycle is 20 s. 4. Use a FTI equipment to record command signals including steering wheel displacement, control column displacement, pedal displacement, brake control handle position, horizontal stabilizer handle location, flaps and slats control handle gears, pitch

472 Chapter 6 Table 6.1: Parameter setting for polarity and transmission ratio inspection under normal working mode. No.

Signal name

Initial value

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23

Control column displacement Steering wheel displacement Pedal displacement Normal overload Lateral overload Pitch rate Yaw rate Roll rate Radio altitude Wheel load Aircraft weight Center of gravity of aircraft Deflection angle of horizontal stabilizer Angle of attack Flaps position Slats position Mach number Dynamic pressure True airspeed Indicated airspeed Pitch angle Slant angle Pressure altitude


20 mm 15 15 mm 1.5 g 0.2 g 5 /s 5 /s 5 /s 100 m 0 130 t 0.32
2 10 0 0 Initial value: 0.5 Initial value: 10,000 Initial value: 400 Initial value: 400 Initial value: 15 Initial value: 30 Initial value: 5000

Setting and change range

(0
0.9) ma linear change (0
25,000) Pa linear change (0
900) km/h linear change (0
700) km/h linear change (
30 to 30) linear change (
60 to 60) linear change (0
12,000) m linear change

rate, slant rate, yaw rate, feedback signals including normal overload, lateral overload, angle of attack sensor, inertial navigation device, working state of the system, command word output by flight control computer and deflection angle of control planes including horizontal stabilizer, elevator, aileron, rudder and spoiler. 5. Analyze, process and record data, generate graphs of each input excitation signal and output response signal, and evaluate test results according to requirements and methods for evaluation of flight control system test results in Section 6.11. 6.9.4.2 Polarity and transmission ratio inspection under simulated backup working mode The methods and process of the polarity and transmission ratio inspection under simulated backup working mode of the fly-by-wire flight control system are as follows: 1. Start up the flight control system and enter normal working mode. 2. Set the flight state and initial values of parameters through the flight simulation system. 3. Use a FTI equipment to record command signals including steering wheel displacement, control column displacement, pedal displacement, brake control handle position, horizontal stabilizer handle location, flaps and slats control handle position,

“Iron bird” integration test of the flight control system 473 pitch rate, slant rate, yaw rate, feedback signals including normal overload, lateral overload, angle of attack sensor, inertial navigation device, working state of the system, command word output by flight control computer and deflection angle of control planes including horizontal stabilizer, elevator, aileron, rudder and spoiler. 4. The test methods and process are as follows. a. Inspection of polarity and transmission ratio from steering wheel displacement to aileron and spoiler control plane: Operate the steering wheel in a sequence of “inertial neutral position - deflect the steering wheel to the left to 20% position - return to neutral position - deflect the steering wheel to the right to 20% position - return to neutral position,” stably and slowly operate the steering wheel to make it have reciprocating motion for a complete cycle. b. Inspection of polarity and transmission ratio from control column displacement to control planes including horizontal stabilizer and elevator and from pedal displacement to rudder: It is similar to the method of polarity and transmission ratio inspection from steering wheel displacement to aileron and spoiler and only the control column or pedal is operated. For selection of limit position, a wider range is available, such as 50% position or limit position. c. Inspection of polarity and transmission ratio from pitch rate to control planes including horizontal stabilizer and elevator: Use a mechanical displacement signal generator to generate triangular excitation signals and corresponding pitch rate changes from “0 /s- 2 10 /s-0 /s-10 /s-0 /s” and a complete cycle is output. d. Inspection of polarity and transmission ratio from feedback signals including slant rate, yaw rate, normal overload, lateral overload, angle of attack sensor, inertial navigation device to control planes including aileron, rudder and spoiler: It is similar to the method of polarity and transmission ratio inspection from pitch rate to horizontal stabilizer and elevator. As the signal generator generates slant rate, yaw rate and signals of airborne equipment such as normal overload, lateral overload, angle of attack sensor and inertial navigation device, as for selection of limit position, a wider range is also available, such as 10 /s or maximum value. e. Process, analyze and record data, generate graphs of each input excitation signal and output response signal, and evaluate test results according to requirements and methods for evaluation of flight control system “iron bird” integration test results in Section 6.11.

6.9.5 Stability margin test Stability is the basic requirement for control system. It refers to the ability to restore to the initial balance state with sufficient accuracy after the external reaction force on the system is removed. The definition of stability actually reflects three requirements for the control system, that is, it shall be able to restore and restore quickly, and the restored state shall be highly consistent with the initial state. This concept just shows the aircraft stability,

474 Chapter 6 maneuverability and control accuracy if reflected in the flight control system, that is, flying quality. The flying quality design specification integrates the flying quality requirements of all types of aircraft above and the stability margin of the flight control system is used to show the flying quality requirements of the aircraft. The development of modeling and simulation technology has created good conditions and made great contributions to the research and evaluation of the stability margin of the flight control system through digital simulation. However, due to the large number of airborne equipment and complex structure of the flight control system, especially the difficulty in establishing accurate mathematical models for hydraulic servo steering engine and its hydraulic energy and the difficulty in realizing the real-time simulation under complex model, it is crucial to establish an integrated test bed including a real flight control system to verify and evaluate the stability margin of the flight control system. The stability margin test includes the stability margin of the fly-by-wire flight control system under normal working mode, simulated backup working mode and autopilot mode. This book only introduces the stability margin test of the fly-by-wire flight control system and the test methods and process under the autopilot mode are similar. The test principle of the stability margin test of flight control system of large aircraft is shown in Fig. 6.33.

Figure 6.33 Principle of stability margin test of the flight control system.

“Iron bird” integration test of the flight control system 475 1. Stability margin test of fly-by-wire flight control system under normal working mode a. On the “iron bird” integrated test bed of the flight control system, connect the test system according to the test principle shown in Fig. 6.42. b. Power on the flight control system and make the fly-by-wire flight control system have closed-loop working under normal working mode. c. Set aircraft configuration and flight state through flight simulation system and trim the aircraft under set flight state. d. Use a dynamic frequency response analyzer to generate sinusoidal frequency sweep signal, input it to ACE through the test port of ACE on the front panel of fly-bywire flight control system, and output to actuator after being added with the simulation command signal after D/A conversion in ACE and servo amplification. e. Input the control plane deflection angle signals caused due to motion of actuator drive control plane to flight simulation system. f. The flight simulation system calculates the aircraft motion state signal to ACE (input to PFC through bus after entering ACE) and PFC and the PFC sends control plane control command to ACE and have D/A conversion. g. Input the control plane control commands (obtained from the test output port on the front panel of fly-by-wire flight control system tester) after D/A conversion and frequency sweep signals to open-loop frequency sweep inverse module for addition and phase inversion and then input to dynamic frequency response analyzer. h. Use a dynamic frequency response analyzer to draw a Bode diagram of control plane control commands after D/A conversion corresponding to open-loop frequency sweep inverse module and analyze the stability margin. 2. Stability margin test of fly-by-wire flight control system under simulated backup working mode a. On the “iron bird” integrated test bed of the flight control system, connect the test system according to the test principle shown in Fig. 6.33. b. Power on the flight control system and make the fly-by-wire flight control system have closed-loop working under normal working mode. c. Set aircraft configuration and flight state through flight simulation system and trim the aircraft under set flight state. d. Use a dynamic frequency response analyzer to generate sinusoidal frequency sweep signal, input it to ACE through the test port of ACE on the front panel of fly-bywire flight control system, and output to actuator after being added with the simulation command signal after D/A conversion in ACE and servo amplification. e. Input the control plane deflection angle signals caused due to motion of actuator drive control plane to flight simulation system. f. The flight simulation system calculates the aircraft motion state signal to ACE, the ACE generates actuator control command and has D/A conversion, the actuator control commands after D/A conversion and frequency sweep signals are input to

476 Chapter 6 open-loop frequency sweep inverse module for addition and phase inversion and then to dynamic frequency response analyzer. g. Use a dynamic frequency response analyzer to draw a Bode diagram of actuator control commands after D/A conversion corresponding to open-loop frequency sweep inverse module and analyze the stability margin.

6.9.6 Closed-loop frequency response test The flight control system of large transport aircraft has a complex composition that involves nearly a thousand electronic, mechanical and electrical and mechanical airborne equipment. The airborne equipment requires power and hydraulic source to provide working energy and contain complex dynamic characteristics and prominent nonlinear factors. In the design stage of flight control system, we adopt simplified linear mathematical model and typical nonlinear link to carry out system design and analysis. However, in the flight control system test, we can no longer separate the linear and nonlinear characteristics of the flight control system, which makes the test result shows the characteristics of a high-order system. In addition, the 6-DOF nonlinear motion, engine thrust and landing gear equations of the aircraft make it more difficult to analyze the test results of the closed-loop flight control system including the aircraft. GJB185
1986 and GJB2874-1997 define the evaluation methods and standards of closedloop characteristics of flight control systems according to the standard second-order system. Therefore, when it is necessary to analyze or verify the closed-loop characteristics of a particular flight control system through tests, the general practice is to equalize the test results (high order only) to the standard second-order system characteristics. The closed-loop frequency response test of the flight control system aims to obtain the frequency response of a high-order flight control system (including aircraft dynamics) and then get the frequency characteristics and equivalent low-order short-period motion characteristic parameters in the form specified by the national military standard with the equivalent matching method. The test principle of closed-loop frequency response test is shown in Fig. 6.34. The closed-loop frequency response test is conducted under the normal and simulated backup working modes of the fly-by-wire flight control system respectively to assess the closed-loop frequency response performance under the two working modes, so as to assess whether the system meets the flying quality requirements. The test method and process are described as follows: 1. Set the signal generator of the dynamic signal frequency response analysis system as sinusoidal frequency sweep signal and set the amplitude and sweep range according to the test requirements.

“Iron bird” integration test of the flight control system 477

Figure 6.34 Principle of closed-loop frequency response analysis test of the flight control system.

2. Connect the excitation signal (or signal generator output) of the dynamic signal frequency response analysis system to the control input end of the mechanical displacement signal generator and introduce the signal to the analyzer reference channel of the dynamic signal frequency response analysis system as the reference signal for frequency response analysis. 3. Start the dynamic signal frequency response analysis system to enter signal analysis and recording state and start the signal generator to generate the output excitation signal. Use a displacement signal generator to apply frequency sweep signal of different amplitude and frequency range to the control column, steering wheel and pedal and analyze the response of actuator output and control plane deflection angle to input. 4. Use a test sensor to test the control signals including steering wheel displacement, pedal displacement and control column displacement, displacement of aileron mechanical backup actuator, rudder mechanical backup actuator, elevator mechanical backup actuator, as well as deflection angle signals of control planes including aileron, rudder and elevator, and introduce the test signals to dynamic signal frequency response analysis system for signal processing and analysis. 5. After the test, analyze the excitation signal, actuator output and control plane deflection angle signal recorded in the test of dynamic signal frequency response analysis system, draw the amplitude-frequency characteristic curve and phase-frequency characteristic curve of the actuator output and control plane deflection angle signal relative to the excitation signal and then obtain the characteristic parameters such as system phase lag from the phase-frequency characteristic curve.

478 Chapter 6 As mentioned above, provide sinusoidal excitation signal to the position signal generator through the dynamic signal frequency response analyzer and test the frequency characteristics of the aircraft motion ωz relative to the command sensor. The actual frequency response of high-order system is obtained in the closed-loop frequency response test and the performance of flight control system can be directly evaluated according to frequency domain analysis technology. To compare with current flying quality design specification, the test results can be equivalently matched to lowerorder transfer function as below:
 s11 K Tθ2 ωz ðsÞ 5 2 Ue2τs XS ðsÞ s 1 2ξsp ωsp s 1 ω2sp wherein, ωz is the pitch rate; XS is the control column input displacement; s is the Laplace operator; K is equivalent gain; 1/Tθ2 is the zero point of equivalent transfer function; ξ sp is the equivalent short-period damping ratio; ωsp is the equivalent short-period undamped natural frequency; and τ is equivalent time delay. The heading closed -loop frequency response test is generally used to calculate the highorder system frequency response of sideslip angle β to heading command input Xy and of roll rate ωx to lateral command input Xr. The form of heading system’s equivalent lower-order transfer function is usually β ðsÞ Kβ 5 2 Ue2τ eβ s Xy ðsÞ s 1 2ξnd ωnd s 1 ω2ns
 2 2 K s 1 2ξ ω s 1 ω ωx ϕ φ φ s ωx ðsÞ    Ue2τ eωx s 5 Xr ðsÞ s 1 1=Ts s 1 1=TR s2 1 2ξ nd ωnd s 1 ω2nd wherein, β is the sideslip angle; ωx is the roll rate; Xy is the input displacement of pedal; Xr is the lateral input displacement of control column; s is the Laplace operator; Kβ is the equivalent gain; ξnd is the equivalent Dutch roll damping ratio; ωnd is the equivalent Dutch roll undamped natural frequency; τ eβ is the equivalent time delay; Kωx is the equivalent gain; Ts is the equivalent helical mode time constant; TR is the equivalent rolling mode time constant; τ eωx is the equivalent time delay; ξ ϕ is the zero point damping ratio of the equivalent transfer function; ωϕ is the zero point undamped natural frequency of equivalent transfer function. The flight control system contains many nonlinear factors and the closed-loop frequency response characteristics under different amplitude input are different. Therefore, the input amplitude shall be carefully determined and its principle is the same as the amplitude selection principle in the stability margin test.

“Iron bird” integration test of the flight control system 479 The input frequency range of frequency response test is generally 0.1
10 rad/s and the frequency interval is divided by logarithm 10 of every 10 octave. If the amplitude curve still does not fall when ω $ 10 rad/s, the upper limit frequency can be extended by one octave.

6.9.7 Time-domain characteristic test The time-domain characteristic test of the flight control system is conducted under closedloop state. Based on the standard input of control column (wheel) or pedal to the flight control system, the flight control system test analysis system records and analyzes the change course of aircraft response (control plane deflection angle, attitude changes) over time. The standard input is generally in the form of a step (or pulse or bidirectional pulse), and the input amplitude shall include small amplitude, medium amplitude and large amplitude. Time-domain characteristic test shall cover full flight envelope and be conducted under all working modes of the flight control system. The flight motion should better be described by 6-DOF nonlinear equation or by simplified linear equation and the analysis of the results shall consider the impact brought by the simplification of the aircraft motion equation. The time-domain characteristic test includes command input test and time-domain disturbance characteristic test. 1. Command input time-domain characteristic test Step input is taken as an example below to introduce the time-domain characteristic test methods and process: a. Set the excitation signal (or signal generator) of the flight control system test analysis system as a square signal and square signal cycle as 20 s. Its amplitude shall be consistent with the requirements of the control displacement. b. Connect the excitation signal (or signal generator) of the flight control system test analysis system to the control input end of the mechanical displacement signal generator as input control signal. c. Start the flight control system test analysis system to enter signal test recording state and start the flight control system test analysis system to excite signal. d. Use a test sensor to test the control signals including steering wheel displacement, pedal displacement and control column displacement, displacement of aileron mechanical backup actuator, rudder mechanical backup actuator, elevator mechanical backup actuator, as well as deflection angle signals of control planes including aileron, rudder and elevator, and introduce the test signals to flight control system test analysis system for signal processing and analysis.

480 Chapter 6 e. After the test, analyze the excitation signal, steering wheel displacement, aileron mechanical backup actuator displacement, aileron control plane deflection angle, pedal displacement, rudder mechanical backup actuator and rudder control plane deflection angle, control column displacement, elevator mechanical backup actuator displacement and elevator control plane deflection angle recorded in the test of flight control system test analysis system, draw the curves of excitation signal and response signal and get the characteristic parameters such as rise time, steady-state time, oscillation times and overshoot. 2. Time-domain disturbance characteristic test The time-domain disturbance characteristic test includes angle of attack disturbance test, sideslip angle disturbance test, pitch angle disturbance test and roll angle disturbance. a. Angle of attack disturbance test: Apply angle of attack disturbance and record signals including elevator deflection angle, horizontal stabilizer deflection angle, control column displacement, pitch rate, normal overload and angle of attack. Meanwhile, introduce the disturbance signal to flight simulation system and calculate the signal of each node of the link and the control plane deflection signal. Compare the control plane response signals recorded in the test with those calculated by the flight simulation system to determine the conformity of the test results. b. Sideslip angle disturbance test: Apply sideslip angle disturbance and record signals including aileron deflection angle, rudder deflection angle, steering wheel displacement, pedal displacement, roll rate, yaw rate, slant angle and sideslip angle. Meanwhile, introduce the disturbance signal to flight simulation system and calculate the signal of each node of the link and the control plane deflection signal. Compare the control plane response signals recorded in the test with those calculated by the flight simulation system to determine the conformity of the test results. c. Pitch angle disturbance test: Apply pitch angle disturbance and record signals including elevator deflection angle, pitch angle, pitch rate, normal overload and angle of attack. Meanwhile, introduce the disturbance signal to flight simulation system and calculate the signal of each node of the link and the control plane deflection signal. Compare the control plane response signals recorded in the test with those calculated by the flight simulation system to determine the conformity of the test results. d. Roll angle disturbance test: Apply roll angle disturbance and record signals including aileron deflection angle, rudder deflection angle, steering wheel displacement, pedal displacement, roll rate, yaw rate, slant angle and sideslip angle. Meanwhile, introduce the disturbance signal to flight simulation system and calculate the signal of each node of the link and the control plane deflection signal.

“Iron bird” integration test of the flight control system 481 Compare the control plane response signals recorded in the test with those calculated by the flight simulation system to determine the conformity of the test results.

6.9.8 Boundary limit and protection function inspection The fly-by-wire flight control system of large transport aircraft under normal working mode is generally set with active control functions including attitude hold through bar loosening (pitch angle hold function and roll angle hold function), stall protection, overspeed protection, overload protection, pitch angle protection and slant angle protection functions to achieve carefree handling of pilots. The flight boundary limit and protection function inspection aims to verify the realization of the boundary limit and protection functions above and check the performance of flight control system and aircraft dynamic characteristics under the condition of deep saturation. The inspection is conducted with the time-domain characteristic test of the flight control system and its basic principle is to take large-amplitude control input to check the effectiveness of the protection functions. In the flight boundary limit and protection function inspection, the flight control system is set under normal working mode, the flight state and initial values of parameters are set through the flight simulation system and FTI equipment is used to test and record signals. Main test content and methods are shown below: 1. Pitch angle hold function a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly and decidedly push/pull control column to 1/4 stroke, 1/2 stroke, 3/4 stroke and full stroke and hold for 10 s. c. Then, quickly release the control column to complete current test. d. During the operation, use a FTI equipment to record flight altitude, airspeed (Ma number), control column displacement, pitch rate, pitch angle and elevator deflection angle and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time. 2. Roll angle hold function a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly and decidedly rotate the steering wheel in clockwise/anticlockwise direction to 1/4 stroke, 1/2 stroke, 3/4 stroke and full stroke and hold for 6 s.

482 Chapter 6 c. Then, quickly release the steering wheel to complete current test. d. During the operation, use a FTI equipment to record slant angle, sideslip angle, lateral overload, left (right) aileron deflection angle, rudder deflection angle, left (right) spoiler deflection angle, pedal displacement, steering wheel displacement, roll rate, yaw angle and yaw rate and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time. 3. Stall protection function inspection a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly pull the control column back to full stroke and hold for 10 s. c. Then, quickly release the control column to complete current test. d. During the operation, use a FTI equipment to record altitude, airspeed (Ma number), column displacement, angle of attack and track angle and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time. 4. Overspeed protection function inspection a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly push the control column forward to full stroke and hold for 10 s. c. Then, quickly release the control column to complete current test. d. During the operation, use a FTI equipment to record flight altitude, airspeed (Ma number), control column displacement, track angle and normal overload and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time. 5. Overload protection function inspection a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly push the control column forward to full stroke and hold for 10 s. c. Then, quickly release the control column to complete current test. d. During the operation, use a FTI equipment to record flight altitude, airspeed (Ma number), control column displacement, track angle and normal overload and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time.

“Iron bird” integration test of the flight control system 483 6. Pitch angle protection function inspection a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly pull the control column back to full stroke, hold for 10 s and then release it to make it return, and keep it at the neutral position for 5 s. c. Quickly push the control column forward to full stroke, hold for 10 s and then release it to complete current test. d. During the operation, use a FTI equipment to record altitude, airspeed (Ma number), column displacement, pitch angle and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time. 7. Slant angle protection function inspection a. Set the flight control system under normal working mode and set the initial flight state through the flight simulation system. b. Quickly rotate the steering wheel in anticlockwise direction to full stroke, hold for 10 s and then release it to make it return, and keep it at the neutral position for 5 s. c. Quickly rotate the steering wheel in clockwise direction to full stroke, hold for 10 s and then release it to complete current test. d. During the operation, use a FTI equipment to record slant angle, sideslip angle, lateral overload, left/right aileron deflection angle, rudder deflection angle, left/right spoiler deflection angle, pedal displacement, steering wheel displacement, roll rate, yaw angle, yaw rate and other signals. e. After the operation is completed, draw the time-domain response curve of each signal and analyze time-domain characteristic parameters such as overshoot, oscillation and adjustment time.

6.9.9 State and alarm display verification test The working state of the flight control system and the deflection angle of the control plane are displayed through the avionics system PFD, MFD, and EICAS. The deflection angle and working mode of control planes including ground spoiler, multifunctional spoiler, aileron, elevator, horizontal stabilizer and rudder as well as ACE state and hydraulic system state are displayed on the MFD flight control system diagram, as shown in Fig. 6.35. 1. Inspection of state and control plane deflection angle display on flight control system diagram Following operations are mainly conducted for the inspection of the state and control plane deflection angle display of the flight control system on flight control system diagram and the display state shall be observed and inspected on MFD:

484 Chapter 6

Actuator Controller

L1 L2 R1 R2 lFy-by-wire mode

Normal

Hydraulic System

1#

2#

3#

Figure 6.35 Flight control system diagram displayed on MFD.

a. Operate the fly-by-wire flight control system to normal, degraded and simulated backup states respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. b. Set the working state of 1#, 2#, 3#, 4# ACE as well as working state of 1#, 2#, 3# hydraulic systems respectively and observe and inspect whether the state display information is consistent with the actual state on MFD.

“Iron bird” integration test of the flight control system 485 c. Set the aileron under normal state, under single-channel control or hydraulic failure, under double-channel control failure and under double-channel hydraulic failure respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. d. Set the elevator under normal state, under single-channel control or hydraulic failure, under double-channel control failure and under double-channel hydraulic failure respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. e. Set the rudder under normal state, under single-channel control or hydraulic failure, under double-channel control failure and under double-channel hydraulic failure respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. f. Set the spoiler under normal state and under hydraulic failure respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. g. Set the horizontal stabilizer under normal state, under single-channel control or hydraulic failure, under double-channel control failure and under double-channel hydraulic failure respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. h. Set the flaps under normal state, under flaps and slats half speed, under flaps and slats control failure and under flaps and slats drive fault respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. i. Set the slats under normal state, under flaps and slats half speed, under flaps and slats control failure and under flaps and slats drive fault respectively and observe and inspect whether the state display information is consistent with the actual state on MFD. 2. Inspection of trim position and flaps and slats position display on EICAS Following operations are mainly conducted for the inspection of the trim position and flaps and slats position display on EICAS and the display state shall be observed and inspected on EICAS: a. Operate the horizontal stabilizer to trim at different positions and observe and inspect whether the state information is consistent with the actual state on EICAS. b. Operate the aileron to trim at different positions and observe and inspect whether the state information is consistent with the actual state on EICAS. c. Operate to release the flaps and slats at different positions and observe and inspect whether the state information is consistent with the actual state on EICAS. d. Operate to withdraw the flaps and slats at different positions and observe and inspect whether the state information is consistent with the actual state on EICAS.

486 Chapter 6 3. Inspection of working way and mode display of the automatic flight control system on PFD Following operations are mainly conducted for the inspection of working way and mode display of the automatic flight control system on PFD and the display state shall be observed and inspected on PFD: a. Operate the automatic flight control system under AP, FD and synchronous modes respectively and observe and inspect whether the state information is consistent with the actual state on PFD. b. Operate the automatic flight control system under pitch attitude hold, attitude hold, takeoff and landing speed, flight level, vertical navigation and approach modes in longitudinal direction respectively and observe and inspect whether the state information is consistent with the actual state on PFD. c. Operate the automatic flight control system under slant attitude hold, heading hold, heading direction selection, horizontal navigation and heading modes in lateral direction respectively and inspect whether the state information is consistent with the actual state on PFD. d. Operate the autothrottle of automatic flight control system under different thrust force and speeds respectively and observe and inspect whether the state information is consistent with the actual state on PFD. e. Set the automatic flight control system under fault state and observe and inspect whether the warning information (characters) is consistent with the actual state on PFD. 4. Inspection of cockpit warning information display Following operations are mainly conducted for the inspection of cockpit warning information display (including lighting, voice and characters) and the display state shall be observed and inspected on the cockpit warning device: a. Set the takeoff configurations including spoiler configuration, horizontal stabilizer configuration, rudder configuration and flaps (slats) configuration respectively and observe and inspect whether the display state is consistent with the actual state on cockpit warning device. b. Set the speed warning states including stall warning and overspeed warning respectively and observe and inspect whether the display state is consistent with the actual state on cockpit warning device. c. Set the fault warning states (including stabilizer failure, automatic air brake failure, aileron fly-by-wire disconnection, elevator fly-by-wire disconnection, rudder fly-bywire disconnection, flight control system degradation, horizontal stabilizer cutoff, simulated backup mode of flight control system, automatic trim disconnection, spoiler failure, autopilot disconnection, autothrottle disconnection, longitudinal operation disconnection, lateral operation disconnection, loss of ILS Cat II, autopilot default mode, AP synchronization, flaps drive, slats drive, flaps and slats control failure, flaps half speed and slats half speed) respectively and observe and inspect whether the display state is consistent with the actual state on cockpit warning device.

“Iron bird” integration test of the flight control system 487

6.9.10 Failure effect test For flight control system of large transport aircraft with high safety requirements, reducing failure and minimizing the impact of failure is the key to the safety design of the flight control system. Therefore, failure effect test has become one of the main items of the flight control system verification test. In general, the fly-by-wire flight control system will adopt measures such as redundancy design and redundancy management, fault reconstruction and degraded safety. The main fault modes of the fly-by-wire flight control system include power supply fault, hydraulic source fault, control plane clamping stagnation (including at neutral position and limit position), control plane uncommanded motion, control plane failure (control free), trim function fault (uncommanded motion and clamping stagnation), unilateral control stagnation (including at neutral position and limit position), airborne equipment fault (including PFC, ACE, command sensor, feedback sensor, actuator, atmospheric data computer and inertial navigation device). The purpose of failure effect test is to check whether the functions and performance of the aircraft and flight control system under the above fault modes meet design requirements and aircraft safety requirements. The failure effect test methods and process are described as follows: 1. Power supply fault Simulate the single faults and fault combinations of each bus bar to make PFC and ACE work under a single power supply failure or power supply failure combinations under redundancy power supply and make the flight control system work under the state of emergency power supply, and check whether the power supply redundancy and safety design of the flight control system can meet the flying safety requirements. 2. Hydraulic source fault Simulate the single low pressure or low pressure combinations of redundancy hydraulic source and check the working state of the fly-by-wire flight control system, and check the working state of the fly-by-wire flight control system under emergency state of hydraulic source. 3. Control plane clamping stagnation Set the flight control system under closed-loop working state and in the flight simulation process, set single control plane clamping stagnation fault or control plane clamping stagnation fault combinations of aileron, elevator, rudder, spoiler and horizontal stabilizer, check the working condition of the flight control system after control plane clamping stagnation as well as the consistency with FHA analysis and evaluation results. The control plane clamping stagnation can be achieved by setting the control plane deflection in flight simulation system. 4. Control plane uncommanded motion Set the flight control system under closed-loop working state and in the flight simulation process, set single control plane uncommanded motion fault or control plane

488 Chapter 6

5.

6.

7.

8.

uncommanded motion fault combinations of aileron, elevator, rudder, spoiler and horizontal stabilizer, check the working condition of the flight control system during control plane uncommanded motion as well as the consistency with FHA analysis and evaluation results. The control plane uncommanded motion can be achieved by setting the control plane deflection in flight simulation system. Control plane failure (control free) Set the flight control system under closed-loop working state and in the flight simulation process, set single control plane failure or control plane failure combinations of aileron, elevator, rudder, spoiler and horizontal stabilizer, check the working condition of the flight control system during control plane failure as well as the consistency with FHA analysis and evaluation results. The control plane failure can be achieved by setting the control plane deflection in flight simulation system. Trim function fault Set the flight control system under closed-loop working state and in the flight simulation process, set trim stagnation or uncommanded motion of rudder, aileron and horizontal stabilizer, check the working condition of the flight control system under trim function failure as well as the consistency with FHA analysis and evaluation results. The trim function fault can be achieved by setting the trim command (or superposing control command) in flight simulation system. Unilateral control stagnation Set the stagnation faults of control column, steering wheel and pedal of the pilot (copilot) and verify the arming/override functions of the system under unilateral control stagnation. Airborne equipment fault Set the flight control system under closed-loop working state and in the flight simulation process, set the PFC, ACE, command sensor, feedback sensor and actuator as faulted and check the working condition of the flight control system under airborne equipment fault so as to evaluate the isolation and field processing capability of the flight control system. The airborne equipment fault can be achieved through equipment power-off and simulating out of tolerance of redundancy signals. In the failure effect test, record the state information and warning information of flight control system reported by PFC, ACE, FSECU, and AFCC, pilot control command, control plane deflection angle and flight state, etc. and analyze whether the working state change of flight control system and whether the aircraft transient state meet the design requirements after removal of the failure of flight control system. In the failure effect test, any of the following circumstances will be deemed as fault: failure to pass the PBIT, failure information of flight control system, airborne equipment affecting the functions, performance and structural integrity of flight control system, and rupture, fracture or damage of structural parts or components.

“Iron bird” integration test of the flight control system 489

6.10 “Iron bird” man
machine combined test 6.10.1 Overview Chapter 4 introduces the control law and flying quality evaluation in the engineering simulator environment with the participation of pilot and Section 6.3 introduces the evaluation of man
machine ergonomics of cockpit control unit with the participation of pilot. In the two evaluation tests above, through the joint work of the pilot and the designer, the pilot control hardware environment that meets the requirements for man
machine ergonomics, the aircraft stability performance that meets the aircraft flying quality design requirements, and the control logic and control law configuration and parameters of the flight control system are confirmed. As we said above, the engineering simulator test is a typical person-in-loop test of the flight control system and the flight control system “iron bird” integration test is a typical hard-inloop test. The engineering simulator test achieves the coupling verification between the flight control system and aircraft and between the pilot and aircraft while the man
machine combined test on the “iron bird” integration test bed achieves the coupling verification between the pilot, flight control system and aircraft (person hard plane). The man
machine combined test is the highest level of ground laboratory test for manned aircraft. In the “Iron bird” man
machine combined test, the scheduled flight missions and flight outline are mainly followed, a pilot pilots the aircraft and the subjective evaluation of the pilot takes the main part. Of course, the pilot must have solid knowledge of flight mechanics and extensive experience in evaluation, preferably a pilot and copilot of the aircraft for maiden flight. The “iron bird” man
machine combined test mainly includes: 1. 2. 3. 4. 5.

Man
machine ergonomics evaluation of cockpit control unit; Control (handling) and display function evaluation; Aircraft control stability evaluation; Fault mode and its effects evaluation; Maiden flight profile adaptability and emergency response and drill.

Aircraft control stability evaluation, as a key point of the “iron bird” man
machine combined test, mainly evaluates the control law and control logic of the flight control system indirectly through the evaluation of the flying quality of the aircraft. Its main content includes: 1. evaluation of basic control functions, including the control functions under various flight states and aircraft configurations during takeoff, landing and flight in air, such as control augmentation, autopilot, navigation and flight director;

490 Chapter 6 2. evaluation of control force and control habits, including the column force performance and automatic trim performance during takeoff and landing; 3. evaluation of conversion between different states, including conversion control and effects of modal conversion to transient state; 4. evaluation of various boundary limits, such as angle of attack limit and overload limit and also the angular rate or angular acceleration limit when entering and exiting a turn; 5. evaluation of fault reconstruction and its effects; 6. PIO check, aiming to check if there will be PIO.

6.10.2 Test principle “Iron bird” man
machine combined test is a ground flight that takes the pilot as the center and the aircraft as the control object. It is conducted in an environment with more aircraft system hardware. The layout of aircraft airborne system and equipment shall be as realistic as possible, ensuring that it is consistent with the state of the aircraft for maiden flight and the aerodynamic load of the aircraft in the flight process shall also be simulated. Generally, the airborne equipment of airborne systems such as the flight control system, hydraulic system, landing gear control system and avionics system on the “iron bird” integrated test bed as well as their layout on the test bed shall be consistent with the state of the aircraft for maiden flight. All airborne equipment shall be installed test pieces (normally called “S” type ground test pieces). For airborne equipment that can only work through the exciter, such as atmospheric data computer, inertial/satellite navigation equipment, radio altimeter, intelligent probe, integrated navigation equipment, angular rate sensor and overload sensor, relevant excitation device shall be configured and put in the test. The basic principle of “iron bird” man
machine combined test is to put a pilot in the cockpit on the “iron bird” integrated test bed and then the pilot will conduct flight control and make display evaluation according to the flight task list prepared according to the test outline and finally give evaluation conclusion. Meanwhile, relevant flight parameters and data are recorded through the flight control system test analysis system. The test principle of the “iron bird” man
machine combined test is shown in Fig. 6.36. The “iron bird” man
machine combined test is a complex and huge test that involves almost all airborne systems and test support equipment on the “iron bird” integrated test bed. Due to the participation of many pilots, higher requirements are proposed to the stability reliability and fidelity of the test system. As a large number of people and host machine, test flight unit, manufacturing unit, user representatives and airborne equipment development unit participate in the test, the test management becomes more complex. The communication between the flight simulation system and the avionics excitation device is realized through the reflective memory optical fiber network. The communication between

“Iron bird” integration test of the flight control system 491

Figure 6.36 Principle of pilot-in-loop test on “Iron Bird”.

fly-by-wire flight control computer PFC, electromechanical management computer and avionics integrated processor is realized by extending the airborne AFDX data bus. The communication between the atmospheric data computer, inertial/satellite navigation equipment, radio altimeter, intelligent probe, combined navigation device and fly-by-wire flight control computer PFC, AFCC, FSECU is realized by extending the airborne ARINC 429 data bus. The pilot controls the aircraft and the flight control system test analysis system detects the aircraft control plane deflection angle signal, landing gear control signal, hatch door signal and engine throttle position signal through the test sensor and transmits them to the flight simulation system. The flight simulation system solves the aircraft dynamics equation and the engine thrust equation and the engine speed signal calculated is sent to the engine drive pump speed simulation system to control the operation of the hydraulic pump driving device and ensure that the working state of the hydraulic system is consistent with the working state of the engine. The parameters such as the motion attitude, motion state and position of the aircraft calculated are taken as the command signals of linear acceleration turntable, three-axis rate turntable, atmospheric data simulation exciter, inertial navigation simulation exciter and intelligent probe simulation exciter to control the operation of various simulation devices, drive three-axis overload sensor and angular rate gyro assembly and excite atmospheric data computer, inertial/satellite navigation equipment, radio

492 Chapter 6 altimeter, intelligent probe and integrated navigation equipment, so as to provide closedloop feedback signal for the flight control system. The parameters such as the motion attitude, motion state and position of the aircraft got through calculation are transmitted by the flight simulation system to excitation devices such as atmospheric data computer, inertial/satellite navigation equipment, radio altimeter, intelligent probe and integrated navigation equipment through the reflected memory network to control the operation of excitation devices. The parameters got through the calculation by the atmospheric data computer such as pressure altitude, true airspeed, takeoff and landing speed, Mach number, true angle of attack, static temperature and sideslip angle as well as the parameters got by the inertial navigation system such as pitch angle, roll angle, true heading angle, aircraft body lateral angular rate, aircraft body vertical angular rate, aircraft body longitudinal angular rate, aircraft body lateral linear acceleration, aircraft body vertical acceleration and aircraft body longitudinal acceleration and ground speed are provided to the flight control system. The atmospheric parameters obtained by the total pressure/static pressure intelligent probe such as total pressure, left static pressure, right static pressure, total temperature and angle of attack are provided to the flight control system. The “iron bird” man
machine combined test mainly includes the takeoff, landing and heading flight under normal working state and fault state. The flight process is briefly described as follows: 1. Takeoff and landing flight simulation test a. Keep the engine in the “maximum” state, take off to run, lift the front wheels off the ground and climb to a certain altitude, and check the takeoff performance of the fly-by-wire flight control system, especially the control force and column displacement to lift the front wheels, the transient response of the aircraft in landing gear withdrawing and the capability to hold and change the aircraft attitude. b. Keep the engine in a “slow” state, glide, level and touch the ground from a given altitude, and check the landing performance, especially the attitude and track control capability, and experience and analyze the PIO trend. c. Possible crosswind conditions shall be added during takeoff and landing flight simulation. d. For flight control system adopting an emergency backup working mode, the approach and landing simulation flight test shall also be carried out under the emergency backup working mode. 2. Flight simulation test under general maneuvering conditions a. General maneuvering conditions include horizontal acceleration and deceleration, rise and fall, S-shaped turning, hover, general aerobatics, control column and pedal step and pulse control.

“Iron bird” integration test of the flight control system 493 b. For test under general maneuvering conditions and typical flight states, the flight state shall be selected according to the analysis of the flight control system and aerodynamic performance. c. All working modes of the flight control system shall be tested and the modal conversion to transient transition under each working mode and different flight states shall be given with special attention. 3. Failure effect test a. Introduce single-channel fault and multichannel fault, check the transient state of faults and study the pilot’s correction action and its consequences. b. The test shall be carried out under normal working mode and fault mode of the flight control system. c. Typical fault modes include failure of elevator at one side, failure of aileron at one side, failure of rudder or failure of flaps (slats) release and shutdown. 4. Software evaluation and flight control system trustworthiness test Given several flight profiles, the pilot (or test flight engineer) simulates the flight according to operation procedures, records the problems exposed in the test and makes evaluation. Confirm that the flight control system has software and hardware working normally when it is in continuous operation and it meets the use requirements. During the flight simulation, use the flight control system test analysis system to record altitude, indicated airspeed, rod displacement, wheel displacement, pedal displacement, rod force, wheel force, pedal force, elevator deflection angle, aileron deflection angle, spoiler deflection angle, rudder deflection angle, horizontal stabilizer deflection angle, pitch rate, roll rate, yaw rate, pitch angle, roll angle, yaw angle, angle of attack, sideslip angle, normal overload, track dip angle, track deflection angle, landing gear position and flaps (slats) control plane deflection angle.

6.10.3 Man
machine combined test of takeoff and landing and free flight The man
machine combined test of takeoff and landing includes normal takeoff, normal landing and deviation-correction landing. 1. The man
machine combined test of normal takeoff evaluates the flying quality of an aircraft during normal takeoff, including the course keeping performance in ground sliding, front wheel lifting performance, climb performance, as well as aircraft response and control performance during landing gear and flaps (slats) withdrawing process. 2. The man
machine combined test of normal landing and deviation-correction landing evaluates the flying quality of an aircraft during approach and landing process, including PIO trend, glide performance, as well as aircraft response and control performance during landing gear and flaps (slats) release process, leveling performance and course keeping performance in ground sliding.

494 Chapter 6 Main test methods and process of the man
machine combined test of takeoff and landing are as follows: Start test support equipment such as the flight simulation system, select the flight state and set the state of flight control system, start the flight control system test analysis system and get prepared for the test. 1. Man
machine combined test of takeoff a. The pilot operates the flaps and slats control handle to gear “3” (flaps 25 ), puts the landing gear control handle at “release” position and engine throttle lever at “slow” position. b. Then, push the throttle lever to takeoff thrust position and when the monitored indicated airspeed reaches vR, pull the control column back to make the aircraft lift the front wheels at 3 /s and withdraw the landing gear under positive rate of rise. c. When the aircraft climbs above 120 m and the airspeed reaches vREF 1 80 km/h, put the flaps and slats control handle at gear “2.” When the airspeed reaches vREF 1 100 km/h, put the flaps and slats control handle at gear “1.” When the airspeed reaches vREF 1 120 km/h, put the flaps and slats control handle at gear “0.” Finally, hold the altitude at 500 m to make the aircraft fly in a straight line stably. d. After the flight, the pilot gives evaluation comments according to the flight experience. 2. Man
machine combined test of landing a. The pilot operates the flaps and slats control handle to gear “0” and puts the landing gear control handle at “withdraw” position and engine throttle lever at specified position. The pilot controls the aircraft to fly in a straight line stably. b. Then, operate the flaps and slats control handle to gear “1,” hold the altitude and decrease the speed. c. When the airspeed of the aircraft reaches vREF 1 100 km/h, put the flaps and slats control handle at gear “2,” put the landing gear control handle at “release” position, hold the aircraft altitude and continue to decrease the speed. d. When the airspeed of the aircraft reaches vREF 1 80 km/h, put the flaps and slats control handle at gear “3,” continue to decrease the speed and align to the runway. e. When the airspeed of the aircraft reaches vREF 1 60 km/h, put the landing gear control handle at gear “4” and continue to decrease the speed. f. When the airspeed of the aircraft reaches vREF 1 40 km/h, put the flaps and slats control handle at gear “5,” control the aircraft to make it glide at 3 glide angle, approach the runway at an altitude of 15 m and speed of vREF, pull the control column back and the main landing gear lands down to the ground, release the control column and the front wheels land down to the ground, withdraw the throttle, pull the thrust reverser and operate the aircraft to make it glide in central runway.

“Iron bird” integration test of the flight control system 495 g. After the flight, the pilot gives evaluation comments according to the flight experience. 3. Man
machine combined test of free flight The man
machine combined test of free flight selects any flight states and the fight states shall cover various weight/center of gravity and configuration combinations. The flight states can cover the boundary points and point of interest of the envelope. The flight control system adopts normal working mode, simulated backup working mode, and the pilot can complete the maneuvers, missions or flight under failures which he is interested in.

6.10.4 Man
machine combined test of mode conversion The man
machine combined test of mode conversion evaluates the transient response of an aircraft during the conversion between normal working mode and simulated backup working mode, between simulated backup working mode and mechanical backup working mode, and between normal working mode and mechanical backup working mode. The man
machine combined test of mode conversion is carried out by controlling the aircraft to take off and land. In the steady state of the aircraft, such as level flight, climbing, gliding and steady turning, the control column, steering wheel or pedal are operated to make the flight control system convert between the normal working mode, simulated backup working mode and mechanical backup working mode so as to experience the transient response of the aircraft during mode conversion. Start test support equipment such as the flight simulation system, select the flight state and set the state of flight control system, start the flight control system test analysis system and get prepared for the test. 1. The pilot operates the flaps and slats control handle to gear “0” and puts the landing gear control handle at “withdraw” position and engine throttle lever at specified position. Select normal working mode for the flight control system. 2. After completion of the automatic trim of aircraft, release the control column to operate the aircraft to fly in a straight line stably under given altitude and speed. 3. During the flight, switch the flight control system to simulated backup working mode and experience the transient response of the aircraft. 4. Under the simulated backup working mode, after completion of the manual trim of aircraft, release the control column to operate the aircraft to fly in a straight line stably under given altitude and speed. Then, switch the flight control system to normal working mode and experience the transient response of the aircraft. 5. After the flight, the pilot gives evaluation comments according to the flight experience and the display during mode conversion of the flight control system.

496 Chapter 6 The test methods and process for conversion between the simulated backup working mode and mechanical backup working mode and between normal fault mode and mechanical backup working mode are similar.

6.10.5 Man
machine combined test of failure effect The man
machine combined test of failure effect mainly evaluates the failures of atmospheric data, radio altitude and angle of attack signals, failures of angular rate and overload signal or control failure, function loss or uncommanded motion of flaps (slats), horizontal stabilizer, elevator, aileron, rudder and spoiler, or the power supply fault, hydraulic source fault, engine fault or the stability and maneuverability of the aircraft in each flight stage under the landing gear control, so as to find out the fault disposal measures. Start up test support equipment such as the flight simulation system, select the flight state and set the state of flight control system, start the flight control system test analysis system and get prepared for the test. 1. Takeoff stage a. The pilot operates the flaps and slats control handle to gear “3” (flaps 25 ), puts the landing gear control handle at “release” position and engine throttle lever at “slow” position. b. Then, push the throttle lever to takeoff thrust position and when the monitored indicated airspeed reaches vR, pull the control column back to make the aircraft lift the front wheels at 3 /s and withdraw the landing gear when the altitude reaches 10.7 m. c. When the aircraft climbs above 120 m and the airspeed reaches vREF 1 80 km/h, put the flaps and slats control handle at gear “2.” d. When the airspeed reaches vREF 1 100 km/h, put the flaps and slats control handle at gear “1.” e. When the airspeed reaches vREF 1 120 km/h, put the flaps and slats control handle at gear “0,” put the landing gear control handle at “withdraw” position and engine throttle lever at specified position. Finally, hold the speed at 600 km/h and make the aircraft fly in a straight line stably and enter the cruise state. 2. Landing stage a. The pilot operates the aircraft to fly at an altitude of 500 m and airspeed of vREF 1 120 km/h, put the flaps and slats control handle at gear “1,” hold the altitude and decrease the speed. b. When the airspeed of the aircraft reaches vREF 1 100 km/h, put the flaps and slats control handle at gear “2,” put the landing gear control handle at “release” position, hold the aircraft altitude and continue to decrease the speed.

“Iron bird” integration test of the flight control system 497 c. When the airspeed of the aircraft reaches vREF 1 80 km/h, put the flaps and slats control handle at gear “3,” continue to decrease the speed and align to the runway. d. When the airspeed of the aircraft reaches vREF 1 60 km/h, put the flaps and slats control handle at gear “4” and continue to decrease the speed e. When the airspeed of the aircraft reaches vREF 1 40 km/h, put the flaps and slats control handle at gear “5,” control the aircraft to make it glide at 3 glide angle, approach the runway at an altitude of 15 m and speed of vREF, pull the control column back and the main landing gear lands down to the ground, release the control column and the front wheels land down to the ground, withdraw the throttle, pull the thrust reverser and operate the aircraft to make it glide in central runway. f. After the flight, the pilot gives evaluation comments according to the flight experience. In the flight test above, the failures of atmospheric data, radio altitude and angle of attack signals, failures of angular rate and overload signal or control failure, function loss or uncommanded motion of flaps (slats), horizontal stabilizer, elevator, aileron, rudder and spoiler, or the power supply fault, hydraulic source fault, engine fault or the landing gear fault are set respectively.

6.10.6 Test task list To better organize the “iron bird” man
machine combined test, a special test task list is prepared for each test item regulated by the test task to clearly show current test flight status, flight control system working mode, flight crew and flight test engineer, describe the flight process, state setting and operating rules in the test and record the test date and time. The basic form and content of the test task list are shown in Table 6.2. Table 6.2: Test task list. No.: 0101 Date Start Time End Time Pilot Test flight engineer Subject: normal takeoff and landing Mode of flight control system: normal working mode Fault Setting: None Weight: 120 t Center of gravity: 35% Takeoff: v1: 192 km/h vR: 200 km/h v2: 220 km/h Flaps 15: 310 km/h Flaps 1: 330 km/h Withdraw: 350 km/h Landing: flaps 1: 380 km/h Flaps 15: 350 km/h Flaps 25: 330 km/h Wheel release: 300 km/h Flaps 30: 300 km/h Flaps 40: 260 km/h vREF: 220 km/h Evaluation comments: Pilot:

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6.11 Test results evaluation of the flight control system “iron bird” integration test During the “iron bird” integration test of the flight control system, a large number of test results such as the static data, dynamic data and flying quality data of the flight control system will be generated. How to evaluate these test results and make a correct evaluation of the aircraft and flight control system is a subject worth studying. Based on previous research and practical experience, this section puts forward some judging criteria, evaluation requirements, content and methods, etc. The results of the “iron bird” integration test of the flight control system shall be analyzed and evaluated according to design requirements of flight control system, design requirements and design scheme of the flight control system and its subsystems, design requirements of aircraft flying quality, technical conditions for installation and debugging of the flight control system and its subsystems, technical conditions for power-on inspection of the flight control system and its subsystems and standards including GJB185
1986, GJB2874-1997, GJB2878-1997, GJB2191-1994, GJB3819-1999, GJB1690
1993 and CCAR 25-R4 shall be taken as references.

6.11.1 Test results evaluation of the machinery control system 1. Evaluation of polarity inspection results The polarity inspection is carried out respectively for aileron, rudder, elevator and horizontal stabilizer control channels. In the polarity inspection, the control column (or steering wheel, pedal, horizontal stabilizer trim handle) is stably and slowly operated and when it moves to half of the full stroke, it will be stopped and kept stable, and then the motion direction of control plane (aileron, rudder, elevator or horizontal stabilizer) will be checked to see whether it is consistent with the design requirements. 2. Evaluation of control limit inspection results The control limit inspection mainly checks the control limit of aileron, rudder and elevator. In the control limit inspection, the control column (or pedal or steering wheel) is stably and slowly operated and when it moves to the full stroke, it will be stopped and kept stable, and then the motion stroke of control plane (aileron, rudder, elevator) will be checked to see whether it is consistent with the design requirements. 3. Evaluation of return performance test results The return performance test is respectively carried out for aileron, rudder and elevator. In the test, the control column in cockpit (or pedal or steering wheel) is stably and slowly operated and when it moves to half of the full stroke, the control column (or pedal or steering wheel) will be released. After the operation is stopped and the control

“Iron bird” integration test of the flight control system 499 plane stabilizes, the stop position of control plane (aileron, rudder and elevator) will be checked to see whether it is consistent with the design requirements. 4. Evaluation of control performance test results The control performance test covers the control command-control plane relation, the control force
displacement relation, system clearance and starting force. Aileron control channel is taken as an example below to introduce the evaluation method of the results of the control performance test. a. Stably and slowly operate the steering wheel in cockpit (or pedal or control the column) in full stroke in a sequence of “clockwise to limit—return—anticlockwise to limit—return.” b. Meanwhile, record the steering wheel (or pedal or control column) control displacement and control force signals, as well as the control plane (aileron, rudder and elevator) deflection angle signals. c. After the test is completed, analyze and process the recorded test data and obtain the steering wheel control force
wheel displacement relation curve, and then get starting force and other parameters through analysis. d. Further get the steering wheel displacement
control plane deflection angle relation curve and analyze to get the aileron control channel clearance and other parameters. e. Compare and analyze the test curve and design curve above to evaluate whether they meet the design requirements. The methods to evaluate the results of the control performance test of elevator control channel and rudder control channel are similar. Fig. 6.37 shows a typical steering wheel control force- wheel displacement relation curve and Fig. 6.38 shows a typical steering wheel displacement
control plane deflection angle relation curve.

6.11.2 Test results evaluation of the fly-by-wire flight control system 6.11.2.1 Evaluation of zero position inspection results The zero position inspection of the flight control system mainly checks the zero position of control column, steering wheel and pedal command displacement and force as well as the aileron, rudder, elevator, horizontal stabilizer and spoiler. The inspection results will be compared with the data required by the technical conditions for power-on inspection (or technical conditions for installation and debugging) and they shall meet design requirements. 6.11.2.2 Evaluation of polarity and stroke inspection results Under different working modes, the polarity and stroke inspection of the flight control system mainly checks the output of control column to the elevator, of steering wheel to aileron, of pedal to rudder of three-axis angular rate gyro assembly to each control plane and of three-axis overload sensor to each control plane. The inspection results will be compared with the data

500 Chapter 6

Figure 6.37 Steering wheel control force
wheel displacement relation curve.

Figure 6.38 Steering wheel displacement
control plane deflection angle relation curve.

“Iron bird” integration test of the flight control system 501 required by the technical conditions for power-on inspection (or technical conditions for installation and debugging) and they shall meet design requirements. 6.11.2.3 Evaluation of display and warning function inspection results In the display and warning function inspection, the information reported by the flight control system to the avionics display control system MFD are checked by setting different working modes or fault modes of the flight control system, and the display results shall be consistent with the working mode of the flight control system. Aileron control channel is taken as an example below to introduce the methods used for evaluate the results of display and warning function inspection of flight control system. The results evaluation methods of other control channels are similar. 1. When the aileron control channel works normally, the green triangular symbol with scale shows the deflection angle of the aileron. 2. If the aileron deflection angle sensor has fault, the triangular symbol of the aileron disappears. 3. If the aileron actuator integrated control valve has fault, the green box will change to a green frame. 4. When the aileron fly-by-wire control channel has failure, the green box will change to yellow box. 5. When the ground hydraulic energy sources 2# and 3# have failure, the green box will change to red box. 6. When the flight control system computer has power failure, the aileron green box will change to carmine box. The flight control diagram of the display and warning function inspection of the flight control system is shown in Fig. 6.35. 6.11.2.4 Evaluation of frequency-domain characteristic test results The frequency-domain characteristic test is also called closed-loop frequency sweep test. The results of frequency-domain characteristic test are usually represented by a Bode diagram, that is, amplitude-frequency characteristic curve and phase-frequency characteristic curve. By analyzing the amplitude-frequency characteristic curve and phasefrequency characteristic curve, the amplitude gain and phase margin used for evaluating the stability of the flight control system are obtained. The results and evaluation of the results of longitudinal and heading frequency-domain characteristic tests are described as below. 1. Evaluation of longitudinal frequency-domain characteristic test results In the longitudinal frequency-domain characteristic test, the frequency response characteristic curve obtained from the test of flight control system test analysis system is compared with the simulation curve (in the same coordinate system) to determine the consistency between them and make further analysis.

502 Chapter 6 Table 6.3: Equivalent matching results of longitudinal control channel. Aircraft configuration

ζ sp

ωsp

Cruise configuration Takeoff configuration Landing configuration

0.82 0.75 0.75

2.82 2.80 2.84

1 Tθ2

1.17 1.70 2.58

τϑ

τ ny

Mismatch degree

0.12 0.11 0.11

0.16 0.14 0.14

4.39 3.08 8.70

Table 6.4: Control anticipation parameter calculation results. Flight state

ny =α

ωnsp

Control anticipation parameter

Quality level

Cruise configuration Takeoff configuration Landing configuration

9.86 14.32 21.74

2.82 2.80 2.84

0.81 0.55 0.37

First level First level First level

a. ζsp ; ωsp ; 1TU2; τU; τ ny : τ s ny ðsÞ 1 1=Tϑ2 Þe- ϑ ðsÞ 5 sk2ϑ1ðs 2ζ Based on ωFezðsÞ ωsp s 1 ω2 and Fe ðsÞ 5 sp

sp

kny e- n 1 2ζ sp ωsp s 1 ω2sp , τ s

s2

equivalent matching is

carried out for test data. Table 6.3 shows the equivalent matching results of longitudinal control channel of the aircraft under given configuration. b. CAP (Control Anticipation Parameter) value ω2sp ω2 In the calculation of CAP, the CAP is simplified as CAP 5 ðny =αÞ  VU sp1 . g Tϑ2

According to the CAP solution formula, the calculated CAP results of the longitudinal control channel of the aircraft under a given aircraft configuration can be obtained, as shown in Table 6.4. CAP evaluation results of A kind of flight stage value are shown in Fig. 6.39. As can be seen from Fig. 6.39, Tables 6.3 and 6.4, the CAP values of longitudinal control channel under the selected aircraft configuration meet the first-level quality requirements in the takeoff, landing and cruise stages. c. Short-period pitch response The short-period pitch response evaluates the effects of pitch attitude to pitch control input response based on ωnspT2Bζ sp. The calculation results of short-period pitch response under selected aircraft configuration in A kind of flight stage are shown in Table 6.5 and the evaluation results of short-period pitch response in A kind of flight stage are shown in Fig. 6.40. According to the requirement of short-period damping ratio, the range of ζ sp corresponding to the first level quality in the takeoff and landing stages is 0.35 # ζ sp # 1.30 and the range of ζ sp corresponding to the first level quality in the cruise stage is 0.3 # ζ sp # 2. The short-period damping under selected aircraft configuration is shown in Table 6.6. It can be seen that the damping ratio under the selected aircraft configuration from longitudinal cruise, takeoff and landing stages meets the first level flying quality requirements.

“Iron bird” integration test of the flight control system 503

Figure 6.39 Control anticipation parameter evaluation results at a kind of flight stage. Table 6.5: Calculation results of short-period pitch response. Flight state

ωnsp

1=Tϑ2

ωnsp Tϑ2

ζ sp

Quality level

Cruise configuration Takeoff configuration Landing configuration

2.82 2.80 2.84

1.17 1.70 2.58

2.41 1.65 1.10

0.82 0.75 0.75

First level First level First level

2. Evaluation of heading frequency-domain characteristic test results In the evaluation of the results of the heading frequency-domain characteristic test, the equivalent matching within the frequency range of 0.1
10 rad/s is carried out for test data. The analog-matching form is shown below: h i 2 2 2τ γ s K S 1 2ζ ω s 1 ω γ γ γ γ e γ    5 Dx s 1 1=TS S 1 1=Tr s2 1 2ξd ωd s 1 ω2d     Aβ s 1 1=Tβ 1 s 1 1=Tβ 2 s 1 1=Tβ3 e2τ β s β    5  Dy s 1 1=Ts s 1 1=Tr s2 1 2ξd ωd s 1 ω2d

504 Chapter 6

Figure 6.40 Evaluation of short-period pitch response in a kind of flight stage. Table 6.6: Damping ratio calculation results. Flight state Cruise configuration Takeoff configuration Landing configuration

ζ sp

Quality level

0.82 0.75 0.75

First level First level First level

Table 6.7: Heading equivalent matching results. Flight state Cruise configuration Takeoff configuration Landing configuration

Tr

ζd

ωnd

ξ d ωnd

τγ

τβ

mis

0.42 0.30 0.38

0.44 0.64 0.55

1.04 1.64 1.32

0.46 1.05 0.73

0.09 0.11 0.10

0.08 0.04 0.06

3.45 4.88 5.09

According to the formula and test results above, the analog-matching results of the equivalent matching of heading test data under selected aircraft configuration are shown in Table 6.7. According to the equivalent matching results in Table 6.7, such as Tr, ζ d, ωnd, ζ dωnd, τ γ and τ β, it can be evaluated whether they meet the flying quality design requirements. If Tr , 1.4, the first level quality requirements of time constant under roll

“Iron bird” integration test of the flight control system 505 mode are met. By the same method, the damping ratio of Dutch roll mode, the natural frequency of Dutch roll mode and the lateral equivalent delay can be evaluated to see whether they meet the flying quality requirements. If the Dutch roll frequency of all aircraft configurations is greater than 0.7 rad/s, it meets the requirement that the frequency of the Dutch roll root is not less than 0.4 rad/s and meets the first level flying quality requirements. 6.11.2.5 Evaluation of modal conversion test results During the modal conversion test of the flight control system, the roll rate, normal overload and lateral overload in the modal conversion process are recorded, the recorded test data are analyzed and processed, and the maximum values of normal overload and roll rate transient changes in 2 s of the modal conversion are calculated. The design specification of the flight control system stipulates that the maximum value of transient motion change shall not exceed following values in at least 2 s after the modal conversion of the flight control system. 1. The normal linear acceleration increment 6 0.5 g at the pilot seat. 2. The lateral linear acceleration increment 6 0.5 g and roll rate 6 10 /s. Analyze and study the data recorded in the test to determine whether the normal overload, roll rate and lateral overload transient changes meet the requirements above. 6.11.2.6 Evaluation of transmission ratio test results In the transmission ratio test, the relationship between the input command signal or the feedback signal of transmission link of flight control system and the actuator control command signal or control plane deflection signal output by the transmission link is recorded. The results of the transmission ratio test can be evaluated by comparing the consistency of the output actuator control command signal or control plane deflection signal when the actual flight control system and the flight control system simulation model are under same input command signal or feedback signal. In the transmission ratio test, the input command signal or feedback signal is applied to the input of the flight control system and introduced into the simulation model of the flight control system. Meanwhile, the actuator control command signal or control plane deflection angle signal output from the transmission link of the actual flight control system and its simulation model, as well as the intermediate point signal of the control link, are excited. After completion of test, recorded data in the test are analyzed and processed to generate graphic curves of the input command signal or feedback signal of the transmission link, the output actuator control commands or control plane deflection angle signal and the intermediate point signal of the control link. Compare the graphic curves of signals of the flight control system and its simulation model under same input, output or intermediate point

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Figure 6.41 Evaluation of transmission ratio test results. Table 6.8: Requirements for amplitude margin and phase margin. Modal frequency fM/Hz fM , 0.06 0.06 , fM , first-order elastic modal frequency fM. first-order elastic modal frequency

From minimum service speed to maximum service speed (as stipulated in GJB185
1986) Gain margin 5 6 4.5 dB Phase margin 5 6 30 Gain margin 5 6 6 dB Phase margin 5 6 45 Gain margin 5 6 8.0 dB Phase margin 5 6 60

of the control link to judge their consistency. If the maximum difference is less than 10% (the allowable difference between test results and simulation results should be determined according to the actual system), it can be judged that the actual test signal is consistent with the simulation signal. Fig. 6.41 shows the transmission ratio test curves of the elevator control channel and the horizontal stabilizer control channel under the selected aircraft configuration. 6.11.2.7 Evaluation of stability margin test results For a fly-by-wire flight control system including the aircraft, stability is extremely important. Generally speaking, the stability margin shall be subject to the GJB185
1986, namely amplitude margin and phase margin, as shown in Table 6.8.

“Iron bird” integration test of the flight control system 507 In the evaluation of the stability margin test results, the consistency of the frequency response characteristics of the output actuator control command signal or control plane deflection angle signal is compared when the actual flight control system and the flight control system simulation model are under the same input command signal or feedback signal. In the stability margin test, the input command signal is applied to the input of the flight control system and introduced into the simulation model of the flight control system. Meanwhile, the actuator control command signal or control plane deflection angle signal output from the transmission link of the actual flight control system and its simulation model, as well as the intermediate point signal of the control link, are excited. After completion of test, recorded data in the test are analyzed and processed to generate frequency response characteristic curves of the input command signal of the transmission link, the output actuator control commands or control plane deflection angle signal and the intermediate point signal of the control link. Compare the frequency response characteristic curves of signals of the flight control system and its simulation model under same input, output or intermediate point of the control link. If the maximum difference is less than 10% (the allowable difference between test results and simulation results should be determined according to the actual system), it can be judged that the actual test signal is consistent with the simulation signal.

6.11.3 Results evaluation of the man
machine combined test The flying quality of man
machine combined test is mainly evaluated with the Cooper
Harper scale, as shown in Fig. 4.5 in Chapter 4 of this book. Its core is considered from two aspects, that is, aircraft maneuverability and the burden of the pilot to complete various flight missions. The characteristics of the aircraft and 10 different evaluation scales for the pilot required in selected missions or operations are described below. These 10 evaluation scales can be classified into 3 large grades: grade 1 (PR 5 1
3), grade 2 (PR 5 4
6), and grade 3 (PR 5 7
9). The specific meanings of the PR values of the Cooper-Harper scale are as follows: 1. PR 5 1, it means an excellent flying quality, very satisfactory, and various links are ideal and coordinated. 2. PR 5 2
3, it means the task can be well done but there is a little discomfort and improvement is expected. The discomfort here may be caused by the pilot’s different hobbies or other less serious issues. 3. PR 5 4
6, it means that the system or aircraft must be improved. Its remarkable feature is that the task can be basically completed through the efforts of the pilot, but the workload is heavy and more energy is required.

508 Chapter 6 Table 6.9: PIOR trend scale. Explanation There is no pilot-induced undesirable motion trend. Undesirable motions occur when the pilot begins to maneuver sharply or to attempt accurate control. These motions can be prevented or eliminated with the use of pilot techniques. It is easy to trigger undesirable motions when the pilot begins to maneuver sharply or to attempt accurate control. These motions can only be prevented and eliminated by sacrificing task performance or with considerable attention and effort of the pilot. If an oscillation trend occurs when the pilot begins to maneuver sharply or to attempt accurate control, the pilot must reduce the gain or abort the task to recover. If a divergent oscillation trend occurs when the pilot begins to maneuver sharply or to attempt accurate control, the pilot must operate the aircraft in an open-loop way such as bar loosening or bar freezing. Vibration or normal operation of pilot may lead to divergent oscillation and the pilot must operate the aircraft in an open-loop way such as bar loosening or bar freezing.

Grade 1 2 3 4 5 6

PIOR, Pilot induced oscillation rating.

4. PR 5 7
9, it means that the flying quality is quite poor and has reached the limit. The pilot has tried his best, but still cannot complete the task well. 5. PR 5 10, it means maneuverability failure. It is a very abnormal phenomenon and only occurs under a certain fault state or under certain conditions, and the result is that the working mode or flight state must be converted to avoid flight accidents. The PIO phenomenon trend is evaluated with the PIOR trend scale, and the scoring criteria and procedures are shown in Table 6.9, which are also described in Fig. 4.6 in Chapter 4 of this book. Among them, “undesired motion” refers to a single overshoot or a small, fast and damped periodic motion, and “oscillation” refers to a motion more than half a cycle or one overshoot.

6.12 Management of the flight control system “iron bird” integration test Scientific test management takes an important part in ensuring high efficiency and high quality of tests of large complex systems. For the “iron bird” integration test of the flight control system, it is also important to ensure the flying safety and reliability of aircraft. The management of “iron bird” integration test of the flight control system mainly covers the following content: 1. Drafting test plan The test plan (including test requirements and test procedures) shall be determined according to the flight control system design specifications, subsystem design requirements, control law design requirements, airborne equipment (hardware) design requirements and software design requirements and related design schemes. The test plan shall include test items, test content, test methods, test procedures, parameters to be tested

“Iron bird” integration test of the flight control system 509

2.

3.

4.

5.

and recorded and requirements and expected test results. Human resources, equipment resources and test environment shall be fully considered for the drafting of the test plan. Testing and recording Complete testing means and recording capability that meet the test and testing requirements are required, including the testing and recording of aircraft motion quantity, system state quantity, various control discrete quantity and internal parameters of flight control computer. In addition to a conventional high-bandwidth pen recorder, recording equipment such as tape and disk shall be used as far as possible to facilitate the reprocessing and long-term storage of test data. Development of test analysis software Test data analysis software that meets the test data analysis requirements should be developed, such as equivalent matching software, test data chart making and display software. Control of test technical state and configuration management The technical state shall be strictly controlled, including the state of the test equipment, software version and interface state. All kinds of faults in the test shall be recorded in detail, including fault nature, hazard degree, occurrence time, corrective measures and effects. When it is necessary to change the technical state of airborne systems or equipment, strict approval procedures shall be followed and the test items related to the change shall be retested. Preparation of test report According to the requirements of the test specification, the main idea of the test report is to truthfully reflect the test results and the key point is to analyze the test deviations so as to provide the basis for the judgment of the test task and results.

6.12.1 Test management requirements According to relevant provisions of Quality Management Requirements of Large Scale Test (GJB1452—92), the test undertaking unit shall formulate and implement strict test management regulations. As the “iron bird” integration test of flight control system of large transport aircraft is one of the most important large-scale tests for aircraft model development, the test quality management shall be planned and implemented according to relevant requirements of GJB1452—92. The quality management of the “iron bird” integration test of flight control system of civil aircraft can be taken for reference. 1. General requirements a. The quality management of large-scale tests shall be serious, meticulous and reliable without omissions. b. The units participating in the test shall jointly establish an effective test quality assurance system and implement whole-process quality management from test preparation, test implementation to test summary.

510 Chapter 6 c. Test activities shall be organized in strict accordance with test outline, quality assurance program, test procedures, operation specification and quality control procedures. d. A quality information network shall be established. e. All operators shall pass on-the-job training and go to the post only after getting the qualification certificate. f. All measuring instruments and apparatus used in the test shall meet the requirements. g. According to the test program, organize test quality inspection and review at different levels and in different stages. 2. Specific requirements a. Quality management in test preparation stage i. Prepare test task list The test task list shall be prepared by the test client and countersigned by test implementation units and relevant units, and shall go through prescribed examination and approval procedures. The content of the test task list generally include: I. name, code, nature and purpose of the test; II. test content and conditions; III. technical status of the test products, attached with the list of supporting products; IV. test items, parameters and accuracy requirements; V. quality assurance requirements; VI. technical requirements for providing test results; VII. judging criteria for test results; VIII. division of test tasks; IX. test schedule requirements and organizational measures. ii. Prepare test outline The test outline shall be jointly prepared by the test implementation units and test client and countersigned by units participating in the test, and shall go through prescribed examination and approval procedures. The content of the test outline generally include: I. source of task, time and place of test; II. name, code, nature and purpose of the test; III. test content and conditions; IV. technical status of the test products; V. technical status of preparation for the test; VI. test items, test methods, test analysis system block diagram, test error analysis and data processing methods; VII. technical difficulties and measures;

“Iron bird” integration test of the flight control system 511 VIII. the quality control and division of labor for junction parts when two or more units participate in the test; IX. plans and disposal principles for major problems at test site; X. technical support conditions and technical support measures for key test items; XI. test success probability or risk prediction; XII. test command procedures; XIII. list of technical documents used at site and requirements. iii. Prepare test plan, test procedures and operation specification According to content of the test outline, the test implementation unit prepares test plan, test procedures and operation specification and has countersignature on quality. iv. Prepare test quality assurance program The test quality assurance program shall be prepared by test implementation unit and the test outline shall be included in test plan, test procedures and operation specification. Generally, the quality assurance program includes: I. requirements for test quality objectives; II. quality responsibilities of the departments and personnel and the specific work evaluation criteria for different posts; III. determination of the quality control items and stipulation of the quality control points and control measures for the staff at different levels and in different stages; IV. content and requirements of staged quality inspection and review; V. fault judgment, response plan and safety assurance measures; VI. fault report and analysis and corrective action requirements; VII. requirements for collection and processing of abnormal data; VIII. evaluation criteria for high-quality test; IX. quality control of status of test products; X. conditions that test products must meet; XI. test products and special tools and test equipment shall be ensured as complete according to test task list and test outline; XII. the state of the test products shall be consistent with the product drawings and technical documents; XIII. the product quality certification documents shall be complete and consistent with the real products; XIV. the products shall have passed product quality review; XV. safety and risk shall have been demonstrated. v. Test products Products shall be inspected, debugged, accepted and handed over in accordance with the list of supporting products, quality certificate documents

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vi.

vii.

viii. ix.

x.

and specified review items. Changes in the test products and test state must be fully demonstrated and strictly approved. Quality control of measuring instruments and apparatus used in the test I. The equipment participating in the test shall be calibrated as conforming. II. The measuring instruments and test apparatus shall be within the validity period of calibration. III. Instruments and equipment repaired can only be used after being calibrated again. IV. The accuracy of test instruments shall satisfy the requirements of test parameters. V. If over two units participate in the test, the measuring standards shall be unified and comparison of test instruments shall be organized. VI. Disposable test instruments and apparatus shall be forcibly calibrated before being put in the test. Prevention of articles irrelevant to the test I. The test implementation unit shall formulate effective measures to prevent articles irrelevant to the test. II. Articles irrelevant to the test shall not be placed at the test site. III. The working media used for the test shall have cleanliness inspection and be disposed. IV. Tools and spare parts used by operators shall be provided with list and operators can only leave after checking the list at site. Establish fault report and analysis and corrective action system Control of quality control points The quality control points and control measures shall be specified, the quality tracking card or dotted line inspection system shall be implemented and inspection at three levels, that is, self-inspection, mutual inspection and special inspection shall be implemented for the quality control items. Organize the inspection on preparation for the test After the preparation for the test is completed, the test implementation unit shall organize an overall quality inspection and an overall joint examination of the test system with the units participating in the test, and then freeze the technical state after it passes the examination. Only after being approved by the commander of the test can it be transferred to the test implementation stage. The preparation state inspection before the test shall mainly include: I. state of test products; II. state of test equipment; III. state of test analysis system; IV. state of operation and command system;

“Iron bird” integration test of the flight control system 513 V. configuration of test documents and original records in test preparation process; VI. technical guarantee and safety measures; VII. environmental conditions of the test; VIII. fault disposal and emergency response plan; IX. qualification review of personnel participating in the test; X. control state of quality control points; XI. logistical support facilities. b. Quality management in test implementation stage The test procedures and operation specification shall be strictly followed and the specific requirements are as follows. i. The commander of the test shall organize the units participating in the test to conduct overall quality review of personnel participating in the test, test products, test instruments and equipment, technical documents, key quality control points and environmental conditions of the test. The test can only be started after being approved by the commander of the test. ii. The personnel participating in the test shall determine the personnel, post, responsibilities, coordination relationship and instrument and equipment before the test. Key positions require two-personnel and three-personnel system that consists of operators and supervisors. iii. For repeatable tests and large-scale tests composed of series test, the results shall be organized in time after the completion of last test item. iv. Analyze, find out and deal with problems in a timely manner and do not leave the problems to the next test. v. Follow the test procedures to unify commands and ensure command accuracy. vi. During the test, the system, key links and test equipment of the tested products shall be equipped with a working state monitoring and fault alarm system. vii. Test data shall be collected accurately and original records shall be made. viii. The test shall be interrupted if the specified test objective cannot be achieved during the test. The test can be continued only after troubleshooting and the order to interrupt or continue the test shall be given by the commander of the test. c. Quality management in test summary stage i. Handle test results Before leaving the test site, collect and sort out the test records and all the original data to ensure the integrity and accuracy of the data, process test data, analyze test results and propose test results evaluation. The evaluation of test results includes the following content:

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ii.

iii.

iv. v.

I. Whether the requirements of the task list and test outline are met; II. Evaluation on test data collection quality, generally including signal acquisition rate, data acquisition rate, key data acquisition rate and instrument integrity rate; III. Handling and handling methods of test results; IV. Test results analysis and evaluation; V. Fault analysis and handling condition. Prepare test summary The test report shall be jointly prepared by the test implementation units and the test client and countersigned by units participating in the test, and shall go through prescribed examination and approval procedures. The content of the test report generally includes: I. test objectives and test conditions; II. brief description of test process; III. analysis of test results; IV. test conclusion; V. improvement opinions. The changes of product status and test status in the test shall go through the examination and approval procedures after review and be timely reflected in corresponding product design drawings and technical documents. Collect, sort and archive quality and reliability information. Organize the summary of the test work, put forward rectification measures, and standardize the successful test experience.

6.12.2 Test measurement requirements The Requirements of Metrological Guarantee and Supervision for Large Scale Tests of Military Products (GJB1309—91) shall be strictly followed and the unit in charge of the test shall formulate the measuring requirements for the test. As the “iron bird” integration test of flight control system of large transport aircraft is one of the most important largescale system tests for aircraft model development, the measuring work shall be planned and implemented in accordance with relevant requirements of GJB1309-91. The requirements for the metrological guarantee and supervision of the “iron bird” integration test of flight control system of civil aircraft can be taken as reference. 1. General requirements a. All measuring instruments, standard materials, special test equipment and related quantity transmission and unification work that participate in large-scale test shall be included into the metrological guarantee and supervision range. The requirements of this standard shall be observed and relevant metrological standards and provisions shall be implemented.

“Iron bird” integration test of the flight control system 515 b. The metrological technical institution that undertakes the task of metrological guarantee and metrological supervision for large-scale tests shall be examined and recognized by National Defense Metrology and obtain the accreditation certificate issued by the metrological management institution of the Commission of Science, Technology and Industry for National Defense or its authorized institution. c. Metrological guarantee and supervision work for a large-scale test shall be included in the test quality control procedure and test plan and the execution of test shall be supervised by the command organization (the unit in charge of the test) that presides over or organizes the large-scale test. It shall be implemented by corresponding national defense metrology institutions. d. The quantity transmission of metrological instruments and standard substances used in the whole process of large-scale tests shall be carried out according to international metrological quantity transmission system or by internationally recognized metrological technical institution. e. Before entering the test site (base), all units participating in large-scale test shall submit a list of measuring instruments, standard substances and special test equipment brought to the test site (base) and submit it to the command organization hosting or organizing large-scale test and corresponding international metrology management institute. They can only be brought to the test site (base) after being approved. 2. Specific requirements a. Organizational management i. The metrological guarantee and supervision work for large-scale test shall be conducted by a chief metrologist or a supervisor in charge of metrological guarantee and supervision. ii. The metrological guarantee and supervision work for large-scale tests shall be included in the test quality assurance system and included in the test outline or test plan as an integral part of the system. iii. The chief metrologist or the supervisor in charge of metrological guarantee and metrological supervision shall, in accordance with the overall and systematic tactical and technical indicators, put forward the specific requirements for metrological guarantee and metrological supervision of largescale test and the configuration of metrological instruments as an integral part of the technical documents for model selection. iv. The general assembly department, the units participating in large-scale test and the international metrological institute at the test site (base) shall formulate the implementation scheme of metrological guarantee and metrological supervision work or working rules respectively according to the specific requirements for metrological guarantee and metrological supervision and organize to implement the metrological guarantee and metrological

516 Chapter 6 supervision. If necessary, the chief metrologist or leader in charge of metrological guarantee and metrological supervision shall coordinate with the superior international metrology management institute. b. Measuring instruments i. The strictest measuring standards of the units participating in the test shall be submitted to superior international metrology technical institute or submitted through internationally recognized metrological technical institution for calibration. ii. The measuring instruments used in the test shall be calibrated by international metrology technical institute of the unit or an internationally recognized metrological technical institution. The conformity certificate or work permit shall be provided and the instruments shall be used within the calibration period or validity period. Standard substances shall be stored according to specified conditions and used in validity period. iii. The calibration period or validity period of the measuring instruments used in the test shall be determined according to relevant military standards or national standards. iv. The measuring instruments that cannot be calibrated by the metrological technical institute of the unit participating in the test or the standard substances that cannot be evaluated shall be submitted to superior international metrology technical institute or through an internationally recognized metrological technical institution for calibration or comparison. v. Measuring instruments that cannot be calibrated or compared at present shall be self-checked by the metrological technical institution of the user according to technical conditions. After they pass the self-check, they can only be used with a use permit and reported to superior international metrological institute for filing. vi. For the test instruments managed by categories which are only used for qualitative indication and have no accuracy requirement, functional inspection shall be carried out by the user and a use permit shall be provided, and they shall be used within the specified period. vii. The personnel participating in metrological calibration and the environmental conditions for metrological calibration at the test site shall comply with the provisions of the Regulations on the Supervision and Administration of National Defense Metrology. The metrological calibration method or calibration method shall comply with relevant calibration regulations or operation specification. 3. Special test equipment a. The special test equipment used in large-scale test shall pass the appraisal or technical review and be provided with conformity certificate or user permit and the period of validity shall be determined.

“Iron bird” integration test of the flight control system 517 b. Technical documents required for the appraisal or technical review of special test equipment shall be provided by the design department and production department of special test equipment. c. It shall be ensured the special test equipment in the test are all within the validity period specified on their conformity certificate or use permit. d. The special test equipment used in the test shall be subject to metrological supervision and metrological technical arbitration carried out by the international metrological institute. e. The international metrological institute shall participate in the appraisal or technical review of the special test equipment used in the test according to test requirements, and undertake the necessary metrological guarantee and metrological supervision work. 4. Test site (base) a. The international metrology management institute and technical institute at the test site (base) shall be responsible for the metrological guarantee and supervision at the test site (base). b. The measuring instruments, standard substances and special test equipment of the units participating in the test used at the test site (base) shall be registered and filed, the list of measuring instruments, standard substances and special test equipment shall be filled in, and the verification certificate, conformity certificate (use permit) and annexes shall be complete. c. The command organization (the general responsible unit for the test) that presides over or organizes large-scale test shall be responsible for organizing measurement, quality, design, testing and other departments to conduct measurement review of all the measuring instruments, standard substances and special test equipment at the test site (base). The conforming ones shall be provided with a use permit label. d. In case of any inconsistency or dispute over the measured value at the test site (base), the metrological institution at the test site (base) shall conduct a metrological technical arbitral verification. e. According to the requirements of the large-scale test command organization (the general responsible unit of the test), the national defense metrological testing and research center, first-level national defense metrological station or other relevant international metrological technical institutes shall assist in the metrological guarantee and metrological supervision of the metrological institution at the test site (base). It shall undertake the measured value arbitration task if necessary.

6.12.3 Test process As the “iron bird” integration test of flight control system of large aircraft has a large amount of complex preparation work and heavy workload of data processing and analysis. To ensure

518 Chapter 6 the safe and efficient completion of system test and verification, the test process design and management shall be strengthened. Generally speaking, the process of implementing each test (item or content) is generally divided into three stages, that is, pre-test preparation, test implementation and post-test inspection. The test process is shown in Fig. 6.42. 1. Pre-test preparation The preparation for “iron bird” integration test of flight control system mainly includes the completion of following tasks: a. Check the ground test support facilities on the “iron bird” integrated test bed of the flight control system, including ground hydraulic source, ground power supply, flight control system test analysis system, test integrated management and control system and flight control system tester and confirm they work normally. b. Check the airborne systems of the flight control system such as PFC, ACE, FSECU, PDU and actuator, ensure that their technical status is consistent with the technical status of the aircraft for maiden flight and they operate normally. c. Check the analog airborne cables and test cables to confirm that they are correctly connected and the conduction and insulation meet requirements of relevant technical conditions for power-on inspection. d. Check the analog airborne hydraulic pipeline and test hydraulic management, confirm the correct connection and ensure the pipeline route, clamps and pipe type meet the requirements of relevant technical conditions for installation. e. Check the test system and test support system, and confirm that the test system status and test support system status are consistent with current test requirements. f. Check the safety assurance measures at the test site and confirm that the test site is safe and equipped with emergency measures and equipment (such as fire-fighting bottle, sandbox and shovel). g. Check the qualification and physical condition of the personnel participating in the test and determine the suitability between the person and post while confirming that they have relevant qualification. h. Check the test communication command system to ensure smooth communication and high communication quality. i. Check the test monitoring system to ensure that the monitoring coverage meets the test requirements, the equipment works normally and the image is clear. 2. Test implementation Following work is mainly completed in the test implementation process of the “iron bird” integration test of the flight control system: a. Prepare for the test, organize the pre-test meeting, introduce the test task, process, personnel responsibility arrangement and emergency response measures, etc. b. Check status of the test, test system, test support equipment, on-job status of test personnel, as well as status of site safety, hydraulic energy and ground power supply.

“Iron bird” integration test of the flight control system 519

Figure 6.42 “Iron bird” integration test process of flight control system.

520 Chapter 6 c. Turn on hydraulic energy (airborne and ground) to supply pressure to the tested system and equipment. d. Turn on the power (airborne and ground) to supply power to the tested system and equipment. e. Operate the test support equipment. f. Set and adjust the status of the tested system. g. Operate the flight control system test analysis system. h. Conduct and observe the test according to current test requirements and test procedures. i. Test and record relevant parameters in real time while executing the test. j. Observe and judge the short range of test in real time. k. Relieve the pressure of hydraulic oil source and disconnect the hydraulic oil source if the test is completed. l. Disconnect the power supply if the test is completed. m. Clear the test site. n. During the evaluation of the test, the make comments on the organization, operation and safety conditions exposed during the test and analyze and solve the problems found in the test. 3. Data processing and analysis Due to the large amount of data detected in the “iron bird” integration test of the flight control system, it is generally impossible to finish data processing and analysis immediately but requires a few days or dozens of days. To guarantee the test process and the validity of the test, only preliminary processing and analysis is conducted for the test data immediately and next test or test in next status can be started according to the test plan if no abnormalities are found.

CHAPTER 7

Onboard ground test of the flight control system 7.1 Overview The onboard ground test of the flight control system is a critical test required before the maiden flight of the aircraft. In a narrow sense, the onboard ground test of the flight control system only includes tests of the flight control system and its subsystems conducted in a more realistic aircraft environment and the cross-linking test with other airborne systems of the aircraft. In a broad sense, the onboard ground test of the flight control system shall also include the structural mode coupling test and electromagnetic compatibility test of the whole aircraft with the participation of the flight control system. The onboard ground test of the flight control system mainly includes the following content: 1. onboard ground test of the flight control system; 2. structural mode coupling test; and 3. electromagnetic compatibility test of the flight control system. The onboard ground test of the flight control system can be divided into the following five levels:

7.1.1 Installation and power-on inspection According to the technical conditions for installation and technical conditions for power-on of the flight control system and its subsystems, the flight control system and its subsystems are installed and debugged (this is the onboard ground inspection work that must be completed by the trial manufacturing unit) to ensure that the installation zero position and clearance of cockpit control unit, machinery control system, and control planes [including main control planes such as aileron, elevator, rudder and horizontal stabilizer, and auxiliary control planes such as ground spoiler and multifunctional spoiler, as well as lift augmentation devices of flaps (slats)] meet the technical conditions for installation inspection and that the cables, connectors, and separation surfaces of electronic airborne equipment of the fly-by-wire flight control system, high lift control system, automatic flight control system, and cross-linked airborne systems are connected correctly. Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00007-3 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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7.1.2 Functional and performance test The flight control system has functions including normal fly-by-wire control, degraded fly-by-wire control, simulated backup operation, mechanical backup operation, automatic control, and lift augmentation control. The first three functions depend on the fly-by-wire flight control system and the last three functions are relatively independent and are realized through independent subsystems. With the cable connector of airborne equipment of the flight control system and the aircraft cable separation surface connector, necessary signals are disconnected or introduced to inspect the static performances of the system, subsystems, and their airborne equipment, such as zero position, dead zone, polarity, transmission ratio, stroke, and speed as well as the cross-linking interface between subsystems. Meanwhile, the modal logic, display logic, protection logic, warning logic, and other functions of the fly-by-wire flight control system, high lift control system, and automatic flight control system are checked. For the key special state points of the aircraft, the dynamic response during typical inputs such as step and pulse as well as the stability margin of each control loop are tested. For the high lift control system, the flaps (slats) control time and control logic under different states shall be checked.

7.1.3 Cross-linking performance inspection between the flight control system and other airborne systems On the aircraft, the most real systems that can be cross-linked with the flight control systems are provided, mainly including the power supply system, hydraulic source system, landing gear system, antiicing and deicing system, atmospheric data system, inertial navigation system, avionics display and processing unit, central warning unit, central maintenance unit, autothrottle actuator, and flight management system. Their cross-linking interface, polarity, and logic, etc. are checked, together with the capability of the flight control system to realize its functions and performance under these environments.

7.1.4 Structural mode coupling test The structural mode coupling test checks the coupling between the flight control system and the aircraft structure. In other words, the control modes of the flight control system shall not cause unstable vibration of the structure and the vibration of the structure shall not cause divergence of the flight control system.

7.1.5 Electromagnetic compatibility test of the flight control system The electromagnetic compatibility test of the flight control system aims to check the mutual electromagnetic interference between the flight control system and other airborne systems.

Onboard ground test of the flight control system 523 In other words, the flight control system shall not affect other airborne systems due to electromagnetic interference, and also other airborne systems shall not affect the flight control system due to electromagnetic interference.

7.2 Onboard ground test of the flight control system 7.2.1 Test principle Compared with the “iron bird” integration test of the flight control system, the onboard ground test of the flight control system focuses more on the validation of the state of the flight control system, aiming to check whether the functions and performance of the onboard flight control system meet the design requirements. In principle, the test shall cover the inspection of all functions, the performance inspection focuses on static performance such as polarity and transmission ratio, and the dynamic performance focuses on inspection of critical design state points or boundary state points. The test results shall correspond to or be consistent with the “iron bird” integration test results and meet the requirements of the flight control system design specification. The onboard ground test of the flight control system shall follow the following six principles: 1. For the inspection on static performance such as polarity and transmission ratio, onboard airborne equipment and cables shall be fully used and only necessary test points or monitoring points are introduced, such as the polarity inspection of rate gyro assembly or overload sensor and inertial navigation device. 2. Simulation or excitation equipment shall be installed at the position of the original equipment of the aircraft to keep as close as possible to the real airborne equipment. 3. On the premise of ensuring the zero position, polarity, measuring range, and accuracy meet the test requirements, the control plane position sensor can connect the measured results to the test system, such as flight simulation system. 4. PEC, flaps (slats) controller, and automatic flight control computer shall all be set with testing or test interfaces to monitor or test the necessary variables of the system in real time and upgrade and maintain the system at the same time. 5. The testing and monitoring of ARINC429, MIL-STD-1553B, and AFDX buses can use a bus simulator or excitation equipment to excite real airborne equipment. For example, dynamic simulation can adopt a bus simulator to simulate an atmospheric data system and inertial navigation system, while the static simulation can use an atmospheric or inertial navigation exciter to set a certain state of the real atmospheric or inertial navigation device. 6. The control column (wheel) and various control handles can be driven by excitation equipment if the space or fixture allows that. For the case that the excitation equipment

524 Chapter 7 cannot be installed, the excitation can be completed through manual control, such as the test of the rod force
rod displacement curve. In the case that manual control is unavailable, the excitation can be realized in the next link and the premise is that the last link shall meet the design requirements. For example, after the neutral position, clearance, polarity, measuring range, and accuracy of control column are confirmed as correct; direct excitation from the sensor is allowed for the bandwidth testing of a single channel; and the excitation frequency sweeping from the control column is not required. However, during the processing of the results, the influence of this link on phase lag of the system shall be considered. The test principle of the onboard ground test of the flight control system is shown in Fig. 7.1. The energy required by the onboard ground test of the flight control system is ground energy, including power and hydraulic energy. The airborne systems/equipment used in the test include the tested system (flight control system) and its associated avionics system, electromechanical management system, antiicing and deicing controller, landing gear control system, hydraulic system, power supply system, and engine control system.

Figure 7.1 The test principle diagram of flight control system on-board test.

Onboard ground test of the flight control system 525 In the onboard ground test of the flight control system, the bus simulator or equipment exciter shall be configured for necessary sensors, such as atmospheric data and inertial navigation devices to achieve system state setting. PEC, flaps (slats) controller, and automatic flight control computer shall be configured with flight test interface equipment (FTI) to read system-related information in real time. To obtain the necessary static transmission ratio and system dynamic information, test equipment shall be installed on necessary interfaces, separation surfaces, or equipment. Similar to the engineering simulator test and “iron bird” integration test, the onboard ground test of the flight control system shall also prepare the test specification, test outline, and test task list before the test and sort out test records, prepare test reports, test summary, and analysis reports after the test completion. The main steps of the onboard ground test of the flight control system are as follows: 1. Complete the installation and debugging of the flight control system and subsystems according to the technical conditions for installation inspection and complete the power-on inspection on the aircraft according to the technical conditions for power-on inspection. 2. According to the principle of onboard ground test of the flight control system as shown in Fig. 7.1, connect the test cables and airborne cables from relevant airborne equipment and separation surfaces and connect to excitation, simulation, and test equipment. 3. According to the selected test state points, set the initial state of flight simulation system, excitation system, and necessary airborne systems. 4. Start the test according to the test content and steps in the test task list. 5. Analyze and deal with the test results. If there are problems, conduct the test again after troubleshooting or optimization design until the results meet the design requirements.

7.2.2 Debugging and preparation before test As the flight control system of large transport aircraft is quite complicated, to ensure the correct installation of the flight control system on the aircraft and the flight control system meets the system design requirements, a series of technical documents such as technical conditions for installation inspection and technical conditions for power-on inspection are specially prepared to assist the aircraft manufacturing department with system installation and debugging. 1. Technical conditions for cockpit control system manufacturing, installation, and debugging This technical document shall describe the general installation requirements on technical status, lubrication, and clearance, etc. related to the design drawings.

526 Chapter 7

2.

3.

4.

5.

Considering that pull rods are nonfinished parts, the manufacturing and assembly processes shall be inspected, including the main processes of steel tube surface, aluminum tube surface, pipe length tolerance, pipe closure and coaxiality, pipe inner diameter, pipe surface treatment, joint bearing pressing, riveting, marking, inspection and printing, and handover. For the installation on the aircraft, specifications shall be made regarding the installation of support, the combined installation of the rocker arm, base and pull rod, the installation and adjustment of the pull rod, and the installation of the neutral locating pin. Finally, an overall inspection shall be carried out for the neutral position of the lateral, heading, and longitudinal channels, zero position and maximum stroke of the command displacement sensor, and zero position and maximum stroke of the command force sensor. Technical conditions for machinery control system debugging This technical document shall describe the installation and debugging of the machinery control system combining cable gearing and the cockpit control system. An overall inspection is mainly conducted for the machinery control system of the aileron, rudder, elevator, and horizontal stabilizer control channels and the content includes neutral position, return performance, polarity, control limit, stroke, mechanical trim, transmission ratio adjustment, relationship between control force and displacement, and system clearance. Technical conditions for cable gearing installation This technical document shall describe the general installation requirements on technical state, appearance, lubrication, neutral locating pin, appearance mark, fastener, bonding jumper, and necessary file and repair related to the design drawings. Meanwhile, requirements are given for the connection between cables, the connection between cables and other airborne equipment and the integration of transmission components such as the base, rocker arm, pull rod, sector gear, pulley and guide part in cable transmission, and the performance indexes such as clearance, seal, tension, and lap performance of the system are specified. Technical conditions for installation and debugging of flight control system airborne sensors This technical document shall give technical requirements for the installation, electrical connection, and power-on inspection of the three-axis angular rate gyro assembly, three-axis linear acceleration sensor, wing tip overload sensor, vertical tail overload sensor, steering wheel displacement sensor, control column displacement sensor, pedal displacement sensor, control column (wheel) force sensor, pedal force sensor, aileron position sensor, elevator position sensor, rudder position sensor, slats position sensor, flaps position sensor, slats tilt sensor, and flaps tilt sensor used in the flight control system and make requirements on the interface, polarity, zero position, and stroke of relevant sensors. Technical conditions for actuator installation and debugging This technical document shall give specific technical requirements for the installation and debugging of actuators such as aileron, rudder, elevator, horizontal stabilizer,

Onboard ground test of the flight control system 527

6.

7.

8.

9.

multifunctional spoiler, and ground spoiler and give requirements for control mode, neutral position, and stroke of different actuators. Technical conditions for power-on inspection of fly-by-wire flight control system The technical document shall give requirements for the preparation before the poweron inspection, the inspection before power-on, the power-on inspection, the polarity inspection, and functions under different modes of the fly-by-wire flight control system. After the installation and debugging of the fly-by-wire flight control system, the power-on inspection of the fly-by-wire flight control system will be conducted according to this technical document. The technical conditions for power-on inspection of the fly-by-wire flight control system shall ensure the overall inspection on the basic functions of the flight control system, that is, power-on, control, display, logic and deviation correction, the overall inspection on interfaces of peripheral equipment, the overall inspection on typical faults including control column (wheel) stagnation, pedal stagnation, power supply fault, and hydraulic energy fault as well as overall inspection on static performances including zero position, polarity, and stroke when the airborne systems work normally. The airborne systems include power supply, liquid supply, atmospheric, intelligent probe, inertial navigation, radio altimeter, flight tube, display and control system, accident recording equipment, central warning, electromechanical integrated management, engine parameters, landing gear, and the antiicing and deicing systems. Technical conditions for installation and debugging of high-lift system actuator system This technical document shall specify technical requirements for the installation and debugging of slats power drive unit (PDU), variable-angle reducer assembly, guide support, torsion bar, slats antiwithdrawing brake device, rotary actuator, and rack and pinion mechanism, as well as the flaps PDU, main reducer, variable-angle reducer, flaps antiwithdrawing brake device, underslung type variable-angle reducer, torsion bar support components, and ball screw lead actuator mechanism, and give requirements for the flexibility of each actuating component and frictional force of the linear system, as well as the initial zero position (control power zero position), mechanical zero position (limit position), and maximum deflection angle of flaps (slats). Technical conditions for power-on inspection of high-lift system The technical document shall give requirements for the preparation before the poweron inspection, the inspection before power-on, the power-on inspection, the polarity inspection, and functions under different modes of the high lift control system. Technical conditions for power-on inspection of automatic flight control system This technical document shall specify the requirements for the inspection before power-on of the test equipment, ground energy, system and equipment state, impedance and voltage, the system power-on steps, and inspection after power-on, as well as the functions and performance covering interface, zero position, polarity, control logic and display, alarm, and BIT.

528 Chapter 7

7.2.3 Test items, test methods, and judging criteria The onboard ground test of the flight control system mainly includes the following aspects. 1. Interface inspection It checks whether the interface between subsystems of the flight control system and between the flight control system and other airborne systems is consistent with relevant definition in the interface control document. 2. Dynamic (static) performance test of fly-by-wire flight control system It mainly includes modal conversion function inspection, trim function inspection, safety protection function inspection, BIT function inspection, inspection on state and warning display and recording correctness, transmission ratio and polarity inspection, inspection on time-domain step performance, inspection on time-domain disturbance performance, openloop stability margin test, and closed-loop frequency response performance test. 3. Dynamic (static) performance test of machinery control system It mainly checks whether the rod force
rod displacement performance, damping performance, step performance, and frequency performance of the machinery control system and the open/closed-loop frequency performance of each channel meet the design requirements. 4. Dynamic (static) performance test of high-lift system It mainly includes flaps (slats) control function and performance test, modal conversion function inspection, safety protection function inspection, BIT function inspection, inspection on state and warning lamp display and recording correctness, and the inspection on the correctness of control logic, function, and time of flaps (slats) under normal, degraded, and backup working states. 5. Dynamic (static) performance test of automatic flight control system It includes modal conversion function inspection, safety protection function inspection, BIT function inspection, and inspection on state and warning lamp display and recording correctness. 6. Inspection on working condition of the flight control system when the engine is running It checks the working stability of the flight control system when the engine is running. 7.2.3.1 Interface inspection 7.2.3.1.1 Objectives and requirements

Before the onboard ground test of the flight control system, relevant mechanical and electrical interfaces shall be inspected according to technical conditions for installation, debugging, and power-on inspection, including neutral position, clearance, polarity, and measuring range. Besides, all bus communication signals shall be checked to see whether their dynamic transmission under different states is normal.

Onboard ground test of the flight control system 529 Requirements for the interface inspection are as follows: 1. Control device or exciter can be used at the interface between airborne equipment to make equipment send required data, but the bus simulator shall not be used to generate data. 2. The data used for communication shall cover the full range of the communication. 7.2.3.1.2 Content and methods

The interface inspection mainly includes: 1. the inspection on the communication interface between PFC, ACE, FSECU, and PDU; 2. the inspection on the communication interface between PFC and angle of attack sensor, atmospheric data equipment, total-pressure intelligent probe, inertial navigation device, central warning system (CWS), audio switching equipment and display processing unit (DPU), central maintenance system, accident recording equipment, radio altimeter, landing gear control system, and engine parameter recording system; and 3. the inspection on the communication interface between FSECU and CWS, central maintenance system, accident recording equipment, and ground proximity warning system. The basic method of the interface inspection is to set parameter values from the data sending device and check the received data information from the data receiving end. 7.2.3.1.3 Judging criteria

The data sent from the data sending end and the data received from the data receiving end shall be recorded respectively to judge whether they meet the design requirements. All test results shall be consistent with the definitions in the interface control document. Table 7.1 shows the record of interface inspection between PFC and FSECU.

Table 7.1: Record of interface inspection between PFC and FSECU. FSECU data sending Flaps (slats) control handle gear 0 1 15 25 30 40



PFC data receiving 

Flaps ( )

Slats ( )

Flaps ( )

Slats ( )

0 0 15.1 26.9 26.9 40.9

0 17.6 17.6 17.6 26.4 26.4

0 0 15.1 26.9 26.9 40.9

0 17.6 17.6 17.6 26.4 26.4

530 Chapter 7 7.2.3.2 Dynamic (static) performance test of fly-by-wire flight control system Main content of the dynamic (static) performance test of fly-by-wire flight control system includes: • • • • • • • • • • •

dynamic (static) performance test of actuator system; modal conversion function inspection; trim function inspection; safety protection function inspection; BIT function inspection; state and warning display and recording function inspection; transmission ratio and polarity inspection; time-domain step performance test; time-domain disturbance performance test; open-loop stability margin test; closed-loop frequency response performance test.

7.2.3.2.1 Dynamic (static) performance test of actuator system

1. Objectives and requirements The dynamic (static) performance test of actuator system aims to check whether the zero position, stroke, and polarity of longitudinal, lateral, and heading channels meet the design requirements and validate the dynamic (static) performance of actuators such as elevator, aileron, rudder, and spoiler. The horizontal stabilizer actuator is a special actuator whose trim function, polarity, and speed shall be checked. Requirements for the dynamic (static) performance test of an actuator system are as follows: 1. The amplitude of stroke inspection shall cover the full measuring range. 2. The frequency sweep range shall cover low-order natural frequency of relevant structures and the frequency sweep amplitude shall be able to stimulate structural vibration but not cause structural damage. In the case of vibration aggravation, there shall be an emergency power cutoff device to ensure the safety of the aircraft. 2. Content and methods The dynamic (static) performance test of the actuator system mainly includes the inspection on the static performance of longitudinal, lateral, and heading control channels and also the inspection on dynamic characteristics of control loops, specifically including: 1. the zero position, stroke, and polarity inspection of longitudinal, lateral, and heading control channels; 2. the maximum speed inspection of elevator, aileron, rudder, and spoiler actuators; 3. the loop step performance test of elevator, aileron, rudder, and spoiler actuators; 4. the loop frequency response test of elevator, aileron, rudder, and spoiler actuators; and 5. the trim function and performance inspection of horizontal stabilizer system.

Onboard ground test of the flight control system 531 All tests above are carried out under no load and the main test methods are as follows: 1. In the zero position, stroke and polarity inspection of longitudinal, lateral, and heading control channels, the left (right) control column, steering wheel, and pedal are operated, the signals of command displacement sensors such as control column, steering wheel, and pedal as well as the deflection angle signals of elevator, aileron, and rudder are recorded and data are processed to get the zero position and limit deviation value, so as to verify if the polarity meets requirements. 2. In the maximum speed inspection of the actuator system, forward and reverse excitation signals are applied to the test command input interface of ACE to test the output displacement of elevator, aileron, rudder, and spoiler and then the maximum speed is calculated. 3. In the step performance test of the actuator system, excitation signals are applied at the test command input interface of ACE to perform step performance test of elevator, aileron, rudder, and spoiler loops. The amplitude of excitation shall ensure that the servo valve of the actuator reaches the maximum opening. 4. In the frequency response performance test of the actuator system, excitation signals are applied at the test command input interface of ACE to test the frequency response performance of elevator, aileron, rudder, and spoilers loops. The amplitude of excitation signals shall ensure that the servo valve of the actuator reaches the maximum opening and the frequency range is 0.1
6 Hz. 5. In the performance test of the horizontal stabilizer, the trim switch is operated to check the angle of upward trim and the angle of downward trim and then the trim speed of the horizontal stabilizer is calculated. 3. Judging criteria According to the recorded data of the test analysis system, the zero position, stroke, and polarity of each control channel are analyzed and the maximum speed of each actuator is calculated. The step performance test data are processed into a time course step performance curve and the characteristic data such as overshoot, rise time, and adjustment time are calculated. The frequency response performance test data are processed into a frequency response performance curve and the bandwidth index is calculated. Table 7.2 shows the results of the step performance test of the multifunctional spoiler at the left side. Through the analysis of the amplitude
frequency characteristic curve and the phase
frequency characteristic curve of the frequency response performance test of the actuator system, the bandwidth of the actuator system can be obtained, namely the characteristic parameters (such as frequency) when the phase lag is 90 . Table 7.3 shows the frequency response performance data of the aileron actuator system.

532 Chapter 7 Table 7.2: Results of step performance test of multifunctional spoiler at left side. Positive excitation

Negative excitation

Excitation signal

σ%

N

tr/s

ts/s

σ%

n

tr/s

ts/s

1.0 V

0.04

0.00

0.10

0.11

0.03

0.00

0.08

0.10

Table 7.3: Frequency response performance data of aileron actuator system. Amplitude 0.5 V

Test parameters

Frequency (Hz)

Amplitude decrease 3 dB Phase lag 90

5.496 5.420

7.2.3.2.2 Modal conversion function inspection

1. Objectives and requirements The modal conversion function inspection mainly checks the modal logic, degraded logic, and the selection logic of the main signals and the content of inspection shall cover all modal conversion logic. 2. Content and methods The modal conversion function inspection mainly includes the following four parts: 1. Inspection on manual model conversion between normal (fly-by-wire) mode, simulated (fly-by-wire) mode, and mechanical mode. 2. Inspection on modal conversion from normal (fly-by-wire) mode to degraded mode under faults of some sensors. 3. Inspection on automatic conversion logic and selection logic under faults of power supply, hydraulic source, and system equipment, such as the inspection on automatic conversion from normal mode to simulated backup mode under power failure of PFC and the inspection on automatic conversion from normal mode to mechanical backup mode under power failure of ACE. 4. Inspection on selection logic of angle of attack, angular rate, and linear acceleration signals. The methods and process of the modal conversion function inspection are as follows: 1. Inspection on modal conversion between normal (fly-by-wire) mode, simulated (fly-by-wire) mode, and mechanical mode a. Set appropriate flight state, put the PFC “auto/off” switch on the fly-by-wire flight control panel at on and off positions, respectively; operate the control column, steering wheel, and pedal to have full-stroke motion; record the working state of the system, the command displacement of control column, steering wheel, and pedal and the deflection angle of elevator, aileron, and rudder; and analyze and record data and curve.

Onboard ground test of the flight control system 533 b. Use a flight control distribution box to simulate four sets of ACE to have power down all; operate the control column, steering wheel, and pedal to have full-stroke motion; record the working state of the system, the command displacement of control column, steering wheel, and pedal and the deflection angle of elevator, aileron, and rudder; and analyze and record data and curve. c. Put the elevator fly-by-wire/mechanical switch on the fly-by-wire flight control panel at “fly-by-wire,” “semimechanical,” and “mechanical” positions, respectively; operate the control column to have full-stroke motion; record the working state of the system, the displacement of control column, the deflection angle of elevator, and the displacement of the transmission ratio adjustment device; and analyze and record data and curve. d. Put the aileron fly-by-wire/mechanical switch on the fly-by-wire flight control panel at “fly-by-wire,” “semimechanical,” and “mechanical” positions, respectively; operate the steering wheel to have full-stroke motion; record the working state of the system, the displacement of steering wheel, the deflection angle of aileron, and the displacement of transmission ratio adjustment device; and analyze and record data and curve. e. Put the rudder fly-by-wire/mechanical switch on the fly-by-wire flight control panel at “fly-by-wire” and “mechanical” positions, respectively; operate the pedal to have full-stroke motion; record the working state of the system, the displacement of pedal, the deflection angle of upper and lower rudder; and analyze and record data and curve. 2. Inspection on modal conversion from normal (fly-by-wire) mode to degraded mode under faults of some sensors Set appropriate flight state, set faults of atmospheric data computer, inertial navigation device, and angular rate gyro assembly, respectively; operate the control column, steering wheel, and pedal; record the maximum deflection angle of elevator, aileron, and rudder; and analyze and record data and curve. 3. Inspection on automatic conversion logic and selection logic under faults of power supply, hydraulic source, and airborne equipment Under the normal working mode, disconnect two sets of PFC and ACE, respectively; record the working state of the system, the command displacement of control column, steering wheel, and pedal and the deflection angle of elevator, aileron, and rudder; and analyze and record data and curve. Under the normal working mode, disconnect No.1, No.2, and No.3 hydraulic systems, respectively; record the working state of the system, the command displacement of control column, steering wheel, and pedal and the deflection angle of elevator, aileron, and rudder; and analyze and record data and curve. 4. Inspection on selection logic of angle of attack, angular rate and linear acceleration signals

534 Chapter 7 Set appropriate flight state, set different values for the angle of attack directly connected with hard wire of the fly-by-wire flight control system, the angle of attack sent by atmospheric data computer, the angular rate of angular rate gyro assembly, and the angular rate of inertial navigation system; record the deflection angle of each control plane; and analyze and record data and curve. 3. Judging criteria The basic condition to judge the modal conversion function inspection results is to see whether the modal conversion meets the design requirements, whether the degraded logic meets the design requirements, and whether the power failure, undervoltage, and voltage cut logic meet the design requirements. Table 7.4 shows the record of inspection on manual conversion from normal mode to mechanical backup mode. 7.2.3.2.3 Trim function inspection

1. Objectives and requirements The trim function inspection aims to check the trim function and trim logic of horizontal stabilizer, aileron, and rudder, etc. and it shall also check all trim functions under different working modes of the fly-by-wire flight control system. 2. Content and methods In the trim function inspection, the trim function and trim logic of the horizontal stabilizer, aileron, and rudder, respectively, under the “normal,” “degraded,” “simulated,” and “mechanical” modes should be checked. The trim logic of the horizontal stabilizer is relatively complex, which mainly includes trim cutoff logic, electrical trim logic, mechanical trim logic, and trim function inspection. Table 7.4: Record of inspection on manual conversion from normal mode to mechanical backup mode. No. 1

2

3

Test conditions

Expected results

Put the elevator mechanical backup changeover switch on the fly-by-wire flight control panel at “semimechanical” position and operate the control column to have full-stroke motion Put the elevator mechanical backup changeover switch on the fly-by-wire flight control panel at “mechanical” position and operate the control column to have fullstroke motion Put the elevator mechanical backup changeover switch on the fly-by-wire flight control panel at “fly-by-wire” position and operate the control column to have fullstroke motion

The elevator is under mechanical backup mode, the transmission ratio adjustment device keeps unchanged, the control column has full deflection angle, the elevator has half deflection angle The elevator is under mechanical backup mode, the transmission ratio adjustment device is withdrawn, the control column has full deflection angle, the elevator has full deflection angle The elevator is under normal mode, the transmission ratio adjustment device is released, when the flaps (slats) are released, the control column has full deflection angle, the elevator has full deflection angle

Test conclusions

Onboard ground test of the flight control system 535 In the trim function inspection, an appropriate flight state is set and different working modes are switched through the fly-by-wire flight control panel to check the trim function and trim logic of the horizontal stabilizer, aileron, and rudder. 3. Judging criteria The results of trim function inspection shall meet the design requirements. Table 7.5 shows the record of trim cutoff logic inspection of the horizontal stabilizer. In the trim function inspection of aileron and rudder, set the “automatic/off” switch on the fly-by-wire flight control panel to “automatic” and “off,” respectively, and set the working mode conversion switch of aileron and rudder to “semimechanical” and “mechanical,” respectively. Conduct trim operation under the states above. Fig. 7.2 shows the curve of the recorded trim position and deflection angle of the aileron in the trim function inspection of the aileron under normal working mode. 7.2.3.2.4 Safety protection function inspection

1. Objectives and requirements The safety protection function is an important function to ensure the carefree handling of pilots and the flight safety of aircraft. It is conducted to ensure the normal condition of all protection functions. The main requirements for the safety protection function inspection are as follows: 1. All limit and protection function inspection shall cover all configurations, states, and modes on logic boundary. 2. The bar loosening attitude hold and spoiler lift destruction/drag increase functions shall be taken as special protection functions and verified. Table 7.5: Record of trim cutoff logic inspection of horizontal stabilizer.

Steps 1 2 3 4 5

Trim operation of fly-by-wire system Press to turn off the left trim switch of the horizontal stabilizer Turn down the left pilot trim switch for about 5 s Turn up the left pilot trim switch for about 5 s Turn down the right pilot trim switch for about 5 s Turn up the right pilot trim switch for about 5 s

Expected control plane response under normal mode

Control plane response under simulated backup mode

—

—

The control plane slowly The control plane slowly The control plane slowly The control plane slowly

moves moves moves moves

The control plane does not move The control plane does not move The control plane moves slowly The control plane moves slowly

Test results

536 Chapter 7

Figure 7.2 Curve of trim function inspection results under normal working mode.

2. Content and methods The functions covered by the safety protection function inspection mainly include: 1. Aileron-assisted lift augmentation function; 2. control plane deflection limit function under normal, degraded, simulated backup working mode and mechanical backup working mode; 3. bar loosening attitude hold and attitude protection function; 4. stall protection function; 5. overspeed protection function; and 6. spoiler ground lift destruction/drag increase function.

Onboard ground test of the flight control system 537 The methods of the safety protection function inspection are as follows. 1. Inspection of aileron-assisted lift augmentation function: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under the normal working mode of the fly-by-wire flight control system, set flaps (slats) to cruise, takeoff, and landing configurations, respectively, and record aileron deflection angle under different configurations. 2. Inspection of control plane deflection limit function: set the fly-by-wire flight control system under normal working mode and provide a cruise flight state point; respectively operate the control column, steering wheel, and pedal in full stroke and record the deflection angle of each control plane. Set failure of the atmospheric data and respectively operate the control column, steering wheel, and pedal in full stroke and record the deflection angle of each control plane. Switch the fly-by-wire flight control system to simulated backup working mode manually and operate the control column, steering wheel, and pedal in full stroke under the two states of flaps (withdrawn and released), and record the deflection angle of each control plane. 3. Inspection of pitch angle holding function: through the flight simulation system, set the normal overload as 1 g and other feedback signals as 0. Set the displacement of control column as 0, start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system; set the wheels to bear no load, set the pitch angle as 0 first and then set it as 15 1 s later; observe if the forward command of the control law longitudinal channel changes through FTI. Set the displacement of control column as 5 mm, set the pitch
15 1 s later, and observe if the forward command of control law longitudinal channel changes through FTI. 4. Pitch angle limit: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system, set the wheels to bear no load, set the pitch angle as 30 , and observe if the elevator moves. Set the pitch angle as 0 and observe if the elevator moves. 5. Slant angle hold: through the flight simulation system, set the normal overload as 1 g and other feedback signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system, set the wheels to bear no load, set the displacement of the pedal as 0, set the displacement of the steering wheel as 0, set the slant angle as 0 first and then set it as 15 1 s later, and observe if the aileron moves. Set displacement of pedal and 2 steering wheel as 2.5 mm and set the slant angle as
15 1 s later, and observe if the aileron moves.

538 Chapter 7 6. Slant angle limit: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system, set the wheels to bear no load, set the slant angle as 65 , and observe if the aileron moves. Set the slant angle as 0 and observe if the aileron moves. 7. Stall protection: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system; set the wheels to bear no load, set the angle of attack as 30 , and observe if the control column shakes and if the elevator moves. Set the angle of attack as 0 and observe if the control column shakes and if the elevator moves. 8. Overspeed protection: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system; set the wheels to bear no load, set the indicated airspeed as 650 km/h, and observe if the elevator and spoiler move. Set the indicated airspeed as 0 and Mach number as 0.8, and observe if the elevator and spoiler move. 9. Inspection of spoiler ground lift destruction/drag increase function: through the flight simulation system, set the normal overload as 1 g and other feedback signals and command input signals as 0. Start the system and make it work under normal working mode (open-loop) of the fly-by-wire flight control system; set the wheels to bear load and set the speed as 150 km/h, put the brake control handle at armed, set the throttle in thrust reverser state, and observe if the spoiler is opened. 3. Judging criteria The functions, including aileron-assisted lift augmentation, control plane deflection limit, bar loosening attitude hold, attitude protection, stall protection, overspeed protection, overload limit, and spoiler ground lift destruction/drag increase, shall meet the design requirements. 1. Aileron-assisted lift augmentation function: when flaps (slats) are under cruise configuration, the left (right) aileron is at a neutral position. When the flaps (slats) are under takeoff and landing configuration, the left (right) aileron deflects downwards simultaneously. 2. Control plane deflection limit function: under normal working mode, the maximum deflection angle of the control plane changes with the airspeed. The greater the airspeed is, the lower the maximum deflection angle of the control plane will be. When the system works under degraded working mode due to atmospheric failure, the airspeed signal is invalid and the control

Onboard ground test of the flight control system 539

3.

4.

5.

6.

7.

8.

9.

plane deflection is not limited. When the system works under simulated backup working mode, the control plane deflection will be limited according to the position of flaps. The control plane deflection is not limited when flaps are released and it is limited when the flaps are withdrawn. Pitch angle hold: when the control column is at the neutral position and the pitch angle is changed, commands are output. When the displacement of the control column is not at the neutral position and the pitch angle is changed, no commands are output. Pitch angle limit: if the pitch angle is set beyond the limit and there is command output, the elevator will move. If the pitch angle is set lower than the value enabling the pitch angle limit function and there is no command output, the elevator will not move. Slant angle hold: when the pedal and steering wheel are at the neutral position and the slant angle is changed, the aileron will move. When the displacement of the pedal and steering wheel is not at the neutral position and the slant angle is changed, the aileron will not move. Slant angle hold: if the slant angle is set beyond the limit and there is command output, the aileron will move. If the slant angle is set lower than the value enabling the slant angle limit function, the slant angle limit module is not started up and there is no command output, the aileron will not move. Stall protection: when the angle of attack exceeds the angle of attack for stall warning, the control column will shake and the elevator will deflect downwards. When the angle of attack is lower than the value enabling the stall warning function, the control column and elevator will have no response. Overspeed protection: when the indicated airspeed or Mach number exceeds the speed protection limit, the elevator will deflect upwards and the spoiler will be opened automatically. Inspection of spoiler ground lift destruction/drag increase function: if the conditions for automatic lift destruction/drag increase are met, the ground spoiler and multifunctional spoiler will be opened automatically.

7.2.3.2.5 BIT function inspection

1. Objectives and requirements The BIT function inspection aims to ensure the correct entry and exit interlock logic of BIT and the normal functions of PUBIT, PBIT, and MBIT. The test conditions shall cover all the judgment conditions for BIT logic. 2. Content and methods The main items and content of the BIT function inspection of the fly-by-wire flight control system are as follows.

540 Chapter 7 1. 2. 3. 4.

BIT interlock condition inspection. PUBIT function inspection. PBIT function inspection. MBIT function inspection.

The methods and process of the BIT function inspection of the fly-by-wire flight control system are as follows: 1. In BIT interlock condition inspection, start the fly-by-wire flight control system normally and set the interlock conditions through the flight simulation system according to Fig. 7.3. Set all wheels to bear load and the indicated airspeed shall be less than or equal to 70 km/h. Send a signal to start the ground test and observe if the BIT inspection is started. Then, change the logic judgment conditions in sequence and check if BIT is stopped. 2. Set the BIT interlock conditions in the ground test of CMS module of the fly-by-wire flight control system tester: software voting WOW_SW $ 3, and software voting speed of at least three channels # 70 km/h. 3. Power on the fly-by-wire flight control system from the power supply simulation equipment of the fly-by-wire flight control system, supply pressure to the fly-by-wire flight control system from ground hydraulic source, wait for 60 s, and observe if there are any PUBIT faults to be reported through the airborne data transmission and processing system and the ground CMS module of the fly-by-wire flight control system tester. 4. If no PUBIT faults are reported, monitor the airborne data transmission and processing system, wait for 10 s, and observe if there are IFBIT faults to be reported. 5. If no IFBIT faults are reported, send PBIT command from MFDS, wait for 150 s, and observe if there are PBIT faults to be reported through the airborne data transmission and processing system. 6. If no PBIT faults are reported, send MBIT command from the CMS module under the ground test and observe if there are MBIT faults to be reported through the airborne data transmission and processing system. 3. Judging criteria The BIT interlock condition inspection shall be consistent with the design requirements of the fly-by-wire flight control system. Table 7.6 shows the combinational logic for the setting of BIT interlock conditions and states of the fly-by-wire flight control system.

Figure 7.3 BIT entry/exit conditions.

Onboard ground test of the flight control system 541 Table 7.6: Combinational logic for setting of BIT interlock conditions and states. State 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16

Wheel load $ 3

Ground test allowed 5 2

Indicated airspeed (#70 km/h) $ 3

Current PFC wheel load signal means loaded

Expected result

True False True True True False True True False False True False True False False False

True True False True True False False True True True False False False True False False

True True True False True True False False True False True False False False True False

True True True True False True True False False True False True False False False False

Enter BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT Exit BIT

7.2.3.2.6 Inspection of state and warning display and recording correctness

1. Objectives and requirements The display of state and warning is an important part of the man
machine interface and it is the only way for pilots to understand the running state of the fly-by-wire flight control system. It aims to ensure that the display of the fly-by-wire flight control system under different configurations, states, and modes is normal. The inspection shall cover all normal states and the warning display. 2. Content and methods The state and warning display inspection mainly checks the reported warnings at warning level, attention level, and prompt level one by one to check whether the warning logic and warning display information and voice information are correct. The main check methods and processes are described as follows: 1. Power on the cross-linked airborne system and ensure it works normally (wheel load signal and wheel speed signal are valid), power on and supply pressure to the fly-by-wire flight control system and ensure it works normally. 2. Set the state of the fly-by-wire flight control system to make PFC send fault information such as “warning level fault,” “attention level fault,” and “prompt level fault” to the CWS. 3. In the case of a “warning level fault” of the fly-by-wire flight control system, set the “stall warning” state and other states and observe the corresponding warning display of different states of the fly-by-wire flight control system through the MFD and EICAS. 4. In the case of “attention level fault” of the fly-by-wire flight control system, set a “simulated backup” state and other states of the flight control system and observe the

542 Chapter 7 corresponding warning display of different states of the fly-by-wire flight control system through the MFD and EICAS. 5. In the case of a “prompt level fault” of the fly-by-wire flight control system, set the “degraded state” and other states of the flight control system and observe the corresponding warning display of different states of the fly-by-wire flight control system through the MFD and EICAS. 3. Judging criteria The state and warning display of the fly-by-wire flight control system shall be consistent with the design specification. Table 7.7 shows the state and warning display ways of the fly-by-wire flight control system. 7.2.3.2.7 Transmission ratio and polarity inspection

1. Objectives and requirements The transmission ratio and polarity inspection aims to check the transmission ratio and polarity of the fly-by-wire flight control system under different states and different feedbacks to ensure they meet design requirements. The inspection shall cover the control stroke of different channels. 2. Content and methods Through the simulation analysis of control law, the flight state points that can reflect the typical configurations of the aircraft are selected, as shown in Table 7.8. Check whether the size of the computer control command and deflection polarity of the control plane caused due to the changes of three-axis control command, three-axis Table 7.7: State and warning display ways of fly-by-wire flight control system. No. 1 2 3

Warning content

Lamplight

Main display area of warning information

Audio

Stall warning Flight control simulated backup Fly-by-wire degraded

Red Yellow

“Stall” “Simulated backup mode of flight control system” “Degraded mode of flight control system”

Stall None

Blue

None

Table 7.8: Selection of aircraft and flight state points for transmission ratio and polarity inspection. No.

Weight Center of Altitude Indicated Landing gear (t) gravity (MAC%) (m) airspeed (m/s) Configuration position

1 2 3 4

120 120 120 —

36 36 36 —

500 500 500 —

90.72 61.44 58.89 —

Cruise Takeoff Landing Takeoff

Withdrawn Released Released —

5

—

—

—

—

Cruise

—

System mode Normal mode Normal mode Normal mode Simulated backup mode Simulated backup mode

Onboard ground test of the flight control system 543 angular rate, normal and lateral overload, attitude angle, and angle of attack feedback signals under given state points are correct. 1. Longitudinal control: set different flight state points and operate the control column steadily and slowly, have reciprocating motion of a complete cycle in the sequence of “neutral and initial position - push to 20% position forwards - return to neutral position - pull back to limit position - return to neutral position,” keep the neutral position and front (rear) limit position stable, record the test curve, and compare with simulation results. 2. Lateral control: set different flight state points and operate the steering wheel steadily and slowly, have reciprocating motion of a complete cycle in the sequence of “neutral and initial position - deflect to 20% position to the left - return to neutral position - deflect to the limit position to the right - return to neutral position,” keep the neutral position and left (right) limit position stable, record the test curve, and compare with simulation results. 3. Heading control: set different flight state points and operate the pedal steadily and slowly, have reciprocating motion of a complete cycle in the sequence of “neutral and initial position - operate the left pedal to 20% position forwards - return to neutral position - operate the right pedal to the limit position forwards - return to neutral position,” keep the neutral position and left (right) limit position stable, record the test curve, and compare with simulation results. 4. Pitch rate input: set different flight state points, input pitch rate signals in the sequence of “0 /s-
10 /s -0 /s-10 /s-0 /s” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 5. Roll rate input: set different flight state points, input roll rate signals in the sequence of “0 /s-
10 /s -0 /s-10 /s-0 /s” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 6. Yaw rate input: set different flight state points, input yaw rate signals in the sequence of “0 /s-
10 /s -0 /s-10 /s-0 /s” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 7. Normal overload input: set different flight state points, input normal overload signals in the sequence of “1g-0-1g-2g-1g” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. Set different flight state points, input normal overload signals in the sequence of “0-0.2g-0-
0.2g-0” through the flight simulation system, change the state

544 Chapter 7 only after the system stabilizes, record the test curve, and compare with simulation results. 8. Angle of attack input: set different flight state points, input angle of attack signals in the sequence of  “0 -10 -20 -0 ” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 9. Pitch angle input: set different flight state points, input pitch angle signals in the sequence of  “0 -10 -0 -
10 -0 ” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 10. Roll angle input: set different flight state points, input roll angle signals in the sequence of  “0 -35 -0 -
35 -0 ” through the flight simulation system, change the state only after the system stabilizes, record the test curve, and compare with simulation results. 3. Judging criteria The transmission ratio and polarity of the fly-by-wire flight control system shall be consistent with the control law design requirements. Fig. 7.4 shows the control plane response curve under longitudinal input.

Figure 7.4 Control plane response curve under longitudinal input.

Onboard ground test of the flight control system 545 7.2.3.2.8 Time-domain step performance test

1. Objectives and requirements The step performance test aims to check whether the step response and simulation curve of the fly-by-wire flight control system are consistent and its core is to validate whether the characteristic parameters such as adjustment time and overshoot meet the design requirements. The polarity and transmission ratio of the fly-by-wire flight control system can also be judged according to the step performance test results. The test shall cover typical flight state points with changes in transmission ratio. 2. Content and methods Through the simulation analysis of the control law, flight state points that can reflect the typical aircraft configurations are selected to check the time-domain step performance of the longitudinal, lateral, and heading control under a given state. Set the flight state points, quickly push forward the control column slightly and keep it still for a moment, and then return it, and then quickly pull it back to full stroke and keep for a moment, and then return it, and finally compare whether the change curves of pitch rate and normal overload in the test are consistent with the simulation results. Set the flight state points, press the left steering wheel to the left to 20% of the full stroke and hold for 10 s, release the steering wheel and hold for 10 s, and then compare whether the change curves of pitch rate, yaw rate, lateral overload, and roll angle in the test are consistent with the simulation results. Set the flight state points, press the left steering wheel to the right to 20% of the full stroke and hold for 10 s, release the steering wheel and hold for 10 s, and then compare whether the change curves of roll rate, yaw rate, sideslip angle, lateral overload, and roll angle in the test are consistent with the simulation results. The methods and process of the time-domain step performance test for heading control are similar. 3. Judging criteria After the analysis and processing of the test data, the obtained step performance test curve and parameters shall be consistent with the simulation test curve and parameters. Fig. 7.5 shows the time-domain step response curve of pitch rate, normal overload, pitch angle, and angle of attack. Fig. 7.6 shows the time-domain step response curve of elevator deflection angle, horizontal stabilizer command, and deflection angle. 7.2.3.2.9 Time-domain disturbance performance test

1. Objectives and requirements The time-domain disturbance performance test aims to check the stability of the fly-by-wire flight control system under the disturbance of different sensors. Disturbance performance of typical sensors such as angle of attack, pitch angle, sideslip angle, and roll angle sensors shall be checked.

546 Chapter 7

Figure 7.5 Time-domain step response curve of pitch rate, normal overload, pitch angle, and angle of attack.

2. Content and methods Under given flight state points, the stability of the fly-by-wire flight control system under disturbances such as angle of attack, pitch angle, sideslip angle, and roll angle are checked. 1. Time-domain disturbance performance inspection under angle of attack: set the flight state points, set the angle of attack to generate 0.1 s 5 disturbance through the flight simulation system, record the change curves of pitch rate, normal overload, angle of attack, and pitch angle, and compare with the simulation results. 2. Time-domain disturbance performance inspection under pitch angle: set the flight state points, set the pitch angle to generate 0.1 s 5 disturbance through the flight simulation system, record the change curves of pitch rate, normal overload, angle of attack, and pitch angle, and compare with the simulation results. 3. Time-domain disturbance performance inspection under sideslip angle: set the flight state points, set the sideslip angle to generate 0.1 s 5 disturbance through the flight simulation system, record the change curves of pitch rate, yaw rate, lateral overload, roll angle, and sideslip angle, and compare with the simulation results. 4. Time-domain disturbance performance inspection under roll angle: set the flight state points, set the roll angle to generate 0.1 s 5 disturbance through the flight simulation

Onboard ground test of the flight control system 547

Figure 7.6 Time-domain step response curve of elevator deflection angle, horizontal stabilizer command, and deflection angle.

system, record the change curves of roll rate, yaw rate, lateral overload, roll angle, and sideslip angle, and compare with the simulation results. 3. Judging criteria After the analysis and processing of the test data, the obtained disturbance performance curve of typical sensors such as angle of attack, pitch angle, sideslip angle, and roll angle sensors shall be consistent with the disturbance performance curve obtained in the simulation. Fig. 7.7 shows the time-domain disturbance performance curve under the effects of the angle of attack. 7.2.3.2.10 Open-loop stability margin test

1. Objectives and requirements Stability margin is a key index to determine whether the fly-by-wire flight control system is stable. The stability margin test aims to confirm whether the stability margin of the fly-by-wire flight control system on a real aircraft meets the design requirements. The stability margin test is conducted under typical flight states. 2. Content and methods The open-loop stability margin test mainly includes the tests on stability margin of longitudinal, lateral, and heading control at different state points of the aircraft. The test principles are shown in Fig. 7.8.

548 Chapter 7

Figure 7.7 Disturbance response curve under effects of angle of attack.

Figure 7.8 Test principles of open-loop stability margin test.

As shown in Fig. 7.8, the flight state points are set to make the fly-by-wire flight control system under closed-loop normal (simulated backup) working mode. The aircraft is trimmed under set flight state and then the relevant parameters are set through a dynamic frequency response analyzer to generate sinusoidal frequency sweep signals, and then the signals are input to ACE through the test interface of ACE on the front panel of fly-by-wire flight control system tester. After they are added with the simulated command signals after D/A conversion in ACE, they are output to the actuator after servo amplification. The drive control plane of the actuator will move and generate a control plane deflection angle signal and input the signal to the flight simulation system, and then the flight simulation system

Onboard ground test of the flight control system 549 Table 7.9: Input signals for open-loop stability margin test. Channel Longitudinal Heading Lateral

Input value 0.1
3 Hz, 10%(1 V) 0.1
3 Hz, 10%(1 V) SP 5 A sinωt, A 5 10%(1 V), ω 5 2πf, f 5 0.1
3 Hz Lateral left (right) aileron frequency sweep input law is as follows:  SP SP $ 0 Left aileron frequency sweep input 5 SP1:5 SP , 0  SPð 21:5Þ SP $ 0 Right aileron frequency sweep input 5 SPð 21Þ SP , 0 G

will calculate the aircraft motion state and send it to ACE (input to PFC through bus after entering ACE) and PFC. The PFC sends control plane control command to ACE and perform D/A conversion. Then, the control plane control command (it can also be obtained through the test output interface on the front panel of the fly-by-wire flight control system tester) after D/A conversion and the frequency sweep signal are input to the open-loop frequency sweep inverse module, added together, and conducted with inversion, then they are input to the dynamic frequency response analyzer. Through the dynamic frequency response analyzer, a Bode diagram of the control plane control command after D/A conversion relative to the open-loop frequency sweep inverse module is drawn to analyze the stability margin (Table 7.9). 3. Judging criteria The results of the open-loop stability margin test shall meet the design requirements of fly-by-wire flight control system. Fig. 7.9 shows the curve of a longitudinal open-loop stability margin test. 7.2.3.2.11 Closed-loop frequency response performance test

1. Objectives and requirements The closed-loop frequency response test aims to check whether the bandwidth of the fly-by-wire flight control system meets the requirements. The sweep frequency range and amplitude shall not cause violent vibration so as not to damage the airborne equipment of the fly-by-wire flight control system and aircraft structures. 2. Content and methods The closed-loop frequency response test checks whether the closed-loop frequency response performance of the aircraft in longitudinal, lateral, and heading control under different flight states meets the design requirements and whether it is consistent with the simulation results. The test principle is shown in Fig. 7.10. As shown in Fig. 7.10, corresponding aircraft model is set in the flight simulation system to make the system work in a closed loop under the working mode of the fly-by-wire flight control system. The aircraft is trimmed under a set flight state and then the sinusoidal frequency sweep signal is sent through the dynamic frequency response analyzer to control

550 Chapter 7

Figure 7.9 Curve of longitudinal open-loop stability margin test.

Frequency sweep command Control displacement Control force Command displacement Command force

Mechanical displacement signal generator

Control column Steering wheel Pedal

PFC

Dynamic frequency response analyzer

ACE

Control surface displacement

Control surface

Actuator

Flight simulation system

Figure 7.10 Test principle of closed-loop frequency response performance test.

the mechanical displacement signal generator, and the control column, steering wheel, and pedal are excited respectively at a frequency of 0.1
3 Hz and amplitude of about 10% of full stroke. The control column, steering wheel, and pedal move to generate a pilot command to PFC and generate pilot command displacement to ACE. After demodulation through ACE, they are input to PFC and PFC outputs the control plane control command to ACE and ACE outputs the control plane control command to the actuator after D/A conversion and servo amplification. Then, the actuator will move to drive control plane and

Onboard ground test of the flight control system 551 generate control plane deflection angle and then the control plane deflection angle will be input to the flight simulation system. The flight simulation system will calculate the aircraft motion state and output the signal to ACE (input to PFC through bus after entering ACE) and PFC through the feedback sensor driven by the turntable. Through the dynamic frequency response analyzer, a Bode diagram of the normal overload, lateral overload, roll rate, yaw rate, and pitch rate relative to frequency sweep excitation is drawn and the test results are compared with the simulation results. 3. Judging criteria The results of the closed-loop frequency response test shall meet the design requirements of the fly-by-wire flight control system. Fig. 7.11 shows the curve of the closed-loop frequency response performance test of roll angle control channel. 7.2.3.3 Dynamic (static) performance test of the machinery control system The machinery control system is a backup control system that guarantees that the aircraft can still return to land when the fly-by-wire flight control system fails to work. The test of the machinery control system mainly includes the test of static performances such as rod force
rod displacement, clearance, friction, neutrality, stroke, return performance, as well as a test of dynamic performances such as step response and frequency response. The onboard ground test of the machinery control system aims to verify that the machinery control system meets the design requirements. The test principles of the machinery control system are shown in Fig. 7.12.

Figure 7.11 Curve of closed-loop frequency response performance test of roll angle control channel.

552 Chapter 7 Manual control Mechanical displacement signal generator

Horizontal stabilizer linear driving system

Cockpit control mechanism

Control surface

Mechanical drive device

Aileron linear driving system

Horizontal stabilizer actuator Aileron actuator

Horizontal stabilizer Left and right aileron

Measuring sensor Signal conditioner Measuring sensor Signal conditioner

Measuring sensor Signal conditioner

Measuring sensor Signal conditioner

Test recording system Frequency response analysis system

Test signal transmission line

Mechanical transmission route

Figure 7.12 The ground test principle for airborne s mechanical backup control system.

7.2.3.3.1 Static performance test of the machinery control system

1. Test objectives and requirements The static performance test of the machinery control system aims to verify that the rod force
rod displacement, starting force, frictional force, and return performance and transmission ratio of the machinery control system meet the design requirements. The control and excitation amplitude shall cover the full stroke of the machinery control system. 2. Test content and methods The static performance test of the machinery control system mainly tests the following items of the lateral, heading, and longitudinal machinery control system of aircraft. a. b. c. d. e.

Control force
control displacement relation. Starting force. Frictional force. Return performance. Transmission ratio.

Except for the return performance test, the testing of other static performances includes two control ways, that is, mechanical state and semimechanical state. All control includes left pilot and right pilot.

Onboard ground test of the flight control system 553 The test methods and process of performances such as control column force
column pilot displacement
control plane deflection angle, steering wheel force
steering wheel displacement
control plane deflection angle, and pedal force
pedal displacement
control plane deflection angle are basically the same. Set the machinery control system in a mechanical or semimechanical state, manually simulate the pilot’s operation in the cockpit at the left and right pilot hand and foot control mechanisms, respectively, or use a mechanical displacement signal generator to simulate the pilot’s operation. Apply a regulated amplitude and periodic excitation signal in the test and record the signals including control force, control displacement, actuator output displacement, and control plane deflection angle. After the test, analyze and process the test data to get the control column force
column displacement
control plane deflection angle, steering wheel force
wheel displacement
control plane deflection angle, and pedal force
pedal displacement
control plane deflection angle curves. The return performance test of the three channels is also basically the same. The methods are similar and the only difference is that an initial displacement meeting the conditions (such as exceeding the half stroke) is applied on the control column, steering wheel, and pedal and then the control column, steering wheel, and pedal are released to measure their return position. 2. Judging criteria The control force
control displacement, starting force, frictional force, return performance and transmission ratio shall meet the design requirements. Table 7.10 shows the results of wheel force
wheel displacement
control plane deflection angle performance test (left pilot position). Table 7.11 shows the results of lateral return performance test (left pilot position). During the analysis of the test results, record all the results in the table and compare them with the design values to finally determine whether the static performance meets the design requirements. 7.2.3.3.2 Dynamic performance test of the machinery control system

1. Test objectives and requirements The dynamic performance test of the machinery control system aims to check the step performance such as adjustment time and overshoot as well as characteristic parameters such as stability margin of the system under the mechanical and semimechanical working modes. In the dynamic performance test of the machinery control system, the excitation frequency range shall cover the low-order natural frequency of relevant structures and the amplitude shall not cause violent vibration that may easily cause structural damage. 2. Test content and methods The dynamic performance test of machinery control system has 12 combinations covering three control directions (longitudinal, lateral and heading), two control

554 Chapter 7 Table 7.10: Wheel force
wheel displacement
control plane displacement performance test (left pilot position). Anticlockwise rotation of steering wheel No.

Measured value

Measured value

Required value

1

Maximum stroke of left steering wheel ( )

58.79

60 6 6

2

Maximum stroke of right steering wheel ( )

56.77

60 6 6

3

Maximum deflection of left aileron ( )


27.89

23011 22

4

Maximum deflection of right aileron ( )

20.17

2011 22

5

Maximum control force (N)

200.23

180

6

Starting force of load mechanism (N)

34.77

# 40

7

Deflection force of left aileron (N)

70.10

# 50

8

Deflection force of right aileron (N)

69.96

# 50

9

Motion clearance of steering wheel relative to left aileron ( ) Motion clearance of steering wheel relative to right aileron ( ) Frictional force (N)

7.46

#3

8.13

#3

26.45

# 25

13.43

#6

13.76

#6

10

11

12

13

Total motion clearance of steering wheel relative to left aileron ( ) Total motion clearance of steering wheel relative to right aileron ( )

Test conclusion In conformity with requirements In conformity with requirements Basically in conformity with requirements In conformity with requirements Not in conformity with requirements In conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Basically in conformity with requirements Not in conformity with requirements Not in conformity with requirements

Clockwise rotation of steering wheel Measured value

Required value


59.22

60 6 6


57.40

60 6 6

20.18

2011 22


28.81

23011 22


221.30

180


51.59

# 40


69.39

# 50


69.68

# 50


5.98

#3


5.63

#3

26.74

# 25

Test conclusion In conformity with requirements In conformity with requirements Basically in conformity with requirements In conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Not in conformity with requirements Basically in conformity with requirements

modes (mechanical, semimechanical), and two kinds of test content (step performance and frequency response performance). 1. Step performance test: in the step performance test of the machinery control system, use a mechanical displacement signal generator to simulate the operation of the pilot and apply a step

Onboard ground test of the flight control system 555 Table 7.11: Lateral return performance test record (left pilot position). Anticlockwise rotation of steering wheel No. 1 2 3 4 5 6

Measured value

First time

Second time

Third time

Average value

Clockwise rotation of steering wheel First time

Second time

Third time

Average value

One-way return
0.222
0.222
0.222
0.222
0.354
0.336
0.336
0.342 distance of steering wheel ( ) Total return distance 0.132 0.114 0.114 0.120 of steering wheel ( )
0.255
0.265
0.272
0.264
0.255
0.239
0.249
0.248 One-way return distance of left aileron ( ) 0.000 0.026 0.023 0.016 Total return distance of left aileron ( )
0.183
0.177 0.177
0.061
0.065
0.095
0.079
0.080 One-way return distance of right aileron ( ) Total return distance 0.118 0.082 0.255 0.152 of right aileron ( )

signal of a certain amplitude (the amplitude shall not cause severe vibration which may easily cause structural damage and shall be greater than the motion clearance between the control mechanism and the control plane) and simultaneously record signals including control force, control displacement, output displacement of actuator, and control plane deflection angle. After the test, analyze and process the test data and draw the step response curve. 2. Frequency performance test: in the frequency performance test of the machinery control system, use a mechanical displacement signal generator to simulate the operation of the pilot and apply a sinusoidal signal of a certain amplitude and frequency (the frequency range shall cover the low-order natural frequency of relevant structures, and the amplitude shall not cause violent vibration that is likely to cause structural damage, and shall be greater than the motion clearance between the control mechanism and the control plane) and simultaneously record signals including control force, control displacement, output displacement of actuator, and control plane deflection angle. After the test is completed, analyze and process the test data and draw the frequency response curve. 3. Judging criteria According to the test data, draw time course curves of each input and output variable and calculate the step performance characteristic parameters according to the curve. The test results shall meet the design requirements. Table 7.12 shows the results of the lateral step performance test of the machinery control system (left pilot position).

556 Chapter 7 Table 7.12: Results of lateral step performance test (left pilot position, mechanical control).

State Allmechanical control Allmechanical control Allmechanical control

Direction

Measured value

Left steering wheel Left aileron anticlockwise control rotation plane Left steering wheel Right anticlockwise aileron rotation control plane Left steering wheel Left aileron control anticlockwise plane rotation

Steering wheel angle ( )

h(N) ( )

13.270


2.713

0

0 0.293 0.713 1.247 1.247

17.161

4.447

0

0 0.210 0.590 1.217 1.217

21.975


6.540

0

0 0.217 0.723 1.787 1.787

σ (%) n

td/s

tr/s

tp/s

ts/s

Table 7.13: Results of lateral frequency performance test (left pilot position, mechanical control). Displacement of left aileron relative to signal generator Amplitude gain (dB) State Mechanical, right wheel control Semimechanical, right wheel control

Amplitude 

10 10

Phase lag ( )

0.10 Hz

0.5 Hz

0.10 Hz

0.5 Hz

7.30
7.91

3.70
15.05


26.65
36.82


66.08
93.72

According to the test data, draw frequency response curves of each input and output variable and calculate the frequency performance characteristic parameters according to the curve. The test results shall meet the design requirements. Table 7.13 shows the results of the lateral frequency performance test of the machinery control system. 7.2.3.4 Onboard ground test of the high lift control system The onboard ground test of the high lift control system is an overall inspection carried out on the basis of the “iron bird” integration test of flight control system in a more realistic aircraft environment. It ensures that the difference between the onboard ground test and the “iron bird” integration test will not affect the normal functions and performance of the aircraft in the aspects of real sensor output, control plane installation and motion position. Its test principle is shown in Fig. 7.13. The onboard ground test of the high lift control system mainly includes the flaps (slats) normal control function and transmission ratio inspection, modal conversion function inspection, safety protection function inspection, BIT function inspection, and display warning function inspection under normal, degraded, and backup working modes.

Onboard ground test of the flight control system 557 Tested system

Flaps and slats control handle

Flaps and slats override control panel

Flaps position sensor

Slats position sensor

Flaps tilt sensor

Flaps and slats controller

Flaps and slats power drive device

Flaps and slats linear driving system

Flaps and slats

Slats tilt sensor

High lift subsystem tester

Photoelectric encoder

Tilt angle sensor

Counter

Signal conditioner

DIFF

Data acquisition system

Test system

Figure 7.13 Test principle of onboard ground test of the high lift control system.

7.2.3.4.1 Flaps (slats) normal control function and transmission ratio inspection

1. Objectives and requirements It checks the correctness of the flaps (slats) normal control function in a real aircraft environment. The high lift control system and relevant aircraft systems shall be under normal working state. 2. Content and methods The flaps (slats) normal control function and transmission ratio inspection includes the inspection of the control logic and function of the high lift control system under normal, degraded, and backup working modes, as well as the inspection of the transmission relationship under various control states. By installing a tilt angle sensor on the flaps (slats) control plane, the testing of the control plane deflection angle is completed. Through the test analysis system, the dynamic motion process of the flaps (slats) is recorded and saved. By installing a photoelectric encoder on PDU execution equipment, the testing of PDU speed and number of rings is completed. 1. Control logic and function inspection under normal working mode Set the flaps (slats) controller and PDU controller to work under normal mode, put the flaps (slats) control handle from gear “0” to gear “40” and record the continuous change curve of the flaps (slats) control plane test sensor signal through the test analysis system. Then, put the flaps (slats) control handle at gear “0” again and have operation step by step

558 Chapter 7 according to the full stroke of gears, namely “0”- “1” “1”- “15” “15”- “25” “25”“30” “30” - “40” and “40” - “30” “30” - “25” “25” - “15” “15” - “1” “1” - “0.” Record the speed of PDU, output shaft stroke (number of revolutions) of PDU execution equipment, and deflection angle of control plane with the counter. 2. Control logic and function inspection under degraded working mode Disconnect one circuit of power supply for flaps PDU and slats PDU to make the high lift control system work under the degraded mode. Operate the flaps (slats) control handle in full stroke step by step according to the operation method in (1) and record the speed of PDU, output shaft stroke (number of revolutions) of PDU execution equipment, and the control time in the whole process. 3. Control logic and function inspection under backup working mode Press the override button on the override control panel, control the flaps (slats) manually, make the high lift control system enter the backup working mode, control the flaps (slats) in full stroke, and record the speed of PDU, output shaft stroke (number of revolutions) of PDU execution equipment, deflection angle of the control plane, and the control time in the whole process. 3. Judging criteria for test 1. Judging criteria for control logic and function inspection under normal working mode Analyze and process the data recorded in the test, calculate the deflection angle of the corresponding flaps (slats) and time and confirm that they meet the design requirements. Compare the number of revolutions of PDU and the deflection angle of control plane and calculate the transmission ratio. The motion curve of normal release of the flaps (slats) is shown in Fig. 7.14, where the long dotted line is the motion change curve of the flaps and the solid line is the motion change curve of the slats. The test record forms and results of the flaps (slats) controlled step by step are shown in Tables 7.14 and 7.15.

Figure 7.14 Time course of flaps (slats) control plane control under normal working mode.

Onboard ground test of the flight control system 559 Table 7.14: Control record of flaps (slats) step by step under normal working mode. Slats Flaps (slats) control handle 0-1 1-15 15-25 25-30 30-40 40-30 30-25 25-15 15-1 1-0

Flaps

Speed (rad/min)

Number of revolutions/r

Deflection angle ( )

Speed (rad/min)

Number of revolutions/r

Deflection angle ( )

604 — — 603 — — 595 — — 595

151.5 — — 68.25 — — 66.75 — — 152.50

17.88 — — 25.74 — — 18.02 — — 0

— 1200 1199 — 1200 1199 — 1198 1199 —

— 230.25 174.25 — 210.75 205.75 — 175.00 227.50 —

— 15.35 26.88 — 41.20 27.11 — 15.39 0 —

Table 7.15: Control record of flaps (slats) step by step under normal working mode.

Flaps (slats) control handle

Slats command control plane deflection angle

0-1 0-18 1-0 18-0 0-1-15 0-18 15-1-0 18-0 0-1-30-40 0-18-26 40-4-1-0 26-18-0 1-15 — 15-1 15-25 — 25-15 25-30 18-26 30-25 26-18 30-40 —

Slats output shaft stroke (number of revolutions) of PDU execution equipment/r

Slats deflection angle ( )

Slats control plane control time requirement (s)

150.4 6 6

18 6 0.5

1612 25

150.4 6 6

18 6 0.5

1612 25

217.3 6 6

26 6 0.5

23 6 2

—

—

—

—

—

—

66.9 6 6

8 6 0.5

712 25

— — 0-27 27-0 0-27-41 41-27-0 0-15 15-0 15-27 27-15 —

—

—

—

27-41

Flaps deflection angle change

Flaps output shaft stroke (number of revolutions) of PDU execution equipment/r

Flaps command control plane deflection angle ( )

Flaps control plane control time requirement (s)

— — 419.6 6 6

— — 27 6 0.5

— — 2513 25

619.7 6 6

41 6 0.5

38 6 3

236.3 6 6

15 6 0.5

1413 25

183.3 6 6

12 6 0.5

1113 25

—

—

—

204.1 6 6

41 6 0.5

1313 25

According to the data recorded in the test above, the transmission ratio can be calculated. Through the comparison with the design value, it can be determined whether the transmission ratio meets the design requirements. 2. Judging criteria for control logic and function inspection under degraded working mode The data processing and judgment for the control function and transmission ratio inspection of the flaps (slats) under degraded working mode are similar to that under normal working mode and the difference is that the control time at each gear doubles.

560 Chapter 7 3. Judging criteria for control logic and function inspection under backup working mode During the normal control process of flaps (slats), if the arm button on the flaps (slats) override control panel is pressed, the flaps (slats) shall stop moving. If the override control panel is under armed status and the “release” button is pressed, it shall be ensured that the flaps (slats) can be released simultaneously. If the override control panel is under armed status and the “withdraw” button is pressed, it shall be ensured that the flaps (slats) can be withdrawn simultaneously. 7.2.3.4.2 Modal conversion function inspection

1. Objectives and requirements Modal conversion function inspection mainly checks whether the working mode of the high lift control system can be converted correctly according to the design requirements. During the modal conversion function inspection, the flight control system and associated aircraft systems shall be under normal working mode. 2. Content and methods The main content of the modal conversion function inspection includes conversion from normal working mode to degraded working mode and override priority logic function inspection. By powering the PDU execution equipment down, set one motor of flaps (slats) PDU as failed and observe the changes of relevant information of flaps (slats) in ECAIS and flight control diagram. Press the button on the override control panel of flaps (slats), observe the changes of relevant information of flaps (slats) in ECAIS and flight control diagram, operate the flaps (slats) control handle, and observe whether the flaps (slats) control plane moves. 3. Judging criteria Confirm the control logic correctness through the aircraft EICAS interface. The logic relation for conversion of working mode of the high lift control system has been introduced in Fig. 6.37 of Chapter 6. On EICAS, the flaps (slats) pointer is green and the flaps (slats) in the flight control diagram are green. On EICAS, the flaps pointer is yellow and the flaps (slats) in the flight control diagram are yellow and the control plane does not deflect. 7.2.3.4.3 Safety protection function inspection

1. Objectives and requirements It checks whether the safety protection function can be enabled normally and whether the displayed warning information can be reported correctly. It shall be ensured that the flight control system and associated aircraft systems are under normal working mode during the safety protection function inspection. 2. Content and methods Main content of the safety protection function inspection includes the inspection of flaps (slats) tilt protection function and flaps (slats) asymmetry protection function.

Onboard ground test of the flight control system 561 Set the flight control system under normal working mode and control the deflection of flaps (slats) through the flaps (slats) control handle. During the flaps (slats) control process, use the high lift control system tester to set the output voltage bias of flaps (slats) tilt detection device to observe the flaps (slats) control plane deflection status and display warning information. Set the flight control system under normal working mode and control the deflection of flaps (slats) through the flaps (slats) control handle. During the flaps (slats) control process, use the high lift control system tester to set the output voltage tolerance of left (right) flaps (slats) position sensor to observe the flaps (slats) motion status and display warning information. 3. Judging criteria From the tilt of flaps to stop of flaps motion, the release angle of flaps shall be less than the flaps tilt protection threshold (such as 4.5 ), PFD displays amber character “flaps drive” and has an alarm light; the flaps in the flight control diagram are shown in red. From the tilt of slats to stop of slats motion, the release angle of slats shall be less than the slats tilt protection threshold (such as 4.5 ), PFD displays amber character “slats drive” and has an alarm light; the slats in flight control diagram are shown in red. From the asymmetry of flaps to stop of flaps motion, the release angle of flaps shall be less than the flaps asymmetry protection threshold (such as 3 ), PFD displays amber character “flaps drive” and has an alarm light; the flaps on EICAS are shown in green and the flaps in flight control diagram are shown in red. From the asymmetry of slats to stop of slats motion, the release angle of slats shall be less than the slats asymmetry protection threshold (such as 3 ), PFD displays amber character “slats drive” and has an alarm light; the slats on EICAS are shown in green and the slats in flight control diagram are shown in red. 7.2.3.4.4 BIT function inspection

1. Objectives and requirements It tests and analyzes whether the BIT function of the high lift control system can be enabled normally and whether current fault information can be reported correctly. Set the flight control system under normal working state and set some faults that can be easily set through the high lift control system tester to check whether the faults can be reported correctly. 2. Content and methods BIT function inspection mainly includes PUBIT, IFBIT, PBIT, and MBIT function inspection.

562 Chapter 7 Power on the high lift control system normally and the PUBIT will be started automatically. After the power on, read the fault code through the high lift control system tester. After the system works normally, IFBIT will be started automatically and observe whether the fault code is reported through the high lift control system tester. Issue PBIT and MBIT inspection command through the ground maintenance equipment of the high lift control system and the high lift control system will be tested. After the test is completed, read the fault code and check whether the fault is fully reported. 3. Judging criteria According to the fault code recorded by the high lift control system tester, judge and confirm whether there is a BIT problem and confirm whether the state information containing the test LRU is reported in each stage of BIT. Due to the restriction of the onboard test environment, the fault is not manually set generally and only whether the BIT can be started normally and whether the system can pass the BIT without faults is checked. Table 7.16 shows the airborne equipment that can be tested in the BIT of the high lift control system. 7.2.3.4.5 State display and warning function inspection

1. Objectives and requirements It checks the cross-linking relationship between the high lift control system and airborne systems such as the CWS and DPU, as well as the correctness of the reported working state information and warning information. Table 7.16: Airborne equipment that can be tested in BIT of high lift control system. No.

LRU Name

PUBIT

PBIT

IFBIT

MBIT

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16

Handle position sensor Flaps (slats) controller Left flaps position sensor Right flaps position sensor Left slats position sensor Right slats position sensor Flaps PDU controller Flaps PDU execution equipment Flaps PDU rectifier Slats PDU controller Slats PDU execution equipment Slats PDU rectifier Left flaps antiwithdrawing brake device Right flaps antiwithdrawing brake device Left slats antiwithdrawing brake device Right slats antiwithdrawing brake device

— O — — — — O — — O — — — — — —

O O O O O O O O O O O O O O O O

O O O O O O O — — O — — — — — —

O O O O O O O O O O O O O O O O

Onboard ground test of the flight control system 563 To ensure the correct display of fault and warning information, it shall be ensured that the high lift control system can be started normally and relevant avionics equipment work normally. 2. Content and methods State display and warning function inspection mainly covers the warning and display of states including normal working, flaps (slats) half speed, flaps (slats) control failure, and flaps (slats) drive fault. The specific inspection methods and process are described below. 1. Check the cockpit display under the normal working mode. 2. Set any three circuits of signals of the four circuits of RVDT signal of flaps (slats) control handle to have failure and observe the fault phenomena and cockpit display. 3. During the motion process of flaps (slats) under normal working mode, set faults of FSECU1 and FSECU2 and observe the fault phenomena and cockpit display. 4. Set the fault of flaps (slats) PDU rectifier and observe the fault phenomena and cockpit display. 5. Set the fault of flaps and slats tilt sensors, respectively, and observe the fault phenomena and cockpit display. 6. Set the state information of flaps control channel (including normal working, flaps half speed, flaps control failure and flaps drive fault), check the state in the flight control system diagram and flaps control plane position display information. 7. Set the state information of slats control channel (including normal working, slats half speed, slats control failure and slats drive fault), check the state in flight control system diagram and slats control plane position display information. 3. Judging criteria Judge whether the state display and warning function are correct by checking whether the aircraft cockpit display screen is consistent with the control state. Table 7.17 shows the flaps (slats) PDU fault state display and warning results. As shown in Table 7.17, all the MFD displays are green, with no fault warning. When any three circuits of signal of the four circuits of RVDT signal of flaps (slats) control handle Table 7.17: Flaps (Slats) PDU fault state form. Test content Flaps PDU rectifier Slats PDU rectifier

Single channel fault Double channel fault Single channel fault Double channel fault

System response

PFD fault information

Flaps half speed Flaps motion stop Slats half speed Slats motion stop

White character “flaps half speed” Amber character “flaps drive” White character “slats half speed” Amber character “slats drive”

Flight control diagram display Flaps green box Flaps red Slats green box Slats red

564 Chapter 7 have failure, two FSECU fail. PFD displays amber character “flaps (slats) control failure” and has a light warning. Normal failure and degraded working mode failure displays an amber “flaps (slats) control failure” light warning. PFD displays amber “flaps driver”/“slats driver” and MFD displays flaps red/slats red. 7.2.3.5 Onboard ground test of the automatic flight control system The onboard ground test of the automatic flight control system is an overall inspection carried out on the basis of the “iron bird” integration test of the flight control system in a more realistic aircraft environment. Its test principle is shown in Fig. 7.15. The onboard ground test of the automatic flight control system mainly includes BIT, basic modal function (performance) inspection, modal conversion logic check, redundancy management inspection, transmission ratio inspection, stability margin test, and state and warning display function inspection. 7.2.3.5.1 BIT

1. Objectives and requirements BIT function inspection of the automatic flight control system is conducted under real aircraft conditions to ensure that the BIT threshold value is set reasonably and the function is effective. Onboard BIT inspection shall be carried out under

Auto-throttle disconnector switch （AT-DISC）

AFCS （tested system）

Automatic flight control unit （AFCU）

Automatic flight control computer（AFCC） Synchronous switch （SYNC） Auto pilot disconnector switch （AP-DISC）

Postback actuator （BDA）

Auto-throttle actuator （ATMA）

Auto-throttle controller （ATMC）

Test system

Auto-throttle transmission ratio changer

Automatic test device （ATE）

Avionics device exciter

Airborne avionics system

Tester

Flight test interface （FTI）

Control device of pilot Steering wheel force sensor Control column force sensor Pedal force sensor Control column displacement sensor Steering wheel angle sensor Pedal displacement sensor Brake panel control sensor

Primary flight control system

Primary flight control computer （PFC）

Actuator controller （ACE）

Linear acceleration turntable

Three-axis acceleration Control surface position signal

Generator simulation model

Elevator Aileron Rudder Spoiler Horizontal stabilizer

Test system

Aircraft equation

Aircraft simulation system Three-axis angular rate gyroassembly RS422

Mechanical connection

Hard wire

Three-axis angular rate turntable 1553B

AFDX

Figure 7.15 The ground test principle for airborne s automatic flight control system.

HB6096

Onboard ground test of the flight control system 565 the normal operation of automatic flight control system, so the aircraft must have conditions necessary for normal operation, including power supply, hydraulic source, avionics system, and fly-by-wire flight control system. 2. Content and methods Main content of the BIT of automatic flight control system is described below. 1. Check the PUBIT function of automatic flight control system. 2. Through PUBIT function inspection, confirm if the automatic flight control system has faults. During BIT of automatic flight control system, confirm relevant aircraft systems have been under normal working state and then power on the flight control system through the power switch (for aircraft without this switch, temporary switch can be made to ensure the flight control system can be powered on after other aircraft systems work normally) of the flight control system, and observe whether PUBIT has faults under the normal state of AFCS through the automatic flight control system tester. 3. Judging criteria Observe the working state of AFCS through the automatic flight control system tester. The automatic flight control system shall be able to start normally, BIT threshold shall be set reasonably, and PUBIT results shall be consistent with the state. Table 7.18 shows the items and record forms of BIT function inspection of the automatic flight control system. 7.2.3.5.2 Basic modal function (performance) inspection

1. Objectives and requirements The automatic flight control system is set with many working modes. The basic modal function (performance) inspection is a key point of the onboard ground test of automatic flight control system, with a purpose to further check the function and Table 7.18: Items and record forms of PUBIT. No.

Items

1 2 3 4 5 6 7 8 9 10 11 12

CPU test RAM test Power supply test Timer test AFCCs synchronization test Double-branch cross transmission test Double-branch No. test Double-channel cross transmission test Double-channel No. test AFCCs software version test Automatic flight control unit Back drive actuator

AFCC1

AFCC1

566 Chapter 7 performance of relevant modes and the correctness of the modal conversion logic in a real aircraft environment on the basis of the “iron bird” integration test of the flight control system. The modal entry condition inspection checks the entry conditions for each working mode. For different entry conditions, the working parameters of the mode after entry may be different. Thus careful verification is required, which mainly covers the roll angle hold mode, heading hold mode, pitch angle hold mode, and altitude hold mode. Modal entry condition inspection mainly determines the mode to be verified first, and then adjusts the tested system to the corresponding state according to the modal entry conditions and then turns on the switch of the corresponding mode. Finally, it is analyzed whether the modal entry conforms to relevant entry conditions according to the test curve. Modal entry condition inspection checks whether the fault test of relevant mode can be connected when the conditions for modal entry are unavailable according to actual conditions on the aircraft. Basic modal function and performance inspection can be conducted with the alarm display test to confirm the display of the mode of automatic flight control system and warning content on the control display equipment of the onboard avionics system. 2. Content and methods The basic modal function (performance) inspection shall cover all basic modes of the automatic flight control system, mainly including longitudinal working mode and lateral working mode. Longitudinal working mode includes pitch angle hold mode, altitude hold mode, vertical speed mode, and track slant angle mode. Lateral working mode includes slant angle hold mode, heading hold mode, track selection mode, track keeping mode, and heading selection mode. The methods and process of the basic modal function (performance) inspection of the automatic flight control system are as follows. 1. Input initial mode from the flight simulation system and ensure the control column, steering wheel, and pedal have no control inputs. 2. Start the automatic flight control system and check whether the state is normal. Check whether the system has faults or warnings in ATE and confirm there is no fault and warning. 3. Start the fly-by-wire flight control system and check if it is under normal working mode. Check from FTI to see whether there is a fault, check the initial position of the control plane, and confirm the initial state is input state. 4. ATE and FTI start recording data. 5. Set to enter relevant working mode and set to input signal to make the test mode generate control command. 6. Exit the working mode, disconnect the automatic flight control system, and clear the warning level warning. 7. ATE and FTI stop recording data and save the data.

Onboard ground test of the flight control system 567 3. Judging criteria Compare the test results with the simulation results to determine whether the test results meet the design requirements, that is, draw two curves in a same coordinate system. At the same time, observe and compare the data to confirm the correctness of the test results. Fig. 7.16 shows the functional (performance) test curve under pitch angle hold mode.

Figure 7.16 Functional (performance) test curve under pitch angle hold mode.

568 Chapter 7 7.2.3.5.3 Modal conversion logic check

1. Objectives and requirements The modal conversion logic check is conducted to further check the functions and performance of relevant modes and confirm the correctness of modal conversion logic in a real aircraft environment on the basis of the “iron bird” integration test of flight control system. The modal conversion logic check verifies each condition of the modal conversion if allowed by the conditions on the aircraft and checks if the alarm display of modal conversion is correct based on the alarm display test. 2. Content and methods The modal conversion logic check of the automatic flight control system includes basic working mode conversion logic and longitudinal and lateral working mode conversion logic. The basic working mode includes director mode, autopilot mode, longitudinal disconnection mode, lateral disconnection mode, and synchronous control mode. The longitudinal/lateral working mode includes takeoff director mode, pitch angle/slope hold mode, altitude hold/heading hold mode, vertical speed/track keeping mode, track slant angle/track keeping mode, flight level change/heading selection mode, and pitch angle hold/track selection mode. The process and methods of the modal conversion logic check of the automatic flight control system are as follows: Set the automatic flight control system in initial state and start the automatic flight control system; set and operate according to the entry conditions of the working method and working mode and observe the response and warning and display content on the aircraft. 3. Judging criteria Compare the test results with the design logic and confirm their consistency. Table 7.19 shows the checklist of the director mode of the automatic flight control system. 7.2.3.5.4 Redundancy management inspection

1. Objectives and requirements The redundancy management inspection further checks the redundancy management logic and strategy of sensor signals of the automatic flight control system in a real aircraft environment on the basis of the “iron bird” integration test of the flight control system to confirm the correctness of functions. In the redundancy management inspection, various airborne sensors on the aircraft shall be excited to check the redundancy management strategy when real airborne sensor signals enter the automatic flight control system.

Onboard ground test of the flight control system 569 Table 7.19: Test steps and results of director mode. No.

Initial conditions

Operation request

1

Initial conditions

Connect left flight director switch

2

Initial conditions

3

Initial conditions, set relative pressure altitude 400 m

4

Initial conditions, set roll angle 10

5

Initial conditions, set roll angle 0 , takeoff stage mark 0

6

Initial conditions, set roll angle 0 , takeoff stage mark 0

7

Initial conditions, set roll angle 0 , takeoff stage mark 0

8

Initial conditions, set roll angle 0 , takeoff stage mark 0

9

Initial conditions

10

11

Initial conditions, under horizontal navigation mode, set straight line mode and ensure effective data, set jaw distance 40 km Initial conditions

12

Initial conditions

13

Initial conditions

Test results

Expected results

Flight director mode, pitch angle hold, heading hold Connect right flight director Flight director mode, switch pitch angle hold, heading hold Connect right flight director Flight director mode, switch pitch angle hold, altitude hold Connect right flight director Flight director mode, switch pitch angle hold, altitude hold “Heading/track” switch button Flight director mode, Heading selection request pitch angle hold, Connect left flight director switch heading hold “Heading/track” switch button Flight director mode, Heading selection request pitch angle hold, Connect left flight director switch heading hold Flight director mode, “Heading/track” switch button Track selection request, connect pitch angle hold, track keeping left flight director switch Flight director mode, “Heading/track” switch button Track selection request, connect pitch angle hold, track keeping left flight director switch flight director mode, Adjust heading selection knob, pitch angle hold, selection request “Heading/ heading selection track” switch button Heading selection request, connect left flight director switch; AFCU slope angle limit gear signal, selected as “automatic” Flight director mode, Connect left flight director pitch angle hold, switch, press “horizontal horizontal navigation navigation” button “Vertical speed/track dip angle selection” button Request to select track dip angle, connect left flight director switch, press vertical speed button Connect left flight director switch, press flight level change button Connect left flight director switch, press flight level change button

Flight director mode, track dip angle, heading hold Flight director mode, flight level change, heading hold Flight director mode, flight level change, heading hold (Continued)

570 Chapter 7 Table 7.19: (Continued) No. 14

15

16

Initial conditions

Operation request

Expected results

Test results

Flight director mode, Connect left flight director switch vertical navigation, Altitude set by AFCU 5 1000 m; heading hold target altitude of FMS 5 2000 m; current relative pressure altitude 5 500 m; press vertical navigation button, FMS indicated airspeed controlled as 1 flight director mode, Set landing heading deviation Relative pressure altitude 400 m, altitude hold, heading effective, landing sliding flight director mode, channel, lower deviation effective, connect longitudinal altitude hold, lateral slideway armed approach and landing armed heading hold Takeoff stage mark 1 Connect left flight director switch Flight director mode, takeoff director Initial conditions, set vertical navigation speed, climbing way or required arrival time or speed decrease way, and the required data shall be effective

2. Content and methods The redundancy management inspection mainly checks the redundancy management logic and strategy when the following sensor signals enter the automatic flight control computer. 1. 2. 3. 4.

Radio altitude signals. Aircraft pitch, roll and yaw rate signals in inertial navigation. Aircraft normal and lateral linear acceleration signals in inertial navigation. Airspeed signals.

The redundancy management inspection is carried out in the following method and process below. 1. Power on the automatic flight control system and make it work normally; and 2. Input various redundancy signals to the automatic flight control system from excited airborne sensor and check the voting results from ATE. 3. Judging criteria Compare the voting value data recorded by ATE with the design values and confirm they meet the design requirements. Table 7.20 shows the redundancy management test record sheet of radio altitude signals when LRA1 and LRA2 are both normal. 7.2.3.5.5 Transmission ratio inspection

1. Objectives and requirements The transmission ratio inspection is carried out in a real aircraft environment on the basis of the “iron bird” integration test of the flight control system to confirm whether the transmission ratio meets design requirements. Before the transmission ratio inspection, the control polarity on the aircraft is confirmed to meet design requirements through control and excitation. In the transmission ratio inspection, typical flight state (transmission ratio is adjusted according to these flight state

Onboard ground test of the flight control system 571 Table 7.20: Redundancy management test results of radio altitude signals when LRA1 and LRA2 are both normal. Expected results LRA1 input/m 500 500 500

Test results

LRA2 input/m

Voting results/m

Signal state

Voting results/m

500 513 516

500 506.5 506.5

Normal Normal Fault

500 506.5 506.5

Signal state

parameters) is selected to check the transmission ratio under various working modes such as longitudinal working mode and lateral working mode of the automatic flight control system. 2. Content and methods The transmission ratio inspection includes the transmission ratio inspection under longitudinal working mode and lateral working mode. The longitudinal working mode includes present pitch angle under pitch attitude hold mode, present altitude under altitude hold mode, target vertical speed under vertical speed mode, present vertical speed under vertical speed mode, target track dip angle under track dip angle mode, and present track dip angle under track dip angle mode. Transverse/lateral working mode includes present slant angle under slant angle hold mode, present heading angle under heading hold mode, present track angle under track keeping mode, target track angle under track selection mode, present track angle under track keeping mode, target track angle under heading selection mode, and present heading angle under heading selection mode. For the transmission ratio inspection, set the flight state required first and then follow the following method and process to conduct the inspection. 1. Input the initial state through flight simulation system and ensure the control column, steering wheel and pedal have no control inputs. 2. Start the automatic flight control system, check if the state is normal, check if there is fault and warning through ATE, and ensure that there will be no fault and warning. 3. Start the fly-by-wire flight control system and check if there is fault through the FTI under the normal working mode. Check the initial position of control plane and confirm if the initial state is input state. 4. Start recording data. 5. Set and enter relevant working mode and input simulation signal to test analysis system. 6. Stop recording data and save the data. 3. Judging criteria Analyze and process test results and compare the test results with the control law simulation results. Compare the input and output of the automatic flight control system with the input and output of the fly-by-wire flight control system and draw the test results and

572 Chapter 7 simulation results in the same coordinate system to judge the correctness of the test results. The input state of the simulation adopts the input state of test data recording. If the simulation results and test results are consistent, it means the test results are correct. Fig. 7.17 shows the present pitch angle transmission ratio test results under pitch attitude hold mode.

Figure 7.17 Present pitch angle transmission ratio test results under pitch attitude hold mode.

Onboard ground test of the flight control system 573 7.2.3.5.6 Stability margin test

1. Objectives and requirements The stability margin test of automatic flight control system further tests and analyzes the stability margin in a real aircraft environment on the basis of the “iron bird” integration test of flight control system to confirm it meets design requirements. In the stability margin test on the aircraft, typical state points shall be selected according to the envelope of the aircraft and the stability margin under all working modes longitudinally, transversely/laterally shall be checked. It should be noted that the signal excitation amplitude shall not be too large to prevent the aircraft shaking violently or even rollover. 2. Content and methods The stability margin test of automatic flight control system checks the stability margin under longitudinal working mode and transverse/lateral working mode. Longitudinal working mode includes pitch angle hold mode, altitude hold mode, vertical speed mode, and track dip angle mode. Transverse/lateral working mode includes aileron channel under slant angle hold mode, rudder channel under slant angle hold mode, aileron channel under heading hold mode, rudder channel under heading hold mode, aileron channel under track selection mode, rudder channel under track selection mode, aileron channel under track keeping mode, rudder channel under track keeping mode, aileron channel under heading selection mode, and rudder channel under heading selection mode. Under a closed-loop state of the flight control system, sinusoidal frequency sweep excitation is input to the actuator from the fly-by-wire flight control system to test and analyze the open-loop stability margin of the system. The test methods and process are shown below. 1. Input system state through the flight simulation system and ensure the control column, steering wheel, and pedal have no control inputs. 2. Start the automatic flight control system, check if the state is normal, check if there is fault and warning through ATE, and ensure that there will be no fault and warning. 3. Start the fly-by-wire flight control system and check if there is fault through the FTI under the normal working mode. Check the initial position of control plane and confirm if the initial state is input state. 4. ATE and FTI start recording data. 5. Set and enter corresponding working mode. 6. Use a dynamic frequency response analyzer to input sinusoidal frequency sweep signals with amplitude as 10% of maximum stroke of actuator and frequency as 0.1
3 Hz to the actuator through ACE test chamber. 7. After frequency sweep is finished, press “Disconnect autopilot” button for two times to disconnect autopilot and clear the warning level warning. 8. ATE and FTI stop recording data and save data.

574 Chapter 7 3. Judging criteria Analyze and process the data recorded by ATE and FTI. The Bode amplitude
frequency and phase
frequency performance curves drawn by the dynamic frequency response analyzer and the stability margin test results under various working modes shall meet the requirements of the amplitude margin greater than 6 dB and phase margin greater than 45 . Fig. 7.18 shows the stability margin test curve under pitch angle hold mode. 7.2.3.5.7 State and alarm display test

1. Objectives and requirements The automatic flight control system realizes the display of its working state and warning through the avionics control display system. The state and alarm display test aims to verify the control display system cross-linking logic design between the automatic flight control system and the avionics system. In the state and alarm display test, the automatic flight control system tester shall be connected to realize the setting and adjustment of its working mode. 2. Content and methods The state and alarm display test of automatic flight control system mainly includes: 1. autopilot disconnection alarm and display; 2. autopilot default mode alarm and display; 3. longitudinal autopilot disconnection alarm and display;

Figure 7.18 Stability margin test curve under pitch angle hold mode.

Onboard ground test of the flight control system 575 4. lateral autopilot disconnection alarm and display; and 5. autopilot synchronous operation alarm and display. The automatic flight control system and avionics system shall work normally. By setting the state changes of automatic flight control system, observe the state changes of AFCU, DPU, audio alarm system, and alarm indicator lamp and judge the system state and alarm display logic. The test methods and process are shown below. 1. Power on the avionics system and automatic flight control system and ensure they work under normal state. 2. Observe the state changes of PFD, AFCU, audio alarm, and alarm indicator lamp according to the logic entry and exit of corresponding states of the automatic flight control system. 3. Press the autopilot disconnection switch on the steering wheel and observe if the alarm indicator lamp and audio alarm disappear. 4. Set and connect autopilot mode on the automatic flight control system tester, push the flight director switch on AFCU to turn it on, repeat step b and c, and observe the changes of PFD display and AFCU signal lamp. 3. Judging criteria Analyze and process the data recorded in the test, compare the test results with the design requirements of the automatic flight control system, and ensure that they are consistent. Table 7.21 shows the test results of the autopilot disconnection alarm and display. 7.2.3.6 Flight control system operation inspection under operation of engine 7.2.3.6.1 Objectives and requirements

Before the inspection of the flight control system operation under the operation of the engine, all tests of the flight control system are carried out under ground power supply, Table 7.21: Test results of autopilot disconnection alarm and display. Alarm and display Alarm lamp PFD alarm display area displays PFD autopilot mode display (Flight director not connected) AFCU display (Flight director not connected) PFD autopilot mode display (Flight director not connected) AFCU display (flight director connected)

Expected results Main alarm lamp is on (red) “autopilot disconnected” (red) Working mode: none Longitudinal mode: none Lateral mode: none The lamp is turned off Working mode: “F/D” Longitudinal mode: unchanged Lateral mode: unchanged Two “autopilot” lamps are off and others are unchanged

Test results

576 Chapter 7 ground hydraulic source, and ground air conditioned vehicle environment and the flight control system of aircraft is confirmed to meet the technical conditions for power-on, and the technical conditions for installation and its functions and performance are confirmed to meet the requirements of the design specification and to work stably. The test aims to, under the condition that the engine is started and the aircraft provides airborne power, hydraulic source, and air conditioning system and a more realistic working environment, observe and inspect the matching degree between the flight control system and aircraft airborne energy systems and other aircraft environment. It is generally required that the test shall be conducted simultaneously with other system tests when the engine is started. The system operation shall cover all working states, including slow operation, takeoff thrust force, cruise thrust force, and thrust reverser. 7.2.3.6.2 Content and methods

When the engine is started, the flight control system operation inspection is a qualitative inspection of functions and no ground test and excitation equipment will be added. In the actual working environment of the flight control system, the correctness of the function execution is checked and confirmed through the function setting of the flight control system. The setting method is to set the state for operation first, and then the actuator is operated and aircraft control plane and display control are observed, and the execution process and result of each function are finally confirmed. The test items include all control functions of the flight control system but the inspection of the parameter adjustment functions related to aircraft configuration and flight state is not involved. 7.2.3.6.3 Judging criteria

When the engine is started, the inspection of the flight control system operation checks the content specified in the technical conditions for power-on inspection. The results of the inspection shall meet the relevant regulations and requirements of the relevant documents. In other words, the modal conversion logic, control logic, and display shall be correct and meet design requirements.

7.2.4 Organization and implementation of the onboard ground test of the flight control system The implementation of the onboard ground test of the flight control system is led by the discipline of the flight control system, which is responsible for confirming the state of the aircraft and the system, preparing documents necessary for the test such as the test task list and the test outline, making and debugging the tester, exciter, simulator, test equipment,

Onboard ground test of the flight control system 577 and cables, implementing the whole test, recording relevant test results, and also preparing the test summary report. As the first coordination unit, the aircraft general manufacturer provides equipment necessary for the test such as test site, tested aircraft, ground power unit, ground hydraulic source, air conditioning vehicle, ground power supply (220 V/AC), and ground highpressure air source, undertakes the operation of aircraft-related equipment, data loading, observation or monitoring of relevant equipment, assists to record the test results, and connects or removes the equipment cables according to the demands. The units participating in the test include the disciplines with regard to the aircraft body, structure, avionics, power, hydraulic source, power source, landing gear control, and environmental control. The technical safety, quality, and customers’ representatives shall attend the whole test and have full-process monitoring and supervision of the safety, process, and results at the test site. In the test process, a detailed plan shall be formulated and test items, requirements, and personnel for the test shall be specified.

7.3 Structural mode coupling test As large transport aircraft present the characteristics of low and dense frequency of low-order elastic mode, the flight control system and aircraft elastic mode may have poor coupling, which brings great difficulties to the design of the flight control system. The aircraft design should satisfy the requirements on both the stability margin of the flight control system and the flutter stability margin of structures, which requires the cooperation between the discipline of aircraft structure and discipline of the flight control system. As it is hard to get a more accurate elastic model of aircraft structure and aircraft design generally adopts rigid aircraft, it is particularly important to verify whether the flight control system and aircraft structure mode coupling performance meets the design requirements and give a conclusion in a real environment. This test is usually called the flight control system and structural mode coupling test (SMI). As SMI usually needs to excite some structural modes of the aircraft and the sweep frequency is high, the structural damage of the aircraft may be easily caused and it is dangerous to complete various tests of the aircraft under different refueling conditions. Based on the above reasons, this test is an important and difficult test item among the ground onboard tests of aircraft.

7.3.1 Test objectives Through the SMI, the dynamic performance of the aircraft flight control system concerning the aeroservoelasticity discipline can be obtained and the coupling degree between the

578 Chapter 7 flight control system and the structural elastic mode can be confirmed, so as to provide a basis for the stability analysis of aeroservoelasticity of large transport aircraft. From the control loop, the flight control system is divided into the control augmentation system with fly-by-wire flight control system as the inner loop and the automatic flight control system with autopilot and others as outer loops. They both need to verify the coupling with the structural mode.

7.3.2 Test principle and methods SMI mainly verifies the dynamic performance of the flight control system of large transport aircraft in the low-order elastic modal frequency domain of aircraft structure and the servoelastic stability margin of the flight control system. It mainly includes the following test items: 1. Longitudinal, lateral, and heading frequency response test. 2. Longitudinal, lateral, and heading closed-loop pulse margin (gain margin) test. 3. Stability test of flight control system under structural resonance mode. The SMI test principle is divided into the frequency response test, closed-loop pulse response test, and resonance test. 7.3.2.1 Frequency response test The frequency response test is generally divided into the open-loop frequency response test and closed-loop frequency response test. The closed-loop frequency response test generally shows the frequency response of output to input when the flight control system has closed-loop operation. The open-loop frequency response test generally disconnects the command input point of ACE and records the curve of the open-loop frequency response of the output of the flight control system at this point to the input. It is difficult to disconnect the system to carry out the open-loop frequency response test. Generally, the open-loop frequency response test is carried out when the flight control system works in the closed loop. As a result, the open-loop frequency performance test is dangerous and can easily lead to structural damage of the aircraft. Therefore the amplitude and frequency range of the excitation signal must be selected reasonably, so as to stimulate the structure mode without damaging the structure. The principle of the open-loop frequency response test is shown in Fig. 7.19. The open-loop frequency response test can realize the testing of the open-loop frequency performance when the system has closed-loop operation by offsetting the forward command of the servo loop. The test principle of the open-loop/closed-loop test of the fly-by-wire

Onboard ground test of the flight control system 579 Steering engine

Data acquistion Cable and analysis PC Interface

Flexible aircraft

+

A/D

Exciation signal generator Stability augmentaion controller Flight control test PC

Figure 7.19 Principle of open-loop frequency response test.

Figure 7.20 Principle of open-loop/closed-loop frequency response test of fly-by-wire flight control system.

flight control system and automatic flight control system is shown in Figs. 7.20 and 7.21. The open-loop transfer function is G(s) 5 VOUT/VSOURCE. Based on the test principles above, the test principles of the open-loop/closed-loop test of the fly-by-wire flight control system and automatic flight control system can be constructed as shown in Figs. 7.22 and 7.23.

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Figure 7.21 Principle of open-loop/closed-loop frequency response test of automatic flight control system.

Figure 7.22 Realization diagram of the open-loop/closed-loop frequency sweep test of fly-by-wire flight control system.

Onboard ground test of the flight control system 581

Figure 7.23 Realization diagram of the open-loop/closed-loop frequency sweep test of automatic flight control system.

7.3.2.2 Stability test of the flight control system under closed-loop pulse response and structural resonance The pulse response is to change the forward gain of the actuator system under the condition of the closed loop of the flight control system and it is usually increased to three times the forward gain. The principle of the stability test of flight control system under the closed-loop pulse response and structural resonance of the fly-by-wire flight control system and automatic flight control system are shown in Figs. 7.24 and 7.25. The gain is changed by changing the parameter K. 7.3.2.3 Resonance test The resonance test has similar test principle with the closed-loop pulse test, that is, excitation signals of certain amplitude and frequency are applied on the wing surface, fuselage, and other structures. The excitation test corresponds to the aircraft structural mode and checks whether the stability performance of the modal coupling between the flight control system and the aircraft structure meets the design requirements. Based on the above test principles, the realization diagrams of the stability test of flight control system under closed-loop pulse response and structural resonance of the fly-by-wire flight control system and automatic flight control system can be constructed as shown in Figs. 7.26 and 7.27.

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Figure 7.24 Test principle of pulse response test of fly-by-wire flight control system.

7.3.3 Implementation of the structural mode coupling test 7.3.3.1 Frequency response test To check whether the stability redundancy of the flight control system meets the design requirements under corresponding structural mode, longitudinal, lateral, and heading open-loop frequency response test is generally carried out under fly-by-wire control and automatic control modes. 7.3.3.1.1 Test items and methods

SMI includes the test under two modes: fly-by-wire control mode and automatic control mode. Fly-by-wire control mode Connect the fly-by-wire flight control system tester, incorporate it into the open-loop frequency sweep reverse module, set different weights and configurations of the aircraft, and connect as shown in Fig. 7.22 (realization diagram of open-loop/closed-loop frequency sweep test). Dynamic frequency response analyzer imposes sinusoidal frequency

Onboard ground test of the flight control system 583

Figure 7.25 Test principle of pulse response test of automatic flight control system.

sweep signals of certain amplitude (the “optimal” amplitude initially determined in the test) and certain frequency band (greater than the first-order bending, first-order torsion and control plane rotational frequency of main wing surface in principle) through the input port of the gain controller and then the dynamic frequency response analyzer records actuator output displacement, control plane deflection angle, angular rate, angular speed, and the frequency response curve of output to input (disconnect point). Automatic control mode Connect the automatic flight control system tester, incorporate it into the open-loop frequency sweep reverse module, set different weights and configurations of the aircraft, and connect as shown in Fig. 7.24 (realization diagram of open-loop/closed-loop frequency sweep test). Dynamic frequency response analyzer imposes sinusoidal frequency sweep signals of certain amplitude (the “optimal” amplitude initially determined in the test) and certain frequency band (greater than the first-order bending, first-order torsion, and control plane rotational frequency of main wing surface in principle) through the output port of the gain controller and then the dynamic frequency response analyzer records angular speed, angular rate, and the frequency response curve of inertial navigation output signal (relevant signals entering control law of the flight control system) to input point.

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Figure 7.26 Realization diagram of closed-loop pulse and resonance test of fly-by-wire flight control system.

7.3.3.1.2 Judging criteria

The open-loop stability margin of lateral, heading, and longitudinal channels shall satisfy requirements of GJB2191—1994, that is, the amplitude margin shall be greater than 6 dB and the phase margin shall be greater than 60 . Analyze and evaluate the test results and draw a frequency response curve (Bode diagram and Nyquist diagram). Fig. 7.28 shows the structural mode coupling test curve. 7.3.3.2 Closed-loop pulse margin test 7.3.3.2.1 Test items and methods

The closed-loop pulse margin test includes the longitudinal, lateral, and heading closed-loop pulse margin test under fly-by-wire control and automatic control modes. Connect the fly-by-wire flight control system tester, incorporate it into the gain adjustment module (it can be realized by using a simulation system construction software module meeting the real-time requirements, such as flight simulation system), set different weights and configurations of the aircraft, and connect as shown in Figs. 7.26

Onboard ground test of the flight control system 585

Figure 7.27 Realization diagram of closed-loop pulse and resonance test of automatic flight control system.

and 7.27 (realization diagram of closed-loop pulse margin test). Dynamic frequency response analyzer imposes pulse signals of certain amplitude (the set amplitude shall ensure the control plane output will not be saturated after it reaches three times the gains of the system) and pulse width not greater than 20 ms through the output port of the gain adjustment module of signal generator, and then the dynamic frequency response analyzer records the time-domain response curve of actuator output displacement, control plane deflection angle, linear acceleration and angular rate. If the coupling test of automatic flight control system is to be conducted, the time-domain response curve of inertial navigation output signals shall be recorded at the same time. 7.3.3.2.2 Judging criteria

When the forward gain is increased to three times the gain, the response of the flight control system shall converge stably. Analyze and evaluate the test data and time-domain response curve and give a conclusion meeting the design requirements. Fig. 7.29 shows the closed-loop stability margin test curve.

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Figure 7.28 Frequency response test curve.

7.3.3.3 Stability test of the flight control system under structural resonance Excite different resonance modes of the aircraft and check whether the stability margin of the flight control system meets the design requirements under corresponding structural modes. 7.3.3.3.1 Test items and methods

The stability margin test of the flight control system usually includes a longitudinal, lateral, and heading stability margin test of the flight control system under the structural excitation of different resonance modes. Its test method is similar to that of the closed loop pulse response test and there is no need to input the control plane pulse excitation. The exciter excites the control plane structural mode (such as the first curve of the vertical tail), increases (lateral/heading) forward gain to three

Onboard ground test of the flight control system 587

Figure 7.29 Closed-loop stability margin test curve.

times, and records the time-domain response curve of control plane, deflection angle, overload sensor, and angular rate gyro assembly under the corresponding state when they output signals. 7.3.3.3.2 Judging criteria

Under the structural excitation, the forward gain increases to three times and the response of the flight control system converges stably. Analyze the evaluate the test data such as control plane deflection angle, linear acceleration, and angular rate and time-domain response curve. Fig. 7.30 shows the stability test curve of the flight control system under structural resonance mode.

7.3.4 Organization and implementation of the structural mode coupling test The implementation of the structural mode coupling test is led by the discipline of flight control, which is responsible for confirming the state of the aircraft and the system, preparing documents necessary for the test, such as the test task list and the test outline, making and debugging the tester, exciter, simulator, test equipment, and cables, implementing the whole test, recording relevant test results, and also preparing the test summary report.

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Figure 7.30 Time-domain response test curve of the first curve of vertical tail.

As the first coordination unit, the aircraft general manufacturer provides equipment necessary for the test such as test site, tested aircraft, ground power unit, air conditioning vehicle, ground power supply (220 V/AC), and ground high-pressure air source, undertakes the operation of aircraft-related equipment, data loading, observation or monitoring of relevant equipment, assists the recording of the test results, and connects or removes the equipment cables according to the demands.

Onboard ground test of the flight control system 589 As the second coordination unit, the aircraft flutter discipline is responsible for pasting and arranging the overload sensor, jacking the aircraft structure excitation equipment and air spring, as well as resonance data processing system. The units participating in the test include aircraft body, structure, and strength disciplines. The confidentiality, technical safety, quality and customers’ representatives shall attend the whole test and have full-process monitoring and supervision of the safety, confidentiality, process, and results at the test site. In the test process, a detailed plan shall be formulated and test items, requirements, and personnel for the test shall be specified.

7.4 Electromagnetic compatibility test Modern large transport aircraft are equipped with a series of electromechanical, electronic, and electrical systems such as flight control system, avionics system, mission system, power supply system and environmental control system. At the same time, the use of a large number of computers and bus transmission technology means that the airborne systems and airborne equipment are more sensitive to electromagnetic interference. The electromagnetic compatibility test of flight control system is conducted to qualitatively and quantitatively test the electromagnetic compatibility between the flight control system and the airborne systems and to find out the sources of interference, interference routes, and sensitive equipment, so as to eliminate the incompatibility and ensure the safety and task execution of the aircraft.

7.4.1 Test items Many test items of the electromagnetic compatibility test are related to the flight control system, which include: 1. Qualitative electromagnetic compatibility test To verify whether the subsystems and airborne equipment of the flight control system are compatible to work, a qualitative electromagnetic compatibility test can be conducted under different working modes of each subsystem and airborne equipment. 2. Quantitative electromagnetic compatibility test To determine the sensitivity of the airborne equipment of the flight control system to a specific signal or to determine the sensitivity threshold value of the equipment, a conducted susceptibility test should be carried out. In order to verify whether the flight control system meets the requirement of electromagnetic compatibility against conducted interference, it is necessary to conduct quantitative conducted susceptibility test with electromagnetically compatible cable bundles for the airborne equipment of the flight control system.

590 Chapter 7 3. Airworthiness conformity verification test It verifies whether the flight control system and other airborne systems of the aircraft can meet the system compatibility requirements of airworthiness clauses CCAR25.1353 (a) and CCAR25.1431(c).

7.4.2 Test methods 7.4.2.1 Qualitative electromagnetic compatibility test Under different power supply states, set the airborne systems or airborne equipment (electrical system, environmental control system, power system, communication/navigation/ monitoring system, weather/ground mapping radar, etc.) causing interference as the sources of interference and start the flight control system to work normally. Under different working modes of the sources of interference, the ground crew has qualitative inspection of the flight control system and its equipment by visual and auditory means and checks whether the flight control system works normally. The qualitative electromagnetic compatibility test can be divided into the following three states. 1. Mutual interference inspection under operation of ground power unit When the ground power unit supplies power and air conditioning vehicle has air conditioning, start or turn off the sources of interference and check the interference of airborne systems to the flight control system. Operate each control plane and control the flaps (slats) to check the compatibility between the flight control system and airborne systems. 2. Mutual interference inspection under operation of APU Operate the APU to normal state and start or turn off the sources of interference to check the interference of airborne systems to the flight control system. Operate each control plane and control the flaps (slats) to check the compatibility between the flight control system and airborne systems. 3. Mutual interference inspection under operation of engine Start the four engines to work slowly and start or turn off the sources of interference to check the interference of airborne systems to the flight control system. Operate each control plane and control the flaps (slats) to check the compatibility between the flight control system and airborne systems. 7.4.2.2 Quantitative electromagnetic compatibility test The quantitative electromagnetic compatibility test of the flight control system is generally carried out through current injection and cable bundles are generally used in the quantitative conducted susceptibility test. According to the interference level value in curve 3 of CS114 in GJB151A-97, interference signals are applied to the interconnected cables of each subsystem of the flight control system.

Onboard ground test of the flight control system 591 Table 7.22: Test level in conducted susceptibility test. Test frequency range

Test level (dBµA)

10 kHz
1 MHz 1
30 MHz 30
400 MHz

49
89 89 89
77

Signal generator Power amplifier Injection probe

EUT

Power meter System controller Figure 7.31 Interconnected cable bundles in conducted susceptibility test.

In this test, through modulated continuous wave signal (pulse modulation with 1 kHz pulse with duty ratio of 50%), electromagnetic energy is injected at the power cable with a current probe. By testing the interference signal of the corresponding level directly coupled on the cable of the probe, the equipment is checked to see if the interference degree meets the requirements. For the performance of the flight control system, under the condition that the ground power unit supplies power, the DC power cable and data cable between the flight control computer and the airborne equipment of the airborne system are selected for the test. The use of a probe can avoid cable disassembly, ensure the integrity of the system to a large degree, and also reduce the test strength and shorten the test time. The test level value is shown in Table 7.22 and the sensitivity test configuration is shown in Fig. 7.31.

592 Chapter 7 7.4.2.3 Airworthiness conformity verification test The airworthiness conformity verification test method is the same as the qualitative electromagnetic compatibility test method. That is to say, under different power supply stages, it is verified whether the flight control system is compatible with airborne systems and conforms to airworthiness clauses CCAR25.1353(a) and CCAR25.1431(c).

7.4.3 Judging criteria 7.4.3.1 Qualitative electromagnetic compatibility test The judgment of the qualitative electromagnetic compatibility test shall be subject to “The system itself shall be electromagnetically compatible to meet the system working performance requirements” in Electromagnetic Compatibility Requirements of Systems (GJB1389A-2005). For the mutual interference inspection of the flight control system and each system, the ground staff judges whether the flight control system or equipment works normally and if there are faults or unallowable response by visual and auditory means under different working modes of the systems or equipment (sources of interference) causing interference. If the system works stably, the information display is stable, the screen has no delay, and graphics and state parameters have no jumping, it indicates that the flight control system is compatible and its electromagnetic compatibility meets the requirements. 7.4.3.2 Quantitative electromagnetic compatibility test In the quantitative electromagnetic compatibility test, the equipment under test and related support equipment (including integrated processing system and display control system) shall be powered on and started to work normally. In the test process, the electronic and electrical equipment under test shall be set in typical state and interference signal shall be applied according to the level in curve 3 of CS114 in GJB151A-97 to draw the calibration curve automatically, record the test information of the equipment under test and judge whether the equipment sensitive frequency and amplitude meet the requirements of GJB151A-97. 7.4.3.3 Airworthiness conformity verification test Under different power supply states of the aircraft and different working modes of the flight control system, if the flight control system and airborne systems work stably, the control plane does not have uncommanded jumping or delay, the information display is stable, the screen has no delay, and graphics and state parameters have no jumping, it indicates that the flight control system and airborne systems are compatible and the electromagnetic compatibility meets the airworthiness requirements.

Onboard ground test of the flight control system 593

7.4.4 Organization and implementation of the electromagnetic compatibility (E3) test The implementation of the electromagnetic compatibility test is led by the E3 discipline, which is responsible for managing the state of the aircraft and the system, preparing documents necessary for the test such as the test task list and the test outline, providing wave-absorbing materials necessary for the test and test instruments and accessories necessary for the quantitative test, organizing pretest review, providing support to the technical work of the test and conducting testing and recording according to the test outline, as well as providing a test report. As the first coordination unit, the general manufacturer provides equipment necessary for the test such as test site, tested aircraft, traction vehicle, ground power unit, air conditioning vehicle, ground power supply (220 V/AC), ground high-pressure air source, and offsite inspection equipment necessary for the test, undertakes the startup of engine, operation of airborne equipment and data loading, observation or monitoring of relevant equipment, assists in the recording of the test results, and connects or removes the equipment cables according to the demands. The units participating in the test include overall, flight control, and electromechanical disciplines. The installed products provided by the product development unit shall meet the electromagnetic compatibility requirements in the technical agreement or fixed-point agreement and relevant technical support shall be provided. The confidentiality, technical safety, quality, and customers’ representatives shall attend the whole test and have full-process monitoring and supervision of the safety, confidentiality, process, and results at the test site. During the test, the detailed plan shall be formulated and test items, requirements, and personnel for the test shall be specified.

CHAPTER 8

Flight test of the flight control system 8.1 Overview The flying quality of an aircraft is jointly determined by its aerodynamic performance and flight control system. The manual flight control system of large transport aircraft is generally composed of the fly-by-wire flight control system and machinery control system. The fly-by-wire flight control system generally has a normal working mode, degraded working mode, simulated backup working mode, and direct chain working mode. The machinery control system is only an important supplement to ensure the safety of the aircraft. With the progress of technology and the continuous improvement of the reliability of fly-by-wire flight control system, modern aircraft have gradually phased out the machinery control system, such as A380, A400M and B787 aircraft. The flight test of the flight control system has become an important part of the aircraft flight test. Almost all items of the flight test of aircraft shall be carried out under the normal working state of the flight control system (the angle of attack and speed protection functions shall be closed for such flight test items as stalling speed and stalling angle of attack). Therefore the flight control system becomes the core system to ensure the safe flight of aircraft, and also the key system to ensure the smooth and effective flight test. The flight test of the flight control system aims to verify the working state, function, logic, operation stability characteristics, and flight performance of the flight control system in a real environment (structural deformation, vibration, electromagnetic interference, etc.), so as to show that the flight control system works safely and reliably with complete functions and the flying quality of the aircraft meets the design requirements. The main test items of the flight control system flight test include: 1. functions, logic, and operation stability characteristics of manual flight control system under various working modes; 2. functions, logic, and performance of high lift control system under various working modes; and 3. functions, logic, and performance of automatic flight control system under various working modes.

Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00008-5 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

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596 Chapter 8 Most flight test items of the flight control system are carried out jointly with the flight test items of other systems or disciplines. For example, the BIT of the flight control system shall be executed and executed correctly for every flight. The flying quality test shall be consolidated with the flight test item of the operation stability discipline. The automatic flight control functions should be combined with airline operation. The items of the flight control system that need independent tests mainly include the stability margin, boundary protection, autopilot, and modal conversion. Besides the subjective evaluation of pilots, the evaluation of the flight control system shall also be based on objective data. To enhance the accuracy of objective evaluation, specific requirements on the flight test items, test actions, test data accuracy, and sampling rate shall be proposed in 1 2 years before the flight test, so that the flight test department can conduct the necessary aircraft testing and refitting, pilot training, and study of flight test methods, including the preparation of the flight test outline and task list.

8.1.1 Requirements and objectives of the flight test As a key system to ensure flying safety, the necessity of the flight test for the flight control system can be mainly seen in four aspects. The first is that the design department needs to test the consistency between the realization of flight control system and design state through the flight test. The second is that problems should be analyzed and exposed and system design should be improved and optimized according to flight test data. The third is that the design department should show to the evaluation and use department that the flight control system has high safety and reliability through a number of flights. And the fourth is that the flight test data necessary for the evaluation should be submitted to the evaluation department to prove the aircraft function, performance, and flying quality meet the design requirements of the aircraft. The test items of the flight control system flight test are as follows. 1. Check the environmental adaptability of the airborne equipment of flight control system (flight control computer, sensor, and actuator). For example, check the correctness and functional integrity of the electrical and mechanical interfaces of airborne equipment, verify the cross-linking relationship between the airborne equipment and other systems or equipment, and check the conformance of airborne equipment performance indexes with design requirements. 2. Verify the control function and flying quality of the manual flight control system. For example, verify the effectiveness of the BIT function of the flight control system, the safety of mode conversion of the flight control system, the safety of air/ground state conversion, the effectiveness of feedback signal filtering, and evaluate the flying quality.

Flight test of the flight control system 597 3. Verify the control function and flying quality of the high lift control system. For example, verify the control time sequence of the lift augmentation device and the aircraft response in the control process. 4. Verify the functions and performance of the automatic flight control system. For example, verify the connection and disconnection logic of the automatic flight control system, the function logic of the automatic flight control system, the functions and performance of the automatic flight control system, and the function conversion transient state of the automatic flight control system.

8.1.2 Basis of the flight test The basis for the preparation of the flight test requirements of the flight control system is mainly from technical documents including the general requirements of aircraft development, requirements of aircraft conceptual design, design requirements of aircraft control stability, aircraft airworthiness requirements, design report of flight control system, control law design report of fly-by-wire flight control system, control law design report of automatic flight control system, and control law design report of high lift control system. To supplement and improve the requirements in the aircraft design documents, military aircraft should also refer to technical documents such as General Specification for Flight Control Systems of Piloted Aircraft (GJB2191—1994), General Specification of Fly-by-wire Flight Control System of Piloted Aircraft (GJB2878— 1997), General Specification for Automatic Flight Control Systems and Stability Augmentation System (SAS), Control Augmentation System of Piloted Aircraft (GJB3819—1999), Flying Qualities of Piloted Airplanes (Fixed Wing) (GJB185—86), and Flying Qualities Standard for Airplane with Fly-by-Wire Control System (GJB2874—97) and their interpretation documents. The preparation of the flight test requirements of flight control system of civil aircraft is described in detail in chapter 9 of this book. According to the flight test requirements of the flight control system, the flight test requirements for the design and type determination of the flight control system shall be prepared. For special missions, the flight test requirements or coordination sheet of the flight control system shall be prepared separately. The flight test requirements document shall mainly include the following: 1. 2. 3. 4. 5. 6.

functional flight test requirements of the flight control system; logic flight test requirements of the flight control system; performance flight test requirements of the flight control system; flight test state requirements of the flight control system; flight test procedure requirements of the flight control system; and flight data recording requirements of the flight control system.

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8.1.3 Objects of the flight test The control functions of the fly-by-wire flight control system of large transport aircraft are generally realized through the inner loop and outer loop. The inner loop, as the core part, realizes the basic control augmentation function and guarantees the flying quality of the aircraft. The outer loop, as a functional module, achieves various control functions such as attitude hold and boundary protection. The automatic flight control system can only work normally and realize different autopilot functions on the premise that the fly-by-wire flight control system and avionics system work normally. Subsystems of the flight control system are interdependent and there is a precedence sequence in the research and development stage. They have a cross-linking relationship with the avionics system, power system, and landing gear control system of the aircraft, and have a functional impact on special test items. To ensure flying safety, the flight test of the flight control system is generally carried out by gradually enabling the functions of each system. For example, in the early stage of the flight test, signals such as dynamic pressure, static pressure, angle of attack, and sideslip angle have not been corrected and some control functions that affect safety will be closed or gain will be reduced. When the aircraft stall characteristic is tested, the angle of attack protection function needs to be closed. In the test of some machinery control systems, the fly-by-wire flight control system should be partially closed. To ensure flying safety, pilots need to master the use restriction and control skills of flight test objects under various conditions. Therefore, for different versions of flight control system hardware and software, that is, different flight test objects, it is necessary to make clearly different descriptions of the flight test objects and make the disclosures to the flight test units and pilots before the test. For high-risk items, the flight test can only be conducted after test verification on the engineering simulator or the “iron bird” integration test bed of flight control system.

8.1.4 Stages and content of the flight test The stage division of the flight test of the flight control system is the same as that of the aircraft development stage. It can be general divided into three categories, that is, research flight test, type determination flight test, and trial and service flight test. Research flight test includes the preresearch flight test and principle flight test. Aircraft type determination flight test can be divided into the maiden flight of aircraft, adjusting flight test, and evaluation flight test. Trial and service flight test refers to the process of delivering smallbatch aircraft to the user, improving the technical conditions of ground use and maintenance, and improving the operation manual and operating skills of the crew under the conditions of customer use, maintenance, and guarantee. Internationally, the trial and service flight test is aimed at forming application performance and it is a process of

Flight test of the flight control system 599 identifying problems in normal flight training and no special flight test items will be arranged. Thus it will not be further described in this chapter. 8.1.4.1 Preresearch flight test The preresearch flight test of the flight control system demonstrates and verifies the application basis of new principles, new technology, and new equipment of the flight control system through a flight test in order to show the application effect and value of new principles, new technology, and new equipment. With the continuous progress of technology, aircraft purchasers tend to put forward high (or new) requirements for the overall aircraft and control system, which means that aircraft designers must constantly study new technologies and new methods to meet the evergrowing requirements for design indicators. However, the application of new technologies and methods will lead to worries about flying safety and aircraft designers must demonstrate and verify the new technologies and methods through flight test platforms. This situation is particularly obvious when the structure of the flight control system changes greatly, for example, the transformation of the aircraft from the machinery control system to the fly-by-wire flight control system. Conclusions obtained in the preresearch flight test are uncertain and sometimes they will deny the prospect of an engineering application of technologies and methods. Take the research on the forward-swept wing aircraft and swept-back wing aircraft as an example, despite their effective technology, they cannot be applied at the large scale due to the lack of practicality. Even if the application prospect is clear, a large amount of engineering application research is still needed before they are used in aircraft design. When the flight control system plans to adopt advanced airborne equipment and control algorithms, it usually adopts the “other aircraft flight test” method to carry out research in advance to improve the technology maturity. Take the newly applied technologies and algorithms such as distributed intelligent sensor, load reduction control algorithm, and hard aerial refueling device as examples, as the design basis (aerodynamic data, etc.) and application conditions (aircraft elastic deformation and vibration form, etc.) may vary largely from the actual state, the technology and algorithm shall be feasible and valid through the flight test first. 8.1.4.2 Aircraft principle flight test The aircraft principle flight test is a flight test conducted during the aircraft engineering development and shall be performed after the designers complete various laboratory tests and the onboard ground test. It is conducted to verify the functions, performance, and technical indicators of new key systems (or important airborne equipment), adjust system parameters, find out and eliminate system design and manufacturing faults and defects, and correct the system technical instructions and operation and maintenance instructions.

600 Chapter 8 If major defects in design and manufacturing are found through the principle flight test or the design indexes still fail to be reached through parameters adjustment, the designers may improve the design and conduct the principle flight test again after the simulation test. Thus the principle flight test may be conducted more than once in the engineering development stage. The system level flight test is mainly conducted in this stage for the flight control system to have overall verification of the system. For example, the J８IIACT demonstration engine is used in China to have overall verification of the control law design and active control technology of fly-by-wire control aircraft. The integrated in-flight simulation test aircraft in the flight test center has the flight test for the aircraft control system and flying quality. 8.1.4.3 Aircraft type determination flight test 8.1.4.3.1 Maiden flight of aircraft

The maiden flight is just the maiden flight test. The maiden flight stage can be roughly divided into preparation for the flight test, ground taxi, and maiden flight. The preparation for the flight test of the flight control system requires the completion of test verification of the monitoring system, maiden flight profile and emergency program training, aircraft control efficiency verification, onboard ground test, and ground taxi test, etc. 1. Test verification of the monitoring system The monitoring system refers to a ground real-time monitoring system. The monitoring page of the flight control system is mainly used for monitoring the working state of the flight control system (including the power supply state, hydraulic source state, state of flight control computer, sensor, actuator, control plane, and aircraft attitude and angular speed) to provide the basis for ground commanders to make correct decisions. Fig. 8.1 shows a typical monitoring screen of the flight control system. The flight control system information page of the monitoring system shall contain at least the working state of the main equipment of the flight control system, control plane motion parameters and aircraft motion parameters, and can display the sources of information (test system or telemetry). 2. Maiden flight profile and emergency program training The maiden flight profile and emergency program training shall be conducted on an engineering simulator or “iron bird” integration test bed of the flight control system. The maiden flight profile training is designed according to the aircraft configuration for the maiden flight (state of landing gear: released generally; and state of lift augmentation device: under takeoff configuration generally) and the working state of the flight control system (direct chain, digital fly-by-wire, etc.). The flight program training for the traffic pattern of the maiden flight covers normal operation program, single engine interrupt program, single engine takeoff program, and single engine

Flight test of the flight control system 601

Figure 8.1 Typical monitoring screen of flight control system.

landing program. The emergency program training covers minimum safe control state training, control plane clamping stagnation, all engine failure and drift down, emergency landing, and landing of lift augmentation device at an abnormal position. 3. Aircraft control efficiency verification Aircraft control efficiency verification verifies whether the control efficiency of aircraft longitudinal control plane is consistent with the state of aircraft taxi test so as to provide a direct basis for predicting the front wheel lifting speed and liftoff speed of aircraft. This work should be conducted on the “iron bird” integration test bed of the

602 Chapter 8 flight control system and the engineering simulator to fully verify the accuracy of the low-speed design data and ensure the pilot’s acceptance of the conclusion of the engineering simulator test (and the “iron bird” integration test of the flight control system) and the establishment of its confidence for the maiden flight. 4. Onboard ground test The onboard ground test has many test items and almost all airborne systems should have the onboard ground test. The content related to the flight control system of the onboard ground test includes the onboard ground test and structural mode coupling test of the flight control system, which are the last steps of the design department fully examining the design of the flight control system. The BIT function inspection, transmission ratio inspection, stability margin test, and man machine closed-loop flying quality inspection of flight control system should be carried out. For details, please refer to chapter 7. 5. Ground taxi test The ground taxi test includes low speed, medium speed, and high speed taxi tests. For the flight control system, the ground taxi test is conducted to check whether the flight control system works normally under the condition of taxiing vibration, check the characteristics of aircraft taxiing, check the front wheel control ability and balance keeping ability of the aircraft, and check whether the ground real-time monitoring system works well. After taxiing, the data of the airborne test system shall be extracted for comparison and analysis to check whether the data can truly reflect the state of the flight control system. The maiden flight is faced with both psychological and technical pressures. To ensure the success of the maiden flight, the quality and safety monitoring department shall hold the maiden flight evaluation meeting. The following aspects are mainly reviewed in the meeting. a. b. c. d. e.

Whether the maiden flight outline and task list are reasonable. Whether the aircraft state is suitable for flight. State of the accompanying aircraft and task list. Whether the test equipment meets the requirements for maiden flight. Whether the pilot and commander have the technical and psychological quality and physical health to meet the requirements. f. Whether ground monitoring conditions and personnel meet the requirements. g. Whether aircraft maintenance and flight guarantee conditions conform to the requirements. h. Whether the documents for the flight test are complete and correct. Among them, the discipline of the flight control system should focus on a, b, d, f, g, and h. No matter how the environment changes and how much the pressure is, the maiden flight must adhere to two principles. The first is the pilot must be confident for the safety of the maiden flight and the second is the maiden flight shall achieve specific goals.

Flight test of the flight control system 603 8.1.4.3.2 Adjusting the flight test

After the success of the maiden flight of the aircraft, the flight test will enter the adjusting flight test stage. The adjusting flight test of the flight control system is mainly carried out to test and improve the function, performance, and flying quality of the flight control system through the flight test and provide the basis and technical support for the application for the aircraft design and type determination flight test. Generally, the following conclusions shall be drawn from the adjusting flight test of the flight control system. a. The airborne equipment of the flight control system has no quality problems and can ensure the safety of the aircraft. b. Air and ground crew are familiar with the use, support, and maintenance of the flight control system. c. The airborne test system works reliably and the test record is complete. d. Technical problems of the flight control system have been solved, airborne equipment works stably and meets performance index requirements. e. The function and logic of the flight control system meet the design requirements. f. The flight control system can ensure the aircraft performance and flying quality meet the design requirements. g. The reliability, maintainability, testability, and supportability of the flight control system are expected to meet the design requirements of the flight control system. h. Aircraft support facilities, equipment, tools, spare parts, technical data, and other support resources are complete. 8.1.4.3.3 Design and type determination flight test of military aircraft

The design and type determination flight test is a process of obtaining flight test data necessary to support military aircraft to obtain the design and type determination certificate. This definition is also applicable to the design and type determination flight test of the flight control system. The design and type determination flight test of the flight control system is mainly conducted to determine whether the functions and performance of the flight control system meet the approved tactical and technical requirements and provide necessary data for the Manual for Pilots and Manual for Flight Crew. The design and type determination flight test must be undertaken by a state-authorized flight test evaluation unit. The airborne systems and equipment for the installed flight test must be in the approved technical state. The outline of the design and type determination flight test shall be drafted and prepared by the flight test evaluation unit according to approved tactical and technical requirements and type determination state and development conditions after

604 Chapter 8 consulting the opinions of the design and development unit, and they are then submitted to the Committee for Type Determination of Aviation Military Products for approval. 8.1.4.3.4 Conformity certification flight test of civil aircraft

The conformity certification flight test of aircraft is one of the conformity certification methods for new (or refitted) civil aircraft, with a purpose to demonstrate the conformity of aircraft and systems to various aircraft airworthiness standards such as civil aviation regulations, to provide the basis for the conformity certification of aircraft and to provide necessary data for the Aircraft Manual and Maintenance Manual. The basis for the conformity certification of China’s civil aircraft is relevant clauses of CCAR25 airworthiness standards. The outline of the flight test is reviewed by the aircraft conformity certification group and approved by the aircraft conformity certification committee. During the flight test, the aircraft conformity certification group shall be informed to observe the flight test on site. Regarding the conformity certification flight test of civil aircraft, please refer to chapter 9 for details.

8.1.5 Methods and requirements of the flight test To obtain good results of the flight test and ensure valid test data, specific flight test methods and requirements shall be determined for each flight test. No matter how flight test methods and requirements change, flying safety is always the priority. In other words, the pilot can stop the test according to the test state at any time in the flight test. For example, as long as the aircraft has obvious shaking in the stalling angle of attack test, the pilot has the right to stop the test immediately and report the condition even if the aircraft does not reach the expected shaking angle of attack. In the longitudinal short-period test, “3211” control actions are adopted and the pilot has the right to stop the test and report the condition immediately once there is uncontrollable divergent oscillation when it gradually increases the input amplitude. The flight test methods and requirements of the flight control system mainly include flight state requirements, flight environment requirements, flight control system working state requirements, test procedures requirements, precautions, data recording requirements, and evaluation results. 1. Requirements for flight state mainly cover aircraft weight, aircraft center of gravity, flaps (slats) state, landing gear state, flight altitude, and flight speed. 2. Requirements for flight environment mainly cover atmospheric temperature, wind speed, wind direction, runway state, and runway slope. 3. Requirements for the flight control system working state mainly cover the working mode of the flight control system, as well as the information of the flight control system functions, use limit, and emergency operation procedure.

Flight test of the flight control system 605 4. Test procedures are just a series of combinations of specific actions made for a specific test item to reach the test objectives. The test objectives of the flight control system should be clear to the flight test unit and the pilot. Thus no special explanation will be made. The test procedures generally have the following characteristics: a. Standard: the test procedures shall clearly define test entry conditions, control actions, test exit conditions, and precautions, which are conducive to the consistency of test conclusions and data analysis. b. Progressive: to ensure flying safety and accurate completion of prescribed procedures, the pilot should conduct several test operations to accurately complete the test procedures. c. Flexible: in the test procedures, the amplitude of the control action should often be determined by the pilot according to actual conditions, which is determined by difference of the aircraft. For example, for longitudinal stability performance, “3211” or “times pulse” control actions (see Figs. 4.3 and 4.4) can be adopted and its control amplitude increases gradually from small to large until the aircraft presents a typical short-period response. 5. Precautions refer to the descriptions that are made to ensure the test conditions or control actions in the test. Precautions that should be noted in the ground static performance test of aileron are as follows: a. Unless otherwise specified, the aileron shall be adjusted and calibrated to the neutral position. b. The SAS shall be closed. c. External power supply and hydraulic source (if necessary) shall be used. d. Ground crew should wear earphones to observe at the tail of the aircraft to confirm if the limit position of control plane is reached. e. A complete scanning (neutral-right limit-neutral-left limit-neutral) should take about 90 s. 6. The data recording requirements are regulations made for the parameter name, variable name, dimension, range, accuracy, and sampling rate of data to be collected as well as video and audio for data analysis. The amount of data that should be collected for the flight control system is about 1000 and the sampling rate of most signals should be consistent with the working frequency of airborne equipment. It can be even lowered to 10 Hz but shall reach at least 25 Hz for important signals (such as attitude angle and angular rate). 7. The initial evaluation of flight test results depends on the subjective evaluation of pilots. In most cases, subjective evaluation is a description of the overall effectiveness of the aircraft and attention should be paid to it by flight control system designers. As the subjective evaluation is affected by many factors, such as pilot’s experience, evaluation ability, psychological and physiological state, different scores may be obtained for the same state and same subjects and the score may even be very different.

606 Chapter 8 Therefore the pilots should have flight experience of similar aircraft, have received the training on subjective evaluation methods such as PR and PIOR, and have mastered the evaluation language for the flying quality and pilot-induced oscillations (PIO) trend shown in Figs. 4.5 and 4.6. In addition to subjective evaluation, objective evaluation is also required for finishing the type determination (“conformity certification” for civil aircraft) of the aircraft. Objective evaluation depends entirely on flight test data. Given the complexity of aircraft motion and the system working state, the flight test methods and requirements of the flight control system above shall be strictly executed to obtain the quantitative expression completely by a mathematical method.

8.1.6 Ground support facilities The ground support facility guarantees the technical advancement, safety, and efficiency of the flight test. The flight test of the flight control system shall have the following flight test facilities. 1. Airborne data acquisition and recording system It is capable of recording and sending all data of flight control system operation to the ground (such as control signal, working state, sensor signal, and redundant voting value). 2. Test signal generation system It is used to generate the command signals required by the flight control system, such as the control plane frequency sweep signals in the stability margin test and the fault signals in the fault test. 3. Ground monitoring system It is used to display the flight test items and safety information of the flight control system in real time, so that the ground monitor and the commander can make correct and prompt judgment and decisions, so as to ensure the flying safety and effective flight test items. 4. Flight simulation system The flight simulation system has aircraft simulation equipment that can truly reflect the functions and performance of the flight control system, such as engineering simulator and the “iron bird” integration test bed of the flight control system. It mainly acts to familiarize the pilot with the control procedures of the flight control system, the basic performance of the aircraft, the fault handling procedures of the flight control system, conduct exercise of the flight test, confirm the flight test methods, and explore flight test technology. The flight simulation system shall have the following functions and performances: 1. The information of the control display equipment in the cockpit shall be consistent with that of real aircraft.

Flight test of the flight control system 607 2. The form and performance of artificial feel system shall be consistent with that of real aircraft. 3. The resolution ratio of the vision system near the aircraft shall not be greater than 1 m and the aircraft shall have solid modeling. 4. The maximum delay shall not exceed 120 ms. 5. The 6-DOF motion simulation of combined shaft system of atmospheric motion shall be considered based on elliptically rotating earth. 6. Simulation of multiple-wheel and multistanchion landing gear system under brake. 7. Simulation of power system considering dynamic response and ambient temperature. 8. Simulation of mass and inertia considering the aircraft mass changes. 9. Simulation of special motion effects (especially transient state), revised according to flight test data. 10. The aerodynamic taxiing data package (lower than 150 km/h) shall be confirmed by the pilot. 11. The aerodynamic data package (greater than 120 km/h) shall be corrected through flight test data and confirmed by the pilot. 12. Simulation of typical faults (such as single engine failure, control plane clamping stagnation and system degrading).

8.1.7 Organization and management of the flight test The aircraft flight test involves the flight test unit, the chief designer unit, the production and manufacturing unit, the management side and the user side and shall be jointly organized and implemented by the parties above. The chief designer unit shall set up a flight test team at the beginning of the design and the flight control system designer shall be a main member of the team (called the director engineer of the flight test of the flight control system). According to the general aircraft development requirements, the flight test requirements analysis of the flight control system shall be carried out and implemented in the design of the airborne system and airborne equipment. In 1 2 years before the start of flight test, the chief designer unit shall clearly put forward flight test requirements of the flight control system and participate in the preparation of the flight test plan and flight test outline, so that the flight test unit can complete flight test refitting and flight test training. The chief designer unit shall also control and guarantee the technical state of the aircraft before the test, independently analyze the flight test results after the test, and timely analyze the causes of and address the problems exposed in the test. In addition, the aircraft flight test shall fully emphasize the ownership of intellectual property, fully share the flight test data, ensure that the wind tunnel test unit can complete the calibration of wind tunnel parameters and aerodynamic correction coefficient, and the chief designer unit can correct aircraft aerodynamic data and system data.

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8.1.8 Team training of the flight test The flight test team plays an important role in the flight test of the flight control system. The flight test team of the flight control system mainly has a membership of pilots, flight test engineers, and flight control system designers. The training of the flight test team aims to make the team members be familiar with the flight control system control procedures, aircraft performance, and flight control system emergency control procedures as early as possible. Pilots are an extremely important part of the flight test team and they play a crucial role in ensuring the safety of the flight test and improving the efficiency of the flight test. They shall have a profound theoretical basis and skilled flight test technology, be familiar with flight test objects, and have strong communication skills. Flight test engineers shall be familiar with flight test methods, master relevant flight test technologies and basic flight test pilot technology, and master aircraft response characteristics and emergency handling measures. As the main force for the fault analysis in flight test, flight control system designers shall be familiar with flight test technologies, participate in flight test data processing and analysis, participate in flight test task arrangement meeting and flight test result analysis meeting every day, and understand the faults and problems in flight test. The flight control system designers shall also participate in the field monitoring of the flight test and assist in analyzing and dealing with the abnormal phenomena in the process of the flight test. The main methods by which the flight test team can understand and get familiar with the flight control system are to participate in the flight control system training, read the design materials, and conduct the engineering simulator test and the “iron bird” integration test of the flight control system.

8.1.9 Flight test plan As the flight test of the flight control system is limited by many factors (such as development stage, calibration of sensor of cross-linking system, data correction for design, state of equipment, and weather conditions), the flight test plan of the flight control system shall follow the overall flight test arrangement of aircraft and the flight test items can only be tested after all conditions are available. Generally, the flight control system of large transport aircraft includes a fly-by-wire flight control system, high lift control system and automatic flight control system and each subsystem has different working modes. The flight test of the flight control system is divided into several stages according to the design state to complete the flight test tasks of

Flight test of the flight control system 609 each subsystem, and the flight test of each subsystem shall first ensure safety and then gradually release functions. The flight test plan of the flight control system can be roughly divided into the following stages: Stage 1: Preliminary basic function and performance test of flight control system in maiden flight. In the first a few flights of flight test, the flight control system will work in a relatively safe mode. In other words, some functions of the flight control system (such as boundary protection function related to angle of attack and sideslip angle, and the automatic flight control function) are closed, the parameters that may affect the safety, such as the angle of attack, and that are not used as the main feedback signal in the control law are set as constant values, and the aircraft configurations are fixed (such as landing gear is kept in the released state and flaps/slats are kept at takeoff position). In these flights, the BIT function of the flight control system, atmospheric data signal, flight control sensor signal, wheel load signal, and actual working state and stability of lift augmentation device are checked and the results are taken as the basis for gradually opening the functions of the flight control system later. Stage 2: Basic function, performance, and stability margin test of manual flight control system within 80% envelope. In the early stage of the flight test, the aircraft speed, altitude, weight, and center of gravity in the flight test will be generally limited to ensure flying safety. Under this condition, most functions that are manually controlled (three-axis control augmentation, attitude hold, takeoff and landing configuration transformation, etc.) have been opened and flight test items related to the flying quality (short-period response and level flight acceleration and deceleration, etc.) are tested. Stage 3: Limit and protection function test of flight control system. After the calibration of atmospheric data and correction of the flight control sensor, the relevant flight control system limit and protection functions can be put to the flight test, such as angle of attack limit and protection, overload limit, and attitude hold. Stage 4: Basic functions, performance, and stability margin test of manual flight control system after extension of envelope. After the flight test within 80% envelope is completed, the limits on aircraft speed, altitude, weight, and center of gravity are gradually released to gradually increase to the service envelope boundary. Under this condition, some functions of the flight control system may need to be closed, such as the overspeed protection function in the maximum speed flight test and the angle of attack protection function in the minimum stalling speed flight test. Stage 5: Flight test of automatic flight control system. After the manual flight control system works stably and the flight test is preliminarily completed, the flight test of the automatic flight control system can be

610 Chapter 8 carried out. The on/off, basic logic, and function test of the automatic flight control system shall be conducted first, and then the combined function and function switch test, navigation function test, and approach and go-around function test shall be carried out successively. The flight director function test of the automatic flight control is carried out synchronously with the tests above.

8.2 Requirements for the flight test of the flight control system 8.2.1 Basis of preparation The main basis used for the preparation of the requirements for the flight test of the flight control system includes aircraft design and type determination document (or conformity certification standards of aircraft), general aircraft development requirements, the requirements to the flight control system in the general technical requirements for the aircraft, requirements for flying quality of the aircraft, requirements of the flight control system design specification to system functions and control law functions, as well as the control law parameters adjustment law.

8.2.2 Items and requirements of the flight test The content of the flight test of the flight control system shall fully cover the requirements for the design and type determination of aircraft (or aircraft conformity certification requirements) as well as consider other aircraft design requirements. The flight test of the flight control system generally includes: 1. 2. 3. 4. 5. 6. 7. 8. 9. 10.

installed adaptability of airborne equipment of flight control system; functions of flight control system; work logic of flight control system; modal conversion of flight control system; flying quality of manual flight control system; functions of high lift control system; functions of high lift control system; logic of automatic flight control system; functions of automatic flight control system; and performance of automatic flight control system.

8.2.3 Requirements of the monitoring system The ground monitoring system is a necessary facility to ensure flying safety and improve flight test efficiency. The ground monitoring system has requirements in two aspects, requirements for the monitoring personnel and requirements for the monitoring system.

Flight test of the flight control system 611 Before the flight test, the monitoring personnel shall receive the training on the engineering simulator and the “iron bird” integration test bed of the flight control system, be familiar with the monitoring objects, judging criteria, and emergency handling measures, and master the air ground communication mode and password. During the flight test, the monitoring personnel shall follow and monitor the pilot’s operation and also the working state of the flight control system and the pilot’s control in the whole process, communicate and interact with the pilot at any time, and remind the precautions for next test task until the pilot leaves the aircraft safely. If it is found that the flight test and aircraft response do not meet the requirements, the pilot shall be asked to make rectification in time. The judging criteria are used to judge whether the aircraft is in a safe flight or whether the flight test data are effective. These can be used to judge whether the main airborne equipment of the flight control system (flight control sensor, flight control computer and actuator, etc.) works effectively, whether cockpit control signals and angular rate signals are effective and within the normal range, whether the main control plane is stagnated, and check the working state of the airborne equipment of the flight control system and the key state information required by the control. The judging criteria are jointly formulated by the flight control system engineer, flight test engineer, flight test commander, and pilot. During the flight test, they can be revised according to the actual situation of the aircraft. Under the following conditions, the flight test shall be stopped. 1. A or a few judging criteria of the monitoring are broken (e.g., the signal redundancy of the angular rate gyro assembly reduces to single redundancy). 2. Abnormal response of aircraft or critical system fault is found (such as occurrence of noncommanded response and fault of high lift control system). 3. Aircraft response deviates from flight simulation results seriously (stalling angle of attack is lower than the design value). 4. The pilot feels physically uncomfortable. 5. Sensor of key test system has fault and cannot be restored in specified time. 6. Important monitoring information is interrupted and cannot be restored in specified time. 7. False alarm occurs during the flight test. Under this condition, the mission shall be stopped at first and it can only be reexecuted after the ground commander confirms that it is a false alarm. 8. The pilot or commander reckons it is an abnormal phenomenon and the flight test shall be stopped. The monitoring system receives airborne test data and telemetry data, which provide monitoring images and judging criteria for professionals and are used as a basis for communication with pilots. Therefore the monitoring system shall have good real-time performance and the time delay of key signals shall not exceed 100 ms. In addition, it shall

612 Chapter 8 have sufficient transmission bandwidth (no less than 10 Mbps) to ensure the quantity of transmitted monitoring signals, have sufficient display images that are easily switched and marked with a judgment criterion boundary obviously, and have a secondary data processing function to quickly calculate the stability margin and flying quality grade of the flight control system.

8.2.4 Requirements of the testing system The testing system mainly refers to the equipment system that collects and records flight test data and it is composed of the airborne test system and telemetry monitoring system. Most of the flight test data comes from the airborne test system and the telemetry monitoring system is mainly used in the early stage of the flight test. After the atmospheric data system and inertial navigation system are modified by the flight test data, the data of the telemetry monitoring system will not be used but the data of the airborne test system are used. About 1000 flight test parameters are required to be collected and recorded by the testing system, mainly including control system parameters (such as control force, control displacement and trim command), operation parameters (such as working mode of flight control system, work description of high lift control and control channel), flight parameter (such as altitude, speed, angle of attack, overload, attitude angle, and angular rate), aircraft state parameters (such as configuration state, landing gear state, power system state, aircraft vibration, weight, and center of gravity), flight control bus data (such as switch/knob control, system parameters, voting parameters, and sensor parameters), control plane parameters (such as control plane deflection angle and control plane motion speed), and flight control alarm data (such as the control plane failure and loss of function). It should be noted that most of the signals required by the flight control system are existing signals from the airborne computer. However, in the early stage of the test, airspeed head parameters, angle of attack, sideslip angle, angular rate, and overload should better be measured with special test equipment because these parameters are closely related to the installation position and have a great impact on the flight control system.

8.3 Outline of the flight test of the flight control system The outline of the flight test of the flight control system is a part of the outline of the aircraft type determination test. According to the design function and performance of the flight control system and the parameter adjustment law, the flight test items and procedures are formulated to prove the flight control system can satisfy the requirements of design and type determination documents (or aircraft conformity certification requirements), general aircraft development requirements, general technical requirements of aircraft, aircraft flying

Flight test of the flight control system 613 quality requirements, and requirements of flight test guidance documents specified in the design specification of the flight control system.

8.3.1 Categories of test outlines As every flight test stage is certainly pertinent and targeted, a specific flight test outline for each stage shall be prepared. As the objective of the flight test in the research flight test stage is relatively simple, its flight test outline is independent and particular. The content of the flight test in trial and service flight test stage is generally deleted on the basis of the items of the design and type determination (or conformity certification) flight test. Therefore only the outline of the flight test in the design and type determination stage is introduced. According to the arrangement of the flight test, the flight test outline in the design and type determination (or conformity certification) stage can be basically divided into the maiden flight test outline, adjusting flight test outline, and design and type determination flight test outline (or conformity certification flight test outline). 1. The maiden flight stage covers the maiden flight and first several flights of the aircraft. The flight control system in the maiden flight stage mainly verifies the working state and manual control functions of the flight control system and compares the test with the engineering simulator test and “iron bird” integration test of the flight control system to make full preparation for the subsequent gradual release of the flight control system functions and flight envelope. The outline of the maiden flight test specifies the flight test tasks in the maiden flight stage. In principle, the flight control system shall use the safest (or the most confident) working mode for the maiden flight. The mode may require closing some functions and sensors. In the flight test outline, the flight control system shall specify the working mode (such as the working state of airborne equipment such as PFC and the version state of flight control software), the working mode of flight control sensor, the verification mode of high lift control system, available functions, and their performance, etc. 2. After it is proved that the working state of each system of the aircraft (which should be changed and improved) is consistent with the design state in the maiden flight stage, the flight test work is transferred to the adjusting flight test stage. In this stage, the flight test should still be carried out within a relatively safe envelope range and the flight control system should gradually release the functions until all functions within the whole envelope are covered. Besides, the functions, performance and flying quality shall be preliminarily evaluated to achieve the conditions for design and type determination flight test. Therefore the adjusting flight test outline shall make overall requirements on the flight test items on the basis of the functions, performance, and flying quality of the flight control system. The number of selected

614 Chapter 8 flight states can be small, but they shall be typical and representative, so as to cover the whole flight envelope. 3. Design and type determination flight test is a stage to have full-state and full-envelope evaluation of the aircraft. Military aircraft are generally evaluated and reviewed by organizations or departments designated by the aviation products type determination committee. The outline of the design and type determination flight test shall specify the flight test items regulated by the design and type determination requirements clearly and clause by clause, specify corresponding design and type determination clauses, and also specify the requirements for the flight test stage, flight test procedures, flight test actions, and flight test data. In this stage, close communication shall be conducted with the evaluation and review department to ensure the interval of the flight test meets the requirements of the department. 4. The outline of a conformity certification flight test is a flight test outline prepared for civil aircraft to obtain an airworthiness certificate. Airworthiness certification institutions in different countries have issued specific airworthiness certification standards and the flight test outline shall be prepared according to the airworthiness certification standards generally. What should be pointed out is that airworthiness certification has very strict requirements on flight test data. Thus experienced flight test engineers and pilots shall be arranged in the flight test to ensure the validity of data. In addition, the airworthiness certification has witness items and the test procedures shall be well coordinated with the airworthiness institution in advance so as to arrange video records.

8.3.2 Basis for preparation of the flight test outline The basis used for the preparation of flight test outline is the aircraft design and type determination document (or aircraft conformity certification standard). It also involves general aircraft development requirements, general technical requirements for aircraft, aircraft flying quality requirements, requirements of flight control system design specification as well as control law design requirements sometimes. The aircraft design and type determination document specifies the functions, performance, and flying quality of the flight control system that must be verified by flight tests. The general aircraft development requirements stipulate the aircraft flying quality requirements and functional requirements of the flight control system. The requirements of aircraft conceptual design specify the functional requirements of the flight control system and a series of flight limit conditions such as aircraft weight-center of gravity envelope, altitude-speed envelope, and weight-overload envelope. The flying quality requirements of aircraft specify the specific division of flight stages of the flight control system and the flying quality provisions that shall be satisfied in each stage. The design

Flight test of the flight control system 615 documents of the flight control system describe the design details such as working conditions, degradation conditions, feedback gain regulation rules, and air/ground logic. For civil aircraft, documents including CCAR25 China Civil Aviation Regulations, FAR25 and 33 Federal Aviation Regulations and JAR25 Joint Aviation Requirements can be referred to. The flight test directly reflects the basis used for the preparation of flight test outline. In the outline, the test objectives shall be clearly specified so that professionals could understand flight test procedures and flight test actions well.

8.3.3 Selection of flight test items Flight test items are specific and executable flight test content in the flight test outline. Flight test items directly determine the quantity, period, and risk of the flight test work. When selecting flight test items, the design state of the aircraft (aircraft weight, center of gravity, and overload, etc.) shall be fully considered. On the basis of the aircraft performance and the design results of flight control system, not only the design and type determination requirements shall be covered, and the flight test items shall be reduced as far as possible to improve the efficiency of the flight test. Generally, the selection of flight test items should follow the following principles. 1. For the items that require the provision of flight test data in the aircraft design and type determination document and general aircraft development requirements, verification by flight tests shall be arranged. For other content used for research of the flight control system, the chief engineer of the flight control system and the director engineer of the flight test of the flight control system can determine through negotiation. For example, the performances of manual flight control system and automatic flight control system are both significant and quantifiable indicators. 2. The content that affect flying safety shall be tested, such as flying quality, stability margin, and modal conversion. 3. For the doubtful items found by the flight control system designers through calculation and analysis or in the ground test [such as the large difference in roll performance under different flaps (slats) configurations of the aircraft, large jumping of design data and occasional oscillation of actuator system] and also items with design data in or close to the critical condition (such as the stalling speed, stalling angle of attack, aerodynamic performance in taxi stage, sensor filtering and control law parameter), they shall be tested in the test flight. 4. The flight test items that shall be verified with emphasis according to use requirements and previous flight experience shall be selected.

616 Chapter 8 5. For flight test items that shall be selected according to relevant specifications and standards, those typical, valid, and realizable flight test items with great influence shall be selected for flight test (such as PIO trend evaluation flight test and ground effect flight test).

8.3.4 Selection of flight test status Due to the large flight envelope (such as speed-altitude envelope, weight-center of gravity envelope, weight-overload envelope) of the fly-by-wire aircraft and many aircraft configurations [such as flaps (slats) configuration, loading configuration and plug-in configuration], an appropriate flight test state shall be selected for each flight test item to cover the whole envelope range. Meanwhile, the flight test state and tolerance range shall be considered comprehensively in order to arrange several flight test items in one flight. The selection of flight test state shall consider the flight stage, aircraft configuration, landing gear state, throttle state, aircraft weight, aircraft center of gravity, flight altitude, flight speed, runway state, flight control system state, and weather conditions. 1. The state of the flight control system must be specified for the flight test of the flight control system. The main reason is that the design and type determination document (or conformity certification standard) requires that the function and performance of the flight control system must be verified by flight test under a designated working mode. For example, the flight control system has an analog control mode and the digital fly-by-wire control function must be closed in the flight test under the analog control mode. 2. The flight stage mainly describes the operation state of the aircraft. For example, the control system performance test of the flight control system shall be completed on the ground. Table 8.1 shows the definition of aircraft flight stages. 3. Aircraft performance varies greatly due to different configurations. In principle, the aircraft configurations shall be fully covered unless aircraft data and flight evaluation show that the change in control plane performance under different configurations does Table 8.1: Definition of aircraft flight stages. No.

Flight stage

Definition

Symbol

1 2 3 4 5 6 7 8 9

Stop Taxi Takeoff Climbing Cruise Descending Approach Go-around Landing

The aircraft is loaded and is stationary to the ground The aircraft is loaded and is moving to the ground The process that the aircraft taxis to reach the safety altitude The process that the aircraft climbs to preset altitude The process that the aircraft flies within the preset altitude range The process that the aircraft descends to the preset altitude The process that the aircraft enters the site according to a normal procedure The process that the aircraft has go-around according to normal procedure The process that the aircraft lands according to normal procedure

STOP TAXI TKOF CLB CRZ DSD APP GOA LND

Flight test of the flight control system 617

4.

5.

6.

7.

8.

not exceed 10% or there is no significant change in pilot control feeling (control force and control displacement). Landing gear state mainly considers the conditions that may be involved under normal operation procedures of the aircraft (if it is required to release the landing gear at 12 km from the airport, the landing gear may be released at different position due to different entry routes. Therefore the performance of the flight control system under this condition shall be verified), but the flaps (slats) configurations when the landing gear cannot be withdrawn or released in the functional hazard analysis of the flight control system shall also be considered. Specific requirements shall be made for the throttle state in the test. For example, in the single engine failure test during takeoff, to ensure the safety of the aircraft and the state of the engine, the single engine failure is usually simulated under a condition that the aircraft operates at a slow speed. To approach to the real situation, pilots are required to quickly pull the critical engine throttle lever to a slow speed. To cover the actual available weight of the aircraft, the weight range of the aircraft is usually divided into 3 5 zones for flight test. The specific number of zones is determined by the weight change range of the aircraft and the characteristic change condition of the aircraft. In the flight test outline, the requirement of the minimum and maximum weight and other limit conditions of the aircraft shall be proposed independently in the test items. The weight division specified in the flight test requirements of a certain type of aircraft is shown in Table 8.2. The center of gravity of the aircraft must be verified in the flight test of the flight control system. Due to the different loading and mounting configurations of the aircraft, the center of gravity of the aircraft can generally change up to 30% MAC. Therefore the center of gravity shall also be divided into 3 5 zones during the flight test. For civil aircraft, the change of center of gravity is relatively small and generally three centers of gravity are used. Table 8.3 shows the definition of the range of center of gravity in flight test of a certain type of aircraft. The requirements for flight altitude in flight test are relatively relaxed. To ensure flight safety, the required flight altitude for many flight test items is far higher than the actual altitude and flight speed is increased to simulate the aerodynamic load. Table 8.4 shows the definition of the altitude used in flight test of a certain type of Table 8.2: Definition of aircraft weight in flight test of a certain type of aircraft.

No. 1 2 3 4 5

Weight description

Definition

Symbol

Light weight Medium weight Heavy weight Heavy takeoff weight Heavy landing weight

Aircraft weight lower than 120 t Aircraft weight being 120 180 t Aircraft weight greater than 180 t Aircraft weight greater than 200 t Aircraft weight being 180 210 t

LIGHT MID HEAVY MTOW MLDW

618 Chapter 8 Table 8.3: Definition of range of center of gravity of a certain type of aircraft. No. Description of center of gravity 1 2 3

Front center of gravity Normal center of gravity Rear center of gravity

Definition

Symbol

Center-of-gravity position lower than or equal to 30% MACFWD Center-of-gravity position being 30% 36% MACMID Center-of-gravity position equal to or greater than 36% MACAFT

Table 8.4: Definition of altitude used in flight test of a certain type of aircraft. No. 1 2 3 4 5 6

Configuration description

Definition

Symbol

Ground Low Middle High Cruise Ground effect

The altitude where the aircraft stops Altitude lower than 3500 m Altitude being 3000 8000 m Altitude greater than 7000 m Cruise altitude of aircraft Aircraft span altitude

GROUND LOW MID HIGH CRZ GEFT

aircraft. For the flight altitude with specific requirements, it shall be shown independently in the flight test items. 9. Speed is the most complex requirement among all flight conditions. In the flight test outline, the flight boundary speed is generally shown in symbols mainly because these speed values are only theoretical calculation results before the flight test and will be corrected with the progress of the flight test. For the intermediate speed, it is usually divided and rounded according to the indicated airspeed and the general speed interval region is about 40 60 km/h to meet requirements. 10. The takeoff and landing runway can be divided into cement runway and soil runway by pavement material, into dry runway, wet runway, and icy runway by pavement humidity, and into runways of different slopes by runway pavement zones. In the flight test outline, the above states of runway shall be clearly specified. If possible, the airport and runway can also be specified. Generally, the flight test is carried out under good weather conditions. However, some flight test items have clear requirements on weather conditions, such as strong crosswind, high temperature, high humidity, and extreme cold. As there will be certain surface wind, that is, wind speed and direction, even under good weather condition, to ensure the validity and availability of flight test data, each flight test item in the flight test outline shall clearly stipulate the weather conditions for the flight test. For general flight test items, the required wind speed shall be lower than 9 m/s and it shall be recorded at any time. As there will be numerous flight test items if to have numerous permutation and combinations of the states above, the flight test director and flight test engineer of the flight control system shall coordinate with each other to select typical actual state and combine with the numerical simulation and calculation methods to expand the range of flight

Flight test of the flight control system 619

Figure 8.2 Example of speed altitude state selection.

envelope and minimize thee flights for flight test as far as possible. For example, if it is confirmed the elevator control performance is directly related to dynamic pressure through the aerodynamic data analysis, the gain of the direct overload control channel of the flight control system is directly related to dynamic pressure and the altitude and speed of the flight test items such as pitch control performance and short-period response can be selected on the speed-altitude envelope, as shown in Fig. 8.2. A total of five combination points of speed and altitude are selected. The dark point is close to the state frequently used by the aircraft, while the light point is close to the use boundary of the aircraft. The calculated gain of the light point is compared with that of dark point and the comparison result is taken as the basis to determine to add states or not. The five points can usually represent the control stability characteristics of the pitch control performance and short-period motion.

8.3.5 Examples of flight test items The angle of attack protection function and longitudinal short-period response flight tests of a medium-sized transport aircraft are taken as examples to introduce the flight test of the flight control system. 8.3.5.1 Angle of attack protection function flight test The general development requirements of a type of medium transport aircraft and the aircraft conceptual design requirements stipulate that the flight control system shall have the angle of attack protection function. The design specification of flight control system

620 Chapter 8 stipulates that the flight control system must realize the angle of attack protection function and warning function. The design document of the control law describes the design principle, implementation method, and design verification result of the angle of attack protection function. The aircraft design and type determination document stipulates that the angle of attack protection function of the flight control system shall be verified by flight test. According to the information above, the test items for the angle of attack protection function flight test should be prepared in the flight test outline. The aircraft design information collected for this purpose are as follows: 1. Range of design weight of aircraft: (60 100) t. 2. Range of design center of gravity of aircraft: (28 36) %MAC. 3. Confirmation by weight discipline: the maximum weight of aircraft corresponds to rear center of gravity and the minimum weight corresponds to front center of gravity, basically in linear relationship. 4. Minimum stalling speed of aircraft: Vs (related to aircraft weight G and configuration FS, obtained through lookup of table). 5. Minimum service speed of aircraft: 1.2 Vs. 6. Stalling angle of attack of aircraft: as (related to configuration of lift augmentation device of aircraft). 7. Confirmation by aerodynamic performance discipline: the effect of landing gear state to the stalling angle of attack is not large and it can be neglected. 8. Three types of flaps (slats) configurations: FS1, FS2, FS3. 9. Control law design state: under (As-2) , start the angle of attack protection function and the gain will increase with the target difference, ensure the angle of attack will not exceed As under level flight maximum hike, it will exceed As in transient dive hike, and the peak value shall not exceed 0.5 . Based on the information above, the angle of attack protection function flight test should be carried out under nine flight states (see Table 8.5). Table 8.6 shows the flight test task list under a flight state.

Table 8.5: Flight states for angle of attack protection function flight test. No. 1 2 3 4 5

Configuration

Weight-center of gravity

No.

Configuration

Weight-center of gravity

FS1 FS1 FS1 FS2 FS2

65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC

6 7 8 9

FS2 FS3 FS3 FS3

95-35MAC 65-29MAC 80-32MAC 95-35MAC

Flight test of the flight control system 621 Table 8.6: Task list of flight test angle of attack protection function. Item No. Flight stage Aircraft weight

2i6a CLB Aircraft configuration 80 t Center of gravity position Dry cement runway Digital fly-by-wire

FS2 32% MAC

Priority Landing gear state Flight altitude

A DN

Throttle TFLF state 2000 m Flight 1.2 speed Vs 1 50 No obvious slope Wind speed lower than 5 km/h

Runway state Slope Flight control Weather state conditions Data file name 2i6a-001 Test requirements: 1. At a predetermined altitude, make the aircraft have straight-line and horizontal flight at a constant speed (1.2 Vs 1 50) and hold for 5 s; 2. Slowly and continuously increase the angle of attack of aircraft and when it approaches to 1.2 Vs, maintain the deceleration rate at about 1 2 km/h; 3. Observe whether the angle of attack protection function of the flight control system works normally; 4. If the angle of attack protection function is enabled, wait it to work continuously for 3 s and then stop the test; 5. If the angle of attack protection function is not enabled and the speed is lower than 1.1 Vs, correct it manually and then stop the test. Precautions: 1. Complete the aircraft’s atmospheric data system before this flight test. 2. Perform an aircraft stalling speed flight test and update the data before this flight test. 3. While increasing the angle of attack, increase the throttle to maintain a horizontal flight with a stable rate of deceleration. 4. If the angle of attack protection function is not enabled, it shall be restored to the preset state and repeated twice. Data supplement requirements: If possible, the flight test shall be provided with external equipment to measure the aircraft’s angle of attack and speed.

8.3.5.2 Longitudinal short-period response flight test The general aircraft development requirements and aircraft conceptual design requirements stipulate that the aircraft shall meet the first level flying quality requirements within the envelope. The aircraft flying quality requirements of the aircraft regulate that the longitudinal short-period response of the aircraft shall meet the first level flying quality requirements. The design document of the control law describes the design principles, implementation methods, and design evaluation results of the control augmentation function. The aircraft design and type determination document stipulates that the longitudinal short-period response of the aircraft shall meet the first level flying quality requirements and it shall be verified by flight test. According to the information above, the test items for the longitudinal short-period response flight test should be prepared in the flight test outline. The aircraft design information collected for this purpose are as follows:

622 Chapter 8 1. Range of design weight of aircraft: (60 100) t. 2. Range of design center of gravity of aircraft: (28 36) %MAC. 3. Confirmation by weight discipline: the maximum weight of aircraft corresponds to rear center of gravity and the minimum weight corresponds to front center of gravity, basically in linear relationship. 4. Minimum service speed of aircraft: 1.2 Vs (obtained through lookup of table). 5. Range of flight speed of aircraft: (1.2 Vs 520/0.68), (FS1), transition altitude 7500 m. 6. Range of flight speed of aircraft: (180 210), (FS2). 7. Range of flight speed of aircraft: (170 200), (FS3). 8. Three types of configurations of lift augmentation device: FS1, FS2, FS3. 9. Control law design state: direct channel, angular rate, and overload feedback are adopted to realize the control augmentation function. Through evaluation by calculation and evaluation by simulator, the longitudinal short-period response of aircraft shall meet first level flying quality requirements within the service envelope. Based on the information above, there are 54 flight states that should be put into the flight test, as shown in Table 8.7. Table 8.8 shows the flight test task list under a flight state. Table 8.7: Flight states for longitudinal short-period response flight test. Configuration FS1 FS1 FS1 FS1 FS1 FS1 FS1 FS2 FS2

Weight-center of gravity

Speed/km

Altitude/m

65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC

220

2000、4000

260

2000、4000

300

2000、5000

350

2000、5000、8000

400

3000、5000、8000

450

5000、8000

500

5000、11,000

Vref 1 10

2000

Vref 1 10

2000

(Continued)

Flight test of the flight control system 623 Table 8.7: (Continued) Configuration FS3 FS3

Weight-center of gravity

Speed/km

Altitude/m

65-29MAC 80-32MAC 95-35MAC 65-29MAC 80-32MAC 95-35MAC

Vref 1 10

2000

Vref 1 10

2000

Table 8.8: Task list of longitudinal short-period response flight test. Item No. Flight stage

2i6a CLB Aircraft configuration

Aircraft weight

95 t Center of gravity position Dry cement runway Digital fly-by-wire

FS1 35% MAC

Priority A Landing gear state UP

TFLF

Flight altitude

220

Throttle state 200 Flight speed

Runway state Slope No obvious slope Flight control Weather Wind speed lower than state conditions 9 km/h Data file name 2i6a-001 Test requirements: 1. At a specified altitude, make the aircraft have straight-line and horizontal flight at a specified speed and hold for 10 s; 2. Input constant-amplitude step of pull, push, pull and push to the control column and hold for 3, 2, 1, 1 s respectively; 3. Release the control column, the aircraft has free response for 10 s and then use the steering wheel to correct the slope; 4. Make the aircraft trim again and keep in level flight and then stop the test. Precautions: 1. Ensure the aircraft is in trim state before operating. 2. The action amplitude shall be increased gradually until typical response is generated. 3. Ensure the flight environment is a quiet atmosphere. 4. If the flight speed cannot reach requirements, the most proximate speed can be used. Data supplement requirements: None

8.4 Test system of the flight test of the flight control system The test system of the flight test plays a crucial role in the research and identification of flight control system, which is mainly reflected in the following aspects: 1. It is conducive to the safety of flight test. As it processes and monitors the ground data in real time, acknowledges the state of aircraft and system in a timely manner, supervises the control and operation of pilots, finds out problems and inform the pilots in a timely manner, and supports the flight test of pilots with the strength of a technical team, it is of great significance to ensure the safety of the flight test.

624 Chapter 8 2. It is conducive to improving the quality of flight test. The general aircraft development requirements have specific technical indexes and aircraft design has relevant design standards and specifications. The conformity of these technical indexes and standards and specifications shall be validated through the flight test. On one hand, the test system of the flight test is required to have good performance and accuracy. On the other hand, the aircraft flight profile and pilots’ maneuver actions are required to comply with the flight test task list. No matter whether in real time or after the test, through data processing, the pilot’s control can be checked to see whether it meets requirements and the technical state and configuration of aircraft can be checked to see whether they are correct. If they do not meet the requirements, the flight test shall be conducted again to improve the quality of flight test. 3. It is conducive to quickening the progress of flight test. The effective use of the test system of flight test can make the flight test process proceed according to the predetermined outline and plan. Through real-time monitoring, the effectiveness of the flight test can be validated to avoid an ineffective flight test and unnecessary repeat of the flight test. Meanwhile, as there are many test items in the flight test, a perfect test system of flight is of great significance for improving flight test efficiency and accelerating flight test progress. The flight control system mainly relies on the test systems of flight test installed on the aircraft to obtain flight test data, such as the data acquisition system, data recording system, and telemetry system. The data acquisition system acquires aircraft parameters and flight control system parameters through sensors, special interfaces, or data buses. After format conversion, the data are stored in the recording system and synchronously transmitted to the data receiving station. The requirements of flight control system for test system of the flight test are embodied in simultaneity, real-time performance, and data storage. 1. Simultaneity The simultaneity requirement is mainly targeted to the synchronization of system clock. In other words, the data acquisition equipment such as the sensor, data bus, and special test equipment used for the test system shall use consistent clock and the difference shall not exceed three sampling steps generally, and the clocks for different flights shall be consistent. The simultaneity requirement is very important for getting correct conclusions from the data analysis after the test. As shown in Fig. 8.3, an aircraft has wheel being loaded and the pilot makes a “weird” action to withdraw the landing gear due to nonsynchronous data clocks. 2. Real-time performance The real-time performance requirement is mainly targeted to the general time delay requirement that “the telemetry system transmits data to the ground monitoring station,

Flight test of the flight control system 625

Figure 8.3 Nonsynchronous data clocks for test system of flight test of a type of aircraft.

processes and analyzes the data in the computer, and displays the results on the monitoring screen in a variety of easy-to-understand forms.” Generally, the total delay shall not exceed 100 ms. Test flight engineers and aircraft designers can decide to continue the test and repeat the test items or not based on these data. 3. Data storage The data storage requirement is proposed to ensure the consistency of the quantity, accuracy, sampling rate, and dimension of the flight test data and it shall be followed by all units participating in the test. Table 8.9 shows the data storage requirements for the test system of flight test of a type of aircraft.

8.5 Data acquisition, processing, and analysis of the flight test of the flight control system Data analysis of the flight test is an objective basis for the flight control system designer to judge whether the actual state of the aircraft is consistent with the design state and is also an effective approach to find out the hidden design dangers of the flight control system and

626 Chapter 8 Table 8.9: Data storage requirements for test system of flight test of a type of aircraft. Discipline Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control Flight control

Variable name

Variable symbol

Dimension accuracy

Sampling

Rate (Hz)

Left/right control column displacement voting value Left/right steering wheel displacement voting value Left pedal brake displacement

X_Pitch

mm

0.10 mm

25

X_Roll

( )

0.10

25

X_LeftBrake

mm

0.10 mm

10

Front wheel turning handle position

X_Tiller

( )

0.10

10

Flaps and slats control handle position

X_FlapHandle

( )

0.10

10

Air brake control handle position

X_SpHandle

( )

0.10

10

Left/right steering wheel force

F_Roll

N

0.5 N

10

Left aileron position

( )

0.10

25

Right aileron position

De_LeftAileron 0.10 De_RightAileron

( )

0.10

25

Flaps position

De_Flap

( )

0.5

10

Wheel 2 brake pressure signal

BrakePressure02

MPa

0.1 MPa

10

mm

0.1 mm

25

Right front wheel column compression X_Strut_RightFwd Right outer engine combustion adjustment valve angle Left outer throttle lever angle

X_Throttle04

( )

0.10

25

X_PowerLever01

( )

0.10

25

Roll angle

Angle_Roll

( )

0.05

25

True heading

Angle_Heading

( )

0.05

25

Indicated airspeed

Vias

km/h

0.1 km/h

25

formulate solutions. The quality and integrity of flight test data acquired are the key factors affecting the work above. In 1 year before the start of the flight test, the flight control system shall prepare a flight test data demand report to clearly put forward the flight test data demand to the flight test unit, so that the flight test unit can add/refit the test. The acquisition of flight test data is completed by the test system of flight test. The chief engineer unit shall in a timely manner request the flight test data from the flight test unit for quick processing and can require conducting the flight test again for the state with problems.

Flight test of the flight control system 627 The processing and analysis of flight test data shall be completed by flight control system designer. It mainly covers the following aspects: 1. Analysis of working state Extract the information about the working state for periodic functions of the flight control system, and analyze whether the working conditions and states are consistent with the design state, such as BIT function, work cycle, working state, redundancy voting, control channel selection, command calculation, and sending. 2. Analysis of work logic According to the requirements of the flight test task list and the pilot’s actual operation, analyze whether the system’s working state conversion during pilot operation is consistent with the design, and whether the time and transient state of the conversion meet the design requirements. 3. Contrastive analysis of control law logic Extract the start, switch, and exit time of logics related to the control law, such as air/ground conversion, configuration conversion, attitude hold, flight level change, and boundary protection, and judge whether they are consistent with the design state. 4. Performance analysis of actuator system Extract control plane control command and control plane deflection angle information, analyze the maximum initial deflection rate and maximum deflection angle of the actuator, correct the power curve of the actuator, and confirm the difference with the design state. 5. Time sequence analysis of actuator system Extract the actuator and control plane deflection angle information with sequential logic requirements, draw their time sequence curve during actuation, and judge whether they are consistent with the design state. 6. Analysis of control plane deflection angle response Extract typical aircraft state and main control plane deflection angle data, have contrastive analysis with the design state and take the results as basis for adjusting the control law parameters. 7. Analysis of sensor Extract the sensor information under different configurations, analyze the influence of sensor location, aircraft deformation, vibration noise, and electromagnetic noise, and optimize the sensor feedback filtering. 8. Statistics of working state According to the pilot comment record form and flight test data, make statistics of the working time and probability under normal, degraded and simulated backup working modes of the flight control system, make statistics of the situation and authenticity of warning information at different levels of the flight control system and correct the judgment logic of warning information.

628 Chapter 8 9. Statistics of working state of actuator Make statistics of the time when the actuators are under different working modes and working rates and improve the power demand database of actuator. 10. Statistics of working state of sensor Make statistics of the work stability and installation position of sensors as well as the aircraft deformation and external noise. As the flight test data volume of the flight control system is very huge (each flight can reach above 500 MB), the data processing and analysis can only be completed smoothly with the help of powerful processing software and a small part of human manpower. Complete data processing software shall at least include data preprocessing software, special data analysis software, flying quality analysis software, and aerodynamic parameter identification software. 1. The main functions of the data preprocessing software are to convert the flight test data into understandable physical quantities that can meet the demand of flight test data (e.g., to improve the accuracy of some of the data, the test system of the flight test will use data bits of two double-precision types to represent the parameter), and decompose the flight test data to data segments that can be quickly processed according to the flight test actions or flight stages. 2. Special data analysis software refers to the data processing software specially compiled according to the actual design state of the flight control system and the items that need to be analyzed. This software mainly analyzes the redundancy voting and management, logic, and working state at the system level. 3. Flying quality analysis software can use general flying quality analysis software to calculate and analyze the items specified in the aircraft flying quality specification. 4. Aerodynamic parameter identification software is mainly used to calculate and extract the control performance data of the aircraft’s main control planes and correct the gain increment brought by the use of wind tunnel test data. In actual aircraft design, wind tunnel test data are used for the design of control law parameters first. Although the robustness of the control law is tested with the aerodynamic data perturbation method, the wind tunnel test data shall still be corrected with the aerodynamic parameter identification method to make the aircraft achieve better controllability, stability, and comfort.

8.6 Organization and implementation of the flight test of the flight control system The direct leader of the flight test of the flight control system is the chief designer of the flight control system. All personnel on the test aircraft shall participate in the flight test under its leadership. The personnel on the test aircraft include the flight test director engineer, flight control system designer, control law designer, control system designer, and actuator system designer.

Flight test of the flight control system 629 All personnel on the test aircraft shall fully digest the technical design data of the flight control system and other related systems. Through participating in the engineering simulator test and the “iron bird” integration test of the flight control system, the personnel shall be familiar with the characteristics and control stability characteristics of the flight control system. Through the training conducted by the flight test unit, they shall be familiar with the working procedures of the flight test monitoring hall and evaluation hall. Before the test, the chief engineer of the flight control system shall take the lead to specify the technical state of the flight control system and all personnel on the test aircraft shall fully understand the purpose of the flight test, flight test outline, and flight test task list. Personnel on the test aircraft in the evaluation hall are responsible for the technical state disclosure to pilots and the evaluation records after the flight test. The personnel on the test aircraft in the monitoring hall are responsible for monitoring the working state of the flight control system during the flight test, judging the impact of the faults of the flight control system, and assisting the monitoring commander to make test decisions. After the flight test, the personnel on the test aircraft in the monitoring hall are responsible for extracting flight test data from the flight test unit, which will be distributed to various disciplines for data processing and analysis with the comment record forms. The major faults, major technical problems, or major design changes of the flight control system during the flight test shall be confirmed by the flight test data. Relevant decisions can only be made after technical personnel and pilots of the development unit and flight test unit have had a thorough discussion and conducted the necessary ground test. Verification by flight test shall also be carried out if necessary. The flight test of the flight control system is not only necessary for the type determination and identification of aircraft, but also necessary for the development of the advanced flight control system design technologies. To ensure the safety and efficiency of verification by flight test, test planning shall be started at the initial stage of aircraft design and airworthiness standards for civil aircraft shall be referenced for the establishment of sound requirements for the aircraft type determination flight test. In 1 2 years before the maiden flight, full communication shall be conducted with the flight test unit to improve the design and testing of the monitoring system and data acquisition system. During the flight test, work closely with the flight test commander to optimize the system monitoring criteria. After the flight test, the test data shall be extracted and analyzed in time. As the workload of flight test data analysis and statistics is huge, a complete institutional framework shall be established and sufficient technical personnel shall be configured. It shall also be ensured that the flight test data analysis and statistics are always completed to a high standard.

630 Chapter 8

Further reading [1] AVIC No. 640 Institute, ARJ21 Aircraft Conformity Certification Flight Test Outline, No.640 Institute, Shanghai, 2003. [2] AVIC No. 1 Aircraft Design and Research Institute, A Test Machine to Verify Key Technologies for Large Aircraft Development, AVIC No.1 Aircraft Design and Research Institute, Xi’an, 2008. [3] H. Beh, G. Hofinger, E. Huber, Control law design and flight test results of the experimental aircraft X-31A, in: M.B. Tischler (Ed.), Advances in Aircraft Flight Control, Taylor and Francis, Oxon, 1996. [4] Z. Defa, Y. Shengli, et al., Ground and Flight Test of Flight Control System, National Defense Industry Press, Beijing, 2003. [5] Z. Liang, et al., Study on configuration evaluation methods for system flight tests of civil aircraft, Aeronaut. Sci. Technol. (5)(2012) 49 51. [6] X. Lixue, Engineering simulator is an indispensable tool in modern aircraft design, Air China (7)(1995) 41 43. [7] Norton W.J. Balancing modeling and simulation with flight test in military aircraft development. AGARD Conference Proceedings, CP-593, December 1997. [8] Z. Tiesheng, et al., Aircraft Flight Test Workbook—Aircraft Flight Test and Data Acquisition, National Defense Industry Press, Beijing, 1998. [9] G. Weihao, et al., Technical Manual for Flight Test of Transport Aircraft, Bureau of Civil Aircraft of Ministry of Aviation Industry, Beijing, 1988. [10] Z. Yongjie, et al., Current Situation and Development of Integrated In-Flight Simulation Test Aircraft (IFSTA), Flight Test Research Institute of China, Xi’an, 2010. [11] Z. Ziquan, Flight Test Engineering, Aviation Industry Press, Beijing, 2010.

CHAPTER 9

Airworthiness verification test of the flight control system 9.1 Overview According to the requirements of Regulations for Conformity Certification of Civil Products and Spare Airborne Equipment (CCAR21), the applicant shall demonstrate and prove that the design of products conforms to applicable airworthiness requirements through conformity certification test and demonstrate such conformity to the airworthiness department during the type conformity certification period.1 Chapters 2 and 4 8 of this book, respectively, introduce the airborne equipment test of the flight control system, control law and flying quality evaluation of the flight control system, combined test of the flight control subsystem, “iron bird” integration test of the flight control system, onboard ground test of the flight control system, and flight test of the flight control system. Chapter 3 introduces the verification and validation of the flight control system airborne software. The test content and test methods introduced are basically based on the development requirements of domestic military aircraft and are rarely involved with airworthiness verification. Although the development of domestic military aircraft requires most of the same test content and requirements as the airworthiness verification test, some special requirements of the airworthiness verification test are not involved, such as the management requirements of the airworthiness authority and the description of relevant test items showing the conformity of airworthiness clauses. This chapter mainly describes the management requirements for the airworthiness verification test of the flight control system and the test items and test methods related to the airworthiness clauses as the supplement to other chapters. Please refer to other chapters of this book for detailed test content and methods. The airworthiness verification test of the flight control system is one of the main conformity certification methods to judge whether the design of the system satisfies the requirements and is also an important means to guarantee the safety of aircraft. It takes up a large part in the whole development work and goes throughout the whole development process of the flight control system. To ensure the test can accurately reflect the inherent 1

CCAR-21-R3. Regulations on Conformity Certification of Civil Aviation Products and Spare Airborne Equipment. Beijing: Civil Aviation Administration of China, 2007.

Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00009-7 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

631

632 Chapter 9 characteristics of the tested system, obtain exact test parameters and results, and provide a reliable basis for conformity certification, the verification test shall be managed according to the procedures stipulated by the airworthiness authority. The test piece shall accurately reflect the requirements of the test and it shall be generally the airborne equipment of the aircraft. Before the aircraft obtains the type certificate or before the type is determined, all products of the flight control system are called test pieces, including the products participating in the test bed test, installed products of aircraft, and products for the verification test. The airworthiness verification test is generally divided into the engineering verification test and flight test (MC6). The engineering verification test includes the laboratory test (MC4), onboard ground test (MC5), simulator test (MC8), and equipment qualification test (MC9). The classification of the airworthiness verification test of the flight control system is same as the classification of the general airworthiness verification test, as shown in Table 9.1 in detail. The test items of the flight control system included into the airworthiness verification test shall be negotiated with the airworthiness examination group according to the specific conditions of the aircraft systems and conformity verification requirements of airworthiness clauses. Section 9.2 covers the certification requirements for verification test of flight control system and introduces the general process, procedures, and requirements of the examination conducted by the airworthiness authority to the verification test. They are the same as the general requirements for the airworthiness verification test and are described mainly from the perspective of airworthiness management. Generally, the certification requirements cover the formulation of the Certification Compliance Plan of flight control system, determination of the conformity certification methods of applicable clauses, determination of corresponding test items, preparation of test plan, preparation of test procedure and submission of it for review, manufacturing of test pieces and test facilities, implementation of conformity inspection and witnessing test, preparation of the test report, and its submission for review. From the perspective of the examination conducted by Table 9.1: Classification of airworthiness verification test of flight control system. Classification

Code

Engineering MC4 verification test MC5 MC8 MC9 Test flight

MC6

Name Laboratory test

Description

Generally the integrated test bed test of subsystems and “iron bird” integration test of flight control system Onboard ground Generally the flight control system test test conducted for aircraft on the ground Simulator test Generally the flight control system test conducted on engineering simulator Qualification Generally the airborne equipment test test Flight test Generally the flight control system that must pass the aircraft flight test

Test items in the book Test items introduced in chapters 5 and 6 of this book Test items introduced in chapter 7 of this book Test items introduced in chapter 4 of this book Test items introduced in chapter 2 of this book Test items introduced in chapter 8 of this book

Airworthiness verification test of the flight control system 633 airworthiness authority, the airworthiness verification test of the flight control system is the same as the airworthiness verification test of other systems and it mainly has the following features: 1. The test is conducted to verify whether the airborne equipment and system design meet the airworthiness requirements. The airworthiness requirements are the applicable requirements regulated by airworthiness clauses, special conditions or other airworthiness documents. 2. The examination representative monitors the whole process of the test. 3. The test procedure, design drawing, and documents of test pieces as well as the test report shall be reviewed and approved by the examination representative. 4. The examination representative will conduct conformity inspection for the test pieces, installation of test pieces, test devices, and environment. 5. The examination representative may witness the test at the scene and have the right to discontinue the test if a major problem is found or occurs.

9.2 Airworthiness verification test certification requirements 9.2.1 Engineering verification test certification requirements 9.2.1.1 Preparation of test plan Generally, the preparation of the test plan starts from the detailed design stage of the flight control system after the airworthiness verification test of the flight control system is determined. After the test demand is put forward to the examination group, the test plan shall be submitted to the airworthiness authority, so that a schedule that can be accepted by both parties can be reached after negotiation and discussion with the examination representative. During the test, if the timeline changes for any reason, it shall be informed to the other party in advance so that corresponding adjustment and arrangement can be made in time. Table 9.2 shows the main content of the schedule of the airworthiness verification test of flight control system. 9.2.1.2 Preparation of test procedure According to the development progress of the flight control system, the preparation of the test procedure of the airworthiness verification test starts in the detailed design stage. According to the requirements of the airworthiness authority, the test procedure shall be submitted long before the airworthiness verification test of flight control system for review and approval by the examination group. According to the characteristics of the test items, the test task list can be prepared first to put forward specific test requirements to the unit undertaking the test. The preparation of the test task list is not a mandatory requirement of the airworthiness authority and the examination representative will not approve the test task list and will only approve

634 Chapter 9 Table 9.2: Schedule of airworthiness verification test of flight control system. No. 1

2

3 4 5

6 7 8 9 10

Test work Prepare test procedure according to airworthiness requirements and system design requirements, and submit it to examination representative for discussion Complete the drawing design of test pieces, let the examination representative to review it, and determine the conformity inspection items Submit the test procedure and let the examination representative to review it Test piece production and procedure conformity inspection Complete the test pieces manufacturing, conduct conformity inspection of test pieces, and obtain the airworthiness approved label Conformity inspection before test Conduct test and the examination representative witnesses the test Complete the test and sort out test data Complete the test report Submit the test report of the applicant, and the examination representative reviews, approves and verifies the report

Completion

Time

Remarks

Test procedure

3 . 3 The test task list can also be prepared first according to demands

Drawings and documents

3.3

33

Process

3.3

33

Process

3.3

33

Statement of Conformity

3.3

33

Statement of Conformity, conformity inspection record form Conformity inspection Record Test observation report

3.3

33

3.3

33

3.3

33

Test report Test report (or test analysis report)

3.3 3.3

33 33

the test procedure. The applicant can also discuss the test task list with the examination representative in advance so that both parties can reach a consensus on the test items and content as early as possible. The outline of the airworthiness verification test of flight control system shall generally include but not be limited to the content in Table 9.3. 9.2.1.3 Conformity inspection 1. Conformity inspection of test pieces When the test pieces are manufactured, in addition to the inspection in the processing process, each test piece shall be submitted to the examination representative for examination. If the applicant promises the test piece is ready and can be submitted

Airworthiness verification test of the flight control system 635 Table 9.3: Main content of outline of airworthiness verification test. No. 1 2 3 4 5 6 7 8 9 10 11 12

Items

Main content

Brief description of scope Reference documents

Main content and scope of application of test procedures Documents, standards, specifications, or drawings referred to by test procedures Test objectives It shall include airworthiness clauses for verification Description of test products Configuration of test products, installation of test products on test devices, and relevant drawing number Test equipment List of all test equipment used in the test, description on calibration and approval and description on testing equipment and their accuracy Conformity inspection Description of test products and requirements on conformity of test products Test description Description on expected conformity of verification clauses Test steps Detailed description of specific operation procedures of the test Judging criteria of test Judging criteria showing that test results comply with requirements Test record Requirements on items and data recording Handling of abnormal Description of handling procedures for special circumstances circumstances Special requirements for Description of qualification of test personnel test personnel

for examination, it shall provide a conformity statement to the examination representative. The conformity statement is generally signed by the applicant’s legal representative or its authorized person. The applicant shall provide all the materials involved in the manufacturing of the test piece, including copies of the data of materials entering the factory, process and process sheet, production records, inspection records, deviation processing documents, and conformity certificates. For the conformity certification test of the flight control system conducted by test units not manufacturing the test facilities of the applicant, the examination representative shall issue authorized release the certificate/airworthiness approval tag for test pieces passing the conformity inspection. What should be noted is that test pieces not obtaining the authorized release certificate/airworthiness approval tag shall not be put to the conformity certification test. 2. Conformity inspection before verification test of flight control system After the preparation for the test is completed, the applicant shall organize an internal inspection before the test and make timely changes to any problems found in the inspection. After passing the internal inspection conducted before the test, the applicant shall submit a conformity statement to state that the test pieces have obtained the authorized release certificate/airworthiness approval tag and the test facilities, test personnel, test fixtures, test procedures, installation of test pieces, and test environment comply with airworthiness requirements. The conformity inspection conducted by the examination representative before the test focuses on the maintenance and effectiveness of the test environment and test facilities, the accuracy and period of validity of

636 Chapter 9 measuring equipment, the installation correctness of test pieces, the qualification certificate of personnel participating in the test, the applicability and rationality of the data acquisition system, the understanding and execution of test procedures by the personnel participating in the test, and the effectiveness of items set for the test record, with the purpose of ensuring that the test products and test devices comply with engineering drawings and the test outline. The inspection items are shown in Table 9.4. The inspection results are recorded in the conformity inspection record form. What should be noted is that unless agreed by the examination group, the test products and test devices shall not be changed during the period from the submission of the conformity statement showing their conformity to type data to their submission for the verification test. In case of any changes, they shall be resubmitted to the examination representative for approval and the conformity inspection shall be conducted again if necessary. 9.2.1.4 Witness test The examination group usually selects some airworthiness verification test items as witness test items. According to the general practice of the airworthiness authority, the selection of witness test items is coordinated between the applicant and the examination group to determine some important test items as the witness test items. The determination of witness test items mainly depends on the evaluation of the system and test condition by the examination group. Some test items of the airworthiness verification test of the flight control system can be selected as the witness test items of the airworthiness authority. For example, the fault simulation test and frequency response test under the “iron bird” integration test of the flight control system and the modal conversion function inspection, system state and alarm display test, and structural mode coupling test under its onboard ground test can be considered as the witness test items. After the witness test items are determined, the applicant should notify the examination representative to arrive at the site of the witness test in advance. During the witness test, the examination representative shall check whether the test follows the test steps specified in the approved test outline and whether the data collected with the test instruments in the test are effective for the test and carry out posttest inspection. If disassembly inspection should be conducted for test pieces, the examination representative shall witness the disassembly inspection and the applicant shall not disassemble and inspect these test pieces before the examination representative arrives at the site of the test. In the process of the witness test, the examination representative will send a list of recorded problems found in the test to the applicant and the leader of the examination group. In the case of any major problems during the test or found by the examination representative, the examination representative has the right to suspend the test. Once the examination representative decides to suspend the test, the applicant must immediately find out the

Airworthiness verification test of the flight control system 637 Table 9.4: Conformity inspection items. No. 1 2 3

4

5

6 7 8

9

Inspection items

Inspection requirements

Test procedure

Current effective version of test procedure shall be approved by the airworthiness authority and placed on the test site Conformity certificate and Catalog of test parts shall be provided, including conformity certificate airworthiness approval tag and airworthiness tag of test pieces. The catalog and certificates shall of tested objects be put on the test site Airborne software Software configuration description list of tested objects shall be provided, including the name of the tested object, part number, software version, software level, and whether there is an upgrade and upgrade record TSOA (Technical Standard If there are spare airborne equipment that have obtained the TSOA Order Authorization) parts among the test pieces, TSOA parts list including the name, part number, TSOA certificate state and airworthiness tag shall be provided The configuration report provided shall accurately describe the test Test configuration evaluation and airworthiness configuration, generally including the list and state of the tested objects and the installation state of the tested objects approval The configuration report shall provide a formal list of the equipment in the test, including the equipment name, equipment model, parameters, and the number and validity period of the calibration certificate. All equipment in the test shall be within the validity period and the list of test equipment and calibration certificate shall be put on the test site together with the test equipment. Type data approval form The test procedure, configuration evaluation report, and approval forms of other airworthiness documents shall be placed on site Qualification of personnel All operators in the test shall be trained and have appropriate participating in the test qualification certificates and these certificates shall be on site for inspection On-site Preparation 1. All operators shall be ready for the operations and all personnel participating in the test shall have a clear division of labor to ensure the smooth progress of the test. 2. The objects shall be installed in place and the installation shall meet the requirements of test procedure. 3. The test equipment shall be in place and the working state shall meet the requirements of test procedure. 4. The operation rules that can be used to guide field operation shall be provided, the operation steps in test procedure shall be combined with the test data record forms, the operation steps and data recording nodes shall be clear. Others Special requirements and technical conditions for some tests shall be prepared in advance

causes and take measures for rectification. After the causes are removed, the applicant shall submit a report on the resumption of the test to the examination group. The test could only be resumed after being approved by the examination group. If the examination representative responsible for the test item cannot witness the test, the conformity inspection representative will be entrusted to the witness test generally. After the test, the

638 Chapter 9 examination representatives who witnessed the test at the site will prepare a test report within 10 working days generally to outline the test results, the problems found, and the treatment measures taken by the applicant. 9.2.1.5 Preparation of test report After the completion of the airworthiness verification test of the flight control system, the applicant shall prepare a verification test report, which shall be a conformity report and approved by the examination representative. The verification test report is used to faithfully record the test process and data, explain whether the test follows the steps in test outline, analyze the test results, and make a judgment on whether the test meets the requirements of airworthiness clauses. The applicant may also divide the verification test report into two reports according to specific conditions of the test, namely a test report and a test analysis report. The airworthiness verification test items of the flight control system shall generally include but not be limited to the content shown in Table 9.5. Only items that must be included in the verification test report are specified here and the specific content of the verification test report of the flight control system shall be prepared according to the actual conditions of the test items. If the airworthiness verification test is carried out at a unit who is not the manufacturer of the test facilities of the applicant, the test report shall be prepared by the unit undertaking the test and recompiled by the applicant. In the conclusion part, it shall be specifically described Table 9.5: Content of airworthiness verification test report. No.

Name

1

Scope

2

Reference documents

3 4

Test objectives Description of test products

5

Test equipment

6

Test procedures

7

Test data and materials

8

Results of disassembly inspection after the test Relevant test analysis report Conclusions

9 10

Main content The main content and scope of application of the test report shall be briefly described The documents, standards, specifications, or drawings referred to in the test report shall be listed The airworthiness clauses used for verification shall be specified The configuration and deviations of test products, conformity inspection as well as the evaluation on effects of configuration and deviations of test products shall be described Complete description attached with photos or the report of same equipment used before (if necessary), the installation way of test products on test equipment, the instruments and their correction status shall be included Test name, test steps and records, the times and causes of delay of tests shall be described The results, curves, and charts after test data sorting as well as data sorting methods and correction methods shall be generally included Important dimensional changes, NDT results, photos of failures, and analysis shall be included Such as the hydraulic oil contamination testing report Specific conclusions on the conformity shall be provided

Airworthiness verification test of the flight control system 639 whether the verification test conforms to the requirements of the approved test outline, whether the test meets the requirements of airworthiness clauses or other airworthiness requirements, and whether the test results reach the intended purpose of the test.

9.2.2 Flight verification test certification requirements 9.2.2.1 Applicant flight test The examination group shall conduct the research and development flight test and inspection before the issuance of the type inspection authorization (TIA). The purpose of the test and inspection conducted by the applicant is to show that the products submitted for ground and flight test to the examination party meet the minimum quality requirements and conform to the type design and they are safe for the planned flight test. The applicant shall first conduct the various flight tests regulated by the airworthiness authority to show that the products conform to airworthiness regulations. Generally, the applicant shall prepare the requirements for the test flight of the flight control system according to the airworthiness requirements and design requirements of the flight control system. The airworthiness requirements are mainly the applicable airworthiness clauses to MC6. When the requirements for the test flight are proposed, it shall be ensured that the test content can fully verify the requirements of the airworthiness clauses and prove the system can accomplish the expected functions and different system configurations, and single fault conditions can be evaluated under the worst case. The unit undertaking the flight test shall prepare the outline of the verification test flight according to the requirements of the test flight. Generally, the outline can be prepared according to the specific requirements of the type development. The specific configurations for the flight test of the flight control system and the overall condition of the test shall be described. The configuration emphasizes the equipment installation that has effects on the flight control system test or that should be considered for the flight control system test and the idea of the test, factors, and principles that should be considered in the overall description of the test shall be specified. The outline of the verification test flight shall also describe the requirements for external environmental conditions for the test flight of the flight control system, such as runway characteristics, visibility, wind speed, wind direction, day and night conditions. The specific items for the test flight can be given in tables or other forms and they shall include the altitude, speed, weight, center of gravity, flaps and slats position, landing gear state, engine power state, and tolerance of various test parameters for the test flight. In the outline of the test flight, the airworthiness clauses and requirements to be verified, content and procedures focused by test items, acceptable criteria, and test flight risk management shall be described. After the research and development flight test of the flight control system of the applicant, the unit undertaking the test shall prepare the test flight report, which shall include the description of the configuration state of the flight control system for the test, the description of the test equipment and their refitting condition, the description of test equipment and installation

640 Chapter 9 position and other information about the flight test items, the description of conformity inspection, and all data and materials about the conformity inspection, including the deviations of the conformity inspection. The flight test procedures shall be described according to the test items and the test procedures can also be shown by referring to relevant documents or in the form of an annex. Generally, information such as test points and test states, as well as the center of gravity, weight, pressure altitude, speed, and other relevant parameters are given in the form of tables. If the actual test procedures are inconsistent with regulated test procedures, an explanation shall be made. The flight test data and their processing shall be described in detail. Usually, the test points and measured data are shown in the form of tables, necessary change curves of physical parameters shall be drawn, the test data shall be analyzed, the test data processing results shall be given, and the basic conditions showing the requirements of airworthiness clauses shall be described. If the test state deviates from the outline, the influence of the deviations to the test results shall be analyzed. In the test flight report, the test pilots’ evaluation on the flight test shall be described, the flight test results shall be finally analyzed, and conclusions conforming to the test flight outline shall be given. According to the test flight report, the applicant shall also prepare a test flight analysis report to give a conclusion that the requirements of the airworthiness clauses are met through verification by test flight. Generally, the test flight analysis report shall describe the airworthiness clauses that should be verified through the test flight, the configuration of the flight control system, test equipment, and their refitting condition for the test flight, the items for and methods of the test flight, the analysis of flight test results, the specific implementation of the flight test, the time, place, and content of the test flight. It shall also contain the curve charts, characteristic parameters, necessary calculation, and comparison of test flight results that can show the conformity as well as the evaluation opinions of the pilots on the test flight. Through the analysis above, a conclusion that the requirements of airworthiness clauses are met through the verification by test flight is given finally. The outline of the verification test flight, the test flight report, and the test flight analysis report shall be provided to the examination group. The examination representative will review their acceptability to confirm whether the aircraft conforms to the type design and then determine the specific flight test items that should be reevaluated by test pilots of the airworthiness authority. These flight tests conducted by the applicant are not included into the content of the certification flight test directly, unless the examination party agrees to conduct a parallel flight test with the applicant and has issued a TIA for these tests. The Civil Aviation Administration of China calls the flight test the applicant flight test and the certification flight test is a parallel test flight. In some specific cases, in order to reduce the burden of the applicant, a parallel test flight can be considered when the examination group considers that it is appropriate and feasible.

Airworthiness verification test of the flight control system 641 The certification flight test can only be started after the TIA is issued. For the flight test data and materials of the test products that cannot represent the type design, the applicant shall understand the importance of the configuration control and the conformity record of the test products for each flight as the validity of these data and materials cannot be confirmed yet. Before conducting the research and development flight test, the applicant must obtain a special flight permit for the test aircraft. The conformity inspection representative is responsible for checking the aircraft before the issuance of the special flight permit. 9.2.2.2 Certification flight test The requirements of the airworthiness authority for the certification flight test of the flight control system are consistent with the common requirements for the certification test flight. The certification flight test is conducted to check and verify the flight test data submitted by the applicant, assess the performance and handling quality of the aircraft and the working condition of equipment, as well as determine the use restrictions, operating procedures, and information provided to pilots. The examination group can only issue the TIA after completing the examination of the applicant’s test data package and confirming its acceptability, and then the conformity inspection representative can conduct a formal ground inspection for the prototype aircraft before the certification flight test. The formal ground inspections shall be conducted before the certification flight test. The applicant shall commit that the aircraft is ready and can be provided to the examination group for examination and flight test by submitting a conformity statement. The conformity inspection representative generally completes the inspection by taking the ground inspection part of the type inspection report as the guidance, and relevant regulations (such as CCAR25) as the basic basis, and follows relevant instructions of the TIA. The conformity inspection representative will witness and check the operation and testing of all ground systems required by the TIA. After the formal ground inspection is completed, if it is confirmed as having no adverse effects on the safety and test effectiveness of the scheduled test flight, a class I special flight permit will be issued. Under this condition, the prototype aircraft is ready to fly. From then on, the applicant shall not do any work on the aircraft without the consent of the conformity inspection representative. A risk management process is required before any certification flight test (regardless of the risk level involved in the test) and a flight test plan shall be prepared before each flight. All certification flight tests shall be conducted in accordance with the issued constraints and restrictions to ensure the safety of the flight test and to determine the compliance with civil aviation regulations. The TIA can be issued by stages or increments to ensure the test aircraft has had basic airworthiness before entering the next stage and to ensure the safety of the certification flight test. The data generated in the flight test conducted by the

642 Chapter 9 applicant before the issuance of the TIA may still be valid on the premise that the following requirements are met. 1. The aircraft used in the applicant flight test is essentially the same as the aircraft used later to indicate the conformity with the model design. 2. There are no significant changes in the period from the applicant flight test to the conformity inspection.

9.2.3 Practices of airworthiness verification test of military aircraft The development of military and civil aircraft is a complex management system with two different backgrounds. The introduction of the airworthiness concept to the development of military aircraft is a major innovation and practice of the management mode of military aircraft development, which will be faced with the docking and integration of the two systems in the aspects of concept, standard, procedure, process, and examination. In the late 1980s, the US military put forward the concept of airworthiness of military aircraft and began to use the experience of airworthiness management of civil aircraft for reference in the type development of military aircraft to improve the safety level of its military aircraft. At present, the military forces of the United States, Britain, France, Germany, Italy, Spain, Netherlands, Poland, Canada, Australia, and other countries have carried out the airworthiness work of military aircraft and paid full attention to the performance and airworthiness requirements of military aircraft. China has also carried out useful exploration in the research and practice of airworthiness of military aircraft. Under the existing military aircraft development management system, the airworthiness mode of civil aircraft is not just simply copied but the development process of military aircraft is not changed but integrated with existing development process. Through the exploration and practice, a standard and systematic airworthiness workflow for military aircraft has been formed preliminarily. Airworthiness, as an inherent property of aircraft, is realized and maintained through the design, manufacture, test, use, maintenance, and management of aircraft through its whole life cycle. Airworthiness is first reflected in technical requirements, including system safety requirements and physical integrity requirements, and then reflected in management requirements, including technical state management and process control management. Airworthiness is not only an inherent property of civil aircraft, but also an inherent property of military aircraft. Airworthiness of military aircraft can be understood as the pronoun of aircraft safety in the field of aviation safety, which represents the characteristic that the aircraft can be used safely under the expected operating environment and operating restrictions. That is to say, no matter how the operating conditions are, good or bad, and even in the case of any fault or accident, the aircraft is controllable and can still return and land safely (minimum safety requirements). The airworthiness of military aircraft is

Airworthiness verification test of the flight control system 643 validated through the performance verification of military aircraft and it is included in the performance verification of military aircraft in the form of the airworthiness examination. The airworthiness of aircraft systems, subsystems, and airborne equipment can be determined by means of tests and test flights. The main purpose is to verify that the aircraft meets the operational safety level and meets the airworthiness requirements of the type within the specified service restrictions. The airworthiness verification test of the flight control system of military aircraft, taking the airworthiness experiences of civil aircraft as reference, includes a laboratory test, onboard ground test, flight test, simulator test, and unit qualification test, with basically the same classifications as the verification test of civil aircraft. The basic process is as follows: to determine the airworthiness clauses applicable to the flight control system, to determine the measures on conformity, to prepare the verification plan, to prepare the test task list, to prepare the test outline, to conduct conformity inspection and witness test, to complete the test report, and submit it for review. The technical requirements for airworthiness verification are the same and the compliance of the airworthiness clauses shall be verified as well. The difference lies in the management methods. In combination with the development process of military aircraft and the airworthiness concept, the method of accompanying examination is generally adopted. In other words, the airworthiness examination is carried out together with the quality review. The test items of the verification test of the flight control system shall be determined according to the conformity examination method and verification plan. Unified planning shall be carried out by combining the development test with the design and type determination test to avoid repetition. Only one set of test documents including relevant airworthiness verification shall be prepared. The test task list shall be prepared first to determine the requirements of clauses to be verified and the verification measures, including test name, test basis, and verified clauses, design requirements of test pieces, requirements of test steps or procedures, and requirements of test equipment. The unit undertaking the test shall prepare the verification test outline according to the test task list, with content including test purpose, test basis, test items, test steps, recorded items, list of test equipment and statement on verification approval, test equipment and their accuracy, the indication of conformity before test, and judging criteria. The conformity inspection is conducted to check the production process of test pieces, the test pieces, the test facilities, and test operators to confirm if the test pieces meet the requirements of the engineering approved design drawings, process specifications, and relevant design documents. In the 2 weeks before the start of the verification test, the quality department shall organize relevant units to conduct conformity preinspection. If any nonconformities are found in the inspection, they shall be rectified within a specified period. After passing the preinspection, the test pieces shall be submitted to the examination

644 Chapter 9 representative for formal conformity inspection. After the verification test, the test unit shall sort out the test data as soon as possible, prepare a test report, and submit it to the examination group for review and approval. Airworthiness verification test flight shall be performed together with aircraft type determination test flight. According to the confirmed airworthiness requirements and conformity examination methods, requirements for the airworthiness verification test flight of flight control system are put forward. On the basis of the requirements for design and type determination test flight, the requirements for type determination/verification test flight of flight control system are formed. The unit undertaking the test flight shall prepare a test flight outline according to the requirements for the type determination/verification test flight and submit it for review according to requirements. Upon completion of the verification test flight, a verification test flight summary report shall be prepared by the unit undertaking the test flight and then submitted for review.

9.3 Technical requirements for the airworthiness verification test 9.3.1 Laboratory test Laboratory test (MOC4) generally refers to the system simulation test, environment test, structure test, static test, and fatigue test of the aircraft. The test may be conducted on airborne equipment, subassemblies, and complete assemblies. For the flight control system, it refers to relevant tests at subsystem level, system level, and aircraft level generally. Of course, some airborne equipment level tests can be regarded as laboratory tests (MOC4). The airborne equipment level laboratory test of the flight control system have been introduced in chapter 2, the subsystem level laboratory test of the flight control system have been introduced in chapter 5, the system level laboratory test of the flight control system have been introduced in chapter 6, and the aircraft level laboratory test of the flight control system have been introduced in chapter 7. This section only summarizes the airworthiness test items in the system level laboratory test of the flight control system (namely the “iron bird” integration test of the flight control system). The main test items include functional test, fault simulation test, and load test of the system. The airworthiness clauses involved include CCAR25.671ac, CCAR25.672ac, CCAR25.683, and CCAR25.1309ad. 9.3.1.1 Functional test The functional test of the flight control system mainly covers system function inspection, system modal conversion correctness, system alarm display, and verification of airworthiness clauses CCAR25.671a, CCAR25.672ac, and CCAR25.1309a. See chapters 5 and 6 for specific test methods and process.

Airworthiness verification test of the flight control system 645 Table 9.6: Requirements for control components, motions, and effects. Control components Aileron Elevator Rudder Horizontal stabilizer Multifunctional spoiler Ground spoiler Flaps Slats

Motions and effects Deflect to the right (clockwise) to make the right wing sink Move backwards to raise the nose up Pedal forwards with right foot to make the nose deflect to the right Move downwards to trim and raise the nose up Deflect to the right (clockwise) to make the right wing sink (assisted roll), move backwards to brake the aircraft (deceleration function) Move backwards to brake the aircraft Move backwards to increase aircraft drag and lift Move backwards to increase aircraft lift

Clauses CCAR25.671a and CCAR25.1309a are general requirements for system functions. CCAR25.671a requires that each control component and control system shall operate simply, stably, and accurately to accomplish its functions. This clause is a general qualitative requirement for the flight control system. The main control components shall meet the requirements of Table 9.6 to ensure that the motion of the pilot’s hands and feet is consistent with human motion instinct. Meanwhile, the relevant requirements of deceleration control component, trim system, and modal conversion switch shall be considered. CCAR25.1309a requires that the system and equipment shall be able to complete the scheduled functions under various expected operating conditions. This clause is a general requirement for the system. CCAR25.672a requires the stability augmentation system (SAS), automatic systems, and control systems with a driving force setting a warning system for any faults that may cause unsafe results, so that the warning system can give a clear and identifiable warning to the pilots and it shall perform at the appearance of the fault that may cause unsafe results to enable the pilots to take corrective actions in time. The failure warning function can be inherent or added during the system design. The warning system shall not directly drive the control system nor be triggered by the normal operation of the aircraft. CCAR25.672c requires that the SAS, automatic systems, and control systems with driving force, can still be safely controlled after the occurrence of any single faults. The functional verification test of the flight control system mainly includes system function inspection, system modal conversion inspection, system alarm display test, system closed-loop characteristic test and man machine combined test, which are detailed as below. 1. Functional inspection a. Basic function inspection. Operate the elevator, rudder, aileron, horizontal stabilizer, multifunctional spoiler, ground spoiler, flaps and slats according to operating regulations, measure relevant parameters such as the control command, control plane position, and control plane motion rate, and confirm that the operation meets the specifications and requirements of control law. The inspection shall cover all working modes of the flight control system.

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b. Trim control inspection. Check and operate the manual pitch trim, roll trim and yaw trim, and automatic trim functions adopted by the aircraft, measure the relevant parameters such as control command, control plane position, and response time, and confirm whether the working polarity, priority, and performance of the trim meet the requirements. c. Work logic check. Check whether the on off logic, override logic, and synchronous control logic of relevant functions of flight control system (including the work logic of automatic flight control system) meet the design requirements of system and control law. d. BIT. For the flight control system with BIT function, check whether the logic conditions for the entry/exit of relevant functions of the flight control system, the equipment fault detection logic, and the correctness of handling the reported information meet the design requirements. e. Redundancy management inspection. For the flight control system adopting redundancy signals and equipment, check whether the redundancy signal voting logic, redundancy signal fault reconstruction logic, and signal switching transient state meet the design requirements. Modal conversion inspection Modal conversion inspection mainly checks whether the working mode conversion logic of the flight control system meets the requirements after the failure of the crosslinked system or bus communication. The modal conversion to transient state test mainly checks whether the transient state of conversion between the working modes of the system meets the design requirements. Alarm display test The system alarm display test mainly checks the displayed information of avionics system, records the correctness of the working state information and warning information of the flight control system, and confirms the logical conditions triggered. Closed-loop characteristic test The closed-loop characteristic test of the system checks the stability of the flight control system, obtains the frequency response of high-order flight control system (including aircraft dynamics) and then obtains short-period motion characteristic parameters of equivalent low-order aircraft with equivalent matching method, and analyzes the change course of various response quantities of the aircraft with time through the closed-loop time-domain characteristic test of the system. Man machine combined test The man machine combined test aims to examine and evaluate the control display of the flight control system, detect various functional performance indexes, evaluate the control law and flying quality, and train the pilots to handle emergency situations.

Airworthiness verification test of the flight control system 647 9.3.1.2 Fault simulation test The fault simulation test mainly examines the impact of the failure of the flight control system on the cross-linked system or the aircraft, as well as the impact of the failure of the cross-linked system on the flight control system or aircraft response, judges whether the modal conversion transient state of the system fault can meet the requirements, and verifies airworthiness clauses CCAR25.671c, CCAR25.672ac, and CCAR25.1309d. CCAR25.671c is the requirement for the fault of the flight control system. If the aircraft can still fly and land safely after a fault occurs, it can be considered as a single fault or a fault combination. CCAR25.1309d is a general clause for safety analysis and evaluation. The fault simulation test shall be generally carried out to evaluate the safety analysis results. The specific test methods and procedures are shown in chapters 5 and 6. Typical fault simulation test items are as follows. 1. Force dispute monitoring test For the primary control planes (rudder, elevator, and aileron) that adopt two actuators, the adjustment error, system gain/offset error, and dynamic error will lead to mismatching actuators on the same control plane, thus causing force dispute among actuators and thus affecting the fatigue life of the aircraft structure. The force dispute equilibrium and monitoring strategies can be adopted to the actuator control system to reduce the effect on the fatigue life of the aircraft structure. Generally, a force dispute test is required. By setting different force dispute states under different conditions, the equilibrium strategy for the primary control planes and the monitoring measures meet the requirements of technical specifications. 2. Oscillation monitoring By setting fault signals, check the warning and display of oscillation detection of rudder system, elevator system, aileron system, and spoiler system as well as system monitoring and cutoff or switch function. After the injection of fault conditions, the actuator corresponding to the fault injection control channel switches to the damping state within the specified time. The control plane displacement time curve shall be smooth in one cycle of the pilot’s operation. 3. Hydraulic source fault simulation test Under different flight states and in the process of manual operation of the aircraft, set faults of the hydraulic system and simulate the faults of one, two, and three sets of hydraulic systems respectively and their fault combinations, and analyze whether the transient state caused by the system fault and cross-linked system fault meets the design requirements. 4. Power supply fault simulation test Under different flight states and in the process of manual operation of the aircraft, set power system faults and simulate the normal power supply, emergency power supply, flight control bus bar, and their fault combination, respectively, and analyze

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whether the transient state of the aircraft caused by the system fault and cross-linked system fault meets the design requirements. Flaps/slats system fault simulation test Under different flight states and in the process of manual operation of the aircraft, set flaps/slats controller fault, flaps and slats asymmetry fault, flaps and slats tilt fault, and flaps and slats override fault, and analyze whether the transient state of the aircraft caused by the system fault and cross-linked system fault meets the design requirements. Horizontal stabilizer system fault test Under different flight states and in the process of manual operation of the aircraft, set feedback signals and operate the horizontal stabilizer to verify whether the trim system feedback signal fault monitoring and display of the horizontal stabilizer meet the design requirements. Automatic flight control system fault test Under different flight states and in the process of manual operation of the aircraft, set autopilot disconnection (normal disconnection, emergency disconnection, faulted disconnection, and override disconnection, etc.) fault, and autothrottle disconnection (normal disconnection, emergency disconnection) fault, and analyze whether the transient state of the aircraft caused by the system fault and cross-linked system fault meets the design requirements. Main equipment fault test Under different flight states and in the process of manual operation of the aircraft, set faults of primary flight control computer (PFCC), actuator controller, inertial navigation equipment, atmospheric data equipment, and engine, and analyze whether the transient state of the aircraft caused by the system fault and cross-linked system fault meets the design requirements.

9.3.1.3 Control load test Clause 25.683 of CCAR specifies the operational test of the flight control system control component after 80% limit load is applied and the flight control system part bearing dynamic load is applied with the expected maximum load under normal conditions. It aims to ensure that the flight control system operating under possible load conditions will not be jammed, undergo excessive friction, or be deformed. As for the description of “no excessive deformation” in this clause, as long as the control mechanism of the cockpit makes full-stroke motion under the action of limited load and makes the maximum deflection angle of the corresponding control plane meet the requirements of the flight characteristics of the aircraft, the system can be deemed to have no excessive deformation. To use lower (less than 80%) pilot force (for operational test), the application for the exemption of clause 25.683 is required. When the operational test is conducted, expected maximum load shall be applied to control planes such as rudder, elevator, aileron, horizontal stabilizer, flaps, slats, and

Airworthiness verification test of the flight control system 649 spoiler and 80% limit load shall be applied to control components such as steering wheel, rudder, and pedal to check and ensure the system works without jam, excessive friction, or deformation. For the fly-by-wire flight control system, the load test of the control system and the maximum load test of the control system with driving force can be carried out, respectively. The two tests can be carried out on different test beds. Meanwhile, the corresponding test bed shall consider the effects of aircraft body structure deformation and vibration load.

9.3.2 Onboard ground test The onboard ground test of the flight control system generally includes the functional and performance test of the primary flight control/automatic flight control system, flight control system and structural mode coupling test, electromagnetic compatibility test, and inspection of the cross-linking interface with the hydraulic system, landing gear system, avionics system, power system, and power supply system, etc. This part mainly describes the airworthiness verification of the tests above. For the onboard ground airworthiness verification test of the flight control system, namely MOC5 test, the range is generally lower than the content covered by the tests above and it is usually completed by the discipline of the flight control system. 9.3.2.1 Onboard functional test of flight control system The airworthiness clauses to be verified in the onboard ground test of the flight control system usually include CCAR25.655, CCAR25.671a, and CCAR25.1301a. The test task can be decomposed according to the airworthiness clauses to be verified. 1. Description of airworthiness clauses to be verified a. Clause 25.655 of CCAR This clause is the installation requirement for the movable tail surface (including elevator, rudder, and movable horizontal stabilizer). Clause 25.655a requires that when one tail surface is at the limit position and other tail surfaces move in the full angle range, there shall be no interference between any tail surfaces. Clause 25.655b requires that if an adjustable horizontal stabilizer is used, a stop dog must be used to limit its stroke to a maximum value indicating that the aircraft can meet the trim requirements in Clause 25.161. Through the onboard functional test, it shall be verified that there is no interference between movable tail surfaces within any deflection angle range and there is sufficient clearance between these movable tail surfaces. Operate the horizontal stabilizer to the maximum angle and verify that the stop dog will limit it to the specified maximum trim position.

650 Chapter 9 b. Clause 25.671a of CCAR CCAR25.671a is a general rule for control systems. It requires that each control component and control system shall realize their functions conveniently, smoothly, and accurately. In this clause, “conveniently” means the control system can be operated simply and directly and the pilot’s motion of hands and feet is consistent with human motion instinct when operating a component. “Smoothly” means that the system shall have no sudden change, feeling of stagnation, jamming, selfvibration, and shall have appropriate bar force gradient and the pilot feels comfortable. “Accurately” means that the system can correctly execute the pilot’s command and can transit from one flight state to any other flight state smoothly according to the command. Especially when one position is selected, it is unnecessary to wait for the completion of the initially selected motion before selecting another different position. In addition, the control system can reach the finally selected position smoothly and the motion of the system and the time for the selection will not cause adverse effects to the aircraft safety. This clause is verified through the onboard functional test of the flight control system. c. Clause 25.1301a(4) of CCAR The clause CCAR25.1301a(4) requires that the equipment shall have normal functions after being installed. In other words, the airborne equipment of the flight control system shall work normally on the aircraft. This clause is verified through the onboard functional test of the flight control system. 2. Airworthiness verification methods Clauses CCAR25.655, CCAR25.671a, and CCAR25.1301a(4) are verified through the onboard functional and performance test of the flight control system. The specific test items are shown below: a. When the flight control system is under normal working mode and the aileron actuator system is under different working modes, operate the steering wheel in full stroke under primary/damping and primary/primary working mode and measure and check whether the control plane control stroke, control rate, and relevant display interface of avionics system are consistent with the working state. b. When the flight control system is under normal working mode and the elevator actuator system is under different working modes, operate the control column in full stroke under primary/damping and primary/primary working mode and measure and check whether the control plane control stroke, control rate, and relevant display interface of avionics system are consistent with the working state. c. When the flight control system is under normal working mode and the rudder actuator system is under different working modes, operate the pedal in full stroke under primary/damping and primary/primary working mode and measure and check whether the control plane control stroke, control rate, and relevant display interface of avionics system are consistent with the working state.

Airworthiness verification test of the flight control system 651 d. When the flight control system is under direct link working mode and the control planes above are under different working modes, check whether the test results meet the requirements of the judging criteria. e. Work logic check, transmission ratio inspection, and control speed test of horizontal stabilizer. f. For aircraft with mechanical backup control system, measure and check whether the control stroke, control rate, and relevant display interface of avionics system are consistent with the working state. g. When the flaps/slats system is under normal working mode, operate the flaps/slats control handle, measure the control angle and rate of control plane under full speed/half speed, monitor whether the corresponding monitoring interface of avionics system is consistent with the test state, check whether the control plane response time and angle are correct, and check whether the control handle can be operated conveniently and easily. h. When the flaps/slats system is under fault mode, operate the flaps/slats emergency switch, measure the control plane control angle and rate, monitor whether the corresponding monitoring interface of avionics system is consistent with the test state, check whether the override switch can be operated conveniently and easily, and check whether the functions are correct. i. Inspection of cross-linking interface between flight control system and avionics system, power supply system, environmental control system, and power system. For the test content above, there are no special requirements for the preparation of the test environment and test implementation. As the basic test content of the onboard ground test of flight control system, they are an important and basic part of the test. Please refer to chapter 7 for specific test methods and test process. What should be emphasized here is that the airworthiness conformity verification test must follow the test process and a test verification plan shall be submitted to the airworthiness authority for approval. Test equipment and test state management shall meet the management regulations of conformity certification test and pass the airworthiness examination. The test purpose, the clauses to be verified, test methods, and judging criteria shall be clear. This is the basic requirement that must be satisfied for all airworthiness verification tests. 9.3.2.2 Electromagnetic compatibility test of flight control system The airworthiness clauses verified through the electromagnetic compatibility test of flight control system include CCAR25.1353a and CCAR25.1431c. 1. Description of airworthiness clauses to be verified a. Clause 25.1431c of CCAR This clause requires that the installed radio and electronic devices, control devices, and their interconnection lines shall be mutually compatible with the

652 Chapter 9 electromagnetic environment formed by other airborne equipment or systems on the aircraft. In other words, the operation of any airborne equipment or systems on the aircraft (including radio and electronic airborne equipment or systems) will not affect and interfere with the operation of radio and electronic equipment or systems at the same time as regulated by civil aviation regulations. This clause is not verified independently and it is generally verified with the aircraft level electromagnetic compatibility test. Please refer to chapter 7 for the content. b. Clause 25.1353a of CCAR This clause requires that the electrical equipment and installed electrical equipment and control devices shall ensure that the operation of any airborne equipment or airborne equipment systems will not adversely affect other systems and airborne equipment that are working at the same time and play a major role in the safe operation. Any possible electrical interference on the aircraft shall not cause danger to the aircraft or its systems. 2. Airworthiness verification methods The electromagnetic compatibility verification test of the flight control system checks whether the flight control system and other airborne systems work stably under different power supply modes, whether the control screen has uncommanded jumping or dithering, whether the information display is stable, and whether the graphics and state parameters have jumping. The specific test items are shown below: a. Measure the stability of electronic equipment including PFCC, automatic flight control computer (AFCC), actuator controller cross-linking command sensor, flight control feedback sensor, and actuator system control valve. b. Check whether the primary flight control/automatic flight control cross-linking bus is susceptible. c. Check whether there is a susceptibility signal between flaps/slats controller and position sensor, tilt sensor, driving device, power distribution box, and avionics cross-linking equipment. d. Check whether the high lift power drive unit and system brake device are susceptible to the power distribution system. The indirect effect test of lightning protection of aircraft and high-intensity radiated field (HIRF) test of flight control system are generally carried out with the aircraft level tests. 9.3.2.3 Flight control system and structural mode coupling test The flight control system and structural mode coupling test is part of the ground vibration test. The ground vibration test is generally carried out by the discipline of aircraft structural strength to verify clause CCAR25.629 “aeroelastic stability requirements.” The flight control system and structural mode coupling test verifies whether the stability margin of the flight control system meets the design requirements when the flight control system works

Airworthiness verification test of the flight control system 653 normally and the primary flight control system and automatic flight control system are coupled with structural modes under certain working conditions. The test purpose, test principle, and test implementation are described in detail in chapter 7.

9.3.3 Flight test The aircraft mathematical model of the flight control system in ground tests such as the “iron bird” man machine combined test and quality simulator test mainly comes from aerodynamic data of aircraft, while the inaccuracy of the aerodynamic data causes the failure of fully simulating the real characteristics of aircraft such as the high angle of attack, air drop, and cabin door opening under stall conditions. Therefore the flight test shall be carried out for the qualification of the flight control system. It can be seen from this that the flight test is the final stage of evaluating the performance of the flight control system and the flight test verification results are the most authoritative. All essential characteristics of the flight control system will be shown only through the test flight of various flight items within the flight envelope, so this stage is crucial for the development of the flight control system. During the flight test, the primary flight control system and high lift control system usually have function verification in the two aspects of normal function of the system and fault simulation. The automatic flight control system mainly completes the automatic flight control of the aircraft and it is the outer loop of the primary flight control system [note: the primary flight control system is the inner loop, and the automatic flight control system realizes the automatic flight of the aircraft through the communication between the AFCC and the PFCC]. When performing the items of the airworthiness verification test flight of the automatic flight control system, the airworthiness clauses in the aspects of general inspection, maneuverability, outer loop performance, fault demonstration, etc. of the automatic flight control system are mainly verified. Some flight test items are described below to introduce how to verify the airworthiness clauses through the flight test, so as to evaluate whether the performance of the flight control system of large transport aircraft meets the airworthiness requirements. 9.3.3.1 Airworthiness clauses verification of the primary flight control system The airworthiness clauses of the primary flight control system include CCAR25.671a, c, CCAR25.672, CCAR25.677a, CCAR25.697c, CCAR25.703, and CCAR25.1309a. 1. Description of airworthiness clauses to be verified a. Clause 25.671a, c of CCAR This clause requires that each control component and control system shall realize their functions conveniently, smoothly, and accurately and this point shall be

654 Chapter 9 verified through analysis, test, or both. The aircraft shall be able to fly and land safety without using special pilot skills or force even when the flight control system and control surface (including trim, lift, drag, and feeling system) have any of following faults or jam within the normal flight envelope. It shall be ensured that possible malfunctions will have only a minor effect on the operation of the control system and it can be solved by the pilot. b. Clause 25.672 of CCAR This clause shows the necessity of the functions of SAS or other automatic systems or control systems with driving force to meet the flight characteristic requirements. c. Clause 25.677a of CCAR This clause requires that the design of the trim control component shall be able to prevent unintentional or rough operation, and the direction of the control shall be within the plane of motion of the aircraft and consistent with the direct motion of the aircraft. d. Clause 25.697c of CCAR This clause requires that under constant or changing airspeed, engine power (thrust), and aircraft attitude, all control surfaces shall respond to the motion rate of control components and characteristics of the automatic positioning device or load limiting device so as to ensure the aircraft has good flight characteristics and performance. e. Clause 25.703 of CCAR This clause requires the aircraft shall be installed with a takeoff warning system and meet a series of requirements. Clause 25.1309a requires that for all equipment, systems, and installations for which the airworthiness standards of aircraft have requirements on their functions, their design shall ensure that the predetermined functions can be completed under various expected operating conditions. 2. Airworthiness verification methods a. Normal state The flight test under normal state mainly verifies whether the function and performance of the flight control system meet the requirements. Under the normal mode or direct mode of primary flight control system, the function and performance of the flight control system can be checked by recording the threeaxis control force and control displacement (such as command force and displacement of control column/wheel and pedal), the command signals and position signals of each control surface (such as the commands of aileron/rudder/ elevator/spoiler, EFCS commands, actuator position and position signals, discrete commands, and position signals of horizontal stabilizer), the trim command signals and position signals (such as rudder trim command and rudder trim position), the system state quantities (such as the state of aileron/rudder/elevator/

Airworthiness verification test of the flight control system 655 spoiler/horizontal stabilizer and the working state of flight control system), relevant signals of primary flight control system (such as attitude angle (angle of attack, pitch angle, roll angle, etc.), angular rate (pitch rate, roll rate and yaw rate), overload, indicated airspeed, radio altitude, dynamic pressure, and static pressure), and signals of other systems cross-linked with the primary flight control system (such as landing gear speed signal and throttle lever position signal), so as to verify the airworthiness clauses 25.671a, 25.677a, 25.697c, 25.703, and 25.1309a. b. Fault simulation i. Fault of PFCC: In the flight test, the PFCC fault is simulated to verify the conversion between the direct mode and normal mode and whether the transient state of conversion meets the design requirements. The display information of the engine indication and crew alerting system (EICAS) is also checked. After the flight test is completed, airworthiness clauses 25.672a, 25.672b, and 25.672c can be fulfilled by verifying the “recorded fault information” of the maintenance system. ii. Fault of primary flight control actuator channel: In the flight test, the fault of the primary flight control actuator is simulated independently or in combinations on the basis of the loss of one or several control channels to verify whether the actuator fault will affect aircraft safety. After the flight test is completed, airworthiness clause 25.671c can be fulfilled by verifying the “recorded fault information” of the maintenance system. iii. Stagnation fault of control column/wheel: In the flight test, the stagnation fault of the control column/steering wheel is simulated to check the operation of the release mechanism of the control column/steering wheel after the control column/steering wheel is stagnated, to record the position information of control column/steering wheel and aileron/ elevator before and after the release, and to check the display information of EICAS at the same time. After the flight test is completed, the airworthiness clause 25.671c can be fulfilled by verifying the “recorded fault information” of the maintenance system. iv. Stagnation of control plane: In the flight test, the stagnation of single aileron/elevator/rudder is simulated to verify the fault will not affect the aircraft safety so as to verify the airworthiness clause 25.671c. v. Loss of trim function: In the flight test, the trim failure of control planes such as aileron, rudder, and horizontal stabilizer is simulated to verify the aircraft safety, so as to fulfill the airworthiness clause 25.671c.

656 Chapter 9 9.3.3.2 Airworthiness clauses verification of the high lift control system The airworthiness clauses of the high lift control system include 25.671a, c, 25.677a, and 25.1309c, d. 1. Description of airworthiness clauses to be verified Refer to section 9.3.3.1 for clauses 25.671a, c and 25.677a. Clause 25.1309c requires that warning information must be provided to indicate unsafe working conditions of the system to the crew and enable the crew to take appropriate corrective actions. Systems, control components, and associated monitoring and warning devices shall be designed to minimize the errors of the crew that may increase the risk. Clause 25.1309d requires that through analysis, or through appropriate ground, flight, or simulator tests if necessary, the high lift control system shall be verified to conform to the regulations of this clause. If the above three tests can be used for conformity verification, the ground test or simulator test is preferred. 2. Airworthiness verification methods a. Normal state In the flight test, the airworthiness clauses 25.671a, 25.677a, and 25.1301d can be verified by checking the flaps and slats release procedures and time, flaps and slats withdrawal procedures and time, the consistency between the flaps and slats control handle and control plane positions, as well as display information of EICAS. b. Fault simulation i. Fault of flaps and slats controller: In the flight test, a single FSECU fault is simulated and the flaps and slats system shall work normally. Record the flaps and slats system withdrawal or release procedures and time and check the consistency between flaps and slats control handle and control plane positions. Meanwhile, the airworthiness clauses 25.671c, 25.1309c, and 25.1309d can be fulfilled by verifying the display information of EICAS and the “recorded fault information” of the maintenance system after the flight test. ii. Fault of flaps/slats: In the flight test, the flaps withdrawal fault and slaps release fault are respectively simulated to verify the slats withdrawal condition and flaps release condition, so as to verify the airworthiness clauses 25.671c, 25.1309c, and 25.1309d. 9.3.3.3 Airworthiness clauses verification of the automatic flight control system The airworthiness clauses of automatic flight control system include 25.672a, b, 25.1309a, c, and 25.1329. 1. Description of airworthiness clauses to be verified Refer to section 9.3.3.2 for clauses 25.672a, b, and 25.1309a. Clause 25.1309c requires that warning information must be provided to indicate unsafe working

Airworthiness verification test of the flight control system 657 conditions of the system to the crew and enable the crew to take appropriate corrective actions. Systems, control components, and associated monitoring and warning devices [autopilot (AP) disconnection switch and AP connection switch, etc.] shall be designed to minimize the errors of the crew that may increase the risk. Clause 25.1329 makes requirements to the flight director (FD) system. 2. Airworthiness verification methods a. General inspection In the flight test, the crew disconnect or override the automatic flight control system in a variety of ways and demonstrate that the autothrottle can always be disconnected or overridden by crew in a variety of possible ways. Meanwhile, the interface between the control and display of the automatic flight control system and personnel is evaluated, normal functions of the test devices in the flight are verified, and the design of the flight control system is judged based on specific criteria to judge whether it meets design requirements. Then, the airworthiness clauses 25.672b, 25.1309a, c, and 25.1329 are verified. b. Maneuverability i. Stall warning disconnection: In the flight test, the airworthiness clauses 25.672b, 25.1309a, c, and 25.1329a, d, e, g, l can be verified by simulating the performance of the autopilot under approach conditions when the test aircraft is approaching the stall conditions and judging whether the stall warning disconnection meets the design requirements according to specific criteria. ii. Threshold value for warning trim position deviation: In the flight test, the airworthiness clauses 25.672a, b, 25.1309a, c, and 25.1329d, e, f, g, l can be verified by demonstrating the acceptable threshold value for pitch out of trim warning (attention: the threshold value shall not be too low to prevent premature warning nor too high to prevent affecting the safety and performance) and the response of aircraft on pitch axis when out of trim warning occurs and judging whether the threshold value for pitch out of trim warning meets the design requirement according to specific criteria. iii. Acceleration/deceleration under changes of configurations: In the flight test, the airworthiness clauses 25.1309a and 25.1329d, e, f, g can be verified by demonstrating the limit value of servo torque (the maximum limit value of torque of back drive actuator) and setting the automatic pitch trim at an appropriate level under changes of configurations and judging whether the acceleration/deceleration under changes of configurations meets the design requirements according to specific criteria. iv. Mach trim: In the flight test, the airworthiness clauses 25.627a, 25.1309a, and 25.1329g can be verified by demonstrating the Mach trim performance in the Mach trim

658 Chapter 9 zone and judging whether the Mach trim meets the design requirements according to specific criteria. c. Performance of FD mode This test is used to evaluate the performance of autopilot’s FD mode and also evaluate and adjust the command response of the autopilot and FD when necessary. i. Evaluation of takeoff (TO) mode: In the flight test, the airworthiness clauses 25.1309a and 25.1329b, e, f, g, h can be verified by verifying the director command of FD, the appropriateness of autopilot control components, and the performance of autopilot roll axis and pitch axis when the mode converts from FD mode to autopilot mode. ii. Evaluation of heading selection (HDG) and altitude hold (ALT) mode: In the flight test, set the aircraft trim as horizontal flight, press the HDG and ALT buttons on automatic flight control panel (AFCU), select “heading selection” and “altitude hold” mode and connect the autopilot, and check the correct display on the primary flight display. The airworthiness clauses 25.1309a and 25.1329b, e, f, g, h can be verified by changing the heading direction through the HDG button on AFCU and the HALF BANK selector switch and verifying the director command of FD, the appropriateness of autopilot control components, and the performance and stability of autopilot HDG and ALT. iii. Heading navigation (NAV) mode: In the flight test, the airworthiness clauses 25.1309a and 25.1329b, e, f, g, h can be verified by verifying the director command of the FD, the appropriateness of autopilot control components, and the performance of AP and FD under LNAV mode after a waypoint is input as a target to which the aircraft flies in various ways. iv. Capture of vertical speed (VS) and altitude: In the flight test, connect the autopilot, select HDG and VS mode, set a preset altitude higher than the current altitude, climb under VS mode at a very high speed, press ALT button before approaching to the altitude selection and capturing point, observe the requirements of the command and get the altitude and the overshoot relative to the reference altitude (the altitude when the ALT is selected). In the same way, set a preset altitude lower than current altitude and repeat the process above. In this way, the director command of the FD, the appropriateness of autopilot control components, and limited altitude capturing and tracking can be verified and the stability of the aircraft in the maneuvering period can be ensured. The airworthiness clauses 25.1309a and 25.1329b, e, f, g, h can be verified through the operations above. v. Vertical navigation (VNAV) mode: In the flight test, set FMS (flight management system) as VNAV descending mode and configure the aircraft to the maximum angle path as

Airworthiness verification test of the flight control system 659 quickly as possible, make the VNAV be in standby status, verify the warning character and color on each FD, observe whether the VNAV path and captured warning can be smoothly obtained, note the descending rate of VNAV and verify the maneuvering performance of AP and FD under VNAV mode. The airworthiness clauses 25.1309a and 25.1329b, e, f, g, h can be verified through the operations above. vi. Evaluation of autothrottle functions: In the flight test, check whether the autothrottle system can be correctly connected and work normally under various flight states and working modes, and verify relevant warning and indication, so as to verify the functions of the autothrottle. The airworthiness clauses 25.1309a and 25.1329a, k can be verified through the operations above. d. Fault demonstration i. Determination of fault detection time: This test procedure shall be completed before the fault test. The test shall cover flight conditions including climbing, cruising, maneuvering, descending, approaching (ILS and nonILS), and autothrottle on/off. The fault can be started on any axis and in any direction and shall not be notified to the pilot in advance. The test aims to demonstrate the fault detection time and warn the pilot to take corrective actions after the occurrence of the autopilot fault. The airworthiness clauses 25.672a, c, 25.1309b, and 25.1329d, e, f, g can be verified in this way. ii. Fault in climbing: In the flight test, pitch axis and roll axis faults are found, respectively, and the test aims to demonstrate that the effects of the fault will not result in the aircraft load exceeding the limit value specified by airworthiness requirements in the climbing process. The airworthiness clauses 25.672a, c, 25.1309b, and 25.1329d, e, f, g can be verified in this way. iii. Fault in cruising: In the flight test, pitch axis and roll axis faults are found, respectively, and the test aims to demonstrate that the effects of the fault will not result in the aircraft load exceeding the limit value specified by airworthiness requirements in the cruising process and determine the altitude loss during cruising process for the Flight Manual. The airworthiness clauses 25.672a, c, 25.1309b, and 25.1329d, e, f, g can be verified in this way. iv. Fault in descending: In the flight test, pitch axis and roll axis faults are found, respectively, and the test aims to demonstrate that the effects of the fault will not result in the aircraft load exceeding the limit value specified by the airworthiness requirements in the descending process. The airworthiness clauses 25.672a, c, 25.1309b, and 25.1329d, e, f, g can be verified in this way.

660 Chapter 9 v. Fault in maneuvering and flight holding: In the flight test, pitch axis and roll axis faults are found, respectively, and the test aims to demonstrate that the effects of the fault will not result in the aircraft load exceeding the limit value specified by airworthiness requirements in the maneuvering process and determine the altitude loss for the Flight Manual. The airworthiness clauses 25.672a, c, 25.1309b, and 25.1329d, e, f, g can be verified in this way.

9.3.4 Engineering simulator test 9.3.4.1 Classification and composition of simulator In the process of aircraft type design, different types of simulation systems can be built according to the requirements of simulation tasks in different design stages and available aircraft data and components. Typical simulation systems in civil aircraft design include desktop simulator, engineering simulator, system test bed, and “iron bird.” The desktop simulator has simple functions and is mainly used for the analysis and preliminary evaluation of the overall performance and operational stability characteristics of aircraft in the early stage of type design. The engineering simulator usually has a relatively complete simulated cockpit, can simulate a real aircraft cockpit control environment and visual scene, and is mainly used for pilot-in-loop aircraft handling quality evaluation, man machine interface, and cockpit layout analysis and evaluation. The simulation of aircraft flight includes the simulation of aircraft dynamic characteristics, simulation of system functions, and simulation of external environment, all of which require the establishing and running of a corresponding mathematical simulation model. The simulation of aircraft dynamic characteristics is the core and foundation to realize flight simulation. For a flyby-wire aircraft, the closed-loop response characteristics of the aircraft to the control input depend on the integration between the dynamic characteristics of the aircraft body and the control law of the flight control system. And the influence of the delay effect of the aircraft data sensor and the dynamic characteristics of the control plane actuator on the dynamic response of the aircraft shall also be considered. Therefore to accurately simulate the dynamic characteristics of the aircraft, core models such as aircraft body dynamical model (including engine model), control law model, sensor model, and actuator model shall be established. In the process of type conformity certification, some content should be verified with the conformity certification method of simulator test (MOC8). The objects of the engineering simulator test are mainly engineering developers and they will participate in the design as engineers. Generally, pilots will also be invited to participate in the design verification. In the R&D and airworthiness certification process of the flight control system of civil transport aircraft, the engineering simulator shall be verified and some potential dangers and failures shall be evaluated. For example, the clauses CCAR25.1309 and

Airworthiness verification test of the flight control system 661 CCAR25.1329b can be verified through the engineering simulator test and the dynamic characteristics of the aircraft’s flight dynamics can be simulated through the digital computer. Thus the performance of the aircraft, the skills and qualities of the pilots, and the man machine system can be evaluated. On one hand, the engineering simulator can get rid of some limitations of ground test verification and carry out more comprehensive verification. On the other hand, some dangerous test items can be preliminarily verified before the test flight of the aircraft, so as to find out problems as early as possible, reduce development risks, and provide a basis for the flight test. Also, the flight control, avionics, and other airborne systems can be comprehensively verified at the same time. In consideration of the composition characteristics of the engineering simulator, MOC8 method is generally adopted to carry out the test for clauses with great impact on the safety of the flight control system such as the clauses involved with system faults or functional failures, such as clauses CCAR25.671, CCAR25.672, CCAR25.1309, and CCAR25.1329. For the safety analysis results of the flight control system, fault simulation is carried out on the engineering simulator and the fault phenomena and fault effects are evaluated to verify the compliance with the clauses. 9.3.4.2 Simulator test verification In the process of aircraft development and airworthiness certification, the simulation tasks carried out on the engineering simulator are mainly the evaluation of pilot-in-loop tests, which are mainly in two categories: simulator R&D test and simulator conformity certification test (MOC8 test). 1. Simulator R&D test There is a relatively wide range of R&D test tasks that can be carried out on the engineering simulator, including the research of specific technologies, the validation and verification of different system requirements, and the preliminary evaluation of the conformity of airworthiness clauses. The requirements for the simulator R&D test are mainly from disciplines of flight control, operational stability, and cockpit layout, etc. The R&D test includes: a. Control law and flying quality evaluation test This test is conducted to analyze the flying quality and mission envelope of the aircraft under various configurations and flight conditions, improve the design of the control law and the aircraft’s control characteristics under the guidance of pilotin-loop test evaluation results, and study the influence of the response characteristics of various actuators and control links on the flying quality. b. Flight control system requirements validation and verification test The flight control system requirements validation test is conducted to validate some functional requirements on the engineering simulator in the development stage of the flight control system requirements. It is used to ensure the correctness and integrity of main requirements of the flight control system and restrict the occurrence

662 Chapter 9 possibility of unintended functions within the system or between relevant systems. The flight control system requirements verification test is conducted to verify that the flight control system meets the system requirements in an integrated simulation environment provided by the engineering simulator after the development of the software and hardware of the flight control system is completed. c. System safety evaluation test By simulating the failure scenarios of the flight control system on the simulator, it evaluates the impact of failure scenarios on aircraft safety and confirms the hazard level of failure conditions to support the hypothesis and conclusion of safety evaluation activities such as FHA of the flight control system and other systems. d. Cockpit evaluation test In different stages of aircraft design, through pilot-in-loop testing, it evaluates the operational characteristics of cockpit control components, verifies the correctness and reasonableness of the design requirements of the cockpit control system and whether the design requirements meet service requirements, and validates the man machine ergonomics of the cockpit control system. The characteristic evaluation test of cockpit control equipment includes the static and dynamic evaluations of cockpit control equipment and control characteristics under different control laws and different modes. 2. Simulator conformity certification test As for the compliance with airworthiness clauses, AC25-7C puts forwards that test flight is still a preferred conformity certification method.2 However, in the following cases, the test flight can be replaced by a simulator test to certify the compliance with airworthiness clauses: a. Test flight under high risks. b. The environmental conditions or aircraft state required by the test can hardly be achieved under the limited test flight conditions. c. The repeatability is demonstrated through simulation or different pilots participate in the demonstration in specific situations. d. Reasonable and wide range of flight test items are supplemented through simulation. The simulator can only be used to certify the compliance with the airworthiness clauses on the premise that it is confirmed that the simulator can represent the characteristics of the aircraft used for the certification. The simulator conformity certification test mainly covers the following five aspects. 1. Evaluation of handling quality under faults of flight control system Clause to be verified: CCAR25.671, CCAR25.672 To meet the requirements of airworthiness regulations that the aircraft shall continue to safely fly and land under faults, the faults that can hardly be realized in test flight or 2

AC25-7C. Flight test guide for certification of transport category airplanes, 2012.

Airworthiness verification test of the flight control system 663

2.

3.

4.

5.

that will cause high risks will be simulated and the pilot will conduct evaluation on the aircraft handling quality on the engineering simulator according to the safety analysis results of the flight control system. Verification of workload of minimum flight crew Clause to be verified: CCAR25.1523 Simulate the working scenarios of minimum flight crew on the simulator and evaluate the workload of the minimum flight crew. The test scenarios include manual flight, entry of standard instruments, departure of standard instruments, nonprecision approach (normal), nonprecision approach (turbulence), generator failure, nonprecision approach, single engine failure after passing V1, failure of two sets of hydraulic system, TCAS warning (failure of one pilot), fuel imbalance, go-around under crosswind, direct mode of flight control system, engine on fire, and icing environment. Verification of minimum flight weight Clause to be verified: CCAR25.25b The minimum flight weight selected by the aircraft can hardly be obtained in the flight test. The operational stability characteristics of aircraft at the minimum weight shall be evaluated on the simulator to verify the compliance with the clauses. The test items include single-engine takeoff, double-engine crosswind takeoff, double-engine crosswind landing, go-around, longitudinal control at low speed, lateral control at low speed, and overloaded bar force gradient-maneuvering characteristics. Evaluation of procedures in Flight Manual Clause to be verified: CCAR25.1585, etc. The pilot evaluates specific procedures in the Flight Manual on the simulator to evaluate the logical rationality of the procedures and the maneuverability. The evaluation carried out on the simulator is usually the evaluation of abnormal procedures and emergency procedures in the Flight Manual. Other difficult tests under test scenarios realized by test flight It includes the verification of handling quality and system function under special atmospheric conditions (such as gust and turbulent flow) and the safety evaluation of aircraft under uncontrollable high thrust of engine. The test method of the engineering simulator is close to the test flight method of real aircraft. The engineer or pilot will form a subjective evaluation while some other test and testing results can be evaluated as per the Cooper Harper scale. In the system definition stage, the engineering simulator can be used to develop the flight control law, and can also be used for the preliminary design and verification of the flight control system. In the detailed design stage, the engineering simulator can be used for the improvement and verification of the flight control law, and can also be used for the evaluation of flying quality and partial design and verification of the flight control system. In the system verification phase, it can be used for the pilot closed-loop test, especially for some airworthiness conformity tests and hazard

664 Chapter 9 analysis of system functions. Accurate test data and test methods can greatly cut the development cost of aircraft. In the airworthiness examination of civil aircraft, MOC8 is mainly used to verify the relevant airworthiness clauses on the engineering simulator and the airworthiness examination applicant shows the conformity by conducting tests on the engineering simulator. For high-risk verification items in the FHA analysis in system design, they will be put to a test flight on the engineering simulator first. By conducting a fault test of the flight control system on the engineering simulator, the influence of the flight control system fault on the aircraft can be evaluated. The fault level defined in the FHA of the flight control system is confirmed and risky items in the test flight are supported to verify the availability of emergency procedures in the Flight Manual. Some items of the simulator fault test of one type of aircraft are shown in Table 9.7. 9.3.4.3 Precautions for the simulator test The engineering simulator is installed with the aircraft motion simulation system, vision system, avionics simulation system, and cockpit control components necessary for the test. In the test, a fault model can be injected as needed to realize fault simulation. The engineering simulator can simulate the response of the aircraft under faults. Table 9.7: Fault levels and modes in MOC8 test verification. Fault levels Level I Level II

Level III

Fault modes Loss of roll control function of two ailerons Control column loses all artificial feel and return force Sudden deflection of single aileron control surface Steering wheel loses all artificial feel and return force Loss of yaw control function of rudder Pedal of two sets of rudder loses all artificial feel and return force Inconsistent actual unwarned flaps/slats position with selected position and predetermined position Loss of brake function of multifunctional spoiler and accompanied with sudden descending Uncommanded enabling of ground lift destruction function of two ground spoilers Loss of single aileron roll control function Sudden deflection of single multifunctional spoiler Uncommanded action of aileron trim unit Loss of pitch trim function of horizontal stabilizer Loss of pitch control limit function of elevator under high speed Control column loses half artificial feel and return force Rudder control right degraded Single-way uncommanded action of rudder trim Warned loss of flaps and slats release function Loss of ground lift destruction function

Airworthiness verification test of the flight control system 665 Before the airworthiness verification test, the test task list and test procedure shall be prepared first and they shall pass the supervision and examination by the airworthiness authority. Before the test is started, the fault models needed for the tests should be established correctly and the inspection before the engineering simulator test shall be completed. The main test content includes: 1. For the inspection of the flight control system model, the aircraft model, engine model, and control plane model of the flight control system are mainly checked to ensure each actuator model meets the design requirements within the specified error range. 2. For the inspection of the functions of the flight control system, the functions of each control channel of the flight control system are mainly checked to verify that the functions of the flight system meet the requirements. 3. For the testing of the flying quality of the simulator, the flying quality of aircraft and the closed-loop response characteristics of the flight control system are mainly verified to meet the requirements under various configurations and flight conditions of the aircraft. 4. Simulation test is conducted for faults of the flight control system to evaluate the impact of the faults on the aircraft. The pilot or designated personnel shall conduct the test on the engineering simulator according to the test items and relevant personnel of the disciplines of engineering simulator. The flight control and operational stability shall provide support to the test. For airworthiness verification items, examination personnel from the airworthiness authority shall participate in the verification. Please refer to chapter 4 of this book for specific test requirements and operating procedures.

9.3.5 Unit qualification test 9.3.5.1 Overview The development specifications of the flight control system are formed according to the system engineering development methods, type development requirements, airworthiness requirements, and industrial standards and specifications. Then, the development specifications are broken down to the airborne equipment to form product specifications of airborne equipment, describing the functions, performance, physical characteristics, and verification requirements of the products. The product specifications, as the top-level documents of the airborne equipment, shall stipulate the basic technical requirements for the products and the inspection procedures, rules, and methods that should be followed to ensure the satisfaction of these requirements. They shall also have the provisions of relevant standards and specifications to meet the requirements in the aspects of use and maintenance, etc. To ensure that the aircraft system can complete its intended functions, the general requirements for airborne equipment are as follows. The airborne equipment shall be

666 Chapter 9 installed on the aircraft and under natural and artificial operating conditions, any equipment installed on the aircraft shall be able to operate without function deterioration and will not interfere with the normal operation of other equipment. Therefore it is necessary to conduct a qualification test for the equipment installed on aircraft to prove they can meet the specified requirements under expected environmental conditions. In other words, they can meet the requirements of product specifications. The qualification test covers functional test, performance test, environmental test, and life/fatigue test. The environmental test standards involved mainly include Laboratory Environmental Test Methods for Military Materiel (GJB150A),3 Environmental Engineering Consideration and Laboratory Tests (MIL-STD-810G) (Environmental Engineering Considerations and Laboratory),4 and Environmental Conditions and Test Procedures for Airborne Equipment (RTCA DO-160G).5 The environmental test standard commonly used for airborne equipment of military aircraft in China is GJB150A (or previous version GJB150), the environmental test standard used by military aircraft in the United States is MIL-STD-810G (or previous version), and the environmental test standard used by civil aircraft in the United States is RTCA DO-160G (or previous version). Due to different missions, the civil aircraft and military aircraft have different requirements. As civil aircraft are mainly used to transport passengers, they mainly consider safety requirements. As military aircraft execute operational missions, they mainly consider performance requirements. Therefore, although the environmental test standards have different characteristics for different types of aircraft and the test parameters and test items will be different, most of the test items are the same or similar, including the natural environment, mechanical environment, and electromagnetic environment test. The environmental test requirements for airborne equipment of the flight control system of transport aircraft are close to those for civil aircraft. Usually, the overall unit in charge of types will tailor the standards GJB150 and DO-160 to form environmental test requirements for airborne equipment of different types of aircraft and the environmental test of airborne equipment of the flight control system shall follow the requirements. The specific test methods are shown in chapter 2 of this book. 9.3.5.2 Environmental test standards for airborne equipment of military aircraft The environmental test standard GJB150 used in the past for airborne equipment of military aircraft in China was prepared on the basis of the military standard of the US Environmental Test Methods for Air Force and Land Equipment (MIL-STD-810C) and in combination with China’s actual situation, which was at the level of the 1984. At present, GJB150A has been officially promulgated and this version is prepared by taking the American military standard MIL-STD-810F as reference. Aiming at the 18 kinds of environmental factors or their synthesis, MIL-STD-810F totally arranged 64 test items. 3 4 5

GJB150A. Laboratory environmental test methods for military materiel, 2009. MIL-STD-810G. Environment engineering consideration and laboratory tests, 2012. RTCA DO-160G. Environmental conditions and test procedures for airborne equipment, 2010.

Airworthiness verification test of the flight control system 667 Considering the characteristics of ships and warships, GJB150A added 8 test items, so there are a total of 72 test procedures. The environmental test in each stage is actually a series of tests composed of one or several test items corresponding to each environmental factor. Due to different test purposes and the influence of other factors, the number and sequence of test items included in the series of tests and the test conditions adopted by the test items are different. All of these should be tailored according to specific conditions. Generally, the overall unit of a specific type of aircraft will tailor the standards and form the environmental test requirements for that type of aircraft. The airborne equipment of the flight control system can be tested according to the environmental test requirements for that type of aircraft. Table 9.8 lists the main environmental test items of the environmental test standards GJB150A and MIL-STD-810F for airborne equipment of military aircraft for reference. Specific test content and methods are shown in GJB150A, MIL-STD-810F, and environmental test requirements for relevant types of aircraft. 9.3.5.3 Environmental test standards for airborne equipment of civil aircraft The environmental test standard RTCA/DO-160 was formulated by the SC-135 branch of the RTCA (Radio Technical Commission for Aeronautics). It defines the minimum standard environmental test conditions and test methods for airborne equipment and provides laboratory methods for the performance characteristics of airborne equipment that will definitely appear under typical environmental conditions during use. Since its initial promulgation, the standard has been updated and improved many times with the new understanding and new requirements for the actual working environment conditions of airborne equipment. At present, the latest version is G version, namely RTCA/DO-160G. Through constant revision and increase of test items, the test methods are becoming increasingly perfected. FAA recommends using DO-160G for the verification test in the form of advisory circular AC-21-16. This standard is also directly adopted for civil aircraft in China. The main environmental test items include temperature and altitude, temperature changes, damp heat, flight impact and crash safety, vibration, explosion, waterproofness, fluid sensitivity, sand and dust, mold, salt spray, magnetic effect, power input, power peak, audio transmission sensitivity, induction signal sensitivity, radio frequency sensitivity, radio frequency energy emission, lightning induction transient susceptibility, lightning direct effect, icing test, electrostatic discharge, and flammability test. In addition to the basic requirements in the aspects of common test equipment, test sequence, and laboratory environmental conditions, for each category of environmental conditions, DO-160G describes from the aspects of test objectives, the basis or principle for the selection of test categories, the test steps of specific categories, the test curves, and judging criteria. Table 9.9 illustrates the changes of previous versions of the DO-160. For airborne equipment of civil aircraft, if the military standard such as MIL-STD-810G was used for the environmental test before, a comparative analysis shall be carried out

668 Chapter 9 Table 9.8: Main environmental test items for airborne equipment of military aircraft. No.

Environmental factors

1

Low air pressure

2

Temperature

Test procedures/test items MIL-STD-810F Storage, air transportation Working/air transportation Quick decompression High-temperature storage High-temperature working

3 4

Contamination fluid Solar radiation

5

Water (rain)

6 7 8 9

Damp heat Mold Salt spray Sand and dust

10

Explosive atmosphere

11

Water (immersion)

12

Acceleration

13 14

Vibration Noise

15

Impact

16

Temperature humidity vibration altitude Vibration, noise, temperature Icing/freezing rain

17 18

Low-temperature storage Low-temperature working Low-temperature disassembly operation Temperature impact based on constant limit value Impact based on hightemperature cycling Fluid contamination Cycling heating effect Steady-state chemical effect Rainfall and rain blowing Waterproofness Raindrop Damp heat Mold Salt spray Sand blowing Dust blowing Dust falling Working in explosive atmosphere Immersion Fording Structural test Performance test Crash safety test General vibration Noise in reverberation field Grazing incidence noise Open resonance noise Functional impact Crash safety impact Qualification

GJB150A Storage and (or) air transportation Working and (or) hanging outside aircraft Quick decompression High-temperature storage (cycling storage and constant-temperature storage) High-temperature working (cycling working and constant-temperature working) Low-temperature storage Low-temperature working Low-temperature disassembly operation Temperature impact based on constant limit value Impact based on high- temperature cycling Fluid contamination Cycling test Steady-state test Rainfall and rain blowing Waterproofness Waterdrop Damp heat Mold Salt spray Sand blowing Dust blowing Dust falling Working in explosive atmosphere Explosion-proof test Immersion Fording Structural test Performance test Crash safety test General vibration Noise in reverberation field Grazing incidence noise Open resonance noise Functional impact Crash safety impact Qualification

Vibration-temperature test

Vibration-temperature test

Rain ice test

Icing and freezing rain test

Airworthiness verification test of the flight control system 669 Table 9.9: Changes of previous versions of DO-160. No.

Version

Time of promulgation

1

DO-138

1968

2

DO-160

1975

3

DO-160A 1981

4

DO-160B 1984

5

DO-160C 1989

6

DO-160D 1997

Main changes The standard is Environmental Conditions and Test Procedures for Airborne Electronic/Electrical Equipment and Instruments. It is mainly used in conjunction with some radio product specifications prepared by RTCA. The test items are relatively complete and they basically constitute a complete set of environmental test standard. The standard is Environmental Conditions and Test Procedures for Airborne Electronic/Electrical Equipment and Instruments. This document is not much different from DO138. The test items are slightly increased. For example, magnetic influence test and radio frequency energy emission test are added. This document consolidates high temperature, low temperature, high pressure, overpressure, and decompression tests into temperature-altitude test. The classification of equipment is more detailed and temperature and pressure level is more reasonable. The standard is Environmental Conditions and Test Procedures for Airborne Equipment. This document is a revised version of DO160, with no major changes in general content. In the revision of the document, opinions of member states of ISO/TC20/SC5 are consulted and the revision is approved by ISO. The standard code ISO7137 is approved and becomes an international standard. The standard is Environmental Conditions and Test Procedures for Airborne Equipment. This document is a revised version of DO160A. Compared with DO160A, it is different in general requirements of test and test procedures and has characteristics of international standards. The name of appendix A of this document is changed from Recommended Practice for Making Equipment Nameplate to Environmental Test Identification, and the environmental test situation is recorded in the format of environmental test record form. The document is a revised version of DO160B. This revision is agreed between the working groups 14, 31 and 33 of EUROCAE and the technical committee 20 of ISO. It is taken as a priority standard for environmental test of TSO. The major changes in this document are as follows. Icing test is added, a graph describing the change of test conditions with time is added in the temperature-altitude, temperature impact, and damp heat test, and the regulations on the nameplate identification of equipment are recovered. This document is a revised version of DO160C, with main changes as follows: 1. The fourth category is added for the category of equipment for temperature-altitude test. (Continued)

670 Chapter 9 Table 9.9: (Continued) No.

Version

Time of promulgation

7

DO-160E 2004

8

DO-160F

9

DO-160G 2010

2007

Main changes 2. The impact pulse waveform of flight impact and crash safety test is changed from half-sine wave to back peak sawtooth wave. 3. High value short vibration method is added for the vibration test. 4. Lightning induction transient susceptibility test, lightning direct effect test and electrostatic discharge test are added. This document is a revised version of DO160D, with main changes as follows: 1. Different test procedures are provided for different test methods. 2. Fireproof test is added. This document is a revised version of DO160E, with main changes as follows: 1. Language description is more complete and specific. 2. Test items do not have changes. This document is a revised version of DO160F, with main changes as follows: 1. Language description is more complete and specific. 2. Some test methods of electromagnetic environmental effect are improved. 3. Content of fireproof test are revised and improved.

according to DO-160G to indicate that it provides an equivalent safety level for the environmental test. If airborne equipment needs a HIRF test, the requirements of DO-160F or DO-160G shall be followed. If the test level required by HIRF test is class B or class C, it is acceptable to follow the requirements in DO-160E. If electronic and electrical airborne equipment have requirements on lightning protection, DO-160D or its later versions shall be used. 9.3.5.4 Precautions for the qualification test 1. Principles to be considered when using the environmental test standards a. For specific airborne equipment, not all test items are required to be carried out but shall be selected according to the specific situation. In addition, in most environmental test methods, several different test procedures or different test values (severity grades) are included and shall be selected reasonably as well. b. For the selection of test items, the aircraft type, aircraft performance, installation position (such as cockpit, cargo compartment, electronic equipment compartment, engine compartment, and aircraft surface), as well as different features of the

Airworthiness verification test of the flight control system 671 airborne equipment, shall be considered. Normally, the cockpit area is sealed, the temperature is controlled, and pressure is increased. The electromagnetic compatibility requirements for electronic equipment compartment are strict. The environment of engine compartment is the harshest, with high temperature, large vibration, and noise. The surface of the aircraft is affected by the solar radiation, rain, and sand and dust. Mechanical products are not sensitive to electromagnetic compatibility effects and sealed products do not need to undergo a waterproof test. c. Generally, there is no rigid regulation on the number of test pieces of the products. It is also acceptable to provide only one product for all performance and environmental tests. However, the test will take a long time and breed high risks. Several test pieces can be provided for the tests, so that different products can be selected for the tests according to the duration of the test, so as to shorten the test cycle. Before the airworthiness verification test, the number of test pieces shall be reasonably determined given the cost of sample, the number of test items, the test sequence, and other factors. d. If a single test piece should be used in a continuous test and multiple tests, it shall be remembered that the environment of different factors often induces or expands the harmful effects of other factors on the product, in addition to its own unique effects and destructive effects. For example, high temperature will exacerbate the corrosive effects of humidity and salt spray, and vibration will expand the cracks in materials caused by low temperature and sudden change of temperature. In the airworthiness verification test, the test sequence shall be reasonably arranged. The following principles shall be followed generally: i. The salt spray test shall not be conducted before the mold test. ii. The sand and dust test shall not be conducted before the mold test, salt spray test, or damp heat test. iii. The explosion-proof test is generally carried out after all tests listed in DO-160. iv. The flammability test is generally carried out after all tests listed in DO-160. e. To reduce costs, some tests can be combined if allowed by the verification requirements. The precondition is that the service environmental conditions specified in the original single environmental test method can be duplicated or even improved. When a combined test is conducted, it is allowed to carry out the single tests alternately. If the test items are conducted alternately, relevant information shall be provided in the environmental qualification form. For the temperature change test, the “ground low temperature withstand test and low temperature shortduration working test” and “low-temperature working test,” and the “ground high temperature withstand test and high temperature short-duration working test” and “high-temperature working test” can be combined together.

672 Chapter 9 f. If the equipment under test is composed of several independent units, each unit can be tested separately on the premise that the functions specified in relevant technical specifications of equipment can be ensured. What should be noted is that each unit of the equipment under test shall be independent and can correctly realize specified functions. For some large equipment, a complete machine test cannot be carried out as the load bearing capacity of the environmental test equipment is limited. For some complex systems, it is not appropriate to take the complete machine for all tests due to complex functions and the large number of test items of the complete machine. If different units of the system are distributed at installation locations in different environments, a complete machine test is not allowed due to different environmental conditions. Under this condition, the independent units can be tested separately. 2. Requirements for inspection of environment test Before the qualification test, the manufacturing status of the airborne equipment to be tested shall be clearly defined, recorded, and reported. The technical status of the airborne equipment shall typically represent its manufacturing status and any subsequent changes in requirements shall be added to the design document and approved by the purchaser or approver. Any faults, defects, and other deviations found in the qualification test shall be documented immediately and reported to the purchaser or approver to ensure the corrective actions agreed between the purchaser (or approver) and supplier can be taken in time. Any changes to the design not approved by the purchaser (or approver) shall not be executed. In the qualification test, the representative of the purchaser (or approver) shall be allowed to enter the test site. Test equipment used for qualification shall be effectively calibrated and traceable to relevant standards. If required by the purchaser (or approver), the supplier shall verify the effectiveness of the calibration by submitting calibration data. 3. Requirements for exemption of environmental test of airborne equipment According to the specific situation of the development of airborne equipment, the environmental test may be exempted in part or in whole and it is unnecessary to conduct the test again. The technical data of the environmental test of other verified products can be quoted, and the specific requirements are as follows: a. There is evidence proving prior qualification or engineering evaluation. b. There is evidence of the qualification of similar equipment and the similar equipment shall be compared in detail with it. c. The analysis judgment showing that the test can be exempted shall be submitted. d. Any exemptions shall be approved by the purchaser. e. The authority accepts the exemption of the test.

CHAPTER 10

Development and expectations on test techniques of the flight control system The test techniques of the flight control system always develop along with the development of aircraft conceptual design technology, flight control system design technologies, and test techniques themselves. From the aircraft conceptual design technology, with the gradual popularization of control layout and active control technology (ACT), the control plane has also developed from the traditional control plane to a generalized control plane and the focus of the test has shifted to the verification and validation of aircraft performance and flying quality under different layouts, different aerodynamic characteristics, and different operation controls. From the flight control system flight technology, the functions of flight control system have developed from control augmentation to integrated control and are developing toward intelligent and autonomous control. The information transmission mode of the flight control system has developed from machine electric hybrid control and fly-by-wire control to optical transmission control and is likely to develop into wireless transmission in the future. The architecture of the flight control system has developed from centralized to distributed and a new generation of network architecture has appeared. Flight control law has experienced the leap from classical control algorithm to state space algorithm and then to nonlinear control algorithm. The flight control actuator has developed from a mechanical actuator and hydraulic actuator to all-electric/more electric actuator and even an intelligent (smart) actuator. The development of these technologies requires an overall evaluation and verification of system safety, reliability, economical efficiency, and maintainability with complex system validation and verification methods from the aspects of system architecture, control algorithm, information transmission, etc. From the development of test techniques themselves, with the continuous development of technologies such as simulation technology, information technology, and computers, a large amount of high-fidelity simulation equipment will be used to gradually replace physical devices and the processing and analysis of the huge amount of data produced in the test process will gradually become automatic with the gradual application of cloud storage and cloud computing technologies, making the traditional centralized physical test become a flexible and intelligent test verification system of a distributed network. The aircraft system Test Techniques for Flight Control Systems of Large Transport Aircraft. DOI: https://doi.org/10.1016/B978-0-12-822990-3.00010-3 © 2021 Shangai Jiao Tong University Press. Published by Elsevier Inc. All rights reserved.

673

674 Chapter 10 integration and verification centered with “virtual iron bird” will become one of the important features of the verification of advanced aircraft management system in the future.

10.1 Role of aircraft conceptual design technology in promoting the flight control system test The design idea of the aircraft adopting control configured vehicle technology (CCV) is to enhance the role of the flight control system in the overall design of aircraft on the basis of the previous “triad” design of aerodynamic layout, engine, and aircraft body structure, so as to form the “quaternity” design idea. It can solve many compromised problems caused by the contradictions in the previous “triad” design idea and realize the transformation from “passive control” of aircraft to “active control” (ACT). Taking aircraft attitude and track control as the design objective of flight control system, the aircraft adopting CCV, according to the requirements of the aircraft to the flight control system, specifically sets up necessary forms of control surfaces or determines the structural parameters of control surfaces to greatly improve the aircraft’s comprehensive performance, reduce the difficulties in overall, aerodynamic, and structural design of aircraft, improve the aircraft’s safety, and obtain great benefits. The flight control technology realizing the control configured layout scheme is just the ACT and integrated control technology. The ACT adopted by large transport aircraft mainly covers direct force control, maneuvering load control, gust alleviation, riding quality improvement, and envelope protection control. The integrated control technology is just the comprehensive utilization of the control plane with the flight mission as the core and the role of ACT is also fully played while ensuring the flying safety of the aircraft. The control surface design of modern large transport aircraft is no longer involved with the aileron, elevator, rudder, spoiler, flaps, and slats in the traditional sense, but is defined according to the force (moment) required for aircraft control, namely the generalized control surface. For example, the power lift augmentation control realizes the power lift augmentation function by changing the up down airflow velocity of flaps after flaps seam opening and the thrust vector control generates the torque consistent with aircraft control by adjusting the airflow direction at the engine tail nozzle to realize the attitude control of the aircraft. Even the traditional control surface does not have a single function. For example, the aileron is used for assisted lift augmentation to deflect downwards and generate lift force at the takeoff and landing stage of aircraft, the spoiler is used for assisted rolling and it will be opened laterally to generate rolling torque when the aileron roll efficiency is insufficient, and the spoiler, aileron, elevator, and rudder are the control surfaces needed to realize different active control functions. It can be seen from the analysis above that the design of the aircraft adopting CCV puts the design of the flight control system ahead of the aircraft requirements demonstration stage;

Development and expectations on test techniques of the flight control system 675 the repeated iteration and optimization of the flight control, aerodynamic, engine, and aircraft body structure may take a long period and be accompanied by the experimental verification work of the flight control system, which may be a simulation verification and principle verification, or a flying quality and control law evaluation on an engineering simulator, or a flight verification. The verification of flight control systems for large transport aircraft with direct force control, maneuvering load control, gust alleviation, riding quality control, envelope protection control, or some active control functions is still a significant challenge. This not only results from the extremely complex configuration and cross-linking relationship of the flight control system and the large number of control modes, but also because the active control functions require the combination of ground test verification and in-flight test flight verification, the research of the ground simulation method of in-flight environment, in-flight test flight method, and the addition of the environment for test flight items, as well as further improvement of the flying quality design specification and evaluation system for aircraft with active control function. The comprehensive use of the generalized control surface reflects the comprehensive consideration of flight control, aerodynamics, and engine and gives full play to the comprehensive control ability of the aerodynamics, engine, and flight control system, which plays a crucial role in improving the comprehensive performance of aircraft. This design idea poses great challenges to the test of the flight control system. The so-called challenge is also reflected in the complexity of the test tasks, the authenticity of the simulation environment and the establishment of the evaluation system.

10.2 Role of flight control system design technology in promoting test techniques Modern large transport aircraft endow the flight control system with more functions and tasks. The application of aircraft integrated control, distributed network control, new information transmission medium, highly reliable nonsimilar redundancy computer, smart actuator, and information fusion technology greatly improves the safety and environmental adaptability of aircraft. The technologies above make the structure and algorithm of the control law become complicated and difficult and the classical control theory of single input and single output adopted in the past is becoming more and more unsuitable. Thus nonlinear theory and modern control theory become the key theoretical support for solving this technology. Through “functional integration” and “physical synthesis,” the integrated control-vehicle management system (VMS) integrates flight control, propulsion, and public equipment management systems that provide critical functions for aircraft flight safety into an organic entirety. To realize the integrated control of aircraft resources and energy, it is necessary to

676 Chapter 10 design new airborne network and data communication, fault detection and health diagnosis, distributed networked control, digital adaptive public equipment control, task-oriented integrated flight/propulsion performance optimization, and integrated flight/propulsion control technologies. Nowadays, the flying environment of aircraft is no longer affected merely by the atmosphere, temperature, humidity, wind, and other natural factors. The increasingly harsh electromagnetic environment, more intense electronic countermeasures, the emergence of laser weapons, and the imminent nuclear radiation all have a huge impact on future aircraft safety. Thus it is imperative to improve the antiinterference ability and data transmission ability of the flight control system. An effective method to solve this problem is to develop the fly-by-light (FBL) system, that is, digital FBL system. The FBL flight control system focuses on the management of FBL control redundancy, complex actuator structure, FBL distributed control, and FBL fault-tolerant data communication technologies. At the same time, with the continuous improvement and wide use of wireless communication, the economy, confidentiality, and convenience of wireless communications are more and more favored by the designers of flight control systems and a kind of flight control system with wireless communication as the medium for information transmission is quietly emerging. Its core is the solution for wireless transmission-based global time-triggered, distributed, redundancy system architecture, distributed multinode synchronization, real-time multitask allocation and scheduling, wireless transmission-based health management (PHM), active/fault emergency control and takeover technology, multinode, composable and intelligent maintenance, special frequency spectrum, identification, and confidentiality technologies. The flight control system of more electric aircraft shows the application of highperformance electric drive/actuation technology, which is based on a large-capacity power supply system and replaces other secondary energy sources with electricity. The highperformance electric drive/actuation technology can simplify the system structure, optimize resource allocation, improve energy utilization efficiency, power-to-weight ratio, reliability, testability, and maintainability, and can also reduce the whole life cost. It has become the development direction of advanced airborne actuator systems of aircraft in the future. From the analysis above, it can be seen that with the development of the control theory, flight control system integration, multimedia (FBL, wireless network) distributed control, and more electric aircraft, the experimental verification will also develop inevitably. In terms of advanced control theory, the problem we are facing now is how to evaluate the control law and the flying quality of the aircraft designed based on modern control theory and nonlinear theory. Therefore we must construct a development verification environment that meets the requirements of flight control system of modern aircraft and determine the judging criteria and specifications through a large number of tests.

Development and expectations on test techniques of the flight control system 677 The flight control system of integrated control aircraft no longer executes attitude control and track control independently, but completes the flight mission together with aircraft engine and public equipment and aims for least oil consumption, shortest time consumption, and lowest labor intensity of pilots. The inner loop and outer loop of the original flight control system as well as the loops cross-linking with the flight management system are no longer obvious and the functional integration and information fusion put higher requirements to designers and test verifiers. Therefore the flight control system of integrated control aircraft does no longer have the conditions for independent verification and cross-linking verification and the test method of the flight control system taking complete function, good flying quality and flying safety of the flight control system as the core is no longer applicable; it is replaced by the idea of aircraft management verification taking the optimal execution of the flight mission as the core. The basic idea of the verification is shown in Fig. 10.1. The multimedia development provides a variety of options for information transmission and the key points of verification are the rationality of system topology, the reliability, stability, and safety of information transmission, as well as the synchronization, failure, and reconstruction between distributed multiple nodes. A key feature of the flight control system of more electric/all-electric aircraft is that it adopts an all-electric or electrohydrostatic actuator rather than the aircraft control surface actuator based on hydraulic sources in the past. Therefore the research and verification technology of high-performance electric drive/actuation technology has become one of the key technologies of modern flight control systems.

10.3 Role of test techniques in promoting the flight control system test Virtual test technology directly brings a new concept of verification and validation. “Iron bird” is an integrated test bed coming up with the control system, which can have full-size simulation of the control mechanism, actuator system, and control surfaces. It is crucial for the verification of mechanically controlled aircraft in the early stage or aircraft with backup machinery control system. Static inspections such as polarity, transmission ratio, rod force, and rod displacement inspections and dynamic tests such as step, pulse response, bandwidth, and frequency sweep testes, and even simulation of structural deformation and system loading can be carried out on the “iron bird” test bed. With the gradual application of the fly-by-wire system, the concept of “electric bird” appears after the design technology of structure and mechanism, etc. is mastered. By cross-linking the cockpit control mechanism, computer, sensor, and actuator, the interface, logic, algorithm, function, and performance of the flight control system can be comprehensively verified. Now, with the simulation technology gradually becoming mature, the “virtual iron bird” is coming into being, which can have the all-digital simulation of structure, mechanism, controller, sensor,

678 Chapter 10

Figure 10.1 Basic idea of vehicle management system verification.

actuator, energy, and cross-linking systems, to form a virtual environment completely consistent with the real “iron bird” environment, and thus further reducing the input of resources, improving the test efficiency, and reducing the risk of the test. Multisystem comprehensive verification is an inevitable development trend of aircraft test technologies. In the early stage, the quality simulator test bed is mainly used to verify the control law of the flight control system, especially the control augmentation system and the flight performance and augmentation characteristics of the aircraft. As the functions of the automatic flight control system of large transport aircraft become more and more powerful, the verification content gradually is expanding to the control law of the automatic

Development and expectations on test techniques of the flight control system 679 flight control system. Meanwhile, the requirements for flight management and automatic navigation systems also become higher and higher. As a result, it is necessary to add the verification of relevant control algorithms into the quality simulator test. As the aircraft management system (VMS) becomes mature, the integrated control of electromechanical, avionics, and flight control systems is becoming a trend and the research and verification of these key systems affecting aircraft performance on a quality simulator also is becoming a trend. From the “iron bird” test bed, the development tendency of the comprehensive verification of systems can be seen from the experimental verification of the machinery control system and the mechanism and structure in early stage, the verification of flight control systems such as primary flight control, high lift control and automatic flight control systems, the cross-linking verification with power supply, hydraulic source, landing gear, and avionics systems, as well as the cross-linking test with fuel, environmental control, and other electromechanical systems. Rapid processing of a huge amount of information is one of the key technologies in the modern flight control system test. Due to characteristics of large transport aircraft such as multiple control surfaces, complex control law, large number of components, multiple interface types, multiple system logic, and high requirements for safety and reliability, with the development of intelligent equipment, Internet of things, cloud computing, and other technologies, a large amount of information and data can be automatically analyzed and processed from excitation simulation, test acquisition to static and dynamic state, criteria can be judged, and test conclusions can be determined on the basis of the expert database. Parallel testing is an effective way to improve the time of R&D and verification. In foreign countries, many new technologies will only be applied in type development after they are mature to some degree. However, in China, many technologies are becoming mature in the process of type development. In the field of large transport aircraft, there are relatively small technical reserves in China. With the further development of “more electric” and other new technologies, especially for the flight control system and other key systems that affect aircraft safety, the situation of parallel design and testing is bound to appear at a certain stage. For subsystems or technologies that are not so mature, the principle verification test should be planned well in the early stage, the degree of maturity should be constantly improved by combining with the development of prototype samples (C-shaped parts) and test pieces (S-shaped parts), and verification at different levels should be carried out for components on the basis of component testing, subsystem testing, and system testing to reduce unnecessary followup reworking. In the testing process, the durability, reliability, and environmental adaptability tests should also be synchronously conducted with the component, subsystem, “iron bird” and onboard ground tests. The planning and coordination on the relationship between these tests is critical for the comprehensive analysis and treatment of abnormal conditions in the test.

680 Chapter 10 In conclusion, the large transport aircraft in China is still in the stage of type development at present. Compared with the international advanced level, it still lags far behind in technical fields including nonsimilar redundancy synchronization, monitoring, voting, balance, integrated management and control of avionics and electromechanical systems, comprehensive control and reconstruction of multiple control surfaces, and active control technologies such as load alleviation and riding quality improvement. But in the fields of architecture design, control law design, flying quality evaluation of flight control system, airborne software design and testing, and system integrated test verification, it has made great achievements and created a precedent for the engineering practice of large transport aircraft in China. With the development of the aircraft conceptual design, flight control system design, and test technologies in China, the flight control system design and testing will complement and promote each other and will definitely catch up with the international advanced level and thus realize the Chinese dream of becoming an aviation power.

Index Note: Page numbers followed by “f” and “t” refer to figures and tables, respectively.

A Acceleration test, 114 115 turntable, 296 Acceptance test procedure (ATP), 36 ACE. See Actuator control electronics (ACE) Active control technology (ACT), 1, 673 674 Actual fly-by-wire flight control system, 310 Actual frequency response of highorder system, 478 Actuator fault return function and performance inspection, 434 435 main/standby conversion function inspection, 433 434 maximum output speed and displacement test methods and steps, 435 436 Actuator system dynamic (static) performance test, 530 531 test, 298 300 Actuator control electronics (ACE), 22, 29, 379 Actuator controller, 291 Adjusting flight test, 603 Advanced control theory, 676 AFCC. See Automatic flight control computer (AFCC) AFCP. See Automatic flight control panel (AFCP)

AFCS. See Automatic flight control system (AFCS) AFCU. See Auto flight control unit (AFCU) Aileron aileron-assisted lift augmentation function, 538 machinery control system, 344f, 346 350 Aileron-assisted lift augmentation function, 255 inspection, 254 Airborne computer and simulated target machine of flight control system, 227 Airborne data acquisition and recording system, 606 Airborne electrical unit, 94 Airborne equipment, 207 208, 239, 268, 271, 287, 290 291, 345 346 fault, 311, 313, 488 effects, 465 Airborne software of FCS, 187, 188t Airborne software testing, 200, 203 Airborne units, 42 development, 41 43 Aircraft aerodynamic performance calculation module, 376 airborne systems, 394 395 and aircraft system, 249 251 aircraft-level verification, 24 25 cockpit, 358 conceptual design technology, 673 675

681

control efficiency verification, 601 602 control stability evaluation, 489 490 control system, 239 6-DOF full amount motion equation, 377 level, 40 mass performance calculation module, 376 motion equation, 292 293 motion sensor, 85 90 performance calculation module, 375 power supply system, 94, 99 principle flight test, 599 600 safety, 642 643 security, 1 type determination flight test, 598 604 adjusting flight test, 603 conformity certification flight test of civil aircraft, 604 design and type determination flight test of military aircraft, 603 604 maiden flight of aircraft, 600 602 monitoring screen of flight control system, 601f Aircraft motion sensor driver, 397 403 linear acceleration turntable, 398 400 single-axis rate turntable, 397 398 three-axis flight simulation turntable, 400 401

682 Index Aircraft motion sensor driver (Continued) total (static) pressure simulator, 401 403 Aircraft simulator cockpit, 361 364 configuration, 362f internal layout of, 363f Airworthiness conformity verification test, 590, 592 regulations, 417 verification methods, 652 Airworthiness verification test, 631 633 certification requirements, 633 644 engineering verification test certification requirements, 633 639 flight verification test certification requirements, 639 642 practices of airworthiness verification test of military aircraft, 642 644 classification, 632t content, 635t schedule, 634t technical requirements engineering simulator test, 660 665 flight test, 653 660 laboratory test, 644 649 onboard ground test, 649 653 unit qualification test, 665 672 Alarm display test, 310 314, 646 Alarm display verification test in flight control system, 483 486 All-digital simulation, 241 242, 677 678 Altitude holding/selection, 244 Amplitude-limit pitch angle hold mode, 459 Analog interface inspection, 445, 457, 468 469 simulation, 214 215 systems, 359

Angle of attack protection function flight test, 619 620, 621t turntable, 296 297 Angle of attack disturbance test, 480 Angular displacement sensor, 367 Angular rate gyroscope assembly, 85 90, 142 Antiicing control system, 411 AP. See Autopilot (AP) Applicant, 633 634 flight test, 639 641 Application layer, 375 APU. See Auxiliary power unit (APU) ARINC429 bus, 33, 457 interface inspection, 444 signals, 378 ARJ21 aircraft, 7 Arming mechanism, 419 420 ATM. See Autothrottle actuator (ATM) Atmospheric data sensor, 92 ATP. See Acceptance test procedure (ATP) Attention level fault, 541 542 Auto flight control unit (AFCU), 404 Automatic control mode, 583 Automatic flight control computer (AFCC), 334, 379 380, 490 491, 652 Automatic flight control panel (AFCP), 334, 658 Automatic flight control subsystem, 2 Automatic flight control system (AFCS), 2, 271, 404, 598, 653. See also Fly-by-wire flight control system; Machinery control system airworthiness clauses verification, 656 660 of background aircraft, 335f fault test, 648 “iron bird” integration test built-in-test, 462 464 control function and performance test, 461 462

control logic and display function inspection, 459 461 failure effect test, 464 466 interface inspection, 456 457 overview, 454 455 polarity and transmission ratio inspection, 457 459 test principles, 455 456 onboard ground test, 564 575 basic modal function (performance) inspection, 565 567, 567f BIT, 564 565 closed-loop pulse and resonance test, 585f modal conversion logic check, 568 open-loop/closed-loop frequency response test, 580f realization diagram, 581f redundancy management inspection, 568 570 resonance test system, 583f stability margin test, 573 574 state and alarm display test, 574 575 transmission ratio inspection, 570 572 system introduction, 334 test items and methods fault simulation test, 342 functional test, 338 340 performance test, 340 342 test objectives, 334 336 test requirements environmental requirements for combined test, 336 337 preparation for test and precautions, 338 requirements for test equipment, 337 requirements for tested objects, 336 test results and judging criteria, 342 343 tester, 379 380 Automatic trim, 244, 276

Index Automatic trim, 244, 276, 495 Autopilot (AP), 656 657 Autopilot entry/exit logic and display function inspection, 460 461 Autothrottle actuator (ATM), 334 Autothrottle functions, 659 Auxiliary power unit (APU), 358 Aviation technology, 5 6 Avionics system, 410 411 exciter, 383 385 interface with, 378

B Back drive actuator, 334 Background aircraft flight control system, 28 33, 29f Backup mode retraction test, 328 329, 333 Ball screw actuator, 65 Big bang integration testing strategy, 205 BIT. See Built-in test (BIT) Boeing 777 aircraft, 22, 27f Bottom-up integration testing strategy, 205 Boundary limit and protection function inspection, 481 483 Brake logic check, 325 326, 332 Built-in test (BIT), 285 in automatic flight control system, 462 464 in-flight built-in-test, 463 maintenance built-in-test, 464 power-up built-in-test, 463 preflight built-in-test, 463 of automatic flight control system, 564 565 detection test, 300 301 function, 357, 646 of flight control system, 596 inspection, 452, 539 540 inspection of high lift control system, 561 562 functional inspection and testing, 440 441 in high-lift system, 452 Bus interface inspection, 457

C Cabin lighting system, 33 Cable tension compensator, 63 64 Calibrator factor, 90 CAS. See Control augmentation system (CAS) Cause analysis in flight control system, 412 414 CCV. See Control configured vehicle technology (CCV) CDR. See Critical design review (CDR) Central maintenance system (CMS), 30 Certification Compliance Plan, 632 633 Certification flight test, 641 642 China National Accreditation Service for Conformity Assessment (CNAS), 163, 171 172 Civil aircraft. See also Military aircraft conformity certification flight test of, 604 environmental test standards for airborne equipment of, 667 670 changes of previous versions of DO-160, 669t Civil aircraft (ATA100), 2 Civil aircraft Boeing 777, 4 5 Civil Aviation Administration of China, 640 Clearance measurement, 347, 350, 352 Closed loop characteristic test, 646 characteristic test and judgment, 317 control, 259 method, 216 Closed-loop frequency response performance test, 549 551, 550f Closed-loop frequency test, 309 310, 310f, 476 in flight control system, 476 479

683

Closed-loop pulse margin test, 584 585 judging criteria, 585 test items and methods, 584 585 Closed-loop simulation, 217 Cloud computing, 679 CMS. See Central maintenance system (CMS) CNAS. See China National Accreditation Service for Conformity Assessment (CNAS) Cockpit evaluation test, 662 heading control channel, 274 275, 275f lateral control channel, 273 274, 274f longitudinal control channel, 275 276, 277f system, 336 warning information display inspection, 486 Cockpit control system, 271, 287, 404 “iron bird” integration test of, 414 423 dynamic performance testing of, 422 423 static evaluation of man machine ergonomics of, 415 417 static performance testing of, 417 421 test principle, 415 Cockpit control unit subsystem, 276 Combined test of flight control subsystem of automatic flight control system, 334 343 of fly-by-wire flight control system, 287 318 of high-lift system, 318 334 of machinery control system, 343 353 of PCUs, 273 287 Command input time-domain characteristic test, 479 480

684 Index Communication interfaces of flight simulation system, 377 378 Component testing, 204 208 methods, 206 plan, 205, 205f results, 207 208 specification, 206 Composition, engineering simulator, 247, 248f Comprehensive test management system, 394 396 Computer and bus transmission method, 162 Computer technology, 5 6 Computer unit, 51 54 power collection data, 54t power supply characteristic test, 53t Configuration item testing, 208 210, 208f methods, 208 209 plan, 208 results, 210 specification, 209 210 Conformity certification, 660 661 flight test, 614 of civil aircraft, 604 Conformity inspection, 634 636, 637t Connection line error, 414 Construction process of software whole life cycle support environment, 229 230 environment evaluation stage, 229 230 implementation stage, 229 planning stage, 229 use and maintenance stage, 230 Control and display unit, 47 51 judging criteria and results handling, 49 51 OFF indication function test, 50t switch quantity output test, 50t test items and test methods, 48 49, 49t test objectives and test requirements, 48

Control augmentation, 5 6, 11, 598 Control augmentation system (CAS), 239 240 Control Augmentation System of Piloted Aircraft, 597 Control column, 285 286, 362, 386, 438, 473, 496 Control configured vehicle technology (CCV), 674 Control displacement force characteristic curve, 418 419 Control function, 443, 445 447 under backup working mode, 447 under degraded working mode, 446 in high-lift system, 445 447 under normal working mode, 445 446 and performance test, 461 462 Control law, 3, 7 8, 32 and flying quality evaluation test, 661 software, 7 8 code, 7 8 component testing, 206 Control load test, 648 649 Control logic check, 301 304, 324 325, 332 Control logic inspection in automatic flight control system, 459 461 autopilot entry/exit, 460 461 mode priority, 460 in high-lift system, 445 447 Control plane aerodynamic hinge moment simulation system, 295 aerodynamic load simulation system, 388 391 clamping stagnation, 312, 487 deflection limit function, 538 539 failure, 488 uncommanded motion, 487 488 Control plane deflection limit function, 255 inspection, 254

Control stiffness test, 421 Control surface motion sensor, 90 92 Control system. See also Flight control system deicing, 411 display and, 377 dynamic (static) performance test of machinery, 551 556 dynamic performance testing of cockpit, 422 423 of machinery, 426 428, 553 556 fault mode verification of machinery, 428 429 landing gear, 409 410 Control test under degraded mode, 328, 333 under normal mode, 326 327, 332 333 Control theory, 5 6 Control-display function inspection of flight control system, 406 Conventional mechanical aircraft, 1 Cooper Harper rating scale, 265, 266f, 357 Coverage of FCS verification, 24 Critical design review (CDR), 20 21 Cross-linking system, 408 412 signal fault, 465 466 Current pitch angle hold mode, 459

D Damp heat test, 151 153, 152t Damping characteristic test, 420 421, 427 428, 428f Data acquisition of flight test, 625 628 and processing system, 294 and recording system, 369 service computer, 226 Data bus interface inspection, 469 Data collection, processing, and evaluation methods, 262 266 objective evaluation methods, 263 265

Index requirements for data collection, 262 263 for data processing, 263 subjective evaluation methods, 265 266 Data display function, 251 Data recording function, 251 Data storage, 625, 626t Debugging, 403 414, 525 527 and preparation for flight control system, 403 414 problems in debugging process and cause analysis, 412 414 static adjustment and inspection, 405 408 and technical status of crosslinking system, 408 412 Debugging equipment (DIF), 228, 295 296 Defense Science and Technology Industry Laboratory Accreditation Committee (DILAC), 163, 171 172 Degraded working mode, 288 Deicing control system, 411 Dielectric strength, 57, 57t test, 61t, 71t, 78t, 112 114 DIF. See Debugging equipment (DIF) Differential transformer sensor, 90 DILAC. See Defense Science and Technology Industry Laboratory Accreditation Committee (DILAC) Discrete interface inspection, 445, 457 Displacement characteristic test, 279 281, 284 286 Displacement feedback control, 386 Displacement relation curve, 425, 500f Displacement signal generator, 423 424 Display control panel, 362 and control system, 377 function inspection in automatic flight control system, 459 461

warning function test in high-lift system, 451 452 Distributed flight simulation system, 373 375 Distributed support core layer, 374 Distributed test analysis system, 370 Driver module, 201 202 Dual-redundancy real-time Internet environment, 385 Durability test, 281 282, 287, 314 315, 315f, 331, 334 and judgment, 317 318 Dynamic (static) performance test of fly-by-wire flight control system, 530 551 of machinery control system, 551 556 Dynamic (static) pressure simulator. See Total (static) pressure simulator Dynamic models of aircraft systems, 375 Dynamic performance testing of cockpit control system, 422 423 of machinery control system, 426 428 of machinery control system, 553 556 Dynamic signal frequency response analysis system, 369 Dynamic testing, 206 Dynamic tests, 677 678

E EICAS. See Engine indication and crew alerting system (EICAS) Electric aircraft, 676 “Electric bird”, 677 678 Electric-driven tab effect mechanism, 239 Electrical interface analysis, 21 Electroexcitation dynamic test method, 86 87 Electrohydraulic actuator, 30 Electrohydraulic servo loading system, 390, 390f

685

Electromagnetic compatibility test, 166 167, 589 593 of flight control system, 522 523, 651 652 judging criteria, 592 airworthiness conformity verification test, 592 qualitative electromagnetic compatibility test, 592 quantitative electromagnetic compatibility test, 592 organization and implementation of, 593 test items, 589 590 test methods, 590 592 airworthiness conformity verification test, 592 qualitative electromagnetic compatibility test, 590 quantitative electromagnetic compatibility test, 590 591 Electromagnetic emission and susceptibility test, 162 166, 163t, 164t, 165t judging criteria and results handling, 164 166 test equipment and environmental requirements, 163 164 test items and test methods, 164 test objectives and test requirements, 163 Electromagnetic environment adaptability, 2 3 Electromagnetic environment protection test, 162 177 electromagnetic emission and susceptibility test, 162 166 electrostatic discharge protection test, 176 177 HIRF protection test, 171 176 lightning direct effect test, 167 lightning-induced transient susceptibility test, 167 170 Electromechanical actuator, 73 77 Electromechanical products, 3, 123 Electronic products, 3 Electrostatic discharge protection test, 176 177 Elevator, 2, 287 288, 312 313, 360 361

686 Index Emergency control displacement force characteristic test, 419 420 Emulation technology, 11 13 Endurance test, 180 181 Engine indication and crew alerting system (EICAS), 483 487, 655 Engine parameter, 30 Engine thrust equation, 292 293, 372 378 performance calculation module, 376 Engineering practice of flight control system test verification, 25 28 Engineering simulator composition, 247 design requirements, 248 253 requirements for engineer analysis and appraisal system, 251 253 simulation requirements for aircraft and aircraft system, 249 251 main functions, 247 248 network topology block diagram, 249f test, 660 665 classification and composition of simulator, 660 661 precautions for simulator test, 664 665 verification, 661 664 Engineering verification test and flight test (MC6), 632 633 Engineering verification test certification requirements, 633 639 conformity inspection, 634 636 preparation of test plan, 633 of test procedure, 633 634 of test report, 638 639, 638t witness test, 636 638 Environment architecture, 223 224 Environment composition and functions, 224 228, 225f composition and functions

of flight control test system, 226 227 of public test system, 226 of software development and testing environment, 228 Environment requirements, 222 223 Environmental Conditions and Test Procedures for Airborne Equipment (RTCA DO160G), 666 Environmental Engineering Consideration and Laboratory Tests (MILSTD-810G), 666 Environmental requirements for combined test, 277, 290 291, 320 321, 336 337, 345 346 Environmental stress screening test (ESS test), 177 178 Environmental test standards, 666 for airborne equipment of civil aircraft, 667 670 for airborne equipment of military aircraft, 666 667 Equipment qualification test (MC9), 632 633 ESS test. See Environmental stress screening test (ESS test) Exciters, 46 47, 52, 359 avionics system, 48 49 Execution structure layer, 374 External interface relationship of AFCS, 33, 34f

F F-22 aircraft, 10 11 Failure effect test, 444, 493 in automatic flight control system, 464 466 in flight control system, 487 488 in high-lift system, 452 454 Fatigue load displacement spectrum block, 281, 282t Fault mode verification of machinery control system, 428 429

Fault return function, 434 435 Fault simulation, 655 Fault simulation test, 310 314, 342, 647 648 Fault test, 331, 333 Fault warning function test, 444 in high-lift system, 451 452 FBL system. See Fly-by-light system (FBL system) FBW. See Fly-by-wire (FBW) FBW flight control system (EFCS), 29 30 FCS. See Flight control system (FCS) FCS airborne software model-based flight control system airborne software development, 212 222 significance of V&V, 187 190 software safety and reliability test, 230 238 software testing, 199 211 software whole life cycle support environment, 222 230 FCS control law and flying quality evaluation test data collection, processing, and evaluation methods, 262 266 design requirements for engineering simulator, 247 253 flight control system control law of large transport aircraft, 243 244 flying quality of large transport aircraft, 242 243 management of control law and flying quality evaluation test, 266 270 control of test process, 269 270 planning of test, 267 268 preparation for test, 268 269 summary of test, 270 stage division and objectives of evaluation test, 244 246 test items and methods, 253 262

Index FCS verification, 18 25 coverage and traceability, 24 requirements definition, 19 22, 19f, 21f management tools, 24 verification and assignment, 22 23, 23f system verification supports aircraft-level verification, 24 25 FCS verification test, 13 18 background aircraft flight control system, 28 33, 31f, 32f content, 34 37 integration and test verification, 36 onboard ground test and flight test, 36 37 validation in definition stage, 35 validation of users’ requirements, 34 35 verification test in design and implementation stage, 35 36 features, 13 15 functions, 15 model of flight control system design, 39f principles for verification test design, 16 procedures, 37 40 technical management of verification test, 17 18 technologies, 7 10, 8f, 9f development, 6 7 development trends, 10 13 software complexity, 10f FD system. See Flight director system (FD system) Flaps (slats) asymmetry test, 329 330, 333 Flaps (slats) normal control function, 557 560, 558f, 559t Flaps and slats controller unit (FSECU), 318, 381, 490 491 Flaps and slats WTB unit, 148

Flaps asymmetry protection function inspection, 449 Flaps tilt protection function inspection, 450 sensor, 90 92 test, 329, 333 Flaps/slats system fault simulation test, 648 Flight boundary limit and protection function inspection, 440 Flight control computer, 22, 51, 648 power characteristic test abnormal voltage transient curve, 109f connection of flight control computer under steady-state voltage limits, 105f connection of spike test equipment, 108f cross-link in, 103f limit values of DC input power voltage spike, 104f limits in normal voltage transient test, 106f spike generator open-circuit voltage waveform, 108f steady-state characteristic limits, 105t test conditions of abnormal voltage transient test, 110t test conditions of normal voltage transient test, 107t test items and test methods, 102 112 test requirements, 101 102 Flight control system (FCS), 1, 522, 673 control law of large transport aircraft, 243 244 diagram, 484f electromagnetic compatibility test of, 522 523, 651 652 flight technology, 673 integrated verification, 357 “iron bird” integration test boundary limit and protection function inspection, 481 483

687

closed-loop frequency response test, 476 479 failure effect test, 487 488 interface inspection, 468 471 management, 508 520 overview, 466 467 polarity and transmission ratio inspection, 471 473 stability margin test, 473 476 state and alarm display verification test, 483 486 test principles, 467 468 test results evaluation, 498 508 time-domain characteristic test, 479 481 onboard functional test of, 649 651 operation inspection under operation of engine content and methods, 576 judging criteria, 576 objectives and requirements, 575 576 requirements validation test, 661 662 and structural mode coupling test, 652 653 test system, 226 227 airborne computer and simulated target machine, 227 signal adapter unit, 227 simulation excitation unit, 227 test control computer of flight control system, 227 test techniques, 673 aircraft conceptual design technology role in promoting test, 674 675 design technology role in promoting test techniques, 675 677 test techniques role in promoting flight control system test, 677 680 tester, 359, 378 383 composition, 380f verification, 674 675

688 Index Flight director system (FD system), 656 657 Flight simulation software package, 375 377 Flight simulation system, 372 378, 390 391, 606 composition principle, 375f frame structure, 373f Flight simulation test under general maneuvering conditions, 492 493 Flight test, 36 37, 597, 653 660. See also Onboard ground test of flight control system airworthiness clauses verification of automatic flight control system, 656 660 of high lift control system, 656 of primary flight control system, 653 660 data acquisition, processing, and analysis, 625 628 data analysis, 629 of flight control system, 595 ground support facilities, 606 607 methods and requirements, 604 606 objects, 598 organization and implementation, 628 629 and management, 607 outline of flight test of flight control system, 612 622 basis for flight test outline preparation, 614 615 categories of test outlines, 613 614 examples of flight test items, 619 622 items selection, 615 616 status selection, 616 619, 616t plan, 608 610 requirements and objectives, 596 597 requirements for flight control system

basis of preparation, 610 items and requirements, 610 requirements of monitoring system, 610 612 requirements of testing system, 612 stages and content, 598 604 aircraft principle flight test, 599 600 aircraft type determination flight test, 600 604 pre-research flight test, 599 team training of flight test, 608 test items, 595 test system, 623 625 Flight test interface (FTI), 359, 371 372, 372f, 525 Flight test platform (FTB), 11 12 Flight verification test certification requirements, 639 642 applicant flight test, 639 641 certification flight test, 641 642 Flow sensor, 368 Flutter suppression, 1 Fly-by-light system (FBL system), 676 Fly-by-wire (FBW), 1 aircraft, 660 control, 240 mode, 274, 582 583 Fly-by-wire flight control system, 187, 404, 436 437, 499 507, 595, 598, 648 649. See also Automatic flight control system (AFCS); Machinery control system accuracy and measuring range of test sensor for test of, 293t damping ratio calculation results, 504t dynamic (static) performance test, 530 551 of actuator system, 530 531 BIT function inspection, 539 540 closed-loop frequency response performance test, 549 551

inspection of state and warning display and recording correctness, 541 542 modal conversion function inspection, 532 534 open-loop stability margin test, 547 549 safety protection function inspection, 535 539 time-domain disturbance performance test, 545 547 time-domain step performance test, 545 transmission ratio and polarity inspection, 542 544 trim function inspection, 534 535 dynamic (static) performance test, 530 551 equivalent matching results of longitudinal control channel, 502t evaluation of display and warning function inspection results, 501 of frequency-domain characteristic test results, 501 505 of modal conversion test results, 505 of polarity and stroke inspection results, 499 501 of stability margin test results, 506 507 of transmission ratio test results, 505 506, 506f of zero position inspection results, 499 heading equivalent matching results, 504t “iron bird” integration test, 429 441 basic status inspection and testing, 430 432 built-in-test functional inspection and testing, 440 441

Index logic function inspection, 437 440 redundancy management function inspection, 441 testing of servo actuator system, 433 437 zero position and stroke inspection, 432 433 for large transport aircraft, 289f open-loop/closed-loop frequency response test, 579f realization diagram, 580f realization diagram of closedloop pulse and resonance test, 584f requirements for amplitude margin and phase margin, 506t resonance test, 582f short-period pitch response calculation results, 503t steering wheel control force wheel displacement relation curve, 500f displacement control plane deflection angle relation curve, 500f system introduction, 287 288 test items and test methods, 297 315 actuator system test, 298 300 BIT detection test, 300 301 closed-loop frequency response test, 309 310 control logic check, 301 304 durability test, 314 315 fault simulation and alarm display test, 310 314 interface inspection, 297 298 modal conversion test, 304 polarity and transmission ratio inspection, 304 305 redundancy management test, 301 stability test, 306 309 time-domain response test, 305 306 test objective, 288 290 test requirements

environmental requirements for combined test, 290 291 requirements for test equipment, 291 297 requirements for tested object, 290 test results and judging criteria results of durability test and judgment, 317 318 results of open-loop test and judgment, 315 316 results of stability margin test and judgment, 317 time-domain and closed-loop characteristic test and judgment, 317 tester, 379 Flying Qualities of Piloted Airplanes, 597 Flying Qualities Standard for Airplane with Fly-by-Wire Control System, 597 Flying quality, 240 241 of aircraft, 595 of large transport aircraft, 242 243 Force dispute monitoring test, 647 Force feedback control, 386 Force displacement curve, 282 283 envelope curve, 63, 63f Formal review, 191 Four-redundancy tested signal, 49 Freezing test, 148 Frequency characteristic test, 437 Frequency performance test, 555, 555t Frequency response method, 87 Frequency response test, 578 580 of SMI, 582 584, 586f judging criteria, 582 583 test items and methods, 582 583 Frequency-domain flight quality criteria design, 12 performance testing, 422 423, 427 response test, 300

689

Frictional force, 282 283, 285 286, 348, 353 FSECU. See Flaps and slats controller unit (FSECU) FTB. See Flight test platform (FTB) FTI. See Flight test interface (FTI) Functional inspection, 645 646 Functional integration, 675 676 Functional/performance test, 46 92 computer unit, 51 54 control and display unit, 47 51 of flight control system, 644 646 manipulator unit, 54 58 mechanical actuating unit, 66 79 mechanical drive unit, 58 62 mechanical transmission unit, 62 66 requirements of test, 46 47 sensor unit, 79 92 Functional test of transmission ratio adjustment device, 346 347, 352

G General Specification for Automatic Flight Control Systems and Stability Augmentation System, 597 General Specification for Flight Control Systems of Piloted Aircraft, 597 GJB185 1986, 476 GJB2874 1997, 476 Glaciation test, 148 Glide slope, 244 Graphics generation computer, 365 Ground hydraulic energy, 391 392 Ground hydraulic source, 391, 404, 409, 423 Ground monitoring system, 606, 610 Ground power supply system, 392 394, 393f Ground spoiler, 2, 30, 295, 312, 434, 521, 645

690 Index Ground taxi test, 602 Gyro assembly. See Angular rate gyroscope assembly

H Hardware integrated testing, 222 Hardware-in-loop test, 355 Hardware-in-the-loop, 217 Heading channel, 276, 309, 530 Heading control channel, 279 280, 284 285 Heading hold/selection, 458 Hierarchical design, 271, 466 High airborne equipment fault rate, 414 High lift control system (HLCS), 2. See also Automatic flight control system (AFCS) airworthiness clauses verification, 656 external interface relationship, 33, 33f onboard ground test, 556 564, 557f BIT function inspection, 561 562 flaps (slats) normal control function and transmission ratio inspection, 557 560, 558f modal conversion function inspection, 560 safety protection function inspection, 560 561 state display and warning function inspection, 562 564 High temperature test, 125t equipment and environmental requirements, 125 example, 127, 128t items and test methods, 125 127 judging criteria and results handling, 127 objectives and test requirements, 124 High-intensity radiated field protection test (HIRF protection test), 171 176

judging criteria and results handling, 175 176 test equipment and environmental requirements, 171 175 test items and test methods, 175 test objectives and test requirements, 171 High-intensity radiated fields (HIRF), 43 test, 652 High-lift control system tester, 381 High-lift system, 319f gears of flaps and slats control handle, 320t “iron bird” integration test built-in-test and redundancy management function inspection, 452 control function and logic inspection, 445 447 display and fault warning function test, 451 452 failure effect test, 452 454 interface inspection, 444 445 modal conversion function inspection, 447 449 overview, 442 safety protection function inspection, 449 451 test principle, 443 444, 443f system introduction, 318 test items and methods, 322 331 backup mode retraction test, 328 329 brake logic check, 325 326 control logic check, 324 325 control test under degraded mode, 328 control test under normal mode, 326 327 durability test, 331 fault test, 331 flaps (slats) asymmetry test, 329 330 flaps tilt test, 329 polarity inspection, 322 324 slats tilt test, 329

unconventional control test of flaps and slats control handle, 330 331 test objective, 318 test requirements environmental requirements for combined test, 320 321 preparation for test and precautions, 321 requirements for tested object, 320 requirements for tester, 321 test requirements, 322, 323f test results and judging criteria, 332 334 High-lift system, 404 HIRF. See High-intensity radiated fields (HIRF) HIRF protection test. See Highintensity radiated field protection test (HIRF protection test) HLCS. See High lift control system (HLCS) Horizontal navigation, 244, 410 411, 460, 486 Horizontal stabilizer machinery control system, 345f system fault test, 648 Human machine combined test, 252, 255 257 Hydraulic actuator, 66 71 Hydraulic source, 47 Hydraulic source fault, 312, 487 simulation test, 647 Hysteresis, 56

I Icing test, 151t judging criteria and results handling, 151 test equipment and environmental requirements, 149 test example, 151 test items and test methods, 149 151 test objectives and test requirements, 148 149

Index IFBIT. See In-flight built-in-test (IFBIT) Impedance check, 431 In-flight built-in-test (IFBIT), 440, 462 463 Inertial slip test of trim mechanism, 419, 420f Integrated control-VMS, 675 676 Integrated test of aircraft system, 355 integrated management system, interface with, 378 of subsystem, 7, 14, 421, 466 467 Integrated verification test, 355 Interface inspection, 297 298, 443 analog, 468 469 ARINC429 bus, 444 in automatic flight control system, 456 457 data bus, 469 discrete, 445, 457, 468 469 in flight control system, 468 471 in high-lift system, 444 445 MIL-STD-1553B bus interface inspection, 444 445 for onboard ground test, 528 529 Interface with flight control system, 377 Interface with turntable, 377 Internet, 247 Internet of things, 679 “Iron bird” human machine combined test, 242 “Iron bird” integrated test bench, 4, 6 7 “Iron bird” integration test, 241 242, 355 356, 523, 556, 568, 598, 600 602, 606 608, 611, 636, 644 of automatic flight control system, 454 466 bed, 210 211, 268 269 of cockpit control system, 414 423, 417f debugging and preparation for flight control system, 403 414

of flight control system, 356, 466 488 “iron bird” integration test of fly-by-wire flight control system, 429 441 of high-lift system, 442 454 of machinery control system, 423 429 test environment and test support equipment, 358 403 aircraft motion sensor driver, 397 403 aircraft simulator cockpit, 361 364 avionics system exciter, 383 385 bed, 359 361 comprehensive test management system, 394 396 flight control system tester, 378 383 flight simulation system, 372 378 FTI, 371 372 ground hydraulic energy and ground power supply, 391 394 mechanical displacement signal generator, 385 391 sensor and test analysis system, 367 371 sound system, 366 367 vision system, 364 366 “Iron bird” man, 490, 492 493 “Iron bird” man machine combined test, 653 of failure effect, 496 497 of mode conversion, 495 496 overview, 489 490 PIOR trend scale, 508t results evaluation, 507 508 of takeoff and landing and free flight, 493 495 test principle, 490 493 test task list, 497, 497t “Iron bird” test, 677 679 Items of test flight, 595 597

691

L “L” profile, 361 362 Laboratory Environmental Test Methods for Military Materiel (GJB150A), 666 Laboratory test (MC4), 632 633, 644 649 control load test, 648 649 fault simulation test, 647 648 functional test, 644 646 requirements for control components, motions, and effects, 645t Landing gear control system, 409 410 Landing gear system performance calculation module, 376 Lateral control channel, 278 279, 282 284 Lateral polarity and transmission ratio inspection, 458 Layer-by-layer synthesis, 466 Lightning direct effect test, 167 Lightning-induced transient susceptibility test, 167 170 Linear acceleration turntable, 398 400, 399f Linear displacement sensor, 367 368 Linearity, 56 Loading system, 295 Logic check brake, 325 326, 332 control, 301 304, 324 325, 332 modal conversion, 568, 569t Logic function inspection, 437 440 flight boundary limit and protection function inspection, 440 lift destruction and drag increase function inspection, 439 440 modal conversion function inspection, 438 439 Logic inspection, 443, 445 447 under backup working mode, 447 under degraded working mode, 446

692 Index Logic inspection (Continued) under normal working mode, 445 446 Longitudinal control channel, 280 281, 285 287 Longitudinal polarity and transmission ratio inspection, 458 Longitudinal short-period response flight test, 621 622, 622t, 623t Low pressure (altitude) test judging criteria and results handling, 123 national environment test on conventional large aircraft, 120t test cases, 123 test data requirements of, 124t test equipment and environmental requirements, 121 test items and test methods, 121 123 test objectives and test requirements, 119 121 test procedures of low-pressure test, 121t Low temperature test, 128 130, 131t

M MA700 aircraft adopt open-loop FBW flight control systems, 7 Machinery control system, 404, 595. See also Automatic flight control system (AFCS); Fly-by-wire flight control system dynamic (static) performance test of, 551 556 dynamic performance test of machinery control system, 553 556 flight control system operation inspection under operation of engine, 575 576

onboard ground test of automatic flight control system, 564 575 onboard ground test of high lift control system, 556 564 static performance test of machinery control system, 552 553 wheel force wheel displacement control plane displacement performance test, 554t “iron bird” integration test, 423 429 dynamic performance testing of, 426 428 fault mode verification of, 428 429 man machine ergonomics evaluation of, 424 static performance testing of, 425 426 study on effects of mechanism support stiffness, 429 test principle, 423 424 system introduction, 343 344 test items and methods aileron machinery control system, 346 350 mechanical backup system of horizontal stabilizer, 350 352 test objectives, 344 test requirements environmental requirements for combined test, 345 346 requirements for test equipment, 346 requirements for tested objects and tested system, 344 345 test results and judging criteria, 352 353 test results evaluation, 498 499 Maiden flight of aircraft, 600 602 Main equipment fault test, 648

Maintenance built-in-test (MBIT), 431 432, 440, 452, 462, 464 Maneuverability, 2 3, 11 12, 27 28, 239 241, 473 474 Manipulator unit, 54 58 judging criteria and results handling, 57 58 test items and test methods, 54 57 test objectives and test requirements, 54 Man machine combined test. See Pilot-in-loop test (PIL test) Man machine ergonomics, 415 evaluation of machinery control system, 424 static evaluation of, 415 417, 417f Man machine interface, 361 functions, 406 Maximum force, 282 283, 285 286 MBIT. See Maintenance built-intest (MBIT) Mechanical actuating unit, 66 79 electromechanical actuator, 73 77 hydraulic actuator, 66 71 judging criteria and results handling, 71 73, 77 79 Mechanical backup system of horizontal stabilizer, 350 352 Mechanical displacement signal generator, 292 293, 385 391 composition principle, 387f Mechanical drive unit, 58 62 judging criteria and results handling, 61 62 test items and test methods, 58 61 test objectives and test requirements, 58 Mechanical environment test, 114 119 acceleration test, 114 115 noise test, 116 117

Index shock test, 117 119 vibration test, 115 116 Mechanical flight control system (MFCS), 29 30 Mechanical products, 3 Mechanical transmission unit, 62 66 ball screw actuator, 65 cable tension compensator, 63 64 judging criteria and results handling, 64 66 spring load mechanism, 62 63 MFCS. See Mechanical flight control system (MFCS) MFD, 483 487, 563 564 MIL-STD-1553B bus interface inspection, 444 445 MIL-STD-810C standard, 666 667 MIL-STD-810F standard, 666 667 MIL-STD-810G standard, 667 670 Military aircraft airworthiness verification test practices of, 642 644 design and type determination flight test of, 603 604 environmental test standards for airborne equipment of, 666 667, 668t Military standard, 230, 242, 667 670 MOC5 test, 649 MOC8 method, 661, 664t Modal conversion function inspection, 438 439, 444, 532 534, 560 in high-lift system, 447 449 inspection, 646 logic check, 568, 569t of high-lift control system, 448f test, 304 Modal entry condition inspection, 566, 567f Mode priority logic and display function inspection, 460

Model-based flight control system airborne software development model-based development methods, 212 213 SCADE model testing and verification features, 214 215 SCADE software testing process, 215 222 Model-in-the-loop, 217 Mold test, 153 156, 156t Motion clearance inspection, 278 282, 284 285, 417 422 Multichannel servo controller, 386 388 Multifunctional spoiler, 521 Multimedia development, 677 Multisystem comprehensive verification, 678 679

N Natural environment test, 119 162 damp heat test, 151 153 high temperature test, 124 127 icing test, 148 151 low pressure (altitude) test, 119 123 low temperature test, 128 130 mold test, 153 156 rain test, 145 148 salt spray test, 156 158 sand and dust test, 159 162 single event test, 140 142 solar radiation test, 142 145 temperature shock test, 130 132 temperature altitude test, 133 137 temperature humidity altitude test, 137 140 Network data transmission, 247 Noise test, 116 117 Nonlinear 6-DOF motion equation, 372 378 Nonsecondary power supply unit, 98 Normal working mode, 288

693

O Object-oriented and modular technology, 375 Objective evaluation methods, 263 265 Off-line flight simulation function, 252 Onboard ground test, 36 37, 242, 602, 632 633, 649 653 electromagnetic compatibility test of flight control system, 651 652 of flight control system, 521. See also Flight test cross-linking performance inspection between flight control system, 522 debugging and preparation before test, 525 527 dynamic (static) performance test of fly-by-wire flight control system, 530 551 dynamic (static) performance test of machinery control system, 551 556 electromagnetic compatibility test, 522 523, 589 593 functional and performance test, 522 installation and power-on inspection, 521 interface inspection, 528 529 organization and implementation, 576 577, 587 589, 593 structural mode coupling test, 522, 577 589 test items, test methods, and judging criteria, 528 576 test principle, 523 525, 524f flight control system and structural mode coupling test, 652 653 onboard functional test of flight control system, 649 651 Open-loop control, 259 frequency response test, 578 579, 579f method, 216

694 Index Open-loop (Continued) simulation, 216 217 stability margin test, 547 549, 548f, 549t, 550f static characteristic test of flight control system, 406 408 test and judgment, 315 316, 316f Open-loop/closed-loop frequency response test, 578 579 of automatic flight control system, 580f realization diagram, 581f of fly-by-wire flight control system, 579f realization diagram, 580f Operating force displacement curve, 56 Oscillation monitoring, 647 Overload protection, 244 function inspection, 482 Overspeed protection, 244, 255 function inspection, 482

P Parallel test flight, 640 Parallel testing, 679 Passive control, 674 PBIT. See Preflight built-in-test (PBIT) PCU. See Pilot control units (PCU) PDR. See Preliminary design review (PDR) PDU. See Power drive unit (PDU) Pedal force sensor, 368 and rudder trim mechanism, 279 280 PFC. See Primary flight computer (PFC) PFCS. See Primary flight control system (PFCS) PFD, 483 487 Physical synthesis, 675 676 PIL test. See Pilot-in-loop test (PIL test) Pilot control units (PCU), 29 30, 272, 358 system introduction, 273 276

cockpit heading control channel, 274 275 cockpit lateral control channel, 273 274 cockpit longitudinal control channel, 275 276 test objectives, 276 test requirements criteria for assessment of test results, 282 287 environmental requirements for combined test, 277 for test equipment, 277 278, 278t test items and methods, 278 282 for tested object, 276 Pilot directive power sensor, 83 84 Pilot instruction displacement sensor, 79 81 Pilot-in-loop test (PIL test), 355, 358, 364, 646 Pilot-induced oscillations (PIO), 252 253, 357, 605 606 PIOR trend scale, 508t Pitch angle, 539 disturbance test, 480 hold function, 481 protection function inspection, 483 Pitch angle holding function inspection, 254 Pitch angle limit, 254 PM. See Pulse modulation (PM) Polarity and transmission ratio inspection test, 436 Polarity inspection, 304 305, 322 324, 332, 418, 542 544 in automatic flight control system, 457 459 in flight control system, 471 473 under normal working mode, 471 472 parameter setting, 472t under simulated backup working mode, 472 473 Power characteristic test, 94 112

of flight control computer, 101 112 judging criteria and results handling, 101 test items and reduction principles, 94 99 and test methods, 101 test objectives and test requirements, 99 101 Power check, 431 Power drive unit (PDU), 30, 58, 318, 443, 527 Power supply fault, 487 effects, 464 465 simulation test, 647 648 Power supply system, 411 Power-up built-in-test (PUBIT), 440, 452, 462 463 Preflight built-in-test (PBIT), 433 434, 440, 452, 462 463 Preliminary design review (PDR), 20 21 Preresearch flight test, 599 Pressure sensor, 368 Primary flight computer (PFC), 29, 490 491 Primary flight control computer (PFCC). See Flight control computer Primary flight control system (PFCS), 19f, 653 airworthiness clauses verification of, 653 660 Problem tracking process, 25 “Prompt level fault”, 542 Protection function inspection, 444 PUBIT. See Power-up built-in-test (PUBIT) Public test control computer, 226 Public test system, composition and functions of, 226 Pulse modulation (PM), 174

Q Qualification test, 4 reliability, 179 unit, 183 184, 665 666

Index Qualitative electromagnetic compatibility test, 589 590, 592 Quality simulator. See Engineering simulator Quality simulator test, 653 bed, 678 679 Quantitative electromagnetic compatibility test, 589 592

R R&D test. See Research and development test (R&D test) Radio Technical Commission for Aeronautics (RTCA), 667 Rain test, 145t, 148t judging criteria and results handling, 147 148 test equipment and environmental requirements, 145 146 test example, 148 test items and test methods, 146 147 test objectives and test requirements, 145 Ram-air turbine (RAT), 30 Random waveforms, 382 Rate turntable, 86 88, 296 RDC. See Remote data concentrator (RDC) Real environment, 8 9 Real-time performance, 624 625 Redundancy management function inspection, 441 in high-lift system, 452 inspection, 568 570, 571t test, 301, 339 Regulations for Civilian Aircrafts, 13 14 Regulations for Conformity Certification of Civil Products and Spare Airborne Equipment (CCAR21), 631 Release force and release angle test, 279, 281, 284, 286, 428 429 Reliability analysis and testing, 232 238, 235f

Reliability test ESS test, 177 178 reliability acceptance test, 180 reliability growth test, 179 reliability preexposure test, 178 reliability qualification test, 179 Remote data concentrator (RDC), 33, 383 384 Requirements for test environment, 59 61 Research and development test (R&D test), 661 662 Resonance test, 581 of automatic flight control system, 583f realization diagram of closedloop pulse and resonance test, 585f of fly-by-wire flight control system, 582f realization diagram of closedloop pulse and resonance test, 584f Return performance test, 346, 350, 352, 553 Risk management process, 641 642 Roll angle disturbance test, 480 481 Roll angle hold function, 481 482 RTCA. See Radio Technical Commission for Aeronautics (RTCA) Rudder, 2, 287 288, 312 313, 360 361

S Safety analysis and testing, 232 Safety braking function inspection, 450 451 Safety protection function inspection, 449 451, 535 539, 560 561 Salt spray test, 159t judging criteria and results handling, 158 test equipment and environmental requirements, 157 test example, 158

695

test items and test methods, 157 158 test objectives and test requirements, 156 157 Sand and dust test, 160t, 162t judging criteria and results handling, 162 test equipment and environmental requirements, 160 test example, 162 test items and test methods, 160 161 test objectives and test requirements, 159 160 SAS. See Stability augmentation system (SAS) SCADE model testing, 218 222, 220f code integration testing, 221 222 model coverage analysis, 219 221 software and hardware integrated testing, 222 test environment, 218 test examples, 218 219 and verification features, 214 215, 215f traditional software testing, 214f SCADE software testing process, 216f testing process, 216 217 closed-loop simulation, 217 open-loop simulation, 216 217 testing strategy, 215 SDR. See System design review (SDR) Self-check, 431 432 Sensor, 545 for purposes, 92 and test analysis system, 367 371 Sensor unit, 79 92 aircraft motion sensor, 85 90 atmospheric data sensor, 92 control surface motion sensor, 90 92 judging criteria and results handling, 82, 85

696 Index Sensor unit (Continued) pilot directive power sensor, 83 84 pilot instruction displacement sensor, 79 81 sensor for purposes, 92 Servo actuator system testing, 298, 433 437, 435f fault return function and performance inspection of actuator, 434 435 frequency characteristic test, 437 main/standby conversion function inspection of actuator, 433 434 maximum output speed and displacement test methods, 435 436 polarity and transmission ratio inspection, 436 time-domain characteristic test, 436 437 Shock test, 117 119 Sideslip angle disturbance test, 480 Signal adapter unit, 227 conditioner, 368 369 isolation amplifier, 294 test board, 228 Simulated backup working mode, 288 Simulation of aircraft flight, 660 excitation unit, 227 scheduling management system, 374 technology, 677 678 Simulator conformity certification test, 662 Simulator test, 632 633 verification, 661 664 Simulink, implementation, method, 213 Simultaneity, 624 Single event test, 140 142, 143f, 144f Single-axis rate turntable, 397 398, 398f Single-phase AC airborne electrical unit, 99

Single-pulse control, 257 Sinusoidal frequency sweep, 258 Slant angle, 539 hold, 254 limit, 254 protection function inspection, 483 Slats asymmetry protection function inspection, 449 450 Slats tilt protection function inspection, 450 test, 333 SMI. See Structural mode coupling test (SMI) Software evaluation and flight control system trustworthiness test, 493 failure modes, 234t integrated testing, 222 process management system, 228 product, 190 reliability test, 232 review, 190 191, 192t safety and reliability test, 231f, 233f, 234f reliability analysis and testing, 232 238 safety analysis and testing, 232 safety and reliability, 230 232 software failure modes, 234t of sensor and test analysis system, 369 371 V&V, 189 190 in software design stage, 197 198 in software implementation stage, 198 199 in software planning stage, 195 196 in software requirements analysis stage, 196 197 validation process, 189 190 verification process, 214 Software development

computer, 228 process, 189 190 and testing environment, 228 DIF equipment, 228 signal test board, 228 software process management system, 228 test data management computer, 228 Software testing FCS airborne software, 199 211 component testing, 204 208 configuration item testing, 208 210 system testing, 210 211 unit testing, 200 204 process, 191 193 V&V in software testing stage, 199 Software whole life cycle support environment, 222 223, 224f construction process of, 229 230 environment architecture, 223 224 environment composition and functions, 224 228 requirements of environment, 222 223 Software-in-the-loop, 217 Solar radiation test, 142 145 judging criteria and results handling, 145 test equipment and environmental requirements, 144 test items and test methods, 144 test objectives and test requirements, 142 143 Solenoid valve and motor, 98 Sound system, 366 367 SOW. See Statement of work (SOW) Speed holding/selection, 244 Spoiler ground lift destruction/drag increase function inspection, 255 Spring load mechanism, 62 63 Square wave modulation (SW modulation), 174

Index Stability augmentation system (SAS), 239 240, 645 647 Stability margin, 547 Stability margin test, 307 308, 308f, 573 574 in flight control system, 473 476 and judgment, 317 Stability test, 306 309 of flight control system, 581, 586 587 Stall protection, 244, 251, 538 function inspection, 482 Standard interactive interface, 374 Standard waveforms, 382 Starting force, 282 283, 285 286 Starting stroke, 282 283, 285 286 State and alarm display test, 574 575 State display, 562 564 verification test in flight control system, 483 486 Stateflow, implementation, method, 213 Statement of work (SOW), 17 Static (dynamic) performance test. See Control function and performance test Static adjustment and inspection of flight control system, 405 408 Static analysis, 206 Static debugging, 407 Static evaluation of man machine ergonomics, 415 417 Static inspections, 677 678 Static performance testing of cockpit control system, 417 421 of machinery control system, 425 426, 552 553 Steady-state long-term photochemical effect test, 143 Steady-state voltage limit test, 104 105 Steering wheel clearance, 282 283 control force wheel displacement relation curve, 500f

displacement control plane deflection angle relation curve, 500f Step characteristic, 57, 70 Step performance test, 554 555, 556t Strength and stiffness test, 92 94 judging criteria and results handling, 93 94 test items and test methods, 93 test objectives and test requirements, 92 93 Structural mode coupling test (SMI), 522, 577 589, 652 653 implementation, 582 587 closed-loop pulse margin test, 584 585 frequency response test, 582 584 stability test of flight control system, 586 587 organization and implementation of, 587 589 test objectives, 577 578 test principle and methods, 578 581 frequency response test, 578 580 resonance test, 581 stability test of flight control system, 581 Stub module, 201 202 Subjective evaluation methods, 265 266 pilot-induced oscillations trend assessment scale, 267f Sudden release operation failure effect test, 453 Sudden withdrawing operation failure effect test, 453 Support stiffness test, 429 SW modulation. See Square wave modulation (SW modulation) System “iron bird” synthesis and verification, 357 System design review (SDR), 20 21 System level flight test, 600 System safety evaluation test, 662

697

System stiffness measurement, 347 348, 350 351, 353 System testing, 210 211 methods, 210 211 plan, 210 results, 211 specification, 211 System-level verification, 18 19 System-wide “iron bird” test bench environment, 2

T Team training of flight test, 608 Temperature shock test, 130 132, 133t Temperature altitude test, 134t, 138t judging criteria and results handling, 136 test equipment and environmental requirements, 133 test example, 137 test items and test methods, 133 136 test objectives and test requirements, 133 Temperature humidity altitude tests, 139t judging criteria and results handling, 140 test equipment and environmental requirements, 137 test items and test methods, 137 139 test objectives and test requirements, 137 Tension pressure sensor, 368 Test content and methods, 481 483, 552 555 Test control computer of flight control system, 227 Test data management computer, 228 Test environment check, 432 Test facilities, 27 28, 606 Test flight profile, 315 Test flight system, 14 15 Test integrated management system, 394 395

698 Index Test items and methods, 253 262 planning for test tasks, 254 preparation of test report, 261 selection of test state points, 254 255 test analysis report, 261 262 test control action, 257 259 test task list, 259 260, 260t Test objectives and test requirements, 163 Test outline categories of, 613 614 unit, 44 46 Test piece selection of unit qualification test, 183 184 Test procedure preparation, 633 634 Test report preparation, 261, 638 639, 638t Test sensor, 367 368 Test sequence of unit qualification test, 183 184 Test signal generation system, 606 Test state point selection, 254 255 Test support systems, 359 Test techniques role in promoting flight control system test, 677 680 vehicle management system verification, 678f Testability test, 181 183 Tester, 379 380 flight control system, 359, 378 383 high-lift control system, 381 requirements for, 321 The fly-by-wire flight control system, 454 Three-axis control augmentation, 244 Three-axis flight simulation turntable, 400 401, 401f “3211” control, 257, 258f TIA. See Type inspection authorization (TIA) Tilt angle sensor, 368 Time-domain characteristic test, 436 437, 479 481 disturbance characteristic test, 480 481

disturbance performance test, 545 547 performance testing, 422, 426 427 response test, 300, 305 306 step performance test, 545, 546f test, 317 Time-pulse control, 257, 257f Torque sensor, 368 Total (static) pressure simulator, 401 403, 402f Traceability, FCS verification, 24 Transmission ratio, 542 544 Transmission ratio inspection, 304 305, 436, 542 544, 557 560, 558f, 570 572 in automatic flight control system, 457 459 in flight control system, 471 473 under normal working mode, 471 472 parameter setting, 472t under simulated backup working mode, 472 473 Transport aircraft, 1 Trial and service flight test, 598 599 Trim function fault, 313, 488 inspection, 534 535 Trim range test, 419 Trim speed, 419 Turning and brake command output test, 421 Type inspection authorization (TIA), 639, 641 Type inspection report, 641

U UAV. See Unmanned aerial vehicles (UAV) UCAV. See Unmanned combat aircraft vehicle (UCAV) Unbalanced force measurement, 348, 353 Unconventional control test of flaps and slats control handle, 330 331, 333 Unconventional operation failure effect test, 453

Unilateral control stagnation, 313, 488 Unit development process, 42 43 Unit qualification test, 665 666 environmental test standards for airborne equipment of civil aircraft, 667 670 of military aircraft, 666 667, 668t precautions for qualification test, 670 672 test piece selection and test sequence of, 183 184 of unit qualification test, 184 Unit test of FCS categories of, 43 classification of FCS unit, 44, 45t dielectric strength test, 112 114 electromagnetic environment protection test, 162 177 endurance test, 180 181 flight control system airborne units by physical properties, 44t functional/performance test, 46 92 mechanical environment test, 114 119 natural environment test, 119 162 organization and implementation of unit qualification test, 184 power characteristic test, 94 112 principles for selection of test items, 43 process of unit development, 42 43 reliability test, 177 180 requirements for preparation of unit test outline, 44 46 strength and stiffness test, 92 94 test piece selection and test sequence of unit qualification test, 183 184 testability test, 181 183 Unit testing, 41, 200 204 methods, 201 202, 203f plan, 200 201

Index results, 203 204, 204f specification, 202 203 Unmanned aerial vehicles (UAV), 9 10 Unmanned combat aircraft vehicle (UCAV), 9 10 Upper limb accessibility analysis, 417

V “V-shaped” development process, 210 212, 211f V&V. See Verification and validation (V&V) “V”-shape development process, 7 8 Validation, 14 in definition stage, 35 of users’ requirements, 34 35 Validation level by level, 272 Vehicle management system (VMS), 7, 675 676, 678 679 Verification and validation (V&V), 189 process, 194 199 airborne software life cycle validation activities, 194f in software design stage, 197 198 in software implementation stage, 198 199 in software planning stage, 195 196

in software requirements analysis stage, 196 197 in software testing stage, 199 in system analysis and design stage, 194 195 requirements, 190 193 software analysis, 191 software review, 190 191 software testing, 191 193 significance, 187 190 Verification methods, 650 652, 654 660 Verification process, 25 28, 189 of flight control system of Boeing 777 aircraft, 26f Verification test, 14 15, 15f, 16f in design and implementation stage, 35 36 development, 6 7 features of, 13 15 functions of, 15 principles for verification test design, 16 procedures of, 37 40 technical management of, 17 18 Vertical navigation, 410 411 Vertical navigation mode (VNAV mode), 658 659 Vertical speed (VS), 658 Vibration test, 115 116 “Virtual iron bird”, 677 678 Virtual prototyping technology, 11 Virtual test technology, 677 678 Visibility, 417

699

Vision system, 364 366 composition principle of, 366f interface with, 378 on “iron bird” integrated test bed of flight control system, 364f VMS. See Vehicle management system (VMS) VNAV mode. See Vertical navigation mode (VNAV mode) Voltage transient, 111 112 VS. See Vertical speed (VS)

W Warning function inspection, 562 564 Warning level fault, 541 Wire pull sensor, 368 Witness test, 636 638 Work breakdown structure (WBS), 17

Y “Y-shaped” development process, 212, 213f “Z” profile, 361 362

Z Zero drift, 69, 75 76 Zero position and stroke inspection, 432 433