Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 1468604694, 9781468604696

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Aircraft Wheels, Brakes, and Brake Controls Key Principles for Landing Gear Design

R. Kyle Schmidt

Aircraft Wheels, Brakes, and Brake Controls Key Principles for Landing Gear Design

Aircraft Wheels, Brakes, and Brake Controls Key Principles for Landing Gear Design R. KYLE SCHMIDT

Warrendale, Pennsylvania, USA

400 Commonwealth Drive Warrendale, PA 15096-0001 USA E-mail: [email protected] Phone: 877-606-7323 (inside USA and Canada) 724-776-4970 (outside USA) FAX: 724-776-0790

Copyright © 2022 SAE International. All rights reserved. No part of this publication may be reproduced, stored in a retrieval system, or transmitted, in any form or by any means, electronic, mechanical, photocopying, recording, or otherwise, without the prior written permission of SAE International. For permission and licensing requests, contact SAE Permissions, 400 Commonwealth Drive, Warrendale, PA 15096-0001 USA; e-mail: [email protected]; phone: 724-772-4028. Library of Congress Catalog Number 2022936970 http://dx.doi.org/10.4271/9781468604702 Information contained in this work has been obtained by SAE International from sources believed to be reliable. However, neither SAE International nor its authors guarantee the accuracy or completeness of any information published herein and neither SAE International nor its authors shall be responsible for any errors, omissions, or damages arising out of use of this information. This work is published with the understanding that SAE International and its authors are supplying information but are not attempting to render engineering or other professional services. If such services are required, the assistance of an appropriate professional should be sought. ISBN-Print 978-1-4686-0469-6 ISBN-PDF 978-1-4686-0470-2 ISBN-ePub 978-1-4686-0471-9 To purchase bulk quantities, please contact: SAE Customer Service E-mail: [email protected] Phone: 877-606-7323 (inside USA and Canada) 724-776-4970 (outside USA) Fax: 724-776-0790 Visit the SAE International Bookstore at books.sae.org

Chief Growth Officer Frank Menchaca Publisher Sherry Dickinson Nigam Product Manager Amanda Zeidan Director of Content Management Kelli Zilko Production and Manufacturing Associate Erin Mendicino

Dedication For my wife, Natalie, and my children, Jacob, Dylan, and Hunter.

©2022 SAE International

v

Contents Acknowledgements

xi

Preface

xiii

A Note on Units

xv

Introduction

xvii

About this Book

xxv

CHAPTER 1

Brakes

1

CHAPTER 2

Aircraft Deceleration

9

CHAPTER 3

Brake Sizing Energy

15 15

Kinetic Energy Calculation

16

Rational Brake Energy Calculation

18

Torque

20

CHAPTER 4

Brake Design

23

Brake Actuation

30

Mechanical Connection to the Landing Gear Structure

34

Weight

36

©2022 SAE International

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Contents

CHAPTER 5

Wheel and Brake Certification and Recommended Practices

43

CHAPTER 6

Brake Issues and Concerns

59

Vibration

59

Failure and Degradation Modes

61

CHAPTER 7

Braking Accessories

65

Brake Cooling Fans

65

Brake Temperature Measuring Systems

65

Retraction Braking

66

CHAPTER 8

Wheels

69

Bearing Selection and Preload

74

Over Temperature and Over Pressure ­Relief

79

Wheel Mass

79

Failure Modes

80

Bearing Failure

80

Wheel Rim Release

80

CHAPTER 9

Brake Control

83

Brake Control Architectures

85

Antiskid and Related Functions

92

Braking Efficiency

95

Antiskid Dynamics

95

Antiskid Hardware

98

Autobrake

101

Failure Modes

101

Contents

ix

References

103

Index

107

Acknowledgements

T

he book you are reading, while itself a stand-alone book, originally constituted a single chapter of The Design of Aircraft Landing Gear. Below are the acknowledgements for that book—they remain true and valid for this excerpt. The idea for this stand-alone book was Sherry Nigam’s and I thank her, Erin Mendicino, Amanda Zeidan, and the other staff at SAE International Books for their assistance with this project. I would like to thank my family: Natalie, Jacob, Dylan and Hunter, for their patience, support, and encouragement, without which I would not have been able to dedicate the time to writing this book. I would also like to thank my father, Bob Schmidt, who was the first to read and comment on each chapter as it was produced. I thank my colleagues in Canada, France, the USA, and the UK who have read sections and chapters of this work and provided me with suggestions, corrections, and encouragement. In particular, I would like to thank those who gave up their time to review and comment: Bruno Aldebert, Steve Amberg, Rod Van Dyk, Andrew Ellis, Jack Hagelin, Dan Hetherington, Marianna Lakerdas, Grant Minnes, Andy Paddock, Michael Saccoccia, Jon Smith, and Peter Taylor. Monica Nogueira at SAE International has supported me from the outset of this project, gently prodding to ensure that it was completed! I would also like to thank the industry expert reviewers who reviewed portions of the book on behalf of SAE International: CB Alsobrook, Gregg Butterfield, David Brill, Bob Knieval, and Henry Steele. Finally, I would like to thank Ian Bennett and Mark Shea who reviewed the entire manuscript in detail and provided a number of excellent comments and suggestions.

©2022 SAE International

xi

Preface

T

he author has been fortunate enough to work in the field of aircraft landing gear for over twenty-five years and in three countries: Canada, France, and the UK, and to have held a variety of engineering roles relating to the development of new landing gears and the sustainment of existing landing gears in service. Landing gear provides an intriguing and compelling challenge, combining many fields of science and engineering. This book is an excerpt of The Design of Aircraft Landing Gear intended to present a specific element of landing gear design in an accessible way. The content here and in the original book was born of the author’s desire to learn ever more about landing gear — their history and the ways in which others have addressed their problems and challenges; in continuously striving to learn more about the field, it was considered advantageous to put these learnings into print in the hope that they can assist others. The book is intended, broadly, for two audiences: experienced aircraft and landing gear design engineers, for whom it is hoped that the book will act as a reference as well as an ‘idea book’, and for those new to the field who are, perhaps, working on their first landing gear design (maybe as part of their education). For the latter, it is hoped that the book provides the information needed to aid in their design and studies, and that they are as intrigued and compelled by the beautiful complexity of landing gear to consider this challenging field for their future employment. No single book can provide all the answers; throughout the chapters there are a number of references to additional documents which can aid in the design, development, and support of landing gears and their associated systems. In particular, documents produced by the SAE International A-5 committees on aircraft landing gear are widely referenced and participation in these committees is highly recommended to readers of the book and practitioners of landing system engineering. The opinions and approaches outlined in this book are those of the author and do not necessarily represent those of his employer (Safran Landing Systems). Although a great deal of care has been taken in the preparation and review of this work to ensure that the approaches, methods, and data provided are accurate, the author and publisher are not liable for any damages incurred as a result of usage of this book, for typographical errors, or for any misinterpretations.

©2022 SAE International

xiii

A Note on Units

W

herever possible, units in this book follow the International System of Units (SI, also known as the metric system) approach. However, aircraft and landing gear are international in nature and many components and analysis approaches are conducted in US Customary units. In particular, some empirical formulas are based on US Customary measures and do not lend themselves to conversion to another system of measure. In general, most calculations can be performed using either SI or US Customary units, provided two different measurement systems are not mixed in the same calculation and that the units utilized are self-consistent. An area where attention needs to be paid is the use of the US customary unit of weight and force, the pound, which is often colloquially used as a unit of mass (with an implicit assumption of earthly gravity); calculations conducted in US customary units which require units of mass can employ the ‘slug’ – which is defined as the mass that is accelerated by 1 foot per second per second when a force of one pound is exerted on it. A familiarity with both systems of measure is recommended due to the international nature of the aircraft business.

©2022 SAE International

xv

Introduction

T

he aircraft landing gear and its associated systems represent a compelling design challenge: the retractable landing gear is simultaneously system, structure, and machine; it supports the aircraft on the ground, absorbs landing and braking energy, permits maneuvering, and retracts to minimize aircraft drag. As the system is not required during flight it represents dead weight and significant effort must be made to minimize its total mass. The landing gear is one of the most complex and diverse systems on an aircraft. An article in Flight magazine [1] in 1940 expressed this, “for on no other part of the aeroplane is there such scope for engineering ingenuity and no other part can boast of so many ways of achieving the desired result”. This remains true today, many decades later. An expert in landing gear must be conversant with a wide range of engineering disciplines including materials, mechanisms, structures, heat transfer, aerodynamics, tribology, and many more. Depending on the given aircraft’s needs a landing system may be little more than wheels and tires attached to suitable aircraft structure or it may be a complicated system enabling performance on unpaved runways, steering, kneeling, retracting, and permitting further aircraft operations. Very few aircraft are designed for no other purpose than to carry the landing gear (perhaps only the Messier Laboratoire test aircraft qualifies); rather, the aircraft is designed to perform a function and the landing system must enable this function with high reliability and low mass.

•• The aircraft landing gear and system provides a number of functions: •• The landing gear, wheels, and tires support the aircraft on the ground •• The tires and shock absorber absorb vertical energy during landing and minimize shocks during ground maneuvering

©2022 SAE International

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xviii

Introduction

•• The brakes absorb forward energy and hold the aircraft when stopped and parked

•• Differential braking and steering permit turning and maneuvering on the ground

•• Specific structure and attachments permit towing, jacking, and tie down of the aircraft

•• The landing gear can retract to minimize airframe drag •• The landing gear can articulate to change the aircraft geometry – assisting takeoff or kneeling for loading

•• The landing gear can include driven wheels to maneuver the aircraft without relying on main engine thrust

•• The landing gear can comprise attachments to permit catapult launch from ships as well as airframe mounted arresting gear

•• The landing gear can include a tail bumper for protection of tail cone structure The wheels, brakes, and braking system of an aircraft perform vital and critical functions. The wheel provides the interface between the tire and the landing gear structure: retaining the tire inflation pressure, carrying bearings, transferring ground loads from the tire to the landing gear structure, providing a mechanical interface to the brake, and transmitting braking torque to the tire. While a number of means exist to provide aircraft retardation, almost all aircraft dissipate a significant amount of their kinetic energy into the wheel brakes when landing, aborting a take-off, and maneuvering on the ground. The brake control system ensures that the deceleration command from the pilot or aircraft system results in appropriate actuation of the brakes, assists in maintaining directional control of the aircraft, and in many cases ensures that tires do not skid and that the tire is maintained at its peak possible friction value. The Wright Brothers did not include wheeled landing gear on their gliders or first successful powered aircraft, preferring a light weight skid arrangement (Figure 1). However, the practicality of wheels meant they were soon included on the Wrights’ aircraft as well as that of others. Bleriot made the first crossing of the English Channel with an aircraft having sprung wire wheels as shown in Figure 2 . For a period of time in the 1930s, aircraft speed was increasing, with the result that the proportion of aircraft drag developed by the landing gears increased. There was reluctance to adopt the complexity and weight of retractable landing gears and a number of aircraft with streamlined fairings around fixed landing gears were produced. An effort to provide a more streamlined form for fixed gear was Dowty’s internally sprung wheel, as used on the Westland Lysander (Figure 3). The internally sprung wheel mounted an oleo-pneumatic shock absorber within the wheel, providing a compact package that could be streamlined by fixed fairings. Little space remained for brake installation.

Introduction

xix

© SAE International.

 FIGURE 1   The Wright Flyer Moments into its First Flight, December 1903.

Gilles Paire / Shutterstock.com.

 FIGURE 2   Shock cord arrangement on Bleriot XI reproduction.

Early aircraft typically used a single tire per landing gear which worked reasonably while aircraft masses were relatively small (single tires per landing gear remains an appropriate configuration for lighter weight aircraft). As overall aircraft size and mass grew, the limits of reasonable single tire capacity were met. The peak of large aircraft on single wheel landing gears was reached with the XB-36, which first flew in 1946. With a maximum weight of around 280,000 pounds (127 000 kg), this aircraft used a single 110 inch (2.8 m) diameter main tire (Figure 4, left) and was only suitable for operation from highly reinforced concrete surfaces (the high point load exerted by

xx

Introduction

Kev Gregory / Shutterstock.com.

 FIGURE 3   Westland Lysander (left) and internally sprung wheel (right).

the single main tires would overload most lower strength ground surfaces). Later versions of the aircraft were designed using one of the first multiple wheel units to enter service. Tracked systems (Figure 4, right) were also tested on this aircraft. Arrangements of multiple smaller wheels improved the ability of ground surfaces to support higher loads and in many cases, stowage of a grouping of smaller wheels was more readily facilitated. The production B-36, Sud Aviation Caravelle, and de Havilland Comet all used four wheel main landing gears where each pair of wheels was fitted to a lever arm, with a mechanism joining the levers to a common shock absorber. Following this brief flirtation with paired wheels on levers, most multiple wheel (more than two) landing gears have mounted the wheels to a rigid bogie beam, pivoted at the bottom of a cantilevered shock strut. Early large high speed aircraft such as the Convair B-58 (Figure 5), Tupolev Tu-144, and Avro Vulcan utilized eight small diameter tires fitted in pairs to four braked wheels, mounted on a bogie beam. Advances in tire technology (as well as changing constraints on the required retracted position volume) have permitted a reduction in the total number of required tire positions; large aircraft today utilize multiple wheels on bogie beams almost exclusively (the rare exceptions being certain military transport aircraft). An early task for the aircraft and landing gear designer is to find the best configuration of tires (size, pressure, and position) to meet the required aircraft performance goals, ensure  FIGURE 4   Convair XB-36 main wheel (left); XB-36 with tracked landing

© SAE International.

gears (right).

Introduction

xxi

© SAE International.

 FIGURE 5   Convair B-58 main landing gear.

appropriate compatibility with the intended airfields, all while minimizing weight, cost, and complexity. As was shown for the B-36 in Figure 4, attempts have been made to use alternative ground interfaces to pneumatic tires: skids, tracks, and air cushions have been explored with limited success. Caterpillar track designs were trialed on a number of aircraft, including the P-40, A-20, C-82, B-36 (Figure 6), and B-50 (Figure 7) with the prime intention to lower the ground contact pressure through a significant increase in the ground contact area. While the feasibility of tracks as alternatives to tires was demonstrated, development was halted – the weight and mechanical complexity of the track suspension and guidance mechanisms as well as increased maintenance and exposure to snow, ice, and debris rendered the tracks sub-optimal when compared to pneumatic tires. However, significant elements of tracked based systems are otherwise similar to conventional aircraft wheels and brakes. As can be seen in the figures,

xxii

Introduction

© SAE International.

 FIGURE 6   Track main landing gear of XB-36.

© SAE International.

 FIGURE 7   Track main landing gear of B-50.

Introduction

xxiii

© SAE International.

 FIGURE 8   Air cushion landing system on XC-8A.

there are a large number of load bearing wheels in track systems, each with their own bearings; wheels in tracked systems do not need to withstand tire inflation pressures, however. Wheel brakes are also applied in track systems. Air cushion landing systems (effectively the marriage of the aircraft and the hovercraft) have been developed and demonstrated, as shown in Figure 8. While an air cushion landing system permits aircraft alighting on virtually any surface (including those with low/no surface strength), power is required to operate the system and the low friction created by the film of blown air results in challenges for directional control and braking, especially with the aircraft at low speed or when stationary. Actuatable braking mechanisms (Figure 9, left) and no-power parking mechanisms (Figure 9, right) have been proposed but not actively pursued. Most wheel braking systems rely on friction as the means of retardation, and Chapter 1 outlines some of the historical developments in frictional wheel braking. As with other approaches for landing systems, alternatives to friction for wheel braking have been investigated but not widely adopted, including electromagnetic

© SAE International.

© SAE International.

 FIGURE 9   Potential air cushion landing system braking and parking mechanisms.

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Introduction

and fluidic approaches to energy absorption. Figure 10 shows two views of a brake which used the wheel to spin an air compressor (through a gearbox) in order to dissipate energy by doing work on the ambient air. As the effectiveness of this device reduces with decreasing speed, a conventional friction brake was installed in parallel. Other means of deceleration have been trialed, such as the retro-rocket installation of the XFC-130H/YMC-130H developed under the project Credible Sport. Despite these forays into wheel-less solutions and braking solutions, the pneumatic tire with a friction wheel brake remains the dominant configuration for aircraft today. As is discussed further in the following chapters, the aircraft wheel and friction wheel brake have evolved into lightweight, long-lasting, and highly reliable pieces of aircraft equipment.

Reprinted from US Patent 3,142,360.

© SAE International.

 FIGURE 10   Turbo brake.

About this Book

T

his book is designed to guide the interested reader through the key principles of aircraft wheel, brake, and brake control system design. Additionally, references to further information are provided when it is available. This book is excerpted from the author’s two volume treatise, The Design of Aircraft Landing Gear [2], which provides details on the entirety of landing system design. Much of the beauty of interesting design problems such as aircraft landing gear is that any one subject could fill an entire book, whether it be the tribology of wearing surfaces, the interaction of gas and oil in shock absorbers, or the kinematic arrangement and analysis of mechanisms. It is therefore impossible to provide every last detail on every problem which the landing system engineer may face, but an effort has been made to tackle many of the subjects one is likely to face when designing new products or supporting the operation of systems in service. This particular volume outlines a key element of aircraft design and landing gear design: how to consider, select, and design wheels, brakes, and brake control systems. In the author’s experience many problems that must be confronted have already been addressed by others in the past, but the information is not widely known or shared, leading to the observation that there are few new problems, but many new people. This book is intended to share much of the information available and provide avenues for further exploration. A career in landing system design, development, and support can be spent while continually facing fresh and interesting challenges. No two aircraft are exactly alike; regulators and customers are always increasing their expectations, elevating the design challenge and making landing system engineering an exciting and rewarding discipline.

©2022 SAE International

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About this Book

The book commences in Chapter 1 with an overview of various types of aircraft wheel brakes before broadening the topic in Chapter 2 with a discussion of aircraft deceleration, including thrust reversers, braking parachutes, and arresting systems. Chapter 3 focusses on brake energy and torque sizing, including worked examples. Design considerations for wheel brakes as well as means to perform the initial design of carbon brakes, including mass estimation are provided in Chapter 4. Recommended practices for wheel and brake design as well as the current certification rules are discussed in Chapter 5. Chapter 6 works through common brake issues and concerns, such as vibration, wear, and oxidation, while Chapter 7 outlines typical braking accessories such as cooling fans, brake temperature measurement system, and functions such as retraction braking. The design of wheels is the focus of Chapter 8, including bearing selection, over temperature and pressure protection, failure modes, and a means for wheel mass estimation. The book concludes with Chapter 9, which outlines typical strategies for brake control and anti-skid braking. The specific names used for various components of the landing gear vary depending on geographic location as well as company history. A consistent set of terms is used throughout this book but as an aid to comprehension, the various components are identified, and their commonly used names indicated as an aid to the reader. Further terminology is explained in document AIR1489 [3]. The common names for a variety of landing gear components that occur in this book are shown in Figure 1.

© SAE International.

 FIGURE 1   C-160 Transall main landing gear (cutaway).

1 Brakes

A

ircraft wheel brakes (Figure 1.1) are installed on all land-based aircraft to provide a number of functions: absorption of the horizontal kinetic energy of the aircraft during landing, absorption of horizontal kinetic energy during a rejected takeoff, holding the aircraft static when parked or during engine testing, and maneuvering the aircraft on the ground (stopping and turning by differential braking). A variety of brake types have been created during the development of aircraft, many of the historical types following similar developments in road vehicles. In general, the wheel brake works by forcing one material that is fixed against rotation against another material that is mechanically arranged to rotate with the wheel. The contact of the two materials generates a frictional force which is proportional to the applied contact force. This frictional force acts at the radius where the clamping force is applied, resulting in a torque about the wheel axle. The wheel brakes are the primary actuators in a safety critical system – they must have appropriate torque capability, energy capacity, and response bandwidth to integrate correctly with the aircraft. A historical example of a wheel brake is the expander tube brake, shown in Figure 1.2. This type of brake uses hydraulic or pneumatic pressure to inflate a tube, which pushes brake blocks radially outward against a brake drum. Many aircraft from the 1940s used expander tube brakes. More recent brake designs are hydraulically operated disk type brakes, either with pads clamping one disk or a stack of disks in a multiple rotor/stator arrangement. The diameter and number of disks required is a function of the torque required whereas the thickness and mass of the disks is a function of the energy that must be stored in the brake; heat generated at the interface of the pad and the disk accumulates predominately in the disk. Figure 1.3 shows small aircraft brakes - on the

©2022 SAE International

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2

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Shutterstock.

 FIGURE 1.1   Cutaway of wheel and brake.

left, the brake used on a single piston engine aircraft and on the right, the brake fitted to a Beech King Air C90, a twin turboprop aircraft with a maximum weight of approximately 4400 kg. These brakes use a steel disk with brake linings made from a variety of sintered metallic and organic materials and are similar to the disk brakes and pads used on passenger cars. As can be seen in the figure, one brake has a caliper with two pistons whereas the other brake has two calipers, each with two pistons. The addition of more pistons (and larger piston diameter) increases the clamping force available for a given hydraulic pressure. This results in greater available braking torque. Multiple disk brakes are used on large aircraft where high energies and high brake torques are required. The arrangement and nomenclature of a typical multiple disk brake is shown in Figure 1.4. Brakes of this form are predominant in large aircraft – significant amounts of energy can be stored in the “heat pack”: the collection of stators and rotors that make up the brake. The stators are disks that are slotted or keyed on their inner diameter such that they may slide, but not rotate, on a torque tube. The torque tube then connects with tie bolts to the piston housing, and the torque may then be taken out from the brake in a number of ways. Figure 1.4 shows that it is taken out to the landing gear structure through a lug or torque arm. The rotors are slotted or keyed on their outer diameter to interface with torque bars (drive keys) in the wheel. The rotors rotate with the wheel and transfer the braking torque to the wheel structure. A variety of materials have been used to store the brake energy as well as to provide the friction surface. In single disk brakes, steel is a common choice. A cross section of a steel brake is shown in Figure 1.5. In a steel brake, the steel forms the structure of the rotors and the stators. Individual lining blocks are attached (riveted, typically) to both sides of the rotors and/or the stators. It is common for metal-ceramic lining blocks on the rotor or stator to run against steel lining blocks on the adjacent disc. An example of how these individual linings are arranged is shown in Figure 1.6. An alternative to steel as the material for the rotors and stators was beryllium, which offered the ability to store a significant amount of energy at a much lower weight than steel. However, due to the toxicity of beryllium, it is unlikely future brakes will use this material.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

3

Reprinted from US Navy NAVEDTRA 14329.

 FIGURE 1.2   Expander tube brake arrangement.

A lightweight alternative to steel brakes is the use of reinforced carbon–carbon as the structural and friction material for both rotors and stators. Carbon brakes offer a number of advantages over steel brakes – lighter weight for the same life and energy requirements, monolithic construction of the stators and rotors (the heat absorbing material is the friction material), the ability to remain functional at very high temperatures (where steel brakes can warp, suffer lining damage, or melt), and typically longer life than an equivalent steel brake. The cross section of a typical carbon brake is shown in Figure 1.7. The stators and rotors have a form similar to that shown in Figure 1.8.

4

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 1.3   Examples of single disk brakes.

© SAE International.

 FIGURE 1.4   Hydraulic, multiple disk brake nomenclature.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 1.5   Steel brake cross section.

Reprinted from Midas Touch, Staff Seargent Chad Chisholm, USAF.

 FIGURE 1.6   Steel brake linings on the stator of a KC-135 brake.

5

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 1.7   Carbon brake cross section.

© SAE International.

 FIGURE 1.8   Reinforced carbon–carbon brake disk.

A relatively recent development in aircraft wheel brakes is the replacement of the hydraulic actuation function with electromechanical actuation. This brake configuration eliminates the risk of hydraulic leakage near hot brakes (reducing the risk of a brake fire), can improve dispatch reliability and brake dynamic performance, and provides a brake solution for aircraft without hydraulic systems. Initially introduced

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

7

© SAE International.

 FIGURE 1.9   Electric brake from Boeing 787.

on military unmanned vehicles, electric brakes are in airline service on the Boeing 787 and the Airbus A220 aircraft. An example of an electric brake is shown in Figure 1.9. On most aircraft, the wheel brakes are fitted only to the main landing gears. As the main landing gears bear the majority of the weight of the aircraft, there is significantly more drag force available to be generated by braking these wheels. However, some aircraft do fit brakes to the nose landing gears. On early jet aircraft, where landing speeds were high and runway lengths short (as they had been designed for piston engine/propeller transports), brakes were fitted to the nose wheels. The Boeing 727 and the Convair CV880 and CV990 are indicative examples. Some tactical aircraft designed to operate from austere or battle damaged runways may be equipped with nose landing gear mounted wheel brakes. Many historical Russian tactical aircraft, such as the MiG-21, were equipped with nose wheel brakes. The Saab Gripen is also fitted with nose wheel brakes to ensure the maximum possible deceleration.

2 Aircraft Deceleration

T

he wheel brakes are not the only source of deceleration of the aircraft on the ground although they are typically the most powerful as well as the most controllable (however, at landing speeds aerodynamic drag may be higher than the available wheel brake force). Airframe drag is a source of deceleration – Figure 2.1 shows a B-1B landing and rolling out in a high drag configuration. Spoilers on top of the wing have been deployed (these are typically deployed once the aircraft senses it is on the ground) that increase the airframe drag and decrease the wing lift – ensuring that the weight of the aircraft is carried by the wheels and making braking more effective. Most jet-powered civil transport aircraft (and some specialized tactical aircraft) include thrust reversers (as shown in Figure 2.2) or reverse pitch on propellers that permit effective retardation at high speeds. Certification regulations for large civil transport aircraft limit the cases where credit can be taken for reverse thrust to landings and rejected takeoffs on wet or contaminated runways1. In this case, although reverse thrust may be used on a regular basis on dry runways, the brakes must be sized to absorb the landing and rejected takeoff energy in the absence of reverse thrust. Improvements in the reliability and availability of reverse thrust systems in the future may introduce the possibility to take credit for reduced stopping distances on dry runways. Military aircraft may rely on reverse thrust for mission performance, and many civilian operators balance the fuel cost of utilizing reverse thrust (and reducing brake wear) against the cost of brake replacements.   14 CFR 25.125 Landing and 14 CFR 25.109 Accelerate-stop distance.

1

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Reprinted from http://www.af.mil/News/Photos/ igphoto/2000672673/.

 FIGURE 2.1   B-1B Lancer braking during landing.

Fasttailwind / Shutterstock.com.

 FIGURE 2.2   Il-62 landing showing thrust reversers deployed.

While not a feature on modern civil aircraft, braking parachutes (Figure 2.3) may be used on tactical and strategic military aircraft to reduce stopping distance, especially on wet or contaminated runways and also on other aerospace vehicles such as the Space Shuttle Orbiter. Modern antiskid braking systems and thrust reverse systems have replaced these systems on civil aircraft as there is a significant cost associated with capturing, repacking, and stowing the parachute system after each use. However, in certain specialty applications, braking parachutes continue to be fitted – Norway requires them on tactical aircraft (such as the F-16 and F-35) to ensure appropriate stopping distances on snow-covered runways. In some cases, braking parachutes can be  used to reduce the required brake sizes, which can be  advantageous on small tactical aircraft where wheel stowage volume is a significant concern.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

11

Fasttailwind / Shutterstock.com.

 FIGURE 2.3   Tu-160 strategic bomber decelerating with braking parachutes.

The advantage of a braking parachute is that it generates its maximum deceleration force at high speed while wheel brakes offer a relatively constant source of deceleration. Reducing the aircraft velocity from high landing speeds using a parachute can significantly reduce the amount of energy that needs to be stored in the wheel brakes. The typical arrangement of a braking parachute is shown in Figure 2.4. A disconnection fitting is provided so that the pilot may select when to drop the parachute after use. Typically, a frangible link is included to ensure that the parachute will separate automatically if the drag load created is too high (for instance, during inadvertent deployment in flight). A stable parachute design with low opening force and low oscillation is desired. Typically, ribbon, ringslot, and cross-design parachutes are used. The parachute is usually sized to have a drag area which is 25%–50% of the aircraft wing area (the lower ratio used for bomber aircraft and the higher ratio used for tactical aircraft) [4]. The drag area of a parachute is the surface area of the parachute multiplied by the drag coefficient (CdS). Typical drag coefficients for ringslot and ribbon parachutes vary between 0.5 and 0.65. Both of these parachute types have

Reprinted from Parachute Recovery Systems Design Manual.

 FIGURE 2.4   Typical braking parachute arrangement.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

low opening shock factors of around 1.05 times the steady-state drag force. The drag force created by a parachute once opened and stable can be calculated as:

Fp

where:

1 V 2 Cd S 2

•• Fp is the drag force generate by an open, stable parachute •• ρ is the air density •• V is the velocity of the aircraft (parachute) •• Cd is the drag coefficient of the parachute •• S is the surface area of the parachute Determining the effective stopping distance also requires an estimation of the opening time of the parachute. ESDU data sheet 09012 [5] provides detailed methods for all parachute types as well as a method for calculating opening times based on the selected parachute type. Aircraft carrier-based airplanes are fitted with arresting hooks that are used for every landing aboard the carrier. In addition, many land-based tactical aircraft are fitted with arresting hooks designed for limited use. These arresting hooks engage with fixed or portable arresting gear (a cable stretched across the runway, connected to straps which unroll from braked drums). Figure 2.5 shows an F-16 trailing the arrestor cable following an arrested landing. These land based arrestor cables are

Reprinted from http://www.af.mil/News/Article-Display/Article/124050/ give-me-a-brake/.

 FIGURE 2.5   F-16 decelerating with arresting system.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

13

typical only for military aircraft and are used in the case of system failures such as braking hydraulic system loss. Commercial aircraft may have to operate at runways fitted with these systems, in which case the only concerns are the ground clearance and the bump loads created when traversing the cable. Chapter 2, in The Design of Aircraft Landing Gear, addresses the size and behavior of the cable when trampled, (see also Airfield Compatibility: Key Principles for Landing Gear Design). A guide for the design of arrestor hook systems for land-based aircraft is provided in document ARP1538 [6]; this guide provides the hook shape and recommendations for the installation. It is recommended that when the specific arresting system performance is unknown then the properties of the BAK-13 system be used. The maximum hook load as a function of engagement speed and aircraft weight is shown in Figure 2.6. Arrestor hook design for US Navy Carrier aircraft is governed by MIL-STD-18717 [7]. Alternatives to wheel brakes have been attempted, such as the “Dowty Drag Brake” – shown in Figure 2.7. This system employed a pad of frictional material that was pushed hydraulically against the runway surface to provide retardation. While effective in principle, only limited usability of the system was achieved, despite over 200 different friction materials having been tested [8]. While the drag brake system offers benefits in terms of energy dissipation (the energy is transferred to the ground surface rather than being stored in the brake), the available stopping force is governed by the dynamic coefficient of friction of the braking material. This tends to be lower than the available coefficient of friction from a tire when operated at limited slip (to achieve the peak coefficient of friction), making the drag brake less effective at stopping than a wheel brake and tire arrangement.

© SAE International.

 FIGURE 2.6   BAK-13 arresting system maximum hook load.

14

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Reprinted with permission from © Safran Landing Systems.

Reprinted from US Patent 2877969A.

 FIGURE 2.7   Dowty drag brake.

3 Brake Sizing

D

etermination of the appropriate size and quantity of brakes is an important step. In most cases, selecting the number of brakes is as simple as choosing all main landing gear wheels to have brakes. Aircraft configurations exist where not every main wheel is braked (such as the A380) and others exist where pairs of wheels were connected to a common brake (such as the XB-70), but these are exceptions. Important considerations in brake selection include the maximum energy to be absorbed and the braking torque to be generated. Energy requirements will size the mass of material required to store the kinetic energy as heat while torque requirements will determine the stopping distance, number of rotors and stators, as well the ability to hold against engine thrust and slopes. Additional material is added to the minimum requirements to achieve the desired brake wear life.

Energy During stops following landing or during a stop following a rejected takeoff, the brakes absorb most of the kinetic energy of the aircraft. Many brake sizing cases exist depending on the conditions of the landing or the rejected takeoff. Often, the rejected takeoff case with the brake completely worn determines the minimum amount of material permitted in the brake (it is representative of an “end of life” brake). However, there are rare cases where the aircraft in an abnormal or emergency condition can have a higher energy than the rejected takeoff design baseline energy. Table 3.1 shows a number of cases with their approximate brake energies compared to the maximum energy rejected takeoff value. ©2022 SAE International

15

16

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

TABLE 3.1  Brake energy cases [9]. Service landing

Design landing

Overweight landing

Flapless landing

Rejected takeoff

15%–20%

30%–50%

100%–110%

90%–120%

100%

A service landing is the typical landing case: reverse thrust may be employed as well as a long rollout (maximizing aerodynamic energy dissipation); in this landing type, the brake energy requirements are minimized. The overweight landing case and flapless landing cases are aircraft failure cases – in the overweight case, a failure immediately after takeoff necessitates an immediate landing with generally less than 1% of the fuel onboard being consumed. A flapless landing occurs at a significantly higher than normal touchdown speed – due to the failure of the flap system. In many cases, an aircraft will be required to perform its design landing stop, taxi-in, refuel and reload, then taxi-out, and perform an RTO, with the only brake cooling period being the refueling and reloading time. US Navy specification SD-24 [10] requires such a turnaround performance, with a 15-minute stop period, 2-mile taxi lengths, and the possibility to perform an RTO at the aircraft maximum takeoff mass. Many civil aircraft may include brake temperature indication systems to indicate to the pilots whether the brake temperatures are sufficiently low to permit departure and the possibility of a rejected takeoff.

Kinetic Energy Calculation The simplest sizing assumption is to consider that all of the kinetic energy of the aircraft is converted to heat in the brakes. The kinetic energy is given by:

Ek

1 mV 2 2

where:

•• Ek is the kinetic energy •• m is the mass of the aircraft •• V is the ground speed of the aircraft at time brakes are first applied In SI units, the value of Ek is given in Joules (equivalent to Newton-meters) where the mass is in kilograms and the speed is in meters per second. In US customary units, Ek is typically calculated in foot-pounds; with velocity in feet per second, the mass must be provided in slugs (one slug weighs 32.174 pounds – in earth’s gravity). The historical certification regulation for light aircraft, Part 23, permitted either a conservative, rational approach to energy estimation or the use of the kinetic energy formula above with the speed at landing equal to the power off stall speed and the speed during a rejected takeoff equal to the maximum value of V1 (the decision speed where a choice must be made to either reject or continue a takeoff). The relevant maximum landing weight and maximum takeoff weight are used.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

17

Light Aircraft Example The de Havilland DHC-6 Twin Otter is a twin turboprop aircraft with one nose wheel and two braked main wheels. It has a maximum takeoff weight of 11,579 pounds and a maximum landing weight of 11,400 pounds. The decision speed (V1) is 72 knots, and the stalling speed at maximum landing weight, with the aircraft in the landing configuration (full flaps), is 54 knots. Calculate the required brake energies for the two cases:

•• Landing energy:

Convert 54 knots to ft / s : 54 knots 1.69 ft / s per knot 91.3 ft / s Convert 11,400 pounds to slugs : 11,400 pounds 32.174 ft / s2

354.3 slugs

Calculate kinetic energy:



Ek



Ek

1 mV 2 2

1 2 354.4 91.3 2

1 354.4 8,335.7 2 E k 1,477,084 ft-lb



Ek



•• Rejected takeoff energy:

Convert 72 knots to ft / s : 72 knots 1.69 ft / s per knot 121.7 ft / s Convert 11,579 pounds to slugs : 11,579 pounds 32.174 ft / s2

359.9 slugs

Calculate kinetic energy:



Ek



Ek



Ek

1 mV 2 2

1 2 359.9 121.7 2 2,665,220 ft-lb

The most critical energy case is the rejected takeoff case (due to both a higher mass and higher velocity). As the aircraft has two brakes, the individual brake energy is:



Ebrake

2,665,220 1,327,610 ft-lb 2

In general, aircraft which have rejected takeoff requirements (typically fixed wing aircraft) will have their brake energy determined by the mass and speed during the rejected takeoff. The trend is even more pronounced on large commercial aircraft where the landing weight is significantly less than the maximum takeoff weight.

18

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Large Aircraft Example A large transport category passenger aircraft has one nose landing gear with two unbraked wheels and two main landing gears, each with four braked wheels. The maximum takeoff mass is 275 000 kg. If the decision speed at that mass and in the worst case “hot and high” takeoff is 180 knots, calculate the required energy per brake:



Convert 180 knots to m / s : 180 knots 0.514 m / s per knot 92.5 m / s

Calculate kinetic energy:



Ek



Ek



Ek

1 mV 2 2

1 2 275000 92.5 2 1 176 484 375 Joules

As the aircraft has eight brakes, the energy per brake is:



Ebrake

1 176 484 375 147 060 547 Joules 8

Considering that the entire kinetic energy of the aircraft is absorbed by the wheel brakes is a conservative assumption. Some of the kinetic energy is dissipated in other ways, and as a result, a more refined calculation of brake energy can be conducted.

Rational Brake Energy Calculation In contrast to the kinetic energy calculation, a rational method for brake energy estimation involves attempting to consider all sources of energy input and removal during the braking event. These include:

•• Kinetic energy of the aircraft at the initiation of the braking event •• Residual thrust from the propulsion system (reverse thrust if applicable) •• Airframe drag •• Tire rolling resistance and wheel bearing rolling resistance •• Potential energy change due to runway slope •• Wheel braking, including the effects of wing lift that may limit the available retardation force

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

19

•• Atmospheric conditions (pressure, wind) •• If allowable, retardation force from reverse thrust or braking parachutes As the determination of the brake energy results from an integration of these parameters, solutions are typically determined by numerical integration of the parameters. A method using a stepwise integration at a time interval of 0.25 seconds is developed in ASD-TR-68-56 [11]. A more refined approach considering more complex models for tire and bearing rolling resistance is developed by Dowden [12]. Accurate estimations of the brake energy require knowledge of the lift and drag behavior of the aircraft as well as the change in lift and drag coefficients during the braking event. A simplified assessment conducted for the large passenger aircraft example given previously indicates that rolling resistance in this case (on eight radial tires) provides approximately 4% of the energy absorption. The analysis considered that:

•• the full weight of the aircraft is borne by the landing gears (reasonably accurate in cases where spoilers deploy)

•• no significant residual thrust and no contribution from reverse thrust •• no runway slope •• constant braking force equal to a third of a g deceleration •• rolling resistance for each main wheel tire equal to 0.015 •• 95% of aircraft weight borne by the main landing gears •• An airframe drag coefficient of 0.04 and a representative area of 443 m2 Aerodynamic drag is proportional to the square of the velocity – as this braking event begins at 180 knots, the initial contribution to deceleration from aerodynamic drag is significant – approximately 10% of the brake generated force. The end result is a stopping distance of 1210 meters and energy per brake of 135.8 MJ. This is an 8% reduction in the required braking energy compared to the kinetic energy approach which for large aircraft and large brakes can be significant in terms of the size and mass of the brake assembly. The same calculation performed for the Twin Otter rejected takeoff example provides a required brake energy which is 5% less than the value computed through the kinetic energy approach. In this case, the difference is only 72,460 foot pounds, which is not likely to change the design of the brake. The rational analysis technique depends strongly on the rate of deceleration resulting from the brake torque: more brake torque generated results in shorter stopping distances and less contribution from the other sources of energy dissipation (and more energy stored in the brakes). For the certification of civil large transport aircraft, a minimum average deceleration rate [13] of 3.1 m/s2 is required for design landings and 1.8 m/s2 is required for a rejected takeoff. Aircraft designers may seek higher performance than these requirements.

20

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Torque The torque delivered by a brake will vary, not only as a function of the applied clamping pressure, but with wheel speed and brake temperature. It is important to ensure that the brake will provide a minimum torque to achieve the desired aircraft performance. The required minimum average torque can be computed from the required deceleration rate (under the assumption that adequate friction is available between the tire and the ground). For instance, to decelerate a 275 000 kg aircraft at an average of 3.1 m/s2 requires a braking force (assuming there is no contribution to stopping from other sources) of:

F



275 000 3.1

F



mA

F

852 500 N

If it is considered that eight 1400×530R23 tires have already been selected to meet the ground compatibility requirements, then the loaded tire radius of each tire can be approximated as the static loaded radius (available in tire catalogs as well as in Chapter 3 in The Design of Aircraft Landing Gear, and Aircraft Tires: Key Principles for Landing Gear Design) and is approximately 23.5 inches (0.6 m). A more accurate calculation of the effective radius during braking would utilize the tire load-deflection curve to ensure the appropriate radius for the applied vertical load (considering weight transfer to the nose wheels and reduced ground reaction on the main wheels) as well as accounting for the distortion resulting from the applied drag load. Assuming all eight wheels are braked, then the drag force to be generated at each tire contact patch is:

F

852 500 106 563 N 8

The brake torque required to generate that drag force is:

F r



106 563 0.6



63 938 N m

Assuming approximately 95% of the aircraft mass on the main landing gears, then the vertical load on each wheel is 320 225 Newtons. The ratio of the individual wheel drag force to the vertical load is then:

106 563 0.33 320 225

which represents the minimum tire to ground friction coefficient in order that achievement of this deceleration rate is not limited by tire friction characteristics. A value of 0.45 is generally considered achievable on dry pavements, so eight brakes capable of generating this average torque would meet the performance minimum requirement. Calculation of the minimum average torque during a rejected takeoff can be performed in a similar fashion. For this hypothetical aircraft, an average brake

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

21

torque of 37 125 N∙m would be the minimum required to meet the rejected takeoff performance criterion. While dynamic torque values are directly related to stopping distance and deceleration performance, the ability of the brake to maintain a static torque can be of importance. The brakes must typically be adequate to restrain the aircraft against an engine test at full thrust. For the large passenger aircraft discussed previously, two high bypass turbofans each producing 375 000 N of thrust are provided. The resultant force on the aircraft is 750 000 N – which is slightly less than the deceleration force required to achieve the 3.1 m/s2 requirement. A static brake torque of 56 250 N∙m is needed on all eight brakes to hold the aircraft against this type of test. Some aircraft manufacturers may choose to perform this type of test with the aircraft restrained mechanically as opposed to ensuring the brakes have adequate torque capability. In some cases, the engine thrust may exceed the static friction force available from the tires, in which case there is no point in designing the brakes to achieve this case – the tires will simply slide along the ground. Over-specifying static torque can lead to high torque gain in the brake, which can make antiskid tuning difficult. When aircraft are parked and power is turned off, a parking brake is usually fitted to ensure that clamping force is maintained on the brakes (either by maintaining pressure to hydraulic brakes or by locking electric brake actuators in position). The available clamping pressure from a parking brake may vary with time (if the energy source is a hydraulic accumulator) and is typically not the full clamping pressure available during service use of the brakes; additionally, while applying the parking brake when the brake is hot is generally discouraged to avoid damage to the brake, parking brakes applied to a warm brake can experience a change in clamping pressure as the brake cools, especially on an electrically actuated brake. The parking brake must be capable of generating sufficient brake torque to resist the tendency of the aircraft to roll away when parked on a slope and this case can often be a driving design factor for rotorcraft due to the high slope requirements. The certification regulations for civil rotorcraft require the brakes to hold the aircraft on a 10° slope; many rotorcraft manufacturers extend this to 12°. Figure 3.1 shows a helicopter parked on a slope.

Reprinted from NASA.gov.

 FIGURE 3.1   Helicopter park brake force requirement.

22

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

The required braking force is F = W sin θ. From the required brake force, the number of brakes, and the tire radius, the required brake torque can be calculated. During the testing and certification of brakes, the brake structural torque must be demonstrated. This is not a torque that can necessarily be generated by the clamping of the friction materials, but rather a torque value used to demonstrate the strength safety margin of the brake, brake components, and the attachment to the wheel. Certification of civil aircraft and helicopter brakes requires compliance to AS5714 [14] or TSO-C135a [15], both of which specify a test to a structural torque of 20% above the vertical static load times the rated tire’s radius at that load for landing gear with a single wheel per leg and 44% above that value for landing gear with more than one wheel per leg. The wheel and brake must withstand these loads without failure for at least 3 seconds.

4 Brake Design

T

he principal function of the brake is to generate a torque on the wheel to decelerate the aircraft; it is the primary device achieving retardation in many safety critical cases. All brakes in use on aircraft today convert the kinetic energy of the airframe into heat through friction between the brake components. The ability to store this energy as heat without malfunction is a key design driver for the brake. Whether the brake is a single steel disk or multiple disk carbon brake, there must be enough thermal mass to absorb the energy without rising to a temperature that damages the brake (or landing gear or starts a hydraulic or tire fire). The relationship between temperature rise, mass, and energy is material specific and given by the material’s specific heat, c, as shown:

where:

E c m

T

•• E is the energy to be absorbed in the brake •• c is the specific heat of the material used in the brake’s heat sink •• m is the mass of material in the brake’s heat sink •• ΔT is the temperature rise of the heat sink material This formulation assumes that all of the material is at the same temperature, which will not be true during braking where the heat is generated at the friction interface. However, the heat input to the brake will conduct through the components and at some time (depending on the thermal conductivity) after braking a steady-state value will be achieved. For initial design and comparison purposes, this formulation ©2022 SAE International

23

24

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

TABLE 4.1  Properties of common brake materials.

Material

Specific heat (J/g-°K)

Density (g/cm3)

Melting point (°C)

Thermal conductivity (W/m-°K)

Tensile strength (MPa)

Steel

0.49

8.05

1370–1540

35–60

410

Copper

0.385

8.96

1083

350–385

240

Beryllium

1.82

1.85

1283

218

255

CarbonCarbon

1.42

1.6–2

3000–3700 (sublimation)

20–160*

66

* Carbon–Carbon materials are strongly anisotropic; properties vary depending on the fiber construction, orientation, type of densification employed, and degree of graphitization/ heat treating of the matrix.

is reasonable. The specific heat of a material depends strongly on the temperature of the material. As a result, only estimates of brake performance can be made when using a constant value of specific heat. Table 4.1 shows the properties of common brake materials. The optimum material would maintain its strength at high temperature (this property is approximately proportional to its melting point), would have a high specific heat (to minimize the mass required), and a high density (to minimize the space required to store the energy). By observing the values in Table 4.1, it is clear that the lightest material to absorb a given amount of energy is Beryllium, provided the target temperature for each material is the same. However, different materials respond differently with increasing temperature. Figure 4.1 shows the reduction in strength for each material at elevated temperatures. Only carbon–carbon composite retains its strength to exceptionally high temperatures. However, carbon begins to oxidize at high temperatures, so a design trade must be made between carbon heat sink life and weight – a lighter heat sink will operate at higher temperatures, leading to oxidation and reduced life (independent of wear).

 FIGURE 4.1   Variation of material strength with temperature.

Variaon of Strength with Temperature

Percentage of Room Temperature Strength

120 100 80

Steel Beryllium Copper Carbon-Carbon

40 20 0 0

1000

2000 Temperature ( °C)

3000

© SAE International.

60

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

25

Steel brakes are typically designed [16] to a maximum temperature of around 500°C for service landings, 780°C for design landings, and 1100°C for maximum energy rejected takeoffs. A steel brake is unlikely to be serviceable after the rejected takeoff stop if designed to this temperature. A beryllium brake may be designed [17] to a service temperature of approximately 530°C and rejected takeoff temperature of 840°C. Carbon brake temperature targets are related to the expected life as time at elevated temperature leads to oxidation of the carbon and loss of mass and strength. A service landing could be designed for a temperature of 300°C–400°C while a design landing value could be in the range of 600°C–700°C. A rejected takeoff could target a temperature of 1800°C–2000°C. The maximum allowable temperature range in carbon–carbon may be more limited by the surrounding wheel, brake, and landing gear structure and the ability to protect nearby metal components from the intense heat.

As an example, the rejected takeoff energy for the large passenger aircraft considered earlier is calculated to determine the required mass and volume of material:

•• For the carbon material with a maximum allowable temperature of 1800°C:

E

T

c m

•• Rearranging to solve for mass:

m

E c

T



•• Considering the specific heat value of 1.42 J/g °K, brake energy of 135.8 MJ, and a starting temperature of the brake of 50°C, then:



m

135 800000 1.42 1750

m 54648 g 54.6 kg

Table 4.2 provides the resulting mass and volume for the different brake materials at the RTO energy of 135.8 MJ. The analysis in Table 4.2 is flawed due to the use of a constant value for specific heat. In practice, brake manufacturers use empirically derived values termed the “mass loading” to determine the required amount of material for a given energy. Each manufacturer will have different design target mass loadings depending on their particular material or combination of materials. Example mass loadings are shown in Table 4.3. In many cases, the full capability of the material is not used to maintain lower temperatures and permit a more rapid turnaround time for the aircraft. In the case of the carbon–carbon material, using the high end of the mass loading range may require long cooling periods or the use of brake cooling fans. Using the highest values for worn carbon–carbon and the largest value for steel brakes, the mass of material required to achieve the RTO with an energy of 135.8 MJ

26

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

TABLE 4.2  Mass and volume of various brake heat sink materials at RTO energy. Material

Maximum temperature (°C)

Mass (kg)

Volume (l)

Carbon

1800

54.6

30.6

Steel

1100

263.9

32.8

Beryllium

840

94.5

51.1

TABLE 4.3  Typical brake mass loading [18]. Copper brake/ Steel brake/ organic pad ceramic pad

Steel brake/ sintered pad

Carboncarbon brake (fully worn)

Carboncarbon brake (new)

Design landing mass loading (kJ/kg)

240

450

390

1250

900–1000

Rejected takeoff mass loading (kJ/kg)

540

750

900

2700–3200

2000–2400

is 42.4 kg of carbon or 150.9 kg of steel. The volume of carbon would be 23.6 liters while the volume of steel would be 18.7 liters. This sizing mechanism confirms the typical trend – that while steel brakes are heavier, they can occupy a smaller volume than a comparable carbon brake due to the higher density of steel. The values provided in the table are indicative only and should be confirmed with the brake vendor. However, they should serve for initial sizing purposes. New to worn masses can be estimated for carbon brakes from the mass loadings provided in Table 4.3; alternatively, typical wear rates are provided in Figure 4.2.

© SAE International.

 FIGURE 4.2   Wear rates for various brake materials.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

27

© SAE International.

 FIGURE 4.3   Carbon brake cross section.

A typical brake heat sink sizing approach is to ensure that in the fully worn condition there is sufficient mass to meet the rejected takeoff case and any other critical sizing cases. Additional material is then added (in the case of carbon–carbon) to address the required wear life (often termed landings per overhaul). In a multiple disk brake, the brake designer must select the number of disks, their outer diameter, inner diameter, and thickness, respecting a number of constraints. Figure 4.3 shows how the rotors and stators of the brake must fit within the wheel (indeed, the slots in the rotors must engage with torque bars attached to the inner diameter of the wheel). Heat shielding is often provided on the inner diameter of the wheel to protect the wheel material from the high brake heat. In addition, it is often advantageous to use a brake rotor outer diameter smaller than the maximum permissible in the wheel (if possible) to improve the natural cooling of the brake heat stack. The inner diameter of the rotors and stators must clear and engage (respectively) with the torque tube, which must be a ­sufficiently large diameter to clear the wheel structure and bearing supports. A method to estimate (for initial sizing purposes) the rotor inner and outer diameter for carbon brakes, as well as the number of rotor and stator drive keys, has been developed by Bailey [9], based on a regression of a number of example brakes and wheels. The resulting formulae are based on the desired wheel rim size (in inches) and provide results in inches. The values are based on a regression of a limited data set with rim diameters of 13 to 18 inches for military and 14 to 23 inches for civil brakes; caution should be employed when extrapolating beyond these diameters. Table 4.4 provides the equations for both military and civil aircraft. For rapid estimation purposes, the civil inner and outer diameter values are shown graphically in Figure 4.4.

28

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

TABLE 4.4  Carbon brake disk sizing estimation. Civil aircraft

Military aircraft

Stator outer diameter (inch)

0.7091D+2.286

0.7797D+0.5491

Stator inner diameter (inch)

0.417D+0.391

0.5279D–1.394

Number of stator drive keys

[1.442D–11.25]

[0.9419D–3.105]

Rotor outer diameter (inch)

0.788D+2.322

0.849D+0.6259

Rotor inner diameter (inch)

0.6645D–2.361

0.6954D–2.704

Number of rotor drive keys

[0.35D+2.217]

[0.4706D+0.706]

D is the wheel rim diameter in inches; size results in inches  FIGURE 4.4   Carbon brake disk size estimation – Civil applications.

Carbon Brake Disk Estimation For Civil Applications

20

15

10 Stator OD Stator ID

5

Rotor OD Rotor ID

0 11

13

15

17

19

Rim Diameter (in.)

21

23

© SAE International.

Carbon Disk Diameter (inch)

25

With the diameters of the brake material selected, the lining loading (or area loading) of the brake can be calculated. The lining loading is the energy to be absorbed by the brake divided by the total swept area of the brake (or the total pad area): .



LL

where:

E Aswept

•• LL is the lining loading •• E is the energy to be absorbed •• Aswept is the swept area of the brake: for carbon brakes, it is the rotor to

stator interface area multiplied by two times the number of rotor to stator interfaces

The range for lining loading for carbon brakes is up to 3000 J/cm 2 for design landings and up to 7000 J/cm2 for rejected takeoff cases.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

29

Lining power (or power loading) is the lining loading divided by the time, t, taken to stop the aircraft:

LP

LL t

The torque generated by a brake is dependent on a number of factors: the available coefficient of friction at the interface surfaces, the clamping pressure applied, and the mechanical losses due to friction arising when clamping the rotors and stators. The torque generated by the brake is calculated by:

2N

F r

where:

•• τ is the torque generated •• μ is the brake efficiency (a lumped term incorporating the coefficient of friction of the braking material and the mechanical losses within the brake)

•• N is the number of rotors •• F is the clamping force; for hydraulic brakes, the clamping force is given by the effective pressure times the total piston area:



F

P A

where:

•• ΔP is the effective pressure (the pressure applied to the brake less the ineffective pressure resulting from friction, return springs, and/or adjusting devices)

•• A is the total hydraulic piston area for the brake •• r is the radius where the combined friction of the surface acts; it is often

estimated as the average of the inner radius and the outer radius; this is often the radius where the electric or hydraulic actuators are placed to clamp the brake. A formula for this radius derived from an integration of the forces acting on the brake disks (assuming a constant coefficient of friction and consistent clamping pressure) is:

r

2 re3 ri3 3 re2 ri2



where:

•• re is the exterior radius of the friction contact area •• ri is the interior radius of the friction contact area

30

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

 FIGURE 4.5   Brake efficiency variation during a stop.

Brake Efficiency vs. Time 0.35 0.25 0.2 0.15 0.1 0.05 0

0

5

10

15 Time (s)

20

25

30

© SAE International.

Brake Efficiency

0.3

From these equations and knowledge of the expected brake efficiency, the size, number, and placement of clamping actuators can be determined. An increase in brake torque is most readily achieved by either increasing the clamping force or increasing the number of rotors. Increasing the number of rotors increases the number of wearing surfaces as well as decreases the brake efficiency (due to frictional forces on the stators and rotors as they slide within the wheel, the clamping force decreases across the stack – the rotors and stators furthest from the actuators see less clamping force than those nearest the actuators). Increasing the clamping force (by increasing the number of actuators or by increasing the piston diameter for hydraulic brakes) will increase the torque but also increases the gain of the brake (amount of torque delivered per unit of clamping force demanded), which can create issues with brake control. As in all design activities, choosing the optimum solution requires balancing all of the requirements. For carbon brakes, the average brake efficiency expected during a rejected takeoff is on the order of 0.2; for steel brakes with sintered pads, a minimum expected efficiency is on the order of 0.3. Different manufacturers with different materials and designs will achieve varying results. The brake efficiency varies during a stop – an example for a sample carbon brake is shown in Figure 4.5. Designing to ensure that the required torque is generated from the brake necessarily requires the use of the minimum expected brake efficiency. It is possible to have significant variation from the minimum value with cold brakes, at low speeds, at different ambient humidity levels, or on new brakes during their bedding-in process. Significant torque peaks are possible under these conditions.

Brake Actuation Hydraulic actuation is the standard form of brake actuation, although electric brake actuation is rapidly advancing, with electric brake actuation on both civil and military applications. The simplest form of hydraulic actuation is to provide one or more

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

31

pistons in cylinders with the pistons acting to clamp the brake materials. On large brakes, a number of pistons are arranged radially around the piston housing. These pistons can be single arrangements (as shown in Figure 4.10) or dual cavity arrangements, where the pistons are provided in pairs around the piston housing (as shown in Figure 4.11). In the dual cavity arrangement, the adjacent pistons are hydraulically independent; two separate hydraulic brake control systems power each brake (providing complete redundancy in the event of any single hydraulic or brake control system failure). Given the significant amount of piston displacement necessary to operate a brake between the new condition and the fully worn condition, an adjusting mechanism is employed to ensure that the gap between the piston and the heat sink remains small (a large gap would lead to an unacceptable delay prior to the application of braking). Depending on the size of the brake, the adjuster(s) may be a separate component on the brake or they may be integrated into the pistons. On large brakes having piston diameters of 32 mm or more, the adjuster is typically integral to the piston (it is difficult to fit an adjuster in pistons smaller than this diameter). An example of a brake piston with integral adjuster is shown in Figure 4.6. Automatic adjusters provide for a constant running clearance when the brake is released. When pressurized, the piston moves toward the brake, causing the adjuster spring to be compressed. An amount of lost motion is provided for by having clearance before the movement pushes on a rod. Further movement of the piston will drive the rod to move along a tube; only brake wear results in the rod moving along the tube. Movement is permitted in one direction only (compressing the heat sink) and restrained in the retraction direction. One method of providing this one-way motion is a friction bush that is arranged in a tightly fitting tube; under certain conditions, it will advance in the tube and close the gap between the piston and the heat sink. As it operates based on close tolerances and friction, it is not necessarily reliable. A more

© SAE International.

 FIGURE 4.6   Cross section of self-adjusting brake piston.

32

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Reprinted from US Patent 6,684,983 B1.

 FIGURE 4.7   Auto-adjusting brake piston.

reliable system that is widely employed uses plastic deformation of a tube to control the length change. A rod with a protrusion (often a split ring) is drawn through a tube; the tube plastically deforms in a controllable manner and gives reliable service. The principle of operation is similar to the way a pop rivet deforms under the progressive action of the mandrel being drawn through. When the brake pressure is released, the spring drives the piston back only the amount of the original lost motion, providing for consistent running clearance. A cross-sectional drawing of a typical arrangement is shown in Figure 4.7. Because plastic deformation of metal is used to achieve the automatic adjustment, the adjuster tube can only be used once. At overhaul of the brake (heat sink replacement), the used adjuster tubes are discarded and new ones fitted. Applied pressure to the brake must overcome the force of the return springs as well as any frictional loss in the system. This pressure is referred to as the ineffective pressure, and it will vary with the specific design of the brake and the design operating pressure of the brake system. For brake systems operating at 206 bar (3000 psi), typical ineffective pressures are in the range of 15–20 bar (220–290 psi). The required stroke of the piston is determined by the wear thickness of the brake (from new to completely worn) plus any additional stroke required to account for lost friction material during a high-energy RTO and an allowance for the flexibility of the brake components. The wear thickness can be estimated from Figure 4.2 but will vary depending on the specific supplier of the brake and their proprietary material. Providing a means to avoid overextension of the brake piston is required to prevent ejection and resulting leakage of hydraulic fluid, which could result in a fire when the fluid leaks onto a hot brake. The piston arrangement must be such that the ineffective pressure is sufficient to ensure that hydraulic system back pressure cannot apply the brake. Determination of the number of pistons and their size is related to the required clamping force (calculated from the required torque and the minimum brake efficiency) and the maximum available braking pressure. Dividing the clamping force by the braking pressure less the ineffective pressure results in the total piston area

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

33

© SAE International.

 FIGURE 4.8   Brake electromechanical actuator.

required. A number of pistons are generally arrayed around the piston housing to meet this requirement. Fewer pistons result in less cost and complexity, but unless the brake is a caliper style, the general minimum is four pistons with five being a more common minimal value. The available space may determine the total number as it is often easier to array a larger number of smaller pistons around the housing. Electric actuation of brakes replaces the hydraulic pistons and adjusters with an electromechanical actuator (EMA): usually a brushless direct current motor, gear reduction box, and ballscrew as indicated in Figure 4.8. This arrangement provides linear motion from the rotary drive available from the motor. Sizing of the EMA is a specialist activity. To achieve the possibility of dispatching the aircraft with a single EMA inactive, each EMA is often sized such that the brake can generate its full clamping force with one of the several actuators in a failed state. Given the large gear reduction ratio between the motor and the linear portion of the actuator, the actuator can be locked effectively with a small brake (either a friction brake or a clutch arrangement) on the motor shaft. This shaft brake is usually solenoid operated with the shaft locked when power is off. The actuator is able to serve the parking brake function by being powered to deliver the required clamping force and then de-energized. The small motor shaft brake keeps the actuator from backing off. As brushless DC motors are used (for their high energy density), electronic commutation of the motor is required. Effective motor control schemes require detailed knowledge of the motor shaft position or resulting clamping force. Position control has the advantage that the brake controller is aware of the position of the actuator (by counting the motor revolutions and dividing by the gear ratio). As the brake controller can precisely control the position of the actuator, no dedicated adjuster mechanism is required. In addition, brake wear status can be reported to the aircraft users without the need for any additional instrumentation. Some disadvantages of electric actuation for brakes are that an estimate or direct measurement of the clamping force delivered must be made – force measurement systems in the brake environment are difficult to make reliable. Additional disadvantages are that

34

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

the unit cost and weight of the brake are increased compared to a hydraulic brake. However, with advances in technology it appears that an electric brake system can be of the same weight or lighter than a hydraulic brake when all the associated system components are taken into account. AIR5937 [19] provides additional details on electric brakes.

Mechanical Connection to the Landing Gear Structure Many methods exist to retain the brake on the landing gear and to transfer the torque from the brake to the landing gear structure. On small brakes, the caliper or calipers are often bolted rigidly to the structure, as seen in Figure 1.3. On large brakes, three primary methods exist to extract the torque from the brake. A torque arm is often fitted to brakes on multiple wheel gears. This is a lug on the piston housing that receives a pin. The arrangement is clearly shown in Figure 4.9. This mounting arrangement is advantageous when rotation of the brake is required, such as when mounted on a bogie beam. It can be used when mounting directly to landing gear structure but requires a dedicated pin or suitable structure to be provided. When using this type of mount directly to structure, care must be taken that axle and structure deflections do not lead to unacceptable loads on the piston housing. A torque take-out lug (Figure 4.10) may provide a more inexpensive solution for the structural attachment – a boss or fitting may be directly machined on the landing gear structure to receive the brake. This arrangement is often utilized on trailing arm

© SAE International.

 FIGURE 4.9   Electric brake with torque arm.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

35

Reprinted from © Safran Landing Systems.

 FIGURE 4.10   Hydraulic brake with torque take-out lug.

landing gears where bosses are fitted on either side of the trailing arm to accept the brake. The torque take-out lug may also be advantageous where relative movement of the brake and structure is expected as the sliding fit of the joint can accommodate differential movement while still resisting the required torque. Differential movement can arise from the use of different materials on the brake housing compared to the landing gear structure as well as local variation in the temperature of the components. Another method to extract the torque is to remove it directly from the brake torque tube or brake flange. This type of assembly, shown in Figure 4.11, connects to a flange provided (visible in Figure 4.12) on the landing gear axle. This type of connection likely represents the stiffest brake to landing gear attachment and is the lightest

© SAE International.

 FIGURE 4.11   Dual cavity hydraulic brake with brake flange mounting.

36

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 4.12   Landing gear axle showing brake flange and brake sleeve.

for the brake as the torque is not seen by the piston housing. In the assembly shown, the torque is taken by a series of pins while the brake is retained by three bolts. While an excellent assembly from the brake’s perspective, it represents an expensive option for the manufacture of the landing gear structure. It also provides the least amount of damping for vibration control. On aircraft where axle mounted brakes are removed and replaced frequently (such as on commercial transport aircraft), a brake sleeve is usually fitted over the axle. The brake sleeve serves to protect the axle (which is typically made of highstrength steel) from abrasion and damage to its protective treatments during the removal and replacement of the brake. Brake sleeves are usually manufactured from corrosion resistant steel or titanium and are matched to the design of the brake; the sleeve can serve to bridge the load bearing lands provided on the axle to the physical interface locations required by the brake – which can be very useful if a pedestal or other support is added to the torque tube to avoid vibration modes such as whirl.

Weight The optimum brake weight estimation approach is to calculate the expected heat sink mass, size, and required clamping force. With this, a mechanical design of the brake components can be conducted and a weight identified based on the materials selected. A lower fidelity approach for initial sizing estimation can be conducted using the curves in Figure 4.13 based on the maximum energy to be absorbed. This approach has significant scatter as it does not consider the required brake torque, brake structure materials, and design brake life. As can be seen, there is little data in the plot for steel brakes, so the prediction of steel brake weights should be done with caution.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

37

 FIGURE 4.13   Weight estimation for steel and carbon brakes.

Brake Assembly Mass vs. Maximum Energy 300

© SAE International.

Brake Assembly Mass (kg)

250

200

150 y = 2.0489x + 7.6956 R² = 0.9637

100

y = 0.8653x + 13.722 R² = 0.9615

50

0 0

20

40

60

80

100

120

140

160

Maximum Energy (MJ)

Worked Example Considering the details provided, size the brakes for a civil aircraft having:

•• a maximum takeoff mass of 32 500 kg •• four braked main wheels •• H38×13R18 tires •• rejected takeoff speed of 120 knots •• requested 0.35 g minimum average deceleration during rejected takeoff •• 3000 psi aircraft hydraulic system pressure Carbon brakes with a wear life of 2000 flight cycles are desired. As little is known about the aircraft, the kinetic energy calculation method should be used to determine the maximum energy required to be stored in the brakes. Using this approach will provide some margin to permit preheating of the brakes during taxi out:



Convert 120 knots to m / s : 120 0.5144 61.7 m / s Calculate kinetic energy:



Ek

1 mV 2 2

38

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

1 2 32 500 61.7 2

Ek



Ek

61 918 122 J

Calculate kinetic energy per brake:



61 918 122 15 479 531J 15 479kJ 4

Ek

Using the more conservative value of mass loading for carbon brake, RTO, and worn condition from Table 4.3, compute the minimum required mass of carbon:



Brake mass

15 479 kJ 5.7 kg 2700 kJ / kg

Using a value of 2700 kJ/kg provides a minimum heat sink mass of 5.7 kg. The rotor and stator inner and outer diameters are estimated using the expressions in Table 4.4 for the 18-in. wheel size that supports the selected tire. The values are shown in Table 4.5. The contact (swept) area between the stators and the rotors is the annular area defined by the outer diameter of the stator and the inner diameter of the rotor. In this case, the annular area is:



A



A



De2 4

Di2 4

2

9.6

15 4

4

2



A 104.3 in.2 Converting square inches to square centimeters:



Area is 104.3 6.4516 672.9 cm2 .

TABLE 4.5  Estimated stator and rotor sizes for 18-in. wheel. Civil aircraft estimation formulae

18 in. wheel size results

Stator outer diameter

0.7091D+2.286

15 inch

Stator inner diameter

0.417D+0.391

7.9 inch

Number of stator drive keys

[1.442D–11.25]

15

Rotor outer diameter

0.788D+2.322

16.5 inch

Rotor inner diameter

0.6645D–2.361

9.6 inch

Number of rotor drive keys

[0.35D+2.217]

9

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

39

For a mass of 5.7 kg, and assuming that the rotor and stator drive key materials are not part of the heat sink (a conservative assumption), an annulus having the inner and outer diameter identified previously would require a length of: Determine the volume of carbon from the mass, assuming a density of 1.8 g/ cm3 from Table 4.1:



Volume

mass density

5.7 1000 3167 cm3 1.8

Volume

With an annular area of 672.9 cm2, the length of the annulus (heat sink) would be:



Length



Volume Area 3167 4.7 cm. 672.9

Length

The estimated length of the completely worn heat sink is approximately 4.7 centimeters. The radius, r, where the frictional force will act must be determined:





r

2 re3 ri 3 3 re2 ri 2



15 2

3

9.6 2

3

2

2

9.6 2

2

3

15 2

r

6.25 in

Converting 6.25 inch to meters: 6.25 × .0254 = 0.159 m. The required torque to provide the requested deceleration of 0.35 g during an RTO is calculated using the minimum loaded tire radius for the tire (15.75 inch) from tire data (available in The Design of Aircraft Landing Gear, Chapter 3, and Aircraft Tires: Key Principles for Landing Gear Design): Force to decelerate the aircraft:



F

ma

40

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design



F

32 500

0.35

9.806

111 543 N

As there are four braked wheels, the individual force to be developed in each tire contact patch is:



F

111 543 27 886 N 4

The torque required to generate this is:



F rrolling

Convert 15.75-inch tire loaded radius to meters: 15.75 × .0254 = 0.4 m. The minimum brake torque is then:



27 886 0.4 11 154 Nm

Assuming 95% of the aircraft mass is supported by the main landing gears confirm adequate tire to ground friction will be available to achieve this deceleration:

•• Mass supported by main landing gears = 32 500 × 0.95 = 30 875 kg 30 875 7719 kg 4 Normal force at each tire contact patch = 7719 x 9.806 = 75 962.5 N

•• Mass supported by a single wheel = ••

Required friction coefficient at ground:



F N



27 886 0.37 75962.5

As this value is less than the 0.45 value typically taken for dry pavement, the deceleration will not be friction limited at the ground. With the required minimum brake torque determined, the clamping pressure and number of rotors remain to be selected:



2N

F r

A value of brake efficiency of μ=0.2 is taken, in the absence of more detailed data. With two unknowns, the selection of the clamping force and number of rotors can be an iterative process. For the purposes of this exercise, two rotors will be selected as an initial value. While increasing the number of rotors is beneficial to increase brake torque, it can lead to reduced brake efficiency and increases the number of

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

41

wearing faces of the brake. During detail design, the brake designer will trade the weight and complexity of the number of rotors against the weight and size of the required pistons and available hydraulic pressure. The clamping force required assuming a two rotor brake:



F

F

2N

r



11,154 87689 N 2 2 0.2 0.159

Determine the number and size of the hydraulic pistons to provide the required clamping force. A minimum diameter of 32 mm is preferred to permit an integral adjuster. The ineffective pressure will be considered to be 290 psi, leaving an effective pressure of: Effective pressure maximum applied pressure ineffective pressure 3000 290 2710 psi



2710 0.006895 18.7 N / mm2

Convert 2710 psi to N / mm2 MPa

The clamping pressure divided by the effective pressure gives the total piston area:



A

A

F P

87689 4689.3 mm2 18.7

The area of a single piston having a 32 mm diameter is:



d2 4

A

A

32 4

2

804.2mm

To achieve the required, total piston area requires:



4689.3 5.8 804.2

Therefore, 5.8 pistons of 32 mm diameter are required. A piston housing with six 32 mm pistons could be utilized. However, a piston housing with five pistons could be used with a larger piston size:

42

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Target piston area for five pistons: 4689.3 937.9 mm2 5



The piston diameter required to achieve an area of 937.9 mm2:



A



d



d

4 937.9

d2 4 4A



34.56 mm

Five pistons of 34.56 mm would achieve the deceleration target. In normal practice, the piston diameter is rounded up to a standard O-ring size. Given the small diameter change, some additional margin may be taken to ensure adequate torque under all circumstances. To determine the length of the pistons requires an estimation of the wear thickness to meet the 2000 flight cycle objective. If the normal landing condition (which was not provided) is assumed to be 30% of the RTO energy (a design landing value as suggested by Table 3.1), then the normal landing energy is approximately 15.479 MJ × 0.3 = 4.6 MJ. The mass loading for this condition is then approximately: 4.6/5.7 = 0.8 MJ/kg. The estimated wear from Figure 4.2 is then 0.0014 mm per face per cycle. A brake with two rotors has eight wearing surfaces (the stators and the rotors both wear at the contact faces). The total wear length for 2000 cycles is:



l 2000 .0014 8 22 mm

The worn heat sink length of 4.7 cm is then extended by 22 mm (2.2 cm) to be 6.9 cm. As there are two rotors (necessitating three stators), the new thickness of the rotors and stators is 6.9 cm divided by 5, for a disk thickness when new of 1.38 cm (13.8 mm). The wear pin on the brake would be set to 22 mm to indicate the fully worn condition. Using the weight estimation curve from Figure 4.13, this brake is likely to have a mass of approximately 27 kg.

5 Wheel and Brake Certification and Recommended Practices

T

he specific certification requirements for the brake and wheel (usually they are certified as a matched pair) depend on the application. For small civil aircraft as well as helicopters, TSO-C26d [20] is the applicable standard (ETSO-C26d is the European equivalent). An update to this TSO is pending which will refer to AS5714 [14] for the minimum performance standard. For civil large transport aircraft, TSO-135a [15] is the applicable standard (ETSO-135a is the European equivalent); future updates are expected to refer to AS6410 [21] for the minimum performance standard. The military standard, MIL-W-5013 [22], is inactive for new design and has no direct replacement with many militaries taking a more performance-based approach to standardization. An updated set of requirements in the absence of a replacement for MIL-W-5013 are provided in ARP1493 [23]. The military standard is significantly more prescriptive than the civil technical standard orders. A comparison of the three major certification regimes is provided in Table 5.1. Not every requirement has been included, and much of the requirement text has been simplified to ease reading - refer directly to the latest standard when required rather than working from this table. A large number of prescriptive requirements from the military standard have not been included in the table, where they do not directly relate to the function of the brake or wheel. Further to the certification requirements, a number of industry recommended practices help ensure reliable performance of brakes and wheels. Those which supplement the certification regulations for large transport aircraft are outlined in ARP597 [24] and are summarized in Table 5.2. Additional recommended practices to ensure good maintainability are outlined in ARP813 [25].

©2022 SAE International

43

Wheels and brakes for amphibious aircraft and/or electric brake actuation shall be sealed to prevent water and contaminant ingress

For tubeless tire applications, overpressure relief means must be provided

TRA and/or European Tyre and Rim Technical Organization (ETRTO) approval of rim and valve dimension is encouraged

Hydraulic brakes with more than one rotating disk shall include a hard stop which prevents fluid escape

Sealing

Burst Prevention

Wheel Rim and Inflation Valve

Brake Piston Retention

The brake must incorporate means to ensure that the actuation system does not allow hydraulic fluid to escape if the limits of piston travel are reached

TRA and/or ETRTO approval of rim and valve dimension is encouraged

Means must be provided to prevent wheel failure and tire burst that might result from over pressurization or from elevated brake temperatures

Wheels intended for use on amphibious aircraft must be sealed to prevent entrance of water into the wheel bearings or other portions of the wheel or brake, unless the design is such that brake action and service life will not be impaired by the presence of sea water or fresh water

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  Wheel and brake certification criteria.

Piston stops shall be provided when structural carbon composites are used for brake heat sinks that limit travel to prevent venting of hydraulic fluid upon stark failure or operation beyond the normal removal condition. The piston stops shall allow sufficient piston travel to permit a maximum designed gross weight, rejected take-off (MDGW RTO) at the worn out brake condition. The stops shall be designed for 1.5 times maximum operating pressure without the disks installed

The wheel rim contour shall conform to the rim contour standard for the particular tire listed in MIL-T-5041. In cases where those standards do not exist, the rim contour shall conform to the specification control drawing or to the one recommended by Tire and Rim Association (TRA)

—

—

MIL-W-5013 (historical)

Military

44 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

—

—

Magnesium and alloys having magnesium as Magnesium and alloys having magnesium as a major constituent must not be used on a major constituent shall not be used on brakes or braked wheels brakes or braked wheels for commuter category applications

—

Wheel Bearings

Fatigue

Magnesium Parts

Steel Parts

—

The design of the wheel must incorporate techniques to improve fatigue resistance of critical areas of the wheel and minimize the effects of the expected corrosion and temperature environment. The wheel must include design provisions to minimize the probability of fatigue failures that could lead to flange separation or other wheel burst failures

Means should be incorporated to avoid mis-assembly of wheel bearings

(A number of additional requirements on steel selection and procurement are provided)

Consumable electrode vacuum melted steel shall be used for parts made from heat treated alloy steel with ultimate tensile strengths of 220,000 psi and above

Free machining carbon steel shall not be used.

Magnesium and magnesium alloys shall not be used

—

The wheel bearings shall be of the tapered roller type Conforming to FF-B-187

Automatic adjusters shall be provided to compensate for brake lining wear. Brake assemblies shall be designed for the most practicable protection of the brake adjusters

The brake mechanism must be equipped with suitable adjustment means to maintain appropriate running clearance when subjected to the rated retraction pressure

When provided, these mechanisms shall maintain appropriate running clearance. Electrically actuated brakes shall provide a system to ensure appropriate running clearance is maintained

Brake Release and Wear Adjustment

Brake lining wear indicators shall be provided

MIL-W-5013 (historical)

Military

A reliable method must be provided for determining when the heat sink is worn to its permissible limit

A visible and reliable wear indicator must be provided

Wear Indicator

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 45

For all tests the wheel must be fitted with a suitable tire and wheel loads must be applied through the tire (some exceptions for ultimate cases exist)

Yield Load: 1.15 time the radial limit load

Wheel Tests

Radial Load

Yield Load: 1.15 time the radial limit load

For all tests the wheel must be fitted with a suitable tire and wheel loads must be applied through the tire (some exceptions for ultimate cases exist)

Determine wheel limit load in accordance with 14 CFR 25.471 through 25.511

Ultimate Load: 2 times the radial limit load Ultimate Load: 2 times the radial limit load for cast wheels, 1.5 times the radial limit load for cast wheels, 1.5 times the radial limit for forged wheels load for forged wheels

Determine wheel limit load in accordance with to 14 CFR paragraphs 23.471 through 23.511, or 27.471 through 27.505, or 29.471 through 29.511, as appropriate.

Wheel capacity

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

The wheel shall support the ultimate load for 10 seconds after which there shall be no cracks in any areas. Upon successful completion of this test, the radial load shall be reapplied to failure

There shall be no yielding of the wheel such as would result in loose bearing cups, air leakage, or interference in any critical clearance areas such as brake interface part (if brake-type wheel)

The maximum limit, yield, and ultimate loads shall be determined in accordance with MIL-A-8863 and shall be equal to the operational condition of maximum radial load reaction with the side load equal to zero

The rated load capacity of each landing wheel or auxiliary wheel on an aircraft shall be equal to or greater than the maximum load that the wheel will be subjected to at maximum towing or taxiing static design gross weight of the aircraft. (Additional information on dynamic loading provided)

MIL-W-5013 (historical)

Military

46 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Combined Radial and Side Load

Requirement

Combined Yield Load: 1.15 times the ground limit loads

Combined Ultimate Load: 2 times the ground limit loads for cast wheels, 1.5 times the ground limit loads for forged wheels

Combined Ultimate Load: 2 times the ground limit loads for cast wheels, 1.5 times the ground limit loads for forged wheels

—

—

Combined Yield Load: 1.15 times the ground limit loads

ETSO-C135a

TSO-C135a

TSO-C26d (AS5714)

ETSO-C26d

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

Ultimate combined radial-side load test. The ultimate combined load shall be applied at the 0° position of the same wheel on which the respective yield combined load tests were performed. The ultimate load shall be sustained for a minimum of 10 seconds after which there shall be no cracks in any area. With the side load component held constant or allowed to proportionally Increase with the vertical load, the vertical load shall be Increased until failure occurs.

Yield combined radial-side load test. The wheel shall support the components of the yield combined load applied consecutively at 0°, 90°, 180°, and 270°, followed by two more load applications at the 0° position

Design landing radial load test. The maximum design landing load shall be determined in accordance with MIL-A-8863. The load shall be supported for not less than 10 seconds, and the resulting permanent set shall not produce loose bearing cups, air leakage, interference in critical running areas, or make the wheel unsuitable for further service. The tire inflation pressure shall be the maximum design operating pressure for the condition being simulated. For Navy aircraft intended for shipboard use, the wheel shall be loaded for this test condition through 1-1/2-in. diameter cable or steel bar that simulates statically the wheel design landing load plus the load imposed by rolling over or landing on a 1-1/2-in. diameter cable. At the completion of the above, the same wheel and brake assembly shall be rolled a minimum of 5000 feet at the rated static load

MIL-W-5013 (historical)

Military

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 47

Overpressure Test

Withstand 3.5 times the rated inflation pressure

Withstand four times the rated inflation pressure

Roll-on-Rim Test (not applicable to nose — landing gear wheels): The wheel assembly without a tire must be tested at a speed of no less than 10 mph (4.6 m/s) under a rated static load. The test roll distance (in feet) must be determined as 0.5VR2 but need not exceed 15,000 feet (4572 meters). The wheel assembly must support the load for the distance defined above. During the test, no fragmentation of the wheel is permitted; cracks are allowed

—

Wheels of land-based aircraft shall be tested to a burst pressure of 3.5 times the rated tire pressure, at the rated static load of the wheel or the burst strength of the tire, whichever is least. Wheels of carrier-based aircraft shall be tested to 4.5 times the rated tire pressure or to the burst strength of the tire, whichever is least. Helicopter wheels shall be tested to a burst pressure producing not less than 3.0 times the axial load that results from the tire pressure required for the static wheel load at the taxi gross weight

The roll test shall consist of a series of landings or a continuous roll of the wheel assembly against a rotating flywheel to complete the roll test spectrum for a total of 3000 miles. Helicopter wheels are not subject to the loading conditions in the roll test spectrum above unless specified in the aircraft detail specification. Instead, helicopter wheels shall be rolled 250 miles minimum with an applied radial load not less than the static wheel reaction based on helicopter maximum taxi gross weight

MIL-W-5013 (historical)

With inflation pressure not less than 1.14 times rated inflation pressure, roll 2000 miles under rated static load, 100 miles under rated load with 0.15 × rated load applied as an outboard side load, and 100 miles under rated load with 0.15 × rated load applied as an inboard side load; No cracking or leakage permitted

Under wheel rated load, with inflation pressure 1.1 times rated inflation pressure, roll a distance of 1000 miles for Part 23 aircraft; 500 miles for parts 27 and 29 rotorcraft. No cracking or leakage permitted

Military

Wheel Roll Test

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

48 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

A tubeless tire and wheel assembly must hold its rated inflation pressure for 24 hours with no greater than 5% pressure drop

For Part 23 aircraft, the wheel and brake assembly must complete 100 stops at the design landing energy, each at the mean distance averaged deceleration (normally defined by the aircraft manufacturer) but not less than 10 ft/s2 (3.05 m/s2); the wheel and brake for Part 27 or 29 rotorcraft must complete 20 stops at the design landing energy, each at the mean distance averaged deceleration (normally defined by the aircraft manufacturer) but not less than 6 ft/ s2 (1.83 m/s2)

—

Diffusion Test

Design Landing Stop Test

Peak Torque

MIL-W-5013 (historical)

Military

—

The wheel and brake assembly must complete 100 stops at the design landing energy, each at the mean distance averaged deceleration (normally defined by the aircraft manufacturer) but not less than 10 ft/s2 (3.05 m/s2)

The peak brake torque during any aircraft braking condition, within the speed and pressure range of the aircraft, shall not produce a peak drag deflection of the strut and axle exceeding that due to a drag load of 0.8 times the maximum wheel static 1 g vertical load, at maximum design gross weight, applied at the corresponding loaded radius

A spectrum depending on aircraft type is provided. Land and carrier based fighters, bombers, attack, reconnaissance, and refueling tankers have to meet 45+5 landings under various conditions. Helicopters have to meet 20 landings. Land-based cargo, patrol, transport, trainer, and liaison aircraft have to meet 100 landings. A rational approach to energy calculation is expressly provided

A tubeless tire and wheel assembly must All wheels intended for use with tubeless tires shall hold its rated inflation pressure for 24 hours show adequacy to retain rated operating tubeless tire with no greater than 5% pressure drop pressure. The tubeless tire and wheel assembly shall hold the normal inflation pressure for 24 hours with no greater pressure drop than 5 psi

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 49

Most Severe Landing Stop Test

—

No sustained fire that extends above the level of the highest point of the tire is allowed before 5 minutes have elapsed after application of parking pressure or brake clamping force as applicable following the accelerate-stop test; until this time has elapsed, neither firefighting means nor coolants may be applied

Applicable to Commuter Category Aircraft and Jets Over 6000 Pounds. The wheel and brake assembly complete the accelerate-stop test at the mean distance averaged deceleration (normally defined by the aircraft manufacturer) but not less than 6 ft/s2 (1.83 m/ s2); for the accelerate-stop test, the tire, wheel, and brake assembly must be tested at energy for a new brake, and it is recommended that this test is conducted on a fully worn brake. While not required by regulation for Part 23, 27, and 29 aircraft it is strongly supported by industry as a safety enhancement, as required for Part 25 airplanes in TSO-C135a

Accelerate-Stop Test

The wheel and brake assembly under test — must complete the most severe landing braking condition expected on the airplane as normally defined by the airplane manufacturer. This test is not required if the accelerate-stop testing is more severe or the condition is shown to be extremely improbable, normally by the airplane manufacturer

No sustained fire that extends above the level of the highest point of the tire is allowed before 5 minutes have elapsed after application of parking pressure or brake clamping force as applicable following the accelerate-stop test; until this time has elapsed, neither firefighting means nor coolants may be applied

MIL-W-5013 (historical) —

The wheel and brake assembly complete the accelerate-stop test at the mean distance averaged deceleration (normally defined by the aircraft manufacturer) but not less than 6 ft/s2 (1.83 m/s2); for the accelerate-stop test, the tire, wheel, and brake assembly must be tested for both a new brake and a fully worn brake

Military

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

50 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Test to a structural torque of 20% above the vertical static load times the rated tire’s radius at that load for landing gear with a single wheel per leg and 44% above that value for landing gear with more than one wheel per leg. The wheel and brake must withstand these loads without failure for at least 3 seconds.

There must be no interference in any critical areas between the wheel and brake assembly (with fittings) up to limit load conditions

Structural Torque Test

Wheel to Brake Clearance Test

There must be no interference in any critical areas between the wheel and brake assembly (with fittings) up to limit load conditions

Test to a structural torque of 20% above the vertical static load times the rated tire’s radius at that load for landing gear with a single wheel per leg and 44% above that value for landing gear with more than one wheel per leg. The wheel and brake must withstand these loads without failure for at least 3 seconds.

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

The brake shall be actuated at twice the normal operating pressure, or force in the case of a mechanical brake, or the maximum safe operating pressure, whichever is higher. Tangential load shall then be applied at the static load radius of the tire until the applied tangential load equals 1.2 times (1.0 for helicopters) the maximum rated static load of the wheel. The friction surfaces of the brake may be bolted or clamped together or otherwise restrained to withstand the required tangential load of 1.2 times (1.0 for helicopters) the maximum rated static load of the wheel. The wheel and brake shall withstand the structural torque test without failure. Co-rotating nosewheels shall also be tested to verify the structural integrity of the co-rotating feature. These shall be subjected to a tangential load of 1.2 times maximum rated load of the wheel applied at the loaded radius of the tire. There shall be no evidence of failure as a result of this test

MIL-W-5013 (historical)

Military

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 51

Hydraulic brakes must not exceed a total leakage of 5 cc during the test.

Endurance Test: For fixed wing airplanes, the brake must be subjected to 100,000 cycles of an application of the average of the peak brake pressures needed in the design landing stop test. The pistons must be adjusted so that 25,000 cycles are performed at each of the four positions where the pistons would be at rest when adjusted to nominally 25, 50, 75, and 100% of the wear limit. The brake must then be subjected to 5000 cycles of application of pressure to rated pressure and release at the 100% wear limit. For rotorcraft, the same requirements exist but only for 50,000 cycles, split in four groups of 12,500 cycles, followed by 2500 cycles of pressure at the 100% wear limit.

Yield and Overpressure Tests: the brake must withstand a pressure equal to 1.5 times the Brake Rated Maximum Pressure for at least 5 minutes without permanent deformation of the structural components under test. For rotorcraft, the brake must withstand the greater of a pressure equal to 1.5 times the Brake Rated Maximum Pressure or 1.5 times Brake Rated Minimum Slope Pressure (20° slope) for at least 5 minutes without permanent deformation of the structural components under test

Hydraulically Actuated Brake Tests

The brake must be subjected to 100,000 cycles of an application of the average of the peak brake pressures needed in the design landing stop test. The pistons must be adjusted so that 25,000 cycles are performed at each of the four positions where the pistons would be at rest when adjusted to nominally 25, 50, 75, and 100% of the wear limit. The brake must then be subjected to 5000 cycles of application of pressure to rated pressure and release at the 100% wear limit

The hydraulic brake shall be subjected to 100,000 cycles (50,000 for helicopters) of application and release of pressure equal to normal operating pressure and 5000 cycles (2500 for helicopters) at a pressure equivalent to the maximum operating pressure. This test shall be conducted using a minimum clearance equivalent to the maximum clearance allowable between adjustments. The first portion of the test may be divided into four parts so that 25,000 cycles (12,500 for helicopters) may be applied at each of four positions of brake piston travel conforming to 25, 50, 75, and 100% travel, respectively

The brake shall be parked for a period of 5 minutes with an applied operating pressure equal to twice the maximum operating pressure. The test shall be conducted with linings having a thickness comparable to the maximum permissible wear. There shall be no leakage or failure during this test. Pressure shall then be increased until failure occurs and the ultimate pressure shall be recorded

Pressure

MIL-W-5013 (historical)

The brake must withstand a pressure equal to 1.5 times the Brake Rated Maximum Pressure for at least 5 minutes without permanent deformation of the structural components under test; the brake, with actuator piston(s) extended to simulate a maximum worn condition, must, for at least 3 seconds, withstand hydraulic pressure equal to 2.0 times the Brake Rated Maximum

Military

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

52 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Impulse Tests

Requirement

Leakage Tests: No leakage after 5 minutes at 1.5 times maximum pressure followed by 5 minutes at 5 psig; Minimal leakage after 25 applications of maximum brake pressure

Leakage Tests: No leakage after 5 minutes at 1.5 times maximum pressure followed by 5 minutes at 5 psig; Minimal leakage after 25 applications of maximum brake pressure

—

The hydraulic pistons must be positively retained without leakage at 1.5 times maximum brake pressure for at least 10 seconds with the heat sink removed

Piston Retention Test: If piston retention means are included in the design, the hydraulic pistons must be positively retained without leakage at 1.5 times maximum brake pressure for at least 10 seconds with the heat sink removed

—

ETSO-C135a

TSO-C135a

TSO-C26d (AS5714)

ETSO-C26d

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

The brake assembly shall be subject to 200,000 impulse cycles from 250 psi to 1.5 times the normal system’s pressure (or braking pressure) to 250 psi. Cycling may be conducted at the maximum rate consistent with achieving the required pressure levels and ensuring the pistons return to their rest position after each pressure cycle. The cycles shall be divided into four parts so that an equal number of cycles may be applied at each of the four positions of brake piston travel conforming to 25%, 50%, 75%, and 100%, respectively. There shall be no deformation or structural failure of the brake assembly

The brake shall be subjected to 25 cycles of the application and release of maximum operating pressure. Leakage at static seals shall not exceed a trace. Leakage at moving seals shall not exceed one drop of fluid per each 3 inches. of peripheral seal length.

The brake shall be parked for a period of 5 minutes with an applied operating pressure equal to 1-1/2 times (1.0 for helicopters) the maximum operating pressure. The brake shall then be parked for a period of 5 minutes with an applied pressure of 5 psi. There shall be no measurable leakage (less than one drop) or permanent set during these tests

The piston stops and brake housing for carbon heat sinks shall demonstrate their ability to withstand one and one-half times the maximum operating pressure for 5 minutes without the brake discs installed. No deformation or performance degradation shall result

MIL-W-5013 (historical)

Military

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 53

—

Extreme Temperature Soak Test

MIL-W-5013 (historical)

Military

Hydraulic brakes must not exceed a total leakage of 5 cc during the temperature soak tests:

Aging and heat test. The brake, filled with operating fluid, shall be subjected to a temperature of 160°F (71°C) for AN seals, 225°F (107°C) for MS seals or better, or higher temperature as required by the Subject the brake to at least a 24-h hot particular application other than energy overload soak at the maximum piston housing fluid conditions for seven days. With the brake and temperature experienced during a design landing stop test, conducted without forced operating fluid being maintained at this temperature, the brake shall be cycled 1000 times at normal air cooling. While at the hot soak operating pressure followed Immediately by 25 cycles temperature, the brake must be subjected at maximum operating pressure. Leakage limits apply. to the application of the average of the Cold test. Upon completion of the aging and heat test peak brake pressures required during the (4.3.31.1), the brake, filled with operating fluid under 100 design landing stops and released for atmospheric pressure, shall be subjected to a 1000 cycles, followed by 25 cycles to the maximum operating pressure and released. temperature of –65°F (–54°C) for a period of 72 h. The brake must then be cooled from the hot There shall be no leakage during this period. With the brake and operating fluid being maintained at this soak temperature to a cold soak temperature, the brake shall be cycled 25 times at temperature of –40°F (–40°C) and normal operating pressure followed immediately by maintained at this temperature for at least five cycles at maximum operating pressure. The brake 24 hours. While at the cold soak temperature, the brake must be subjected clearance shall be checked between each cycle at maximum operating pressure to insure that the brake to the application of the average of the releases completely peak brake pressures required during the design landing stops and released for 25 Leakage limits apply cycles, followed by 5 cycles to maximum operating pressure and released

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

54 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

— 100,000 cycles to the design landing load, equally divided into five or more equally incremented wear positions, including the new and fully worn positions, followed by 5000 cycles to maximum brake load conducted at five equally incremented wear positions, including the new and fully worn positions

Endurance Test: Aircraft – 100,000 cycles to the design landing load, divided into four groups of 25,000 at different wear levels, followed by 5000 cycles to maximum brake load conducted at five different wear positions; Rotorcraft – 50,000 cycles to the design landing load, divided into four groups of 12,500 at different wear levels, followed by 2500 cycles to maximum brake load conducted at five different wear positions

MIL-W-5013 (historical)

Limit and Ultimate Load Test: The brake — must withstand for at least 5 seconds a force equal to the Brake Limit Load without permanent deformation; the brake, with EMAs extended to simulate a maximum worn condition, must for at least 3 seconds withstand a force equal to 1.5 times the limit load

Limit and Ultimate Load Test: The brake must withstand for at least 5 seconds a force equal to the Brake Limit Load without permanent deformation; the brake, with EMAs extended to simulate a maximum worn condition, must for at least 3 seconds withstand a force equal to 1.5 times the limit load

Military

Electrically Actuated Brake Tests

ETSO-C135a

TSO-C135a

ETSO-C26d

TSO-C26d (AS5714)

Requirement

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design 55

Requirement

MIL-W-5013 (historical)

— Subject the brake to at least a 24-h hot soak at the maximum actuator housing temperature experienced during a design landing stop test (paragraph 3.3.2), conducted without forced air cooling. While at the hot soak temperature, the brake must be subjected to the application of the Brake Design Landing Load required during the 100 design landing stops and release to the brake off position for 1000 cycles, followed by 25 cycles of Maximum Brake Load and release to brake off position. The brake must then be cooled from the hot soak temperature to a cold soak temperature of –40°F (–40°C) and maintained at this temperature for at least 24 hours. While at the cold soak temperature, the brake must be subjected to the application of the Brake Design Landing Load required during the design landing stops and release to brake off position, for 25 cycles, followed by five cycles of Maximum Brake Load and release to brake off position. The brake assembly shall then meet the functional test requirements (acceptance tests) established to assure continued airworthiness

—

Military

ETSO-C135a

TSO-C135a

TSO-C26d (AS5714)

ETSO-C26d

Large Transport Aircraft

Light Aircraft, Helicopters

TABLE 5.1  (Continued) Wheel and brake certification criteria.

56 Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

57

TABLE 5.2  Recommended practices for large transport aircraft wheels and brakes – ARP597. Recommended practices

Large transport aircraft ARP597

Wheel endurance tests

Additional wheel endurance: Consideration should be given to performing a spectrum of roll testing (based on the intended aircraft mission profile) for distances up to 50,000 miles. Accelerated roll testing at higher than anticipated loads is an effective way to reduce roll test duration Corrosion Testing: leaving sensitive areas of the wheel unprotected against corrosion, subject the wheel to salt spray exposure to yield corrosion pit depths of 0.02–0.03 inch. Then, roll test the wheel in this condition to multiple lives to demonstrate the tolerance of the wheel to corrosion damage Missing Tie Bolt Endurance Test: The missing tie bolt endurance test objective is to demonstrate that loss of one wheel tie bolt should not lead to failure of other tie bolts or to wheel failure before the number of cycles corresponding to at least one tire life. The test should be performed on a wheel with one bolt removed, with the same load spectrum as used in the above-described endurance test. An accelerated roll test procedure could be used. The test should preferably be performed on a wheel that has previously completed a roll test (e.g., corrosion testing or TSO-C135 roll test).

Vibration

Vibration assessment and testing (if required) should be conducted to ensure that the brake vibration modes will not interfere with passenger comfort, antiskid behavior, landing gear dynamics and brake stability. Brake integrity must be maintained throughout vibration testing.

Brake wear tests

Brake wear tests: A spectrum of additional wear testing for carbon brakes and a different spectrum for steel brakes is provided. The number of taxi stops and snubs is particularly important in assessing carbon brake life. For steel brakes, landing energies play a significant role in characterizing brake life. The different spectra focus on the important variables to yield life estimates for each brake type. Static torque: It is recommended that one static pull be performed during every other landing test sequence to permit reconditioning of friction surfaces. The brake pressure applied during all those tests should be the parking pressure. These tests to be conducted for the new and worn-in condition, for dry and high humidity, and at cold and hot temperatures. Expanded endurance tests: Experience indicates that increased brake structure endurance requirements are desirable for long life components in field service. An endurance test of 105,000 cycles from 50 psig to maximum brake pressure is recommended. The 105,000 cycles from 50 psig to maximum brake pressure can be used to meet the requirement of the TSO-C135.

Temperature/creep

Wheel and brake design should include temperature/creep analyses to demonstrate compliance with aircraft manufacturer requirements and endurance objectives. Examples of typical components to be analyzed include wheel rotor drive beams and brake torque tube/ backing plate assembly.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

TABLE 5.2  (Continued) Recommended practices for large transport aircraft wheels and brakes – ARP597. Recommended practices

Large transport aircraft ARP597

Static torque

This test requirement is recommended in particular for carbon brakes that exhibit lower average and minimum static friction coefficients in comparison to steel brake friction couples. This test will verify that the brakes have adequate static torque when parked and during engine run up conditions to meet the requirements of 14 CFR 25.735(d).

Dynamic torque

It is recommended that requirements include dynamic verification of wheel and brake assembly structure to an endurance spectrum taking into account the highest braking load(s) and numbers of load cycles that are expected to occur on the airplane. If a torque limiting system is used on the airplane, it is recommended that evaluations be conducted with the torque limiting system inoperative, and with failure modes as determined by the aircraft manufacturer.

6 Brake Issues and Concerns

Vibration As the brake is not a static component – it does most of its work when spinning and being loaded – it can be a significant source of noise, vibration, and dynamic disturbance. There are five main forms of brake and landing gear instability [26] that can arise: squeal, whirl, chatter, gear walk, and shimmy. The first three forms are purely brake and wheel instabilities while the latter two are landing gear instabilities that can be provoked by brake and brake control system behavior. The brake components move, shift, and deflect under load (as shown in exaggerated form in Figure 6.1) that along with the frictional behavior of the brake can lead to vibration. Squeal vibration is characterized by the torsional motion of the stators, torque tube, and the piston housing (Figure 6.2a). It typically occurs in a frequency range between 100 Hz and 20 kHz. Squeal usually occurs during landing stops, but it can also be observed during low-speed maneuvers. Higher contact pressures and higher energies increase the severity of this type of vibration. Squeal can be excited by the characteristics of the friction material and by modal coupling between axial and tangential degrees of freedom of the brake system. Whirl vibration shows up as out of plane “wobble” motion involving the brake disks, torque tube, and piston housing (Figure 6.2b). The frequency range of the whirl mode vibration is typically 100 to 300 Hz, approximately the same as the first squeal mode. Coupling between the squeal and whirl modes is often observed. Whirl is a particularly destructive vibration mode for the brake. Methods to control it include introducing hydraulic orifices between the cylinders in the piston housing; as the brake rotates in the whirl mode, it will attempt to differentially pump fluid and this ©2022 SAE International

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 6.1   Exaggerated brake positions under loading.

© SAE International.

 FIGURE 6.2   Squeal (a) and whirl (b) modes of brake vibration.

can be damped hydraulically by the orifices. When using this approach, a balance between whirl stability and sufficient hydraulic response for antiskid control must be found. Another approach is to introduce an additional support (called a pedestal) between the torque tube and the axle. This pedestal does not react significant torque but provides additional radial support to the brake, and the increased stiffness reduces the whirl effect. An example of squeal and whirl modes is shown in a power spectrum in Figure 6.3. The chatter vibration mode involves torsional motion of the rotors and wheel, typically coupled with fore and aft motion of the landing gear. Chatter is largely controlled by tire stiffness and it can occur at the end of low-speed stops. The frequency of chatter vibration is low – often in the range between 10 and 100 Hz.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

61

© SAE International.

 FIGURE 6.3   Power spectrum analysis showing whirl and squeal.

Gear walk is defined as a fore-aft motion of the landing gear, typically coupled with torsional motion of the brake rotors and wheel. Gear walk oscillations can build up to significant levels, creating passenger discomfort and have caused structural failure of landing gears. Typical gear walk vibration frequency is dominated by the structural mode of the landing gear and is often between 10 and 50 Hz. Gear walk can be induced by antiskid behavior but also by friction material induced instabilities in the brake; the avoidance of gear walk must be a primary element in antiskid tuning. The shimmy vibration mode involves a coupling between the torsional and lateral motion of the wheel/tire and the landing gear. Because of the low frequencies (10–50 Hz) and high energies involved, it can be very destructive. In many cases, the brake is not a primary contributor to shimmy, but in two wheel gears asymmetrical brake application should be considered as an initiating event for shimmy. The shimmy phenomenon is dominated by landing gear stiffness, rolling stock mass, and tire properties. Further information regarding shimmy is provided in The Design of Aircraft Landing Gear, Chapter 13. Shimmy can be excited by antiskid braking, but it is generally beyond the capability of the brakes and brake control system to stop shimmy once it has begun. As the appearance of these phenomena during equipment or aircraft testing can result in expensive delays and redesign, it is recommended that an analytical understanding of these behaviors is developed early in the aircraft development program to minimize the possibility of them occurring in any damaging fashion on the aircraft.

Failure and Degradation Modes The most obvious degradation mode for a brake is the wear mode, which is expected. For steel brakes, the level of wear per flight is highly related to the energy absorbed in the stop. For carbon brakes, the level of wear is related more closely to the number of brake applications. However, the wear and frictional behavior of a carbon brake is also strongly linked to the amount of water vapor in the atmosphere. Carbon absorbs and releases moisture and other gases depending on the temperature. A significant body of work exists studying the frictional behavior of carbon. Work by Tanner and

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Travis [27] specifically related to aircraft brakes suggests, “airline operators who consistently operate aircraft in climates with dew points lower than –7°C might experience an increase in carbon brake wear rates. Aircraft operations in climates with dew points higher than about 12°C might experience weak braking during cold taxi operations with carbon brakes due to excessive water vapor surface contamination. This weak braking condition under cold taxi conditions is sometimes referred to as morning sickness.” High temperature leads to oxidation of carbon brake disks. Oxidation results in a loss of brake mass as the solid form of carbon in the brake is converted to carbon monoxide or carbon dioxide in gaseous form. Carbon brake disk oxidation can be classified into two types: thermal oxidation and catalytic oxidation. Thermal oxidation of carbon brake disks is a chemical reaction between the carbon disk and oxygen at elevated temperatures, typically in excess of 400°C and becoming significant above 1200°C  – occurring in the absence of other oxidizers or contaminants. Oxidation of this type is most severe on the brake disks of the heat sink that typically see the highest repeated peak temperatures in service. Efforts are made in the design of the brake heat sink to minimize the rate of thermal oxidation by managing brake temperatures (increasing the mass of the brake) and by using anti-oxidation coatings that work to limit the contact between the exposed portions of the carbon and oxygen. Catalytic oxidation is thermal oxidation which is accelerated due to the presence of a catalyst. Catalysts which have a significant oxidation acceleration effect on carbon include potassium and sodium salts and oxides, salts and oxides of calcium, barium, and magnesium, and the transition and noble metals such as copper, iron, chromium, silver, platinum, and gold. While gold is not commonly in contact with wheel brakes, many of the other materials (such as potassium and sodium salts) are constituents in aircraft and runway deicing fluids. As the demand for more environmentally friendly deicing fluids has grown, the industry has shifted from urea-based compounds to acetate and formate-based compounds (typically sodium and potassium acetates and formates) that have accelerated the amount of catalytic oxidation observed. AIR5490 [28] outlines the concerns, and AIR5567 [29] provides a standard test method to determine the impact of various chemicals on carbon brakes. An indication of the severity of different chemicals and their exposure time is shown in Figure 6.4. Brakes that have had failures in the brake adjuster or in the control system can apply torque when not commanded. This “dragging brake” will lead to elevated brake temperatures and increased wear in the affected brake. If not detected, dragging brakes can lead to longer takeoff distances, potentially resulting in an unsuccessful takeoff. The use of brake temperature monitoring can aid in the detection of a dragging brake. The amount of torque generated by a dragging brake varies depending on the nature of the failure mode. In the worst case, a brake can lock, either due to mechanical failure of a brake component, failure of the brake control system, or due to freezing of the heat stack. Locked brakes will result in a skid-through failure of the associated tire, especially if the brakes are locked during touchdown and rollout. The phenomenon of frozen brakes, outlined in AIR4762 [30], is generally related to operation on contaminated runways where standing water or slush can infiltrate into the brake area. Wheel wells that are open to the atmosphere can contribute to freezing of the contaminants and jamming of the brake rotors and stators. If operations in very cold,

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© SAE International.

 FIGURE 6.4   Relative impact of deicer types with time and temperature.

contaminated environments are anticipated, then certain operational procedures (such as ensuring firm landings and considering preheating of the brakes) should be considered. Some Beechcraft aircraft (such as the 1900D) have an optional brake anti-ice system that directs warm bleed air to the brakes to ensure they are able to operate in winter conditions.

7 Braking Accessories

Brake Cooling Fans Brake cooling fans, typically powered by an electric motor, are sometimes fitted when space constraints drive the brake to operate at high temperatures or where a trade study has shown that a benefit in reduced brake size and mass is achieved with cooling fans to restore an acceptable turnaround time. Figure 7.1 shows a cutaway view of an early brake cooling fan installation – modern installations are similar in principle. Not every aircraft manufacturer chooses landing gear mounted cooling fans; some promote the use of external ground service cooling fans or prefer instead to have a higher brake mass to keep temperatures lower. Higher brake mass and lower resultant temperatures can help reduce oxidation on carbon heat sinks as at high temperatures the cooling action of the brake fan also brings fresh oxygen to the brake – while the brake cools more quickly, the presence of a higher flow rate of oxygen can increase oxidation rates until the temperature drops.

Brake Temperature Measuring Systems As the kinetic energy of the aircraft is stored in the brake as heat, brake temperature is a good indication of the ability of the brake to store additional energy. Measuring and reporting brake temperature to the aircraft systems and cockpit is important for a number of reasons: limits must be placed on the maximum brake temperature before a flight may be authorized (to ensure adequate RTO energy absorption capability is available), a dragging brake can be detected (a dragging brake will be hotter ©2022 SAE International

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 7.1   Brake cooling fan on TSR.2.

than neighboring brakes), and landing gear retraction can be inhibited if the brakes are above a safe temperature limit to be in the bay. Measuring systems typically involve using a “K-type” thermocouple (a junction of Chromel and Alumel metal wires) which is inserted near the brake heat sink. Other temperature sensors may be used but the K-type thermocouple has the advantage of simplicity in the sensor (it is simply two wires welded together, which generate a voltage proportional to the temperature at the junction) and a wide temperature measuring range of approximately –200°C to 1250°C. As thermocouples work by generating a voltage between dissimilar metals, specific alloys are used in the wire. When these wires are connected to a measuring system, the temperature at that connection must be known to permit temperature compensation (to identify the voltage difference between that generated by the sensor and that generated by the contact between the thermocouple wires and the copper of the measurement system or wiring harness. To avoid having thermocouple wire and special contacts in the landing gear wiring harness, many installations mount the temperature compensation unit near the brakes. This permits the use of standard wire and connectors in the harness but puts the compensation electronics in an environmentally challenging location. Many brake temperature measurement systems are designed in accordance with AS1145 [31]. Additional recommendations for brake temperature measurement systems are found in ARP6812 [32].

Retraction Braking For wheels fitted with wheel brakes, it is commonplace to apply a small snubbing pressure to the brakes after takeoff to decelerate the spinning wheels and tires prior to retraction. This practice reduces the risk of damage to components in the wheel bay if there is any tire damage. For wheel locations without wheel brakes, spin down frictional pads are typically installed to decelerate the tires during the retraction sequence. Figure 7.2 shows two examples – a historical approach used on the DC-3 and a modern approach implemented in the nose landing gear bay of the KC-135. Spindown brakes are typically installed in nose landing gear applications to reduce

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© SAE International.

Reprinted from http://www.af.mil/News/Article-Display/ Article/497013/new-part-to-improve-old-plane-savemoney-along-the-way/.

 FIGURE 7.2   Wheel Spindown brakes; DC-3 MLG (left) and KC-135 NLG (right).

noise and vibration from the wheels and tires. The A320 aircraft uses spring loaded belts that contact the tires and decelerate them; early versions of the A330 used a transverse bar fitted to the drag stay such that contact with the tires was made during retraction. The Fokker F100 (Figure 7.3) uses spring mounted pads, similar to the KC-135 implementation.

© SAE International.

 FIGURE 7.3   Fokker F100 nose gear showing spin down brake installation.

8 Wheels

T

he wheel is the essential interface between the tire, brake, and landing gear structure. The wheel must retain the tire inflation medium (for tubeless tires), accept the load applied by the tire, and transfer it through bearings to the axle. The wheel must also accept the torque loading applied by the brake and transfer it to the tire. Aircraft wheels are generally two piece assemblies joined by tie bolts or a lock ring. This type of assembly permits the use of tires with relatively inflexible bead wires. An overview of the typical assembly and nomenclature for a bolted wheel is shown in Figure 8.1 and for a lock ring wheel in Figure 8.2. In both those figures, wheels for braked applications are shown. These wheels offset the wheel web (the “spokes” of the wheel) to create a cavity to house the brake. In unbraked wheel applications, the spokes are usually positioned centrally in an “A frame” arrangement, which offers the most symmetrical and lightest topology, as shown in Figure 8.3 (left). Braked wheels that house the heat sink within the wheel usually have an offset arrangement, as shown in Figure 8.3 (right). Most wheels are manufactured from aluminum alloys (historically some wheels were made from magnesium to achieve a weight reduction but the flammability of magnesium has resulted in a near total ban on magnesium wheels – some magnesium wheels are still available in small sizes for light aircraft). Typical aluminum alloys include 2214, 2014, 2618, 7010, and 7050. Most modern wheels are forged and machined to final size to achieve the best material properties. Cast wheels exist, but a casting factor is applied by most certification regimes (cast wheels must withstand an ultimate load twice the limit load while forged wheels have an ultimate load 1.5 times the limit load); the result is that forged wheels offer a lighter solution at the cost of forging lead time and additional machining. Corrosion protection has historically

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 FIGURE 8.1   Bolted wheel assembly and nomenclature.

been through the use of anodization followed by paint. While anodizing of the aluminum increases its corrosion resistance, depending on the specific formulation, it also decreases the fatigue life of the material. Alternative corrosion protection schemes are being developed which reduce or eliminate the use of hexavalent chromium in the protection process as well as minimize or eliminate the fatigue abatement. The certification rules and recommended practices (Tables 5.1 and 5.2) outline the minima in terms of wheel life. However, many large aircraft manufacturers specify a wheel life of around 10,000 flights according to a rational fatigue spectrum. Most wheel structural design is dominated by fatigue loading rather than static loading (although a static load check should be performed) due to the high number of stress reversals seen by the wheel (a stress reversal every revolution of the wheel). The alloy selected for the wheel must have good fatigue performance when hot as well as when cold as the wheel sees significant temperature variation with braking. The critical areas of the wheel tend to be the flange, the tie bolt and bolt hole areas, and the spoke areas. The form of the interface with the tire is standardized by the TRA and the ETRTO. The standard form is shown in Figure 8.4; dimensions depend on the tire selected and are shown in the tire tables. Further information, including tolerances, is available in

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

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 FIGURE 8.2   Lock ring (boltless) wheel assembly and nomenclature.

Reprinted from © Safran Landing Systems.

 FIGURE 8.3   Symmetrical (left) and offset (right) wheel designs.

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Copyright 2020 by The Tire and Rim Association, Inc. Reprinted with permission of The Tire and Rim Association, Inc. Although The Tire and Rim Association, Inc. has authorized reproduction of these materials, there is no express or implied endorsement or sponsorship by The Tire and Rim Association, Inc. of either the author of this publication or of the conclusions or information contained in this publication.

 FIGURE 8.4   Standard rim contours for aircraft tires.

the TRA Aircraft Year Book [33], which is published annually. The TRA also standardizes the forms of the valves that are appropriate for use in aircraft wheels. The loads seen by the wheel depend on the tire, and the type of tire construction influences the manner in which the force from the tire bead is transferred to the wheel. Bias ply tires drive significant load outward into the flange as shown in Figure 8.5. A radial tire drives more load directly into the wheel and less into the flange. As tires age their contact zone evolves, increasing the contact area between the tire and the rim. A wheel optimized for one tire type may not adequately support another tire type; conversely, a wheel design compatible with both tire types is unlikely to be optimized for weight. The certification loads and process for wheels are linked to the brake certification process (when the wheel has a brake they are certified as a pair). Certification regulations and recommended practices are outlined in Tables 5.1 and 5.2. Large transport aircraft require a “roll on rim” test procedure for main landing gear wheels, which simulates the wheel rolling on the ground with the tire absent. Recommendations for the conduct of the test are given in ARP1786 [34]. The intention of the roll on rim test requirement is to prevent fragmentation of the wheel; the requirement drives

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 FIGURE 8.5   Bias tire forces on rim.

significant robustness into the wheel design at the cost of some additional weight. For single, two wheel, and bogie configurations, the roll load should be the wheel maximum static rated load and it should be rolled for a distance equivalent to:

d

where:

1 2 Vr 2

•• d is the roll distance in feet, up to a maximum of 15,000 feet (4572 m) •• Vr is the aircraft takeoff rotation velocity in knots at the aircraft maximum gross weight

For more complex landing gear configurations where flat or no-tire conditions result in wheel loads greater than the static rated wheel load, a rational load as determined by analysis performed on the specific gear configuration should be used in lieu of the static rated wheel load. The wheel should be rolled in the most adverse attitude that it is likely to adopt, relative to the runway, in a tire deflated or missing case. The required roll distance is to be completed without fragmentation or sudden loss of load carrying capacity. Cracks are permitted. Separation of load carrying portions of the wheel including removable flanges or flange retaining devices from the wheel is not permitted. In addition, it should be demonstrated during roll or verified by analysis that there is no interference between the wheel and brake that would inhibit free rolling of the wheel. In all multipart wheels, attention must be paid to ensuring there is good protection against fretting as micro-movement between components is almost impossible to eliminate. It is critical in bolted wheel assemblies to ensure that the bolt preload is correct. The “snug-angle” procedure outlined in ARP5481 [35] is the recommended bolt torque assembly procedure. This process involves torqueing the bolt to a predetermined angle rather than to a measured torque. A measured torque has high variability due to variation in friction in the threads and surfaces in contact whereas torqueing to a fixed, predetermined, angle provides much higher bolt preload repeatability.

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 FIGURE 8.6   Locations to apply anti-seize compound.

The bolt is coated with anti-seize compound (typically to AMS2518 [54] or MIL-PRF83483 [55]) on the thread and contacting surfaces (Figure 8.6) and snugged to a bedding in torque. The bolt is then rotated from the snugged position to its final angle. Determination of this angle is performed during the wheel development phase through instrumentation of the bolt to determine the angle to achieve the desired preload. Preload is required in the bolts to ensure the wheel halves are secure against movement and to maximize the fatigue performance of the bolts. Tie bolts are typically selected from high-strength steel or Inconel materials. As outlined in Table 5.2, designing and testing a wheel to operate safely with a tie bolt missing is good practice. Braked wheels usually include heat shields to protect the aluminum of the wheel from the intense heat of the brake. The aluminums selected typically retain usable strength to a temperature of 200°C while the brakes can be at temperatures nearly ten times that value following a maximum energy rejected takeoff. In normal use, the wheel temperature is maintained less than or equal to 200°C for energies up the “fuse plug no melt” energy. To provide some protection to the wheel, shields are fitted as shown in Figure 8.7. Heat shields are typically sheets of corrosion resistant steel fastened to the wheel structure. An example is shown in Figure 8.8.

Bearing Selection and Preload Aircraft wheel bearings are universally of the tapered roller bearing type and are mounted in the wheel and onto the axle as shown in Figure 8.9. This form of assembly is called indirect mounting by the bearing industry – it requires that the running clearance of the bearing be set externally. Clearance adjustment, or setting, is accomplished by using an adjustable nut on the axle. The selection of appropriate wheel bearings is an iterative activity; landing gear structural designers typically want a large diameter axle to optimize stiffness and reduce weight while the brake designer may seek a small inner brake diameter to maximize the amount of brake material in the wheel. The arbitration of these competing demands must keep in mind the possible wheel bearing options. Often, wheel bearings are selected from existing bearing part numbers which have been developed for aircraft wheel applications and which have shown good performance in service. Bearings made to Anti-friction Bearing

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

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 FIGURE 8.7   Braked wheel cross section.

http://www.jber.jb.mil/News/Photos/igphoto/2001864532/

 FIGURE 8.8   C-17 heatshield and drive key attachment.

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 FIGURE 8.9   Typical bearing arrangement.

Manufacturers Association (ABMA) standard/commercial precision class 2 and ISO metric class N or K are used for aircraft wheel applications. Many manufacturers have developed customized requirements for aircraft wheel applications and provide a specific part number code to designate these bearings (for instance, Timken uses code 629 for aircraft wheel bearings). In general, bearings are selected from the manufacturers’ catalogs respecting the allowable radial and thrust load limits and ensuring that the loads applied by the wheel will yield the bearing life expected by the aircraft in service. In general, wheel bearings are inspected and greased at every tire change. The overall life of the bearing is often selected to be equal to the anticipated wheel life, although in demanding applications this may not be possible and the bearings may have to be replaced during the life of the wheel. Document AIR4403 [36] provides information on the selection and application of aircraft wheel bearings. Bearing fatigue life calculations (described in bearing manufacturer catalogs) can be made to assess the static and dynamic load capacity of the bearings so that the landing gear assembly will perform adequately. Similarly, refined fatigue life calculations can also be made for a load spectrum that includes the effects of load, speed, reliability, environment, temperature, alignment, and lubrication. Consultation with a bearing manufacturer is recommended to ensure the appropriate bearings have been selected and to confirm the expected life given a defined usage spectrum. The Timken company provides a dedicated tapered roller bearing selection approach based on a bearing fatigue life estimation process outlined by Dominik [37]. The bearing industry refers to a “rating life,” Ln, which is the number of revolutions that (100–n) percent of a group of apparently identical bearings will complete or exceed. The standard rating life for the bearing industry is L10 or 90% reliability life.

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 FIGURE 8.10   Tapered roller bearing angles.

For tapered roller bearings, the Timken approach is to provide a basic load rating, Cm, which is the calculated constant load (radial or thrust) which a group of apparently identical bearings can theoretically endure for a rating life of m million revolutions. A widely used rating is defined at m = 90 million; that is, C90 is the reference for that load rating. Catalog values provide the radial load rating, C90, as well as the thrust load rating, Ca90, for each bearing. Due to the tapered shape of the bearings (Figure 8.10), a load applied radially (normal to the surface of the axle) will generate a thrust load. Tapered roller bearings are designed to carry a combination of radial and thrust loads. Manufacturer catalogs will also provide a value, K, which is the ratio of basic dynamic radial load rating to basic dynamic thrust rating in a singlerow bearing. Additional information on the calculation and selection of rolling element bearings is available in ESDU sheet 81005 [38], and particular information relevant for tapered roller bearings is provided in ESDU sheet 81037 [39]. Many large civil aircraft have wheels designed to a rolling life of 50,000 miles, and bearings are typically replaced twice throughout this lifetime (three sets of bearings are consumed in one wheel life). In this case, the minimum required bearing life corresponds to 16,670 miles. It is recommended to refer to the bearing catalog and the associated manufacturer’s engineering manual in order to aid in selection, load calculation, and life estimation. No bearing will meet its expected life if it is not adequately lubricated in service. Aircraft wheel bearings are greased, typically with greases such as those meeting MIL-PRF-81322 [40] and more recently, MIL-PRF-32014 [41]. Attention should be paid to ensuring that different greases (including greases meeting the same specification but from different manufacturers) are not mixed. Depending on the expected usage of the aircraft, consideration should be given to including grease dams or seals in the wheel to protect the bearing and grease. A number of good recommendations regarding bearing maintainability are made in ARP813. To ensure optimum bearing performance, bearings should be slightly preloaded. The preload setting is usually 30–50% of the Ca90 rating of the smaller bearing (which is usually the bearing in the outboard position). In almost all cases, the preload is achieved (and the wheel retained) by a nut that is adjusted into position and then locked. A seating torque is generally applied (while rotating the wheel) to twice the

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https://www.166aw.ang.af.mil/News/Photos/igphoto/2000266061/

 FIGURE 8.11   C-130 tapered roller wheel bearing.

 FIGURE 8.12   Three wheel nut retention schemes: split pin (left), single lock bolt (center), dual lock bolt (right).

nominal torque value. The nut is then loosened and torqued to the design value (typically set to be the lower end of the acceptable preload range). The nut is then rotated to the next available locking feature and locked. Care must be taken in the design to ensure that locking features are sufficiently numerous to ensure that the maximum preload value is not exceeded as excessive preload will reduce bearing life rapidly. On some aircraft, the wheel nut locking feature also retains equipment (such as tachometers) in the axle. Many wheel nuts and locking features are shown in Figure 8.12. Despite the best efforts of many, wheel separations resulting from incorrect torqueing and locking do occur. In the interests of improving aircraft safety, consideration should be given to fitting the port side axles with left-hand threads and the starboard side axles with right-hand threads such that when the aircraft is rolling forwards, the tendency will be to tighten the nut rather than loosen it.

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Over Temperature and Over Pressure ­Relief During a rejected takeoff or other high-energy brake event, the heat from the brakes will warm the wheel, tire, and tire inflation medium. The result is an increase in pressure in the tire at the same time as reduced strength from the wheel and tire, which will result in the explosive rupture of the wheel if not avoided. The means of avoidance is the inclusion in the wheel of fuse plugs that contain a eutectic mixture of metals designed to melt at a precise temperature. The eutectic mixture is usually selected to have a melt temperature near 200°C. Fuse plugs are designed to meet the AS707 [42] standard for performance, requiring that one device melting will permit the release of inflation gas to 50% of the initial pressure within 2 minutes. Typically more than one device is fitted per wheel. Many accidents have occurred where tires have been inflated using unregulated pressure from a nitrogen bottle (a new bottle usually contains gas at a pressure of 3000 psi (207 bar), well beyond the required burst pressure for wheels and tires). To protect against these events, wheels are now often fitted with an over pressure relief valve designed to meet the recommended practice ARP1322 [43], which calls for a valve which releases at a pressure above the expected normal service pressures (which result from normal loading and normal brake heating). The valve, when releasing, should have sufficient flow rate to ensure that no pressure increase occurs in the tire (beyond the release pressure) when the tire is inflated by an infinite volume of 3000 psi gas.

Wheel Mass While the optimum method for estimating the wheel mass is to perform a preliminary design exercise and estimate the weight based on the size and shape of the wheel as well as its constituent parts, for preliminary sizing a parametric estimate based on the rim diameter is possible:

Mass 0.0202 D 3 0.3936 D 2

3.1364 D 5.707

where D is the rim diameter in inches and mass is given in kilograms. This relationship is shown graphically in Figure 8.13. Considerable scatter in the data is to be expected as different wheel loads, roll lives, and certification requirements will drive different masses. Care should be taken with any attempt to extrapolate this data (although the curve covers the full range of aircraft wheel sizes). The data included in the plot include both nose wheel and braked main wheel masses. For a given wheel diameter, main wheels that provide internal space for brakes will weigh more than an equivalent size nose wheel as the structural load paths within the wheel are not as efficient.

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 FIGURE 8.13   Estimated wheel mass vs. rim size. 120

100

Mass (kg)

80

60

20

0

3

4

5

6

7

8

9

10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 Rim Diameter (in.)

© SAE International.

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Failure Modes Bearing Failure Appropriate maintenance and protection of wheel bearings is required to avoid bearing failure. A number of bearing failure modes (brinelling, spalling, corrosion, and heat damage) can be detected by visual inspection of the bearings. Bearings in this degraded state must be removed and replaced. For properly sized and installed bearings, inadequate lubrication is typically the precursor to bearing failure. This inadequate lubrication can result from extensive lubrication periods, ingress of water or contaminants, or incorrect lubrication. If bearing degradation is not identified, it can lead to excessive heat generation in the bearing and failure of the bearing, wheel housing, or axle. Axle failure has been known to result from overheat softening of the axle material and also from cadmium embrittlement of the axle, resulting in fast fracture of the axle. Many main landing gear axles are protected with metallic-ceramic anti-corrosion coatings to avoid concerns regarding cadmium embrittlement.

Wheel Rim Release One of the wheel failure modes that can be damaging to the aircraft is a fracture and release of a portion of the wheel flange. Figure 8.14 shows an example of a wheel flange separation. As the wheel flange is ejected by the force of the tire and its inflation pressure, it is typically propelled away from the wheel at high speed and can inflict significant damage on items it impacts. The certification of large transport aircraft

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Reprinted from Australian Transport Safety Bureau licensed under CC BY 3.0 AU.

 FIGURE 8.14   Fractured wheel rim.

requires that consideration for, and protection from, wheel flange debris is included in the design and qualification of the aircraft. This failure mode is considered to occur when the landing gear is extended. However, some rim release and wheel separation events have occurred with the landing gear retracted – in this case, the aircraft floor and pressure vessel can be compromised. The introduction of roll on rim criteria to the certification of wheels is intended, in part, to improve robustness against this type of occurrence. A model for the expected size and velocity of the wheel flange debris is provided by EASA [44]. This model, shown in Figure 8.15, requires that a 60° arc of wheel flange be considered to depart the wheel at a velocity of 100 m/s (328 ft/s). For landing gears with multiple wheels on one axle, lateral release of only the outboard wheel flange needs to be considered. For single wheel gears, release of either wheel flange is to be considered. The orientation of the debris at the point of impact is to be considered as the most damaging to the item impacted. Vertical release of wheel debris is considered to be covered by the tire models (outlined in The Design of Aircraft Landing Gear, Chapter 3, and Aircraft Tires: Key Principles for Landing Gear Design). This certification model considers effectively that the rim releases while the aircraft is rolling. Some aircraft manufacturers choose to also consider cases where the wheel is static and the complete rim detaches. This case is typically considered with the landing gear extended for all wheels and both extended and retracted for braked wheels. The logic for considering a complete rim release with the landing gear retracted is that heat from the brakes increases the tire gas pressure while simultaneously reducing the static strength properties of the wheel.

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Reprinted from © EASA CS25.734.

 FIGURE 8.15   Model for wheel flange release.

9 Brake Control

T

he control of braking spans the entire gamut from simple systems where the brake is directly and manually controlled by the pilot to electronically controlled systems that can provide fully automatic brake control, antiskid, and autonomous braking to prevent runway overruns or to target a preselected runway exit. For piloted aircraft, regardless of how the brake control system is implemented, the pilot must be presented with an interface that provides progressive control over the brake torque generated. Parking brake systems are provided to hold the aircraft stationary when shut down but also during engine testing or often in the case of rotorcraft, during landing. Early in the development of aviation, the primary pilot interface to control the brakes was not standardized. Many British aircraft of the 1930s and 1940s used a hand lever attached to the control column (the silver item behind the control yoke in Figure 9.1). Some modern ultralight aircraft continue to use hand-operated brakes. However, most of the world has standardized on foot-operated brakes (toe brakes) that are linked to the rudder pedals. In this system, translation of the rudder pedals forward and aft controls the yaw of the aircraft; rotation of the rudder pedal (the top of the pedal moving relative to the bottom) applies brake torque. Differential braking (the ability of the left rudder pedal to control torque of the left-hand landing gear wheel brakes and the right pedal to control the right-hand brakes independently, to achieve aircraft directional control) is routinely implemented in this system. A fighter aircraft cockpit is illustrated in Figure 9.2 showing rudder pedals with toe brakes. While individual implementations vary, the pedals in this picture can be seen to have a pivot near the cockpit floor, permitting the pedals to rotate individually when pressed by the pilot’s toes. ©2022 SAE International

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Reprinted from http://www.nationalmuseum.af.mil/Upcoming/Photos/ igphoto/2001258119/.

 FIGURE 9.1   Hawker Hurricane cockpit.

Reprinted from http://www.nationalmuseum.af.mil/Upcoming/Photos/ igphoto/2001508160/.

 FIGURE 9.2   Fighter aircraft cockpit showing rudder pedals.

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Braking control is concerned with how the braking command is communicated to the brake and the means of applying the brake clamping force. In some cases, the two can be the same: a mechanical system can transmit braking force from the cockpit to the brake, as in many bicycles. However, purely mechanical systems are limited to the smallest and lightest aircraft due to the limited forces that can be applied. Many systems employ a mechanism to directly generate hydraulic pressure, which is then communicated directly to the brake – these systems are widely used on light aircraft. Mechanical links from the cockpit to a metering valve that proportions pneumatic or hydraulic pressure have broad applications. Pneumatic systems appear primarily on some historical aircraft such as the Antonov An-2 and the Fairchild F-27, but the vast majority of braking systems use hydraulic power. The state of the art in braking is to communicate the braking command electrically (brake by wire) with the braking power delivered hydraulically or electrically.

Brake Control Architectures Many brake control architectures are shown graphically in the following figures, including simple single-wheel brakes for sailplanes (Figure 9.3) and the master cylinder arrangement typical of light aircraft (Figure 9.4) to more complex architectures involving brake metering and park brake valves. The master cylinder system is a hydromechanical system where the brake pedals are connected to hydraulic cylinders (as shown in Figure 9.5) that convert the applied force to hydraulic pressure which is then routed to the brakes. Differential braking is provided by having the left pedals control the left brake and the right pedals control the right brake. Park brake functionality can be provided by a valve that is engaged once the pedals are depressed (as indicated in Figure 9.5) or by a system that separately actuates the master cylinders via a cable or mechanism, or by provided latching brake pedals that are capable of being locked once depressed. Master cylinder arrangements offer mechanical simplicity but are not usually compatible with antiskid devices. Antiskid valves require the ability to release fluid

Reprinted from USAF technical order 09-45-AA-2.

 FIGURE 9.3   Sailplane wheel brake system from Frankfort TG-1A Cinema.

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 FIGURE 9.4   Master cylinder system with in-line park brake.

from the brake (and direct it to the hydraulic return line) which would quickly deplete the reservoir on a master cylinder arrangement. Powered brakes using either a dedicated or the aircraft’s hydraulic system are the usual solution. There are many possible arrangements possible. Figure 9.6 shows a powered system that is a direct extension of the master cylinder arrangement of Figure 9.4. An accumulator provides stored energy in the case of the loss of the supply pressure (to permit emergency braking). A metering valve is used to provide pressure proportional to the pedal position to the brake cavities. The park brake valve locks pressurized fluid in the brakes. An optional antiskid valve dumps fluid to return in the event of skidding (under control of a skid control system). The park brake valve can be controlled in a variety of ways. An arrangement providing more redundancy is shown in Figure 9.7. This system employs shuttle valves near the brake to permit the brake to be pressurized by two different hydraulic systems. The primary system is similar to that discussed previously, but the park brake system is moved to an alternate source of supply. The park brake valve can either be an on/off type valve or it can be a proportional valve permitting manual modulation of the brakes in the event of a failure of the primary system. To provide a high level of braking availability in the event of system failures, a dual independent system such as that shown in Figure 9.8 can be fitted. This type of system duplicates most functionality. Shuttle valves near or on the brakes ensure the brake is fed pressure from whichever system is providing the higher pressure. An implementation similar to this is provided on the Boeing 767 – the hydraulic circuit for this braking system is shown in Figure 9.9. The fitting of an accumulator

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Reprinted with permmission from ©Diamond Aircraft Industries, Inc.

 FIGURE 9.5   Pedal and master cylinder arrangement from Diamond DA20.

downstream of the parking brake valve permits parking brake pressure to be maintained despite variations in temperature and small amounts of system leakage. Some designs prefer to avoid fluid mixing between hydraulic systems. In these systems, dual cavity brakes are used where two sets of brake pistons are provided in the brake, either of which is capable of providing the full clamping pressure. Examples of these system types are shown in Figures 9.10 and 9.11. The A320 is an example of a dual cavity brake system with a variety of operational and alternate modes. Brake by wire systems, where the braking command is transmitted electronically, are shown in Figures 9.12 and 9.13. Brake by wire systems can have a variety of implementations and typically result in a reduced number of components compared to hydromechanical or brake by wire with hydromechanical backup arrangements, especially when implementing features such as autobrake. Further information on brake by wire systems is available in AIR5372 [45], and information on parking brake implementations is available in AIR6441 [46]. Systems with electrically actuated brakes are conceptually similar to dual cavity brakes – with multiple independent actuators on a brake, a group of them can be driven from one power supply and control system while the other group is driven from a separate

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 FIGURE 9.6   Powered system with in-line park brake.

© SAE International.

 FIGURE 9.7   Powered system with independent emergency/park brake.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 9.8   Dual independent system.

© SAE International.

 FIGURE 9.9   B767 brake system hydraulic schematic.

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 FIGURE 9.10   Isolated dual independent system.

© SAE International.

 FIGURE 9.11   Brake by wire with dual cavity brakes and hydromechanical backup.

Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

© SAE International.

 FIGURE 9.12   Brake by wire system, single cavity brakes.

© SAE International.

 FIGURE 9.13   Brake by wire system, dual cavity brakes.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

supply and control system. To enhance availability, each actuator is typically designed with overdrive capability such that full braking can be delivered with a single unit on each brake failed, which can permit aircraft dispatch with a single actuator inactive. Operation from 28 VDC buses ensures that the system can operate on battery power alone. As sufficiently powerful motors running on 28 VDC would be large and heavy, some systems step the 28-V input up to high voltage (±270 VDC, for instance) to more efficiently drive the brushless direct current motors mounted on the brakes.

Antiskid and Related Functions There are a number of different antiskid systems that have been developed and implemented. In large aircraft, especially those with powered brake systems, it is difficult or impossible for the pilot to detect whether one or more tires are skidding. A skidding tire can rapidly lead to tire failure and early systems detected skidding and alerted the pilot so that brake pressure could be released manually. Systems that automatically adjusted brake pressure were developed, starting with hydromechanical systems such as Dunlop’s Maxaret that used a flywheel running against the braked wheel. The flywheel is connected to a clutch arrangement such that when the braked wheel decelerates sharply, the internals of the Maxaret unit continue to rotate and actuate a valve to relieve the brake pressure. Figure 9.14 illustrates the principle. An installation of a Maxaret unit is shown in Figure 9.15.

© SAE International.

 FIGURE 9.14   Maxaret antiskid unit.

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© SAE International.

 FIGURE 9.15   Maxaret unit on CF-100 main wheel.

Electromechanical adaptations of the system such as the Hydro-Aire Hytrol system and the Messier Ministop were developed which replaced the mechanical actuation of the valves with electrical control. Placing these systems in the axle and driving the system mechanically through the hubcap resolved issues of icing which could cause reliability problems with the Maxaret. These systems were all of the “on-off” type – metered pressure was either applied to the brakes or relieved when a skid was detected. While these systems resulted in improved braking on wet or contaminated surfaces, the focus of further antiskid development was on ensuring optimal stopping performance. Systems that modulate the brake pressure, rather than release it completely, were developed – typically using a tachometer to drive an electrical signal that was used to modulate a release valve. This type of modulating system provided a significant improvement over the “on-off” systems. Latter developments are systems that continuously seek the maximum deceleration force available from the tires. These adaptive systems form the state of the art for antiskid brake control. Example behaviors of the three different types of system are shown in Figure 9.16. Although these results are for different aircraft on different runways, the difference in performance is clear: the on-off system has deeper skids (larger changes in wheel speed) and larger excursions in brake pressure while the modulating system improves both; the adaptive system achieves the smallest excursions in both pressure and wheel speed. This behavior results in the best braking efficiency.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

 FIGURE 9.16   Wheel speed and brake pressure examples for wet runway braking -

© SAE International.

ON/OFF (top), Modulating (middle), and Adaptive (bottom).

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Braking Efficiency When specifying or discussing antiskid braking, the braking efficiency is an important parameter. The certification regulation for large transport aircraft explicitly requires the use of braking efficiency in addition to the prescribed tire to ground friction coefficients to calculate wet runway stopping distances. While the calculation of braking efficiency can be complicated, it can be considered simply as the fraction of the available tire to ground friction which can be utilized. An antiskid control system that succeeded in maintaining the tire at the point of maximum friction throughout the stop, with no variation, would achieve a braking efficiency of 100%. Many adaptive antiskid systems achieve efficiencies above 90%. In the absence of determining antiskid efficiency through flight testing, the regulation [47] dictates an efficiency of 0.3 for on-off systems, 0.5 for modulating systems, and 0.8 for adaptive systems. A number of methods to calculate efficiency based on flight test data or dynamometer testing are given in document AIR1739 [48]. While the certification regulations require a demonstration of antiskid efficiency through flight testing, most development of antiskid routines occurs using simulation employing hardware in the loop testing – this method is faster and less expensive than attempting to develop the antiskid routine on the aircraft. Some automatic tuning routines have been developed to achieve optimum performance. Guidance for the determination of the antiskid efficiency through flight testing is provided in Advisory Circular 25-7C [49].

Antiskid Dynamics The objective of an adaptive antiskid system is to rapidly control the brake torque (by modulating the brake clamping pressure) to maintain the tire slip at the point of maximum tire to ground friction. This point is at approximately 6–20% slip, depending on the specifics of the tire. The notional shape of the curve of friction developed versus tire slip is shown in Figure 9.17; this is a notional curve because the shape and location

© SAE International.

 FIGURE 9.17   Notional μ-slip curve.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

of the peak are unknown and variable. The specific shape depends on the type of tire, construction of the tire, speed, runway properties, and more. Some tires exhibit a relatively flat curve with an ill-defined peak. An adaptive antiskid system must continuously seek the peak, and the only way to determine the peak is to increase the wheel slip and monitor for the increase and then decrease of the resistive force. The controller then continually repeats this, aiming to minimize the disturbance away from the peak. Operating too frequently on the “backside” of the curve leads to deep skids, tire wear (and possible blowouts), loss of cornering power (aircraft directional control), and passenger discomfort. However, operating too far on the “frontside” of the curve (permitting the wheel speed to approach synchronous speed) leads to very low braking efficiency due to the sharp drop in friction coefficient. Operation in this region results from releasing the clamping pressure too much (which is what occurs in on-off systems). Fast response from the braking controller and from the brake is required to achieve optimum performance, with modern systems having a frequency response in the range of 6–12 Hz (and sometimes achieving response frequencies up to 20 Hz). The step response of the system should also be confirmed to be appropriate. With hydraulic brakes, the hydraulic fluid must flow into the piston housing to displace the pistons, taking up the gap between the pistons and the heat stack and then compressing the heat stack. An example of this behavior is shown in Figure 9.18 (the inflection point in the curves is due to the automatic wear adjuster). To achieve the optimum performance, the brake volume should be minimized. In addition, the length of hydraulic lines between the control valves and the brakes should be minimized as the fluid entering and leaving the brake must travel through

© SAE International.

 FIGURE 9.18   Example brake pressure vs. volume plot.

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the lines, leading to losses as well as the potential for water hammer effects. Minimizing the release time of the brake is a key element to providing the potential for good antiskid performance. Release times of less than 40 ms are needed for highly efficient braking control [50] and to provide robustness against changing frictional surfaces. Release times of less than 60 ms can provide acceptable performance; caution should be taken with release times greater than 70 ms. In general, longer release times lead to deeper skids, which decrease braking efficiency. Electric actuation systems offer the potential for very rapid brake release and actuation. SAE document AIR1739 provides greater detail on the dynamics, design options, and pitfalls for antiskid braking; document ARP1070 [51] provides the industry recommended practices for ensuring antiskid systems provide the desired functionality and integration with the aircraft. Achieving the optimum antiskid efficiency typically requires individual wheel control. In this architecture, every wheel brake is provided with its own wheel speed sensor and brake control circuit. An alternative approach is paired wheel braking, where each pair of wheels is controlled by a single valve or single control signal. This can provide a lower cost implementation (fewer valves) at the cost of some efficiency (as both wheels in the pair will have brake torque reduced during a skid even if one of them is capable of delivering more torque). This type of architecture has been employed to ensure adequate directional control in cases where individual wheel control can produce significant yaw moments and as an alternate system for aircraft with multiple braked wheels to reduce the number of required components in the brake control system. Most modern antiskid systems incorporate a locked-wheel protection feature. The wheel speed signals on a number of wheels are used to produce a reference signal that the aircraft is rolling on the ground. In the event one of the wheels should lock up, a comparison of that wheel speed with the reference signal provides a basis for the release of the brake pressure of the slow wheel. A reference memory may be provided to maintain the reference should all the wheels in a comparison group lock up at the same time. Typical comparison groups are: inboard and outboard wheels across the aircraft, adjacent wheels on a landing gear, and fore/aft wheel pairs on a four-wheel bogie. Locked wheel protection can protect against hydroplaning where the friction characteristics could otherwise fool the brake control into slowly controlling the wheel into a deep, prolonged skid. Touchdown protection is usually incorporated into antiskid brake control to ensure that braked wheels are free to rotate at touchdown even if the pilot has inadvertently commanded braking. This protection is normally implemented using the wheel speed signal and a signal from the aircraft’s weight on wheels logic system. If the aircraft is weight off wheels and the wheel speed is lower than a predefined level, a full brake release signal is applied to the antiskid valve. Transition from weight off wheels to weight on wheels indicates that the braking system can permit brake pressure, usually following a time delay to permit spin-up of the wheels during landing. Measured wheel speeds higher than a predetermined threshold also indicate that braking is permitted, even if the aircraft systems are still indicating weight off wheels.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

Antiskid Hardware Antiskid braking control requires electronic control (either analog or digital), a source of wheel speed information, a source of aircraft ground speed information (this is sometimes derived from wheel speed information), and the means to control the brake clamping force. For hydraulic brakes, a servo valve is usually employed to provide proportional pressure control. An example valve is shown in Figure 9.19. The image on the left of Figure 9.19 is a view of a first- and second-stage antiskid valve cartridge. On the right is a sectional view of the first- and second-valve stages. A typical antiskid valve is a two-stage valve with a flapper/nozzle first stage and spool and sleeve second stage. A permanent magnet torque motor in the first stage, driven by an applied current, operates the flapper. Three-way and four-way valves are both used in antiskid systems. With a three-way valve, the application of an electrical signal to the torque motor from the skid control avionics results in the flapper moving from its zero command position against the return nozzle (maximum pressure). Movement of the flapper unbalances the control chamber pressure, with the result that a command pressure is applied to the second stage spool. The forces on the spool act to position it until an equilibrium position is reached, and the feedback brake port pressure balances the command pressure. The output of the antiskid valve provides the control pressure to the brakes. In a classical antiskid system, movement of the flapper from the zero command position serves to reduce pilot metered pressure

© SAE International.

 FIGURE 9.19   Antiskid servo valve.

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to the brake. Due to the desire to fail in a fashion that retains braking capability, the valve is normally designed to port the pressure commanded by the pilot to the brake when no current is present. As current is increased, the valve reduces the brake pressure. In a brake by wire system, brake pedal position is communicated to a brake controller that provides automatic braking, antiskid, and other functions. For these systems, the functions of pressure application and reduction are usually provided by a single valve. In this case, movement of the flapper from the relaxed position typically serves to increase pressure to the brake: increasing current to the valve increases brake pressure. Brake by wire systems usually require a method to ensure that pressure is not applied at times when it is not commanded. A pressure transducer included at some point in the brake lines can be used to detect pressure different than that commanded. A solenoid shut off valve may be used to remove pressure from the system if braking pressure greater than commanded is detected. Direct drive antiskid valves drives also exist, in which the first stage is replaced by an electric motor which drives the second stage spool directly. Many aircraft systems use both types of valves with the servo valve on the primary system and the direct drive valve on the alternate system. This architecture provides benefits in that dissimilarity between brake control circuits avoids common mode failures. A key element in antiskid systems is the wheel speed sensor. It is this sensor that is used to compare each wheel’s tangential speed to the aircraft’s ground speed (to determine the amount of tire slip). In most cases, the wheel speed sensor is mounted in the hollow axle of the landing gear with the rotor shaft of the sensor connected to a hub cap on the wheel. Historical wheel speed sensors were direct current generators – producing a voltage proportional to wheel speed. To resolve signal noise and reliability concerns, these direct current systems were replaced by alternating current sensors that produce a signal with a frequency proportional to wheel speed. The sensor shaft must be connected mechanically to the hub cap. This is often performed with a splined shaft, bellows coupling, or “dog bone” coupling. With both direct and alternating current sensors, there is typically a speed below which there is insufficient signal to perform antiskid – often around 20 knots. Care must be taken with the mechanical coupling to the wheel speed transducer to ensure that flexibility does not introduce spurious signals that can disrupt antiskid control. An approach to providing non-contact speed sensing that is used frequently in automotive anti-lock braking systems is to use a sensor and exciter ring arrangement or a Hall Effect sensor with a series of magnets. There is a recent interest in antiskid control to locate as much of the control electronics close to the brakes to reduce landing gear wiring and to interface more efficiently with distributed avionics systems. The data rate achievable on most avionics communications busses is not sufficiently high to permit antiskid control loops to operate, so dedicated electronics are usually provided. An example of an antiskid system that is integrated with Hall Effect sensors and located in the axle is shown in Figure 9.20. This system is indicative of that installed on the Boeing 787 aircraft where the Axle Remote Data Concentrator (ARDC) provides wheel speed sensing, antiskid control, and communication to the aircraft data bus. Adaptive antiskid systems require an aircraft speed reference signal. In early adaptive systems, this was provided by measuring the wheel speed of an unbraked wheel. On Concorde, the nose landing gear wheel speeds were measured as the

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

 FIGURE 9.20   Wheel speed sensing and brake control electronics mounted

© SAE International.

in hubcap.

© SAE International.

 FIGURE 9.21   XB-70 main landing gear showing reference speed sensing wheel.

reference. The XB-70 supersonic bomber employed a dedicated sensing wheel among the four main landing gear wheels (Figure 9.21). On modern systems, the reference speed is typically determined directly from the braked wheel during the spin-up phase following each brake clamping pressure release. Another option is to take the reference speed from the aircraft inertial navigation system.

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Autobrake To provide automatic braking upon landing or in the event of a rejected takeoff event, autobrake systems have been developed. These systems provide a constant (or scheduled) deceleration rate during the landing roll and are capable of applying full braking as soon as the throttles are retarded during a rejected takeoff. The systems reduce pilot workload and provide more consistent braking. A cockpit interface similar to those shown in Figure 9.22 is provided. ARP4102/2 [52] recommends that the minimum deceleration level be approximately 0.1–0.2 g and the medium deceleration level be 0.2–0.3 g. The maximum level commands the maximum braking torque available. The recommended conditions for arming and disarming the system and for the activation of the brakes are also included in ARP4102/2. Further recommended practices related to the implementation of automatic braking are included in ARP1907 [53]. The existence of automatic braking capability on aircraft has led to further functions such as “Brake to Vacate” where the desired runway turnoff is identified prior to landing. The avionics then command the required deceleration through the autobrake system to arrive at the turnoff at the correct exit speed. Another feature enabled by autobrake is “Runway Overrun Protection” that uses the avionic system’s airport database to identify the remaining runway distance available. If the system predicts an impending runway departure at the selected or manually applied braking level, then it will apply higher autobraking to ensure the aircraft stays on the runway.

Failure Modes Braking systems can be complex – a complete failure mode and effects analysis should be completed to understand the possible ways that failures can occur and their impact on the aircraft. In well-designed systems, some failures can arise due to maintenance

© SAE International.

 FIGURE 9.22   Autobrake cockpit control options.

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Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design

or other issues. To the greatest extent possible, these should be protected against by inherent design features. One issue which can lead to a loss of braking is cross wiring of the wheel speed sensors. For individual wheel control systems on two wheel or multiple wheel landing gears, if the inboard and outboard wheel speed sensors are cross wired, then the antiskid system can relieve pressure on a non-skidding wheel while driving a skidding wheel further into skid or into lock. This type of failure has been known to occur and lead to runway excursions. As the wheel speed sensors themselves are typically the same part number at all locations, it is difficult to prevent this type of occurrence. Providing different part number sensors with unique connectors is a robust but expensive solution. Mechanically limiting the length of wiring harnesses is a partial solution (it relies on the harness being installed correctly). A visual indication (color coding, etc.) does not provide a solution but can augment other maintenance procedures to reduce these occurrences. A related failure type is a cross connection of hydraulic fittings. Depending on where the system is cross connected, it can lead to the same failure types as cross connecting the wheel speed sensors. Avoidance of this failure mode is typically performed by mechanically limiting the routing of hydraulic hoses and pipes. Where hoses could be cross connected, they should be protected against incorrect attachment through the use of different fitting sizes. Of key concern in the design of braking systems is ensuring that uncommanded braking is not applied during takeoff, either through the normal brake control system or through the parking brake system. The application of braking during takeoff can result in the aircraft not becoming airborne and departing the runway at high speed – which is usually fatal. An outline of the system safety analysis process which should be utilized in the design of braking and other complex systems is found in The Design of Aircraft Landing Gear, Chapter 13.

References



1. Foster, B. “Undercarriages,” Flight, February 8, 1940, 131.



2. Schmidt, R.K., The Design of Aircraft Landing Gear (Warrendale, PA: SAE International, 2020)



3. Aerospace Information Report, “Aerospace Landing Gear Systems Terminology,” AIR1489, Revision C, SAE International, May 2017.



4. Knacke, T.W., Parachute Recovery Systems Design Manual, AD-A247-666 (China Lake, California: Naval Weapons Center, March 1991).



5. Aerodynamics of Parachutes, ESDU Data Sheet 09012, Engineering Sciences Data Unit, August 2009.



6. Aerospace Recommended Practice, “Arresting Hook Installation, Land-Based Aircraft,” ARP1538, Revision B, SAE International, April 2013.



7. MIL-STD-18717(AS), Design Criteria for Naval Aircraft Arresting Hook Systems.



8. Highley, F.H., “An Analysis and Evaluation of Unconventional Methods of Passenger Car Braking,” SAE Technical Paper 760790, October 1976, https://doi.org/10.4271/760790.



9. Bailey, D.A., “Investigation of Improvements in Aircraft Braking Design,” PhD thesis, Cranfield University, October 2004.

10. Department of the Navy, “General Specification, Performance, Design Characteristics and Construction of Aircraft Weapon Systems,” SD-24M, Naval Air Systems Command, Arlington, VA, February 18, 1994. 11. Creech, D.E., “Aircraft Brake Energy Analysis Procedures,” Technical Report ASD-TR-68-56, Aeronautical Systems Division, US Air Force, October 1968. 12. Dowden, J., “Aircraft Braking Energy Model,” MSc thesis, Cranfield University, August 2015.

©2022 SAE International

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13. European Aviation Safety Agency, “Certification Specifications and Acceptable Means of Compliance for Large Aeroplanes,” CS-25, Amendment 19, European Aviation Safety Agency, May 2017. 14. Aerospace Standard, “Minimum Performance Standard for Parts 23, 27, and 29 Aircraft Wheels, Brakes, and Wheel and Brake Assemblies,” AS5714, Revision A, SAE International, August 2017. 15. Technical Standard Order, “Transport Airplane Wheels and Wheel and Brake Assemblies,” TSO-C135a, Federal Aviation Administration, July 2009. 16. Sforza, P.M.Commercial Airplane Design Principles, Butterworth-Heinemann, Oxford, 2014. 17. Gilson, R.D., Beryllium Brake Experience on the C-5A Airplane, SAE Technical Paper 710427, 1971, https://doi.org/10.4271/710427. 18. Devi, G.R. and Rao, K.R., “Carbon-Carbon Composites – An Overview,” Defence Science Journal, 43, no. 4 (October 1993): 369–383. 19. Aerospace Information Report, “Information on Electric Brakes,” AIR5937, SAE International, May 2012. 20. Technical Standard Order, “Aircraft Wheels, Brakes and Wheel/Brake Assemblies for Parts 23, 27 and 29 Aircraft,” TSO-C26d, Federal Aviation Administration, October 2004. 21. Aerospace Standard, “Minimum Performance Standard for Part 25 Transport Aircraft Wheels, Brakes, and Wheel and Brake Assemblies,” AS6410, Revision Draft, SAE International, October 2019. 22. Military Specification, “Wheel and Brake Assemblies,” Aircraft General Specification for, MIL-W-5013L, October 1991. 23. Aerospace Recommended Practice, “Wheel and Hydraulically Actuated Brake Design and Test Requirements for Military Aircraft,” ARP1493, Revision C, SAE International, November 2013. 24. Aerospace Recommended Practice, “Wheels and Brakes, Supplementary Criteria for Design Endurance Civil Transport Aircraft,” ARP597, Revision E, SAE International, May 2015. 25. Aerospace Recommended Practice, “Maintainability Recommendations for Aircraft Wheel and Hydraulically Actuated Brake Design,” ARP813, Revision C, SAE International, April 2012. 26. Hamzeh, O.N, Tworzydlo, W.W., Chang, H.J, and Fryska, S.T., “Analysis of Friction-Induced Instabilities in a Simplified Aircraft Brake,” SAE Technical Paper 1999-01-3404, September 1999, https://doi.org/10.4271/1999-01-3404. 27. Tanner, J.A. and Travis, M., “Adsorption and Desorption Effects on Carbon Brake Material Friction and Wear Characteristics,” SAE Technical Paper 2005-01-3436, October 2005, https:// doi.org/10.4271/2005-01-3436. 28. Aerospace Information Report, “Carbon Brake Contamination and Oxidation,” AIR5490, Revision A, SAE International, April 2016. 29. Aerospace Information Report, “Test Method for Catalytic Carbon Brake Disk Oxidation,” AIR5567, Revision A, SAE International, August 2015. 30. Aerospace Information Report, “Compilation of Freezing Brake Experience and Potential Designs and Operating Procedures to Prevent Its Occurrence,” AIR4762, Revision A, SAE International, May 2016. 31. Aerospace Standard, “Aircraft Brake Temperature Monitor Systems (BTMS),” AS1145, Revision C, SAE International, September 2016.

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32. Aerospace Recommended Practice, “Aircraft Brake Temperature Measurement,” ARP6812, Draft Revision, SAE International, April 2019. 33. The Tire and Rim Association, 2018 Aircraft Year Book, The Tire and Rim Association, Inc., Copley, Ohio, 2018. 34. Aerospace Recommended Practice, “Wheel Roll on Rim Criteria for Aircraft Applications,” ARP1786, Revision C, SAE International, September 2013. 35. Aerospace Recommended Practice, “Recommended Wheel Tie Bolt Preload Procedure,” ARP5481, Revision A, SAE International, August 2018. 36. Aerospace Information Report, “Selection, Testing, Lubrication, and Sealing of Single Row Tapered Roller Bearings for Aerospace Wheel Applications,” AIR4403, Revision B, SAE International, May 2017. 37. Dominik, W.K., “Rating and Life Formulas for Tapered Roller Bearings,” SAE Technical Paper 841121, September 1984, https://doi.org/10.4271/841121. 38. Engineering Sciences Data Unit, “Designing with Rolling Bearings. Part 1: Design Considerations in Rolling Bearing Selection with Particular Reference to Single Row Radial and Cylindrical Roller Bearings,” ESDU Item 81005, Amendment A, June 1982. 39. Engineering Sciences Data Unit, “Designing with Rolling Bearings. Part 2: Selection of Single Row Angular Contact Ball, Tapered Roller and Spherical Roller Bearings,” ESDU Item 81037, Amendment A, June 1982. 40. Performance Specification, “Grease, Aircraft, General Purpose, Wide Temperature Range,” NATO Code G-395, MIL-PRF-81322G, Department of Defense, January 2005. 41. Performance Specification, “Grease, Aircraft and Instrument,” MIL-PRF-32014A, Department of Defense, September 2006. 42. Aerospace Standard, “Thermal Sensitive Inflation Pressure Release Devices for Tubeless Aircraft Wheels,” AS707, Revision C, SAE International, November 2011. 43. Aerospace Recommended Practice, “Overpressurization Release Devices,” ARP1322, Revision B, SAE International, August 2014. 44. AMC25.734, “Protection against Wheel and Tyre Failures,” Certification Specifications and Acceptable Means of Compliance for Large Aeroplanes, CS-25, Amendment 19, European Aviation Safety Agency, May 2017. 45. Aerospace Information Report, “Information on Brake-By-Wire (BBW) Brake Control Systems,” AIR5372, Revision A, SAE International, July 2014. 46. Aerospace Information Report, “Information on Parking Brake Systems,” AIR6441, SAE International, July 2015. 47. CS25.109(c)(2), “Certification Specifications and Acceptable Means of Compliance for Large Aeroplanes,” CS-25, European Aviation Safety Agency, Amendment 19, May 2017. 48. Aerospace Information Report, “Information on Antiskid Systems,” AIR1739, Revision B, SAE International, November 2016. 49. Advisory Circular, “Flight Test Guide for Certification of Transport Category Airplanes,” AC25-7C, Federal Aviation Administration, October 2012. 50. Butterfield, G., “Scaling a Skid - The Impacts of Tire/Wheel/Brake Size and Hydraulic System Design on Skid Control Performance,” Presentation to the A-5 Committee on Aircraft Landing Gear, September 2017.

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References

51. Aerospace Recommended Practice, “Design and Testing of Antiskid Brake Control Systems for Total Aircraft Compatibility,” ARP1070, Revision E, SAE International, July 2019. 52. Aerospace Recommended Practice, “Automatic Braking System (ABS),” ARP4102/2, Revision A, SAE International, July 2012. 53. Aerospace Recommended Practice, “Automatic Braking Systems,” ARP1907, Revision C, SAE International, January 2016. 54. Aerospace Material Specification, Thread Compound, Anti-Seize, Graphite-Petrolatum, AMS2518, Revision D, SAE International, February 3, 2012. 55. MIL-PRF-83483E, Performance Specification:, Thread Compound, Antiseize, Molybdenum Disulfide-Petrolatum, July 25, 2014.

Index

A

Aircraft actuatable braking mechanisms, xxiii air cushion landing system, on XC-8A, xxiii brake control system, xviii brakes (See Brakes) caterpillar track designs, xxi Convair B-58 main landing gear, xxi Convair XB-36 main wheel, xx deceleration (See Deceleration) Dowty’s internally sprung wheel, xviii, xx function, xvii–xviii landing gear, xvii no-power parking mechanisms, xxiii pneumatic tire, with friction wheel brake, xxiv retro-rocket installation, of XFC-130H/ YMC-130H, xxiv shock cord arrangement, on Bleriot XI reproduction, xix single tire per landing gear, xix sprung wire wheels, xviii streamlined form, for fixed gear, xviii tracked systems, xx track main landing gear of B-50, xxii of XB-36, xxii turbo brake, xxiv Westland Lysander, xx Wrights’ aircraft, xix, xviii XB-36 with tracked landing gears, xx Anti-friction Bearing Manufacturers Association (ABMA), 74, 76 Axle Remote Data Concentrator (ARDC), 99

B

B767 brake system hydraulic schematic, 89 Beechcraft aircraft, 63 Beryllium, 24, 25 Braked wheel cross section, 75 Brakes beryllium, 2 carbon brakes, 3, 6 clamping force, 2 design (See Designs, in brakes) electric brake from Boeing 787, 7 expander tube brake, 1, 3 frictional force, 1 functions, 1 hydraulic pneumatic pressure, 1, 4 issues and concerns deicer types, with time and temperature, 63 exaggerated brake positions under loading, 60 failure and degradation modes, 61–63 vibration, 59–61 multiple disk brakes, 2 nose wheel brakes, 7 reinforced carbon, 3, 6 Saab Gripen, 7 single disk brakes, 2 sizing in (See Sizing, in brakes) sleeves, 36 steel brake cross section, 5 steel brake linings, stator of KC-135 brake, 5 steel brakes, 3 types, 1

107

108

Index

C

Carbon brakes, 61, 62 Catalytic oxidation, 62 Chatter vibration mode, 60 C-17 heatshield, 75 Control system, in brakes antiskid and related functions antiskid dynamics, 95–97 antiskid hardware, 98–100 autobrake, 101 braking efficiency, 95 failure modes, 101–102 architectures, 85–92 autobrake cockpit control options, 101 differential braking, 83, 85 with dual cavity brakes and hydromechanical backup, 90 hand-operated brakes, 83 Hawker Hurricane cockpit, 84 mechanical systems, 85 piloted aircraft, 83 pneumatic systems, 85 rudder pedals, 83, 84 Cooling fans, brakes higher brake mass, 65 lower resultant temperatures, 65 retraction braking, 66–67 temperature measuring systems, 65–66 Cross-design parachute, 11 C-130 tapered roller wheel bearing, 78 D

Deceleration airframe drag, 9 arresting hooks, 12, 13 BAK-13 arresting system maximum hook load, 13 B-1B Lancer braking during landing, 10 braking parachutes, 10 drag force, 12 typical arrangement of, 11 Dowty drag brake, 13, 14 dynamic coefficient of friction, of braking material, 13 F-16 decelerating with arresting system, 12 Il-62 landing, thrust reversers deployed in, 10

jet-powered civil transport aircraft, 9 military aircraft, 9 reverse thrust systems, 9 tactical aircraft, 10 Tu-160 strategic bomber, decelerating with braking parachutes, 11 wheel stowage volume, 10 De Havilland DHC-6 Twin Otter, 17 Designs, in brakes brake actuation, 30–34 brake efficiency variation, during a stop, 30 carbon brake cross section, 27 disk size estimation, 28 carbon–carbon composite, 24 clamping force, 29, 30 common brake materials, properties of, 24 friction interface, 23 heat shielding, 27 landing gear structure, mechanical connection to, 34–36 lining power, 29 mass loadings, 25, 26 material strength variation, with temperature, 24 multiple disk brake, 27 rejected takeoff energy, 25, 26 size, brakes for civil aircraft, 37–42 torque generation, 29 wear rates, brake materials, 26 weight, 36–37 Differential braking, 83, 85 Dowty drag brake, 13, 14 “Dragging brake,” 62 Dual cavity hydraulic brake, 35 E

Electromechanical actuator (EMA), 33 F

Fokker F100 nose gear, spin down brake installation, 67 Fractured wheel rim, 81 Frozen brakes, 62 G

Gear walk, 61

Index

H

Helicopter park brake force requirement, 21 Hydraulic actuation, 30 Hydro-Aire Hytrol system, 93 I

Ineffective pressure, 32 K

K-type thermocouple, 66 L

Locked brakes, 62 M

Maxaret antiskid unit, 92 on CF-100 main wheel, 93 Messier Ministop, 93 P

Pedestal, 60 Powered brakes independent emergency/park brake, 88 with in-line park brake, 88 R

Ribbon parachute, 11 Ringslot parachute, 11 S

Sailplane wheel brake system, Frankfort TG-1A Cinema, 85 Shimmy vibration mode, 61 Sizing, in brakes energy brake temperature indication systems, 16 flapless landing case, 16 kinetic energy calculation, 15–17 large aircraft example, 18 light aircraft example, 17 overweight landing case, 16 rational brake energy calculation, 18–19 rejected takeoff design baseline energy, 15 main landing gear wheels, 15 torque, 20–22 Squeal vibration, 59, 60 power spectrum analysis, 61 Steel brakes, 25, 61

109

T

Tapered roller bearing angles, 77 Thermal oxidation, 62 Three wheel nut retention schemes, 78 Timken approach, 76, 77 Torque, 20–22 W

Wheels “A frame” arrangement, 69 anti-seize compound, locations to apply, 74 bearing selection and preload, 74–78 bias ply tires, 72, 73 bolted wheel assembly and nomenclature, 70 and brake certification criteria, 44–56 large transport aircraft wheels and brakes, ARP597, 57–58 military standard, 43 TSO-C26d, 43 braked wheels, 74 corrosion protection, 69–70 ETRTO, 70 failure modes bearing failure, 80 wheel rim release, 80–82 forged wheels, 69 “fuse plug no melt” energy, 74 lock ring (boltless) wheel assembly and nomenclature, 71 magnesium wheels, 69 over temperature and over pressure relief, 79 “roll on rim” test procedure, 72 “snug-angle” procedure, 73 standard rim contours, for aircraft tires, 72 symmetrical and offset wheel designs, 71 tie bolts, 74 tire inflation medium, 69 TRA, 70, 72 wheel mass, 79–80 Wheel spindown brakes, 66, 67 Whirl vibration, 59, 60 power spectrum analysis, 61

Aircraft Wheels, Brakes, and Brake Controls Key Principles for Landing Gear Design R. Kyle Schmidt

The author’s two volume treatise, The Design of Aircraft Landing, was the inspiration for this book. The Design of Aircraft Landing is a landmark work for the industry and utilizes over 1,000 pages to present a complete, in-depth study of each component that must considered when designing an aircraft’s landing gear. While recognizing that not everyone may need the entire treatise, Aircraft Wheels, Brakes, and Brake Controls: Key Principles for Landing Gear Design is one of three quick reference guides focusing on one key element of aircraft design and landing gear design. This volume features an overview of brakes, aircraft deceleration, brake sizing, brake design, braking accessories, wheels, brake control as well as brake issues and concerns. R. Kyle Schmidt has over 25 years’ experience across three countries and has held a variety of engineering roles relating to the development of new landing gears and the sustainment of existing landing gears in service.

RELATED RESOURCES BY R. KYLE SCHMIDT: The Design of Aircraft Landing Gear 978-0-7680-9942-3 Aircraft Tires: Key Principles for Landing Gear Design 978-1-4686-0463-4

Airfield Compatibility: Key Principles for Landing Gear Design 978-1-4686-0466-5 more related resources inside...

ISBN: 978-1-4686-0469-6

Cover image used under license from Shutterstock.com

Landing gear provides an intriguing and compelling challenge, combining many fields of science and engineering. Designed to guide the interested reader through the fundamentals of aircraft wheels, brakes and brake control design systems, this book presents a specific element of landing gear design in an accessible way.