Jahrbuch der Wissenschaftlichen Gesellschaft für Luftfahrt e.V. (WGL). 1961: Mit den Vorträgen der WGL-Tagung in Freiburg im Breisgau Vom 10. bis 13. Oktober 1961 [Reprint 2021 ed.] 9783112527160, 9783112527153

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JAHRBUCH

1961 DER WISSENSCHAFTLICHEN GESELLSCHAFT FÜR LUFTFAHRT E.V. (WGL)

M I T DEN V O R T R Ä G E N DER W G L - T A G U N G IN F R E I B U R G IM B R E I S G A U VOM 10. BIS 1 3 . O K T O B E R 1 9 6 1

H E R A U S G E G E B E N VON H E R M A N N B L E N K SCHRIFTLEITUNG:

WERNER SCHULZ HELMUT N A B L

FRIEDR. VIEWEG & S O H N BRAUNSCHWEIG

Alle Rechte, auch die des auszugsweisen Nachdrucks, der fotomechanischen Wiedergabe und der Ubersetzung, vorbehalten. 1962 Wissenschaftliche Gesellschaft für Luftfahrt e. V. (WGL) Satz und Druck: A C O D R U C K GMBH, Braunschweig Bindearbeiten: Buchbinderei Schaare, Braunschweig Bezug des Jahrbuches für Mitglieder der WGL durch die Hauptgeschäftsstelle der WGL, Köln, Martinstraße 40—42

VORWORT

D a s Jahr 1961 hat für die Geschichte der deutschen Luft- und Raumfahrt eine besondere Bedeutung. Denn im Sommer 1961 erklärte die Bundesrepublik ihre grundsätzliche Bereitschaft, auf dem Gebiete der Raumfahrt im internationalen Rahmen mitzuarbeiten. Dieser Entschluß der Bundesregierung hat die WGL bestärkt, ihre Arbeit auf das Gebiet der Raumfahrt auszudehnen und bei ihrer Jahresversammlung 1961 Probleme der Raumfahrt in den Vordergrund zu stellen. Die Jahresversammlung der WGL wurde im Auditorium maximum der Albert-Ludwig-Universität in Freiburg mit zwei wissenschaftlichen Festvorträgen eröffnet. Den ersten von ihnen hielt Prof. Dr. phil. Julius Bartels, Ordinarius für Geophysik an der Universität Göttingen und Direktor des MaxPlanck-Instituts für Stratosphären-Physik; er berichtete über die wissenschaftlichen Ergebnisse, die die Erforschung des Weltraumes uns in den letzten Jahren gebracht hat. Als zweiter Redner sprach Prof. Dr.-Ing. Otto Lutz, Ordinarius an der Technischen Hochschule Braunschweig, Präsident der Deutschen Forschungsanstalt für Luft- und Raumfahrt und Direktor des Instituts für Strahltriebwerke dieser Anstalt; er gab in seinem Vortrag einen umfassenden Uberblick über die moderne Energietechnik, die für die Raumfahrt eine entscheidende Bedeutung hat. Die Probleme der Raumfahrt sind außerordentlich vielseitig und erstrecken sich auf die verschiedensten Wissensgebiete. Es erschien daher zweckmäßig, in der Jahrestagung aus der Fülle dieser Gebiete einige Problemkreise herauszustellen, wie z. B. die Themen „Antriebe und Treibstoffe", „Regelung und Steuerung" von Raumflugkörpern. Da diese Gebiete in der Bundesrepublik heute nur in geringem Umfange bearbeitet werden, wurde es mit großer Dankbarkeit begrüßt, daß zahlreiche Wissenschaftler aus dem Ausland sich sowohl für Übersichtsvorträge als auch für Vorträge zu speziellen Problemen zur Verfügung gestellt hatten. Neben diesen Vorträgen wurde eine größere Anzahl Referate gehalten, die sich mit Fragen der Strömungsmechanik, der Flugmechanik und Flugregelung, der Werkstoffe und der Bauelemente befaßten. Die Jahresversammlung fand ein lebhaftes Interesse, das sich durch die hohe Teilnehmerzahl ausdrückte. Insgesamt hatten sich mehr als 900 Wissenschaftler und Ingenieure bei der Tagung eingefunden, davon über 100 Herren aus dem Ausland, so aus den USA, Großbritannien, Frankreich, Holland, Schweden, Belgien, Österreich und der Schweiz. Neben den Vorträgen, die auf der Jahresversammlung gehalten wurden, enthält das Jahrbuch einen Bericht über die Arbeit der WGL; hierbei sei besonders auf die Sitzungen der Fachausschüsse, die Sprechabende, die Preisausschreiben und auf die Ludwig-Prandtl-Gedächtnis-Vorlesung hingewiesen, bei der alljährlich der Ludwig-Prandtl-Ring verliehen wird. In dem Berichtsjahr verstärkte die WGL weiterhin ihre Verbindungen zu ausländischen wissenschaftlichen Gesellschaften. Die Otto-Lilienthal-Gedä.dntms-'Vodesung, die gemeinsam mit der Association Française des Ingenieurs et Techniciens de l'Aéronautique et de l'Espace alljährlich abgehalten wird, fand am 24. November 1961 in Braunschweig statt. Ferner wurde anläßlich der Ludwig-PrandtlGedächtnis-Vorlesung in Würzburg am 7. April 1961 mit der Royal Aeronautical Society verabredet, in zweijährigem Abstand eine Reynolds-Prandtl-Vorlesung abzuhalten, die inzwischen erstmalig am 16. April 1962 in London veranstaltet wurde. Auch mit den Gesellschaften, die in der Bundesrepublik auf verwandten Gebieten arbeiten, wurde enge Fühlung gehalten. Dies gilt in besonderem Maße für die Deutsche Gesellschaft für Raketentechnik und Raumfahrt e. V., unter deren Mitwirkung die Jahrestagung der WGL stattfand, sowie für die Deutsche Raketen-Gesellschaft. G. B O C K Darmstadt, im August 1962

PREFACE

PREFACE In the history of German Aviation and Space Travel, the year 1961 is of special significance in so f a r as the Federal Republic declared its readiness in principle to take an active p a r t in the field of space travel at an international level. This decision has instigated the W G L to extend its activities into the fields of space travel and to give priority to problems of astronautics at the Annual Conference in 1961. The Annual Conference of the W G L began with two special science lectures read in the Main H a l l of the Albert-Ludwig University of Freiburg. The first of these lectures was held by Professor Dr. phil. Julius Bartels, H e a d of the Department for Geophysics at the University of Göttingen and Director of the Max-Planck-Institute for Physics of the Stratosphere; the lecturer reported on the scientific results obtained f r o m space research in recent years. T h e second lecture was presented by Professor Dr.-Ing. Otto Lutz, H e a d of Department at the Technical University of Braunschweig, President of the German Establishment for Aeronautics and Astronautics, and Director of the Institute for Jet Propulsion at this Establishment; the lecturer gave a comprehensive survey of modern power engineering, which plays a decisive p a r t in space travel. Problems of space travel being extremely manifold, and involving a great variety of fields of specialised knowledge, it was logical on the occasion of the Annual Conference to set apart from the multitude of different fields, some general problems, such topics as for instance "Propulsion and Combustion" and "Control and Guidance of Spacecraft". As relatively little work is being done in these fields within the Federal Republic today, the fact that many scientists from abroad had made themselves available to lecture not only on general problems but also on distinctly formulated problems was greatly appreciated. A p a r t from these lectures, a considerable number of informatory talks were given, dealing with questions arising in fluid mechanics, stability and control, material and elements of structure.

Considérant le développement de l'aéronautique et de l'astronautique en Allemagne, l'année 1961 prend une signification particulière. En fait, la République Fédérale déclara sa volonté en principe de participer aux recherches de l'espace dans un cadre international. Cette décision incita la W G L à étendre son activité sur le domaine de l'astronautique et à donner priorité, à l'occasion de son Congrès Annuel 1961, aux problèmes de l'espace. Le Congrès Annuel de la WGL, ayant eu lieu au grand amphithéâtre de l'université Ludwig Albert à Fribourg, f u t inauguré par deux conférences scientifiques. La première conférence, présentée p a r le Professeur Dr. phil. Julius Bartels, professeur titulaire de géophysique à l'Université de Gottingen et Directeur de l'Institut Max Planck des Physiques de la Stratosphère traita des résultats scientifiques fournies p a r l'exploration de l'espace au cours des dernières années. La deuxième conférence f u t lue par le Professeur Dr.-Ing. Otto Lutz, professeur titulaire de l'Université Technique de Braunschweig, Président de l'Etablissement de Recherches Aéronautiques et Spatiales et Directeur du Groupe de Recherches de la Propulsion par Jet de cet Etablissement. Cette conférence rendit compte du développement actuel de la technique de l'énergie, celle-ci jouant un rôle décisive dans l'astronautique. Les problèmes de l'astronautique étant extrêmement variés et touchant une multitude de branches spécialisées de la science, il parut indiqué de n'en faire ressortir que certains groupes de problèmes, comme p a r exemple : la propulsion et la combustion, le contrôle et la commande des véhicules d'espace. Relativement peu de travail est fourni dans ces domaines en Allemagne Fédérale et les conférences et communications traitant de problèmes spéciaux et présentées p a r de nombreux hommes de science de l'étranger furent vivement accueillies. Outre ces conférences il y avait un nombre de rapports ayant trait aux questions de la mécanique des fluides, de la stabilité et du contrôle ainsi que des matériaux et éléments de construction.

The Annual Conference aroused great interest, as was evident from the high figure of attendance. The Conference was attended by a total of 900 scientists and engineers, including 100 visitors f r o m foreign countries such as USA, Great Britain, France, the Netherlands, Sweden, Belgium, Austria and Switzerland.

Le grand nombre de participants prouve l'intérêt qu'incita le Congrès Annuel. Plus de 900 hommes de science et ingénieurs y participaient, dont plus de 100 aérotechniciens venant de l'étranger : des Etats-Unis, de la Grande-Bretagne, de la France, des Pays Bas, de la Suède, de la Belgique, de l'Autriche et de la Suisse.

Apart f r o m the lectures presented at the Annual Conference, the Yearbook also comprises a report on the activities of the W G L ; special attention is d r a w n to the meetings of subcommittees, the evening discussion groups, the prize competitions and the Ludwig Prandtl Memorial Lecture, on the annual occasion of which the Ludwig Prandtl Signet Ring is awarded. In the current year the W G L continued to strengthen its relations with scientific organisations of other countries. The Otto Lilienthal Memorial Lecture, an annual event in connection with the Association Française des Ingénieurs et Techniciens de l'Aéronautique et de l'Espace, was held on November 24th, 1961 at Braunschweig. In addition, it was agreed, on the occasion of the Ludwig Prandtl Memorial Lecture on April 7th, 1961 at Würzburg, to organise a Reynolds-Prandtl Lecture every other year which, meanwhile, took place in London on April 16th, 1962.

O u t r e les conférences du Congrès Annuel, cet annuaire contient un r a p p o r t sur l'activité de la WGL ; dans ce contexte, il f a u t spécialement mentionner les sessions des comités spécialisés, les groupes de discussion, les compétitions ainsi que la conférence à la mémoire de Ludwig Prandtl, à l'occasion de laquelle une bague à cachet est conférée annuellement. Dans l'année courante, la W G L continua à renforcer ses liens avec d'autres organisations scientifiques étrangères. Le 24 novembre 1961 eut lieu à Braunschweig la conférence Otto Lilienthal organisée annuellement en commun avec l'Association Française des Ingénieurs et Techniciens de l'Aéronautique et de l'Espace. A l'occasion de la conférence Ludwig Prandtl, le 7 avril 1961 à Wiirzburg, la W G L et la Royal Aeronautical Society convinrent d'organiser tous les deux ans une conférence ReynoldsPrandtl. La première de ces conférences eut lieu le 16 avril 1962 à Londres.

Close contact was also entertained with Organisations working in allied fields within the Federal Republic. This applies to a special degree to both the German Association for Rocket Engineering and Space Travel, an Organisation that played an active part in the Annual Conference of the WGL, and the Geman Rocket Society.

U n contact étroit fut également gardé avec des associations de la République Fédérale travaillant dans des domaines alliés. Ceci s'applique particulièrement à l'Association Allemande de la Technique des Fusées et de l'Espace, ayant activement participé au Congrès Annuel de la WGL, et à l'Association Allemande de la Technique des Fusées.

Darmstadt, August, 1962

Darmstadt, août 1962 G. B O C K

INHALT V o r t r ä g e der Freiburger Tagung Festvortrag O.Lutz,

R a u m f a h r t f o r s c h u n g u n d m o d e r n e Energietechnik

Gemeinsame B. H. Goethert

11

Sitzungen

and

A. E. Lennert,

D e v e l o p m e n t t r e n d s in rocket p r o p u l s i o n systems

30

0. H. Lange,

T r e n d s in t h e d e v e l o p m e n t of U . S. space vehicles

40

D. ]. Lyons,

Satellite launching possibilities

75

K. H. Schirrmacher,

L e n k u n g s - u n d S t e u e r u n g s p r o b l e m e v o n Satelliten u n d Raumfahrtvehikeln

85

W. ]. Jacobi a n d

F o r m u l a t i o n of g u i d a n c e a n d c o n t r o l e q u a t i o n s , their

C . S. Bridge,

mechanization and instrumentation

C. A. Traenkle,

P r o b l e m s of t h e mechanics of i n t e r p l a n e t a r y space t r a v e l

102

93 113

G.Ulbricht,

A n t e n n e n z u m Nachrichtenaustausch m i t R a u m f l u g k ö r p e r n

1. Estermann,

S i m u l a t i o n of high a l t i t u d e f l i g h t c o n d i t i o n s

122

R. Hermann,

H y p e r s o n i c f l o w p r o b l e m s d u r i n g r e - e n t r y i n t o the a t m o s p h e r e

131

E. H. Graul,

R a u m f a h r t m e d i z i n als biophysikalisches u n d anthropotechnisches Forschungsproblem'

145

J. Fabian,

Raketenwerkstoffe

155

Th. Peters,

P l a s m a p h y s i k in d e r H y p e r s c h a l l - A e r o d y n a m i k u n d in der Triebwerkstechnik

166

Strömungsmechanik W. Gretler,

N e u e r e M e t h o d e z u r Berechnung d e r ebenen U n t e r s c h a l l -

K. Kraemer,

K r a f t - u n d D r u c k v e r t e i l u n g s m e s s u n g e n a n einem Deltaflügel

strömung an dünnen Profilen

171

bei i n k o m p r e s s i b l e r S t r ö m u n g u n d Vergleich m i t d e r T h e o r i e (Übersicht)

179

H. Ludwieg,

Z u r E r k l ä r u n g der I n s t a b i l i t ä t der über angestellten D e l t a f l ü g e l n

H. Friedet,

Berechnung d e r A u f t r i e b s v e r t e i l u n g a n T r a g f l ü g e l n bei

a u f t r e t e n d e n f r e i e n W i r b e l k e r n e (Übersicht)

180

Überschallgeschwindigkeit

181

M. Enselme,

E i n neues A n a l o g i e v e r f a h r e n z u r U n t e r s u c h u n g der S t r ö m u n g s -

B. Lasebka,

D i e i n s t a t i o n ä r e n L u f t k r ä f t e an harmonisch schwingenden

v o r g ä n g e im Uberschallbereich

203

T r a g f l ü g e l n endlicher S p a n n w e i t e bei U n t e r - u n d Überschallgeschwindigkeit

207

F. Bartlmä,

I n s t a t i o n ä r e S t r ö m u n g s v o r g ä n g e bei Überschreiten d e r kritischen

Ting-Yi

D e r E i n f l u ß d e r D r u c k v e r t e i l u n g a u f die s t a r k e Z ä h i g k e i t s -

W ä r m e z u f u h r (Übersicht) Li u n d

J. F. Gross,

Wechselwirkung im H y p e r s c h a l l g e b i e t

220 a n einer P l a t t e

mit

Stoffaustausch F. Hindelang, P. Carrière

221

D e r E i n f l u ß der Eigenschaften realer Gase auf hypersonische Strömungen

228 234

et

M. Sirieix,

E f f e t s a é r o d y n a m i q u e s de l ' é c l a t e m e n t d ' u n jet de fusée

L. F. Crabtree,

T h e f l o w f i e l d associated w i t h a r o c k e t jet e f f l u x at v e r y

F. Thomas,

U n t e r s u c h u n g e n ü b e r die E r h ö h u n g des A u f t r i e b e s von T r a g -

high a l t i t u d e s

240

flügeln m i t t e l s Grenzschichtbeeinflussung durch Ausblasen (Übersicht)

243

R. Michel,

C o n d i t i o n s d e d é c o l l e m e n t de la couche limite à g r a n d e a l t i t u d e a u x vitesses h y p e r s o n i q u e s

244

E. Becker,

P s e u d o s t a t i o n ä r e , k o m p r e s s i b l e Grenzschichten (Übersicht)

254

W. van

Nes,

Bericht über M e ß m e t h o d e n und Meßergebnisse bei Flugmessungen in der Grenzschicht Die Bestimmung der hydrodynamischen Größen einer Kärmanschen Wirbelstraße aus Hitzdrahtmessungen bei kleinen Reynoldsschen Zahlen (Obersicht) Spoiler im Uberschall Experimentelle Untersuchungen an festen Spoilern und Strahlspoilern bei Machschen Zahlen von 0,6 bis 2,8 (Ubersicht)

E. Berger,

H.-G. Knoche, A. Heyser u n d F. Maurer,

F l u g m e c h a n i k, E. E,

Perret,

D. H.

Schmieder,

}. Tschauner, A. Becker, O. Weber, K. Ernst, G.

Brüning,

K. B.

Kindermann,

E. M. Vischel, G. W. Braun, M. G.

Jaenke,

C. L.

Emmerich,

C. R.

Himmler,

R. J. E. E.

Dinjar, Domhrowski, Rehhock,

M. R.

Nagel,

H.

Kleinwächter,

Th.

Heller,

A.

Kutzer,

F.

Wiegand,

F 1 u g r e g e 1 u n g, N a v i g a t i o n ,

255

260 262 268

Elektronik

Relativbewegung v o n K ö r p e r n in einem zentralen Gravitationsfeld Use of calculus of variations methods for t r a j e c t o r y optimization and a d v a n c e d guidance concepts Die Stabilität der Abtastsysteme Eine AM-FM-Klein-Telemetrieanlage f ü r Flugmessungen E r f a h r u n g e n mit dem R a d a r g e r ä t der D F L Beurteilungsmöglichkeiten von Regelkreisgliedern auf G r u n d von Flugmeßergebnissen Frequenzgangmessungen an einem instabilen Flugzeug mittels

269 272 277 281 285 291

der Systemtheorie f ü r regellose Vorgänge (Ubersicht) Über die E r p r o b u n g eines drehmomentfreien u n d getriebelosen

298

Hubschrauberantriebs Ü b e r Probleme der N a v i g a t i o n Calibration af inertial guidance systems in an operational environment A u t o m a t i c d a t a processing and computation techniques for the evaluation of inertial guidance systems A n w e n d u n g der Schwimmkreisel bei Lenkgeschossen und in der R a u m f a h r t Mesures sur le comportement statique et d y n a m i q u e des systèmes de guidance électro-hydraulique p o u r engins et considérations sur les valeurs optima possibles A d v a n c e d control techniques for solution of structural bending and elastic feedback problems of large, flexible boosters A u f b a u elektronischer Geräte f ü r die R a u m f a h r t t e c h n i k Elektromagnetische Strahlung als Verbindungsmittel zwischen Erdoberfläche und R a u m f a h r z e u g e n Some applications of i n f r a r e d techniques in space flight and space research A n w e n d u n g der Polarisation von R a d a r - und Lichtwellen zur Messung der D r e h u n g v o n Flugkörpern Mikrowellen-Antennen zur Positionsbestimmung von Satelliten f ü r die N a v i g a t i o n

299 302 308 314 319

323 328 331 335 339 344 347

Untersuchungen zum Verhalten des Menschen in Lenksystemen nach optischer Sicht Sichtbarmachung von Bordgeräten mit neuentwickelten Betastrahler-Dauerleuchtfarben

361

Entwicklung der Strahlruder bei der S N E C M A Gasstrahlablenkung

364 368

356

Antriebe G. Ernst, W. Eick, G. Kuers, G. Au, H. Prechter, W. 0.

Trommsdorff, Conrad,

Einige Eigenschaften der von TL-Triebwerken u n d ihrem Gasstrahl emittierten U l t r a r o t - E n e r g i e I n d u k t i v e Plasmaerzeugung in H o c h f r e q u e n z f e l d e r n Berücksichtigung neuerer Forschungsergebnisse f ü r die Auslegung von Unterschalldiffusoren (Übersicht) Beitrag zur Auslegung von Triebwerkseinläufen im mittleren und hohen Uberschall Wirtschaftlich optimale Zweikreis-Triebwerke

378 382 392 393 399

A, Leprince, Influence du taux de travail des enveloppes sur les performances R. Féraud, R. Aider, d'un engin à propergol solide W.Buschulte, Fragen der Lithergolantriebe L.Nadaud, La combustion des propergols solides métallisés

408 414 418

Festigkeit G. Czerwenka,

Krafteinleitung in zylindrische oder konisch veränderliche Schalen Über das elastische Stabilitätsverhalten von dünnwandigen Zylinder- und Kegelschalen Contribution à l'étude des récipients sous pression Messung der elastischen Deformationen an einem Kunststoffflügel Schwellfestigkeitsuntersuchungen an Nietverbindungen

435 441 444 454

H. H. Weigand, E. Fitzer und

Eigenschaften von Titanlegierungen für Flugkörper

461

O. Vohler,

Graphit als Werkstoff in der Raketentechnik

467

W. Schnell, O. Brodln, L. Schmieder, G. Hoffmann,

425

Werkstoffe

Allgemeines Ehrentafel Vorstand und Vorstandsrat der WGL (Stand vom 31. März 1962) Nachrufe auf Peter Supf, Erich Kleinemeyer, Hanns Klemm, Erich Syamken, Ernst Koschel, Günther Kempf, Ludwig Schiller, Fritz Roßmann, Erich Siebel, Ernst Wilhelm Pleines, Ludwig Weickmann Bericht über die Tätigkeit der WGL im Jahre 1961 Bericht über die Tätigkeit der WGL-Ausschüsse Fünfte Ludwig-Prandtl-Gedächtnis-Vorlesung Verleihung des Ludwig-Prandtl-Ringes 1961 Zweite Otto-Lilienthal-Vorlesung Preisträger der Preisausschreiben der WGL Wissenschaftliches Preisausschreiben 1959 Otto-Lilienthal-Preis 1961 Ludwig-Prandtl-Preis 1961 Projekt-Wettbewerb „Schul- und Übungsflugzeug" Preisausschreiben 1961 der WGL Studienrat-Schmitz-Preis Bericht über die Jahrestagung der WGL in Freiburg Protokoll über die ordentliche Mitgliederversammlung der WGL in Freiburg Namen- und Sachregister

477 479

481 490 495 500 501 502 504

505 506 510 513

O T T O LUTZ,

D K 620.92 621.455 629.19

BRAUNSCHWEIG*)

RAUMFAHRTFORSCHUNG

UND

MODERNE

1. E I N L E I T U N G Nur widerstrebend habe ich diesen Vortrag übernommen. Man sollte im allgemeinen über Dinge reden, an denen man selbst aktiv mitgewirkt hat; aber auf dem Gebiet der Raumfahrt sind wir schließlich alle noch Lernende. Andererseits beginnt die Raumfahrttechnik unserer Meinung nach eben jetzt, ihre eigenen technischen und wissenschaftlichen Grenzen abzustecken, ihre besonderen Verfahren zu entwickeln, Einfluß auszuüben auf andere technische Zwecke, kurz ein technisches Eigenleben zu führen, das sie weit hinaushebt über die bloße Hilfeleistung für Aufgaben, wie sie sich etwa bei der extraterrestrischen Forschung ergeben. Es ist deshalb berechtigt, die sich bildenden Säulen europäischer Zusammenarbeit auf diesem Gebiet — die European Space Research Organization (ESRO) und die European Launching Development Organization (ELDO) — strukturell auseinanderzuhalten und nur an der Spitze zusammenzuführen. Es ist aber hier nicht der Ort, darauf näher einzugehen. Erwarten Sie nun nicht, daß ich Ihnen einen Katalog der wissenschaftlichen und technischen Aufgaben der Raumfahrttechnik vorlege; dies wäre von meiner Seite her vermessen, für Sie langweilig. Mein ganzes Leben habe ich mich dem Triebwerksbau verbunden gefühlt, also der Energieerzeugung oder noch schärfer der Konversion von Wärme in mechanische Energie unter Verwendung thermodynamischer Prozesse. Auf dem erweiterten Gebiet der Energiekonversion — also auch auf dem elektrischen, physikalischen, chemischen und kerntechnischen Gebiet — beginnt mit der Raumfahrt geradezu ein neues Zeitalter heraufzuziehen. Die gegebenen Aufgabenstellungen verlangen den Einsatz aller nur erdenklichen Möglichkeiten; es ist bereits ein großer und vielfarbiger Strauß, der auf diesem Gebiet dargeboten werden kann. So möchte ich Ihnen ganz allgemein zeigen, welche Möglichkeiten der Energie-Darbietung und -Wandlung ins Auge gefaßt werden; die Ausblicke auf Anwendungen für andere Zwecke allgemeinster Art sind überraschend und eindringlich. Man hat zu unterscheiden zwischen den Antriebsenergiequellen für das Fahrzeug und den Hilfsenergiequellen für die Bordversorgung. Dazwischen stehen die elektrischen Antriebe, die zwar hohe Ausströmgeschwindigkeit, aber nur kleinen Schub ergeben. Man spricht deshalb auch von „high thrust" und „low thrust" propulsion, besonders pointiert sogar von „microthrust". Wo ist die Grenze? Um aus einem Gravitationsfeld zu entkommen, wendet man am günstigsten eine große, nur kurzzeitig wirkende Beschleunigung an. Die aufzubringende Arbeit muß um so größer sein, je geringer die Beschleunigung ist und je länger sie im Gravitationsfeld wirken muß. Für die vorliegenden Aufgaben (Erde, Mond, Venus, Mars) zeigt sich, daß die Grenze bei 0,3 bis 0,5 g liegt (naturgemäß nicht für einen *) T H Braunschweig und Deutsche Forschungsanstalt für Luftund Raumfahrt (DFL).

ENERGIETECHNIK

Start von der Erdoberfläche); unterhalb dieser Grenze muß zunehmend mehr Arbeit aufgebracht werden. Es ist zweckmäßig, die Triebwerke ebenfalls dieser Grenze entsprechend einzuteilen. Zwei Größen kennzeichnen im wesentlichen ein Antriebssystem: der spezifische Impuls in kp sec/kg, d.h. der Impuls je kg Treibstoff, der 9,81 mal kleiner als die wirksame Ausströmgeschwindigkeit ist, und der „propellant loading factor" X, das Verhältnis des Treibstoffgewichts zum Gesamtgewicht von Treibstoff und Triebwerk. Um Werte zu geben: der spezifische Impuls liegt bei heutigen Feststofftriebwerken am Boden bei 250 kp sec/kg, im Vakuum bei 310 kp sec/kg, bei Flüssigsauerstoff-KerosenTriebwerken bei 285 bzw. 335 kp sec/kg; bei den in Entwicklung begriffenen Flüssigsauerstoff-FlüssigwasserstoffTriebwerken beträgt er 350 bzw. 450 kp sec/kg. „Lowthrust"-Triebwerke liegen um Größenordnungen höher. Bei Kernantrieben erwartet man in nächster Zukunft Werte zwischen 1000 und 2000 kp sec/kg. Der Konstruktionsfaktor l bewegt sich zwischen 0,90 und 0,95, bei „lowthrust"-Triebwerken liegt er sicherlich weniger hoch. Schließlich soll noch die Tatsache erwähnt werden, daß die im Triebwerk zu installierende Leistung dem Produkt aus Schub und spezifischem Impuls verhältig ist, was sich besonders gravierend auf die elektrischen Antriebe auswirkt; die Ausstoßleistung ist m* w 2 /2 (rrf in der Zeiteinheit ausgestoßene Masse, w Ausströmgeschwindigkeit), wobei man rrv' w zum Schub zusammenfassen kann und das verbleibende w/2 dem spezifischen Impuls verhältig ist. Um die Wichtigkeit des spezifischen Impulses zu unterstreichen, sei angeführt, daß man bei den mit Kernenergie arbeitenden Triebwerken mit spezifischen Impulsen von 1000 kp sec/kg und darüber für die gleiche Aufgabe mit einem Fahrzeug auskommt, das nur 1/10 bis 1/20 so groß wie ein heutiges ist; ja, die Firma Douglas hat bereits ausgerechnet, daß die „direct operating costs" bei weniger als 1000 Dollar je Person für eine Mond-Umrundung liegen werden, und daß ein Drei-Personen-Fahrzeug die gleiche Größenordnung annimmt wie die heutige Douglas DC-8. Nach diesen vorausgeschickten Bemerkungen sollen nun nacheinander die einzelnen Antriebsformen behandelt werden, wobei die konventionellen Antriebe nur kurz gestreift werden. Meinem Mitarbeiter Dipl.-Phys. G. Au bin ich zu Dank verpflichtet; er hat mich bei der Abfassung wesentlich unterstützt. 2. A N T R I E B S E N E R G I E Q U E L L E N 2.1. Chemische Energie Im Mittelpunkt des Interesses stehen die HochenergieTreibstoffe, wie die Knallgasreaktion H 2 + 0 2 , die Reaktion H 2 + O s , sowie die Verbrennung von Fluor und Hydrazin oder Fluorverbindungen. Bei letzteren kann man spezifische

12

O. Lutz, Raumfahrtforschung und Energietechnik

1000

Jahrbuch 1961 der WGL Industrie z u r Erweiterung bzw. z u m N e u a u f b a u von A n lagen z u r P r o d u k t i o n v o n Treibstoff w u r d e n im J a h r e 1960 etwa auf 30 bis 40 Millionen D o l l a r geschätzt. A u d i f ü r den Einsatz v o n Kernenergieantrieben w i r d flüssiger Wasserstoff in erheblicher Menge erforderlich sein. Fortschritte sind auch auf dem Gebiet der Kälteisolierung der verflüssigten Gase zu verzeichnen, die bei T e m p e r a t u r e n v o n —183 b z w . — 257 ° C gehalten werden müssen. Die Linde C o m p a n y entwickelte eine aus einer Reihe von Klebeschichten bestehende Superisolierung, durch die der U m g a n g mit Flüssigwasserstoff ebenso ungefährlich gew o r d e n ist wie mit Flüssigsauerstoff. Festtreibstoffe k ö n n e n wegen der höheren Molekulargewichte der Treibgase grundsätzlich keine so hohen I m pulswerte bereitstellen; d e n k t m a n etwa an LithiumH y d r i d u n d L i t h i u m - F l u o r a m i d , so sind v o n einer solchen Zukunftsentwicklung doch nicht mehr als e t w a 350 k p sec/kg zu e r w a r t e n . Kernenergie

2.2.

Durch die R a u m f a h r t f o r s c h u n g erwächst der A t o m k e r n forschung u n d Kerntechnik in verschiedener Hinsicht ein mächtiger Bundesgenosse, bezüglich der Verteilung der Mittel ein starker K o n k u r r e n t . Die Erfordernisse d e r Höchstleistungsantriebe f ü h r t e n bereits v o r einigen J a h r e n in den U S A dazu, K e r n r e a k t o r e n f ü r R a k e t e n t r i e b w e r k e in Angriff zu nehmen, da hiermit spezifische Impulse von über 1000 k p sec/kg realisierbar sind. In der jüngsten Zeit w u r d e in den U S A h i e r f ü r eine beträchtliche Summe in das Jahresbudget eingesetzt, um möglichst bald K e r n r e a k t o r e n f ü r R a k e t e n t r i e b w e r k e z u r V e r f ü g u n g zu haben. -3000

Bild 1. ¿*, T-Tafel für a) H 2 G a s + O S G a s , b) H 2 t l + 0 3 G a s , c = ) H 2 G a s + Ogfj, d) Hgfj + 0 2 f i ; Reaktionsenthalpien i* nach den Reaktionen und Entspannung ins Vakuum (T — 0 °K) in Abhängigkeit vom Reaktionsverhältnis f . Impulse v o n r u n d 400 k p sec/kg erwarten, w ä h r e n d die Knallgasreaktion theoretisch über 500 k p sec/kg ergibt. In Bild 1 ist die Reaktionsenthalpie i* in Abhängigkeit v o m ReaktionsVerhältnis r = « o / ( i " h + «o) f ü r eine E n t s p a n n u n g ins V a k u u m u n d einen Druck v o n 100 k p / c m 2 aufgetragen. H i e r b e i bedeuten wo u n d « h die A t o m d r ü c k e f ü r Sauerstoff b z w . Wasserstoff. Die K u r v e f ü r 0 ° K gilt f ü r alle Drücke; die Linien f ü r die verschiedenen Verbrennungsgemische entsprechen verschiedenen Reaktionsverhältnissen nach der R e a k t i o n . M a n erkennt, d a ß die B r e n n k a m m e r t e m p e r a t u r e n selbst f ü r O z o n noch unter 4000 ° K bleiben, die theoretische Verbesserung bei O z o n gegenüber Sauerstoff w ü r d e einen Schritt von 520 k p sec/kg nach 580 k p sec/kg bringen. Bei Sauerstoff glaubt m a n auf e t w a 450 k p sec/kg zu k o m m e n . Durch die Erfordernisse beim S a t u r n p r o j e k t erhält die Gasverflüssigung neue Impulse. D e r S a u e r s t o f f t a n k einer T h o r r a k e t e f a ß t beispielsweise zwischen 38 000 u n d 50 000 1, w ä h r e n d bei der S a t u r n r a k e t e f ü r die S t a r t s t u f e allein etwa 240 t Sauerstoff benötigt werden. Die V e r w e n dung v o n flüssigem Wasserstoff f ü r Höchstleistungstriebwerke, z. B. f ü r die C e n t a u r , die O b e r s t u f e der Saturn, macht die Verflüssigung dieses Gases im großtechnischen M a ß s t a b bereits lohnenswert. Die Investitionen der US-

Die außergewöhnlichen Fortschritte, die sich f ü r die Erf ü l l u n g der gleichen A u f g a b e ergeben (chemisch: dreistufig, 10- bis 20facher G e w i c h t s a u f w a n d ; kernenergetisch: einstufig), sind bereits in der Einleitung geschildert w o r d e n . Es steht heute fest, d a ß durch das v o n der N A S A v o r gesehene P r o g r a m m f ü r die b e m a n n t e R a u m f a h r t die Forschungen u n d Entwicklungen über K e r n r e a k t o r e n f ü r Triebw e r k e u n d elektrische Energieerzeugungsanlagen stark v o r w ä r t s getrieben werden. Seit September 1960 w i r d das

1260 kg L210kg

=.

V

1 260 kg

^

IV

I 960 kg

[—IV

2 530 kg

—III

10 330 kg

Kernreaktorleistung I1725MWI

7100kg -III

17 260 kg

Kernreaktor— leistung !3iS0MWI

10670 kg

-II

\ /

-II

7!l.i0kg

122310 kg 226030 kg

Schub=2l7x10°kp Chemischer Antrieb

§

-II

2UU0kg

^I

USLOkg 60 8 70 kg

ip*1trkp

\ /

V

26 170 kg

Kernreaktor— leistung ItOOOOMWl

/

33 S70 kg 72 000 kg

0,7x10° kp

Chemischer und Kernenergieantrieb

Bild 2. Saturn-Trägerraketen für bemannte Mondlandung mit rein chemischen Stufen (links) und chemischen und Kernenergiestufen (rechts). Für jedes System ist die zur Erde zurückkehrende Fahrzeugmasse etwa 3600 kg. Bei der Anwendung von zwei Kernenergietriebwerken wird der erforderliche Schub um etwa 25°/o kleiner. Nach H. B. Finger [2],

J a h r b u c h 1961 der W G L

O . L u t z , R a u m f a h r t f o r s c h u n g u n d Energietechnik

Programm f ü r das Projekt Rover von dem Direktor des A E C - N A S A Nuclear Propulsion Office ( N P O ) , Harold B. Finger, geleitet, der dieses Programm auch zu organisieren begann. Die hierfür vorgesehenen Prüfstände werden in Jackess Fiats in N e v a d a errichtet. Das N A S A - P r o g r a m m sieht f ü r die nächsten zehn Jahre folgende Großprojekte vor: 1. Bemannte Mondumrundung gegen 1970; 2. Bemannte Mondlandung nach 1970; 3. Bemannte Planetenlandung gegen 1980. Bild 2 zeigt den Vergleich einer Saturn-Trägerrakete mit zwei chemischen und einer Kernenergiestufe mit der reinchemischen Version f ü r die erste der genannten Aufgaben, die Mondumrundung. Bezieht man auf gleiche Nutzlasten, so erkennt man die große Überlegenheit des Kernenergieantriebes. Für die bemannte Mondlandung hat die N A S A übrigens folgende Schwerpunkte aufgestellt: a) Zusammenbau oder Treibstoffversorgung eines Antriebssystems auf einer Erdsatellitenbahn; b) Direkter Flug von der Erde zum Mond mit einer reinchemischen Trägerrakete; c) Direkter Flug mit einer Kombination aus chemischen und Kernenergiestufen. Ein Antriebssystem, das nur aus Kernenergiestufen bestehen soll, wird f ü r die bemannte Mondlandung mit Rückkehr noch einige Jahre auf sich warten lassen.

Die Gesamtmasse beträgt dabei 424 000 kg. Wieder ergibt sich eine Gewichtsverringerung auf ein Zehntel. Die experimentellen Untersuchungen beim Kiwi-A-Programm in N e v a d a zeigten in den letzten zwei Jahren große Fortschritte, können aber keineswegs als so weit fortgeschritten angesehen werden, daß heute schon ganz konkrete Vorstellungen über einen R e a k t o r t y p bestehen, der f ü r ein Kernenergietriebwerk eingesetzt werden könnte. Von den zahlreichen ungelösten Problemen stehen die folgenden beiden im Vordergrund:

_

1. Reaktorbetrieb im Hochtemperaturbereich bei hoher Leistungsdichte; 2. Zündung und Wiederzündung von Reaktoren. In N e v a d a sind z. Z. vorbereitende Arbeiten über Untersuchungen der Kiwi-B-Reaktoren im Gange, die schon als Flugkonfigurationen ausgebildet werden. Bisher wurden auf diesem Gelände wertvolle Untersuchungen an den Reaktoren Kiwi-A (Juli 1959), Kiwi-A' (Juli 1960) und Kiwi-A3 (Oktober 1960) unternommen. Die Kiwi-AReaktoren (vgl. Bild 4) enthalten den Kernbrennstoff Uran in Graphitplatten eingebettet. Diese sind in einer Ringzone

Nozzle outlet

M3in-coolant inlet manifold

30000

kg O) N 20 000

>--

Nozzle inlet

Core-support-liner inlet

Control-rod

=3

7 0000

13

3Optima er ehernicher Saturni ntrieb 3 i 103MW Reaktorleistung 25

50 80 Mp 100 Schub der Hemenergiestufe

Bild 3. Mögliche N u t z l a s t einer T r ä g e r r a k e t e zur Erreichung des M o n d g r a v i t a t i o n s f e l d e s in A b h ä n g i g k e i t v o n der R e a k t o r l e i s t u n g u n d v o m Schub der K e r n e n e r g i e s t u f e . N a c h H. B. Finger [ 2 ] .

D2O outlet In Bild 3 ist die mögliche Nutzlast einer Trägerrakete zur Erreichung des Mondgravitationsfeldes in Abhängigkeit von der Reaktorleistung und vom Schub des Kernenergietriebwerkes aufgezeichnet. Zum Vergleich ist die optimale Nutzlast f ü r die Saturn-Trägerrakete angegeben, die mit chemischen Triebwerken realisierbar ist. Der Kernenergieantrieb läßt also eine doppelt so schwere Nutzlast zu wie die Saturn-Version mit chemischen Triebwerken. Bei einem Dreistufenbahnstart f ü r eine bemannte Fahrt zum Mars und zurück würde mit einem chemischen Antrieb eine Gesamtmasse von 4 125 000 kg erforderlich sein. Dieselbe Aufgabe läßt sich mit den folgenden drei Kernenergiestufen erfüllen: Stufe

Reaktorleistung

1 2 3

10 900 M W 6 160 M W 1 890 M W

Bild 4.

K i w i - A - R e a k t o r auf d e m T e s t w a g e n m i t obenliegender D ü s e . N a c h R. E. Schreiber [3],

um ein Zentralgebiet von etwa 38 cm Durchmesser angeordnet, das zirkulierendes schweres Wasser D a O und die Kontrollstäbe enthält. Außerhalb der Brennstoffzone befinden sich ein Flußtrennzylinder, ein äußerer Reflektor, die beide aus Graphit bestehen, und ein wassergekühlter doppelwandiger Aluminiumdruckkessel, der von der Firma Rocketdyne hergestellt wurde. Die Graphitscheibe, die Kühlkanäle enthält, dient als Neutronenreflektor an der Einlaßseite des Reaktorkerns. Das Arbeitsmedium Wasserstoff strömt unter Druck durch die im Reflektor befindlichen Löcher in den Reaktorkern zwischen die einzelnen Brennstoffplatten. Hierbei wird der Wasserstoff durch die thermische Reaktorleistung stark erhitzt, in Atome dissoziiert und durch eine Laval-Düse entspannt. Die experi-

14

O. Lutz, Raumfahrtforschung und Energietechnik

Jahrbuch 1961 der W G L

mentelle Untersuchungsphase von Kiwi-A begann am 1. Juli 1959 zunächst mit relativ kleiner Reaktorleistung. Dann wurde die Reaktorleistung vergrößert und mehrere Minuten lang die volle Leistung gefahren. Der erste Flugversuch mit einem Kernenergieantrieb soll in der letzten Stufe einer Saturnrakete nach 1965 durchgeführt werden. Dazu sind folgende Hauptarbeiten erforderlich: Reaktorentwicklung für Flugkonfiguration, Reaktoreinbau in das Triebwerk, Triebwerksversuche, Entwicklung der oberen Stufe, Einbau des Triebwerkes in die Stufe, Verbindung der Kernenergiestufe mit dem Booster. Am 2. Februar 1961 wurde die Entwicklung eines Kernenergietriebwerkes an die Industrie unter der Bezeichnung Nerva vergeben. Für Flugversuche sollen die Vorschläge der Firmen Lockheed Aircraft Co., Martin Co., Douglas Aviation und Convair Astronautics dem Nuclear Propulsion Office (NPO) eingereicht werden. Für die ersten Versuche ist ein Reaktor von 1000 MW Leistung mit festem Reaktorkern vorgesehen. Reaktoren mit gasförmigen Reaktorkernen werden in Los Alamos, beim Lewis Research Center und bei verschiedenen Industrien untersucht. Zweifellos werden diese zu leichten Höchsttemperatur-Reaktoren führenden Entwicklungen für die allgemeine Energietechnik von großer Bedeutung werden. Wir in der Bundesrepublik sollten uns dabei nicht ausschließen.

Bild 5 zeigt die Nutzlast, die von einer 500-km-Bahn auf eine 24-Stunden-Bahn gebracht werden kann, wobei eine elektrische Leistung von 60 kW vom Kernreaktor SNAP-8 bereitgestellt wird. Diese Nutzlast ist bei einem elektrischen Antriebssystem bedeutend größer als beim chemischen Antrieb (H 2 + 0 2 ) , jedoch ist der Zeitaufwand erheblich größer. Von einer 24-Stunden-Bahn kann mit drei gleichmäßig verteilten Satelliten, die jeweils einen Kernreaktor SNAP-8 enthalten, jeder Punkt der Erde eingesehen werden, und es läßt sich gegebenenfalls ein weltweites Fernsehprogramm errichten (Bell).

3. ELEKTRISCHE ANTRIEBE Diese Antriebssysteme, über die in den USA bereits heute industrielle Entwicklungen im Gange sind, gehören nach den bisher bekannten Versionen zu den „low-thrust"- bzw. „microthrust"-Antrieben, sie haben daher einen beschränkten Anwendungsbereich in Räumen, wo relativ geringe Gravitationskräfte wirken. Sie lassen sich also nicht für den Start von der Erdoberfläche aus einsetzen, sondern müssen erst durch chemische oder Kernenergie-Trägerstufen aus dem Gravitationsfeld der Erde gebracht werden. Es treten immer wieder Zweifel auf, ob die elektrischen Antriebe in der Raumfahrt überhaupt eine Bedeutung erlangen können. Sind die realisierbaren Schubbeschleunigungen auch relativ klein (etwa 10—4 bis 10 —3 g), so lassen sich damit andererseits hohe Austrittsgeschwindigkeiten erzielen, die den Treibstoffverbrauch gegenüber den chemischen Antrieben beträchtlich reduzieren. Die Forschungen und die Entwicklungen müssen auf elektrische Antriebssysteme und Energiequellen mit möglichst geringem Leistungsgewicht zielen. i-500 's

®4ffflff 13500

+

s 'Capacitor NCurrent path ¡breakdown Jahr später sollen für diese Flugversuche elektrochemische —nrpWjTRTnir] banks region-plasma} Batterien von 1 k W Leistung verwendet werden. Der erste Flugversuch mit einem Kernreaktor als Energiequelle in der < < i j Größenordnung von 30 kW elektrischer Leistung ist für das Frühjahr 1965 vorgesehen. Hierfür wird der Kernreaktor Bild 8. Plasmaantrieb (schematisch). SNAP-8 eingesetzt, der der erste Reaktor im Raum sein Nach Douglas Report SM-37915. und mit einer Atlas-Centaur, zusammen mit den genannten Ionen- und Plasmaantrieben, auf eine Erdumlaufbahn geBeim elektrothermischen Plasmaantrieb wird ein Arbeitsbracht werden soll. Die Forderungen der N A S A sind aber gas im elektrischen Feld aufgeheizt, dissoziiert und ionisiert auf elektrische Antriebe ausgerichtet, für die elektrische und dabei ein Plasmastrahl erzeugt, der durch eine kurze Leistungen im Megawatt-Bereich vorgesehen werden müssen Düse expandiert wird (Bild 8). und die Austrittsgeschwindigkeiten zwischen 20 000 und Beim Ionenantrieb (Bild 9) wird durch eine der zahl60 000 m/sec haben sollen. reichen bekannten Methoden ein vorgeheizter Treibstoff,

z. B. eines der Alkalimetalle, ionisiert und das entstehende Ionenbündel in einem Linearbeschleunigersystem auf eine gewisse Geschwindigkeit gebracht. Hierbei entsteht der Schub durch die Wechselwirkung des Ionenstrahles mit einem elektrostatischen Beschleunigungsfeld. Beim Ionenbeschleuniger ist die Ausbildung positiver Raumladungen einmal im B'eschleunigersystem und zum anderen in der Strahlaustrittsebene ein Effekt, der heute bei den Forschungen noch eine Vielzahl von Schwierigkeiten bereitet. Es läßt sich keinesfalls ein Ionenantriebssystem realisieren, bei dem ein elektrisch geladenes, austretendes Ionenbündel zur Schuberzeugung dient. Es ist vielmehr erforderlich, den austretenden Ionenstrahl in der Höhe der Austrittsebene durch Elektronenbeschuß zu neutralisieren. Dieses Problem bildet heute in den Forschungslaboratorien der USA einen besonderen Schwerpunkt.

Die Forderungen, die an diese Kernreaktoren zu stellen sind, lauten: 1. Hohe elektrische Leistung, bei einigen 100 kW; 2. Einsatz bei Fahrten, die von der Sonne weggerichtet in interplanetare Räume führen, wo die Strahlungsleistung der Sonne viel zu gering wird, um Sonnenzellen oder parabolische Hohlspiegel noch mit großem Wirkungsgrad einsetzen zu können; 3. Lange Lebensdauer. Den Angaben der N A S A ist klar zu entnehmen, daß der Bedarf an elektrischen Leistungen für die Raumfahrt stark ansteigen wird und insbesondere nach 1965 für den Einsatz elektrischer Antriebssysteme eine erhebliche Steigerung erfahren muß. Hierfür müssen vor allen Dingen in der Kernreaktortechnik bei der Realisierung hoher spezifischer

O. Lutz, Raumfahrtforschung und Energietechnik

Jahrbuch 1961 d e r W G L

Leistungen ( > 0 , 1 kW/kg) bedeutende Fortschritte erzielt werden. Besonders deutlich traten die Fortschritte der Amerikaner auf der ersten Fachtagung der American Rocket Society über Ionenantriebe in Erscheinung, die am 3. und 4. September 1960 in Monterey in Kalifornien stattfand. Als interessantes Ergebnis stellte sich die Entwicklungstendenz des Cäsium-Ionenmotors heraus, der mit zunehmender Größe einen ansteigenden Antriebswirkungsgrad aufweist. Die beiden Erzeugungsmechanismen für Ionenstrahlen, die Gasentladungs-Ionenquellen und die Kontaktpotential-Ionenquellen, waren auf dieser Tagung etwa in gleicher Zahl vertreten.Ein besonderes Schwergewicht bildete dielonenstrahlneutralisation, die man heute allgemein als das schwierigste Problem bei der technischen Realisierung des Ionenantriebes erkannt hat. Es wird angenommen, daß bei der Neutralisation unter Weltraumbedingungen wesentlich andere Faktoren auftreten werden als bei den Laborversuchen in Hochvakuumbehältern von etwa 10—6 bis 10 ' s Torr.

zur Entwicklung von Raumradiatoren und Kondensatoren. Diese Anlage, die eine Raumklimakammer von etwa 2,5 m Durchmesser und etwa 5,5 m Länge einschließt, soll bis Oktober 1962 fertiggestellt sein. Auch diese Entwicklungen werden von fundamentaler Bedeutung für die allgemeine Energietechnik sein.

16

Von ganz besonderer Bedeutung ist die Erstellung einer Raumklimaanlage beim Lewis Research Center der NASA in Cleveland, deren Hochvakuum mit 20 öldiffusionspumpen auf einen Druck bis zu 10—6 Torr gebracht werden kann. Der Behälter besitzt zylinderförmige Gestalt und hat bei einem Durchmesser von 7,5 m eine Länge von etwa 24 m. Diese Größe ist einmal zur Ausschaltung der Ionenstrahlwechselwirkungen mit der Umgebung und auf der anderen Seite zur Erzielung exakter Messungen über die ganze Strahllänge erforderlich. Schließlich besteht auch noch ein Zusammenhang zwischen der zur Ionenstrahlerzeugung benutzten elektrischen Leistung, die von der Kernenergiequelle SNAP-8 mit etwa 70 kW gegeben ist, und den geometrischen Abmessungen des Vakuumbehälters. Es werden Schubwerte zwischen 1 und 2 kp erwartet. Über die SNAP-Kernreaktoren wird im nächsten Kapitel noch zu berichten sein. Offiziellen Stellen der NASA wurde Ende Juni berichtet, daß in der UdSSR beträchtliche Fortschritte bei elektrischen Antrieben gemacht wurden, für die Kernenergiequellen vorgesehen sind. Bereits im Jahre 1964 soll ein elektrisches Triebwerk mit einer Leistung von einigen 100 kW, die von einem Kernreaktor bereitgestellt wird, einsatzbereit sein. Demgegenüber steht das Programm der NASA, das erst für das Jahr 1966 ein elektrisches Antriebssystem von nur 30 bis 60 kW elektrischer Eingangsleistung (SNAP-8) vorsieht. Es wird daher bei der NASA die Notwendigkeit einer starken Intensivierung der Forschungs- und Entwicklungsarbeiten für äußerst vordringlich und wichtig gehalten. Diese Perspektiven veranlaßten die NASA, die Entwicklungen von Kernreaktoren mit 300 kW bis 1 MW Leistung stark zu beschleunigen, die als Oberstufen der Trägerraketen Saturn oder Nova verwendet werden sollen. Es wurden daher auch mehr Aufträge von der NASA an die amerikanische Raumfahrtindustrie erteilt, die insbesondere kernenergetisch-elektrische Hochtemperatur-Antriebssysteme betreffen. Für die Erteilung der Entwicklungsaufgaben ist das Lewis Research Center maßgebend, das folgende Firmen beauftragte: D e v e l o p m e n t D i v i s i o n of U n i t e d N u c l e a r C o r p . (früher NDA) mit der Entwicklung eines Wärmeübergangskreises von 1 MW f ü r alle die Aufheizung von Kalium und Natrium betreffenden Untersuchungen; A M F A t o m i c s mit dem Bau, der Installation und Untersuchung eines •Natriumverdampfers und Turbinenkreislaufes; P r a t t & Whitney's A i r c r a f t and Nuc l e a r E n g i n e L a b o r a t o r y ( C A N E L ) mit der Erstellung eines Prüfstandes beim Lewis Research Center

4. HILFSENERGIEQUELLEN Hierzu gehören alle Energiequellen, die elektrische Leistungen für die Stromversorgung der elektrischen Anlagen und Meßgeräte in Satelliten, Raumfluggeräten und Raumfahrzeugen bereitstellen sollen; sachlich gehören auch die Energiequellen für die elektrischen Antriebssysteme dazu. Dieses Feld ist besonders groß, reizvoll und aussichtsreich. Die Vielzahl der Möglichkeiten und die verschiedenen Beträge der erzeugten elektrischen Leistungen geben den einzelnen Energiequellen in der Raumfahrt ganz spezifische Aufgaben. Es werden sich in der nahen Zukunft noch viele neue Möglichkeiten eröffnen, die zum Teil schon heute in den ersten Versuchsstadien stehen. Die konventionelle Art der elektrischen Energieerzeugung wird durch völlig neue Wege verdrängt, bei denen der Prozeß auf direkte Weise erfolgt; Zwischenstufen werden entbehrlich, mechanische Prozesse nicht mehr erforderlich sein, so daß keine beweglichen Teile vorkommen. Schon Auswirkungen beschränkter Art auf unsere heutige Fahrzeug- und Energietechnik würden einer Revolution gleichkommen. Die Energiequellen kann man in folgende Gruppen untergliedern: 1. Chemische Reaktionen, 2. Atomkerne, a) Sonne, b) Radionuklide (Zerfallsenergie radioaktiver Kerne), c) Kernspaltung, d) Kernsynthese (Kernfusion). 1Ql kW 10 3

C3, I

y>

102

Qj - J

1 ff'

10"

10 '

10~2

IMin 5Min

IStd

Hag

W.

IMon. Zeit

1Jahr

10Jahre

Bild 10. Die verschiedenen Möglichkeiten der elektrischen Energieerzeugung mit charakteristischen Leistungs- und Zeitintervallen. 1 = elektrochemische Batterien, 2 = chemische Verbrennungsprozesse, 3 = Wasserstoff-Sauerstoff-Verbrennungsmaschine, 4 = Brennstoffelemente, 5 = Sonnenzellen und Radioisotope, 6 = Kernspaltung und Sonnenspiegel, 7 = Kernspaltreaktor.

Jahrbuch 1961 der W G L

102

Eine übersichtliche Einteilung ist in Bild 10 enthalten. Die bisher verwendeten Energiequellen in der R a u m fahrt waren auf elektrochemische Batterien, Sonnenzellen und auf einen einmaligen versuchsweisen Einsatz eines Radionuklide-Generators beschränkt. Mehr versprechende Systeme für die Zukunft sind Kombinationen eines Sonnenkollektors oder Kernreaktors mit einem Turbogenerator oder thermoelektrischen bzw. Thermionen-Wandlern. Für Betriebszeiten zwischen 4 Stunden und 1 J a h r liegen die Anforderungen zwischen 1 und 200 k W elektrischer Leistung. An der unteren Grenze dieser Arbeitsaufgabe liegt unser heutiger Personenkraftwagen. Die chemischen Energiequellen lassen sich in zwei H a u p t gruppen einteilen: a ) Elektrochemische Methoden: Batterien oder Brennstoffelemente; b) Verbrennungsprozesse: Ausnützung der thermischen Energie chemischer Reaktionen. In Erdsatelliten wurden bisher nur Ni-Cd-Batterien eingesetzt; f ü r das Projekt Mercury waren AgO-Zn-Batterien mit einer elektrischen Leistung von 70 W und einer Masse von 70 kg vorgesehen. Bei mehrmaliger Aufladung kann man bei diesen Batterien in Verbindung mit Sonnenzellen 100 bis 140 Wh/kg realisieren. Brennstoffelemente sind in der letzten Zeit ganz besonders interessant geworden, da hierbei die elektrische Leistung ohne Umwege unmittelbar freigemacht werden kann. Beim Brennstoffelement mit Sonnenenergie werden die Photonen von der Sonne dazu verwendet, die Reaktionsprodukte des Brennstoffelementes zu dissoziieren. Die Reaktanten laufen dann nach der Trennung im Brennstoffelement um und erzeugen direkt elektrische Energie. Bei der chemischen Energiequelle mit offenem Kreislauf w i r d Wasserstoff mit Sauerstoff verbrannt. Die Verbrennungsprodukte werden zum Antrieb einer Turbine verwendet, die mit einem Elektrogenerator verbunden ist. Bei der chemischen Energiequelle mit geschlossenem Kreislauf w i r d die Verbrennungsenergie von Wasserstoff und Sauerstoff zur Verdampfung von flüssigem Schwefel benutzt. Das in Umlauf gesetzte Arbeitsmedium w i r d dann zum Antrieb einer Getriebeturbine verwendet. Bei der Sonnenenergiequelle mit geschlossenem Kreislauf w i r d die Strahlungsenergie der Sonnenstrahlung durch einen parabolischen Hohlspiegel gesammelt und in einen Boiler fokussiert, in dem eine Arbeitsflüssigkeit aufgeheizt werden soll. Reaktoren, die den radioaktiven Zerfall einiger R a d i o nuklide verwenden, liefern für manche Anwendungen ausreichend hohe thermische Leistungen, die sich durch thermoelektrische bzw. Thermionen-Wandler in elektrische Leistungen umsetzen lassen. Dabei spielt für die heutigen Erfordernisse der R a u m f a h r t der turboelektrische Leistungsw a n d l e r noch eine bedeutende Rolle, w i e w i r später bei einigen in der Entwicklung befindlichen Kernreaktoren sehen werden. Es sind aber bereits eine Reihe von Methoden in der Untersuchung und teilweise auch in der Entwicklung, die die beweglichen Teile eines turboelektrischen Wandlersystems durch andere Möglichkeiten ersetzen. W a s bereits erreicht ist, zeigen die Bilder 11 und 12; Bild 13 gibt besonders deutlich die steigende spezifische Leistung mit wachsender Größe wieder. 4.1. B r e n n s t o f f e l e m e n t Die direkte Erzeugung von Elektrizität aus einem Brennstoff und einem O x y d a t o r ohne U m w a n d l u n g der diemischen Energie in W ä r m e ist das Ziel der Wissenschaftler seit über hundert Jahren. 2

17

O. Lutz, Raumfahrtforschung und Energietechnik

¡n

Verbindung Silizium p-Typ ^

\

'

Q © Silizium n-Typ \ P-n-

~?5i-mm - J [lektronen+luR

Bild 17.

ZBelastungsj kreis

Silizium-Sonnenzelle. Nach E. Justi

[70].

Kontaktfläche ruft eine kontinuierliche Verschiebung der Elektronen und Löcher innerhalb des vollständigen Gitters hervor. Durch diese eingeleiteten Verlagerungen der Ladungsträger tritt ein elektrischer Stromfluß auf. Der Wirkungsgrad einer solchen Zelle liegt z. Z. bei etwa 1 1 % . Gegenwärtige Untersuchungen zielen auf höhere Wirkungsgrade hin, die einmal durch Variation der chemischen Zusammensetzung und zum anderen durch die Technik der sogenannten Doppelschichten erreicht werden. Bei der zuletzt genannten Methode wird der Lichtdurchgang durch eine zweite Zelle benutzt, die genau hinter der ersten liegt und die für die einfallende Lichtfrequenz empfindlich ist, für welche die erste Zelle vollkommen transparent erscheint. Um eine elektrische Leistung von 100 W zu erzeugen, ist ein Silizium-Sonnenzellensystem von etwa 0,75 m 2 erforderlich. Bei der genannten Doppelschichtmethode vergrößert sich natürlich auch die Masse des Sonnenzellensystems. Für eine Leistung von 3 kW sind etwa 100 000 Zellen erforderlich, die eine Masse von 270 kg besitzen. 4.3. Thermoelektrischer Wandler Jede Erscheinung, bei der ein Austausch von thermischer und elektrischer Energie auftritt, nennen wir einen thermoelektrischen Effekt. Den Prozeß bezeichnen wir als reversibel, wenn beide Richtungen durchlaufen werden können. Das Interesse gilt hier ganz besonders den reversiblen Erscheinungen, die an Verbindungsstellen verschiedener Metalle auftreten. Der Peltier-Effekt kann als reversible Transformation potentieller elektrischer Energie und Wärme an einer Verbindungsstelle verschiedener Leiter bezeichnet werden. Unter dem Thomson-Effekt verstehen wir die reversible Transformation potentieller elektrischer Energie und Wärme bei einem endlichen Temperaturgradienten, der in jedem einfachen Leiter vorliegen kann. Da Peltier- und Thomson-Effekt reversibel sind, stellt jede Verbindung verschiedener Leiter oder jeder Teil eines Leiters, an dem ein Temperaturgradient vorliegt, eine Quelle elektromotorischer Kraft (EMK) dar. In einem vollständigen Leiter ist die Summe dieser elektromotorischen Kräfte die sogenannte Seebeck-EMK des Kreises, die man die thermoelektrische Kraft nennt. Die thermoelektrische Leistung erhält man als Quotienten aus der in einem Temperaturintervall auftretenden Änderung der thermoelektrischen Kraft und der betreffenden Temperaturdifferenz. Für ein einzelnes Element ist die thermoelektrische Leistung sehr klein. Es sind daher Hunderte von Elementen mit hohen Temperaturdifferenzen zu verwenden, um die geforderten Leistungen zu erzeugen. Nach dem heutigen Stand der Forschung sind als thermische Energiequellen Sonnenkollektoren und Kernreaktoren am besten geeignet.

20

O. Lutz, Raumfahrtforschung und Energietechnik

Jahrbuch 1961 der WGL

Die neuen Fortschritte in der Technologie des Festkörpers schufen die G r u n d l a g e n f ü r die E n e r g i e w a n d l u n g nach der thermoelektrischen Methode. Stoffe v o m P - u n d N - T y p mit hohem Atomgewicht, wie Blei, Wismut, Tellur u n d Quecksilber, w u r d e n d a z u untersucht. Die experimentellen Arbeiten zeigten, daß sich mit Wismut- u n d Wismut-TellurLegierungen W i r k u n g s g r a d e zwischen 5 u n d 7%> realisieren lassen. In den letzten J a h r e n w u r d e n insbesondere U n t e r suchungen über Stoffe in den V o r d e r g r u n d gestellt, die eine hohe thermoelektrische Leistung bei minimalem spezifischen elektrischen W i d e r s t a n d u n d minimaler thermischer Leistung ergeben. Es ist wünschenswert, diese P a r a m e t e r zu optimieren, um eine möglichst große elektrische S p a n n u n g zu erhalten u n d u m die innere Leistung u n d die Verluste durch thermische Leitfähigkeit möglichst klein zu halten. Eine Methode, die z u r Bereitstellung der thermischen Leistung einen K e r n r e a k t o r u n d z u r Erzeugung der niederen T e m p e r a t u r einen K ü h l e r b e n u t z t , ist heute von ganz besonders großem Interesse. Bei einer anderen Anlage w i r d die thermische Leistung benutzt, die durch Bündelung der Sonnenstrahlung mittels eines parabolischen Hohlspiegels erhalten w i r d u n d bei der die einfallende Strahlungsleistung auf die eine Verbindungsseite eines thermoelektrischen Elementes fokussiert w i r d . Eine elektrische Energiequelle, die das P r i n z i p des thermoelektrischen W a n d l e r s verwendet, ist beispielsweise auch der R a d i o n u k l i d e - G e n e r a t o r S N A P - 3 . Z u r Bereitstellung der thermischen Leistung w i r d hier der A l p h a - S t r a h l e r P o l o n i u m - 2 1 0 u n d z u r Konversion in elektrische Leistung P - u n d N-Blei-Tellur verwendet.

die einander in einem geringen A b s t a n d gegenüberstehen (Bild 18). D e r zwischen K a t h o d e u n d A n o d e a u f t r e t e n d e Strom ist v o n der K a t h o d e n t e m p e r a t u r , v o n der Austrittsarbeit des Kathodenmetalles u n d v o n der R a u m l a d u n g zwischen beiden Elektroden abhängig. D e r thermische W i r k u n g s g r a d liegt f ü r eine K a t h o d e n t e m p e r a t u r von 1535 ° K u n d eine A n o d e n t e m p e r a t u r v o n etwa 800 ° K bei 13°/o. D i e hierbei a u f t r e t e n d e elektrische S p a n n u n g beträgt etwa 0,5 V u n d die Stromdichte e t w a 1 A/cm 2 Kathodenfläche.

4.4.

T h e r m i o n e n -

oder

G l ü h k a t h o d e n -

w a n d l e r

Bei diesem System w i r d die thermische Energie dazu benutzt, aus einer beheizten K a t h o d e mit möglichst kleiner Austrittsarbeit Elektronen freizumachen. Dabei sind hohe K a t h o d e n t e m p e r a t u r e n erforderlich, u n d m a n braucht eine große A n z a h l solcher Systeme, u m W i r k u n g s g r a d e zu erzielen, die mit denen anderer Quellen vergleichbar sind. D a s System besteht aus einer K a t h o d e u n d einer Anode, ißlühkathode I

[Anode

~ -

AMA— lig

—

SS,'

Betastung

4.5.

Beim R a d i o n u k l i d e - G e n e r a t o r w i r d die thermische Leistung durch Zerfallsprozesse erzeugt u n d durch einen der Energiewandler in elektrische Leistung t r a n s f o r m i e r t . Beim K e r n r e a k t o r w i r d die thermische Leistung durch Spaltung der A t o m k e r n e freigemacht u n d einem Arbeitsmedium übertragen, das gleichzeitig beim Strömen durch den R e a k t o r k e r n die A u f g a b e des Kühlmittels zu erfüllen h a t . D i e thermische Kernenergie ist auch deshalb f ü r die R a u m f a h r t von g r ö ß t e r Bedeutung, weil hier die spezifischen Leistungen viel höher liegen als bei allen anderen Energiequellen u n d insbesondere die Leistungshöhe k a u m beschränkt ist. Solche R e a k t o r e n w e r d e n f ü r die elektrischen Antriebssysteme beim S t a r t v o n einer E r d u m l a u f b a h n eine g r o ß e Bedeutung erlangen. D i e Forderungen, die die N A S A a n die elektrischen Energiequellen f ü r die nächsten zehn J a h r e stellt, sind 1. h o h e spezifische Leistung (etwa bei 0,1 k W / k g ) u n d 2. Hochtemperatur--Ranfeme-Kreislauf v o n e t w a 1400 ° K . Tabelle 1 zeigt die v o n der N A S A geplanten thermoelektrischen R a d i o n u k l i d e - G e n e r a t o r e n . D i e

S N

A P - S

erie

Tabelle 1. Thermoelektrische Radionuklide-Generatoren.

Elektrische Leistung Maximale Eingangsleistung Masse Länge Durchmesser Lebensdauer Brennstoffkerne

[W] [kW] [kg] [cm] [cm]

K e r n -

D a s wissenschaftlich-technische Arbeitsteam, das f ü r die Entwicklung eines elektrischen Antriebes mit dem K e r n r e a k t o r S N A P - 8 verantwortlich ist, setzt sich aus folgenden Institutionen z u s a m m e n : N A S A : Entwicklung des Ionen- u n d P l a s m a antriebes,

Bild 18. Schema einer Thermionenquelle.

Charakteristische Daten

und

Nach den bisher erzielten Ergebnissen werden diese beiden T y p e n von Energiequellen in der zukünftigen R a u m f a h r t von ganz f u n d a m e n t a l e r Bedeutung werden. W i r d das erste System nur f ü r elektrische Leistungen bis 30 k W vorgesehen werden können, so lassen sich hingegen mit K e r n r e a k t o r e n weitaus höhere elektrische Leistungen erzeugen. Beide Systeme sind gleichermaßen v o n großer Wichtigkeit, da f ü r die einzelnen F l u g a u f g a b e n elektrische Leistungen in verschiedenen Leistungshöhen erforderlich werden.

4.6.

1!

R a d i o n u k l i d e - G e n e r a t o r r e a k t o r

SNAP-1A

SNAP-3

SNAP-9

SLLG

125 6,5 90,5 86 61 1 Jahr Ce-144

3 0,07 1,81 14 12

14,5 0,26 5,4 17,8 31,7 5 bis 10 Jahre Pu-238

19 0,65 7,5 21,3 19 4 Monate Cm-242

—

Po-210

Jahrbuch 1961 d e r W G L N A S A Jet P r o p u l s i o n L a b o r a t o r y in Kalifornien: Untersuchung der R a u m f a h r t f o r d e r u n g e n , L e w i s - R e s e a r c h - C e n t e r : Systementwicklung, A t o m i c E n e r g y C o m m i s s i o n : Kernreaktorentwicklung. Z u s ä t z l i c h z u m elektrischen R a u m f a h r t p r o g r a m m d e r N A S A m i t einem 1 - M W - R e a k t o r , der von der A t o m i c E n e r g y C o m m i s s i o n m i t Z i r k o n h y d r i d als M o d e r a t o r f ü r das S N A P - P r o g r a m m b e a r b e i t e t w i r d , h a t die U S A i r Force an die Firmen Aerojet-General u n d Air Research D i v i s i o n of G a r r e t C o r p o r a t i o n S t u d i e n a r b e i t e n ü b e r e l e k trische E n e r g i e q u e l l e n z w i s c h e n 300 u n d 1000 k W v e r g e b e n . F ü r k l e i n e L e i s t u n g e n v o n einigen W a t t bis 30 k W sind Reaktoren und Radionuklide-Generatoren vorgesehen. V o m N A S A J e t P r o p u l s i o n L a b o r a t o r y w i r d ein 1 5 - W C u r i u m - 2 4 2 - G e n e r a t o r f ü r das S u r v e y o r - P r o j e k t v o r g e schlagen, das im J a h r e 1 9 6 3 d i e weiche M o n d l a n d u n g d u r c h f ü h r e n soll. D i e F i r m a H u g h e s A i r c r a f t u n t e r s u c h t z. Z . in einer S t u d i e , welche d e r f o l g e n d e n drei E n e r g i e q u e l l e n a m v o r t e i l h a f t e s t e n ist: 1. R a d i o n u k l i d e - S y s t e m , 2. S o n n e n e n e r g i e q u e l l e , 3. B r e n n s t o f f e l e m e n t m i t B a t t e r i e n . R a d i o n u k l i d e - G e n e r a t o r e n v o n 1 bis 100 W k ö n n e n f ü r folgende Raumfahrtaufgaben verwendet werden: 1. f ü r e i n e n S a t e l l i t e n , dessen L e b e n s d a u e r auf einer E r d u m l a u f b a h n m e h r als ein J a h r b e t r ä g t , 2. M o n d f a h r t f ü r e i n e n M o n d a u f e n t h a l t l ä n g e r als einen Tag, 3. R a u m f l u g g e r ä t e in die V e n u s - A t m o s p h ä r e , 4. R a u m f l u g g e r ä t e , d e r e n F l u g r i c h t u n g v o n d e r S o n n e w e g g e r i c h t e t ist. N a c h den heutigen Erkenntnissen bieten A l p h a - u n d Beta-Strahler ausreichend hohe Leistungsdichten z u r E r z e u g u n g t h e r m i s c h e r E n e r g i e . D a b e i sind d i e A l p h a - S t r a h l e r den Beta-Strahlern überlegen, da erstere eine geringere R e i c h w e i t e h a b e n u n d d e s h a l b auch eine g e r i n g e r e A b s c h i r m u n g e r f o r d e r l i c h ist. F ü r d i e elektrische E n e r g i e g e w i n n u n g n a c h d e m t h e r m o e l e k t r i s c h e n P r i n z i p sollen diese A n l a g e n so a u s g e l e g t w e r d e n , d a ß d i e m i n i m a l e T e m p e r a t u r e t w a 800 ° K u n d d i e m a x i m a l e T e m p e r a t u r 1200 ° K b e trägt. Als t e m p e r a t u r f e s t e Werkstoffe stehen I n c o n e l - X , Hastelloy-B und Rene-41 zur Verfügung.

O. Lutz, Raumfahrtforschung und Energietechnik

21

v o n r e i n e m P l u t o n i u m ( 0 , 9 5 7 r e m / h b z w . 2 0 , 5 r e m / h in einer E n t f e r n u n g von etwa einem Meter). D i e meisten N e u t r o n e n r ü h r e n v o n d e r R e a k t i o n m i t C - l l her, e i n e m K o h l e n s t o f f i s o t o p v o n d e r A t o m m a s s e 11. SLLG (Soft Lunar L a n d i n g G e n e r a t o r ) H i e r w i r d als K e r n b r e n n s t o f f eine G o l d - C u r i u m - L e g i e r u n g (5 G e w i c h t s t e i l e A u , 1 G e w i c h t s t e i l C m ) v e r w e n d e t . D i e b e s o n d e r e n V o r t e i l e dieser L e g i e r u n g liegen in d e r leichten H e r s t e l l u n g u n d in d e r r e l a t i v g ü n s t i g e n R e a k t o r leistung. D e r B r e n n s t o f f b i l d e t e i n e n H o h l z y l i n d e r , w o b e i die B r e n n s t o f f l e g i e r u n g z u m chemischen Schutz in eine T a n t a l A u s f ü t t e r u n g e i n g e k l e i d e t ist. D a s G a n z e b e f i n d e t sich in H a s t e l l o y - C , w o d u r c h eine h o h e F e s t i g k e i t erreicht w i r d . D e r t h e r m o e l e k t r i s c h e T e i l b e s t e h t aus 30 P a a r B l e i - T e l l u r i d - E l e m e n t e n , d i e eine G e s a m t s p a n n u n g v o n 3 V l i e f e r n . R a d i o n u k l i d e - G e n e r a t o r S N A P - 1 1 2 5 W L e i s t u n g ( v g l . B i l d 19)

A

mit

D i e s e r R e a k t o r w i e g t 9 0 k g u n d l i e f e r t eine elektrische S t r o m s t ä r k e v o n 4,5 A bei 28 V S p a n n u n g a n e i n e m ä u ß e r e n W i d e r s t a n d v o n 6,3 O h m . D i e Schale, d i e die t h e r m o e l e k trischen E l e m e n t e e n t h ä l t , b e s t e h t aus z w e i h e r m e t i s c h a b geschlossenen O b e r f l ä c h e n , d i e durch einen n e u e n t w i c k e l t e n Isolierstoff M i n - K - 1 3 0 1 g e t r e n n t s i n d . D e r h e i ß e P o l ( i n n e r e Schicht) b e s t e h t aus k o h l e n s t o f f a r m e m S t a h l u n d d e r k a l t e P o l ( ä u ß e r e Schicht) aus A l u m i n i u m . D i e z y l i n d r i s c h e B r e n n s t o f f k a p s e l m i t 28 c m L ä n g e u n d 10 c m D u r c h m e s s e r ist aus I n c o n e l - X g e f e r t i g t . D i e C e r o x y d - K e r n b r e n n s t o f f e s i n d in 7 d ü n n w a n d i g e k o h l e n s t o f f a r m e S t a h l r o h r e eing e b e t t e t u n d in d e r B r e n n s t o f f k a p s e l g e l a g e r t . V o r d e m E i n s a t z dieses R e a k t o r s w i r d das I n n e r e des R e a k t o r s aus G r ü n d e n der Abschirmung mit Quecksilber gefüllt. Eine um die Brennstoffkapsel angebrachte Wasserkühlung n i m m t die thermische Energie v o m Quecksilber auf. K u r z vor dem S t a r t soll d a s Q u e c k s i l b e r abgelassen w e r d e n , so d a ß d i e

F ü r d i e U m w a n d l u n g v o n t h e r m i s c h e r in elektrische E n e r g i e k a n n f ü r L e i s t u n g e n , d i e k l e i n e r s i n d als einige 100 W a t t , auch noch d i e T h e r m i o n e n - M e t h o d e v e r w e n d e t w e r d e n , d i e a b e r erst z u e i n e m s p ä t e r e n Z e i t p u n k t g u t a u s g e r e i f t sein w i r d . F ü r d i e g e n a n n t e n R e a k t o r t y p e n s i n d f ü r die t h e r m o e l e k t r i s c h e E n e r g i e u m w a n d l u n g als M e t a l l e B l e i - T e l l u r i d e v o r g e s e h e n . Z . Z . l a u f e n auch noch U n t e r suchungen über Cadmiumsulfid, Cobaltsilizid u n d G a d o linium-Selenid. D i e R e a k t o r e n S N A P - 3 sind bereits einer R e i h e v o n P r ü f u n g e n unterzogen worden, wie Beschleunigungsversuchen in d e r G r ö ß e n o r d n u n g v o n 15 g ü b e r e i n e n Z e i t r a u m v o n e t w a 5 M i n u t e n , d i e auf d e n R e a k t o r k e i n e n E i n f l u ß z e i g t e n . S t ö ß e v o n 50 g m i t A n s t i e g z e i t e n k l e i n e r als 1 msec in d r e i z u e i n a n d e r s e n k r e c h t e n R i c h t u n g e n w u r d e n a l l e g u t überstanden. B e i m R e a k t o r S N A P - 9 , d e r f ü r 5 bis 10 J a h r e eine L e i s t u n g v o n 14,5 W l i e f e r n soll, w i r d als K e r n b r e n n s t o f f Pu 238 b z w . P l u t o n i u m k a r b i d verwendet. P l u t o n i u m k a r b i d w i r d d e s h a l b v o r g e s e h e n , w e i l d a m i t eine h ö h e r e N e u t r o n e n a u s b e u t e möglich w i r d , d e n n i m a n w e s e n d e n K o h l e n stoff findet noch e i n e z u s ä t z l i c h e ( a , ?z)-Reaktion s t a t t . D i e Strahlendosis von P l u t o n i u m k a r b i d beträgt das 20fache der

Bild 19. S N A P - l - A , ein Hochleistungs-Radionuklide-Generator mit thermoelektrischen Elementen zwischen zwei äußeren Schalen. Nach D. G. Harvey und / . G. Morse [91].

22

J a h r b u c h 1961 der W G L

O . L u t z , R a u m f a h r t f o r s c h u n g u n d Energietechnik

thermische Energie von den Brennstoffen durch Strahlung zum Innern der zwei isolierten Schalen gelangt. Kernreaktoren zur Energ iee rzeugung im R a u m In den ersten vier Jahren der R a u m f a h r t brauchten nur einige Watt elektrische Leistung f ü r einige Tage oder Wochen bereitgestellt zu werden. In der Z u k u n f t werden diese Erfordernisse ganz beträchtlich ansteigen. Als Hilfsenergiequellen werden die Reaktoren vom T y p S N A P eine Ausgangsbasis f ü r die technische Weiterentwicklung darstellen. Dabei sind folgende Kriterien in den Vordergrund zu stellen: 1. Extremer Leicht- und Kleinbau; 2. Reaktorbetrieb im Hochtemperaturbereich; 3. Minimale Kosten; 4. Leistungswandler mit höchstem Wirkungsgrad; 5. Lange Lebensdauer; 6. Einfachheit und möglichst keine beweglichen Teile. Die gewöhnlichen Moderatoren wie Wasser, Beryllium, Berylliumoxyd, schweres Wasser und Kohlenstoff zeigen bei ihrer Anwendung ganz bestimmte Charakteristiken. Die Reaktormasse steigt mit der Temperatur von Wasser-, Beryllium- und Zirkonhydrid-Moderatoren an. Die Wassermoderierung ergibt den leichtesten raumtemperierten Reaktor, aber oberhalb von 600 bis 700 ° K wird wegen des erforderlichen hohen Druckes, der das Sieden unterdrücken soll, der Druckbehälter zu schwer. Bis zu einer Temperatur von 1100 ° K läßt sich Beryllium verwenden, aber der Reaktor wiegt dann 250 kg. Die Übergangsmetallhydride können Wasserstoffdichten liefern, die mit kaltem Wasser vergleichbar werden (6,7 • 1022 Wasserstoffatome/cm 3 ), und somit einen Reaktor von derselben Größe ermöglichen, wie sie ein mit Wasser moderierter besitzt, wobei jedoch bei höherer Temperatur der hohe Druck entfällt. Zirkonhydrid ist bis etwa 950 ° K stabil. Oberhalb dieser Temperatur kann die Wasserstoffdissoziation durch eine Beimischung eines anderen Stoffes herabgesetzt werden. Diese Methode wird z. B. beim Kernreaktor S N A P - 2 angewendet, wo nur 1 °/o Wasserstoff pro J a h r verloren geht.

Bild 20. S N A P - 2 - E i n h e i t (schematisch). 1 = N a K - P u m p e , 2 = K e r n r e a k t o r , 3 = Boiler, 4 = Quecksilberpumpe, 5 = Turbine, 6 = Kondensator-Radiator, 7 = Wechselstrom-Generator.

Skala:

o

wo 200 300mm Beryllium 2

Bild 21.

Kernreaktor S N A P - 2 , Längs- und Querschnitt. Nach H. M. Dieckamp, R. Balent, J. R. Wetch [ 9 3 ] . 1 = Reaktorkern, 2 = Kernbrennstoff, 3 = Be-Kontrolltrommel, 4 = Be-Sicherheitselement, 5 = E n d r e f l e k t o r , 6 = K e r n g e h ä u s e mit 61 Z i r k o n h y d r i d - B r e n n s t o f f m o d e r a t o r s t ä b e n von 25 cm Länge.

Die Kernreaktoren der SNAP-Serie haben vollangereichertes U r a n in homogener Mischung mit dem Moderator, damit das Gewicht möglichst minimal bleibt. Als Reflektor verwendet man vorzugsweise Beryllium, da dieses Element leicht, temperaturfest und nur in einer Dicke von 5 cm erforderlich ist. Wegen der kleinen Größe treten etwa 40 °/o der bei der Kernspaltung freiwerdenden Neutronen aus dem Reaktor aus. Zur Energiewandlung werden die thermoelektrische oder die turboelektrische Methode angewendet. Die erste Methode ist f ü r Leistungen zwischen 0,1 und 1 k W von Vorteil, da keine beweglichen Teile dazu nötig sind. Der Gesamtwirkungsgrad ist klein, aber die Quelle ist von relativ langer Lebensdauer. K e r n r e a k t o r mit geschlossenem Kreislauf (turboelektrisches System) SNAP-2 (vgl. Bild 20, 21 und 22) Dieses Energieerzeugungssystem besteht aus einem kleinen Kernreaktor von kleiner Masse, der durch flüssiges N a K Eutektikum gekühlt wird, das mit einer Temperatur von etwa 920 ° K aus dem Reaktorkern ausströmt. Die vom N a K - E u t e k t i k u m im Reaktor aufgenommene thermische Energie wird zur Aufheizung des Hg-Arbeitsmediums verwendet. H g wird dadurch in einem Boiler verdampft und strömt dann in die Turbine. Das Hg-Energiewandlersystem besitzt als einziges bewegliches Teil die Welle, die

Bild 22. K e r n r e a k t o r S N A P - 2 , G e s a m t a n s i c h t . N a c h F. D. Anderson, D. J.Cockeram, H. M. Dieckamp, J.R.Wetch [94]. 1 = Kernbrennstoff, 2 = Kühlmitteleinlaß (NaK-Eutektikum), 3 = A u s t r i t t des K ü h l m i t t e l s , 4 = K o n t r o l l e l e m e n t , 5 = BeSicherheitselement, 6 = A n t r i e b s m o t o r f ü r die K o n t r o l l t r o m m e l , 7 = thermomechanische K o n t r o l l e .

O. Lutz, Raumfahrtforschung und Energietechnik

Jahrbuch 1961 der W G L

die N a K - P u m p e , die H g - P u m p e , die H g - T u r b i n e u n d den Wechselstromgenerator antreibt. Die übrigen Teile der A n lage sind der direkte K o n d e n s a t o r - W ä r m e a u s s t o ß k ü h l e r , verschiedene Kontrollsysteme, H a l t e r u n g e n u n d Abschirmungen. D i e Masse der gesamten Anlage beträgt 270 kg. S N A P - 2 besteht also aus zwei H a u p t s y s t e m e n :

23

Reactor

Reactor shield

1. R e a k t o r als thermische Energiequelle, 2. Leistungswandler. D i e thermische Energie w i r d im K e r n r e a k t o r durch Spaltung v o n U 235 erhalten u n d v o m flüssigen N a K E u t e k t i k u m beim Strömen durch den K e r n a u f g e n o m m e n . In einem Boiler w i r d die thermische Reaktorenergie vom K ü h l m i t t e l auf die Hg-Arbeitsflüssigkeit übertragen. D a durch w i r d H g v e r d a m p f t u n d durch eine T u r b i n e geleitet. Die gewonnene mechanische Energie w i r d d a n n im Wechselstromgenerator in elektrische Energie t r a n s f o r m i e r t . D e r aus der T u r b i n e austretende H g - D a m p f w i r d im R a d i a t o r Kondensor kondensiert. Diese A n l a g e soll ein Teil des R a u m f a h r z e u g e s sein. Wegen der R a u m u m g e b u n g m u ß die A u s s t o ß t e m p e r a t u r des Kreislaufes durch Abstrahlen von S t r a h l u n g s w ä r m e allein aufrechterhalten werden. Das H g K o n d e n s a t w i r d z u m Boiler zurückgeführt, was durch eine E i n l a ß p u m p e am Boiler erfolgt. Für das flüssige N a K E u t e k t i k u m w i r d eine elektromagnetische R o t o r p u m p e verwendet. Tabelle 2 gibt einige charakteristische D a t e n der K e r n r e a k t o r e n S N A P - 2 u n d S N A P - 8 . Bild 23 zeigt eine A u s f ü h r u n g von S N A P - 2 , die f ü r Flugversuche eingesetzt werden soll. D e r K e r n r e a k t o r befindet sich hier oberhalb

Radiator condenser

Boiler

Turboalternator Bild 23. SNAP-2, Flugkonfiguration. Nach H. M. R. Baient, J.R.W etch [93],

Dieckamp,

Tabelle 2. Kernreaktoren mit Turbogenerator. SNAP-2

Charakteristische Daten Ausgangsleistung Thermische Leistung Wirkungsgrad Wandler-Typ Verfügbar Lebensdauer Reaktorlänge Durchmesser Kernvolumen Thermische Leistungsdichte Reaktormasse System-Masse Spezifische Leistung Zahl der Moderatorstäbe U-235 Maximaltemperatur Neutronenfluß

[kW] [kW] [°/o]

35 bis 70 3 600 50 11,5 6 Hg (Rankine) H g (Rankine) 1965 1964 [Jahr] 1 1 [cm] 40,6 48,2 [cm] 35,5 38,0 [cm 3 ] 0,8 • 104 1,3 • 104 [W/cm 3 ] 46,2 6,1 181 113 [kg] Tkg] 340 634 [kW/kg] 0,055 bis 0,11 0,0088 37 211 Tkg] 4,3 7,0 920 1045 [°K] 2 [«/cm sec] 2,3 • 10 12 3,1 • 10»

einer konischen Abschirmung, w ä h r e n d die Turbo-Wechselspannungsgeneratoreinheit im unteren Teil angeordnet ist. Bild 24 zeigt eine Hilfsenergiequelle, bei der thermoelektrische W a n d l e r v o m S N A P - 1 0 A-System mit einem S N A P - 2 - R e a k t o r über einen elektromagnetischen N a K Kreislauf gekoppelt sind. Mit diesem System w i r d es möglich sein, eine elektrische Leistung bis etwa 1 k W bereitzustellen. Nach Erreichen einer bestimmten F l u g b a h n im R a u m kontrolliert d e r negative T e m p e r a t u r k o e f f i z i e n t der R e a k t i v i t ä t den K e r n r e a k t o r ohne weitere Bewegungen durch irgendwelche Kontrollsysteme. 4.7. Na-Kreislauf

an

der

TH

SNAP-8

Braunschweig

Z u r D u r c h f ü h r u n g von Versuchen mit flüssigem N a t r i u m ist an der T H Braunschweig ein Natrium-Versuchskreislauf (Bild 25) erstellt w o r d e n . Es handelt sich um eine Anlage,

in der eine N a t r i u m - M e n g e v o n etwa 50 k g geschmolzen und dann bei T e m p e r a t u r e n bis zu 500 ° C durch eine elektromagnetische P u m p e in einem Kreislauf umgewälzt werden k a n n . D e r K e r n r e a k t o r ist hierbei durch eine elektrische Beheizung ersetzt. A n dieser Anlage werden zunächst grundlegende E r f a h r u n g e n auf dem Gebiet der N a t r i u m Technologie gewonnen u n d das Verhalten v o n N a t r i u m Bauelementen wie P u m p e n , Ventilen, Wärmeaustauschern, Erhitzern, Kühlern, Durchfluß- u n d Füllstandsmessern studiert. D a r ü b e r hinaus lassen sich in einem Versuchsabschnitt spezielle Untersuchungen bei variabler Strömungsgeschwindigkeit u n d T e m p e r a t u r durchführen. H i e r erfolgt z. 2 . die E r p r o b u n g einer thermoelektrischen P u m p e . Dies ist eine P u m p e n f o r m , die selbstregelnd allein auf G r u n d der T e m p e r a t u r des N a t r i u m s o h n e E n e r g i e z u f ü h r u n g von außen arbeitet.

24

O. Lutz, R a u m f a h r t f o r s c h u n g u n d Energietechnik

Bild 24. Thermoelektrische W a n d l e r im S N A P - 1 0 - A - S y s t e m , die über einen elektromagnetischen N a K - K r e i s l a u f mit einem S N A P - 2 - R e a k t o r gekoppelt sind. Nach H. M. Dieckamp, R. Balent, J. R. Wetch [93].

Jahrbuch 1961 der W G L

ist kein neues Konzept. Es basiert auf dem bekannten Prinzip, daß ein elektrischer Leiter, der magnetische Feldlinien schneidet, eine E M K induziert. Die Anwendung flüssiger elektrischer Leiter ist beim Induktionsflußmesser und bei elektromagnetischen Pumpen ebenfalls bekannt. Beim konventionellen Turbogenerator wird zuerst die Dampfenergie in der Turbine in mechanische Energie umgewandelt und dann erst eine Konversion in elektrische Energie durchgeführt. Das M H D - P r i n z i p begeht f ü r eine Vielzahl von Anwendungsmöglichkeiten den Weg der unmittelbaren elektrischen Energieerzeugung. Die grundlegende Bedeutung des stationären Betriebes liegt beim thermischen Wirkungsgrad, der sich durch die vollständige Nutzbarmachung der hohen Temperaturen ergibt. Beim kombinierten offenen Turbinenkreislauf treten die aus der Turbine austretenden heißen Verbrennungsgase noch in einen M H D - W a n d l e r ein und geben dabei einen Energiebetrag ab, der sich aus einer Enthalpieabnahme ergibt. Das austretende Gas kann dann zur Vorheizung des Verbrennungsgemisches verwendet werden. Als Hilfsenergiequelle wird dieses Prinzip von größtem Interesse sein; die Temperatur kann dabei durch eine Verbrennung in einer Brennkammer oder durch das ionisierte Gas bzw. Plasma im Strömungsfeld eines Hyperschallflugkörpers bereitgestellt werden. Tabelle 3 gibt einen Oberblick über die verschiedenen Methoden zur Erzeugung strömender ionisierter Gase und eine Reihe charakteristischer Größen, die f ü r die elektrische Energieerzeugung nach dem magnetohydrodynamischen Prinzip von Bedeutung sind. Tabelle 4 stellt zwei MHD-Generator-Experimente einander gegenüber, bei denen das Edelgas Argon und ein Gemisch aus Luft und Kohlenstoff benutzt wurden. In

Bild 25. Na-Kreislauf im Institut f ü r Maschinenelemente und Fördertechnik der T H Braunschweig.

Die Arbeiten an derartigen Natrium-Anlagen sind sehr aufwendig und erfordern umfangreiche Sicherungs- und Schutzmaßnahmen wegen der Feuer- und Explosionsgefahr. 4.8. Magnetohydrodynamische Wandler Die Erzeugung elektrischer Leistungen beim Strömen eines elektrisch leitfähigen Mediums durch ein Magnetfeld

Bild 26. Schema des linearen M H D - G e n e r a t o r s . Nach R.J. Rosa [ I I I ] ,

Tabelle 3. Versuchsanlagen zur M H D - F o r s c h u n g . 2500-kWKohlelichtbogen Temperatur Geschwindigkeit Druck Kanalbreite Erzeugte Spannung Elektrische Leitfähigkeit Kanallänge Versuchszeit Erzeugte elektrische Leistung Druckabfall

[°K] 3000 bis 5000 [m/sec] bis 1500 [Atm] 1 bis 10 10 [cm] 60 [V] 1 bis 4 [1/ßcm] 60 [cm] [sec] 20 50 [kW] [Atm] 1

*) Uber einen weiten Bereich variabel.

Raketenausströmgase 3300 bis 2100 20 30 300 0,1 bis 0,7 60 30 10 0,1

LichtbogenWindkanal

Plasmastrahl

2600 3000 bis 5000 4500 900 3 IO" bis 10" 8 20 0,5 600 5 4 4 60 5 120 60 Hall-Effekt 0,001 10" 4 0,06

Beheizter Windkanal

Stoßrohr

2000

3000 bis 7000

*)

2

20 *) 0,04 *) 5

*) 15 *) 60 0,001 •••')

}

Jahrbuch 1961 der WGL

O. Lutz, Raumfahrtforschung und Energietechnik

25

Tabelle 4. Vergleich zweier MHD-Generator-Experimente. Charakteristische Größen Gas Gesamttemperatur Statische Temperatur Druck Massendurchsatz Geschwindigkeit Machzahl Alkalimetall-Zusatz Generatordimensionen Versuchsdauer Elektrodenmaterial Isolatormaterial Magnetische Feldstärke Spannung (offener Kreis) Stromstärke Leistungsdichte Elektrische Leitfähigkeit

[°K] [°K] [Atm] [kg/sec] [m/sec] [cm] [sec]

Nach R. ]. Rosa [105]

Nach F. Foshag und A.E. Were [104]

Argon 3260 2800 0,815 bis 1,36 0,333 688 0,7 1 °/o K 2 CO, 2,5 • 7,5 • 50 5

Luft + Kohlenstoff 5500 5000 5,5 0,453 1350 1,0 l°/oK 0,94 - 5 - 5 9 Graphit Silikon 8600 31 150 63 350

— —

[Gauß] [V] [A] [W/cm 3 ] [1/Q cm]

14000 55 800 20 65

H i e r a u s zeigt sich schon eine wesentliche Vereinfachung der Wirkungsweise, wobei noch der ganz besonders große Vorteil beim M H D - G e n e r a t o r herausgestellt w e r d e n m u ß , daß er keine zusätzlichen beweglichen Teile erforderlich macht. In Bild 28 ist die lineare A u s d e h n u n g eines magnetohydrodynamischen Generators von 100 M W Leistung in Abhängigkeit von der elektrischen Leitfähigkeit des Gases u n d der magnetischen Feldstärke dargestellt. D i e eingezeichneten Felder bedeuten den u n g e f ä h r e n T e m p e r a t u r bereich f ü r das ionisierte Gas. Die elektrischen Eigenschaften ionisierter Gase u n t e r scheiden sich v o n metallischen Leitern in drei P u n k t e n : 1. Ionisierte Gase sind bei 3000 ° K u n d 3 a t m Druck keine guten elektrischen Leiter; Luft h a t z. B. mit einem Zusatz v o n 1 °/o K a l i u m eine elektrische Leitfähigkeit, die 10 6 mal kleiner ist als die v o n K u p f e r . 2. Ein M H D - G e n e r a t o r m u ß grundlegend andere Techniken f ü r den Stromeingang u n d -ausgang aus dem G e n e r a t o r benutzen. 3. D e r Hall-ESekt ist bei Gasen viel stärker ausgeprägt als bei festen elektrischen Leitern. Bei diesen Untersuchungen ist es sehr vorteilhaft, die Verbrennungsgase mit einem Stoff zu mischen, der eine kleine Ionisationsenergie besitzt (Cäsium oder Kalium), damit eine ausreichende hohe elektrische Leitfähigkeit erhalten wird, die f ü r diese Anlage von ganz besonders großer Bedeutung ist. Nach G. W. Satton und L. Steg [106]. beiden Fällen w u r d e mit einem Alkalimetallzusatz gearbeitet, u n d z w a r einmal mit K a l i u m k a r b o n a t u n d im anderen Fall mit 1 % K a l i u m . Bild 26 zeigt das P r i n z i p eines linearen m a g n e t o h y d r o dynamischen Wandlers. Hierbei w i r d einem strömenden ionisierten Gas beim Durchgang durch ein M a g n e t f e l d ein gewisser E n t h a l p i e b e t r a g entnommen. In Bild 27 ist der konventionelle D a m p f - T u r b o g e n e r a t o r einem M H D - G e n e r a t o r gegenübergestellt. Beim k o n v e n tionellen G e n e r a t o r w i r d die D a m p f e n e r g i e zunächst in mechanische Energie u m g e w a n d e l t u n d d a n n erst im Gener a t o r in elektrische Energie t r a n s f o r m i e r t . Beim M H D G e n e r a t o r hingegen w i r d die elektrische Energie dem elektrisch leitfähigen ionisierten Gas direkt e n t n o m m e n .

D i e Lösung der A u f g a b e n f ü r einen M H D - G e n e r a t o r verlangen große Anstrengungen auf den Gebieten T h e r m o gasdynamik, E l e k t r o d y n a m i k u n d Quantenmechanik. D e r sogenannte Hall-Effekt verursacht eine sehr k o m p l e x e Erscheinung der elektrischen Leitfähigkeit. H i e r ist ein bedeutender Aspekt im Leistungserzeugungsmechanismus zu erkennen, d a durch die Anwesenheit eines Magnetfeldes der Stromfluß in Richtung des elektrischen Feldes reduziert w i r d . Das elektrische Feld beschleunigt Elektronen in entgegengesetzter Richtung u n d das M a g n e t f e l d b e w i r k t eine senkrecht z u r Bahn v e r l a u f e n d e Ablenkung. D a d u r c h verlaufen die Elektronenbahnen nicht m e h r geradlinig, sondern g e k r ü m m t . D i e f r e i e Weglänge der Elektronen entlang der Bahn ist k o n s t a n t ; daher w i r d die freie Weglänge in Richtung des elektrischen Feldes v e r k ü r z t u n d es tritt eine m i t t l e r e E l e k t r o n e n d r i f t o r t h o g o n a l zu den elektrischen u n d magnetischen Feldlinien auf.

26

O. Lutz, Raumfahrtforschung und Energietechnik

Diese Untersuchungen werden für die Fragen der Werkstoffkunde bedeutende neue Aussichten darstellen.

/

10 2

W'

Jahrbuch 1961 der WGL

10° Siek frische

w'

10! Vilm w3 Leitfähigkeit

Bild 28. Länge des MHD-Generators in Abhängigkeit von der elektrischen Leitfähigkeit. Angenommen ist eine Leistung von 100 MW. Die eingezeichneten Felder bedeuten die erforderlichen Gastemperaturen. Nach R.J.Rosa [111].

Verluste Die folgenden Faktoren verkleinern den Wirkungsgrad eines MHD-Generators: 1. Ohmsdie Aufheizung des Gases durch die relativ kleine elektrische Leitfähigkeit; 2. Wirbelstromverluste beim Ein- und Austritt des Gases aus dem Magnetfeld; 3. Grenzschichtverluste, viskose Druck V e r l u s t e ; 4. Anoden- und Kathodenspannungsverluste; 5. //¿//ströme. Bei der zirkulären Anordnung des magnetohydrodynamischen Generators w i r d die elektrische Leistung einem elektrisch leitfähigen Gas entnommen, das eine zum Zentrum hin gerichtete spiralförmige Bewegung zwischen zwei Elektroden in Gegenwart eines Magnetfeldes ausführt (siehe Bild 29).

•I TO

REGENERATOR

POWER FLOW

ELECTRODES FIELO COILS

6 A S FLOW

Bild 29. Schema des zirkulären MHD-Generators. Nach R.J.Coerdt, W.C.Davis, R.T.Craig, J.E.McCune [108].

Bild 31. Der MHD-Generator von Bild 30 in Einzelteile zerlegt. Nach R.J.Coerdt, W.C.Davis, R.T.Craig, J.E.McCune [108], Bild 30 und 31 zeigen eine Ausführungsform eines z. Z. in der Untersuchung stehenden MHD-Generators. Der wesentliche Bestandteil dieser Anlage ist ein hohler Flachzylinder, in den ein heißes ionisiertes Gas durch eine konventionelle Düse eingelassen wird. Die Bewegung des ionisierten Gases erfolgt spiralförmig zum Zentrum des Zylinders; es tritt dann durch zwei symmetrische axiale Öffnungen aus. Die elektrische Spannung w i r d zwischen den aus Metall bestehenden Austrittsöffnungen und einer dünnen kreisförmigen äußeren Elektrode erzeugt.

Bild 30. Foto eines kleinen MHD-Generators. Nach R.J.Coerdt, W.C.Davis, R.T.Craig, J.E.McCune

[108],

Gegenwärtige MHD-Generatoren machen hohe Arbeitstemperaturen und daher einen guten Korrosionsschutz notwendig. H i e r f ü r geeignete Materialien sind Silikone, O x y d e und Graphite.

5. S C H L U S S In den U S A hat bereits eine Umstellung von der Luftzur R a u m f a h r t in großem Ausmaße stattgefunden. In den amerikanischen Flugzeugwerken ging beispielsweise zwischen 1957 und Sommer 1960 die Produktionsfläche um etwa 5 0 % zurück (von 5,54 auf 2,77 Millionen m 2 ). Bis zum J a h r e 1961 werden nur noch 1,4 Millionen m 2 erforderlich sein, wobei allerdings auch die fortschreitende Auto-

Jahrbuch 1961 d e r W G L m a t i s i e r u n g d e r P r o d u k t i o n eine R o l l e spielt. D e r H a u p t g r u n d ist a b e r d e r , d a ß e i n e v ö l l i g e U m s t e l l u n g i n d e r Ausrüstung der L u f t w a f f e u n d die eintretende Schwerpunktv e r l a g e r u n g v o n der Flugzeug- auf die R a k e t e n p r o d u k t i o n e r f o l g t e . D e r U S - E x p o r t a n L u f t f a h r t e r z e u g n i s s e n ist v o n e t w a 1 M i l l i a r d e D o l l a r in den J a h r e n 1956 u n d 1957 auf e t w a 7 5 0 M i l l i o n e n D o l l a r z u r ü c k g e g a n g e n , w ä h r e n d in d e r R a u m f a h r t i n d u s t r i e bereits m e h r Menschen beschäftigt werd e n als in d e r A u t o m o b i l i n d u s t r i e ( 5 0 0 0 0 0 g e g e n 2 5 0 0 0 0 ) . I m Z u g e d e r s c h n e l l e n E n t w i c k l u n g s e h e n sich I n d u s t r i e v e r b ä n d e u n d Berufsorganisationen in den U S A g e z w u n gen, a u f G r u n d d e r n e u e n a n sie h e r a n k o m m e n d e n A u f g a b e n i h r e N a m e n z u ä n d e r n . Bis j e t z t ist d i e a m e r i k a n i s c h e R a u m f a h r t i n d u s t r i e fast völlig v o n A u f t r ä g e n der Regier u n g b z w . des V e r t e i d i g u n g s m i n i s t e r i u m s u n d d e r N A S A abhängig, o b w o h l genug Initiative z u r E n t w i c k l u n g eigener privatwirtschaftlicher P r o j e k t e in reichhaltiger A n z a h l v o r h a n d e n ist. E i n e R e i h e n o c h i n d e r Z u k u n f t l i e g e n d e r P r o j e k t e , z. B. d i e M o d e l l e u n b e m a n n t e r L a b o r a t o r i e n z u r E r f o r s c h u n g des Mondes u n d anderer Planeten sowie die Modelle b e m a n n t e r Außenstationen oder bemannter Raumfahrzeuge zum M o n d u n d z u d e n P l a n e t e n , w e r d e n v o n so b e d e u t e n d e n F i r m e n w i e Boeing, Douglas, Lockheed, M a r t i n , G o o d y e a r , Convair und N o r t h r o p ernsthaft durchgeführt. Die Indus t r i e z e i g t , d a ß sie b e r e i t u n d in d e r L a g e ist, m i t A n g e b o t e n a u f z u w a r t e n , noch b e v o r d i r e k t e A u f t r ä g e a n sie e r g a n g e n sind. I m gegenwärtigen S t a d i u m liegt das Schwergewicht noch a u f d e n m i l i t ä r i s c h e n P r o j e k t e n . D i e F o r s c h u n g e n u n d E n t w i c k l u n g e n k o m m e n aber nach u n d nach auch der z i v i l e n T e c h n i k z u g u t e , u n d sie w e r d e n i n s b e s o n d e r e , w i e ich h o f f e g e z e i g t z u h a b e n , d e r E n e r g i e w i r t s c h a f t b e d e u t e n d e Impulse verleihen. D i e B u n d e s r e p u b l i k d a r f sich d i e s e r E n t w i c k l u n g n i c h t v e r s c h l i e ß e n , w i l l sie n i c h t i n z e h n J a h r e n z u d e n „ u n t e r entwickelten" N a t i o n e n gehören. W i r sind f r o h , d a ß das B u n d e s k a b i n e t t b e r e i t s beschlossen h a t , sich a n d e r E L D O , dem erweiterten „ B l u e - S t r e a k " - P r o g r a m m zu beteiligen, u n d d a ß a u c h d e r B e i t r i t t z u r E S R O e n t s p r e c h e n d d e n bish e r i g e n V e r h a n d l u n g e n w a h r s c h e i n l i c h ist. D a n e b e n b l e i b t die u n e r h ö r t wichtige Z u s a m m e n a r b e i t mit unseren Freund e n in d e n U S A u n d — w i r h o f f e n es — a u c h e i n e k o l l e g i a l e Aussprache mit den Ingenieuren u n d Wissenschaftlern R u ß lands. S c h o n d i e s e w e l t w e i t e n K o n t a k t e w ä r e n es w e r t , sich m i t R a u m f o r s c h u n g u n d R a u m f a h r t zu beschäftigen. Sinn u n d Z w e c k m e i n e r A u s f ü h r u n g e n a b e r m ö c h t e n sein, z u z e i g e n , d a ß die R a u m f a h r t t e c h n i k , u n a b h ä n g i g v o n allen sozialk r i t i s c h e n Ü b e r l e g u n g e n , als m ä c h t i g e r M o t o r u n d K a t a l y s a t o r u n s e r e s technischen Z e i t a l t e r s z u g e l t e n h a t , als d e r g r o ß e Beweger u n d A n r e g e r schlechthin. 6. S C H R I F T T U M 6.1. K e r n e n e r g i e a n t r i e b [1] R. E. L. Adamson; Washington foresees m a j o r role for nuclear space p o w e r . N u c l e o n i c s 19 (1961), N r . 4 , S.54—57. [2] H. B. Finger: T h e case f o r nuclear energy. Nucleonics 19 (1961), N r . 4, S. 58—63. [3] R. E. Schreiber: K i w i tests pave w a y to Rover. Nucleonics 19 (1961), N r . 4, S. 77—79. [4] R. R. Möhler and J.E.Perry, J r . : Nuclear rocket engine control. Nucleonics 19 (1961), N r . 4, S. 80—84. [5] K. R. Pinckney: N u c l e a r rocket pneumatic control system. A R S - P r e p r i n t 1433-60 (1960). [6] D. P. MacMillan: H i g h - t e m p e r a t u r e materials for rocket reactors. Nucleonics 19 (1961), N r . 4, S. 85—88. [7] I. L. Brauch and ]. A. Connor, J r . : N u c l e a r safety in space. Nucleonics 19 (1961), N r . 4, S. 64—68. [8] G.A.Graves: Safety aspects of propulsion reactor flight testing. A R S - P r e p r i n t 1520-60 (1960).

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[9] G. F. Au: Entwicklungsstand zukünftiger R a u m f a h r t antriebe. Luftfahrttechnik 7 (1961), S. 15—25. [10] G.F.Au: R a u m f a h r t a n t r i e b e , Stand der G r u n d l a g e n f o r schung und der Entwicklung (Deutsche Aufgabenstellungen). Bericht 144 der Deutschen Forschungsanstalt f ü r L u f t f a h r t , Braunschweig 1961. [11] ]. L. Kerrebrock and R.V. Meghrehlian: An analysis of vortex tubes f o r combined gas-phase fission heating and separation of the fissionable material. O R N L - C F 57-11-3 (1957). [12] F. E. Rom: A d v a n c e d reactor concepts for nuclear rocket propulsion. Astronautics 4 (1959), N r . 10, S. 20—22 und 46—50. [13] ]. Grey: A gaseous-core nuclear rocket utilizing h y d r o dynamic containment of fissionable material. ARSP r e p r i n t 848-59 (1959). [14] J. Grey: Gaseous-core nuclear rockets. Astronautics 4 (1959), N r . 10, S. 23—25, 111 und 112. [15] R. G. Ragsdale: Analysis and experiment on an airbromine v o r t e x system. N A S A T N D - 2 8 8 (1960). [16] R. V. Meghrehlian: Gaseous fission reactors for spacecraft propulsion. J e t Propulsion L a b o r a t o r y , Pasadena, Calif., Techn. R e p . 32—42 (1960). [17] J. ]. Keyes, J r . and R. E. Dial: A n experimental study of vortex f l o w for application to gas-phase fission heating. O R N L 2837 (1960). [18] 5.7". Nelson: The plasma core reactor. Space Technology Laboratories G M 60-7630-2-9 (1960). [19] J. J. Keyes, J r . : An experimental study of f l o w and separation in vortex tubes w i t h application to gaseous fission heating. A R S - P r e p r i n t 1516-60 (1960). [20] M. L. Rosenzweig, W. S. Lewellen, J. L. Kerrebrock: The feasibility of turbulent v o r t e x containment in the gaseous fission rocket. A R S - P r e p r i n t 1516A-60 (1960). [21] H.Weinstein and R. G. Ragsdale: A coaxial f l o w reactor — a gaseous nuclear-rocket concept. A R S - P r e p r i n t 1518-60 (1960). [22] R. V. Meghrehlian: Gaseous propulsion reactors. N u c l e onics 19 (1961), N r . 4, S. 95—99. [23] R. V. Meghrehlian: Gaseous fission reactors f o r booster propulsion. Jet Propulsion L a b o r a t o r y , Pasadena, Calif., Techn. R e p . 32—56 (1961). [24] R. G. Deissler a n d M. Perlmutter: A n analysis of the energy separation in laminar and turbulent compressible vortex flows. University of C a l i f o r n i a , Berkeley, Calif. (1958). [25] J. M.Kendall, Jr.: Vortex separation experiments. J.P.L. Research S u m m a r y 36-3 (1960), Vol. 1, P a r t 2, S . 3 1 . — Experimental study of a compressible viscous vortex with applications to gaseous separation. Jet Propulsion L a b o r a t o r y , Pasadena, Calif., Techn. Release 34—168 (1960). [26] R. R. Burley and L. Bryant: Experimental investigation of coaxial jet mixing of t w o subsonic streams at various temperature, Mach-number, and diameter ratios for three configurations. N A S A M E M O 12-21-58E (1959). [27] R. W. Bussard and R. DeLauer: N u c l e a r Rocket P r o pulsion. M c G r a w - H i l l Book Comp., N e w Y o r k / T o r o n t o / London 1958, S. 322. [28] ]. L. Kerrebrock and P. G. Lafyatis: Analytical study of some aspects of vortex tubes f o r gas-phase fission heating. O R N L CF-58-7-4 (1958). [29] J. L. Kerrebrock and ]. J. Keyes, J r . : A preliminary experimental study of v o r t e x tubes for gas-phase fission heating O R N L 2660 (1959). [30] J. }. Keyes, J r . : H e a t T r a n s f e r and Fluid Mechanics Institute. S t a n f o r d University Press, S t a n f o r d , Calif., 1960, S. 31—46. [31] R.D.Evans: The Atomic Nucleus. M c G r a w - H i l l Book Comp., N e w Y o r k / T o r o n t o / L o n d o n 1955. [32] R. H. Fox: A study of the nuclear gaseous reactor rocket. Lawrence R a d i a t i o n L a b o r a t o r y , University of California, U C R L 4 9 9 6 (1957). [33] R. W. Bussard: Concepts for f u t u r e nuclear rocket propulsion. Jet Propulsion 28 (1958), S. 223—227. [34] G. Safanov: E x t e r n a l l y m o d e r a t e d reactors. Reactor Physics, Vol. 12, P a p e r 625. Proc. Second U N Int. C o n ference on Peaceful Uses of Atomic Energy, Genf 1958, S. 705—718.

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[35] R. G. Ragsdale and R. E. Hyland: Some criticality aspects of gaseous reactors. Erscheint als NASA T N . [36] C. B. Fulmer: Fission fragment studies by magnetic analysis. O R N L 2320 (1957). [37] H. Aroeste and W. C. Benton: Emissivity of hydrogen atoms at high temperatures. J. Appl. Phys. 27 (1956), Nr. 2, S. 117—121. [38] S. S. Penner: Quantitative Molecular Spectroscopy and Gas Emissivities. Addison-Wesley Publ. Comp., Inc., Reading, Mass./London 1959, S. 359—361. [39] J. W. Miser, R. E. Hyland, D. Fieno: Computer program for solving nine-group diffusion equations for cylindrical reactors. NASA MEMO 12-24-58E (1959). [40] G. V. Gordew and A. I. Gubanov: Acceleration of plasma in a magnetic field. Soviet Phys. JETP 3 (1958). [41] C. S. Wu: Transient characteristics of a rotating plasma. Jet Propulsion Laboratory, Pasadena, Calif., Techn. Release 43—122 (1960).

6.4. MHD-Antrieb [61] A. S. Penfold: Two-fluid model for the motion of a rarefied plasma accelerated by induction. ARS-Preprint 153360 (1960). [62] A. Schaffer: Plasma propulsion with a pulsed transmission line. ARS-Preprint 1534-60 (1960). [63] G. Au: Hydromagnetische Antriebs- und Steuerungssysteme für Raumfluggeräte. Raketentechnik und Raumfahrtforschung 5 (1961), S. 90—101. [64] B. Podolsky and A. Sherman: Some aspects of the Hall effect in crossed field M H D accelerators. ARS-Preprint 1531-60 (1960). [65] T. T. Reboul and C. D. Gordon: Plasma acceleration by a quasistatic R F electric field gradient. ARS-Preprint 153260 (1960). [66] S. T. Demetriades and R. W. Ziemer: Direct thrust and efficiency measurements of a continuous plasma accelerator. ARS Journ. 31 (1961), S. 1278—1280. [67] T. B. Reed: Induction-coupled plasma torch. J. Appi. Phys. 32 (1961), S. 821—824. [68] G. Lehner: Das dynamische Verhalten des Hohlpincheffektes. Z. Naturforschung 16 a (1961), S. 548—562. [69] G. Lehner: Die hydromagnetischen Instabilitäten des Hohlpincheffektes. Z. Naturforschung 16 a (1961), S. 700 bis 711.

6.2. Elektrothermischer Plasmaantrieb [42] R. R. John, / . F. Connors, A. Mironer, H. Macomber: Arc jet engine performance. Experiment and theory. ARSPreprint 1505-60 (1960). [43] R. J. Page, R. D. Buhler, W. A. Stoner, P. S. Masser: Arc plasma thrustor studies. ARS-Preprint 1508-60 (1960). [44] R. M. Spongberg: Performance characteristics of the 3000 cycle electrothermal rocket. IAS-ARS-Preprint 61-991793 (1961). [45] R. R. John, A. Mironer, H. Macomber, ]. F. Connors: Arc jet engine performance. Experiment and theory, II. IAS-ARS-Preprint 61-101-1795 (1961). [46] F. Martinek, G. F. Vaughn, W. A. Geideman, Jr.: Dynamics of arc jet system. IAS-ARS-Preprint 61-98-1792 (1961). [47] P. S. Masser: Recent experimental plasma thrustor results. IAS-ARS-Preprint 61-100-1794 (1961). [48] J. R. Jack: Regenerative and radiation cooling of electrothermal thrust generators. IAS-ARS-Preprint 61-97-1791 [49] ]. F. Cassidy, F. Martinek, M. L. Ghai: High temperature tunnel using plasma generators. ARS-Preprint 1509-60 (1960). [50] H. A. McGee, Jr. and G. Heller: Thermodynamic properties of the hydrogen, helium, and lithium plasmas. ARSPreprint 1507-60 (1960). [51] A. Q. Eschenroeder and J. W.Daiber: Ionisation nonequilibrium in expanding flows. ARS-Preprint 1458-60 (1960). 6.3. lonenantrieb [52] R. J. Hayes, Ch. A. Huebner, J. M. Glasmeyer: The duoplasmatron ion rocket. ARS-Preprint 1588-60 (1960). [53] P. D. Reader: Experimental effects of scaling on the performance of ion rockets employing electron-bombardment ion sources. IAS-ARS-Preprint 61-87-1781 (1961). [54] J. W. Davis, A. P. Walch, R. G. Meyerand, Jr., F. Salz, E. C. Lary: Theoretical and experimental description of the oscillating electron ion engine. IAS-ARS-Preprint 61103-1797 (1961). [55] R. A. Hubach and G. D. Seele: Properties of porous tungsten and ionization of cesium. IAS-ARS-Preprint 61-86-1780 (1961). [56] H. Shelton: Surface diffusion studies of cesium on tungsten. IAS-ARS-Preprint 16-85-1779 (1961). [57] M. P. Ernstene, A, T. Forrester, E. L. James, D. Telec, R. M. Worlock: Development of high efficiency cesium ion engines. IAS-ARS-Preprint 16-83-1777 (1961). [58] R. J. Cybulski and D, L. Lockwood: An experimental evaluation of a twodimensional ion engine with charge neutralization. IAS-ARS-Preprint 61-82-1776 (1961). [59] J. M. Sellen and R. F. Kemp: Cesium ion beam neutralization in vehicular simulation. IAS-ARS-Preprint 61-841778 (1961). [60] J. E. Etter, S. L. Eilenberg, J. R. Anderson, J. W. Ward, G. R. Brewer: The development of a flight test ion engine. IAS-ARS-Preprint 61-81-1775 (1961).

6.5.

Hilfsenergiequellen

6.5.1. Brennstoffelement [70] E. Justi: Höhere Energieumwandlung. Referat f ü r den Ausschuß f ü r angewandte Forschung der Deutschen Forschungsgemeinschaft (DFG). [71] E. Justi: ökonomische Erzeugung und Speicherung von Elektrizität durch Brennstoffelemente. Elektrotechn. Z. (B) 13 (1961), S. 377—386. [72] H. A. Liebhafsky and W.T.Grubb, Jr.: The fuel cell in space. ARS Journ. 31 (1961), S. 1183—1190. 6.5.2. Sonnenzelle [73] N. W. Snyder and R. W. Karcher: Solar cell power systems for space vehicles. In: Progress in Astronautics and Rocketry, Vol. 4: Space Power Systems. Academic Press, New York/London 1961, S. 3—10. [74] R.W. Karcher: Solar cell power system for advent. In: Progress in Astronautics and Rocketry, Vol. 4; Space Power Systems. Academic Press, New York/London 1961, S. 11—18. [75] R. C. Hamilton: Ranger spacecraft power system. In: Progress in Astronautics and Rocketry, Vol. 4: Space Power Systems. Academic Press, New York/London 1961, S. 19—27. [76] S. H. Winkler, I. Stein, P. Wiener: Power supply for the Tiros I meteorological satellite. In: Progress in Astronautics and Rocketry, Vol.4: Space Power Systems. Academic Press, New York/London 1961, S. 29—47. [77] W. C. Scott: Development of the power supply for the transit satellite. In: Progress in Astronautics and Rocketry, Vol. 4: Space Power Systems. Academic Press, New York/London 1961, S. 49—78. [78] W. H. Evans, A. E. Mann, I. Weiman, W. V. Wright: Solar panel design considerations: In: Progress in Astronautics and Rocketry, Vol. 4: Space Power Systems. Academic Press, New York/London 1961, S. 79—109. 6.5.3. Thermoelektrischer Wandler [79] A. C. Beer: Physics of thermoelectricity. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 3—25. [80] R. C. Miller and R. W. Ure, Jr.: Thermoelectric generator materials. In: Progress in Astronautics and Rocketry. Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 27—51. [81] S.W. Kurnick, R. L. Fitzpatrick, J.F. heavy: High-temperature semiconductors for thermoelectric conversion. In: Progress in Astronautics and Rocketry, Vol. 3:

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[82]

[83]

[84]

[85]

Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 53—61. S. E. Mayer and I. M. Ritchie: The use of high temperature thermoelectric materials (silicides) for power generation in space. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 63—72. / . C. Danko, G. R. Kilp, P. V. Mitchell: Irradiation effects on thermoelectric materials. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 73 bis 83. D. L. Kerr: Thermoelectric elements in space power systems. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 85—109. E. E. Gardner and E. L. Woisard: Thermoelectric materials for space cooling. In: Progress in Astronautics and Rocketry, V o l . 3 : Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 111—121.

6.5.4. Thermionen-Wandler [86] N. S. Rasor: Parametric optimization of the emissionlimited thermionic converter. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961. S. ARS-Preprint 1283-60 (I960). [87] A. Schock: Magneto-thermionic power generation. Electrical Engineering 79 (1960), S. 973—978. 6.5.5. Nukleare Hilfsenergiequellen [88] J. G. Morse: Nuclear Flight. Duell, Sloan and Pearce, New York 1960, S. 138. [89] S. H. Clark: Radioisotopes as sources of heat. Atompraxis 6 (1960), S. 8—11. [90] G. M. Anderson, M. Barmat, E. W. Bollmeier: SNAP-IIIelectricity from radionuclides and thermoelectric conversion. Nucleonics 17 (1959), Nr. 5, S. 166—174. [91] D.G.Harvey and J.G.Morse: Radionuclice power for space missions. Nucleonics 19 (1961), Nr. 4, S. 69—72. [92] W. Hargis et al.: Nuclear safety analysis of SNAP I I I for space missions. ARS-Preprint 1422-60 (1960). [93] H. M. Dieckamp, R. Baleni, J. R. Wetch: Compact reactors for space power. Nucleonics 19 (1961), Nr. 4, S. 73 bis 76. [94] F.D. Anderson, D. J.Cockeram, H. M. Dieckamp, J.R. Wetch: Nuclear space power — SNAP II. In: Ballistic Missile and Space Technology, Vol. I I : Propulsion and Auxiliary Power Systems. Academic Press, New York/ London 1960, S. 347—372. [95] D. P. Ross, E. Ray, E. G. Rapp, J. E. Taylor: A one-megawatt nuclear electrical power plant for space applications. In: Ballistic Missile and Space Technology, Vol. I I : Propulsion and Auxiliary Power Systems. Academic Press, New York/London 1960, S. 373—382. [96] R. ]. Rosa: A propulsion system using a cavity reactor and magneto-hydrodynamic generator. ARS-Preprint 1519-60 (I960). 6.5.6. MHD-Generator Magnetohydrodynamic [97] R. J. Rosa and A. R. Kantrowitz: energy conversion technique. Proceedings of Seminar on Advanced Energy Sources and Conversion Technique, Pasadena, Calif. (1958). [98] R. J. Rosa: Experimental magnetohydrodynamic power generator. J . Appi. Phys. 31 (1960), S. 735—736. [99] H. T. Yang: Low-speed plane Couette flow of a rarefied conducting gas in a uniform transverse magnetic field. ARS-Preprint 1457-60 (1960). [100] P. Sporn and A. Kantrowitz: Magnetohydrodynamics — future power process? Power 103 (1959), Nr. 11, S. 62 bis 65. [101] G. Wood and A. F. Carter: Dynamics of conducting gases. Proceedings of the Third Biennal Gas Dynamics Symposium "Transport Properties in Gases". Northwestern University Press, Evanston, Illinois, 1960, S. 201 bis 212.

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[102] R. ]. Rosa: Engineering magnetohydrodynamic and shock wave spectroscopy. Dissertation Cornell University

(1956)'

.

,

. ,

[103] L. Brewer: The thermodynamic and physical properties of the elements. In: The Chemistry and Metallurgy of Miscellaneous Materials, herausgegeben von L. L. Quill. McGraw-Hill Book Comp., New York/Toronto/London 1950, S. 13—39. [104] F.Foshag and A. E. Were: Magnetohydrodynamic power experiment. General Electric Comp., Report T I S R 59 SD 447 (1959). [105] R. J. Rosa: An experimental MHD power generator. Avco Everett Research Laboratories, Report AMP-42 (1960). [106] G. W. Sutton and L. Steg: The prospects for MHD power generation. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 625—661. [107] S. Way: Experiments relating to generation of power by magnetohydrodynamics. In: Progress in Astronautics and Rocketry, Vol 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 671—694. [108] R. J. Coerdt, W. C. Davis, R. T. Craig, J. E. McCune: A vortex MHD power generator. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 695—714. [109] J. E. McCune and C. du P. Donaldson: On the magnetogasdynamics of compressible vortices. In: Progress in Astronautics and Rocketry, Vol. 3: Energy Conversion for Space Power. Academic Press, New York/London 1961, S. 715—741. [110] E. E. Covert and C. W. Haldeman: The traveling wave pump. ARS-Preprint 1460-60 (1960). [111] R.J.Rosa: Physical principles of magnetohydrodynamic power generation. The Physics of Fluids 4 (1961), S. 182 bis 194. Übersicht Die extremen Forderungen, die von der Raumfahrt an alle Sparten von Forschung und Technik gestellt werden, machen gewaltige Anstrengungen notwendig, wenn die offenliegenden Aufgaben auch in der Praxis realisiert werden sollen. Andererseits sind die Befruchtungen des allgemeinen Standes der menschlichen Technik, welche die Luftfahrt in der letzten Generation gebracht hat, so außergewöhnlich und unübersehbar, daß auch auf dem Gebiet der Raumfahrt in gleichem Maße ein Wechselspiel zu erwarten ist. Dies an Hand von Beispielen aus der Triebwcrkstechnik zu zeigen, ist Inhalt des Vortrages. Gerade auf dem Gebiet der Antriebe und Energiequellen beschreiten Forschung und Entwicklung heute neue Wege, die einerseits durch die Erfolge auf dem Gebiet der Kernreaktortechnik, andererseits durch die physikalische Grundlagenforschung ganz allgemein eingeleitet wurden. In den kommenden Jahren werden die Antriebe und Antriebsquellen für die Raumfahrt prinzipiell neue Charakterzüge aufweisen, höhere Leistungen ermöglichen und die Anwendungsbereiche stark vergrößern. Die spezifischen Leistungen der Brennkammern und Energiequellen müssen durch die Erfordernisse extremer Leicht- und Kleinbauweise stark erhöht werden. Die elektrischen Antriebe ihrerseits haben bereits große Fortschritte zu verzeichnen, so daß mit ihnen bereits praktische Erprobungen vorgesehen sind und Erwartungsgrößen sich angeben lassen. Weniger rasch schreiten die Forschungen und Entwicklungen auf dem Gebiet der elektrischen Energieerzeugung vorwärts. Aber auch hier wird in einem Umfang Neuland unter den Pflug genommen, welches das Interesse jedes ernsten und aufgeschlossenen Wissenschaftlers und Ingenieurs wachrufen muß, wird doch auch auf diesem Gebiet eine Reihe bedeutender industrieller Anwendungen sichtbar. Summary The extreme demands of astronautics to all branches of research and engineering require great endeavours for realizing in practice the hitherto unsolved problems. On the other hand, the effects of aeronautics on the general technical state of art

30

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

in the last generation are so e x t r a o r d i n a r y and incalculable t h a t an interaction might be expected in astronautics as well. This p a p e r aims at showing this by means of examples taken f r o m propulsion technics. It is the field of propulsion and energy sources in which a new trend of research and development is to be felt; this being due to the success obtained in the field of nuclear chemistry and t o f u n d a m e n t a l research in physics. In the years to come propulsion units and energy sources for astronautics will have new characteristics, thus producing higher p e r f o r m a n c e and extending the applicabilities. The specific p e r f o r m a n c e of combustion chambers and energy sources has to be increased because of the requirements f o r light-weight construction and construction in small dimensions. G r e a t progress can already be registered in the field of electrical p o w e r units so t h a t practical tests are to be carried out and expectations can be indicated. Research a n d development in the generation of electric energy are lagging behind. This field, h o w e v e r , is tackled to an extent t h a t has to arouse the interest of every serious and open-minded scientist and engineer. I t becomes obvious t h a t this field o f f e r s a series of i m p o r t a n t applications to industry. Résumé Les demandes extremes posées p a r le vol spatial à tous les domaines de la recherche et de la technique imposent de grands

Jahrbuch 1961 der W G L

e f f o r t s a f i n de réaliser aussi en pratique des problèmes actuellement encore insolus. D ' a u t r e p a r t , l'influence de l'aéronautique sur la technique en général s'est montrée tellement extraordinaire et d'une portée incalculable que l'on peut s'attendre, pour le domaine de l'astronautique, à une interaction analogue. Cette conférence a pour but d'illustrer ceci au moyen de quelques exemples pris du domaine de la technique de la propulsion. C'est justement dans ce domaine et dans celui des sources d'énergie que la recherche et de développement m o n t r e n t des tendances nouvelles dues aux succès obtenus dans la chimie nucléaire et dans la recherche f o n d a m e n t a l e de physique. D a n s les années prochaines, les moyens de propulsion et les sources d'énergie p o u r le vol d'espace présenteront des caractéristiques nouvelles. Ils p e r m e t t r o n t d'obtenir des rendements plus grands et d'étendre l'applicabilité. Le rendement spécifique des chambres de combustion et des sources d'énergie doit être augmenté étant donné les exigences posées p a r une construction légère et de faibles dimensions. En ce qui concerne les ensembles de puissance électrique, de grands progrès sont à enregistrer p e r m e t t a n t de prévoir des épreuves pratiques et d'indiquer des valeurs anticipées. Les recherches et études dans le domaine de la génération d'énergie électrique sont moins avancées. Mais ici on attaque également beaucoup d'études nouvelles qui doivent inciter l'intérêt de tout ingénieur et homme de science. Il est évident que ce domaine o f f r e une série d'applications importantes à l'industrie.

B E R N H A R D H. G O E T H E R T A N D A N D R E W E. L E N N E R T , ARNOLD AIR FORCE STATION, TENNESSEE/USA*)

D E V E L O P M E N T

1,

T R E N D S

IN

R O C K E T

P R O P U L S I O N

INTRODUCTION

D u r i n g t h e past t h i r t y years, rocket propulsion systems h a v e been developed at a p h e n o m e n a l rate. F r o m t h e small r e s e a r c h r o c k e t s d e v e l o p e d b y t h e p i o n e e r s s u c h as R. H. Goddard a n d H. Obertb ( F i g . 1) t h a t c o u l d p r o d u c e o n l y a f e w p o u n d s of t h r u s t a n d b u r n i n g t i m e s of a f e w s e c o n d s w e h a v e a d v a n c e d to rocket systems h a v i n g a thrust c a p a -

Fig. 1. Goddard

and rocket.

*) A R O , a Subsidiary of S v e r d r u p and Parcel, Inc.

D K 621.455 662.2 533.9

SYSTEMS

b i l i t y of 1.5 m i l l i o n p o u n d s a n d s e v e r a l m i n u t e s b u r n i n g t i m e . T h e f u t u r e h o l d s e v e n g r e a t e r p r o m i s e . I t is i n t e r esting to r e v i e w t h e m o t i v a t i n g forces t h a t led t o present day rocket systems a n d postulate the trends f o r f u t u r e developments. T h e s u c c e s s f u l d e v e l o p m e n t of t h e i n i t i a l r e s e a r c h rockets led to t h e single m a j o r e f f o r t t h a t t r u l y spearh e a d e d t h e a c c e l e r a t e d g r o w t h of r o c k e t p r o p u l s i o n . T h i s w a s t h e d e v e l o p m e n t of t h e V - 2 m i s s i l e i n G e r m a n y d u r i n g the l a t t e r 1930's a n d early 1940's (Fig. 2). R o c k e t s h a v i n g sufficient thrust to hurl sizeable p a y l o a d s over great distances at supersonic speeds w e r e i m m e d i a t e l y recognized b y t h e m i l i t a r y as t r e m e n d o u s w e a p o n s , f o r n o e f f e c t i v e defense against them was k n o w n at t h a t time. T h e comp l e x i t i e s of t h e n e w w e a p o n s y s t e m w e r e s u c h t h a t t h e V-2 did n o t b e c o m e o p e r a t i o n a l until t h e closing m o n t h s of t h e W o r l d W a r I I . A f t e r h o s t i l i t i e s c e a s e d , o t h e r m a j o r p o w e r s , r e c o g n i z i n g its m i l i t a r y i m p o r t a n c e , c o n t i n u e d f u r t h e r d e v e l o p m e n t of t h e V - 2 p r o t o t y p e . D e t a i l e d e v a l u a t i o n of t h e V - 2 p e r f o r m a n c e c h a r a c t e r istics s u b s t a n t i a t e d t h e p r e d i c t i o n s of f a r - s i g h t e d p h y s i c i s t s t h a t missiles c o u l d b e b u i l t w i t h s u f f i c i e n t t h r u s t t o h u r l tremendous p a y l o a d s o v e r transcontinental distances. T h e U S A and Russia both developed I C B M s predicated u p o n the V - 2 design concepts; f o r example, the R o c k e t d y n e E n g i n e f o r t h e A t l a s M i s s i l e ( F i g . 3) h a v i n g a t h r u s t c a p a b i l i t y of 1 8 0 , 0 0 0 p o u n d s a n d t h e R u s s i a n c o u n t e r p a r t reportedly having greater thrust capabilities.

30

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

in the last generation are so e x t r a o r d i n a r y and incalculable t h a t an interaction might be expected in astronautics as well. This p a p e r aims at showing this by means of examples taken f r o m propulsion technics. It is the field of propulsion and energy sources in which a new trend of research and development is to be felt; this being due to the success obtained in the field of nuclear chemistry and t o f u n d a m e n t a l research in physics. In the years to come propulsion units and energy sources for astronautics will have new characteristics, thus producing higher p e r f o r m a n c e and extending the applicabilities. The specific p e r f o r m a n c e of combustion chambers and energy sources has to be increased because of the requirements f o r light-weight construction and construction in small dimensions. G r e a t progress can already be registered in the field of electrical p o w e r units so t h a t practical tests are to be carried out and expectations can be indicated. Research a n d development in the generation of electric energy are lagging behind. This field, h o w e v e r , is tackled to an extent t h a t has to arouse the interest of every serious and open-minded scientist and engineer. I t becomes obvious t h a t this field o f f e r s a series of i m p o r t a n t applications to industry. Résumé Les demandes extremes posées p a r le vol spatial à tous les domaines de la recherche et de la technique imposent de grands

Jahrbuch 1961 der W G L

e f f o r t s a f i n de réaliser aussi en pratique des problèmes actuellement encore insolus. D ' a u t r e p a r t , l'influence de l'aéronautique sur la technique en général s'est montrée tellement extraordinaire et d'une portée incalculable que l'on peut s'attendre, pour le domaine de l'astronautique, à une interaction analogue. Cette conférence a pour but d'illustrer ceci au moyen de quelques exemples pris du domaine de la technique de la propulsion. C'est justement dans ce domaine et dans celui des sources d'énergie que la recherche et de développement m o n t r e n t des tendances nouvelles dues aux succès obtenus dans la chimie nucléaire et dans la recherche f o n d a m e n t a l e de physique. D a n s les années prochaines, les moyens de propulsion et les sources d'énergie p o u r le vol d'espace présenteront des caractéristiques nouvelles. Ils p e r m e t t r o n t d'obtenir des rendements plus grands et d'étendre l'applicabilité. Le rendement spécifique des chambres de combustion et des sources d'énergie doit être augmenté étant donné les exigences posées p a r une construction légère et de faibles dimensions. En ce qui concerne les ensembles de puissance électrique, de grands progrès sont à enregistrer p e r m e t t a n t de prévoir des épreuves pratiques et d'indiquer des valeurs anticipées. Les recherches et études dans le domaine de la génération d'énergie électrique sont moins avancées. Mais ici on attaque également beaucoup d'études nouvelles qui doivent inciter l'intérêt de tout ingénieur et homme de science. Il est évident que ce domaine o f f r e une série d'applications importantes à l'industrie.

B E R N H A R D H. G O E T H E R T A N D A N D R E W E. L E N N E R T , ARNOLD AIR FORCE STATION, TENNESSEE/USA*)

D E V E L O P M E N T

1,

T R E N D S

IN

R O C K E T

P R O P U L S I O N

INTRODUCTION

D u r i n g t h e past t h i r t y years, rocket propulsion systems h a v e been developed at a p h e n o m e n a l rate. F r o m t h e small r e s e a r c h r o c k e t s d e v e l o p e d b y t h e p i o n e e r s s u c h as R. H. Goddard a n d H. Obertb ( F i g . 1) t h a t c o u l d p r o d u c e o n l y a f e w p o u n d s of t h r u s t a n d b u r n i n g t i m e s of a f e w s e c o n d s w e h a v e a d v a n c e d to rocket systems h a v i n g a thrust c a p a -

Fig. 1. Goddard

and rocket.

*) A R O , a Subsidiary of S v e r d r u p and Parcel, Inc.

D K 621.455 662.2 533.9

SYSTEMS

b i l i t y of 1.5 m i l l i o n p o u n d s a n d s e v e r a l m i n u t e s b u r n i n g t i m e . T h e f u t u r e h o l d s e v e n g r e a t e r p r o m i s e . I t is i n t e r esting to r e v i e w t h e m o t i v a t i n g forces t h a t led t o present day rocket systems a n d postulate the trends f o r f u t u r e developments. T h e s u c c e s s f u l d e v e l o p m e n t of t h e i n i t i a l r e s e a r c h rockets led to t h e single m a j o r e f f o r t t h a t t r u l y spearh e a d e d t h e a c c e l e r a t e d g r o w t h of r o c k e t p r o p u l s i o n . T h i s w a s t h e d e v e l o p m e n t of t h e V - 2 m i s s i l e i n G e r m a n y d u r i n g the l a t t e r 1930's a n d early 1940's (Fig. 2). R o c k e t s h a v i n g sufficient thrust to hurl sizeable p a y l o a d s over great distances at supersonic speeds w e r e i m m e d i a t e l y recognized b y t h e m i l i t a r y as t r e m e n d o u s w e a p o n s , f o r n o e f f e c t i v e defense against them was k n o w n at t h a t time. T h e comp l e x i t i e s of t h e n e w w e a p o n s y s t e m w e r e s u c h t h a t t h e V-2 did n o t b e c o m e o p e r a t i o n a l until t h e closing m o n t h s of t h e W o r l d W a r I I . A f t e r h o s t i l i t i e s c e a s e d , o t h e r m a j o r p o w e r s , r e c o g n i z i n g its m i l i t a r y i m p o r t a n c e , c o n t i n u e d f u r t h e r d e v e l o p m e n t of t h e V - 2 p r o t o t y p e . D e t a i l e d e v a l u a t i o n of t h e V - 2 p e r f o r m a n c e c h a r a c t e r istics s u b s t a n t i a t e d t h e p r e d i c t i o n s of f a r - s i g h t e d p h y s i c i s t s t h a t missiles c o u l d b e b u i l t w i t h s u f f i c i e n t t h r u s t t o h u r l tremendous p a y l o a d s o v e r transcontinental distances. T h e U S A and Russia both developed I C B M s predicated u p o n the V - 2 design concepts; f o r example, the R o c k e t d y n e E n g i n e f o r t h e A t l a s M i s s i l e ( F i g . 3) h a v i n g a t h r u s t c a p a b i l i t y of 1 8 0 , 0 0 0 p o u n d s a n d t h e R u s s i a n c o u n t e r p a r t reportedly having greater thrust capabilities.

Jahrbuch 1961 der W G L

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

31

Fig. 3b. Atlas missile. Fig. 2. The V-2 missile.

The decade following 1945 was one that may be characterized as a period of quiescent research on rocket systems. The basic performance of the V-2 was extended by utilizing propellant and oxidizer combinations other than liquid oxygen — alcohol resulting in increasing the specific impulse f r o m 240 seconds to 420 seconds, which is apparently the upper limit of present day chemical propulsion systems. In addition, research was focused upon improvements in the storage and handling of cryogenic fluids such as liquid oxygen and hydrogen that required intricate refri-

Fig. 3a. Rocketdyne engine for the Atlas missile.

geration systems to minimize boil-off. Propellants such as nitric acid, ammonia, and the hydrazines are stable at normal temperatures and pressures; however, because of their acidic nature they require special storage and handling facilities. This period was also marked by refine-

Fig. 4. M i n u t e m a n missile w i t h three stage solid-propellant rockets.

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B. H . Goethert and A. E. Lennert, Rocket propulsion systems

Jahrbuch 1961 der W G L

ments of h a r d w a r e for the different liquid propellants and oxidizer combinations such as valves, pumps, etc. Solid propellant rocket systems were also being developed to the point where they were competing with the liquid systems. Examples of successful solid propellant rockets are the USA Polaris and Minuteman (Fig. 4), which carry tons of payload over long distances.

known and was a natural consequence of the continued V-2 development program. This period can best be described as a refinement and perfection phase of V-2 development. It was not until the advent of Sputnik I that rocket propulsion received the added impetus that sparked unprecedented extension of space efforts. Those same boost systems that were used f o r launching missiles were found suitable to launch research payloads into varying earth orbits and also into outer space. The results of the initial successes furnished invaluable research information. Higher performance rockets were required as that larger payloads could be orbited — resulting in clustering techniques such

Vehicle staging succeeded the single stage rocket during this period. Removing some of the initial first stage tankage, structure, and h a r d w a r e resulted in an increase in performance. Theoretically, this approach was well

as the Saturn booster having eight individual Atlases to yield 1.2 to 1.5 million pounds thrust (Fig. 5). In addition larger single units are being developed such as the Rocketdyne 1.5 million pound thrust liquid propellant rocket,

Fig. 7. Bell Agena rocket engine with 15,000 lb thrust, nitric a c i d - U D M H propellants.

Jahrbuch 1961 der WGL

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

F-l (Fig. 6). Simultaneously, promising developments are being reported by the solid propellant research workers in the million pound thrust booster class. Upper stage rocket systems are being developed having long burning times and high specific impulse to provide the lower thrust-to-weight ratios when drag forces and the gravitational forces are negligible. Illustrations of these are the Bell-Agena system (Fig. 7) utilizing nitric acid-UDMH and the Pratt and Whitney using liquid oxygen-liquid hydrogen (Fig. 8). To supply the basic rocket propulsion

33

where ultimately payloads of several hundred thousand pounds, or perhaps even a million pounds, will be in evidence. 2. NUCLEAR ROCKETS In the 1950's the rocket propulsion research workers started to harness the tremendous nuclear fission energy potential for rocket propulsion in order to extend the limits of the specific impulse of 420 seconds of the chemical system. The conversion of the fission energy to useful thrust consisted merely of heating a suitable propellant to as high a temperature as possible and exhausting it through a normal converging-diverging nozzle. Hydrogen was the obvious choice in view of its low molecular weight. Initial analysis indicated the specific impulse that could be obtained with the use of the fission source to heat hydrogen was at least a factor of two better than the chemical system. A series of design studies were made, a prototype nuclear rocket built, and to date three feasibility tests of the KIWI A series have been successfully completed. 7 8

Fig. 9. Schematic of nuclear rocket. 1 = propellant tank, 2 = ejector booster pump, 3 = main pump, 4 = main control, 5 = control rod actuators, 6 = reflector (cooled), 7 = control, 8 = Vernier rocket for directional control, 9 = chamber coolant control, 10 = nozzle, 11 = reactor core, 12 = turbine bleed control (moderate temperature), 13 = turbine.

Fig. 8. Pratt and Whitney liquid-hydrogen engine, 15,000 lb thrust.

systems for the extended missions anticipated required solutions to a number of basic problems. Clustering of already proven rocket systems tended to resolve the requirement of high initial thrust; however, problems such as supporting structures, control, propellant supply and flow, and backflow of the exhaust gases still had to be solved. Other problems are those attributable to gimballing of nozzles to provide flight guidance and control, precise nozzle design to match the required expansion ratio as altitude is increased, and proper shutdown and restart sequences. If the current research and development rate continues, chemical boost capabilities will be extended to the point 3

The nuclear rocket is described in Fig. 9. The rocket is composed of uranium-loaded graphite fuel elements in an annular region. Control rods are judiciously dispersed throughout the system, although the KIWI A has a central island containing circulating D a O and control elements. Outside the fuel region a coolant, channel-pierced reflector is made of graphite, BeO, or other suitable materials exhibiting favorable high temperature material properties in addition to good neutron reflecting capabilities. A pressure vessel and other supporting structure surround the reflector. Appended to the exit end of the reactor is a convergent-divergent nozzle through which the heated gaseous propellant expands. Hydrogen is pumped around the external passages of the nozzle and then passes through the reflector coolant channels into the entrance plenum, where it initiates its major pass through the reactor. Upon leaving the reactor exit plenum, the gas, having been heated to approximately 4500 °F, expands through the nozzle into space. The specific impulse of 4500 °F hydrogen is approximately 820 seconds, at least a factor of two over the chemical counterpart. Fig. 10 compares the performance characteristics of the nuclear heated rocket with the chemical system. Even at an exit gas temperature of 3000 °F the nuclear rocket out-performs the chemical system. A number of basic problems and disadvantages are inherent in the nuclear system. Some of these problems are caused by the high temperature limitations of materials. Research and development is centered on extending these

34

B . H . G o e t h e r t a n d A . E . L e n n e r t , R o c k e t propulsion systems

limits. Thermal stresses caused by nuclear heating within the pressure vessel and other structural materials, in addition to material property changes caused by absorbed radiation, pose stringent limitations upon the nuclear rocket. The adequate removal of the nuclear afterheat attributed to fission product decay is another problem that has to be resolved. 1200

J a h r b u c h 1 9 6 1 der W G L

vehicle is required, the electrical propulsion mechanism offers more favorable characteristics than its chemical counterpart. T h e unusual capability of the electrical propulsion systems is attributed to the high values of specific impulse that characterize them. O f course, the performance of the electrical system is decreased because an excess dead weight in the form of a nuclear powerplant must be carried with vehicle which is not required of the chemical system. However, the increase in specific impulse more than compensates for the added weight. Three basic electric propulsion systems are currently undergoing extensive research and development. These are, in the order of state-of-the-art development: the electrostatic propulsion system, such as the ion and colloid; the electrothermal propulsion system, such as the arc jet and the resistance heater; and numerous magnetohydrodynamic systems. The magnetohydrodynamic propulsion systems are still in the laboratory stage and considerable research, development, and evaluation are required before they can be considered competitive.

sec 1000 t > I L

1 BOO t > L

600

3.1.

Chamber gas

temperature

F i g . 10. C h e m i c a l versus nuclear propulsion. 1 = nuclear rocket (hydrogen), 2 = nuclear rocket (ammonia), 3 = oxygen-hydrogen, 4 = ozone-hydrogen, 5 = fluorine-hydrogen, 6 = f l u o r i n e - a m m o n i a , 7 = f l u o r i n e - h y d r a z i n e .

Another basic problem is that of control. This problem is further complicated because the reactor control system has to be intimately tied into the propellant control system. Furthermore the exhaust gas temperature is dependent upon the propellant flow rate — a situation not encountered in chemical systems. Thus, the reactor power level and hence, the thrust may be independently varied by controlling the flow rate. A nuclear powerplant therefore has three parameters that must be carefully controlled: the chamber pressure, temperature, and propellant flow rate in addition to the complicated fission process control.

Electrostatic

The electrostatic propulsion system has probably received more attention in research and development than any other electrical device. This system is extremely reliable and has high specific impulse and high payload capability. I t functions in the following manner (Fig. 11): the propellant is permitted to flow into an ion-forming region with the use of proper control devices. In the ionforming region the neutral particles are vaporized, strike a tungsten surface, and give up an electron to the tungsten surface. With the aid of proper negative fields the ions are extracted from the forming region and focused into an accelerating region where they are accelerated out into space.

Flight control may not be as readily effected with the nuclear rocket as with the chemical gimballing system. A series of auxiliary nozzles are required to perform the gimballing function; hence, additional performance losses are evident. In addition to conventional problems, the harmful effects of nuclear radiation upon human beings and on electrical equipment must also be considered when evaluating the nuclear rocket system. As a second stage vehicle, in which case the harmful effects of radiation on human beings may be neglected, the nuclear heated system is a serious contender as the desirable propulsion system. Therefore, the performance advantage gained with the use of nuclear power for extended space missions is not being overlooked, and the basic nuclear rocket is rapidly approaching the flight test stage.

1 = electric valve, 3 = 5 = cesium 7 = heater, 10 = exit

3. E L E C T R I C A L P R O P U L S I O N S Y S T E M S Basically, the trend toward electrical propulsion for upper stage systems stems from the requirements of the extremely lengthy mission that ordinarily cannot be completed economically with all-chemical or chemical-nuclear propulsion systems. In addition, when continuous guidance and control throughout the operational lifetime of the

One of the basic problems is ion beam neutralization. I f a large mass of positive ions are ejected from a space vehicle, a negative charge would accumulate quite rapidly and effect reversal of the beam, thereby reducing the thrust characteristics. It is believed that this basic problem can be overcome by injecting an electron beam close to the posi-

F i g . 11. I o n engine. generator c o m p a r t m e n t , 2 = e l e c t r i c a l l y c o n t r o l l e d reservoir support insulators, 4 = cesium reservoir, feed a n d ionizer support structure, 6 = ionizer, 8 = r a d i a t i o n shields, 9 = a c c e l e r a t i n g e l e c t r o d e , and n e u t r a l i z i n g e l e c t r o d e , 11 = insulator, 12 = neutral e f f l u x m o n i t o r .

Jahrbuch 1961 der WGL

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

tive ion exit region. The problem might also be alleviated with simultaneous acceleration of both positive and negative ions [1]. Problems associated with the ion thrust chamber are related to such items as the ion source, electron source (necessary for space charge neutralization), and the ion accelerator system. With regard to propellants, the over-all propellant weight for an alkali ion system appears to vary between limits that are not too great, whether cesium or sodium is used. That is, alkali metals have atomic weights between 23 and 133, so that at constant current the propellant flow rate, hence total propellant weight, may vary by a factor of six. It may be possible to utilize heavy molecules for the propellant, such as Sbl 5 having molecular weight of approximately 757 provided they may be ionized readily. Electrode erosion is another serious problem. Considerable erosion was detected on a one-millipound thrust engine with a beam current of 60 milliamps and at an accelerating voltage level of two kilovolts [2]. The engine might have failed after 50 hours of continuous operation. Accelerator drain is another problem that must be overcome to acquire greater beam power efficiency. This deleterious effect may be remedied by careful adjustment of the accelerator voltages, the cesium supply rate, and the temperature of the ionizing surface.

35

An ion engine currently under development is schematically illustrated in Fig. 12, and operating hardware ion engines are shown in Figs. 13, 14, and 15.

Fig. 14. 7-beam ion engine.

Fig. 15. 19-beam ion engine.

Fig. 12. Schematic of multi-beam ion engine. 1 = cesium reservoir, 2 = valve, 3 = cesium, 4 = heater, 5 = radiation shields, 6 = porous Tungsten ionizer, 7 = accelerating electrode, 8 = decelerating electrode, 9 = neutralizing electrode, 10 = electrons, 11 = neutralized ion beam.

3.2. Electrothermal The second major system that has undergone extensive research and development for propulsion purposes originated from the welding, chemical processing, and flame spraying of metal fields. The basic phenomenon of electric arcs is relatively well understood [3], Basically the arc jet works in the following manner (Fig. 16): the propellant is injected into the arc chamber, and thermal energy is transferred to the propellant by joule or ohmic heating in the arc discharge effecting an increase in the enthalpy of the gas. Downstream of the arc region is a settling chamber

Specific

Fig. 13. Photograph of multi-beam ion engine.

3"

impulse

Fig. 16. Performance characteristics of the arc jet. 1 = D C electric power converter, 2 = propellant inlet, 3 = arc column, 4 = settling chamber, 5 = sonic throat, 6 = expansion nozzle, 7 = exhaust.

36

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

that permits uniform mixing of the hot and cool gases. The gas approaches the nozzle, is choked at the inlet, and expands through the nozzle where the thermal energy is converted to directed thrust. The conversion efficiency from electrical to directed thrust must be considered. Fig. 17 [4] is a plot of the experimentally determined efficiency as a function of specific impulse. The decrease in efficiency is attributed to the following: the voltage drop across an arc discharge is composed mainly of the drop in the arc plus the voltage drop in the electrodes. The only mechanism of heat transfer to the propellant is by means of the positive column drop. The electrode losses are considered as wasted heat. With a sustained arc the gas propellant becomes more highly dissociated and ionized, indicating an increase in temperature, and thus becomes a better electrical conductor. As the conductivity increases, the voltage drop decreases through the positive column, and the electrode voltage drop becomes a larger fraction of the total voltage.

Jahrbuch 1961 der W G L

be required to insure continuous arc stability. A —c arcs have been substituted for the d —c systems; however, phasing problems are introduced that further complicate the design. In addition, electrode arrangements within the arc chamber become critical. Nevertheless, the arc-jet propulsion system is in a high degree of development, and it will be flight tested simultaneously with the electrostatic propulsion system in 1962.

WOO

1500

2000 Specific

2500 sec impulse

3000

Fig. 18. Frozen flow efficiency versus specific impulse (arc jet). I 80

3.3.

Helium

% i

iO

M Specific

BOO impulse

sec

1200

Fig. 17. Overall efficiency versus specific impulse (arc jet). Mass flow

o A

V

•

2.35 4 70 1.07 1.35

gm/sec gm/sec gm/sec gm/sec

Another major factor in the conversion efficiency is that attributed to frozen flow. The frozen flow efficiency is defined as the ratio of the power available for thrust conversion to the total power delivered to the propellant. As the specific impulse is increased, the temperature of the gas is raised. Increasing the gas temperature results in an increase in dissociation and ionization. If the ions do not recombine, then the energy required to effect this process is lost, and the conversion efficiency is further reduced. Fig. 18 is a plot of the frozen flow efficiency as a function of the specific impulse of the arc jet [5], It appears that for specific impulse values greater than 1500 seconds, the arc jet may be limited to a maximum efficiency of 50 percent. Another of the important problems limiting the lifetime of the arc jet is the electrode and nozzle wear; considerable work is further required to minimize the arc electrode ablation problem at higher power levels for long periods of operation. Another problem inherent with the arc jet, particularly with the d —c arc system, is the matching of the power source to the d —c arc requirements. External controls will

Electromagnetic

The third major category in the electrical propulsion systems is the magnetohydrodynamic system (MHD), which is in the laboratory stage at present. One important feature of the magnetohydrodynamic concept is that the plasma, essentially a neutral field, overcomes one of the basic problems inherent in the electrostatic devices, that is, space charge neutralization. Furthermore, the M H D device is essentially a low temperature device and permits extension of the arc jet specific impulse because the frozen flow losses are low. Since the plasma is never heated to the stagnation temperature of the gas, the gas does not become highly ionized. The electrode losses and hence the ablation problem are also low because of the back electromotive force developed by the motion of the gas through the magnetic field. One fundamental mode of operation of the magnetohydrodynamic system using an arc source is as follows (Fig. 19): The propellant gas is preheated to a temperature of at least 4000 °K by an electric arc. The gases are choked

700

1000

2000

U000 sec 7000 10000 Specific impulse

Fig. 19. Low temperature M H D accelerator and performance.

in the nozzle to accelerate them to supersonic velocities at the entrance into the acceleration chamber. Electromagnetic field coils wound around the acceleration chamber create magnetic fields. An electric current is passed through the gas between the electrodes, and the interaction of the electric and magnetic fields effect a net force upon the plasma through the exhaust duct. A number of basic problems are inherent with the plasma M H D devices: since the energy is generally stored by capacitors, special controls are required to match the

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B. H . Goethert and A. E. Lennert, Rocket propulsion systems

load. In addition, magnetic fields are required for acceleration of the plasma, and therefore weight penalties are incurred. There are also field losses in the system that must be considered, and all these features must be resolved to ascertain the feasibility of the M H D devices. If superconductive materials are used, a number of these problems may be resolved. In general, the complete physical processes inherent with the plasma systems are quite complicated and difficult to control. The development trend has been toward the r. f. pulsed plasma system and also the magnetic induction engine. Different devices in existence are the rail gun series motor, shunt motor, a magnetic-induction plasma engine, and the shock tube [6]. 4. POWER SOURCE The basic requirement for electrical propulsion system operation is a light-weight, reliable, continuously operating powerplant. The nuclear reactor is the basic energy source for this power. A number of conversion systems have been proposed to produce electrical power from the basic nuclear energy source; however, this discussion will be limited to (1) the initial and second generation flight test vehicle powerplants of the SNAP series, and (2) direct nuclear-electrical conversion by thermionic means.

37

liquid by radiating heat directly to space and then subcooling the liquid to 420 °F for return to the mercury pump. The mercury boiler is required to transfer approximately 50 kW of heat to preheat, boil, and superheat 18.6 lb per min of mercury. The interesting feature of the system is the N a K pump which is a permanent magnet mounted in a rotor at the end of the combined rotating unit shaft, and a centrifugal pump passage for the NaK. Shaft rotation produces the swirl in the N a K in the annular passage, and the swirl provides the pumping action similar to the centrifugal pump. Even though the efficiency of the pump is approximately two percent, it was the lightest system to pump the N a K [9].

r0

WOpsia 11S0°F

kl200°F

Boiler super - lOOpsia heater 900 °F

Reactor 50Kwt

1

1000 °F NaK pump

-o 3Kwe -o Alternator

- Spsia - B00°F

TÍ50°F Condenser radiator

Mercury pump

Fig. 21. SNAP 2 cycle schematic.

The second generation SNAP powerplant is the SNAP 8 predicated upon the same design concepts as the SNAP 2 with an order of magnitude increase in the power output. The general characteristics of the two basic SNAP units are outlined in Table 1. Table 1. SNAP characteristics. SNAP 2

6 ^ Beryllium Fig. 20. Schematic of S N A P 2 core. 1 = Be safety element, 2 = reactor core, 3 = end reflector, 4 = core vessel, 5 = control element, 6 = safety element.

The nuclear powerplant that will be used for flight testing the electrical propulsion system in space is composed of homogenous mixtures of zirconium hydridide and 10 weight percent U 235 in a cylindrical bundle [7], A series of bundles are clustered within a cylindrical core surrounded by a beryllium reflector. The system is further characterized [8] by a NaK-tomercury heat exchanger, a combined turbogenerator rotating unit, and a condenser. The liquid metal loops function as follows: the primary loop, NaK, removes the heat from the reactor and conducts it to the boiler superheater where the heat is transferred to the secondary fluid, mercury. The primary coolant is at an exit temperature of 1200 °F and rejects the heat to the mercury and exits the boiler super-heater at 1000 °F. The heat transferred to the secondary coolant preheats, boils, and superheats the mercury to 1150 °F. The combined rotating unit contains on a single shaft the mercury vapor turbine, the alternator, the mercury pump, and the primary N a K pump. The condensersubcooler combination converts the mercury vapor to

Power output

[kW]

Weight without shield

[lb]

Weight with shield

[lb]

3 or 6 600

SNAP 8

SNAP 8

35

70

1,500

3,000

1,900

Unfortunately, the specific weight (lb per powerplant weight per electrical kilowatt output) is quite high for the SNAP 2 (see Table 1), approximately 100 lb per kilowatt or greater. These weight-to-power ratios may be satisfactory for the initial flight test vehicle; however, for the second generation SNAP 8 type systems even 22 lb per kilowatt is relatively high, and the specific power-to-

Fig. 22. SNAP 2 reactor. 1 = turbine generator, 2 = condenser radiator, 4 = reactor, 5 = boyler.

3 = shield.

38

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

w e i g h t ratios will h a v e to b e reduced. T h e influence of the specific p o w e r r a t i o u p o n vehicle p e r f o r m a n c e is clearly s h o w n in T a b l e 2 [ 1 0 ] which indicates that f o r a given p a y l o a d w e i g h t as the p o w e r p l a n t w e i g h t increases f o r a given electrical p o w e r o u t p u t , the p e r f o r m a n c e of a given mission decreases in terms o f increasing the time of f l i g h t . S o m e of the basic p r o b l e m s inherent with the r o t a t i n g conversion systems m a y be t r a c e d to w e a r a n d tear of the r o t a t i n g e q u i p m e n t , l e a k i n g of seals u n d e r high temp e r a t u r e conditions, a n d r a d i a t i o n i n d u c e d m a t e r i a l p r o p e r t y changes. A c o n s i d e r a b l e a m o u n t of w o r k is f o c u s e d u p o n resolving these p r o b l e m s . C o n s i d e r a b l e research a n d d e v e l o p m e n t e f f o r t is c o n c e n t r a t e d on the use of direct conversion devices such as the nuclear thermoelectric conversion a n d nuclear thermionic conversion systems. T h e thermoelectric devices will not b e discussed a t this t i m e b e c a u s e their weights b e c o m e p r o p o r t i o n a l l y l a r g e in the k i l o w a t t a n d higher p o w e r ranges. T h e direct nuclearelectric thermionic device merits f u r t h e r c o n s i d e r a t i o n .

Table 2. Flight parameters for round trip to a lunar orbit. Propulsion system

Specific impulse

Payload/ weight ratio

[sec] Chemical

Flight time, if specific weight

10lb/kW 40 lb/kW 100 lb/kW [days]

[days]

[days]

_

350 400

0.20 0.24

—

—

—

1,100 1,500

0.37 0.50

10 12

20 35

45 76

MHD

2,000 5,000 15,000

0.60 0.81 0.94

15 69 580

50 275 2,300

126 690 5,800

Ion

10,000 15,000

0.90 0.94

260 580

1,050 2,300

2,600 5,800

Arc jet

T h e thermionic-conversion system s c h e m a t i c a l l y illus t r a t e d in a vehicle ( F i g . 23) consists b a s i c a l l y of thermionic

Fig. 23. Direct conversion nuclear-electric propulsion system (ion/arc-jet thrust units). p l a s m a diodes m o u n t e d in a series of circular rings. T h e c a t h o d e s a r e in c o n t a c t w i t h the r e a c t o r f u e l element, a n d the a n o d e s f a c e r a d i a l l y o u t w a r d f r o m the p e r i p h e r y of each ring. A n u m b e r of rings a r e connected in series ( p a r a l l e l c o m b i n a t i o n ) to synthesize the p o w e r p l a n t a t the

Jahrbuch 1961 d e r W G L

desired rating. T h e w a s t e heat is rejected f r o m the a n o d e of the converters into a v a c u u m ; thus, the necessity of a s e p a r a t e cooling system a n d r a d i a t o r is o b v i a t e d . T h r e e d i f f e r e n t systems a t t a i n i n g p o w e r levels u p to 70 k i l o w a t t s with specific p o w e r ratings of 100, 60, a n d 20 lb per electrical k i l o w a t t h a v e been r e p o r t e d [ 1 1 ] . T h e i n d i v i d u a l converter efficiencies v a r y f r o m 10 to 17 p e r cent; h o w e v e r , the c a l c u l a t i o n s w e r e p r e d i c a t e d f o r 10 percent conversion efficiencies. T h e o u t p u t v o l t a g e is a p p r o x i m a t e l y 28 volts d —c at a current r a t i n g u p to a p p r o x i m a t e l y 3 0 0 0 amperes. T h e thermionic conversion unit is i d e a l l y suited f o r the arc jet system b e c a u s e no f u r t h e r t r a n s f o r m i n g devices other t h a n series-parallel m a y be required. D e v e l o p m e n t in the p r i m a r y s p a c e p r o p u l s i o n p o w e r p l a n t systems are o c c u r r i n g r a p i d l y , a n d it is believed t h a t specific p o w e r p l a n t weights will soon reach 10 lb p e r kilow a t t electric. 5. M I S S I O N T h e j u s t i f i c a t i o n f o r d e v e l o p m e n t of the electrical p r o pulsion systems is, of course, d e p e n d e n t u p o n the mission it is c a p a b l e of p e r f o r m i n g . B e c a u s e the thrust to initial w e i g h t of the electrical p r o p u l s i o n system is m u c h less t h a n unity, it is o b v i o u s that these systems are i n c a p a b l e of p e r f o r m i n g as a b o o s t system. H o w e v e r , as the vehicle gains a l t i t u d e where the d r a g forces b e c o m e negligible a n d the o r b i t a l velocity of the vehicle is s u f f i c i e n t to o v e r c o m e g r a v i t a t i o n a l forces, the thrust-to-weight r a t i o requirements a r e r e l a x e d , a n d the l o w thrust-to-weight systems such as the three electrical p r o p u l s i o n systems p r e v i o u s l y described p l a y m o r e i m p o r t a n t roles. S o m e o f the m o r e i m p o r t a n t missions that m a y b e p e r f o r m e d b y these systems a r e o r b i t a l m a n e u v e r s , position a n d a t t i t u d e control, acceleration f r o m earth o r b i t a l v e l o cities to e s c a p e velocities, l u n a r missions, s u p p l y missions to distant planets, a n d deep s p a c e t r a v e l . T h e r e is n o d o u b t that the chemical systems a n d the chemical nuclear rockets c o u l d a l s o p e r f o r m these missions; h o w e v e r , w h e n e v e r l a r g e p a y l o a d s a r e d e m a n d e d a n d when mission t i m e is not a b a s i c requirement, the electrical p r o p u l s i o n systems will perform economically. T h e d i s a d v a n t a g e of c a r r y i n g a b u l k y p o w e r p l a n t does n o t d e t r a c t f r o m the o u t s t a n d i n g p e r f o r m a n c e of these l o w thrust-to-weight r a t i o devices. T h e specific i m p u l s e a n d energy utilization of the electrical p r o p u l s i o n systems are of s u f f i c i e n t m a g n i t u d e to o v e r c o m e the a d d e d initial weight p e n a l t y of the p o w e r p l a n t . A n u m b e r o f missions h a v e been selected to m o r e clearly illustrate this p o i n t ; a n d p e r f o r m a n c e f a c t o r s in terms of mission time, p a y l o a d to initial w e i g h t ratios as a f u n c t i o n of the mission to b e p e r f o r m e d , the specific i m p u l s e required, a n d the t y p e vehicle that c o u l d b e used to e f f i c i e n t l y p e r f o r m the mission are used as p a r a m e t e r s . O n e of the possible number of missions that is being considered is a r o u n d trip to our l u n a r orbit. T h e p a y l o a d weight ratio is d e f i n e d as the useful p a y l o a d , e x c l u d i n g the electric p o w e r s u p p l y d i v i d e d by the weight, which must b e initially p l a c e d into an 150-mile c i r c u l a r orbit. F r o m T a b l e 2, it is interesting to n o t e that the arc jet w o u l d p e r f o r m f a v o r a b l y f o r a 2 0 - d a y l u n a r r o u n d trip mission. T h e p a y l o a d weight r a t i o is a p p r o x i m a t e l y 37 percent. F u r t h e r i m p r o v e m e n t s in p a y l o a d w e i g h t r a t i o m a y be o b t a i n e d if the f l i g h t t i m e is e x t e n d e d to a 5 0 - d a y mission a n d the specific i m p u l s e is increased to a p p r o x i m a t e l y 2 0 0 0 seconds ( T a b l e 2). T h e M H D device w o u l d y i e l d greater p e r f o r m a n c e . T h e use of electrostatic p r o p u l s i o n devices is not e c o n o m i c a l l y f e a s i b l e f o r short missions.

Jahrbuch 1961 der W G L

B. H . Goethert and A. E. Lennert, Rocket propulsion systems

I t is r e a d i l y e s t a b l i s h e d f r o m t h e s e d a t a t h a t e x t e n d e d missions m a y b e m o r e e f f i c i e n t l y p e r f o r m e d w i t h t h e h i g h e r / s p p r o p u l s i o n systems, p a r t i c u l a r l y w h e n t i m e is n o t t h e critical criterion. T h e second generation nuclear p o w e r p l a n t f o r space application will be the S N A P 8 system supplying a p p r o x i m a t e l y 6 0 - k W e l e c t r i c a l f o r t h e p r o p u l s i o n s y s t e m . I t is i n t e r e s t i n g t o n o t e w h a t t y p e of mission this p o w e r level w o u l d b e c a p a b l e of p e r f o r m i n g . T h e M a r s o r b i t i n g v e h i c l e could logically p r o v e to be a v a l u a b l e mission. T h e AtlasC e n t a u r B o o s t e r c o u l d p l a c e 7 5 0 0 lb i n t o a 1 , 0 0 0 - n a u t i c a l mile e a r t h orbit, a n d t h e electrical propulsion system w o u l d then be started. T h e vehicle w o u l d slowly spiral out f r o m the 1,000-mile orbit and ultimately transfer into a M a r s o r b i t a f t e r a p e r i o d of c o a s t i n g . A f t e r m a t c h i n g M a r s velocity the electrically propelled rocket w o u l d spiral down to a 500-nautical-mile Mars orbit [12].

5000 lb m 5500

7500 lb in 1000-n. mi. initial orbit lEa rthl 500- n. mifinal orbit (Mar r Thrust lbA5 .0M0 ,035 ht \—-f-6000

r

S .c:

e^

5000

—

—

j—5000

—¿000

-

iSOO 350

i-00

i50 days 500 Flight time

Fig. 24. Terminal mass versus flight time for a Centaur electric Mars orbiter capability. Fig. 24 c o n t a i n s t h e p e r f o r m a n c e c a p a b i l i t i e s of a n elect r i c a l p r o p u l s i o n s y s t e m . T h e t e r m i n a l m a s s is c o m p o s e d of t h e c o m p l e t e v e h i c l e i n c l u d i n g t h e p r o p u l s i o n s y s t e m a n d t h e p a y l o a d . I t is seen f r o m this f i g u r e t h a t if a n y t y p e of p a y l o a d is t o b e p l a c e d in t h e M a r s o r b i t , a specific i m p u l s e of a t least 4 0 0 0 seconds is r e q u i r e d . As t h e s p e c i f i c i m p u l s e increases, t h e t e r m i n a l mass increases, a n d t h e i n c r e a s e in t h r u s t level r e d u c e s t h e f l i g h t t i m e . T h e a r c jet w o u l d o b v i o u s l y n o t b e c a p a b l e of p e r f o r m i n g this mission, a n d t h e mission w o u l d b e r e l e g a t e d t o t h e ion p r o pulsion system f o r m a x i m u m p e r f o r m a n c e . T h e M H D devices w i l l p r o b a b l y h a v e t o b e c o n s i d e r e d as t h e y b e c o m e more fully developed. It should b e understood at the outset that the chemical n u c l e a r - h e a t e d p r o p u l s i o n systems m u s t h a v e g u i d a n c e a n d control capabilities to correct f o r launch errors a n d other u n c e r t a i n t i e s t h a t m a y n o t b e p r e d i c t e d , s u c h as, m e t e o r i t e i m p a c t s , etc. T h i s w a s n o t c o n s i d e r e d in t h e p e r f o r m a n c e c o m p a r i s o n s a n d t e n d s t o d e c r e a s e t h e e f f e c t i v e n e s s of t h e p r e s e n t d a y p r o p u l s i o n systems f o r e x t e n d e d missions i n t o space. T h e e l e c t r i c a l l y p r o p e l l e d s y s t e m s d o h a v e c o n tinuous guidance a n d control capability inherent with the p r o p u l s i o n u n i t . I n a d d i t i o n , t h e r e is s u f f i c i e n t p o w e r available f o r communications a n d other electrical requirem e n t s if n e e d e d . T h e t r e n d s f o r f u t u r e r o c k e t p r o p u l s i o n systems a p p e a r t o b e d i r e c t e d t o w a r d t h e use of f i r s t s t a g e c h e m i c a l a n d u p p e r stages c o m p o s e d of a n u c l e a r h e a t e d s y s t e m w h e n t i m e is a d i c t a t i n g p a r a m e t e r , o r a n u c l e a r electric s y s t e m w h e r e l a r g e p a y l o a d s a r e d e m a n d e d a n d t i m e is n o t t h e critical p a r a m e t e r .

6.

39

REFERENCES

[1] M. A. Gilleo and S. W. Kash: Propulsion by composite beams of negative and positive ions. ARS Preprint 1157-60 (1960). [2] M. P. Ernstene et al.: Development of high efficiency cesium ionen engines. IAS-ARS-Conference, Los Angeles, California, June 1961. IAS-ARS Preprint 61-83-1777 (1961). [3] T. R. Brogen: Electric arc gas heaters for re-entry simulation and space propulsion. American Rocket Society 13th Annual Meeting, N e w York, 1958. ARS Preprint 724-58 (1958). [4] M. Camac: Plasma propulsion devices. I n : Advances in Space Science and Technology, Vol. 2. Academic Press, N e w York/London 1960. [5] R. R. John, A. Mironer, H. Macomber, J. F. Connors: Arc jet engine performance; experiment and theory, II. ARS Preprint 61-101-1795 (1961). [6] W. H. Bostick: Potential aircraft applications of closed gas cycle nuclcar power plants. I n : Advanced Propulsion Systems. Pergamon Press, O x f o r d / L o n d o n / N e w York/Paris 1960. [7] R. D. Kean and R. R. Eggleston: The status of the S N A P 2 reactor. ARS Preprint 1325-60 (1960). [8] H. M. Dieckamp: The S N A P 2 concept. A R S Preprint 1324-60 (1960). [9] D. J. Cockeram and R. L. Wallerstedt: S N A P 2 svstem and vehicle integration. ARS Preprint 1329-60 (I960). [10] M. Camac: Plasma propulsion of spacecraft. Astronautics 4 (1959), N o . 10. [11] Nuclear News (September 1961). [12] D.G.Elliot and / . H. Molitor: Thrust unit requirements for electrically propelled spacecraft. Jet Propulsion Laboratory, Pasadena, Calif., TN-32-117 (1961). Übersicht Ein kurzer historischer Uberblick über die Entwicklung von chemischen Raketensystemen des Flüssig- bzw. des Festtreibstofftyps wird zu A n f a n g gegeben. Einige der wichtigsten Entwurfsparameter, wie Treibstoffkombination und Zusätze, mechanische Konstruktion, Einzel- und Vielfachdüsen, Anordnung usw., werden an H a n d von speziellen Beispielen besprochen und ihr Einfluß auf Leistungssteigerung und Leistungsgrenzen angegeben. Der Betrieb von Raketensystemen in großen H ö h e n und im Raum wird besprochen, und Versuchsergebnisse über einige Höhen- und Raumeffekte werden angeführt. Im Bereich der elektrischen Antriebe wird ein Überblick über die verschiedenen Typen solcher Antriebe (elektrothermische, elektrostatische und elektromagnetische) gegeben. Konstruktionsund Betriebscharakteristiken solcher Systeme werden besprochen. Es wird betont, daß die Entwicklung einer betriebssicheren Energiequelle geringen Gewichts von entscheidender Bedeutung f ü r die erfolgreiche Anwendung elektrischer Triebwerke ist. Möglichkeiten für die weitere Entwicklung elektrischer Triebwerke über den Stand der jetzigen „ersten Generation" hinaus werden angegeben. Der Vortrag schließt mit einer kurzen Betrachtung der Anwendungsgebiete von elektrischen Triebwerken. Summary A brief historical review of chemical propulsion systems, of both the liquid and solid propellant type, is presented. Some of the significant design parameters, like propellant combinations and additives, mechanical design features, single- and multinozzle arrangements, etc., are discussed on the basis of specific examples, and their influence on performance improvements and performance limitations is indicated. The influence of highaltitude and space environment on the operation of rocket systems is surveyed, and test results for some high-altitude effects are presented. In the field of electric propulsion, the various types of such systems (electrothermal, electrostatic, and electromagnetic) are

40

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Jahrbuch 1961 der W G L

reviewed and the construction and operational difficulties of each type are presented. I t is stressed that the need for a suitable light-weight power source is foremost for the successful development of such systems. Possible advances in the development of electric propulsion systems beyond the present first generation engines under development now, are indicated.

parle de leur influence sur l'augmentation de puissance et sur les limites de puissance. En outre, les auteurs traitent le fonctionnement de fusées en altitudes élevées et dans l'espace et donnent des résultats y obtenus.

The paper closes with a brief analysis of the fields of applications of electric propulsion systems.

Résumé D'abord les auteurs donnent un aperçu du développement des fusées à propergol liquide et solide. En donnant quelques exemples particuliers, on discute les paramètres importants (combinaison de propergol et corps additionnels, construction mécanique, tuyères séparées et multiples, disposition etc.) et on

OSWALD

H. L A N G E , H U N T S V I L L E ,

Quant au domaine des commandes électriques on fait un rapport sur les différents types de telles commandes (électrothermiques, électrostatiques et électromagnétiques). Les caractéristiques de construction et de fonctionnement de tels systèmes sont discutés par les auteurs. On souligne que le développement d'une source énergétique sûre de petit poids est très important eu égard à l'application profitable de commandes électriques. Les possibilités concernant le développement futur des commandes électriques actuelles sont indiquées. A la fin, on considère les champs d'application des commandes électriques.

ALABAMA/USA*)

D K 629.19 (73)

T R E N D S I N T H E D E V E L O P M E N T O F U . S. S P A C E V E H I C L E S

1.

INTRODUCTION

1.1. P r o g r a m s Since the last half of the 1950's, a tremendous growth has been experienced in the development, construction, and launching o f rockets. T o some extent, this growth has been stimulated by the rather glamorous connotation o f space flight, which offers so much potential adventure in those areas previously denied to man's investigation. But a f a r more important stimulation to the growth of the rocket programs has been the unexpected expansion of our knowledge and techniques. Q u i t e literally, our findings h a v e propelled us forward. E a c h area of investigation has disclosed other areas demanding to be probed. A n d each of these, in turn, draws us still further into more investigation and more discovery.

period, therefore, we can find the roots of the current rocket vehicle programs. I t should be b o r n e in mind that large carriers are normally developed from surface-to-surface ballistic missile weapons. In the period following W o r l d W a r I I , the United States placed primary emphasis upon surfaceto-air and air-to-air guided missiles. As a result, development of these classes progressed very rapidly, while, on the other hand, ballistic missile activity received very little emphasis until approximately six years ago. Thus when the

A t the root o f all this activity lies the hard headed w o r k of the engineers. I t is to the engineers, coping with day-today frustrations, problems, and set backs, that we owe the advent of the Space Age. I t is difficult, at best, to reduce theory to hardware, even when theory has been fairly well established. F a r more difficult is the building of hardware b y means of theory that is incomplete — theory in which the most significant portions are yet undefined. In this case, empirical information must be generated, then extrapolated into the future in an effort to define conditions never before experienced by man-made objects. It's just short o f a miracle that equipment designed under these conditions works at all. But it does; and to such a satisfying degree that rocket-powered vehicles capable of sending man, himself, into the fringes o f space and returning him safely to the earth, are now feasible. Approximately eleven years ago, with the establishment of the Huntsville, Alabama, development complex, the first significant steps were taken in the United States which led, ultimately, to the development of rocket launch vehicles. Approximately six years ago, industry began a full-scale participation in vehicle development. Within this time *) Marshall Space Flight Center, N A S A , Huntsville.

Fig. 1. View of Bumper research vehicle in service structure (V-2 plus W A C Corporal upper stage) at White Sands Proving Grounds, New Mexico.

40

O. H . Lange, Development of U . S. space vehicles

Jahrbuch 1961 der W G L

reviewed and the construction and operational difficulties of each type are presented. I t is stressed that the need for a suitable light-weight power source is foremost for the successful development of such systems. Possible advances in the development of electric propulsion systems beyond the present first generation engines under development now, are indicated.

parle de leur influence sur l'augmentation de puissance et sur les limites de puissance. En outre, les auteurs traitent le fonctionnement de fusées en altitudes élevées et dans l'espace et donnent des résultats y obtenus.

The paper closes with a brief analysis of the fields of applications of electric propulsion systems.

Résumé D'abord les auteurs donnent un aperçu du développement des fusées à propergol liquide et solide. En donnant quelques exemples particuliers, on discute les paramètres importants (combinaison de propergol et corps additionnels, construction mécanique, tuyères séparées et multiples, disposition etc.) et on

OSWALD

H. L A N G E , H U N T S V I L L E ,

Quant au domaine des commandes électriques on fait un rapport sur les différents types de telles commandes (électrothermiques, électrostatiques et électromagnétiques). Les caractéristiques de construction et de fonctionnement de tels systèmes sont discutés par les auteurs. On souligne que le développement d'une source énergétique sûre de petit poids est très important eu égard à l'application profitable de commandes électriques. Les possibilités concernant le développement futur des commandes électriques actuelles sont indiquées. A la fin, on considère les champs d'application des commandes électriques.

ALABAMA/USA*)

D K 629.19 (73)

T R E N D S I N T H E D E V E L O P M E N T O F U . S. S P A C E V E H I C L E S

1.

INTRODUCTION

1.1. P r o g r a m s Since the last half of the 1950's, a tremendous growth has been experienced in the development, construction, and launching o f rockets. T o some extent, this growth has been stimulated by the rather glamorous connotation o f space flight, which offers so much potential adventure in those areas previously denied to man's investigation. But a f a r more important stimulation to the growth of the rocket programs has been the unexpected expansion of our knowledge and techniques. Q u i t e literally, our findings h a v e propelled us forward. E a c h area of investigation has disclosed other areas demanding to be probed. A n d each of these, in turn, draws us still further into more investigation and more discovery.

period, therefore, we can find the roots of the current rocket vehicle programs. I t should be b o r n e in mind that large carriers are normally developed from surface-to-surface ballistic missile weapons. In the period following W o r l d W a r I I , the United States placed primary emphasis upon surfaceto-air and air-to-air guided missiles. As a result, development of these classes progressed very rapidly, while, on the other hand, ballistic missile activity received very little emphasis until approximately six years ago. Thus when the

A t the root o f all this activity lies the hard headed w o r k of the engineers. I t is to the engineers, coping with day-today frustrations, problems, and set backs, that we owe the advent of the Space Age. I t is difficult, at best, to reduce theory to hardware, even when theory has been fairly well established. F a r more difficult is the building of hardware b y means of theory that is incomplete — theory in which the most significant portions are yet undefined. In this case, empirical information must be generated, then extrapolated into the future in an effort to define conditions never before experienced by man-made objects. It's just short o f a miracle that equipment designed under these conditions works at all. But it does; and to such a satisfying degree that rocket-powered vehicles capable of sending man, himself, into the fringes o f space and returning him safely to the earth, are now feasible. Approximately eleven years ago, with the establishment of the Huntsville, Alabama, development complex, the first significant steps were taken in the United States which led, ultimately, to the development of rocket launch vehicles. Approximately six years ago, industry began a full-scale participation in vehicle development. Within this time *) Marshall Space Flight Center, N A S A , Huntsville.

Fig. 1. View of Bumper research vehicle in service structure (V-2 plus W A C Corporal upper stage) at White Sands Proving Grounds, New Mexico.

Jahrbuch 1961 der W G L

O . H . Lange, D e v e l o p m e n t of U . S. space vehicles

ballistic missile programs were accelerated, approximately six years ago, in 1955, the state of the art in rocketry was confined to the Army's single-stage Redstone and Corporal missiles, and authorization for the design initiation of the Jupiter IRBM had just been granted. (Some research vehicles, such as Viking and Bumper, had, however, already been flown by this time.) In 1955, the Air Force's Atlas ICBM was approaching the hardware stage, with flight testing still two years in the future. The Thor and Titan programs were just getting underway, while the Navy's Polaris was still two years in the future (Fig. 1).

The largest rocket flown by the United States in 1955 was the Army Redstone. This missile, burning liquid oxygen and alcohol as propellants, delivered a sea-level specific impulse of 220 pounds-seconds per pound, for 110 seconds, with a launch thrust of 75,000 pounds. Engines under development, using non-cryogenic propellants, promised increases in specific impulse from 240 to 250 poundsseconds per pound. In 1955, solid propellant rocket motors were restricted to short action times and low specific impulses of from 210 to 215 pound-seconds per pound. Since solid propellant motors had generally been employed on vehicles operating in the sensible atmosphere, aerodynamic control surfaces were used and little or no development had occurred in thrust vector control. Launching and handling in 1955 were concerned with relatively small missiles, such as Redstone and Corporal. Techniques and equipment were for the most part based on those developed and used for the V-2 rocket. These were but a few of the problem areas facing the designers when the demand for ballistic missiles and space launch vehicles erupted. Fortunately, ballistic missile systems were developed at such speed that demands for satellite and space craft carriers could readily be met. An early proposal was the use of the Redstone Missile as a satellite carrier — the Orbiter program. This proposal was superseded by the initiation of the Vanguard project. Designed purely as a research vehicle, the three-stage Vanguard was capable of injecting a 23-pound satellite into an orbit having an apogee of approximately 800 statute miles and a perigee of not more than 300 statute miles. Fourteen vehicles (2 Vikings and 12 Vanguards) were launched during the course of this program (1956—1959). Both Vikings and one of the Vanguard vehicles were used to study staging and ejection problems. The remaining eleven Vanguard vehicles were launched with the intention of injecting a satellite into orbit. Forty-three percent of the vehicles launched, successfully performed the mission

The Jupiter, Thor, and Titan were, therefore, concurrent with the Vanguard project. U p until this time, vehicle designs were, in general, highly conservative, with relatively high safety factors (1.5), and low ratios of propellant to launch weight (mass ratio). The Viking research vehicle had the highest mass ratio (0.8) of any pump-fed vehicle system flown. In the field of aerodynamics, little was known of transonic flow. Flight data was limited to that gained from the V-2, Viking, Corporal, Aerobee, and artillery shells. The heating problem was not well understood, since theory could not be confirmed until tests had been made and results studied. This was particularly true in the case of the re-entry problem, which was ultimately solved at Huntsville during Jupiter development. Multi-staging, required to reach satellite injection velocities, had been used only on the Bumper vehicle in 1949, the Aerobee (a small upper-atmosphere probe) and various small artillery-type solid-propellant rockets. While studies typified by the Atlas vehicle had proceeded to the design phase, problems such as high-altitude ignition were avoided by igniting all engines at launch. Guidance and Control hinged around continued use of carbon jet vanes and aerodynamic control surfaces. The Viking was the only vehicle which had employed gimballed engines, augmented by fixed fins and control tabs for guidance and control purposes. The gimbaled engine approach was considered superior to jet vanes, since these reduced performance and often burned or broke off, while the electro-hydraulic gimbal actuating systems were rapidly reaching high levels of efficiency and reliability. Coasting flight control had been used in the Viking system; however, the payload capability and performance required in terms of a spacecraft launch vehicle were not available. Altitude reference for Viking was provided by a conventional, medium-precision, vertical-directional gyro combination with drift rates of 1/8-degree per minute. Verticaldirectional gyro units of appreciably lower drift and capable of performing in the rocket environment were not in the hardware stage. Gimbaled inertial platforms with greater precision were in development, although they weighed in excess of 120 pounds, including electronics. However, low-drift, single degree-of-freedom integratingrate gyros were rapidly becoming available. Certain vehicles employed auto-pilots, using d-c lead circuits, a-c gyro pickoffs, d-c transfer valve and follow-up excitation, and miniature vacuum tube circuits for the proportional channels. These auto-pilots were approaching a state of high reliability. At the same time, two new approaches to autopilot design were under consideration. Magnetic amplifiers and transistor circuits were used to replace the vacuum tubes. Magnetic amplifiers appeared to offer longer life and greater reliability. Transistors promised smaller weights and volume, as well as a high degree of reliability. However, the state of development of basic transistor circuits lagged behind both vacuum tubes and magnetic amplifiers.

41

(Fig- 2).

Redstone IJunoIl

Redstone IMercuryj

Vanguard

Scout

Fig. 2. First generation of orbital launch vehicles including the solid p r o p e l l a n t Scout research vehicle.

As is well known, development problems in the Vanguard project prompted the initiation of a back-up program at Huntsville, Alabama. Here, it was decided to press the Redstone into service to launch the Explorer series of satellites. This version of Redstone, designated Jupiter C (Juno I), was originally developed as a re-entry test vehicle. Nine of these vehicles were launched from 1956 to 1958: The first three were for re-entry tests; the remaining six were used to carry Explorer satellites. All of the Explorer launchings which took place between January and October

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O. H . Lange, Development of U . S. space vehicles

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1958 were with Jupiter C. The Juno II vehicle, based on the Jupiter missile, has been used to launch Pioneer I I I and IV, the Explorers (6), VII, (8) and V I I I and Beacon (1). O u t of these seven vehicles launched f r o m Cape Canaveral, Florida, between December 1958 and November 1960, the vehicles marked with arabic numbers in parenthesis failed to orbit. The Pioneer (Solar Orbiter), launched by Juno II, permitted comprehensive investigation of cislunar space, as well as establishing a miniature space laboratory to send back data concerning the interplanetary magnetic field. It was this satellite which discovered a ring of current measuring 25,000 miles in diameter and of five million amperes which was located 40,000 miles f r o m the earth.

communication system named Project SCORE, an abbreviation of "Signal Communication Orbit Relay Experiment". T h e vehicle (Atlas-Able) had been adapted to the lunar satellite orbit projects, with four attempted launchings through December 15, 1960, all of which failed to accomplish the objective. In the Agena series, attempts were made to orbit Midas I and II and Samos I and II. Of these attempts, Midas II and Samos II were successful (Fig. 4).

The Thor is another of the family of launch vehicles adapted f r o m IRBMs. This vehicle system has been used more extensively than any other in the United States. It has been adapted to a number of exploratory missions, as evidenced by the satellites and probes carried (Fig. 3).

[I

Juno II j— j Able AgenaA ThorAgena B Delta I, Fig. 3. The second generation of orbital launch vehicles based on the J u p i t e r and T h o r ballistic missiles showing various u p p e r stages.

In illustration, the T h o r - A b l e was used for re-entry testing and to launch the following payloads: Transit I-A (navigation satellite); Explorer V I ; Pioneer I (space probe), II, and V; and Tiros I (weather satellite). The T h o r - A b l e S t a r was used to carry the following satellites: Transit IB, IIA, I I I A , and IIIB; Courier I-A (communications satellite); and Lofti I (radiation satellite). The T h o r - A g e n a was used for the Discoverer series of earth satellites, as was the Agena-B. The T h o r D e l t a was used f o r launching Tiros II (weather satellite) and Echo I (passive communications satellite). Of the 122 launch attempts of the Thor vehicle (through February 21, 1961), 86 were successful; 36 launches either failed to perf o r m the desired job or were destroyed. The S c o u t v e h i c l e has been designed specifically f o r small satellite launchings. An all-solid propellant vehicle, it has been launched eight times through February 16, 1961, on which date it placed Explorer I X into earth orbit. Of these launchings, two failures were experienced in second-stage ignition. From the engineering stand-point, the Scout vehicle family is representative of the trend to reduce vehicle cost without compromising the job to be accomplished (Fig. 2). The family of vehicles based on the Atlas represents a marked advance in space research programs. The power of Atlas permittel injection of the entire vehicle (December 18, 1958). This exercise was a test operation of a satellite

Atlas Atlas Atlas Atlas Atlas Atlas Score Mercury Able AgenaA Agena B Centaur Fig. 4. The third generation of orbital launch vehicles based on the Atlas missile showing various configurations for manned and u n m a n n e d payloads.

All of these launches have been used to place instruments into space in order to investigate the characteristics of the space environment, and, in addition, to study the earth's atmospheric environment, magnetic fields, and allied subjects. This was, of course, a logical path to follow: Learn as much as possible about space and determine the circumstances under which human survival could be assured. All of this preliminary work has enabled engineers to design man-carrying spacecraft with an extremely high probability of successfully returning the passengers. It is all well and good to visualize the man of action venturing into the unknown, for as long as he is earth-bound or operates within the confines of the atmosphere, he has a good chance of success. But to venture into hostile space, an environment to which he is not accustomed, is quite another story. Thus, it is necessary to collect as much information as possible — radiological, biological and (as you might say) humanological information — f o r this is necessary to assure survival in space. Aside f r o m the spacecraft itself, the launch vehicle presents a new challenge to the design engineers. Redundancy with mechanical over-ride capability is a must which leads to the popular phrase of "man-rating". O u r first man-in-space program, designated Project Mercury, was initiated as a result of studies conducted by N A C A in early 1956. Official program authorization was provided in August 1958. The two launch vehicles selected to support the program were the Army Redstone and the Air Force Atlas. At its outset, Mercury program had the following specific objectives: To place a manned space capsule in orbital flight around the earth; to investigate man's performance capabilities and ability to survive in a true space environment; and to recover the capsule and the man safely. Through February 21, 1961, the Mercury-Redstone 1 and 2 were test flown. Because of a minor p a d malfunction, causing ignition of the escape rocket, MR-1 left the capsule

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O . H . Lange, Development o f U . S. space vehicles

behind on the pad. Since that time, Commander Shepard and Major Grissom have successfully flown through ballistic trajectories. The Mercury-Atlas has also been tested, MA-1 malfunctioning andMA-2 performing successfully (Figs. 2 and 4). As more information about space is acquired, more is desired. The limits of vehicles designed primarily for military usage is, quite naturally, rapidly being approached. To secure additional information about the deep space environment, about the moon, Mars and Venus, more sophisticated instrumentation is required, in addition to larger and heavier spacecraft. These requirements necessarily translate themselves into larger and more powerful launch vehicles. To obtain such vehicles, thereby permitting more extensive scientific investigations of space, is the objective of Project Saturn.

figurations, the C-1, the C-5, and the giant Nova. Current program plans call for the development of the C-1, C-5, and Nova; the C-3 being eliminated because of its limited payload capability (Fig. 5).

In implementing the Vanguard back-up program (1957), the idea of a growth series on vehicle family (the Juno Series) was conceived. The Juno I launched the first Explorer satellite. The Juno II launched the first U. S. payload into solar orbit. The Juno III and I V were not developed; their missions were assigned to the Able, Delta, Agena, and Centaur series. The Juno V design was a clustered-engine, high-thrust booster. In 1957, preliminary design studies had already begun, investigating the concept of high-thrust boosters using clustered engines. One original concept included the clustering of four Rocketdyne E - l engines. This 400 kilopound thrust engine was the largest single engine under development at the time. During 1958, however, the Advanced Research Projects Agency of the Department of Defense, authorized development of a 1.5million pound thrust booster, based on the clustered engine concept. The original idea was modified to eight H - l engines, because of the cost required to bring the E - l engine to an operational status. The initial objective of the program was to demonstrate the feasibility of a high-thrust, clustered-engine booster; however, in the latter part of 1958, the program objectives were supplemented by the requirement to develop a multistage research vehicle. A static test booster and four flight boosters were authorized. The last two flight vehicles were to use an unsophisticated upper stage (preferably from "off the shelf") to achieve an early and large payload capability. Based on this new objective, studies were made of two- and three-stage configurations using IRBM and ICBM hardware. These offered reasonable capabilities but did not produce optimum performance. Continued study of the problem of Saturn upper stages evolved into the Saturn vehicle family, consisting of several basic flight con-

no ledstone Atlas

Saturn C-1 Saturn C-5

Nov:,

Fig. 5. Comparative growth sequence of spacecraft launch vehicles from the Redstone through a conceptual arrangement o f the N o v a .

43

The first Saturn configuration, the C-1, is scheduled to undergo a ten-vehicle flight test program, which will be completed during 1964. The first four C-1 vehicles consist of a live booster (S-I) stage, plus inert or ballasted upper stages (S-IV and S-V) and payload assembly. The first flight test of the C-1 configuration will take place in late 1961. The fifth vehicle of the C-1 series will use a live second stage (S-IV), replacing the inert stage used on the preceding vehicles. The fifth and succeeding vehicles of the series will be used to flight-test the Apollo spacecraft. The S-I booster stage delivers 1,500,000 pounds of thrust from a cluster of eight rocket engines developed by Rocketdyne division of North American Aviation. The engines, attached to an eight-legged thrust frame on the aft end of the vehicle, are arranged in a square pattern. The four inner (inboard) engines are rigidly attached on a threedegree angle, canted relative to the longitudinal axis of the stage. The outer engines are canted six degrees and mounted on gimbals which permit them to be turned 10 degrees, thus providing vehicle control during the first stage of powered operation. Propellants (kerosene and liquid oxygen) are stored in eight, 70-inch diameter tanks clustered around a 105-inch diameter central tank. The kerosene tanks are pressurized by gaseous nitrogen carried in 48 fiberglass spheres which are mounted to the forward structural support assembly. The liquid oxygen tanks are pressurized by gaseous oxygen bled from the heat exchangers of the propulsion system. The second stage of the C-1 configuration (the S-IV), under development by Douglas Aircraft Corporation, is powered by six rocket engines burning liquid oxygen and liquid hydrogen and will develop a total thrust of 90,000 pounds. All engines gimbal for control purposes. The S - I V stage is connected to the booster by an interstage structure which remains with the lower stage upon flight separation. The S - I V stage structure, itself, is an integral, tandem-arranged propellant tankage, with the liquid hydrogen section forward and the oxygen section aft. These compartments are separated by a single insulated bulkhead. The stage represents a major break-through in the development of rocket vehicles. For many years, it has been known that the combination of hydrogen and oxygen represented the most desirable thermochemical propellant combination. To harness the power of these propellants and to obtain the high performance necessary for currently planned space flight research programs, large-scale hardware development has been initiated by the Saturn program. The initial Saturn vehicle will use a guidance and control system adopted from early ballistic missiles. The system consists basically of a guidance computer, control computer, and stabilized platform. The guidance and control scheme of the Saturn anticipates the so-called "man-rating" requirement necessitated by the use of Saturn as a launch vehicle for man-carrying spacecraft. In this connection, if one of the booster engines fails and is extinguished, the vehicle must continue to perform its mission. The guidance system will automatically give corrective signals necessary to compensate for the thrust loss. The initial Saturn configuration has the capability of placing 10 tons of payload into earth orbit and/or carrying three men into orbit.

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Jahrbuch 1961 der W G L 1.2. S t r u c t u r e s

The next configuration of Saturn is designated C-5 — a large step f o r w a r d in the development of launch vehicles. The clustered tanks of the C - l booster stage will be replaced by an integral tank design to which five huge kerosene/liquid oxygen burning engines will be attached. These engines (designated F - l ) are under development by Rocketdyne Division of N o r t h American Aviation and will each produce 1,500,000 pounds thrust. As arranged on the C-5 booster, they will produce a total of 7,500,000 pounds thrust, the new booster stage being five times as powerful as the stage employed with the C - l configuration.

The rocket engineer's dream is to convert virtually all of the launch vehicle's weight into propulsive energy. There is, of course, a practical limitation to this dream, f o r the structure of the launch vehicle, itself, contributes the major portion of the so-called "inert" weight. It is extremely difficult to reduce structural weight, because the structural design of the vehicle must resolve a number of interrelated problem areas. Some of the areas which must be considered are:

The second stage (S-II) of the C-5 configuration will be powered by five engines, each developing 200,000 pounds thrust and using liquid hydrogen/liquid oxygen propellants. This stage is currently under development by N o r t h American Aviation. The third stage of the C-5 configuration will be a modified S-IV stage, using one engine developing 200,000 pounds thrust. In consequence, the C-5 version of the Saturn launch vehicle will have a total thrust output of almost 9,000,000 pounds. It will be capable of placing about 120 tons into earth orbit or to escape almost 45 tons. Following this configuration of Saturn is the N o v a vehicle, an even greater step f o r w a r d in the development of launch vehicles. The first stage of N o v a will consist of a cluster of eight 1,500,000-pounds thrust, single-chamber kerosene engines, developing total thrust of 12,000,000 pounds.

As a result of past launch experience, a set of rather standard criteria have been evolved f o r flight and ground loads. Rigidity is frequently considered a design characteristic, rather than a design requirement. For Saturn, however, stiffness is a design requirement, based on the ultimate use of the vehicle f o r launching rather sophisticated spacecraft, rather than the capsule-type craft in current use.

The second stage might use four 1,200,000 pound thrust engines, burning liquid oxygen and liquid hydrogen propellants. Total thrust developed will be 4,800,000 pounds. The third stage might use two of the same 200,000-pound thrust engines. The total thrust generated by this vehicle is estimated at 17,600,000 pounds. (To realize the growth represented by this vehicle, it should be noted that the Mercury-Redstone develops a total of 78,000 pounds thrust.) With such power available, it will be possible to place a three-man spacecraft on the moon and return it to earth. It would also be feasible to send three men on a voyage around Mars and return them to earth. With these basic Saturn-family configurations ( C - l , C-5 and Nova), we will have the potential to conduct a comprehensive exploration of space. First, however, we must use instrumented payloads to gather data necessary to determine man's survival requirements in space. A short time afterward, we will witness the expedition of man, himself, into the cosmos. The three basic Saturn-family configurations must be manufactured, tested, transported to the launch site, then launched. These operations are tremendous undertakings in themselves, and have required development of special tooling, handling equipment, check-out and testing equipment, and launching facilities. The "Programs" aspect leading to large launch vehicle development could not have been possible without the support of the research and development activity which made the necessary components available. This aspect of the development activity, begun in the 1940's, was built upon work conducted in Germany and the United States during 1920/1930 period: rocket engines, turbines, pumps, electrical generators and batteries, propellants, guidance and control equipment, hydraulic systems, pressurizing equipment, and materials f r o m which these components were constructed. Even gaskets and seal materials and configurations were exhaustively investigated under both cryogenic and extremely corrosive environments.

Flight and ground loads, Rigidity requirements, Safety factors and allowable stress, Analysis requirements (and at what stage of design these should be finalized), Structural test program.

Safety factors and allowable stress are always difficult to predetermine, especially when man-carrying spacecraft are to be launched. The flight test program is a very imp o r t a n t adjunct to this phase of design. For example, on one of the large launch vehicles, the design engineers spared every ounce f r o m the structure to w h a t they thought was an absolute minimum. However, during the flight test program, one vehicle malfunctioned and actually flew sideways f o r several seconds before command destruction was carried out. It was observed that the vehicle did not break u p during this time. This observation led to revisions of safety factors and allowable stress permitting marked improvements in mass ratios. The structural analysis requirements are therefore a continuing function of development. U p o n completion of the flight test program, sufficient information gathered under actual conditions permits finalization of specification

Multiple

containers

r\ ÄWIi

Fig. 6. Schematic representation of various p r o p e l l a n t t a n k configurations.

Jahrbuch 1961 der WGL

O. H . Lange, Development of U. S. space vehicles

requirements. In other words, the criteria for analysis must be based on the specific job to be done; reasonable generalizations in this area are virtually impossible to develop. The launch vehicle itself is basically cylindrical in shape, compartmentalized for convenience into three sections, the forward or payload section, the center section, and the aft section. The largest section by far — the center section — is occupied by the propellant tanks. The general dimensions of this section are determined primarily by the type of propellants used and by the performance requirements of the vehicle (Fig. 6). Three types of tank geometry immediately suggest themselves: tandem, concentric, and multiple. The tandem arrangement is no stranger to rocket engineers, for it is used on virtually every current major launch vehicle. The concentric type of tank arrangement is a very reasonable means of packaging considerable performance into a vehicle without unduly extending the fineness ratio beyond practical limits — particularly when propulsive spacecraft are desired for deep space probes. The clustered arrangement, a possible scheme, is used for S-I of the Saturn C - l vehicle and is planned for a Nova-type booster. The approach has the advantage of obtaining high thrust with existing hardware and technology, thus permitting early launchings of advanced spacecraft and instrumented probes. In determining the overall shape of the launch vehicle, many considerations must be carefully weighed. From the aerodynamic point of view, these considerations are related to air loads, stability, performance, and duration. In the case of Vanguard, the procedure was straight-forward, because the entire vehicle was designed as an entity. However, with the evaluation of plans for more extensive space missions than could be performed by a single-stage vehicle the multi-staging of vehicles became necessary. Problems attendent to growth potential, as planned for the Saturn family, makes the problem more complex. As additional stages are added to the basic vehicle, mechanical and electrical mating, structural rigidity, longitudinal loadings, vibrational characteristics, and guidance and control problems must be anticipated and solved.

helium sphere, due primarily to reduction of maximum sphere pressure. Aluminum, rather than steel, was used for the thrust chamber of the second stage engine, thus saving 60 pounds but at the price of fabrication problems and a marginal chamber lifetime.

In addition to carrying the various dynamic and static loads, the structure is subjected to thermal loads, particularly when cryogenic propellants are used. The Saturn Stage I, for example, has forward fuel tank mounts on expansion joints which allow for contraction of the structure as liquid oxygen is loaded aboard. Because of the penalty that weight imposes on performance, every effort is made to keep the vehicle weight to a minimum. The Vanguard vehicle is a typical example of the weight savings programs conducted during vehicle development. In this case, as in most others, "rules-ofthumb" are used to gage the general effectiveness of possible weight savings. These "rules" are for the most part of an empirical nature being originated by experience in handling the hardware. For Vanguard, the reduction in final payload injection velocity caused by the addition of one pound to the burnout weight of a given stage is as follows: First Stage: 1 ft/sec, Second Stage: 8 ft/sec, Third Stage: 80 ft/sec. The weight savings that were effected on the Vanguard vehicle were done at the expense of increased systems complexity. Two specific examples will serve to illustrate the difficult "trade-off" in reaching a decision to save weight. A savings of 71 pounds in structural weight was indicated by the use of a heating charge in the second-stage

1.3.

S t a g i n g

and

45

S e p a r a t i o n

The use of multistage rocket vehicles has been rather common over the past few years. The problems solved by multi-staging represents a major advance in the development technology of rockets. The velocity limitation of single-stage rockets, even when using all modern designs, is shown in Fig. 7. To supply the incremental velocities (Fig. 8) necessary for space exploration, additional stages are required. The use of a multiplicity of stages presents many unique problems in such areas as manufacture, alignment, separation, and thrust allocation (size optimization). The first successful staging of a research vehicle was accomplished at White Sands Proving Ground with the Bumper Project. This vehicle consisted of two stages; the V-2 topped by the Jet Propulsion Laboratories WAC-Corporal (Fig. 1). The first launching took place May 13, 1948, the last launching (the eighth), July 29, 1950. This vehicle established a record by attaining velocities of 5,000 miles per hour and reaching an altitude of 250 miles. 100

I1

50

"o nj o < ci0

05

10

5

2

1 0

5

10

15

20 • W3 ft/sec

Velocity

Fig. 7. Ratio of vehicle launch weight/payload weight versus cutoff velocity reflecting the design growth potential of singlestage rocket vehicles.

The theoretical aspect of staging, or of obtaining optimum thrust distribution throughout a desired flight profile, was studied by H. S. Tsien and R. C. Evans. The general problem was formulated as determining the optimum time variation of thrust for maximum altitude. This theoretical approach is complicated by the purely engineering aspects of hardware. The structural integrity of the vehicle must support the spacecraft. This means that loads, such as bending moments, wind shear, dynamic pressure distribution over the surface of the vehicle — that is, dynamic, thermal, and static loads — must be taken

46

O . H . Lange, D e v e l o p m e n t of U . S. space vehicles 12,000 ami-,

Jahrbuch 1961 der W G L

i Moon -

30 Burnt

Í0 ft/sec velocity

50-103

Fig. 8. C u t o f f velocity requirements increased with extended missions requirements f r o m ballistic missiles through space vehicles.

into consideration. Consequently, optimizing thrust and determining its theoretical distribution over the desired flight path is not enough in determining the number of stages or their size. When the reality of a carrier vehicle enters these considerations, a less than optimum condition generally results. However, Tsien and Evans did point the way by indicating the relationships of mass ratio and acceleration factors in relation to engine thrust at altitude. Generally, the engine thrust should increase with altitude. H e r e it should be noted that, in practice, large vehicles should be boosted at fairly low velocities in the heavy atmosphere in order to avoid aerodynamic heating and loading problems. Upon reaching the less dense atmosphere, extremely high speeds can be reached with little complication due t o the reduction in atmospheric density and gravity. For large rocket vehicles, the acceleration (total thrust divided by the total weight) ranges in value f r o m 1.0 to 2.0. The near optimum value of initial acceleration is approximately 1.7. Staging experience was gained not only f r o m the Bumper vehicle, but f r o m surface-to-air guided missiles. These missiles were boosted to near flight velocity and the booster then separated. Separation was comparatively simple, since it occurred in the dense atmosphere and aerodynamic drag pulled the used booster away. Sounding rockets, the first generation ballistic missile, were also staged. Again the problem was not as severe since stage separations were at comparatively low altitudes, where drag and gravity could render a considerable assistance. These two examples may be classed as passive separation techniques. For large carrier vehicles, the problem is more difficult, since stage separation usually occurs above maximum "q", in the thin upper atmosphere where drag is ineffective. The designer's only recourse is the use of positive separating devices. The different separation techniques, themselves, pose problems of varying degrees of difficulty, which arise primarily f r o m low initial accelerations, precise guidance and control requirements, and small margins of safety. The positive separating techniques (Fig. 9) generally used are small retro-rockets arranged to fire in opposition to the line of flight, and ignition of the second stage engine prior to separation. The retro-rockets add an incremental velocity to the upper stage, thus permitting the f o r w a r d portion of the vehicle to fly off under full power f r o m the lower, already spent section. This full power take-off at separation is commonly referred to as "fire-in-the-hole". Selection of one of these techniques is usually contingent upon such related factors as trajectory, control system,

Fig. 9. T w o types of positive separation techniques used in separating multi-step rockets.

stage stability, weight and propulsion systems characteristics. Special variations of positive separation, such as used by the Atlas vehicle, is termed "stage and half". In this case, common propellant tanks store propellant for both first and second stage flight. Since all engines are ignited at launch, separation sequenceing involves cut-off of the first-stage engine system and discarding this over board. The "fire-in-the-hole" technique of stage separation requires very careful design of the interstage adaptor. The hot gases must escape f r o m the interstage area in all directions simultaneously, thus eliminating misalignment problems to prevent formation of destructive over-pressure. Of equal importance is the need to prevent flame f r o m recoiling back into the upper stage's a f t structure. Here are located vital wiring, hydraulic lines, actuators, and many other items of h a r d w a r e that must be protected f r o m heat. As might be expected, tank dome configuration plays an important p a r t in pressure distribution. The two basic types of dome design used are conical and elliptical. Of these two types, the conical head shows less induced bearing pressure than does the elliptical head. If a deflector is added to the dome, the pressure distribution over its surface appears to be greater. Thus, the problem of practicable hardware and available space are factors of considerable importance and could very well rule their use out as a practical design solution. The use of auxiliary jets on the f o r w a r d stage and retrorockets on the a f t stage presents a second very important method of positive separation. In this case, the problems encountered using the "fire-in-the-hole" method are circumvented. In addition, the resultant relative positions of the stages are more accurately defineable — the attitude of the upper stage is known prior to ignition of its engines. Moreover, the auxiliary jets furnish power f o r propellant seating and roll control. 1.4. Propellant Sloshing The sloshing of liquid propellants in their tanks is an important source of oscillatory loads. T h e effect on structural loads is not great, but, by coupling back through the control system, the effects can be very detrimental to the

Jahrbuch 1961 der W G L

O . H . Lange, Development of U . S. space vehicles

vehicle. This is one of the major problems in missile design — the crosscoupling effects between the control systems and the structure of the vehicle. The problem lies primarily in the area of control, but the strong interaction with the structural loads must be carefully treated and sympathetic frequencies avoided.

Table 1. Injection errors for S a t u r n - m a n n e d flight systems. Three sigma errors at injection.

1.5. Guidance and Control For most large vehicle designs, the control system receives programmed commands from the guidance system (Fig. 10), based on acceleration time, and altitude for a particular trajectory. In this manner, the altitude trajectory points are determined and fed into an autopilot; however, due to bending of the vehicle, the true altitude is not measured by the autopilot's gyroscope. In consequence,

System

47

Velocity [m/sec]

ST-90 p l a t f o r m

2.25

Four-gimbal p l a t f o r m

0.95

which have been conservatively projected to the time and conditions applying to the Saturn manned flight. The data given in Table 1 include contributions from gyro drift, accelerometer bias and scale factor errors and initial platform leveling errors.

Staging, timing ignition

Fig. 10. Schematic of the control portion of the Saturn launch vehicle guidance and control system.

the control system may make false corrections which lead to vehicle instability. The problem then resolves itself into locating a position for the gyroscope and circuit design for the control system which will make the vehicle system stable. Usually, the gyroscope is located at the point where the first bending frequency crosses the neutral axis of the vehicle. The problem becomes even more complex for multistage designs requiring controls for each stage during its period of operation. The problem is far from simple even for a single-stage vehicle, with weight and bending characteristics continually changing, and many alterations are made as a result of flight test program data, until an operational vehicle is completed. For manned flight, schematic errors will be held low enough to be negligible. However, sufficient work has been done in the area of component errors to make an engineering approximation, based on current experimental data

1.6. Stabilized Platform The stabilized platform for the Saturn vehicle is the ST-90, previously used in the Juno II vehicles. The ST-90 platform is a three-axis, inner-gimbal design, which has a carrier ring for three air-bearing integrating gyro-accelerometers. At second-stage burnot, the roll system is switched from propane to the residual helium remaining in the fuel tank by actuation of a solenoid-operated three-way valve. The one-roll amplifier actuates the first-stage, secondstage powered flight, and second-stage coasting flight roll systems. The operation of the roll jet amplifier is similar to that previously described for the pitch/yaw amplifier because the system deadspot in roll is ± 3 degrees. Again, a high lead a-c network (a-c = 40) is used to provide the least amount of jet-on time («sO.l second) and the maximum jet-off time (2.2 seconds) commensurate with adequate system stability. The roll amplifier output plate relay also

48

O . H . L a n g e , D e v e l o p m e n t o f U . S. s p a c e v e h i c l e s

Jahrbuch 1961 der W G L

operates a slave relay. One set of slave relay contacts actuates the second-stage system after an arming signal is received at second-stage ignition.

The S-IV stage is powered by six Pratt and Whitney RL 10 engines arranged in a star pattern. Control forces are obtained from all six gimbaled engines. Switching the flight control computer command signals to the S-IV gimbaled engines occurs during S-I/S-IV staging.

The associated electronics are housed in a separate package. To allow a large amount of freedom in the roll axis, the platform is mounted on a ring which provides this freedom in the early part of the flight, during which the vehicle is roll-programmed to the proper firing azimuth. Since the platform does not require shock mounting, it can readily provide accurate angular displacement signals for the attitude control system. Three a-c gimbal pick-off devices (resolvers or microsyns) provide signals for the pitch, yaw, and roll axis. The Saturn platform was selected for the following reasons: Proven reliability through a large number of ballistic missiles and space flight, N o development or tooling costs necessary, Accuracy, weight, and size is acceptable for Saturn firings. Since some of the later mission-oriented Saturn flights will require larger gimbal freedoms in all axes, a fourgimbal platform will be developed. Operation times (assuming two parallel arithmetic elements): Multiply Add/subtract Store, transfer, etc.

Guidance

S i g n a l

Basic

F l i g h t

C o n t r o l

Mechanization

A simplified block diagram of the basic vehicle flight control pitch system is shown in Fig. 11. For simplicity, no distinction is made between the control torques achieved Servo loop^v amp box

Angle of attack meters

I

p

Y

~

\

Accelerometers I I T Bui dance repeater

T r r

[flaiel [ffi^ii

Velocity Imeas. . I

signals only!

Control computer

V Electrical -differentia ti on net works

Processor

A primary consideration in the Saturn guidance system is that there should be only one system, adaptable to many missions. The guidance computer is to be the most general purpose machine practical, limited by availability for the early flights. All functions peculiar to small groups of vehicles will be handled in the signal processor, and the computer will have no changes between flights. This philosophy will apply to the first guidance flights and also to those using the advanced platform and computer. In addition to the adaptability feature, the processor will include those types of equipment which do not lend themselves to packaging in a high-density electronic box, e. g., servo systems, power amplifiers, relays, and miscellaneous equipment. The box, itself, may be redesigned for groups of missions, but many of the modules will not require such redesign. The primary purpose of the flight control system is to stabilize the vehicle during the first stage powered flight, separation, and the second stage powered flight. The S-I stage is powered by eight North American Aviation H - l engines. Control forces are obtained from the lateral thrust component achieved when the four outboard engines are gimbaled in response to various control and steering commands from the flight control computer. In addition, the S-I stage includes fins to achieve additional aerodynamic stability during powered flight.

Sain 'servo Tilt

St-90

1>Gyros

Shaping

Magnetic

networks

amplifier

Hydraulic • actuator ¡/aiv'e/

2.0 msec, 0.2 msec, 0.2 msec.

The Saturn computer represents a considerable advancement in computer technology. It will be a general-purpose machine with greater memory capacity, and higher computing speed than the one used for the early flights. The greatest emphasis will be placed on reliability. The approaches considered most promising for improvement of computer reliability are minimization of electrical connections through thin film and solid state circuitry, triple modular redundancy, and solid state memory. 1.7.

1.8.

settingsloops

program

—Program r ~ device /t/r

bearings

Flight sequencer

f

1 Sequencing signals I to missile system \Engine-out •signal

Feedback poten tiometer F i g . 11.

A block d i a g r a m of the c o n t r o l system f o r early singlestage S a t u r n launch vehicles.

from the gimbaled engines and the S-I/S-IV interstage air vanes. The yaw channel is similar to the pitch channel. Roll is mixed on both the yaw and pitch channels as shown in Fig. 12. To take care of the two additional control engines on S-IV, pitch and yaw signals are mixed. No roll signals are used for these engines. The principle inputs to the flight control system are attitude angular position (0), attitude angular rate (&), and lateral acceleration (v). The attitude position signals provide an angular reference for the vehicle flight control system. The pitch attitude signal may be altered by a programmed command to tilt the vehicle over a desired trajectory. Dynamic stability of the vehicle is obtained by the attitude rate signals to provide the necessary damping of the rigid body attitude mode. Lateral forces acting on the vehicle due to angle-ofattack caused by winds are sensed by the lateral control accelerometers. These signals are used to bias the vehicle into the winds to reduce structural loading, angle-of-attack, and control deflections. The ratio of the attitude position gain (a0) to lateral acceleration gain (g2) determines the degree that the vehicle turns into the wind. The a0-g2 gains are continuously programmed by the flight control computer. 1.9.

A t t i t u d e

Angular

Rate

During the S-I stage powered flight, attitude rate information is obtained from a three-axis rate gyro package. Rate gyros earmarked for Saturn have proven to be highly reliable in application on other vehicle projects. The rate gyros include an electromagnetic torquer and a spin motor sensor to provide a means of remotely checking the rate gyros after installation on the vehicle to assure proper operation before flight.

Jahrbuch 1961 derWGL

O. H . Lange, Development of U. S. space vehicles

Max £i=±BMA

[ f H I Yawliwo

BNB transducers} ÉFW fö) !«tel. meas.

A

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±30 V.D.C.

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from pot)'

!ß tel. from meas. pot.I

3y]--j--T,—r--\nm\—I 1 1 Filter |wJ,flf«| ! V-7VJ I .,, Uitn |-[ rFilter ~T sm \DEM\—I Fitter

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:

BND DEM

BOV DC

AX

Sub

Mark ite

pot.5k max

locai

a - p f e s ® " ® - , f ^ - t i t r ^ P ^ H Filtsr '

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z=AMmI—I Filter |-

49

JCant bias L Four times_ ' reg'd -Control computer -

ÌTit

sm St-90and seri/o amp. box

Actuator ~package~

Fig. 12. Circuit diagram of an early Saturn control system.

After S-I/S-IV staging, attitude rate signals are obtained by electrically differentiating the attitude position signals. This is indicated within the flight control computer by the lead networks in the attitude position channels. Rate gyros are used during the S-I stage flight due to the extremely low elastic body bending frequencies encountered. During the S-IV stage flight, the relative high separation between the elastic body frequencies and the rigid body control frequency allows satisfactory rate information to be achieved by differentiating the attitude position. These higher body frequencies can easily be filtered out of the attitude position loop and the location of the stabilized platform is not critical. Attitude position differentiation leads to a more reliable flight control system rather than providing for an additional rate gyro package and associated electronics located in the S-IV stage. 1.10. Lateral Acceleration (Control Accelerometers ) Lateral acceleration information during the S-I stage powered flight is obtained from the pitch and yaw body fixed control accelerometers located at the top of the S-I stage booster. The Statham A-311 type used has proven to be successful for artificial stabilization (angle-of-attack control). The accelerometer is a linear type, fluid damped, with an inductive pick-off. Similar to the rate gyro package, the accelerometers also have a remote checkout capability. By exciting a separate electromagnetic force coil within the accelerometer, it is possible to deflect the seismic mass a calibrated amount and check the electrical output of the unit without subjecting it to an actual acceleration. This provides a means of remotely checking the accelerometers after installation on the vehicle to assure proper operation before flight. 1.11. Attitude

Angular

Position

The attitude angular position signals are provided by the stabilized platform. These signals are a-c, 400 cps modulated and are routed directly to the flight control system. The signals are amplified and demodulated in the flight 4

control computer. In addition, isolation amplifiers are included to provide signal outputs for telemetering, ground support equipment, and pilot displays if required for manned payloads. As shown in Fig. 13, the output signals from the rate gyro package and the control accelerometers are routed directly to the control signal processor. The signals are amplified and demodulated and routed to the flight control computer for additional shaping and gain programming. The combination of the control signal processor, the rate gyro package, and the control accelerometers comprises a subsystem of the flight control system, being required only during the S-I stage powered flight. The control signal processor, being the common unit of the subsystem, provides the following: 1. Electrical power requirements for the rate gyros and accelerometers; 2. Isolation demodulators to provide signal outputs for telemetering ground support equipment, and pilot displays, as required; 3. Ground support equipment commands for rate gyro substitute, and torque signals and spin motor speed sensor output signals for ground support equipment monitoring; 4. Ground support equipment commands for accelerometer substitute torque signals. The construction of the control signal processor is identical to the flight control computer. 1.12. Flight Control Computer The flight control computer provides for the following operations, as required by the flight control system: 1. Attitude amplifier and demodulators; 2. Filter networks for gain and phase shaping of the xp, yj, and y input signals; 3. Lead networks for yj differentiation after S-I/S-IV staging; 4. Continuous a0-g2 gain program; 5. Step gain programming of the yj and roll signals; 6. 12 summing and servo-loop amplifiers for the S-I/S-IV gimbal engines.

50

O. H . Lange, Development of U. S. space vehicles

Jahrbuch 1961 der WGL

r * 0 Digital guidance computer

Time pulses for step program ' changes, staging computer output from $-1 to S-IV, etc.

Xr

Guidance signal processor

\ Shape

Amp Stabilized platform attitude position error

nets

i 4»

Typical pitchjaw and roll control commands to S~Ior S-IVswivel eng. lNo.l23andil,fourlil required

ag Gain program

•MLead nei H

^

fy

Amp

Shape neis

Amp

Shape nets

l>

aD Gam 4*. program r|lead

4 * }

-

arffi

program

i

Command signals routed to corresponding summing and servo loop amplifiers Shape nets

a, Step program

Shape nets

a, Step program

Shape nets

a, Step program Typical pitch and yaw control commands to S-IV swivel eng. fNo. 5 and 8)t two 12) required

Control accelerometer body fixed

Shape

a? Gain

nets

program

Shape nets

a? Gam program

/ E >

Control Flight control

signal processor ^

Signals from isolation amp to

telemetering and gase

locationJstage

computer

S-I only

Block diagram of flight control

4

Swivel engines fß actuator

location g andc

compartment

above stage

S-IV

system

systemsI

L Fixed engines 4 Fixed

^

.6 Swivel engines{6 actuator systems}

fins

Stage S-I arrangement

Stage S-IV arrangement

Fig. 13. Schematic diagram of an early Saturn flight control system. A l l of the v a r i o u s subassemblies r e q u i r e d f o r the a b o v e o p e r a t i o n s u t i l i z e p l u g - i n m o d u l e e m p l o y i n g p r i n t e d circuit techniques f o r e a s e of m a n u f a c t u r e a n d assembly. This t y p e of construction is essential in p e r f o r m i n g f a s t changes to the s h a p i n g a n d l e a d n e t w o r k s a n d gain m o d u l e s to a c c o m m o d a t e the v a r i o u s c o n f i g u r a t i o n s a n d missions that m a y be used f o r the basic S a t u r n S - I / S - I V vehicle.

T h e s u m m i n g a n d s e r v o - l o o p a m p l i f i e r consists of a m a g n e t i c p r e - a m p l i f i e r f o r s u m m i n g , which p r o v i d e s f o r c o m p l e t e electrical isolation between all channels. This is f o l l o w e d b y a t r a n s i s t o r i z e d p o w e r a m p l i f i e r to drive the electro-mechanical a c t u a t o r v a l v e of the g i m b a l engines. This basic s u m m i n g a n d s e r v o - l o o p a m p l i f i e r h a v e a history of success in several other vehicle d e v e l o p m e n t p r o g r a m s .

Jahrbuch 1961 der WGL

O. H . Lange, Development of U. S. space vehicles

The flight control computer consists of three basic sections: The left section is the amplifier demodulators and the servo-loop amplifiers; the center section is the shaping networks; and the right section is the lead networks and gain programmers.

operation. When the gas is released in the propellant tank's ullage area (above the liquid), the gas pressure forces the propellants through the plumbing system to the injector and into the combustion chamber. It is apparent that the propellant tanks must be able to withstand the pressure of the gas. In addition, the pressurizing gas must not be soluble in the liquid propellants, and, as mentioned, for long engine burning time, a large quantity of gas is required, hence, large reservoirs. It is obvious that, at some specific performance requirement level, this system becomes unduly bulky and heavy. When graphically compared with the pump-fed system, the weights and thrust curves "cross over" at the point of diminishing returns. For the large rocket carrier vehicle's operating regime, the pump-fed system is unquestionably the only practicable system to use.

1.13. Hydraulic Actuator System The S-I and S-IV stage of the Saturn vehicle contain four and six, independent, closed loop hydraulic servo systems designed to eliminate the need for an external hydraulic source. Each system consists of two electricallycontrolled, hydraulically-operated actuators, a main pump and an auxiliary pump. Each gimbaled engine is positioned by the two actuators in response to electrical command signals from the flight control computer. The actuators are a linear, equal-area, double-acting piston type. The main hydraulic pump provides hydraulic fluid to operate the actuators during flight. The pump is operated by the engine turbopump. The electrically operated pump is used to supply hydraulic fluid during "cold" gimbaling and checkout operations before flight. 2. R O C K E T PROPULSION The great historic division in rocket propulsion systems lies between the use of liquid and solid propellants. In recent years, however, investigations began of many hybrid types of propulsion. The more unusual of these propulsion systems operate on energy derived from nuclear and electrical power sources — for example, nuclear, thermonuclear, ion, plasma, and photon engines. At the present time, such engines (Fig. 14) possess low-acceleration capabilities, and some engine types could present problems in atmospheric contamination. In consequence, the application of such engines as boosters for large launch vehicles is still in the future. Therefore, in the following paragraphs, only the liquid and solid types of rocket propulsion will be discussed. ?

1

0.B

5 OA

y

Scrub on 18 til SOS

17 th

20 th

19th

18 th

"Si •s1 §s eg ^

ST3

/ lime

-J. §

Os.

Fig. 46. P l o t of launch w i n d o w f o r a typical space probe launched by the T h o r vehicle indicating probabilities of success for specific launch times.

The curve of Fig. 46 shows the composite "probability of success" versus time. As long as the countdown continues on schedule, the solid line prevails. If, on a given day, the flight is held, the dashed curve is used. For an August launch, the calculation of optimum launch cutoff times (after this time it is better to wait until the following day) gave the results in Table 4. Table 4. Day 17 18 19 20

August August August August

N o m i n a l launch time (T — 0) 0714 E S T 0818 E S T 0919 E S T 1017 E S T

C u t o f f time 0734 E S T 0842 E S T 0945 E S T —

N o t e : The actual allowable launch interval f o r each d a y was in excess of 20 minutes.

8. R E C O V E R Y Historically, rocket vehicles have been "single-shot" carriers. All the development, fabrication, and test effort expended in producing the vehicles were "lost" as soon as the launch operation was performed. Thus, if the booster could be recovered and reconditioned f o r f u r t h e r flight operations, the advantages would be appreciable. In addition to the advantages of reuse, study of the recovered flight instrumentation would indicate the validity of the telemetered flight data. As recovery is related to the increasing accentuation of economy on vehicle design, several methods of recovering booster stages have been studied. Some of these include use of parachutes, use of wings, and the newer concept of adapting a Rogallo kite (paraglider) (Fig. 47). All recovery schemes proposed (except the paraglider) necessitate either water landings or significant redesign of the stage structure to withstand landing impact. Since the paraglider's flight paths can be controlled through the vehicle guidance system, ground landings are feasible. Lines connecting the glider to the booster can be pulled in or let out, adjusting the glide path as the landing strip is approached. (For launches f r o m C a p e Canaveral, the two possible landing sites are the Grand Bahama Islands or back at the Cape itself.) Additions of vehicle guidance, landing skids, and packaging provisions of the glider assembly are relatively minor redesign problems. 9. B E N E F I T S A R I S I N G F R O M L A U N C H V E H I C L E DEVELOPMENT The benefits derived f r o m the development of launch vehicles for space exploration fall into two general areas: first the longrange benefits accruing from information concerning the space environment and the extra-earth bodies therein; second, the skills, materials, processes, and advances secured during vehicle development.

Jahrbuch 1961 der W G L

O. H . Lange, D e v e l o p m e n t of U. S. space vehicles

O n e of the most interesting aspects of large launch vehicle developments is the extent to which the national industrial complex is involved, drawing together the fields of electronics, metals, fuels, ceramics, machinery, instruments, plastics, textiles, cryogenics, and many others. Consequently, the benefits of launch vehicle development has been felt directly or indirectly in virtually every area of the national economy. The advanced exploration programs will place even greater demands f o r technological advancement upon the American industrial complex. These new advances will create even more benefits in new products and jobs. It is estimated that at least 5,000 companies or research organizations are a p a r t of the missile-space industry. They have contributed the 3,200 different products thus f a r required. Future programs will probably require many more items designed to study and overcome problems whose existence is unknown today and whose commercial application will surely follow.

electronic brain, currently used, will soon be replaced by a unit only a fraction as large. T h e application of such a small unit to general use is virtually limitless. Even at the present time, computers are used in the conduct of scientific investigations and business operations, providing extreme speed, efficiency and accuracy. Simple but time consuming operations, such as stock control, accounting, and records storage, can now be handled by computer operations, markedly reducing overhead expenses. In the fabrication shops, computers have been adopted to lathe, millingmachine, and drill-press operations, turning out work of extreme accuracy. Such an arrangement will permit hitherto-undreamed increases in production.

Fig. 47. A suggested method for recovery of the Saturn booster is shown. Recovery and reuse of vehicle h a r d w a r e o f f e r s a possibility of reduction in cost of expensive h a r d w a r e used in rocket vehicles.

The area of electronics and electrical equipment is perhaps the most outstanding example of the immediate benefits gained f r o m vehicle development programs. Micro-miniaturization, molecular electronics, and micromodular engineering are all terms indicating the construction of extremely small electrical equipment. The electrical power to operate the equipment aboard a spacecraft has great potentials in the commercial market. O n e development is the fuel cell, which converts fuel directly into electrical power. In England, for example, a 40-cell unit has been used to drive a forklift truck and to operate an electric welding machine. The cell develops up to 5 kilowatts. In the United States, a 30-cell portable powerplant, producing 200 watts, and a 1,000-cell unit, with a 15-kilowatt output, have been developed. The 1,000-cell unit has been used to drive a tractor. Plasma jets, investigated as a source of rocket thrust, can be used as a conductor of electricity and, when used as a "magnetohydro-dynamics" generator, have a variety of applications. Some day, it may prove an economical power source. Currently ten municipal areas are considering experiments with electrical power f r o m this source. It has been estimated that, "as much as one-million watts could be generated by shooting a stream of plasma at Mach number 3 through a magnetic field three feet long with the magnetic poles six inches a p a r t " . The field of high-speed digital computers has been developed as a direct result of requirements f r o m research in both nuclear physics and rocket development. The trend in size, weight, and volume reduction is marked. The giant

73

The ceramics and glass industries have also benefited f r o m the discoveries made during rocket vehicle development. Pyroceram, developed f o r radomes, is currently used to make pots and pans. Light-weight, unbreakable glass is used f o r food containers and f o r fabrics. The plastics industry has progressed even f u r t h e r in developing new materials for rocket vehicle application. Among the many plastics developed is polyethylene, used for packaging everything f r o m food stuffs to chemicals. Polyurethane, originally developed by I. G. Farben as a substitute for nylon, has been adapted to solid propellants, automobile tires, fabrics, and seat cushions. Civil engineering also profits f r o m advances made in development of highly accurate homing guidance. These principles, applied to surveying techniques, promise to revolutionize the surveying industry. The plasma jet has also been adapted to the fabrication of ultrahard metal. Developing heats of 30,000 °F, the torch can work within tolerances of two-thousandths of an inch. Ultra-precise f l o w meters (having no packing or bearings) f o r measuring f l o w of cryogenic propellants can readily find application in the chemical industry. The field of mining, long limited to mechanical or hydraulic cutting away the earth to reach ore bodies, has adopted the jet of the rocket engine. Jet drilling opens the door to inaccessible deep-lying oil deposits or ore bodies. T h e famous Mesabi iron ore range is using jet mining to uncover the bodies of taconite ore. In quarrying stone, jet spalling and channeling ore are proven techniques, reducing the waste characteristic of this industry. Rocket flame permits cutting along natural cleavage planes or crystal boundaries — hence, stone is cut thin without cracking. In addition, jet cutting produces a fine finish that is unobtainable when cutting is performed with steel or abrasive tools. Film-resistance thermometers or tape-recorded electrocardiograms are a few of the tools made available for medical research. Miniaturized electronic equipment can be used to telemeter the critical physiological data of hospitalized patients to a centralized control panel, thus eliminating observation by a nurse or disturbance of the patient, while maintaining unimpeachable records of the patient's condition. Medical research may also be aided by biological findings f r o m the Mercury and Apollo programs, including studies of the reuse and recycling of water (including sea water), to make it suitable f o r human consumption. These studies may lead to practical industrial methods of removing valuable elements f r o m sea water. The effect of the absence of familiar noises on the human being has been extensively studied during the Mercury and Apollo programs. In other studies, high-frequency noise generated and reflected by a parabolic reflector has produced startling results, f r o m destroying masonry walls to starting fires.

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O. H . Lange, D e v e l o p m e n t of U . S. space vehicles

Jahrbuch 1961 der W G L

The vehicle development programs have resulted in a number of new metallurgic processes and materials. Titanium for example, long considered an exotic material that could be processed only under heat, can now be "cold" worked without destruction of the metal's physical properties. Welding of aluminum, long considered difficult if not impracticable, is now done continously by machines especially designed f o r vehicle fabrication.

important applications throughout the national economy. As our investigations continue, as our research probes ever more deeply into the environments around us, we receive a constant stream of unexpected by-products f r o m our development programs. At the same time, our horizons are expanding inexorably outward. With increasing payloads and increasing vehicle reliability, we are progressing toward man in space, man on the moon, and, possibly, man on the planets.

It is possible to list hundreds of tangible items that have been developed as a direct result of space work. The intangible benefits f r o m vehicle development programs are of equal or greater interest, f o r here the imagination finds unlimited horizons: mining operations on the moon; colonies of men in space stations designed as way stops t o w a r d other planets. But other advantages in the immediate f u t u r e are far less visionary. Weather is now being studied by such meteorological earth satellites as Tiros. With study, comes knowledge; and through k n o w ledge we may develop ways and means to control or influence the weather. The field of navigation will be greatly aided by use of the Transit satellite. Satellite communication (illustrated by Echo and the f u t u r e "24-hour" satellite) will permit development of long-distance earth-leased communications f o r both telephone and television. Telestar represents the first generation of communication advantages gained by using satellites. More advanced concepts will permit continuous T V viewing instead of the current restricted viewing as possible with Telestar. Pioneer V, for example, proved the feasibility of extended, long-distance communication by successfully sending "readable" signals across a distance of 23-million miles.

Why should man go to these places? O u r best answer may be, after all, quite simple. We are going o u t w a r d to these places f o r the good and sufficient reason that they are there. Übersicht Das R e f e r a t gibt einen Überblick über den gegenwärtigen Stand der Saturn-Raketen-Entwicklung. Die hauptsächlichen Raumforschungsprogramme werden nach Familien von Trägerraketen eingeteilt. A m A n f a n g stehen die f ü r den Verwendungszweck veränderten ballistischen Geschosse unterschiedlicher T r a g k r a f t mit hinzugefügten höheren Stufen, am E n d e stehen die gegenwärtig in Entwicklung befindlichen, f ü r bemannten Flug gedachten Raketen der S a t u r n - N o v a - K l a s s e . Die technologischen Fortschritte in wissenschaftlicher und technisch-industrieller Hinsicht werden auf den folgenden Gebieten beschrieben: Zelle, Steuerung, G r u n d z ü g e im E n t w u r f mehrstufiger Systeme, flüssige und Feststoff-Antriebe, F a b r i k a t i o n und Zusammenbau, Abschuß sowie allgemeine O p e r a t i o n e n . D e r Fortschritt in der Entwicklung von F l u g k ö r p e r n f ü r instrumentierte und bemannte Flüge ist auch auf wirtschaftlichem und allgemein technologischem Gebiet sowie f ü r die Ausdehnung von Weltnachrichten- u n d Wetterdienst von N u t z e n . Summary

The demand f o r trained personnel to use the data gathered f r o m space exploration and to design and build the space vehicles of the f u t u r e should result in a massive stimulation of the national educational system. Every field of human knowledge will expand, as new facts, new theory, and new techniques are developed. In consequence, technical schools and universities will be given new impetus to maintain dynamic and constantly up-rated curriculums in which the latest findings and techniques are presented.

A brief i n t r o d u c t o r y status report of the first Saturn booster is given. In terms of m a j o r exploration programs, families of space vehicles are classified, beginning at the lower end with modified ballistic missiles of varying p o w e r and a d d e d u p p e r stages and leading to the m a n - r a t e d vehicles of the S a t u r n - N o v a series. The technological advances on a b r o a d scientific-technical-industrial f r o n t are shown in the areas of structures, guidance and control, design features of composite vehicles, associated liquid and solid propulsion systems, fabrication, launching as well as handling.

The present trends in launch vehicle development have crystallized in response to the needs of an ever-expanding space research program. As payload-carrying capabilities constantly increase, vehicles grow constantly more reliable. As we improve the probability of mission success, we simultaneously reduce the cost per pound of payload. In illustration, payload cost f o r the Vanguard was about one billion dollars per pound. For f u t u r e payloads, the cost will be on the order of one thousand dollars per pound. When payload costs are reduced to about one hundred dollars per pound, wide-scale commercial use of space can be expected. One step t o w a r d payload cost reduction is the application of solid-propellant motors to launch vehicles. Still another step is the possibility of recovery and reuse of liquid-propelled boosters. N e w fabrication techniques also o f f e r avenues to substantial cost reductions.

A d d i t i o n a l benefits arising f r o m the progress in t h e development of vehicles f o r instrumented exploration and manned flight are economical, technological, of a general nature, an expansion of the global communications and weather services.

Cost reduction is thus a major element in our space vehicle development programs. Yet we should not allow f u t u r e possibilities of reduced cost to obscure the fact that the large sums already expended in our space programs have resulted in equally large rewards — in useful and

Résumé Le r a p p o r t donne u n aperçu sur l'état actuel d u p r o j e t S a t u r a . Les principaux programmes d ' e x p l o r a t i o n de l'espace sont classifies d'après les familles des véhicules spatiaux, allant des engins balistiques modifiés de puissance d i f f é r e n t e avec des étages supérieurs ajoutés jusqu'aux fusées de la série S a t u r n - N o v a actuellement en développement et destinées aux vols spatiaux occupés. Les progrès technologiques en ce qui concerne la science, la technique et l'industrie sont décrits dans les domaines suivants : structure, commande, caractéristiques de construction de systèmes à plusieurs étages, propulsion à c o m b u r a n t liquide et solide, fabrication et assemblage, lancement ainsi que des opérations générales. Le progrès dans le développement d'engins balistiques p o u r des vols aux instruments ainsi que p o u r des vols occupés est d'un grand bénéfice dans le domaine économique ainsi que dans celui de la technique des communications globales et des services météorologiques.

D. J. L Y O N S , F A R N B O R O U G H ,

HANTS/ENGLAND*)

D K 521.6 521.3 629.19

SATELLITE L A U N C H I N G

POSSIBILITIES

1. I N T R O D U C T I O N This paper is a discussion of the practical aspects of launching satellites into terrestrial and circum-lunar orbits. It comprises two parts: the first dealing with the problems and possibilities in a general form and the second, showing what launching capability will be available for European use assuming that the joint European/Commonwealth plan to develop the Blue Streak satellite launching vehicle goes ahead. 2. O R B I T S A N D P A Y L O A D S 2.1. General Limitations on Orbits Possible With any given satellite launching system, that is with given stages, a variety of orbital versions can be accomplished. This can be done in two main ways. a) The p a y l o a d can be altered up or down in weight. This will cause a corresponding decrease or increase in the final energy/lb of p a y l o a d plus spent last rocket stage and the possible orbits will be changed.The take-off acceleration of the rocket will be a little changed by the increase or decrease in weight, but with more or less conventional rocket launchers this will not b e appreciable. There may, however, be limitations to this process arising from stressing considerations of the upper stages, and the optimization of weight distribution through the stages might be upset. b) T h e p a y l o a d can be altered up or down in weight accompanied by an equal decrease or increase in weight of the rocket stages, usually in fuel loading. The fuel weight alteration can be distributed through the stages in an attempt to i) keep the first stage flight path constant in order to keep aerodynamic heading and stressing constant and within design limitations during exit from the earth's atmosphere and ii) to give reasonably optimum weight sharing between the stages. Of the two methods a) and b) the second is the one that will be more usually applied. N o w there are some obvious limitations that will apply to the process of changing orbit; when the p a y l o a d is reduced to a very small amount a particular launching vehicle will have reached its maximum capability in terms of energy of the orbit, that is in the total energy/lb of p a y l o a d (potential plus kinetic). This possible maximum is quite different for different types of launching vehicle. There are, however, a number of less obvious restrictions. The first of these is the limitation on maximum orbit energy due to restrictions in direction of launch. This arises from the fact that at the moment of launch the rocket possesses a space velocity relative to the centre of the earth equal to the earth's surface velocity (or an aeroplane Royal Aircraft Establishment.

launcher's velocity relative to the earth's centre) at the point of launch. Since the orbital characteristics are measured in space axes, the fixed (for one configuration) additional velocity of the rocket launcher has to be added vectorially to the initial velocity; so greater energy orbits can be obtained when firing eastwards and taking advantage of earth rotation than when firing north or south or firing westwards. Secondly, there will be, in general, problems of range safety that will limit possible direction of fire, and possibly effect limitation to the optimization between the various stage weights. It is quite clear that the reliability of rockets is not 100 %>, so that steps have to be taken to prevent d a m a g e to people and property in the event of a failure of any one of the stages to perform correctly. A part of the precautions taken is the provision of a good radar tracking system to follow the vehicle as it accelerates, and to watch for any large deviation from the expected path, so that a command signal can be sent to the missile either to destroy the missile or stop its engines so that no further deviation from the ballistic path can take place. So, for any given line of fire a range safety funnel must exist on the ground within which vehicles may land which have misbehaved. This funnel ideally must be clear of persons and property or sufficiently sparsely occupied so that the estimated chance of damage is acceptable to the nation which owns the property. If the funnel or part of it is over the sea, the sea area should ideally be clear of ships. Further as each stage burns out and dispatches the next stage on its normal flight, the spent stages (certainly the earlier ones) will have insufficient velocity to orbit and they will fall back to the earth. T h e expected points of impact will have to be in planned impact areas, and the direction of launch from a given launching site will be limited so that acceptable impact areas can be provided. Some of the later stages may, of course, travel a very large distance (thousands of miles) before they impact and this could make the choice of impact area more difficult; this is offset, however, by the fact that much if not all of the spent stage will be reduced to metallic equivalent of matchwood under the combined action of the aerodynamic forces, and the heating effects during re-entry. These pieces or dust m a y have very low terminal impact velocities and may not incommode property or persons in the area of their fall. There are further limitations which may be imposed by the type of guidance system and propulsion system; but these will be dealt with in more detail in Section 3 where methods of injection are described. One consideration of orbital characteristics which may be very important in choosing a desirable orbit is the lifetime of a satellite in a particular orbit, or the rate of change of orbit shape, arising from the loss of energy due to atmospheric drag. The lifetime is both a function of the drag/weight ratio of the satellite vehicle in the orbit considered and also the shape of the starting orbit. Fig. 1 illustrates the very large variations which occur. As

D. J. L Y O N S , F A R N B O R O U G H ,

HANTS/ENGLAND*)

D K 521.6 521.3 629.19

SATELLITE L A U N C H I N G

POSSIBILITIES

1. I N T R O D U C T I O N This paper is a discussion of the practical aspects of launching satellites into terrestrial and circum-lunar orbits. It comprises two parts: the first dealing with the problems and possibilities in a general form and the second, showing what launching capability will be available for European use assuming that the joint European/Commonwealth plan to develop the Blue Streak satellite launching vehicle goes ahead. 2. O R B I T S A N D P A Y L O A D S 2.1. General Limitations on Orbits Possible With any given satellite launching system, that is with given stages, a variety of orbital versions can be accomplished. This can be done in two main ways. a) The p a y l o a d can be altered up or down in weight. This will cause a corresponding decrease or increase in the final energy/lb of p a y l o a d plus spent last rocket stage and the possible orbits will be changed.The take-off acceleration of the rocket will be a little changed by the increase or decrease in weight, but with more or less conventional rocket launchers this will not b e appreciable. There may, however, be limitations to this process arising from stressing considerations of the upper stages, and the optimization of weight distribution through the stages might be upset. b) T h e p a y l o a d can be altered up or down in weight accompanied by an equal decrease or increase in weight of the rocket stages, usually in fuel loading. The fuel weight alteration can be distributed through the stages in an attempt to i) keep the first stage flight path constant in order to keep aerodynamic heading and stressing constant and within design limitations during exit from the earth's atmosphere and ii) to give reasonably optimum weight sharing between the stages. Of the two methods a) and b) the second is the one that will be more usually applied. N o w there are some obvious limitations that will apply to the process of changing orbit; when the p a y l o a d is reduced to a very small amount a particular launching vehicle will have reached its maximum capability in terms of energy of the orbit, that is in the total energy/lb of p a y l o a d (potential plus kinetic). This possible maximum is quite different for different types of launching vehicle. There are, however, a number of less obvious restrictions. The first of these is the limitation on maximum orbit energy due to restrictions in direction of launch. This arises from the fact that at the moment of launch the rocket possesses a space velocity relative to the centre of the earth equal to the earth's surface velocity (or an aeroplane Royal Aircraft Establishment.

launcher's velocity relative to the earth's centre) at the point of launch. Since the orbital characteristics are measured in space axes, the fixed (for one configuration) additional velocity of the rocket launcher has to be added vectorially to the initial velocity; so greater energy orbits can be obtained when firing eastwards and taking advantage of earth rotation than when firing north or south or firing westwards. Secondly, there will be, in general, problems of range safety that will limit possible direction of fire, and possibly effect limitation to the optimization between the various stage weights. It is quite clear that the reliability of rockets is not 100 %>, so that steps have to be taken to prevent d a m a g e to people and property in the event of a failure of any one of the stages to perform correctly. A part of the precautions taken is the provision of a good radar tracking system to follow the vehicle as it accelerates, and to watch for any large deviation from the expected path, so that a command signal can be sent to the missile either to destroy the missile or stop its engines so that no further deviation from the ballistic path can take place. So, for any given line of fire a range safety funnel must exist on the ground within which vehicles may land which have misbehaved. This funnel ideally must be clear of persons and property or sufficiently sparsely occupied so that the estimated chance of damage is acceptable to the nation which owns the property. If the funnel or part of it is over the sea, the sea area should ideally be clear of ships. Further as each stage burns out and dispatches the next stage on its normal flight, the spent stages (certainly the earlier ones) will have insufficient velocity to orbit and they will fall back to the earth. T h e expected points of impact will have to be in planned impact areas, and the direction of launch from a given launching site will be limited so that acceptable impact areas can be provided. Some of the later stages may, of course, travel a very large distance (thousands of miles) before they impact and this could make the choice of impact area more difficult; this is offset, however, by the fact that much if not all of the spent stage will be reduced to metallic equivalent of matchwood under the combined action of the aerodynamic forces, and the heating effects during re-entry. These pieces or dust m a y have very low terminal impact velocities and may not incommode property or persons in the area of their fall. There are further limitations which may be imposed by the type of guidance system and propulsion system; but these will be dealt with in more detail in Section 3 where methods of injection are described. One consideration of orbital characteristics which may be very important in choosing a desirable orbit is the lifetime of a satellite in a particular orbit, or the rate of change of orbit shape, arising from the loss of energy due to atmospheric drag. The lifetime is both a function of the drag/weight ratio of the satellite vehicle in the orbit considered and also the shape of the starting orbit. Fig. 1 illustrates the very large variations which occur. As

76

D . J . L y o n s , Satellite launching possibilities

examples for a normal dense satellite, say an active radio satellite, a lifetime of three years would require a circular orbit of greater than 300 miles altitude or an apogee height up to 1000 or 2000 miles with a perigee height of 180 miles. A balloon type satellite would only last one day in a 300 miles circular orbit.

J a h r b u c h 1961 d e r W G L

city" requirements for Moon circular orbit missions starting from earth; included in these figures are the retro-rocket velocity requirements to stabilize the satellite in the Moon orbit. It is clearly demonstrated that the rocket performance required for Moon orbits is, at the least only slightly greater (3000 ft/sec) than highly elliptical orbits to the altitude of the Moon's orbit, but the low transit times may require the provision of a much better performance rocket. 10 ft/sec

Life Fig. 1. Satellite l i f e as a f u n c t i o n of initial a p o g e e a n d perigee. SCD/m = 0.1 f t 2 / l b (typical satellite), SCD/m = 100 f t 2 / l b (baloon). S = cross section a r e a , CD = d r a g c o e f f i c i e n t , m = mass.

Summarizing, therefore, a variety of orbits may be obtained with one basic launching vehicle with a limit on the maximum orbital energy. Depending on the launch site there may be limits on direction of orbit, e.g. polar orbits may not be possible from some sites and equatorial from others purely on the score of safety. 2.2. Pa y I o a d — Performance Exchange If a rocket combination of one or several stages is fired in vacuum and well away from any heavenly body (i. e. in a zero gravity field) the rocket will reach a given velocity relative to its starting condition; if the same rocket is fired still in vacuum but in a gravity field of any strength and along any path curved or straight an integrating accelerometer fixed in body axes on the longitudinal axis of the body will read the same velocity as that which would be reached in zero gravity field. This in rocket technology is known as the apparent velocity. The real velocity of the end stage of the rocket is of course less, when travelling away from the earth, say, as some of the thrust is being used to cancel out partly or wholly the gravitational attraction. This loss is called "gravity loss"; in addition the aerodynamic forces during the flight through the planetary atmosphere cause further losses of energy. This loss will cause a decrease also in the apparent velocity as measured by the missile fixed integrating accelerometer. These losses could vary considerably with the flight path and thrust over mass of the rocket combination and with the staging, but if optimum values are chosen for the parameters such as take-off acceleration, flight path made against time and so on, it appears that there is generally a fairly constant relationship between the apparent velocity required and the energy of the orbit chosen for a whole range of possible rocket combinations. Thus assuming a certain orbit is required, a first estimate can be made of the performance of the rocket required without recourse to detailed trajectory calculation. This approximation is particularly good when estimating the change in performance required between a rocket launcher giving one orbit to the same launcher modified to give another orbit. Fig. 2 shows the variation of "apparent velocity" capability required from the launcher for a variety of earth equatorial orbits both circular and elliptic (1500 ft/sec would have to be added for polar orbits to all these figures). Fig. 3 is a companion picture to Fig. 2 and shows the "apparent velo-

10"

2

5

height or circular height Fig. 2. Booster a p p a r e n t v e l o c i t y r e q u i r e m e n t s f o r e a r t h o r b i t missions ( e q u a t o r i a l o r b i t f r o m e q u a t o r i a l launch) in easterly direction.

W

ft/sec 52

18

ii

L0

500

WOO

1500 n.nti

Orbit height above moon Fig. 3. Booster a p p a r e n t v e l o c i t y r e q u i r e m e n t s f o r l u n a r circular o r b i t missions ( e q u a t o r i a l launch). 1 = high t r a n s i t v e l o c i t y ( t r a n s i t t i m e 30 hours), 2 = m o d e r a t e t r a n s i t v e l o c i t y ( t r a n s i t time 48 hours), 3 = m i n i m u m t r a n s i t v e l o c i t y ( t r a n s i t t i m e 100 hours).

N o w it is a fairly simple process for the rocket designer to compute the change in payload for a given change in orbit; curves of the type given in Figs. 2 and 3 give the change in velocity required and the designer can interpret this change in terms of the change in propellant required and payload alteration. It must be realised, however, that the variation in payload weight with orbit is not a simple, say, percentage change. Fig. 4 illustrates this point; there it is shown that for typical launching vehicles one may be more capable than the other of putting a larger payload into a low energy orbit but the reverse may apply for higher energy orbits.This situation may be simply attributed to the different all-burnt weights of the end stage without the inclusion of the payload. This dead weight also has to be put into orbit with the payload, and so, if it is heavy, most of any extra fuel energy may go in accelerating this weight leaving a less and less proportion for the payload; if the dead weight is relatively light in comparison with the payload in the low energy orbit a larger proportion of the extra fuel energy goes in producing extra payload energy and so a greater payload can be achieved. Since an increase in the number of stages results in lower dead

Jahrbuch 1961 d e r W G L

D. J. Lyons, Satellite launching possibilities

weights of the last stage, it is in general an easier task to provide useful payloads in the higher energy orbits with these than with say 2-stage vehicles. In the cases illustrated the 3-stage vehicle shows a better capability at higher altitude orbits than the 2g-stage (which is very little different to a 2-stage in the satellite launching role). Increase of percentage structural or dead weight, and decrease of specific impulse will also cause further degradation in payload as the orbit energy is increased. 5000

-a

lb iOOO

f ; 3000

i

2000 WOO

"

W2

2

s

103

Circular orbit

2

s

10in.mi

attitude

Fig. 4. Effect of orbit mission on payloads obtainable from typical boosters (circular polar orbits).

2.3 S t a g i n g Obviously, for a given number of stages, with given energy fuels and given structural efficiency, the maximum possible velocity is limited and even an infinite increase in size will lead asymetrically to a limiting apparent velocity. Fig. 5 illustrates this situation, by plotting the all-up weight 300n. mi circular

specific impulse of fuel in the first stage is also demonstrated for typical examples. It can be shown by simple theory that the performance of a rocket is a maximum for an infinite number of stages; however, inclusion of a number of practical considerations, such as separation devices, increased structural loads, systems for tracking, instrumentation and control which do not scale with size of the stage, and all of which increase the dead-weight proportion of the smaller stages, bring the optimum number of stages down to a reasonably small figure. Graham [1] suggests this number is 4 to 5, but this must depend on the particular case being examined. For cases where the payload is a very large percentage of the take-off weight a much smaller number of stages is more efficient than when the payload is a very small percentage of the all-up weight. However, there are other many practical factors which will tend to favour the rocket with the smaller number of stages; complexity is less, reliability is greater, development cost is smaller (due to the smaller number of items to be developed as well as fewer failures to be paid for) and manufacturing cost is smaller as the number of stages is reduced for a given maximum weight rocket. Of course, in reality the large cost of booster vehicles and their development, means that the greatest possible use of existing vehicles will be made and the rocket designer's problem is usually one of making the optimum use of existing vehicles, or those planned for a variety of tasks, rather than designing the optimum vehicle for each application. Another related and important factor is the inclusion of some "growth factor"; this depends on many factors inherent in the design of particular rockets but some general considerations apply. The greatest growth potential for the least cost will generally be available when the launching vehicle is originally planned to have a smaller number of stages and/or the end stages have a relatively low efficiency: then rapid improvement in performance can be obtained by either increasing the number of stages or by improving the efficiency of the upper stages, typically by re-designing for higher energy propellents. 2.4.

Apparent velocity

attainable

Fig. 5. Apparent velocities attainable, as a function of payload and stage performance.

at take-off necessary to give a 500 lb. payload a series of given apparent velocities. Assuming a mean specific impulse for the propellants of 280, and a mass fraction of 0.9, — i. e. fuel weight/(totaI weight — total carried load of the stage) — for each stage, three stages rather than two are needed if apparent velocities in excess of 36,000 or 37,000 ft/sec are required. The sensitivity of the performance to changes in the mass fraction from 0.9 to 0.85 and to changes in the

77

C i r c u l a r

O r b i t s

f o r

Radio

Use

A number of references have been made to circular orbits because of the increasing importance of these orbits particularly for use in radio applications. There are a number of reasons for this: i) Satellites on circular orbits stabilised with respect to the earth possess constant angular rotational speeds so that orientation of satellite aerial systems to make use of maximum gain aerials becomes generally much easier. ii) The motion of the satellite round the orbit, except for earth rotation under the orbit, gives reasonably consistent conditions at tracking/transmitting stations and for higher circular orbits relatively constant range. This eases the design of both ground and satellite borne systems. In particular equatorial circular orbits where earth rotation effects are more or less absent, give almost ideally constant tracking conditions. iii) Perturbations of circular orbits are much smaller so that prediction and control of tracking systems is much easier. iv) If the orbit time divides exactly into twenty four hours, this further eases ground aerial programming. 3. METHODS OF I N J E C T I O N 3.1.

A l l

Rocket

Systems

Practically all known satellite launching attempts have been made with all rocket systems. In such a system the initial launching is vertical, as maximum efficiency is

78

D . J . L y o n s , S a t e l l i t e l a u n c h i n g possibilities

J a h r b u c h 1961 d e r W G L

achieved with very low initial acceleration (sometimes as low as j g), but as the orbital velocity is required in a more or less horizontal direction the flight path is curved towards the horizontal as soon as it is possible to start acquiring horizontal velocity. The rate of turn is limited by requirements to keep the speed down whilst the booster is still in the atmosphere in order to keep aerodynamic heating and aerodynamic loads within reasonable proportions. As the stages burn out they are jettisoned and the succeeding stages ignited just before, during or after separation. In the most usual system all but one of the stages are burnt in rapid succession and have fairly high mean accelerations, firstly so as to derive the maximum possible energy increase f o r given momentum increases and secondly to minimize the gravity losses. Since, however, this results in burn out at a fairly low altitude and the launching vehicle must enter the required orbit under power, a coast period is usually inserted between the penultimate stage and the end stage while altitude is being gained, and the final stage is ignited near orbital altitude to increase the energy to that needed to remain in the desired orbit. This might be termed the thrust-coast-thrust t y p e of injection. It has the added attraction that on a simple mathematical approach it has the virtue of being the most efficient method of injection f r o m the point of cessation of thrust of the first stages; its disadvantages are also clear orientation systems have to be carried to hold spatial axis references f o r controlling direction of the end stage thrust, and there may be additional problems caused by the necessity to start the end stage engines f r o m conditions of free fall when liquids will not be nicely placed in their required position in the "bottom" of the tanks. An alternative method which has been evaluated in U. K. has been the elimination of the coast period and the substitution of a continuous but very low thrust end stage so that cessation of this thrust and attainment of initial orbit conditions are met at the same time. Fig. 6 illustrates

i) The low thrust lowers the engine weight very considerably which in turn permits increase of nozzle size, so that greater propellent efficiency (higher specific impulse) can be obtained. ii) The continuous thrust motor can and must be used to orient the stage so that a separate orientation system is unnecessary. iii) For reasons of i) and ii) the dead weight is reduced and the resulting increase in performance offsets the simply mathematically calculated loss in performance referred to above. iv) Light up problems during free fall can be eliminated if desired. v) Loss of cryogenic propellents due to solar heating, heat leakage, etc. can be avoided during the continuous thrust programme (this is not completely possible during coast). vi) If r a d a r guidance is used, the very low acceleration permits very long smoothing times both during climb and cut-off so that radar guidance becomes very accurate and early correction of orbit, possible. 3.2.

E f f e c t of H i g h R e q u i r e m e n t

P e r i g e e

If a high perigee is required for the orbit for either the thrust-coast-thrust system or the continuous thrust system, it is clear that during the injection phase the whole of the perigee altitude must be gained. Since for the sake of efficiency the climb angle must be small this inevitably means t h a t : a) the end stage or stages must possess orbital or very near orbital velocity at the beginning of the climb and b) that the point of final injection must be a long way round the earth f r o m the point of launch. End of / boost

phase

Privet

50 000 ft/sec iS 000 F i g . 6. D i a g r a m s h o w i n g d i f f e r e n c e b e t w e e n t h r u s t — coast — t h r u s t a n d c o n t i n u o u s t h r u s t . M e t h o d s of i n j e c t i o n i n t o o r b i t (300 n . miles c i r c u l a r o r b i t , t h r e e - s t a g e b o o s t e r ) . a) C o n t i n u o u s t h r u s t . 1 = m e a n a c c e l e r a t i o n 3 g d u r i n g f i r s t t w o stages; 2 = h e i g h t 120 n. miles, v e l o c i t y 19,750 f t / s e c , c l i m b a n g l e 1 2 ° ; 3 = m e a n a c c e l e r a t i o n 0.2 g i n t h i r d s t a g e ; 4 = e n t r y i n t o o r b i t ( v e l o c i t y 2 4 , 8 8 0 f t / s e c , h e i g h t 300 n . m i l e s ) ; 5 = g r o u n d r a n g e 2 7 0 0 n . miles. b) T h r u s t — c o a s t — t h r u s t . 1 = m e a n acceleration 3 g during first t w o stages; 2 = 125 n . miles, v e l o c i t y 1 9 , 7 0 0 f t / s e c , c l i m b a n g l e 1 4 ° ; 3 = a c c e l e r a t i o n 4 g; 4 = e n t r y i n t o o r b i t ( v e l o c i t y 24,880 h e i g h t 300 n . m i l e s ) ; 5 = g r o u n d r a n g e 1500 n . m i l e s ; 6 = c i t y 18,300 f t / s e c ; 7 = c o a s t .

height mean ft/sec, velo-

U2 000

38 000

3i 000

Is' boost phase cut-off lOOn.mi aLt. final orbit 300n.mi alt. i Coasting phase covers revolution of earth

t

30 000 0

1000

8000

half

(optimum}

12 000 n. mi.

Ground range covered during coast

the general features of the two methods f o r typical optimized launching rockets, assuming constant first and second stages. The advantages of the latter method are numerous:

F i g . 7.

Loss of e f f i c i e n c y d u e t o r e d u c t i o n in t o t a l l e n g t h of boost trajectory.

D . J . L y o n s , Satellite launching possibilities

J a h r b u c h 1961 d e r W G L

This general effect is demonstrated by Fig. 7 in which the thrust-coast-thrust system is assumed (though the same general results apply to continuous thrust trajectories) where the variation in apparent velocity required with ground range during injection is presented. For low orbits (300 n. miles) it is seen that down to 3000 miles range there is little change in rocket performance required, but for 5000 n. miles circular orbits a sharp rise in rocket performance is required if the coasting phase is reduced below one-half of a revolution round the earth. This has two major effects; any guidance or tracking stations must cover a much larger flight path if no uncovered periods are to be left, and final cut-off has to be observed from an area remote from launch; and secondly, if a continuous thrust system is used, extremely low values of motor thrust are necessary perhaps only tens of pounds. The use of a very low thrust in the continuous thrust system has a very attractive feature in that it permits an even larger radar smoothing time than for injection into low orbits; thus even though the radar range is very much greater for the high perigee orbits than the low, the required angular accuracy of the radar used is virtually the same in order to achieve roughly constant velocity accuracy for the two classes of orbits. 3.3. Aeroplane Launch Though the major portion of the speed needed for injecting a satellite into orbit must be given outside the atmosphere, in order to avoid too high aerodynamic heating effects, there are some potential advantages in using an aeroplane-like device for the first stage. The most obvious of these is the possibility of easier recovery and re-use of the first stage by making the first stage an aeroplane which can land in the normal way. Employment of present day aircraft is possible in this role though the performance gains are not striking; an aircraft, say with a high subsonic speed of M = 0.8, will only reduce the apparent velocity requirement of the rocket stages by 1500 to 2500 ft/sec. This reduction is larger than the actual velocity of the aeroplane stage by virtue of the altitude of the aeroplane at rocket launch (taken to be 40,000 ft) with the consequential increase in specific impulse and the advoidance of much of the aerodynamic drag losses in the rocket stage. The reduction in weight of the rocket stages required to be carried by the aeroplane as compared with an all-rocket vehicle depends a lot on the number of stages assumed and to some extent on orbit desired: if the aeroplane is counted as one stage, for the subsonic aircraft quoted above the total 10000 ft/sec

8000 — 6000

•2 e

1.000 2000

—

Launch Mach number

Fig. 8. E f f e c t of air-launch o n satellite v e l o c i t y i n c r e m e n t req u i r e m e n t a n d increase in p a y l o a d (300 n. miles c i r c u l a r o r b i t : t h r e e rocket stages). climbing launch, h o r i z o n t a l launch.

79

rocket weight may be about the same or even increased, but a 30 to 50 °/o rocket weight decrease may be obtained by keeping the same number of rocket stages as before (see Fig. 5). An increase in the aircraft speed to M = 5 will materially improve matters the weight reduction for an equal number of rocket stages being of the order of 65 to 75 °/o. Fig. 8 illustrates the drop in rocket performance required with varying aircraft speed. The economics involved in the use of aircraft first stages need a lot more study. 3.4. Guidance The task of guidance system is to monitor the behaviour of the launching vehicle under autopilot and to apply superimposed command heading changes designed to modify the injection path in the best possible way, subject to certain boundary conditions, meet the required orbit. There are at present two main ways of performing the navigation function, by radar means and by internal inertial navigation, though there are various combinations of the two, and, in fact, the autopilot and heading programme can be regarded as a form of perhaps inaccurate inertial navigation. It should not be forgotten, however, that some form of accurate radio tracking means is needed even if an all-inertial guidance system is used, to provide information to ground control. It might be thought that an ideal flight path could be pre-calculated before flight and the vehicle fired to follow this path; there are, however, a number of perturbations which in general do not permit the achievement of this state. Wind disturbances, thrust variations if a variable thrust device is not fitted, mass flow of propellent variations all cause alterations in the flight path, so that a computer is necessary to work out the best compromise course to steer when it is supplied with measurements of the deviation from the flight path. For an inertial system this computer is in the launching vehicle for a radar monitored system it will be on the ground. The advantages of the radar system are that the most complex part of the system is on the ground and the vehicle borne parts of the system can be fairly simple and light, beacons or transponders; the disadvantages are that complex and usually expensive ground equipments are necessary, and that fairly strong limitations are placed on directions of fire, in order to avoid the installation of multiple guidance systems. The advantages of the inertial system are that the system is self-contained in the vehicle and there is virtually no limitation on direction of fire; the disadvantages are that it is relatively heavy, is fairly expensive and is lost at each firing. It is difficult to predict the accuracy potential of both systems but since for most existing injection requirements both radar and inertial systems can be made sufficiently accurate, accuracy is not usually a criterion for the choice between the two systems. This is illustarated by Table 1 where the sensitivity of various orbit parameters to velocity errors are tabulated, a few tens of feet per second are usually adequate for most missions. An exception must be made for station keeping satellites which may need extremely accurate initial velocities; a form of radio or optical tracking would seem essential for this task, and then the objective is only likely to be achieved by orbit adjustment rather than by accurate initial injection and the problem is mostly a satellite problem rather than entirely a launching problem. It is of interest to draw special attention to the guidance of high circular or any high perigee launchings; as pointed out in Section 3.2 the high orbit is obtained most economically by an initial launch into a low orbit followed by an orbit transfer to the high orbit (coast or continuous thrust). This orbit transfer

80

D. J. Lyons, Satellite launching possibilities Table 1.

Jahrbuch 1961 der WGL

Errors in space mission due to velocity errors at injection.

I. C i r c u l a r

orbits

of

the

Earth

Circular orbit height

Apogee or perigee height error due to horizontal velocity error

Apogee and perigee height error due to vertical velocity error

Orbit time error due to horizontal velocity error

[n. mi]

[n. mi/(ft/sec)]

[n. mi/(ft/sec)]

[sec/(ft/sec)]

300 1000 5000

0.15 0.20 0.51

0.60 0.78 2.04 II. E l l i p t i c o r b i t o f t h e Cut-off at perigee

0.69 0.98 3.52

Earth

Apogee height

Apogee height error due to horizontal velocity error

[n. mi]

[n. mi]

[n. mi/(ft/sec)]

300

100.000

Perigee height

180

Apogeee and perigee height error due to vertical velocity error

very small

III. L u n a r i m p a c t t r a j e c t o r i e s Cut-off at 300 n. mi altitude above Earth Transit time

Permissible horizontal velocity error to graze Moon"')

Permissible vertical velocity error to graze Moon'1)

[h]

[ft/sec]

[ft/scc]

± 12

± 900

± 180

± 150

±50

±90

100 (minimum transit velocity) 48 (moderate transit velocity) 30 (high transit velocity) Moon radius = 940 n. mi.

when d o n e economically will be roughly a half revolution r o u n d the earth f o r t h e coast system ( H o h m a n n transfer) a n d r a t h e r m o r e f o r a continuous thrust system. A coasting system using r a d a r guidance will p r o b a b l y need a m i n i m u m of t w o guidance stations in t h e launch area (one in t h e immediate vicinity of launch, one at transfer orbit injection)

a n d one at final orbit injection: a conventional inertial guidance system could be used. A continuous thrust system m a y need an additional r a d a r guidance station f o r mid-course guidance (see Fig. 9) which can be o m i t t e d if t h e thrust p r o g r a m m e is accurate enough (within ± 1 fl/o in the end stage); in an inertial guidance system a special low acceleration system is needed during the low thrust phase, though there are p r o b a b l y no f u n d a m e n t a l obstacles p r o v i d i n g this. 3.5. Launching

Fig. 9. Launch trajectory of vehicle for high altitude circular orbit (5000 n. miles). Liquid hydrogen/liquid oxygen third stage. 0 = launch; 1 = first stage; 2 = second stage; 3 = third stage (first propulsion phase); 4 = acceleration 1.0 g changing to 0.013 g (altitude 200 n. miles, velocity 28,000 ft/sec, time 15 minutes, climb angle 0°); 5 = third stage (second propulsion phase) ; 6 = third stage cut off (acceleration 0.02 g, altitude 5000 n. miles, velocity 16,550 ft/sec, time 2.5 hours); 7 = down range tracking station (Manus Islands); 8 = possible site of mid-course tracking station; 9 = final tracking station.

Environment

It must be of f u n d a m e n t a l interest to designers of satellites to k n o w t h e environment p r o v i d e d b y the launching vehicle during t h e boost phase and the differences between this e n v i r o n m e n t and t h a t the satellite will have in orbit. T h e most i m p o r t a n t differences are in acceleration a n d vibration. T h e m a x i m u m longitudinal accelerations depend a lot on booster design and m a y v a r y between about 5 a n d 12 g; lateral accelerations are usually f a i r l y low, u n d e r + 1 g except p e r h a p s in a i r c r a f t launch w h e n this f i g u r e m a y rise by a f a c t o r of 2 o r more. T h e vibration e n v i r o n m e n t again is f a i r l y dependent on vehicle design, but a typical design v i b r a t i o n envelope which is given in Fig. 10 shows t h a t it is believed t h a t t h e level of vibration in t h e area of t h e satellite is unlikely to rise above an effective level of ± 1 to 2 g over t h e n o r m a l spectral b a n d . These vibration levels bearing in m i n d the impressed d. c. accelerations, should cause f e w design problems. T h e a e r o d y n a m i c heating during exit t h r o u g h the earth's a t m o s p h e r e does p r o v i d e a serious problem. Fig. 11 shows typical t e m p e r a t u r e variations w i t h time f o r t h e skin at

Jahrbuch 1961 der WGL

D. J. Lyons, Satellite launching possibilities

varying positions along the missile, with varying material also included. It is clear that at the forward end high temperature skins are practically unavoidable and the satellite has to be protected from both convection and radiation heating. Fortunately the heating stops on typical trajectories at a time of 3 minutes or so, so that the heat protecting devices, which are heavy, can be discarded if desired: the Blue Streak satellite launching vehicle proposals include provision for jettisoning satellite heat protecting covers at about 3 j minutes from launch. f

stage

2"'stage

\

r"

stage

1

V -^rtyrn Satellite

50 100

500

2000

81

binations have been considered in assessment of possible third stages. The first, hydrogen peroxyde (H.T.P.)/kerosene, a conventional, relatively low energy propellent combination, was considered for its comparative ease of development. The second, liquid hydrogen/liquid oxygen, yielding a greatly improved performance in terms of payload, especially in distant orbits, demands more extensive development. Only the final stage need be varied to meet a widely varied series of orbit requirements without significant departure from optimum staging conditions. Fig. 12 shows the overall configuration. At the pay load position, for illustration, is shown the outline of jettisonable fairings covering a typical low density solar cell powered satellite for communications work.

c/sec

Frequency

Fig. 10. Typical vibration envelopes for various positions during boost. 1 = position I, during first stage propulsion; 2 = position II, during second stage propulsion; 3 = position II, during first stage propulsion; 4 = position II, during third stage propulsion.

I. Fibreglass 0.2k"thick ''(6 "radius nose }

•III. 0.010 steel II. Light alloy j| 100 Time after

f

150 sec

stage

0.050"thick

cut-off

200

launch

Fig. 11. Typical wall temperatures resulting from kinetic heating during boost.

4. T H E BLUE STREAK SATELLITE L A U N C H I N G VEHICLE 4.1. Description This section gives an outline of the type of launching vehicle and its capabilities which will be available if the Anglo-French proposals for a joint European/Commonwealth multi-stage satellite launching go ahead. The present proposals are based firmly on the use of the British rocket Blue Streak as the first stage and the use of a French rocket, as yet unnamed, for the second stage. At the time of writing no firm joint European proposals have been made regarding the design and construction of a third, end stage for final injection into orbit, but in order to demonstrate the capabilities, two propellent com6

Fig. 12. Satellite launching vehicle. 1 = emplacement level, 2 = equipment panniers, 3 = rate gyros, 4 = kerosene tank, 5 = break up charges, 6 = liquid oxygen tank, 7 = W.R.E.B.U.S. aerials, 8 = slot transponder aerial, 9 = 465 telemetry aerials, 10 = unpressurised equipment bay, 11 = eight explosive bolts, 12 = U.D.M.H. tank, 13 = N 2 0 4 tank, 14 = equipment bay 15 = four explosive bolts, 16 = H.T.P. tank, 17 = pressurised equipment bay, 18 = second stage separation, 19 = third stage separation, 20 — beacon slot aerial, 21 = satellite separation.

The first stage, Blue Streak (Fig. 13) stripped of the equipment associated with its military guidance and warhead, requires minor modifications of the separation bay to carry the second stage.

82

D. J. Lyons, Satellite launching possibilities

It is propelled by two pump-fed liquid oxygen/kerosene engines, each rated at 137,000 lb. thrust at sea level, and in course of being uprated to 150,000 lb. thrust. The mass fraction of 0.93 of Blue Streak which includes an allowance for unburnt propellent has been achieved by the use of very thin stainless steel tank walls, and reliance upon internal pressurization for structural stiffness.

Tank section ox/dant tank and fuel tank with oxidant feed pipe running through fuel tank Equipment fairing compressed nitrogen bottles,etc Propulsion bay engine comprising two rocket motors

Road transportation

to produce relatively low temperature gas for the turbine drive. This turbine operates both fuel and oxidant combustion chamber injection pumps. Each chamber is gimballed in two planes to provide control of the vehicle in pitch, yaw and roll. The chambers are moved by hydraulic actuators, high pressure oil being derived f r o m pumps on the main propellent supply turbine gearboxes. The proposed French second stage will be propelled by a liquid propellent engine using nitrogen tetroxide ( N 2 0 4 ) and unsymmetrical dimethyl hydrazine (U.D.M.H.) with a sea level thrust of 25 tons. The propellents are pressurefed to the single combustion chamber by means of a solid charge gas generator. The thrust chamber is gimbalmounted for control in pitch and yaw, roll control being achieved by auxiliary jets.

Suidance bay gyros1 forward telemetry equipment ^tc

Equipment autopilot aft telemetry etc

Jahrbuch 1961 der W G L

Erection into tower

Fig. 13. Blue Streak (main external features).

The main structural sections, shown in Fig. 13, are: i) Separation bay (originally the guidance bay). ii) Tanks, oxidant and fuel. iii) Propulsion bay. iv) Equipment fairings. The propulsion bay is of aluminium alloy. In flight the oxidant tank is pressurized by gasification of liquid oxygen from the main tank, the fuel tank by nitrogen obtained by gasification of liquid nitrogen f r o m storage bottles. These cryogenic liquids are converted to the gaseous state by passage through heat exchangers located in the turbine exhausts. The R Z 2 engine, of which there are two (Fig. 14), consists basically of an integral power pack comprising gas generator, propellent pumps, gear box and turbine together with the thrust chamber assembly. Liquid oxygen and kerosene in a fuel-rich mixture, are fed to the gas generator

In addition to the propulsion system, control system actuators and electronics, the second stage contains equipment for the initiation of engine light up on separation f r o m the first stage, equipment for initiation of second/ third stage separation, and rate gyros for autopilot stabilization during the period of second stage operation. A third stage using the continuous thrust principles would need an engine working at a relatively low thrust/ level of between 1000 lb. and 3000 lb. (for the low perigee orbits), which could be started during separation f r o m the second stage. A proposal for an H.T.P./kerosene third stage is shown in Fig. 12.The engine is a four-chamber design, each chamber pivoted about one axis for steering in pitch, yaw and roll as in Black Knight. With low thrust and four chambers, very high nozzle expansion ratios, of 1000/1, are possible without undue chamber size, length and weight. To minimize overall stage weight, by reducing propellent tank weight, pump feeding of propellents was preferred in this particular design. The propellent pumps would be driven by a hydrogen peroxide steam turbine. It is possible to meet the varying orbital requirements by exchanging payload weight for propellant weight in the third stage whilst maintaining constant the overall weight of the third stage plus satellite payload at some 5000 lb. that is, the third stage incremental velocity can be increased at the expense of payload. The tank volume, only, is altered to suit the orbital mission, allowing the remainder of the third stage, including the engine and all equipment, to remain sensibly unchanged. A weight breakdown of the whole launching vehicle is given in Fig. 15. Nominal weight at launch 230783 ib

Is'stage

2" stage 13156 lb 3233kg

20757m

Si/5%

Propeäents 1931321b 37913 kg

All burnt weight 133371b 3210kg

Propellents 15i32lb 7330kg

Alt burnt weight 27131b 1233kg

3rdstage (space probe missioni

50401»

2237kg

Propellents ¡.3031b 1952kg

All burnt weight 7001b 317kg

Satellite fairings I jettisonable j iOlb 13kg

r

Fig. 15. Weight s u m m a r y .

Fig. 14. First stage propulsion bay.

Payload (nominal! 2001b 91kg

—i Vehicle 5001b 223kg

4.2. Performance Fig. 16 shows the estimated payload capability of the launcher with the illustrative H.T.P./kerosene end stage. These are based upon weight breakdowns as in Fig. 15, and a

Jahrbuch 1961 d e r W G L

D . J. Lyons, Satellite launching possibilities

sea level take-off thrust of 300,000 lb. (i. e. 2 X 150,0001b.) and with exchange between payload and propellent weight with mission in the third stage, as mentioned earlier. As will be seen, the payloads vary f r o m 1 ton actual payload in a 300 n. mile circle to 500 lb. in a 6000 to 7000 mile circle or in an elliptical orbit out to about 50,000 n. miles f r o m a 300 mile perigee. In obtaining these results, stage weights and flight programmes have been optimized as f a r as possible. In general the stage weight/payload curves are fairly flat around the optimum point, and it has been found advantageous to choose end weights slightly off optimum on the "light" side f o r a number of reasons, such as, for example, the location of the impact point of the discarded first stage in an acceptable area.

83

It is clear that the capabilities of the Blue Streak satellite launching vehicle as shown in Figs. 16 and 17 will cover an enormously wide variety of possible satellite requirements but if it is found necessary to use it, still further development potential exists by converting the second stage to liquid hydrogen, first by itself and later with two such stages above Blue Streak. 4.3.

F a c i l i t i e s

The major facilities for Blue Streak already exist and a large capital expenditure has been invested to obtain this situation. As examples, Fig. 18 shows the engine test area

Equatorial orbits '

(10)

Jahrbuch 1961 der W G L

Der eingeschlagene Weg hat zu einer großen Vereinfachung geführt, und obgleich die aufgezeigten Lösungen nicht beanspruchen, ein praktisch verwertbares Lenkgesetz zu liefern, so soll doch an dem Beispiel gezeigt werden, daß es notwendig ist, diese einfacheren Lösungen zu finden. 5. B A H N K U R V E N Die notwendigen Anhaltspunkte für die Bestimmung des Lenkgesetzes muß man durch Studium der Bahnkurven und der gestellten Aufgabe suchen. Das zeitliche Verhalten der Fehlergrößen und deren relative Bedeutung für das Endergebnis werden dann die optimale Auslegung des Lenksystems bestimmen. Bei der Erwähnung von Bahnkurven muß auf W. Hohaus dem Jahre 1925 [2] hingewiesen werden, in der gezeigt wird, daß vom energetischen Standpunkt aus der optimale Ubergang zwischen kreisförmigen Umlaufbahnen mit zwei impulsiven Schüben erreicht werden kann. Bild 6 ist aus Hohmanns Arbeit entnommen. Die Symbole V und E deuten darauf hin, daß eine Reise von der Erde zur Venus beschrieben wird. Die eingezeichneten Fahrstrahlen markieren die relativen Stellungen. Nimmt man an, daß das Fahrzeug sich nicht mehr im Gravitationsfeld der Erde befindet, aber noch in der Kreisbahn der Erde sich bewegt, dann erreicht man durch eine Geschwindigkeitsreduktion eine elliptische Bahn. Die Größe der Reduktion muß so bemessen werden, daß das Perigäum der Ellipse in der neuen Kreisbahn, in diesem Falle der der Venus, liegt. Eine weitere Geschwindigkeitsverringerung führt das Fahrzeug in eine Kreisbahn um die Venus.

manns Arbeit

Die letzte Gleichung besagt, daß in diesem Zustand der Flugkörper frei fällt, die eingeprägte Beschleunigung also Null sein muß. Diese Bedingung ist notwendig, aber nicht hinreichend. Weitere Bedingungen sind durch die Beziehungen zwischen Lage und Geschwindigkeit gegeben. Die beliebige Konstante C ergibt sich nicht aus den Anfangsbedingungen, da die Gleichungen nur für große Zeiten t zu gelten brauchen. Es kann gezeigt werden, daß, wenn die scheinbare seitliche Beschleunigung a y der Integralgleichung t

2

(11)

2g

S

-

av(s)ds

J say(s)ds

=

H

genügt, die obigen Bedingungen erfüllt sind. Von den möglichen Lösungen sei die einfachste, die Exponentialfunktion, herausgegriffen. (12)

a

/

yo

;

Für (13)

¿2/0 — y

€

v

1+

/1 +

ist Gl. (11) erfüllt.

a

2 a?H g

Bild 5 zeigt die Verwirklichung dieser Lösung für einen perfekten Autopiloten. Es kann weiterhin gezeigt werden, daß die Integration durch ein einfaches R-C-Glied ersetzt werden kann; 1 — a T ist eine Multiplikationskonstante.

Beschleunigungs-*Y f messer Autopilot uod Rugkörper^desponse

(Kommando)

SescMeun/gungs messer

3y (Kommando)

Bild 5.

Akernativlösung.

Bild 6.

Hobmann-Ubergang.

Die von der Sonne zu den beiden Orten der Schubänderung gezogenen Fahrstrahlen schließen notwendigerweise einen Winkel von 180° ein. In letzter Zeit sind weitere Arbeiten veröffentlicht worden [3], die sich auf Ubergänge in einer Ebene beziehen. Das allgemeine Problem von koplanaren Ubergängen zwischen Kegelschnitten verschiedener Art kann nur numerisch gelöst werden. Einige Sonderfälle sind behandelt worden: Ubergang zwischen Kreis und Ellipse, zwischen Ellipsen, deren große Achsen dieselbe Richtung haben, zwischen Ellipsen gleicher Exzentrizität und zwischen identischen Ellipsen verschiedener Orientierung. Die Ergebnisse scheinen zu bestätigen, daß die optimale Methode das Verfahren der zwei Schubimpulse ist. Die mathematische Behandlung dieser Probleme ist für die Auslegung des Lenkgesetzes wichtig. Das Verfahren der zwei Schubimpulse erfordert aber in jedem Fall ein Steuersystem, welches das Fahrzeug vor der Neuzündung in die

Jahrbuch 1961 d e r W G L

K. H . Schirrmacher, Lenkungs- und Steuerungsprobleme

geforderte Richtung dreht. Der im Bild 2 gezeigte Vorgang, einen Satelliten in eine Kreisbahn zu bringen, kann in gewissem Sinne als eine Form der zwei Schubimpulse aufgefaßt werden. 6. ÜBERGANG MIT K O N T I N U I E R L I C H E M SCHUB Untersuchungen in der Ballistic Missile Group des Royal Aircraft Establishment haben ergeben, daß die in Bild 2 gezeigte ballistische Phase und der impulsive Geschwindigkeitszuwachs durch einen sehr geringen kontinuierlichen Schub ersetzt werden können, ohne daß ein Verlust an Nutzlast entsteht. Dies gilt zumindest für das Projekt Blue Streak. Die Methode hat große Vorteile für die Konstruktion des Flugkörpers und besonders für die Lenkung und Steuerung. Ein System dieser Art scheint in den USA nicht in Erwägung gezogen worden zu sein, ist jedenfalls, soweit bekannt, nicht verwendet worden. Ein Vergleich zwischen beiden Methoden ist schwierig, da neue Parameter, die beim Fall der zwei Schubimpulse nicht auftreten, wie z. B. der Anstellwinkel, berücksichtigt werden müssen. Bild 7 zeigt die Bahnkurven für einen Übergang von einer Kreisbahn von 1600 km auf eine solche von 16 000 km Höhe. Hohmann— Übergang : Minimum-Energie Wallistische Phase zwischen A undB)

Bild 7. Vergleich zwischen Hohmann-Übergang Schub.

und konstantem

Bei diesem Beispiel würden — auf Grund der Energiebilanz — 2,5 °/o Nutzlast verloren werden. In der Praxis jedoch kann die Nutzlast sogar größer sein, da erstens das zusätzliche Steuersystem während der ballistischen Phase gespart und zweitens das Totgewicht durch Verwendung kleinerer Triebwerke verringert werden kann. Bei Radarsystemen ergibt sich ein anderer großer Vorteil. Mißt man die Lage, so erhält man die Geschwindigkeit durch zwei aufeinanderfolgende Lagemessungen, dividiert durch das Zeitintervall. Die Lagemessung ist mit Fehlern behaftet. Je größer das Zeitintervall, desto kleiner der Geschwindigkeitsfehler. Während des Zeitintervalls wird der Flugkörper beschleunigt. Solange die Beschleunigung bekannt ist, kann dies bei der Messung kompensiert werden. Unbekannte Beschleunigungskomponenten würden aber einen Geschwindigkeitsfehler hervorrufen, der das Produkt der unbekannten Größe und der Zeit ist. Es kann nun als sicher angenommen werden, daß diese unbekannten Beschleunigungskomponenten etwa proportional der Beschleunigung selbst sind. Ein kleiner kontinuierlicher Schub ist deshalb besser als der f ü r die Methode der zwei Schubimpulse notwendige.

89

7. AUSNUTZUNG DER V O R H A N D E N E N INFORMATION Kurz sei noch die Tatsache erwähnt, daß eine geschickte Ausnutzung der von dem Navigationssystem gelieferten Information das B'ordgewicht des Lenksystems wesentlich beeinflussen kann. Information bedeutet für ein Radarsystem Bandbreite; Bandbreite bedeutet elektrische Leistung und Leistung Batterie- und Gerätegewicht oder — bei gleichem Gewicht —: Reichweite. Systeme, die ein Minimum an Bandbreite benötigen, sind deshalb vorzuziehen. Moderne Dopplersysteme benutzen kleine Bandbreiten. Andererseits ist es schwierig, aus einer reinen Dopplermessung die für die Bahnkurvenbestimmung notwendige Lage abzuleiten. Es wäre denkbar, zusätzliche Information zu benutzen, z.B. die Kenntnis des Gravitationsfeldes, welches Geschwindigkeiten in bekannter Weise verändert. Probleme dieser Art sind noch nicht vollständig bearbeitet worden. Die sich abzeichnenden Lösungen lassen vermuten, daß die notwendigen Rechenoperationen sehr kompliziert sind. Sie wären aber für ein Radarsystem am Boden durchführbar. 8. D I E L E N K U N G AUF UMLAUFBAHNEN U N D VERWANDTE PROBLEME Im folgenden seien einige Lenkaufgaben erwähnt. Für Nachrichtensatelliten ist es notwendig, mehrere Satelliten in einer Umlaufbahn zu haben. Es ist wichtig, daß diese ihre Position zueinander über beträchtliche Zeiten behalten. Kein zur Zeit bekanntes Lenksystem ist in der Lage, die dafür erforderlichen Werte beim ersten Einlauf zu erreichen. Deshalb ist eine nachträgliche Korrektur in der Weise nötig, daß (wahrscheinlich mehrere) genaue Messungen und die daraus folgenden Adjustierungen vorgenommen werden müssen. Man kann diesen Prozeß immer noch als geschlossenes Regelsystem betrachten; die Filterzeiten allerdings erstrecken sich über Tage anstatt Sekunden, und die Korrekturen werden in täglichen oder wöchentlichen Abständen vorgenommen. Jeder einzelne dieser Prozesse bedeutet einen Übergang von einer auf eine andere Umlaufbahn. Es sei ein solches System von Nachrichtensatelliten mit Umlaufbahnen über die Pole angenommen. Man wird in der endgültigen Form in jeder Ebene etwa sechs Satelliten mit genauen Abständen voneinander brauchen und etwa drei bis vier Flugebenen, die 60° oder 45° zueinander versetzt sind. Eine derartige Anordnung kann überhaupt nicht beim ersten Einlauf erreicht werden, da die Umlaufzeiten nicht ganzzahlige Vielfache der Erdumdrehung sind, es sei denn, daß die Abschußbasen sich an den Polen befänden. Lenkaufgaben entstehen durch folgende Fehler: 1. Die Umlaufbahn ist gegen die gewünschte Meridianebene geneigt und läuft nicht über die Pole. 2. Die Umlaufbahn ist um die Erdachse gegen die gewünschte Meridianebene verdreht. Im ersten Falle muß eine Korrektur rechtwinklig zur Flugbahn vorgenommen werden, wenn der Satellit den Äquator kreuzt. Im zweiten Falle hat eine solche Korrektur am Nord- oder Südpol zu erfolgen. Natürlich können beide Korrekturen auch von einer beliebigen geographischen Breite gleichzeitig ausgeführt werden. Die Richtung und Größe der Korrektur hängt dann noch von der Breite ab.

90

K. H . Schirrmacher, Lenkungs- und Steuerungsprobleme

3. D e r Satellit ist in der richtigen U m l a u f b a h n , aber in i n k o r r e k t e r Position. 4. D e r Satellit h a t eine i n k o r r e k t e U m l a u f z e i t oder eine falsche Exzentrizität. D i e beiden letzten Fälle sind reine Übergangsprobleme. Die Ü b e r g ä n g e sollten idealerweise mit minimalem Brennstoffverbrauch durchgeführt werden. Die allgemeine Lösung ist noch nicht b e k a n n t . Die notwendigen Gleichungen können aufgestellt werden, es ist aber geraten, n u r spezifische, praktische Lösungen zu versuchen. Alle M a n ö v e r im R a u m sind solche Ü b e r g ä n g e zwischen Kegelschnitten jeglicher A r t . D i e P r o b l e m e w e r d e n weiterhin dadurch kompliziert, d a ß in der Praxis die Ü b e r g ä n g e sich im Dreidimensionalen abspielen. Reisen z u m M o n d o d e r zu den Planeten e r f o r d e r n mehrere Übergangsoperationen, da m a n v o n der interp l a n e t a r e n B a h n k u r v e auf eine U m l a u f b a h n u m den P l a neten oder auf eine L a n d e b a h n k u r v e übergehen m u ß . Sogar das Starten eines Satelliten (Bild 2) k a n n als ein Ü b e r g a n g angesehen werden, nämlich von der degenerierten Ellipse in der Startposition. Die Scheitel der Ellipse sind die Startstelle u n d der E r d m i t t e l p u n k t . Berechnete B a h n k u r v e n f ü r U m l a u f b a h n e n in größerer H ö h e erstrecken sich tatsächlich über fast 180°. Eine V e r k ü r z u n g w i r d durch aerodynamische Verluste in der A t m o s p h ä r e bedingt. Jeder Ü b e r g a n g bedeutet eine Lenkung, f ü r größere Reisen sind deshalb mehrere Lenkphasen notwendig. D i e Genauigkeitsanforderungen an jede einzelne Phase w e r d e n durch die A n f a n g s w e r t e u n d die M a n ö v r i e r f ä h i g k e i t des Fahrzeuges an der entsprechenden Stelle bedingt. Die A n f a n g s w e r t e hängen v o n der Genauigkeit der L e n k u n g der vorhergehenden Phase ab. D i e M a n ö v r i e r f ä h i g k e i t w i r d vom F a h r z e u g k o n s t r u k t e u r festgelegt.

Jahrbuch 1961 der WGL

fachheit m u ß leitendes K o n s t r u k t i o n s p r i n z i p sein. Alle erw ä h n t e n Lenksysteme benötigen ein Winkelbezugssystem, das durch den A u t o p i l o t e n das Fahrzeug auf K u r s hält. Dies k a n n n u r im F a h r z e u g selbst geschehen; die erreichte Genauigkeit beeinflußt die des Lenksystems. Meistens w e r d e n Kreisel oder stabilisierte P l a t t f o r m e n benutzt. Es sei aber darauf hingewiesen, d a ß über lange Zeit in solchen Systemen durch natürliche Driften Fehler entstehen, wenn nicht durch ä u ß e r e Bezugspunkte Rückk o p p l u n g e n b e n u t z t werden können. Dieser P u n k t ist von Wichtigkeit, da z. B. die Antriebsphase f ü r U m l a u f b a h n e n großer H ö h e (10 000 bis 15 000 k m ) mehrere Stunden d a u e r n k a n n . Auf die bek a n n t e n Steuermethoden sei hier nicht n ä h e r eingegangen. Es sei aber ein Vorschlag e r w ä h n t , Erdsatelliten durch den G r a d i e n t e n des Gravitationsfeldes zu stabilisieren. Beim Betrachten des dynamischen Verhaltens eines Satelliten in seiner Bahn findet m a n eine stabile Orientierung in E r d p o l a r k o o r d i n a t e n , vorausgesetzt, d a ß die H a u p t t r ä g h e i t s m o m e n t e verschieden sind. D i e erhaltenen M o m e n t e scheinen f ü r E n t f e r n u n g e n zwischen 500 u n d 15 000 k m größer zu sein als andere a u f t r e t e n d e Störmomente. Bild 8 e r k l ä r t das P r i n z i p . D e r Satellit befindet sich in einer Kreisbahn u m die E r d e m i t einer gleichförmigen Winkelgeschwindigkeit a) 0 . D i e Rollachse ist die T a n g e n t e an die Kreisbahn, die Gierachse zeigt zum E r d m i t t e l p u n k t u n d die Nickachse ist die N o r m a l e auf der U m l a u f e b e n e . M a n k a n n die A n o r d n u n g als einen Kreisel auffassen, dessen R a d sich mit der Frequenz w 0 um die Nickachse dreht. Gierachse Nickachse

Dies sei als Beispiel f ü r die N o t w e n d i g k e i t , alle Aspekte des Projektes gemeinsam zu behandeln, a n g e f ü h r t . Für längere Reisen mag es n o t w e n d i g sein, K o r r e k t u r e n auf halbem Wege a u s z u f ü h r e n . Es ist wahrscheinlich notwendig, d a f ü r eine F o r m v o n I n e r t i a l f ü h r u n g zu verwenden. Die A n f a n g s w e r t e m ü ß t e n mit H i l f e v o n Astronavigation erhalten werden. D e r Zweck einer solchen I n e r t i a l f ü h r u n g ist weniger die erreichbare Genauigkeit, als die o p t i m a l e A u s n ü t z u n g des v o r h a n d e n e n Treibstoffes. Weitere Lenkungsprobleme treten bei einem geforderten Rendezvous im R a u m zwecks B r e n n s t o f f a u f n a h m e u n d dergleichen auf. Probleme, die wahrscheinlich nicht v o n Lenkstationen auf der E r d e gelöst werden können. H i e r müssen Zielsuchtechniken angewendet werden. D i e b e k a n n t e n M e t h o d e n arbeiten mit der A n n a h m e , d a ß das G r a v i t a t i o n s f e l d N u l l oder k o n s t a n t ist. D i e bestehenden Theorien müssen deshalb auf veränderliche G r a v i t a t i o n s felder ausgedehnt werden. Eine weitere Erschwerung entsteht dadurch, d a ß a u ß e r dem Geschwindigkeitsvektor auch die relative Lage auf N u l l gebracht w e r d e n m u ß . Die erforderlichen N a v i g a t i o n s g e r ä t e sind, je nach der Aufgabe, nach verschiedenen Gesichtspunkten auszulegen. Astronavigationsmethoden, Horizontsucher, Magnetometer u n d die M e t h o d e n der Zielsuchtechnik müssen auf ihre Verw e n d b a r k e i t geprüft werden. 9. W I N K E L S T A B I L I S I E R U N G Z u r Lösung v o n Steuerungsproblemen der A s t r o n a u t i k w e r d e n alle v e r f ü g b a r e n regeltechnischen E r f a h r u n g e n herangezogen w e r d e n müssen. Die eigentlichen P r o b l e m e sind technischer A r t , nämlich zwischen kleinem Gewicht u n d Zuverlässigkeit einen K o m p r o m i ß zu finden, d. h., die Ein-

Flugrichtung-^o

! + 4 « V

(B

- C) 0 , - cü0 (A

B

02 + 3 aV

(A

-

-

B +

C) 0 8 ,

Nickachse: T2

=

C) 0 2 ,

Gierachse: T3 =

C 0

3

+

w02(B

- A ) 0

3

+

w0(A

~ B

+

C ) 0

l

(A, B, C sind die Trägheitsmomente um Rollachse, Nickachse u n d Gierachse). Die Winkelabweichung von der Rollachse ( 0 , ) u n d von der Gierachse ( 0 3 ) verursachen die Glieder ojn- (B — C) 0 , u n d co02 (B — A) 0 3 , weil die „Kreisel"achse aus ihrer N o r m a l l a g e v e r d r e h t w i r d . Dieser Effekt tritt bei einer D r e h u n g um die Nickachse nicht auf. D i e durch den G r a d i e n t e n des Gravitationsfeldes bedingten Glieder sind 3 « 0 2 (B - C) 0 j u n d 3 w 0 2 (A - C) 02.

K. H. Schirrmacher, Lenkungs- und Steuerungsprobleme

Jahrbuch 1961 der WGL

D i e A n o r d n u n g ist f ü r B > A > C stabil u n d f ü h r t Schwingungen aus, u m die Nickachse z. B. mit der Frequenz

o \

B

*

91

Es w i r d im allgemeinen unmöglich sein, das Problem u n t e r V e r w e n d u n g eines einzelnen Kreisels zu lösen. Eine Lösung wäre, mehrere Kreisel an der H ü l l e zu verteilen u n d die sich ergebenden W e r t e in geeigneter F o r m zu k o m binieren, eine andere, die verwendeten Kreisel w ä h r e n d des Fluges nach Bedarf umzuschalten.

Die praktische A n w e n d u n g dieser M e t h o d e ist von Bedeutung, da wegen der Einfachheit ihre Zuverlässigkeit groß ist.

In jedem Falle m u ß ein mathematisches Modell der Gesamtkonfiguration geschaffen werden, welches alle das P r o b l e m beeinflussenden P a r a m e t e r enthalten sollte. Die E r f a h r u n g h a t gezeigt, d a ß die Resultate der Berechnung eines solchen Modells experimentell geprüft w e r d e n sollten. Natürlich sind solche E x p e r i m e n t e mit großen Kosten verk n ü p f t , aber es ist ratsam, sie d u r c h z u f ü h r e n . Es ist weiter anzuraten, das „ T r a n s i e n t " - V e r h a l t e n der g e f u n d e n e n Lösung zu untersuchen, um ü b e r g r o ß e Belastungen der R a k e t e zu vermeiden.

Falls solche Stabilisierungsformen f ü r die Fein-Lenkphase, wie sie f ü r die Nachrichtensatelliten beschrieben w u r d e , gebraucht w e r d e n sollten, ist zu bedenken, d a ß die v e r w e n d e t e n Schübe sehr klein sein müssen.

A d a p t i v e Systeme machen die genaue Untersuchung der Schwingungsformen z u m Teil überflüssig. I h r Nachteil ist aber größere Kompliziertheit u n d d a m i t eine H e r a b s e t z u n g der Zuverlässigkeit des Gesamtsystems.

E r w a r t u n g s g e m ä ß liefern die Gleichungen u n g e d ä m p f t e Schwingungen. Geeignete D ä m p f u n g s g l i e d e r müssen hinzugefügt w e r d e n . Eine A n z a h l Lösungen sind vorgeschlagen w o r d e n , deren Ergebnisse recht ermutigend sind [ 4 ] . Die a u f t r e t e n d e n M o m e n t e sind klein, bei den üblichen A n o r d n u n g e n sind sie v o n der G r ö ß e 10 bis 100 d y n cm.

D e r Schub m u ß nämlich genau durch den Schwerpunkt gehen, u m Streumomente klein zu halten. D a der K o n s t r u k t i o n der mechanischen A n o r d n u n g Grenzen gesetzt sind, k a n n dies n u r durch kleine Schübe erreicht werden. Ein Hohmann-t)bergang w ä r e d a m i t ausgeschlossen. Es w ä r e also auch aus diesem G r u n d nötig, das Übergangsverhalten bei Benutzung von kleinen kontinuierlichen Schüben theoretisch weiter auszubauen. Die Alternativlösung, Kreisel u n d aktive Stabilisierungsmittel zu verwenden, die von Sonnen-, Horizontsuchern u n d dergleichen Gebrauch machen, w i r d sicherlich einen großen Anteil der N u t z l a s t beanspruchen. 10. A N D E R E S T E U E R P R O B L E M E F ü r große, leistungsfähige R a k e t e n m i t mehreren Stufen ergibt sich ein weiteres schwieriges Stabilisierungsproblem. Das Massenverhältnis ist groß u n d das Leergewicht klein. Die Bauweise ist leicht u n d gibt bei den in Frage k o m m e n den Längen zu Biegeschwingungen A n l a ß , welche das Arbeiten der z u r D ä m p f u n g der Steuerung notwendigen Wendezeiger erschweren. D a v o n ist natürlich auch der Lagekreisel betroffen. D i e E r f a h r u n g zeigt aber, d a ß die W i r k u n g auf diesen v o n untergeordneter Bedeutung ist. Steuerbewegungen der Motoren regen diese Schwingungen an. D i e A n r e g u n g s f u n k t i o n enthält zwei K o m p o n e n t e n . D i e eine w i r d durch die Trägheit des Motors bedingt u n d ist p r o p o r t i o n a l der Winkelbeschleunigung der M o t o r a u s lenkung, die andere ist dem M o t o r w i n k e l p r o p o r t i o n a l u n d ist durch die seitliche K o m p o n e n t e des abgelenkten Schubes bedingt. Die beiden K o m p o n e n t e n w i r k e n gegeneinander. Bei sinusförmiger Erregung ä n d e r t sich die Phase um 180°, u n d z w a r bei der Frequenz, f ü r die beide K o m p o n e n t e n gleich sind. Diese Frequenz w i r d „tail-wags-dog"-Frequenz oder I n versionsfrequenz genannt. Die Frequenzen der Schwingungsformen variieren w ä h rend der Antriebsphase wegen der Ä n d e r u n g der Massenverteilung u n d gehen im allgemeinen durch die „tail-wagsdog"-Frequenz. Das P r o b l e m ist nun, f ü r die Dämpfungskreisel entlang der H ü l l e eine Stelle zu finden, f ü r welche alle H a r m o nischen der Schwingungsformen w ä h r e n d der gesamten A n triebsphase stabil sind, d. h. keine u n g e d ä m p f t e n Schwingungen im Steuersystem erzeugt w e r d e n . Es sei bemerkt, d a ß die natürliche D ä m p f u n g praktischer T a n k k o n s t r u k tionen sehr klein ist.

Bewegungen des Antriebsstoffes im T a n k rufen ähnliche P h ä n o m e n e hervor. Diese müssen theoretisch untersucht werden. Es k a n n aber angenommen werden, daß mit verhältnismäßig einfachen Mitteln die Schwierigkeiten beseitigt werden k ö n n e n . 11. E I N I G E B E M E R K U N G E N Z U D E N VERWENDETEN GERÄTEN In diesen A u s f ü h r u n g e n ist fast gänzlich auf die Behandlung der verwendeten G e r ä t e verzichtet w o r d e n . Dabei sollten jedoch die mit der K o n s t r u k t i o n solcher Geräte verbundenen Schwierigkeiten nicht verkleinert werden, sondern m i t Rücksicht auf den Zweck dieser Arbeit, der sich darin erschöpft, die g r u n d l e g e n d e n P r o b l e m e aufzuzeigen. D e r technische S t a n d der Geräteentwicklung w i r d zweifellos die W a h l der L e n k - u n d Steuermethoden beeinflussen. Andererseits sollte aber auch das Studium möglicher Lenku n d Steuerverfahren f ü r die Geräteentwicklung richtungweisend sein. Alle physikalischen Prinzipien, welche es ermöglichen, Geschwindigkeit, Lage u n d Winkel zu messen, müssen in E r w ä g u n g gezogen werden. Die W a h l der M e t h o d e oder einer K o m b i n a t i o n v o n M e t h o d e n w i r d im wesentlichen durch die gestellte A u f g a b e bestimmt u n d h a t sich den Gegebenheiten des Gesamtprojektes zu unterstellen. Abschließend sei bemerkt, d a ß es die Absicht w a r zu zeigen, d a ß alle A s p e k t e des Problems berücksichtigt w e r den sollten, b e v o r m a n sich auf die mit derartigen Projekten verbundenen großen Kosten festlegt. 11. S C H R I F T T U M [ l j E. G. C. Bart: Theoretical principies of guided missile svstems. J. Roy. Aeron. Soc. 63 (1959), S. 455—464. [2] W. Hohmann: Die Erreichbarkeit der Himmelskörper. R. Oldenbourg, München 1925. [3] R. N.A.Plimmer: Fuel requirements for inter-orbital transfer of a rocket. In: Xth International Astronautical Congress, London 1959. Proceedings. Springer-Verlag, Wien 1960. [4] E. G. C. Burt: On the attitude control of earth satellites. Lecture given at the Eighth Anglo-American Aeronautical Conference, London, l l t h Sept., 1961. Ubersicht Die Grundlagen von Lenkungsverfahren werden diskutiert, ohne daß auf praktische Systeme (z. B. Radar oder Inertialsysteme) näher eingegangen wird. Besonders hervorgehoben wird die Notwendigkeit, die entstehenden Probleme nicht im einzelnen, sondern in ihrer Wirkung auf den gesamten Regelkreis, der das

92

K. H . Schirrmacher, Lenkungs- und Steuerungsprobleme

Verhalten des Fahrzeugs einschließt, zu behandeln. N u r so kann eine optimale Lösung gewährleistet werden, die einen Kompromiß zwischen Genauigkeit, Zuverlässigkeit und minimalem Fluggewicht darstellt. Der Zweck der Lenkung in der Astronautik wird definiert als das Mittel, jegliche Störung während der Antriebsphase zu erfassen, so daß die notwendigen Endbedingungen für das Erreichen einer vorgegebenen U m l a u f b a h n erfüllt werden. Ein vereinfachtes Beispiel wird angeführt. Die gestellte Aufgabe f ü r eine bestimmte U m l a u f b a h n bestimmt die Randbedingungen f ü r die Auslegung des Lenkungsund Steuerungsverfahrens. Als nächster Schritt sind Bahnkurvenberechnungen und eine Optimierung der Nutzlast auszuführen. Diese geben die G r u n d lagen für die Formulierung der Lenkungsgleichung. Die Wahl des Lenkungssystems hängt von allgemeinen Erfordernissen ab. Das Problem wird diskutiert mit dem Ergebnis, daß (zur Zeit) f ü r ein Radar-Lenkungssystem das Gewicht kleiner ist als f ü r ein Inertialsystem. Nach Festlegung des Verfahrens ist es notwendig, den gesamten Regelkreis zu betrachten, um größtmögliche Vereinfachungen ohne Genauigkeitsverlust zu erreichen. Unumgänglich ist es, Probleme des Überganges von einer U m laufbahn auf eine andere zu untersuchen. Als Beispiele werden folgende Aufgaben angeführt: die Aufrechterhaltung der relativen Lage von Satelliten zueinander (Nachrichtensatelliten), das Abfangen von Flugkörpern und Rendezvous-Probleme im Raum. Obgleich im allgemeinen die klassische Methode des Überganges {Hobmann) theoretisch als die wirtschaftlichste angesehen wird, wird gezeigt, daß in manchen Fällen der Praxis ein Nutzlastzuwachs erhalten wird, wenn die letzte Stufe einen sehr kleinen, aber langanhaltenden Schub erhält. Für Radar-Lenkungsverfahren hat diese Lösung den beträchtlichen Vorteil einer größeren Meßgenauigkeit, wodurch unter Umständen ein geringeres Fluggewicht ermöglicht wird. Eine weitere H i l f e f ü r die Vereinfachung ist beste Ausnutzung der von einem Navigationsverfahren erhaltenen Information, ein mathematisches Problem, das noch seiner allgemeinen Lösung bedarf. Bei R a d a r v e r f a h r e n wird man die Lösung bevorzugen, die die geringste Bandbreite benötigt. Eines der H a u p t p r o b l e m e der Lenkung ist die Notwendigkeit eines stabilen Winkelbezugssystems, so daß der Schub in der korrekten Richtung erfolgen kann. Die Möglichkeit passiver Stabilisierungsverfahren von Erdsatelliten wird auf Grund ihrer Bewegungsgleichungen diskutiert. Erwähnung findet die Schwierigkeit, aktive Stabilisierungsverfahren zu konstruieren, die über lange Zeit arbeiten. Steuerungsprobleme bei mehrstufigen Raumfahrzeugen werden diskutiert. Biegeschwingungen erschweren die Winkelstabilisierung. Die Probleme, die hier zu lösen sind, werden erwähnt. Lenkungsprobleme zum Mond und den Planeten erfordern im Prinzip keine neuen Verfahren, außer daß man bei Annäherung an das Ziel wahrscheinlich Endkorrekturen wird vornehmen müssen. Summary A discussion of the principles of guidance is attempted with only passing reference t o the actual equipment (radar, inertial guidance etc.). The discussion stresses the need to treat the problem as a system study of the complete loop including the behaviour of the missile. O n l y in this w a y can it be ensured that the most efficient solution can be found, which is the optimum in terms of accuracy, reliability and lowest airborne weight. The purpose of guidance in astronautics is defined as the means to overcome the effect of disturbances during the propelled trajectory with the aim to impart certain prescribed end conditions to the vehicle which are necessary t o achieve the required orbits. To illustrate the method of insertion of a vehicle into orbit and possibilities of simplification an idealised example is chosen. The mission will provide the boundary conditions within which the assessment of the guidance and control system has to be made.

Jahrbuch 1961 der W G L

The general background knowledge for the consideration of guidance is provided by extensive performance studies optimising on payload and suitable trajectories. The choice of the navigational part, i. e. radar, or inertia is discussed in terms of the general requirements, and it is shown that a radar method, at least at present, provides lighter airborne equipment. H a v i n g defined the requirements of the mission and the available and most suitable navigation equipment, problems of the complete loop have to be studied in order to simplify the system without loss of accuracy. It is essential to study methods of orbital transfer, problems which are recurring in astronautics. Examples of this are station keeping for communication satellites and for such missions as satellite interception and rendezvous in space. The standard method of transfer is generally considered theoretically to be the best from the noint of minimum fuel consumption. It can be shown, however, that a low continuous thrust has advantages which in practice result in an increase of payload in comparison with Hohmann's two thrust-impulse method. In addition, considerable advantages are obtained for a r a d a r guidance system, as the low thrust enables a more accurate measurement and in consequence a lower airborne weight of the equipment. The most efficient w a y to use the available information from a navigation system is also a means of simplifying the equipment. In a r a d a r method the best way is to use information which requires the lowest b a n d w i d t h for transmission. One of the major problems in guidance of satellites and spacec r a f t is the need for a stable angular reference so that thrust can be applied in the correct direction. The possibility of passive stabilisation of earth satellites in earth axes is discussed by considering the dynamics of their motion. The difficulties of providing suitable reference systems for active stabilisation are mentioned. The control problems of multi-stage vehicles are discussed. Body bending modes make it difficult to stabilise the angular motions and the need for a comprehensive study of this problem is mentioned. Guidance problems f o r journeys to the moon or planets impose no new requirements in principle except that it will probably be necessary to use terminal corrections. Résumé Les principes des méthodes de guidage sont exposés sans considérer de plus près les systèmes de guidage soit p a r r a d a r soit p a r inertie. La nécessité d'examiner ce problème p a r une étude de tous les systèmes faisant partie du boucle entier y compris le comportement du véhicule est mise en évidence. Ce n'est que de cette façon qu'une solution efficace du point de vue précision, sécurité de fonctionnement et poids total peut être obtenue. Dans l'astronautique on entend p a r guidage les moyens permettant de compenser l'influence de perturbations pendant le vol propulsé de l'engin afin de lui communiquer certaines conditions terminales nécessaires pour le faire graviter sur l'orbite voulue. La mise en orbite d'un engin et les possibilités de simplification de ce procédé sont expliquées à l'aide d'un exemple idéalisé. La mission prévue représentera les conditions limites de l'évaluation du système de guidage et de commande. La connaissance de base pour l'étude du problème de guidage est dérivée de résultats d'études étendues de performances dans des conditions optima de charge utile et de trajectoires convenables. Le choix des moyens de navigation, à savoir, guidage par r a d a r ou p a r inertie est considéré dans le cadre des conditions générales et il est démontré que la possibilité d'utiliser des équipement plus légers est actuellement encore fournie par le radar. Après détermination de la mission et des systèmes de navigation les plus appropriés, les problèmes du boucle entier doivent être étudiés afin de simplifier le système sans diminuer sa précision. Une étude des procédés de transfert orbital, problème se posant souvent dans l'astronautique, s'impose. O n cite à titre d'exemple la nécessité pour les satellites de communication et d'interception de tenir leur positions respectives et rendez-vous

W . J . J a c o b i a n d C . S. B r i d g e , G u i d a n c e a n d control equations

J a h r b u c h 1961 d e r W G L

dans l'espace. Bien que l a m é t h o d e classique d e t r a n s f e r t (Hohmann) semble la p l u s économique, il est d é m o n t r é qu'une a u g m e n t a t i o n de la c h a r g e utile est q u e l q u e f o i s obtenue p a r une poussée f a i b l e mais continue a u dernier étage. E n p l u s , des a v a n t a g e s i m p o r t a n t s sont réalisés p o u r un système de g u i d a g e p a r r a d a r , la p o u s s é e réduite p e r m e t t a n t une mesure p l u s précise, d o n c un p o i d s réduit des équipements de bord. L a f a ç o n la plus e f f i c a c e d'utiliser les i n f o r m a t i o n s disponibles d'un système de n a v i g a t i o n représente en m ê m e temps un m o y e n d e s i m p l i f i c a t i o n du matériel. D a n s le cas d ' u n système r a d a r il sera plus a v a n t a g e u x d'utilisier p o u r émission la largeur de b a n d e minimum. L ' u n des p r o b l è m e s p r i n c i p a u x que pose le g u i d a g e des satellites et v a i s s e a u x s p a t i a u x est la nécessité de disposer d'une

référence a n g u l a i r e stable p e r m e t t a n t l ' a p p l i c a t i o n d e la poussée dans la direction correcte. L a possibilité d ' e f f e c t u e r la stabilisation p a s s i v e des satellites terrestres dans les axes de la terre est traitée p a r une considération de la d y n a m i q u e de leur mouvement. Les d i f f i c u l t é s de p r é v o i r des systèmes de référence convenables a u x fins de la stabilisation a c t i v e sont indiquées. Les p r o b l è m e s de c o m m a n d e des véhicules à plusieurs étages sont e x a m i n é s . Les m o d e s de f l e x i o n du corps rendent d i f f i c i l e la stabilisation des m o u v e m e n t s angulaires et la nécessité de mener une étude c c m p r é h e n s i v e de cette question est mentionnée. Les p r o b l è m e s de g u i d a g e p o u r les v o y a g e s vers la L u n e o u les planètes n'imposent p a s en p r i n c i p e d e conditions nouvelles, sauf qu'il sera p r o b a b l e m e n t nécessaire d ' a p p l i q u e r des corrections terminales.

W. J. J A C O B I A N D C. S. B R I D G E , W O O D L A N D H I L L S , C A L I F . , U S A * )

F O R M U L A T I O N OF G U I D A N C E A N D C O N T R O L E Q U A T I O N S , THEIR MECHANIZATION AND

This paper describes the derivation, mechanization and instrumentation of guidance and control equations f o r a class of vehicles that is subjected to the following environment: Altitude: 0 to 1,000,000 ft. Velocity: 0 to 26,000 ft/sec. Acceleration: 0 to ± 1 0 g. Time: u p to three orbits without external inputs.

M a j o r system output functions are: 1. 2. 3. 4. 5. 6.

Velocities in earth-fixed and space-fixed coordinates. Altitudes above earth's mean surface. Angles relative to local vertical. Angular position in earth-fixed coordinates. Automatic flight control information. Energy management parameters.

The problem is approached in three distinct phases. They are: 1. Preliminary Systems Engineering Study. II. Development of Guidance and Control Equations. I I I . Mechanization of Equations and System Definition. 2. D E F I N I T I O N O F S Y M B O L S U S E D mass of vehicle effective wing area of vehicle pitching moment of inertia moment arm from pitch reaction jet to vehicle center of gravity Cla lift slope of vehicle Cdo parasitic drag coefficient d C n l d C j } induced drag coefficient

ma S ]y x

*) Litton Division.

Systems,

Inc.,

Guidance

and

D K 621—52/-S5 533.6.011.55.015 681.142 627.62

INSTRUMENTATION

1. I N T R O D U C T I O N

(1) (2) (3) (4)

93

Control

Systems

Q K k m Kv ktp At mv x 6e V y R B 45°),

(67) = arc s i n — — (if i > < 4 5 ° ) , v ' cosyjr where N and D are given by Eqs. (57) and (58). Whether ¡T is greater or less than 90° can be determined from cos f j ' = sin T sin + cos cos 0g cos (X3 — Xr)-

Since no propulsive force is available during re-entry, the only control energy is a pitching moment and a rolling moment capability. These moments may be accomplished through the use of any type of reaction jets or aerodynamic control surfaces. The block diagram shows how the commands are produced and the various override functions that are incorporated. This method of control requires the following gain constants which must be specified according to the vehicle and stored in the computer: K = K,f, = k 0 ,

( e n+ i)nc

(73)

V e = ß j t COS g,

(93)

0X

(74)

V,.2 = V 2 + V /

Ar = arc t a n

=

2

V2sin2y

-

-

Vf

2

— 2 K (e n +')nr. =

±

viim

-

1

'

1

(94)

KrtAf-'

(95)

sin

for c o s < M , < 0 , dAr = 9 0 ° - > 1 8 0 ° . (90)

2 V V e cos 7 sin B COS

sin @f + + cos 0g cos r

(102)

A small only),

= 0, L

=

CLaa,

1

=

ÌQVr2,

(83)

L

=

qSCL,

(84)

CD

-

(85)

D

=

(86)

rt

r

0

+ -

3 C d

dCL2

r

2

T

"

If r ,07.

(87)

T

t

= \&„~

,>

Vr = -

t

D

= -

e,pr = r^,

r2 > r2s.

With reference to the earth, rt is the local orbit of the launching space station. When escaping from planet 1 or earth base, the vehicle commutes at a certain distance from the local to the central field; correspondingly, at the capture maneuver at planet 2, this border line is at r2. The transition between local and central field is gradual, of course. But the abstraction of this so-called "radius of activity sphere" rp1, P = 1 or 2, is a useful one. This radius is determined by computing the dynamical equilibrium between the attractive forces of sun and planet on a probe (vehicle) in the rotating coordinate frame of the planet revolving around the sun; it is found to be identical with the question of the libration points for the collinear case of Langrange's three-body problem [3]: r

P_ Rp

mp Y meJ

The numerical evaluation shows that rp1 Rp, but that also rp1 is many times larger, up to some order of magnitudes, than both, rp and rps, for all planets of the solar system; therefore, this condition is on the safe side, as it should be, and need not be considered further. 4. CLASS A TRAJECTORIES The objective is to transfer a vehicle from the base orbit Rx to the target orbit R2 (see Figs. 2 and 3). It is natural to ask for a transition with a minimum amount of pro-

Fig. 3. Transfer trajectory, class A.

rocket power burst (impulse) is applied tangentially at PtL, the launching position, conferring on the vehicle the velocity increase or velocity kick A V t . It arrives at the target position P2 with the velocity V2 at the same time as the target station P2, moving with the circular velocity V2c > V2. Therefore, on arrival, the vehicle has to be synchronized by a second velocity kick A V2.

4.2. Velocity

Terms

The computation of the velocity terms is performed by applying the above basic formulae, especially II and IV. The numerical calculations are simplified by normalizing the terms to corresponding reference quantities of Pt:

MC

where

V2 Q km/sec, 1™/-»,1C = ( ^^ ^ I = = IQ 29.8

V

Ri

»

=

The resulting velocity terms are, step by step:

Vl.

1 + l/R, A V4 = Vj - 1,

v 2 = R*

A V2 = V 2 c - V2,

V2C = V2-° = AV1 + AV2.

The sum of the velocity kicks is the so-called velocity sum Vz° of the basic case and is a characteristic measure for the rocket size and load ratio at take off. The notation V i is always used for the single trip; this amount has to be doubled for the return trip.

Jahrbuch 1961 d e r W G L

Allowance has now to be made for the influence of the local potential holes. If the vehicle is required to ascend and descend in those holes, one would expect that additional contributions for the velocity sum are necessary. But the so-called paradox of rocket energy conversion shows that it is best to apply the impulse of the rocket at as high a rocket velocity as possible. Consequently, it appears favorable to go into the gravity hole to a certain degree, pick up velocity therewith and apply the rocket impulse under this optimum condition [6, 7], 4.3.

G e n e r a l

105

C. A. Traenkle, Mechanics of interplanetary space travel

a n d

O p t i m a l

H y p e r b o l i c

T r a n s f e r

The geometry and kinematics of the capture in, or correspondingly, of the escape from the local holes is represented in Fig. 4.

The velocity sum V /' for the optimal case follows from that by substitution: 1 V„* = =Vy 1/2 However, to establish the optimal case, one has to guide the vehicle into the pericenter distance rp* or, equivalently, when approaching from afar, setting it on the collision parameter b* = ]/2 rp"'; but this is a separate problem of terminal navigation. Another condition to be checked is that rp must not fall under the solid surface rps of the planet: rp*

->1.

rP

The formulae for numerical evaluation follow after some transformation as: C A V

AVJ

t 2

where C =

2

V i f V. 1 c

7.91 29.8

M • s

The velocity difference A Vp (P = 1 and 2) is now not executed immediately as velocity kick; this is delayed until the vehicle reaches the intended distance rp of the local circular orbit. Therefore, A Vp is now the relative velocity with which the vehicle crosses the local activity sphere rp1. As rp1 > rp, practically, A Vp is the approach velocity at infinity of the local field, and the corresponding orbit is a hyperbola. By applying the energy theorem, the velocity vp at pericenter A is: =

2

A VP2

+

v

100

going down for R

the phase angles or constellations of the start and target planets at departure and arrival of the vehicle. According to the Third K e p l e n a n Law, the periods of the R and R orbits and of the A ellipse are, respectively: t

R r -

G©

T

2

- 2

-~jlt

ti

—

2

ti

2

T * J-

t

20

50

100

n,

Fig. 6. Velocity sum V j , class A , single trip.

actual values V ^ " which can be attained for the planets, allowing for the limitations on rp*, exhibited in Fig. 5, are embedded between the 0 and * curves: For the far planets, they lie close to the curve, for the near planets Venus and Mars close to the 0 curve, because of the reasons explained above. The curve branches for the inferior planets R2 < 1 can be imaged from those of the superior planets R 2 ^ > 1 by reciprocity relations, as will be shown below. For the case of the moon, only its own gravity hole can be utilized; therefore: V /
). —Water —Gas

flow flow

P l a s m a generator (schematic diagram).

4.2. The Tunnels at the University of Toronto Institute of Aerodynamics, Armament Research and Development Establishment, Fort Halstead, and National Physical Laboratory, Teddington These tunnels are in many respects similar to the Berkeley tunnel. The main difference is in the pumping system. UTIA and ARDE use six three-stage oil diffusion

124

I. Estermann, Simulation of high altitude flight conditions

booster pumps of the type shown in Fig. 5 giving a total pumping speed of 7000 1/sec. The pumps are backed by two 200 cu f t / m i n mechanical pumps. T h e operating pressure range is f r o m about 10 to 70 ¿tHg. N P L has just completed a somewhat larger facility of the same general design, using five four-stage booster pumps with a total pumping speed of 20,000 1/sec, backed by five Roots blowers and two mechanical pumps. The test section pressure ranges f r o m 10 to 100 ; «Hg, nozzles f o r Mach numbers f r o m 2 to 10 are contemplated. The complete arrangement of the N P L tunnel, which is very similar to those mentioned previously, is shown in Fig. 6.

4.4.

T h e

Jahrbuch 1961 der W G L T w o - p h a s e

v e r s i t y

of

g i n e e r i n g

T u n n e l

a t

t h e

S o u t h e r n , C a l i f o r n i a C e n t e r ,

L o s

U n i E n -

A n g e l e s

In this tunnel, a novel and radically different pumping system, namely the condensation of the tunnel air on a surface cooled to T < 20 ° K ("cryopumping") is applied. If every gas molecule striking such a surface is condensed

GaS in

Fig. 7. Schematic view of the low-density w i n d tunnel of the Laboratoire Méditerranéen in Nice.

Fig. 5. Booster p u m p .

or trapped, each cm 2 provides an equivalent pumping speed of 10 1/sec. It is, therefore, feasible to reach pumping speeds f r o m 10» to 10 6 1/sec at pressures of 1 ,wHg or less in the test section. The condenser used in the U S C tunnel has six plates, each having an area of 10 ft 2 , which are cooled to about 15 to 20 ° K by means of helium gas precooled in a helium refrigerator placed outside the tunnel. The latter is a Collins type cryostat employing two reciprocating expansion engines capable of removing 350 watts at 20 ° K ,

Fig. 6. Low-density wind tunnel at N a t i o n a l Physical L a b o r a t o r y , T e d d i n g t o n ( C r o w n copyright — reproduced with permission of H . M. Government).

4.3.

T h e

T u n n e l

a t

M é d i t e r r a n é e n ,

t h e

L a b o r a t o i r e

N i c e

This tunnel is designed for lower operating pressures than those mentioned previously, f r o m about 0.1 to 10 /iHg. The pumping system consists of four three-stage 011 diffusion pumps with a total capacity of 30,000 1/sec at 0.1 //Hg, backed by booster and mechanical pumps. It has been reported that Mach numbers of 4 have been obtained with a nozzle diameter of 4 cm, producing a uniform f l o w over a core of about 1 cm diameter. A schematic view of the installation is given in Fig. 7.

with a rated power consumption of 50 hp. The tunnel itself consists of a large steel tank measuring approximately 9 feet in diameter and 33 feet in length, with a sump-like appendage at the downstream end housing the condenser (Fig. 8). Before entering the nozzle, the nitrogen gas used for tunnel operation is heated by means of a graphite heater to 800 ° K , it then passes through a large settling chamber before entering the nozzle (Fig. 9). T h e conical nozzle has an angle of 4 0 ° ; to exercise control over the boundary layer developing inside the nozzle with high Mach numbers at low pressures, the nozzle walls can be

I. Estermann, Simulation of high altitude flight conditions

Jahrbuch 1961 d e r W G L

cooled with liquid nitrogen. After about 10 hours of operation, the thickness of the solid nitrogen deposit on the condenser plates becomes so large that thawing becomes necessary. The nozzle has an exit diameter of 19" and is 12^" long, giving a Much number of 8. It is interesting to note that only a core of 2" diameter is filled with uniform flow (see Section 4.5).

125

for the UTIA tunnel). Attempts have been made to reduce the boundary layer thickness by cooling of or by suction through slots in the nozzle walls, but it appears certain that complete exploration of the free molecular flow regime requires additional equipments.

Reynolds number

Fig. 8. Schematic layout of the low-density wind tunnel at the University of Southern California, Los Angeles. 1 = jack, 2 = stainless bellows seal, 3 = modulating valve 10" travel, 4 = floor line, 5 = cryostat, 6 = condenser, 7 = plate coils, 8 = heater plug in, 9 = aftercooler, 10 = access port, 11 = probe mount, 12 = nozzle, 13 = valve.

Leading egde flat plate K\\\\\1 Base pressure EHSSi Cone pressure I 1 Cone drag

Fig. 10. Mach and Reynolds number range, wind tunnel in Berkeley. Characteristic dimension = 1 ft.

-Access port

, 9. Nozzle assembly of the University of Southern California wind tunnel, probe mount, 2 = nozzle, 3 = stagnation chamber, 4 = valve. 4 . 5 . General Wind

Comments Tunnels

on

Low

Density

Low density wind tunnels as described in the preceding section have made it possible to study the basic flow and heat transfer problems in the slip and transition range, with a few stabs into the direction of free molecular flow, for flows up to about M = 8. (Fig. 10 shows the flow regions attainable with the Berkeley tunnel.) The application of plasmajets has made it possible to begin simulation of the temperature and ionization conditions of high altitude flight. Extension of this work into the free molecular flow regime will require much larger installations than those presently in existence or in the construction stage. These, however, do not appear to be impossible. Diffusion pumps with speeds of 50,000 1/sec (32" diameter, 4 stage) are becoming available, and it is to be hoped that cryopumping will be developed sufficiently to permit reliable and continuing performance below the 1 ¿¿Hg pressure range. One of the inherent difficulties is the growth of the boundary layer thickness inside a nozzle with decreasing pressure, which reduces the diameter of the uniform core of the airstream flowing from the nozzle to a small part of the nozzle diameter (as shown in Fig. 11

1.0 inch 1.5 Distance off axis

Fig, 11. Mack number profiles at exit plane at various operating

pressures of the Toronto wind tunnel. Operating static pressure [i" Hg]

o A

V

• •

45 41.1 40.0 34.8 28.8

Another fundamental problem concerning simulation of flow conditions by wind tunnels refers to actual velocity of the air stream with respect to the model. The high Macb numbers in super- and hypersonic tunnels do not necessarily indicate high speeds of the air flow, but only a large ratio between this speed and the local velocity of sound. The latter, however, is proportional to the square root of the absolute temperature. The expansion of air through a nozzle, being largely isentropic, produces a very strong cooling of the effluent gas, and a large part of the high

126

I. Estermann, Simulation of high altitude flight conditions

Macb numbers produced is due to the corresponding reduction of the speed of sound and not to the acceleration of the gas flow. This effect can be compensated, at least in part, by the installation of plasmajet or other heaters upstream of the nozzle, but together with the growth of boundary layer thickness mentioned earlier, it severely limits the usefulness of wind tunnels for the simulation of high altitude flight conditions to relatively low speeds and altitudes. In the following paragraphs, we shall give examples of other equipments which are potentially useful for the solution of this problem for higher speeds and altitudes. 5. T H E "MOLECULAR G U N " OF T H E LABORATOIRE MEDITERRANEEN This newly installed instrument is a combination of a molecular jet and a wind tunnel and has as its objective the simulation of the interactions between a body moving at a very high speed and a rarefied gas in the free molecular flow regime. While a "normal" molecular beam employs molecules of thermal velocities, which are approximately equal to the speed of sound, this apparatus produces particles of much higher speeds by acceleration of electrically charged particles and their subsequent neutralization. The general design of the apparatus is shown in Fig. 12.

Jahrbuch 1961 der WGL

6. REVOLVING ARMS, DISCS, A N D CYLINDERS These instruments have been built in various laboratories and are useful for aerodynamic measurements in the free molecular flow regime under certain precautions. Their main advantage is that there is no fundamental limitation to the Knudsen number which can be obtained since they are, from a vacuum technology standpoint, static systems which can be pumped down without much effort to pressures of 10 - 2 ,wHg and below, providing mean free path lengths of the order of meters. Their main drawback is ' their inability to reach high linear velocities, 800 m/sec being about the top speed that has been reached in actual use. Revolving cylinders are useful for drag and heat transfer measurements in rarefied gases, and revolving arms and discs allow the investigation of flow patterns over models in various media including ionized gases in the free molecular flow regime. As an example, we show the revolving arm of the Laboratoire Méditerranéen de Recherches Thermodynamiques (Fig. 13), which is arranged for

Fig. 13. Revolving arm at the Laboratoire Méditerranéen. 1 = quartz tube, 2 = RF coil, 3 = pyrex tube, 4 = model, 5 = revolving arm, 6 = probes, 7 = gas inlet, 8 = pump, 9 = shaft, 10 = seal. Fig. 12. Molecular gun at the Laboratoire Méditerranéen. 1 = argon intake, 2 = nozzle, 3 = RF coil, 4 = accelerator, 5 = magnet, 6 = valve, 7 = pump, 8 = decelerator, 9 = magnet, 10 = observation chamber, 11 = detector.

The experimental gas, e. g., argon, is admitted through a control valve into a supersonic nozzle where a molecular jet is formed. From there, it passes through a quartz tube, where it is partly ionized by a high-frequency discharge, into the first vacuum chamber. The resulting mixture of electrons, positive and negative ions, and neutral atoms is passed through an electrostatic accelerator and lens system from which a reasonably homogeneous positive ion beam emerges. This beam is deflected through an angle of 90° and thereby separated from the neutral and negative particles by means of an electromagnet. A second electrostatic lens and decelerator system refocusses the positive ions and reduces their velocity to the desired value, corresponding to energies in the range of 20 to 100 eV. The ion beam is then intersected by a second beam of neutral argon atoms. Since the effective cross section for charge exchange is much larger than that for momentum transfer, a fair portion of the ions will be neutralized without an appreciable change of velocity. A second magnetic field removes the remaining ions, and the final beam entering the test section consists of nearly mono-energetic neutral atoms. Beam densities of 3 X 106 molecules per square cm per sec have been reported and their application to aerodynamic measurements is in prospect.

operation at pressure levels between 0.25 and 5 //Hg. The vacuum tank has a diameter of 1.5 m and can be evacuated by means of an oil diffusion pump. The arm, made of hightensile strength duraluminum, has a diameter of 1.25 m and is driven by an external motor at speeds up to 9000 rpm. Ionization up to 1 0 % is obtained by means of an electrodeless high-frequency discharge in a quartz or pyrex tube attached to the top of the tank. Gas is admitted through a controlled leak which also provides pressure regulation. The arm may be used to carry models for aerodynamic tests or Langmuir probes for the exploration of charge exchange. 7. MOLECULAR BEAMS While direct simulation of flight conditions in the free molecular flow regime appears to present severe difficulties, the theoretical treatment of this regime, based on the kinetic theory of gases, is much simpler than that of the transition region. It requires, however, the knowledge of certain details of the energy and momentum exchange between individual molecules and solid surfaces. These details, which are usually expressed as accommodation coefficients, vary from substance to substance, and must be determined by experimental methods. In addition to classical methods involving measurement of heat transfer and viscosity in rarefied gases, the molecular beam method has recently begun to be employed for this purpose. Though simple in principle, the actual experiments are quite difficult and have produced only few results thus far.

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I. Estermann, Simulation of high altitude flight conditions

A molecular beam consists of a stream of particles m o v i n g in a highly evacuated space in rectilinear trajectories defined geometrically b y a set of apertures. A

Oven room

to.

2

5 cm

A

Beam room

;

rp

- 35 cm Fig. 14. Molecular beam apparatus (schematic). 1 = reflector box, 2 = liquid nitrogen cooled surface, 3 = reflector, 4 = collimating slits, 5 = oven, 6 = shutter, 7 = detector filament.

Fig. 15. Cosine law.

escaping f l o w t h r o u g h a hole or slit obeys the laws of f r e e molecular f l o w . A second orifice or collimating slit selects f r o m the escaping molecules those whose velocity vectors lie within t h e solid angle determined by these orifices; this angle is usually so small t h a t t h e motion of t h e beam molecules can be considered as unidirectional. In t h e experiments of interest to this discussion, t h e beam molecules are directed t o w a r d a scattering surface, and their spatial or velocity distribution is analysed a f t e r w a r d s . If t h e spatial distribution is diffuse, i. e., obeys t h e cosine law, according to which the intensity scattered in a direction N O / — N02 +

h v

f o r which t h e O atoms are p r o d u c e d by 0 2 dissociation in t h e discharge tube a n d t h e N O molecules b y collisions between these atoms a n d N atoms, which a r e also p r o d u c e d in t h e discharge. T h e glow is enhanced if N O is introduced upstream of t h e nozzle as shown in Fig. 2. If a plasmajet is used, no f u r t h e r excitation is required since t h e gas stream is hot enough to become luminescent by t e m p e r a t u r e excitation. A t t e m p t s to visualize l o w density f l o w p a t t e r n s t h r o u g h light absorption in the U . V. h a v e met w i t h only limited success.

128

I. Estermann, Simulation of high altitude flight conditions

8.2. F o r c e Under conditions existing in most low density wind tunnels, aerodynamic pressures are of the order of g/cm 2 and their measurement requires more sensitive balances than are usually employed. If one wants to reach the free

Jahrbuch 1961 der W G L

molecular flow regime, targets should have dimensions of the order of I, i. e., approximately 1 mm or less. As a result, forces of the order of milligrams will have to be measured. A small torsion balance capable of the proper sensitivity was used in Berkeley f o r this purpose. It consisted of a tungsten torsion wire supported by a movable f r a m e which carried a quartz fiber to which the target plate was attached. The free end of this fiber served as a pointer f o r indicating the angular twist of the wire. The support fiber was protected f r o m the air stream by a shield which also restricted its motion to a small deviation f r o m the vertical. Forces exerted by the gas flow on the target plate were compensated by applying a twist on the torsion wire. The assembly of the balance is shown in Fig. 18, a representative calibration curve in Fig. 19. For the proper interpretation of the results, knowledge of the accommodation coefficients (see Section 7) is necessary. 8.3. D e n s i t y The most direct approach to the measurement of this quantity is the attenuation in beams of photons or particles according to the differential equation dl/dX

Fig. 18. Schematic of drag balance.

= — q ¡x X ,

where I is the intensity of the beam at the point X inside the gas stream, ¡j, the mass absorption coefficient, and p

Fig. 20. Schematic d r a w i n g of the electron beam a p p a r a t u s . A = supports, B = cathode, C = heat shield, D = leads, E = anode, F = defining orifice, G and H = aperture plates, 1 = electron collector.

0

20°

i0°

60° Angle of twist Fig. 19. N u l l balance calibration. Tungsten torsion fiber diameter = 0.00064 in. O increasing weights, 0 decreasing weights.

the local density. N o n e of the gases used in wind tunnels have a suitable absorption coefficient in the visible, and as mentioned before, attempts to use the absorption band of oxygen in the vacuum ultraviolet have only been partly successful. Better results have been obtained with electron beams of about 10 kV energy f o r which the mass absorption coefficient in air is 7.4 X 10 5 cm 2 /g. It has also been shown that the differential equation listed above can be integrated for the traverse of an axisymmetric air stream. In Fig. 20, we show a schematic drawing of the electron-beam apparatus, in Fig. 21 the results of a traverse along the axis of the supersonic air flow around a sphere f r o m a point upstream to the stagnation point. It can be seen that the electron current at the collector decreases sharply when the shock wave region is reached, then more slowly to the stagnation point of the model. O n e would expect this behavior since the beam must traverse an ever increasing thickness of relatively dense gas as it moves along the stagnation line. The beam intensity drops sharply to zero as the model intercepts the beam. It appears possible to use the attenuation of a soft X - r a y beam, f o r which

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I. E s t e r m a n n , Simulation of high altitude flight conditions

^ « i l 0 3 c m 2 / g , in a similar way. In another variation the intensity of the light emitted by the atoms which are excited by the electron beam is used for local density determinations.

0

88100

88200

86300

Axial 'position

881,00

129

placed with its axis perpendicular to a uniform stream of gas, is heated to an equilibrium temperature which is a function only of the Mach number, the stagnation temperature, and the specific heat of the gas. These equilibrium temperatures become insensitive to changes in flow velocity at Mach numbers of about 2; beyond this range, up to

88500

coordinates

iLast figure =0.001 inchI Fig. 21. 8.4.

D e n s i t y d i s t r i b u t i o n in f r o n t o f a s p h e r e .

Pressure

The main difference between static pressure measurement at normal and low densities lies less in pressure-measuring instruments, which differ more in their sensitivity than in other properties from those used at higher pressures. Among the factors to be considered are the response time of the measuring system, thermal transpiration, and adsorption and desorption of gases. A more fundamental difficulty affecting dynamic measurements arises from viscous effects which require the application of substantial corrections to impact pressure measurements. 8.5.Free

M o l e c u l e

Probes

These instruments have a limited value for measurements in the free molecular flow regime. They must be so designed that their characteristic dimensions are small compared to the mean free path in the gas flow. For the usual test section pressures of 1 to 100 /,