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AFRPL, MAR 15, 1971

THIS PAGE IS UNCLASSIFIED

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AD 500 578

NEW LIMITATION CHANGETO DISTRIBUTION STATEMENT - A

Approved for public release;

distribution is unlimited.

LIMITATION CODE: 1

FROM DISTRIBUTION STATEMENT - B

Distribution authorized to U.S.

Gov't. agencies only.

LIMITATION CODE: 3

AUTHORITY

AFRPL; FEB 5, 1986

THIS PAGE IS UNCLASSIFIED

REPRODUCTION QUALITY NOTICE--!,',

This document is the best quality available. The copy furnishedto DTIC contained pages that may have the following qualityproblems:

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,bNfLASSIFIED

PATENT SECRECY NOTICE

Material In this publication relating to

LAMINATED' CHAMBER COOLING MEANS

reveals subject matter contained In U. S. Patent Application SerialNo. 319,047 entitled "High Pressure Rocket and Cooling Means" whichhas been placed under a Secrecy Order issued by the Commissioner ofPatents. This Secrecy Order has been modified by a SECURITY REQUIRE-MENTS PERMIT.

A Secrecy Order prohibits publication or disclosure of the invention,or any material information with respect thereto. It is separate anddistinct, and has nothing to do with the classification of Governmentcontracts.

By statute, violation of a Secrecy Urder is punishable by a finenot to exceed $10,000 and/or imprisonment for not more than two years.

A SECURITY REQUIREMENTS PERMIT authorizes disclosure of the inventionor any material information with respect thereto, to the extent set forthby the security requirements of the Government contract which imposes thehighest security classification on the subject matter of the application,except that export is prohibited.

Disclosure of this invention or any material information with respectthereto is prohibited except by written consent of the Commissioner ofSPatents or as authorized by the permit.

The foregoing does not in any way lessen responsibility for thesecurity of the subject matter as imposed by any Government contractor the provisions of the existing laws relating to espionage andnational security.

STATLT9 P

n addition to ,..-. -- uirementc which must be met, thisdocu tont t•i '- :. 11 expCr' -::trols and eachtransmittal tý) , •,,o •~ ~ ~ A Vnl -•. S ,

be made OUIY w t-1 Pr.~ I` le! : t5v$.S' 3

UNCLASSIFIEI _ __

(UNLASSIFED TMli)

'AIR fORCE REUSABLE ROCKET ENGINE PROGRAM /XLR129.P-1 //

ENGINE PERFORMANCE (.). ( )

.. /F-GROUP 4

364 ,O

PATENT SECRECY NOTICE

PORTON IS OF THIS DOCUMENT CONTAIN SUJ.IECT HATTONCOVERD iMV A U.S. PATIENT OFFICE 92COOCv ORDoER WITHMODIFYING OCCURITY REQUIREMENTo PERMIT.HANDLING SINALL SE IN ACCORDANCE WITH THE PERMITAS DESCRISEGD ON PAGE A ANl' INDICATED HEREI19N. VIG.-I.ATIMIG MAY on SJUBJECT TO THE P.-ALTICS PRESCRIDEDBY, TITLE $S. U. a. 4. (less). sEcTIohj $111 AND 5A--5.

THIS DOCUMENT CONTAINS INFORMATION AFFIGI THINATIONAL DEaENSEw OF THE UNITED SfATyE WITHIN THEMEANING OTHE SIIlONAGE LAW6. TITLE IS U, a. .. .sEcTING 'I'S AND 714. ITS TRANSmISeION On THEIREVELATION OF ITS CONTENTS IN ANY MANNE ToANd UA•OAUINNGIDPER•SO ISn is m . AWN. .MY LAW.

s -/

UNCLASSIFIEDFOREWORD

This Technical Report presents the design and operational character-istics of the XLR129-P-1 reusable rocket engine. In addition, it presentsengine parametric performance, size, and weight data for future flightengine3 that could result from an engineering development program basedon this engine concept. This report is issued as a special report inaccordance with the requirements of Contract F04611-68-C-0002.

This publication was prepared by the Pratt & Whitney AircraftFlorida Research and Development Center as PWA FR-3108.

Classified information has been extracted from (asterisked) documentslisted under References.

This Technical Report has been reviewed and is approved.

Ernie D. BraunschweigMajor, USAFProgram ManagerAir Force Rocket Propulsion Laboratory

ii

UNCLASSIFIED

I-.

UNCLASSIFIEDUNCLASS IFIED ABSTRACT

This special technical report presents information and data on theXLRI29-P-L rocket engine. Information is presented for both the demon-strator engine and flight engine versions of thisi rocket engine. Ageneral description and pertinent technical information are presentedfor the demonstrator engine. The demonstrator engine program scheduleis also presented. Parametric design, performance, cost, and scheduledata are presented for the flight engine, This technical report hasbeen prepared for the use of airframe manufacturers and governmentpersonnel who are conducting mission and vehicle studies.

hil/iv

UNCLASSIFIED

UNCLASSIFIED

CONTENTS

SECTION PAGE

ILLUSTRATIONS ....... .................. .... vi

TABLES ................ ....................... x

LIST OF SYMBOLS AND ABBREVIATIONS ..... ......... xi

I INTRODUCTION ........ .................... . .. I

II SUMMARY ..................... ...................... 3

III DEMONSTRATOR ENGINE ............... ................ 5

A. Engine Characteristics ....... ............. 5B. Engine Installation Drawing ...... .......... 6C. Engine Flow Schematic ........ ............. 6D. Engine and Component Operating Parameters . . . 14E. Summary Weight Table .................. .... 20F. Start and Shutdown Curve ................ .... 20G. Engine Inlet Condition Curves ......... .... 21H. Program Schedule ..................... .... 21

IV FLIGHT ENGINE ...... ................... ..... 25

A. IntroductLon ....... .................. .... 25B. Description ...... ................. ..... 25C. Parametric Engine Data ................ ..... 32D. Engine/Vehicle Interface Data ..... ......... 59

v

UNCLASSIFIED

CONRDENTIALILLUSTRATIONS

FIGURE PAGE

1 XLR129-P-L Demonstrator Engine ....... ........... 7

2 Engine Installation Drawing ........ ............ 9

3 Simplified XLR129-P-1 Demonstrator EnginePropellant Flow Schematic .... ............. .... 13

4 Detailed XLRI29-P-l Demonstrator EnginePropellant Flow Schematic .... ............ .... 15

5 Estimated Start and Shutdown TransientCharacteristics .......... ................ .... 21

6 Fuel Inlet Operating Regior . ..... .......... .... 22

7 Oxidizer Inlet Operating Region ............ ... 22

8 Propellant Temperature Limits for FullTrim Capability ........ .................. .... 23

9 XLR129-P-l Demonstrator Engine ProgramSchedule ............. ... ...................... 23

10 Engine Configuration With Two-Position Nozzle . . . 26

I1 Perfect Nozzle Contours .... .............. .... 27

12 Perfect Nozzle Contours .... .............. .... 28

13 Nozzle Contour ....... ................... .... 29

14 Contour Optimization ..... ................ .... 30

15 Base Truncations of Perfect Nozzles ........ ... 30

16 XLR129-P-1 Demonstrator Engine Program ........... 35

17 Estimated Engine Development Costs, Oxygen/Hydrogen Engines ....... .................. .... 35

18 Vacuum Specific Impulse vs Nozzle ExpansionRatio (Maximum Performance Nozzle Contour,Fixed and Two-Position Nozzle EngineConfigurations) ........ .................. .... 37

19 Vacuum Specific Impulse vs Nozzle ExpansionRatio (Base Nozzle Contour, Fixed and Two-Position Nozzle Engine Configurations) .... ....... 37

20 Vacuum Specific Impulse vs Nozzle ExpansionRatio (Minimum Surface Area Nozzle Contour,Fixed and Two-Position Nozzle Engine Con-figurations) ....... .................... .... 38

21 Altitude Performance (Maximum PerformanceM!ozzle Contour, Two-Position Nozzle EngineConfiguration (Ep - 35)) ..................... .... 39

viCOHFIUEHTIAL

IL' ,USTRATIONS (Contnued)

FIGURE PAGE

22 Altitude Performance (Base Nozzle Contour,Two-Positlon Nozzle Engine Configuration

(.p a 35)) ........... ..................... ..... 39

23 Altitude Performance (MInimum Surface AreaNozzle Contour, Two-Position Nozzle EngineConfiguration (ep - 35)) .... .............. ... 40

24 Altitude Performance (Maximum PerformanceNozzle Contour, Two-Position Nozzle EngineConfiguration (ep * Minimum)) ............. .... 40

25 Altitude Performance (Base Nozzle Contour,Two-Position Nozzle Engine Configuration(tp a Minimum)) ........ .................. .... 41

26 Altitude Performance (Minimum Surface AreaNozzle Contour, Two-Position Nozzle EngineConfiguration (tp a Minimum)) ............. .... 41

27 Altitude Performance (Fixed Nozzle EngineConfiguration) ....... ................... .... 43

28 Sea Level Specific Impulse vs Nozzle ExpansionRatio (Maximum Performance Nozzle Contour,Fixed and Two-Position Nozzle EngineConfigurations) ........ .................. .... 44

29 Sea Level Specific Impulse vs Nozzle ExpansionRatio (Base Nozzle Contour, Fixed and Two-Position Nozzle Engine Configurations) .... ....... 44

3J Sea Level Specific Impulse vs Nozzle ExpansionRatio (Minimum Surface Area Nozzle Contour,Fixed and Two-Position Nozzle EngineConfigurations) ........ .................. .... 45

31 Vacuum Specific Impulse for ThrottledConditions ............... ..................... 46

32 Total Engine Weight vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, Two-Position Nozzle Engine Configuration (4p - 35)) 47

33 Total Engine Weight vs Nozzle Expansion Ratio(Base Nozzle Contour, Two-Position NozzleEngine Configuration (0p a 35)) ...... .......... 48

34 Total Engine Weight vs Nozzle Expansion Ratio(Minimum Surface Area Nozzle Contour, Two-Position Nozzle Engine Configuration (ep - 35)) . 48

35 Tr':al Engine Weight vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, Two-Position Nozzle Engine ConfigurationUp a Minimum)) ...... .................. .... 49

viigi € ONF1DONIA.M

CONIDENTIALILLUSTRATIONS (Continued)

FIGURE PAGE

36 Total Engine Weight vs Nozzle Expansion Ratio(Base Nozzle Contour, Two-Position NnzzleEngine Configurations (ep a Minimum)) ....... ... 49

37 Total Engine Weight vs Nozzle Expansion Ratio(Minimum Surface Area Nozzle Contour, Two-Position Nozzle Engine Configuration(tp a Minimum)) ........ .................. .... 50

38 Total Engine Weight vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, FixedNozzle Engine Configuration) ............... .... 50

39 Total Engine Weight vs Nozzle Expansion Ratio(Base Nozzle Contour, Fixed Nozzle EngineConfiguration) ....... ................... .... 51

40 Total Engine Weight vs Nozzle Expansion Ratio(Minimum Surface Area Nozzle Contour, FixedNozzle Engine Configuration) ............... .... 51

41 Overall xit Dlaifamer vs Nozzle .... u RaLlu 53

42 Overall Length vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, Fixedand Two-Position Nozzle Configurations) .... ...... 53

43 Overall Length vs Nozzle Expansion Ratio(Base Nozzle Contour, Fixed and Two-PositionNozzle Configurations) .... ............... ..... 54

44 Overall Length vs Nozzle Expansion Ratio(Minimum Surface Area Nozzle Contour, Fixedand Two-Position Nozzle Configurations) .... ...... 54

45 Stowed Length vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, Two-Position Nozzle Engine Configuration (ep a 35)) . . 55

46 Stowed Length vs Nozzle Expansion Ratio(Base Nozzle Contour, Two-Position NozzleEngine Configuration (ep - 35)) ...... .......... 55

47 Stowed Length vs Nozzle Expansion Ratio I(Minimum Surface Area Nozzle Conto.r, Two-Position Nozzle Engine Configuration (Op a 35)) . 56

48 Minimum Stowed Length vs Nozzle Expansion Ratio(Maximum Performance Nozzle Contour, Two-Position Nozzle Engine Configuration(9p M Minimum)) ........ .................. .... 56

49 Minimum Stowed Length vs Nozzle Expansion Ratio(Base Nozzle Contour Two-Position Nozzle EngineConfiguration (fp - Minimum)) ............. .... 57

CONFIDENTIAL

UNCLASSIFIEDILLUSTRATIONS (Continued)

FIGURE PAGE

so Minimum Stowed Length vs Nozzle Expanvion Ratio(Minimum Surface Area Noz,.le Contour, Two-Positicn Nozzle Engine Configuration(ep a Minimum)) ........ .................. .... 57

51 Primary Nozzle Expansion Ratio vs NozzleExpansion Ratio (Maximum Performance NozzleContour, Two-Position Nozzle Engine Configura-tion (ep - Minimum)). . .. ................ .... 58

52 Primary Nozzle Expansion Ratio vs NozzleExpansion Ratio (Base Nozzle Contour, Two-Position Nozzle Engine Configuration(OP a Minimum)) .............. .................. 58

53 Primary Nozzle Expansion Ratio vs NozzleExpansion Ratio (Minimum Surface Area NozzleContour Two-Position Nozzle Engine Configura-tion (ep a Minimum)) ..... ................ .... 59

54 Engine/Vehicle Interface Locations ..... ......... 60

ix

UNCLASSIFIED

I-

UNCLASSIFIEDTABLES

TABLE PAGE

I Demonstrator Engine Characteristics.........

II XLR129-P-1 Demonstrator Engine OperatingCharacteristics, Booster .... .............. .... 16

III Enigine Weights (Dry) ........... ................ 20

IV Flight Engine Characteristics .... ............. 31

V Estimated Flight Engine Operating Character-istics Upper Stage: Nozzle Extended ........ ... 33

VI Location of Engine Inlets and ActuatorArm Attach Points ..... ................. .... 60

VII Power Package Diameter, Oxidizer and FuelInlet Diameters ...... .................. .... 61

VIII Gimbal Loads and Auxiliary Power Availablefrom the Engine ...... .................. .... 62

x

UNCLASSIFIED

UNCLASSIFIEDLIST OF ýYMBOLS AND ABBREVIATIONS

ADP Accessory Drive Pad

F,. Thrust at Altitude

Fvac Thrust at Vacuum

FVM Fuel Vent Manifold

2 Gravitational Constant

Is Specific Impulse

Isalt Specific Impulse at Altitude

Ivac Specific Impulse at Vacuum

L/Dr Length/Throat Diameter

MCi Maximum Performance Nozzle

NSA Minimum Surface Area

NPSH Net Positive Suction Head

OVBD Overboard

OVM Oxidizer Vent Manifold

PFRT Preliminary Flight Rating Test

r Mixture Ratio

TBO Time Between Overhaul

TVC Thru-t Vector Control

* Nozzle Expansion Ratio

op Primary Nozzle Expanuion Ratio

xLIUNCUASSIFIED

This Document

Reproduced From

SUCTON I .IN1RODUCfION

CONFIDENTIAL.iCTION I

INTRODUCTION

(U) This technical report provides a general description, pertinenttechnical information, and the program schedule for thp XLRI29-P-ldemonstrator engine being designed, fabricated, and tested underContract F04611-68-C-0002. In addition, parametric data on the per-formance, weight, and size of reusable rocket engines which could resultfrom an engineering development program is provided for use in vehicleand mission studies.

(C) Data are presented for high pressure (3000-psia chamber pressure)staged-combustion, pump-fed, oxygen-hydrogen, engines with transpiration-cooled thrust chambers and regeneratively and dump-cooled nozzles.

(U) By combining interchangeable nozzle extensions with a basic turbo-pump and combustion chamber module, the engine can be tailored tospecific rocket stage requirements. Data are provided for conventionalfixed nozzle engine configurations and two-position nozzle enginc con-figurations for various nozzle expansion ratios and contours. The two-position nozzle concept is based on part of the nozzle being retractedover the forward portion of the thrust chamber during low altitude opera-tion and extended to the high area ratio position for high altitude opera-tion. This principle enables the high pressure engine to have an exhaustnozzle that is optimized for both high altitude and low altitude operation.In addition, the two-position nozzle will provide high specific impulsein vehicle installations where the length of the engine is limited. Dumpcooling is used for the extendible portion of the nozzle, which alsopermits lightweight nczzle construction.

1/2

CONFIDENTIAL

*U'� t. � - - -- - .-- '----- ".-.,'--'- -

I.

I

ECTION USUMMARY

I

* -

'I

�K.

�1p * i

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I,

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UNCLASSIFIEDSECTION II

SL hMORY

(U) Information prc.ented on the XLR129-P-1 demonstrator engine include:the overall engine design characteristics, an installation drawing, aflow schematic, the engine internal design parameters, the componentand engine weight, a schedule of the demonstrator engine program, startand shutdown thrust versus time curves, and engine/vehicle interfacepropellant requirements.

(U) The parametric performance data, which Are presented for theLRl29-P-1 flight rocket engine, include delivered vacuum specific im-pulse versus nozzle expansion area ratio for three nozzle contours andfive mixture ratios; specific impulse versus altitude for fixed nozzleand two types of two-position nozzle engine configurations. In addi-tion, parametric data are presented for engine weight, diameter, andlength versus nozzle expansion ratio for three nozzle contours. Alsoincluded is a throttling curve that presents delivered vacuum specificimpulse at various throttling conditions for five different mixtureratios.

3/4

UNCLASSIFIED

J

*

UCUONE9DEMONSTRATOR ENGINE

Page

A. Ingsne Characteristics. . . . . .. .. 0 "I. Ingine Inetallation Drawing . . . . . . 6C. XngLne flow Schematic . * . . . ... a. 6D. gngine and Component Operating

Paramer. e.. . . .... . 14B. Sumary Weigh Table .. ....... 201. Start and Shutdown Curve. . . . . . . . 20G. ngsine Inlet Condition Curves ..... 21H. Progrm Schedule. . . . . . ... . 21

.JI'.

p. I<I

I ,

It.. -', . . ..... . , ,

",LONFIDENTIAL

SECTION IIIDEMONSTRATOR ENGINE

A. ENGINE CHARACTERISTICS

(U) The design and demonstration characteristics for the XLR129-P-ldemonstratcr engine are shown in Table I.

(C)(U) Table I. Demonstrator Engine Characteristics

Nominal Thrust 250,000-lb vacuum thrust with area ratio of 166:1244,000-lb vacuum thrust with area ratio of 75:1209,000-lb sea ievel thrust with area ratio of 35:1

Minimum Delivered 96% of theoretical shifting I1 at nominal thrust;Specific Impulse 94% of theoretical shifting Is during throttlingEfficiency

Throttling Range Continuous from 100 to 207. of nominal thrustover the mixture ratio range

Overall Mixture Engine operation from 5.0:1 to 7.0:1Ratio Range

Rated Chamber 2740 psiaPressure

Engine Weight 3520 lb (with flight-type actuators and engine(with 75:1 nozzle) command unit)

3380 lb (less flight-type actuators and enginecommand unit)

Expansion Ratio Two-position booster-type nozzle with area ratiosof 35:1 and 75:1

Durability 10 hours time between overhauls, 100 reuses,300 starts, 300 thermal cycles, 10,000 valvecyc les

Single Continuous Capability from 10 seconds to 600 secondsRun Duration

Engine Starts Multiple restart at swa level or altitude

Thrust Vector Amplitude: ±7 deg;Control Rate: 30 deg/sec;

Acceleration: 30 rad/se€2

Control Capability ±3% accuracy in thrust and mixture ratio at nominalthrust. Excursions from extreme to extreme inthrust and mixture ratio within 5 seconds.

CONFIDENTIAL(C)(U) Table I. Demonstrator Engine Characteristics

(Continued)

Propellant L02 : 16 ft NPSH from 1. atmosphere boilingConditions temperature to 180*R

LH2 : 60 ft NPSH from I atmosphere boilingtemperature to 45*R

Environmental Sea level to vacuum conditionsConditions Combined acceleration: 10 g's axial

with 2 g's transverse, 6.5 g's axialwith 3 g's transverse, 3 g's axialwith 6 g's transverse

Engine/Vehicle The engine will receive no external power,with the exception of normal electrical powerand 1500-psia helium from the vehicle

B. ENGINE INSTALLATION DRAWING

(U) The general component arrangement of the XLRi29-P-l demonstratorengine is illustrated in Figure 1. An installation drawing with envelopedimensions, including the retracted (stowed) length and extended lengthof the engine, is provided as Figure 2. The primary interfaces, such asthe propellant inlet connections and gimbal attachment locations, arealso ahown. It La cati...tcd that tho mmaxium actuar loaduringimbaling of the demonstrator engine is 50,000 lb. The power to gimbalthe engine is approximately 80 horsepower.

C. ENGINE FLOW SCHEMATIC

(U) A simplified propellant flow schematic illustrating the propellantflow paths and functional component arrangement of the engine is shownin Figure 3. The XLR129-P-1 high pressure rocket engine uses a stagedcombustion cycle in which most of the fuel is burned with a portion ofthe oxidizer in the preburner to provide turbopump power before combus-tion with the remainder of the oxidizer in the main burner chamber.Fuel and oxidizer enter the engine through the engine driven low-speedinducers. The low-speed inducers are used to minimize the fuel tankpressure requirements, while allowing high-speed main propellant pumpsfor high turbopump efficiencies. The fuel low-speed inducer is a singleshaft unit with an axial flow inducer driven by a two-stage, axial-flow,partial-admission impulse turbine. The oxidizer low-speed inducer isalso a single shaft unit with an axial flow inducer driven by a singlestage radial inflow turbine.

6

CONFIDENTIAL

CONFIDENTIAL

CONFIDENTIAL

(U) Figure 1. XLR129-P-1 Demonstrator Engine FOC 27532

7/8 NPIGsfiaMT N O 116 As alvIEEPIIPN 0fl 0mai WIT: A 040IOIP~IM -. 66CamiTV ffloIJIuSeNvU PuMIYrCONFIDENT IA S `69" w.S. CUMNIPSSIOABU or PAGNS.

- - _ . . . I IIII t I II .... I 11 l .- _- .. . . _ . . .

UNCLASSIFI[I

Oxidizer Inlet 1.D"10.8 ia Fuel Inlet-

• •.4 5.4

Gimbal Points 90 dog

Apart 12 dog Off AxisActuator Action Lines

(U) Figure 2. Engine Installation Drawing

UNCLASSI

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FT 27724

9/10

ICLASSIRiEDL

UNCLASSIFIED

MU Figure 2. Engine Installation Drawing (Concluded)

UNCLASSIFIE~

ICLASSIFIED

, 41

NoTle Retaate

FD 27725

11/12

CLASSIFIED

"CONFIDENTIAL

(U) Figure 3. Simplified XLR129-P-l Demon- FD 21002Cstrator Engine Propellant FlowSchematic

(U) The main fuel turbopump is a single shaft unit with two back-to-backcentrifugal pump stages driven by a two-stage, pressure-compounded tur-bine. The £ful flow is pu1tmped Lu WLet ,ysLen upetitLiig piessure levt l1by the main fuel pump. The hydrogen is then ducted to cool the regenera-tive sections of the nozzles. The principal fuel flow path from the pumpis through the upstream portion of the primary nozzle, and then into thepreburner chamber through the preburner injector. The remainder of thefuel flows through the downstream portion of the primary nozzle and thenthrough the fuel low-speed inducer drive turbine prior to being passedinto the main chamber as transpiration coolant. A small amounz .,f fuelis bled off between the main fuel pump stages to provide coolant torthe two-position nozzle. This coolant flows to the nozzle through aregulating orifice and a shutoff valve that is provided to stop the flowwhen the two-position nozzle is in the retracted position. The arearatio, at which the fixed primary nozzle ends ard the two-position,translating nozzle starts, is varied to optimize the performance for thespecific application.

(U) The oxidizer turbopump is a single shaft unit with a single, centrif-ugal pump stage driven by a two-stage, pressure-compounded turbine.After being pumped to the system operating pressure levels by the mainoxidizer pump, the oxidizer is divided between the preburner and themain burner chamber. The principal oxidizer flow passes through and isthe working fluid for the oxidizer low-speed itiducer turbine before beinginjected into the main chamber. The remainder, a smaller scheduledportion of the oxidizer, is ducted to the preburner where it is burnedwith the fuel. The resulting combustion products flow through ducts to

13

CONFIDENTIAL(This page is Unclassified)

S " ado

CONFIDENTIALthe two main turbines, which are arranged in parallel. Tile energyrequired to drive the main pumps is extracted from these combustionproducts, which then exhaust from the turbines and mix in a commonpassage of the transition case. These gases then pass through the mainbuv'ner injector and into the main burner combustion chamber where theymix and burn with the principal oxidizer flow. These combustion gasesare then expanded through the bell nozzle to provide thrust.

(U) Tile preburner injector consists of dual-orifice, tangentital-swirieroxidizer injector elements and fixed area fuel injector elements. Apreburner oxidizer valve is incorporated at the inlet to the injectorassembly to vary the total oxidizer flow rate for turbine inlet tempera-

ture control and to adjust the relative flow of the primary and secondar>oxidizer elements. Tile preburner combustion chamber is contained within

the transition case, which also contains the turbine drive gas ducts.Tile main turbopunmps are attached to the transition case with a singleflange and bolt circle arrangement to provide ease of access for turbo-pump maintenance.

(C) Tile main chamber injector consists of fixed-area, tangential1'swirleroxygen injection elements arranged in radial spraybars The fuel side(preburner gas after expansion through the turbine) is a fixed area design

that ducts fuel-rich gas around each row of oxidizer injecter points. A

small portion of the fuel-rich gas flows through a porous face to providecooling. The combustion chamber wall is composed of a fuel cooled liner

extending from the injector face through the throat region to a nozzlearea ratio of 5.3. The liner is composed of porous wafer plates, whichprovides the transpiration cooling.

(U) The nozzle, which attaches downstream of the throat, is composed oftwo regeneratively cooled primary sections and a low-pressure, dump-cooled,two-position nozzle. The regeneratively cooled sections are conventionaltubular heat exchangers, the two-position nozzle employs lightweight sheetmetal construction.

(U) A more complete system schematic, including all main propellant lines,recirculation lines, electrical Interconnections, and the helium systemsis shown in Figure 4.

D. ENGINE AND COMPONENT OPERATING PARLAMETERS

(C) The component and engine system steady-state operating parametersare presented in Table II for mixture ratios of 5, 6; and 7 at 100% and

20% thrust.

14

CONFIDENTIAL

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19

CONFIDENTIAL

CONFIDENTIALE. SM4ARY WEIGHT TABLE

(U) The estimated dry weights of the major components of the demonstratorand flight engines are shown in Table III. The demonstrator dry engineweights are based upon lightweight component designs with the additionalmargins required for a low risk demonstrator engine program. The flightengine dry weights are based upon the normal improvements and weightreduction that would materialize as a result of an engineering develop-ment program.

(C)(U) Table III. Engine Weights (Dry)

Item Demonstrator FlightEngine (0 - 75) Engine (f u 75)

Preburner and Hardware 90 70Transition Case and Gimbal 370 285Main Burner Injector and Hardware 115 85Main Burner Chamber 425 330Nozzle and Actuation 640 460Fuel Turbopump 480 380Oxidizer Turbopump '135 250Low-Speed Inducers i35 185Controls 305 240Plumbing 310 240Miscellaneous 75 55

Total 3380& 2580*Does not include valve actuators

F. START AND SHUTPOWN CURVE

(C) The start and shutdown transient curve from the XLRI29-P-1 demon-strator engine is shown in Figure 5. This curve presents percent thrustversus time, and shows four distinct modes of operations (1) start tominimum thrust in approximately 1 second, (2) accelerations to 100%thrust, (3) decelerations to idle, and (4) shutdowns.

2C

CONFIDENIAL

CONFIDENTIAL

too

M

(U) Figure 5. Estimated Start and Shutdown DFC 69931

Transient Characteristics

G. ENGINE INLET CONDITION CURVES

(U) The ranges of temperature, pressure and NPSH conditions requiredat the inlet to the fual and oxidizer low-speed inducers are shown inFigures 6 and 7.

(U) The relationship required between fuel temperature and oxidizer tem-perature, so that the engine thrust and mixture ratio will remain attheir set points withii the specified control accuracy (t3%4), is shownin Figure 8.

H. PROGRAM SCHFDULE

(U) The XLRI29-P-1 demonstrator engine program schedule is shown inFigure 9. This is a 54-month program that began on 6 November 1967.The program has been divided into five phases. The first phase, whichhas already been completed, generated test and analytical data to com-plete the technology necessary to design the engine and components.During the second phase all the components will be designed. The com-ponents will be fabricated and tested to qualify them for engine useduring the third phase. The fourth phase is the integration of thecomponents into the demonstrator engine and the testing of the demon-strator engine. The fifth phase is the definition of the flight engine.

21

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60

0%

N 0%

N aW4 19:1VNI LL OU 131

22

CONIOERA

UNCLASSIFIED

(U) Figure 8. Propellant Temperature Limits DF 70473for Full Trim Capability

F (Supporting Test)

Component Development

IEngineTs

Flight Engine 'udinition

I t . , i. . . i , I ... I ,

I Nov1967 1968 1969 1970 1971 1972

CALENDAR YEARS

(U) Figure 9. XLRI29-P-l Demonstrator Engine FD 27857Program Schedule

23/24UNCLASSIFIED

S*', ON IV

FIOHT ENGINE epaeE

A. Introduction . . 25B. Description ..... ................ * . . 25

1. General . . . . . .. .. .. .. . ... . 252. Noszle Concepts . . ... . ........... . 25

a. Two-Position Nozzle Conceptr ...... 25b. Nozile Co:•tour. ... . . . . . . . . 27

3. Flight Engine Characteristics ......... ... 314. Cycle Balance ............... 325. Schedule With Cost ............. . . . 32

C. Parametric Engine Data .............. 32

1. General . . . . . . . . . . . .. 322. Engine Performance. ............. 36

a. General ........... .......... ... 36b. Vacuum Performance . ..... ... . . . 36c. Altitude Performance. ............ .... 38

(1) General ........... 38(2) Tvo-Position Nozzle Engine

Configurations ............. .... 38(3) Fixed Nozzle Engine

Configurations. ........... 42

d. Sea Level Prrformance ..... ......... 42

(1) General ............... . . . 42(2) Two-Position and Fixed

Nozzle Engine Con-"figurations..... .... . ...... . 42

e. Throttled Performance . ........ 45

3. Engine Weight. . . . . . . . . ....... 46a. General ................ 46b. Two-Position Nostle and Fixed

Nozsle Engine Configurations ...... .... 47

4. Engine Envelope . . . . . . . . . . . . . . 52

a. General . . . . . . . . . . . . . . . . 52b.* Diameter. . . . . . . . . . . . . . . 52c. Length. ........ ............... 52

(1) Overall Lengt........... 52(2) Stowed sand MinL m Stave Length. 52

D. Zngine/Vehtcle Interface Data ........ 59

to General . 0 0 . . 0 592. Engine Inlet Conditions.u.... 593. InetgineVeicle Interface Locations. 594. Inlet Line and Power Package Diameters. 61

*-5. Gimbal loaods. .* , a o * o - - # & .# @ 616. huiiliary lower .............. 61

SECTION IV

FLIGHT ENG)JNE

A. INTRODUCTION

(U) The flight engine configuration will be based on the results of theXLRI29-P-1 demonstrator eg.,ine program. During a subsequent engineeringdevelopment program further tests would be conducted to refine the engineto meet the flightweight criteria and to develop the level of systemmaturity required for fIight operation. The additional developmenteffort is a logical extension of the demonstrator program.

B. DESCRIPTION

I. General

(C) The staged-combustion, high-pressure oxygen/hydrogen rocket engineis a versatile, high performance propulsion system for use in both upperand lower stages of advanced vehiLles. This reusable engine is capableof maintaining constant thrust over a mixture ratio range of 5 to 7.Nozzle interchangeability and the use of the two-position nozzle con-cept permit operation of the same engine system with optimum nozzlearea ratios for improving the performance of the lower stages withinthe atmosphere as well as providing the high performance attainable withvery high area ratio nozzles in the upper stages. Interchangeabilityis achieved by attaching the desired nozzle to a fixed turbomachine-vpower package. The area ratio at which the primary (fixed) nozzle endsand the two position (translating) skirt starts, can be varied toopr it ptriforanro, for tht ip - app)-_I- it o-n, The generml pr.-lliantflow schematic of the LR129-P-l flight engine is the same as shown forthe demonstrator engine in Figure 3. Engine nomenclature is illustratedin Figure 10.

2. Nozzle Concepts

a. Two-Position Nozzle Concept

(U) The two-position nozzle concept consists of a translating two-posi-tion nozzle and a fixed primary bell nozzle. The two-position nozzle isin the retracted position for sea level operation thus eliminating over-expansion losses associated wiLh the larger area ratio section of thenozzle. This concept, combined with high chamber pressure, provides acompact engine package that provides superior low altitude performance.The two-position nozzle is translated to the extended position for highaltitude operation. On upper stage applications, the two-positionnozzle is in the retracted position before stage separation to providea compact engine package, and is then extended after staging.

S' ' ,!t5

CONFIDENtiAL""Envelop*Power packape,

Diameter Gimbal Asia

Two- Poit ion

Nozzle

OverallEn Kine I. -Two- PositionLength Ntal*PLegh I •/ Nozzle

II

/V

(U) Figure 10. Engine Configuration With FU 2,L152ATwo-Position Nozzle

(C) Data for two types of two-position nozzles are presented in thisreport; (1) a configuration with a 35,1 primary nozzle expansion ratioand various overall expansion ratios and (2) alternatives which produceminimum stowed length of the engine with the nozzle retracted.

(C) In general, lower stage engines, which operate both at low-altitudeand in vacuum, will have a relatively low primary expansion ratio.Pratt & Whitney Aircraft studies have shown the 35:1 primary expansionratio to be near optimum. These studies have also shown thaL overall(extended) nozzle expansion ratios for these applications are also some-what low and in the 50:1 to 150:1 range. High efficiency contoursusually produ4e the best performance in these applications.

(C) When the primary nozzle expansion ratio Ep (or the breakpoint be-tween the stationary and the movrible portions oi the nozzle), is in the35:1 to 80:1 range, it is possible to move the nozzle far enough forwardto provide the altitude compensat.ng feature. The two-position nozzle,however, cannot be moved completoly forward to the gimbal axis becauseof interference with the turbopumps and plumbing. Where short lengthand nigh oxpandion ratios arc primary installation goals (e.g., upperstage), it is generally desirable to accomplish the breakpoint at anarea ratio greater than 80:1 so that the translating portion of thenozzle can be moved back as close as poss~ble to the engine gimbal axis.

26

CONFIDENTIAL

CONFIDENTIALBecause In upper stage applications the engine would not generally betLred with the nozzle retracted, the breakpoint could be selected to pro-vide the shortest possible engine stowed length and the maximum engineexpansion ratio. The engirne stowed length curves presented in thisreport reflect ttese considerations.

b. Nozzle Contour

(W) Engine length, weight, and performance aru functions of the exhaustnozzle contour. In general, shorter length contoirs yield lower nozzleperformance. Therefore, in addition to optimization of the area ratio,the optLmiza•ion of a bell nozzle engine includes the selection of theshape or contour of the nozzle. The bell nozzle contours used by Pratt &Whitney Aircraft are selected from a family of truncated perfect nozzles.Perfect nozzles are defined as those that, at a prescribed area ratio,expand a gas flow from the throat of the nozzle to a uniform and parallelflow at the nozzle exit. Using tho method of characteristics, a seriesof perfect nozzle contours may be computed as a function of this "design"area ratio. The contour surface at any diameter and length along thenozzle may be plotted in nordimensional form as shown in Figure II. Theintegrated thrust ard surface area can also be calculated at axial loca-tions along the nozzle. The calculation procedure includes the eff.oct offriction and varying thermodynamic properties of the reacting gases.Representative results of this detailed analysis can be plotted as shownin Figure 12, which presents contour coordinates for perfect bell nozzleswith lines of constant surface area and constant vacuum thrust coefficientssuperimposed.

2,H 2.0OTP,• = 3000 piai

ra6

K10

0 ••0 10 2D 30 40LENGTH/THMOAT DAMETER~

(U) Figure 11. Perfect Nozzle Contours FDC 27862

27CONFIDENTIAL

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'[UUZVC LUHUsHWI 'a28-CONF*DOM A

CONFIDENTIAL(W) • perfect nozz.le contour does not necessarily produce optimum engineor vehicle performance. because a perfect nozzle is constrained to pro-uue completely axial flow at the exit, a considerable part of the rearsection of the nozzle is involved in the final flow turning process.In a real nozzle, the friction losses here are greater than the perform-ance gains that accrue from the final flow srralghtening. Therefore,rlaxirmum nozzle pertormance is obtained by shortening or truncating tileperftecL nozzle. Further, in real vehicles there are engine longthpenalties such that additi•nal truncation may be required to producemaximum vehicle performance. Pratt & Whitney Aircr,.ft '-s definada series of three nozzle length truncations of the ab,. 'efined per-foct nozzle for application to real vehicles. These t,Ar.catiuns(referred to as nozzle contours in this report) are deocribed asfollows:

1. Maximum Performance Nozzle (NCs)

2. MInimum Surfact. Area Nozzle (MSA)

3, Base Nozzle.

These three nozzle truncations are shown schematically in Figure 13.

S................. ..... .n..........

•-• Lengh

(U) Figure 13. Nozzle Contour FD 22438A

(C) So that the method used in establishing MC., MSA, and Base Nozzlecontours can be more easily identified, a representative portion of theInformation given in Figure 12 is shown in Figure 14.

(U) Nozzles with minimum surface area (MSA nozzles) for a given thrustare defined by the locus of points for which a line of constant thrustcoefficient (CFvac) and a line of constant surface area-to-throat arearatio (As/At) are tangent. This is shown as point A on Figure 14.Maximum nozzle efficiency (NCs) for a given thrust Are defined by thelocus of points for which the lines of constant thrust coefficient(CFvac) have zero slope (point B on Figure 14). It should be notedthrt the maximum performance nozzle (point B) is still short of thefull length perfect nozzle because frictional drag has been included.

29

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Constant Thrust

0Constant Swface

LENGTH/THROAT DIAMETER, L/D1

(U) Figure 14. Contour Optiamization FD 6263C

(U) A third type of nozzle truncation considered is referred to as abase nozzle contour. This truncation has resulted from experience withvarious optimization studies and generally produces nearly optimumbalance of weight and performance, particularly for lower stage applica-tions. While this contour is not established directly from analysis ofFigure 14, It falls approximately half-way between points A (MSA) and"B (.CS). The relationship between trtincated area ratio and "design"area ratio for base nozzle contours is shown in Figure 15.

1000-IH2-Oh Propellant.

Pc - 3000 psimo • Equilibrium Flow

Z -C2

10 100 1000TRUNCATED AREA RATIO

(U) Figure 15. Base Truncations of Perfect FDC 27861Nozzles

30

CONFIDENIA

CONFIDENTIA.(L) Nozzle truncations considered in the p,-rametric data of this reportare: maximum performance, base, and minimum eurface area nozzle contours.Pratt & Whitney Aircraft preliminary engine/vehicle sLudies have shown

that the best upper stage performance is generally obtained by using thebase nozzle or minimum surface area nozzle contour.

3. Flight Engine Chacacrtristics

(L') Preltminary design characteristics for the flight engine are pro-vided in Table IV.

(C)(U) Table IV. Flight Engine Characteristics

Nominal Thrust To be determined from parametrLc dati

Minimum Delivered LO0. nominal thrust - 96.77.Specific Impulse 20:'. nominal thrust 95.4k,Efficiency (All of these conditions art, wiLh an expansion

ratio of 100:1, mixture ratio of 7 and ISA nozzle)

Throttling Range Continuous from 100 to 20'X of nominal thrust overthe mixture ratio ranSe

Overall Mixture Engine operation from 5.0:1 to 7.0:1Patio Range

Rated Chamber 3000 psiaPressure

Durability LO hours :ime between overhauls, 100 reuses,300 stares, 300 thermal cycles, 10,000 valvecycles

Single Continuous Capability from 10 seconds to 600 seconds

Run Duration

Engine Starts Multiple restart at sea level or altitude

Thrust Vector Amplitude: ±7 deg;Control Rate: 30 deg/sec;

Acceleration: 30 rad/sec 2

Control Capability -37. accuracy in thrust and mixture ratio atnominal thrust. Excursions from extreme toextreme in thrust and mixture ratio within5 seconds

Propellant L0 2 : 16 ft NPSH from 1 atmosphere boilingConditions temperature to 180*R

LH2 : 60 ft NPSH from I atmosphere boilingtemperature to 45•R

31

COIHFIDEN1AL

CONFIOENTIAL(C)(UW Table IV. Flight Engine Characteristics (Continued)

Environmenntal Sea level to vacuun conditionsConditions Combined accelerati)n: 10 g's axial with

2 g's transverse, 6.5 g's axial with3 g's transverse, 3 g's axial with6 g's transverse

Engine/Vehicle The engine will receive no external power,with the exception of normal electrical powerand 1500-psia helium from the vehicle

4. Cycle Balance

(C) A 250K LR129-P-l flight engine system and component operating param-eters at 1004, and 20'4 thrust and mixture ratios of 5, 6, and 7 arepresented in Table V.

5. Schedule With Cost

(U) The schedules of the flight engine development program to PFRT asrelated to the XLR129-P-1 demonstrator engine program are shown inFigure 16. The estimated development costs to PFRT and Qualificationas a function of rated thrust between lOOK and 500K are shown in Fig-ure 17.

C. PAR-METRIC ENGINE DATA

1. General

(U) Parametric engine data are presonted in this section for two basicconfigurations: (1) engines equipped with two-position exhaust nozzlesand (2) engines using conventional fixed exhaust nozzles.

(C) The two-position nozzle engines may be configured to provide either(1) high sea level performance or (2) minimum stowed length; the selec-tion of the primary nozzle expansion ratio (dp) depending upon theapplication. The above objectives are achieved by having the two-position nozzle retract from a nozzle expansion ratio of 35 (denotedby 9p a 35) mainly for lower stage applications or from the expansionratio that will provide minimum stowed length (denoted by ep a minimum).

(C) Engine data provided for the fixed and two-position nozzle engineconfigurations are: (1) performance (specific impulse), (2) weight,and (3) envelope. These data are for a vacuum thrust range of 100,000 lb(lOOK) to 500,000 lb (500K), mixture ratios from 5 to 7, and nozzleexpansion ratios (E) 35 to 400 for three nozzle truncations or contours),as applicable. Nozzle -runcations considered in these data presentationsare maximum pertormance (MC.), base, and minimum surface area (MSA) nozzldcontours. A discussion of nozzle contours is provided in Section II.Engine dimensions and nomenclature are illustrated in Figure 10.

32

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t~flP COPNN~ O01114119 Kfl, IL•IV D3enonstrattiriT t nXLine Pp rognm I

iini Development to PIrRT

II I 1 ''I I LR 29

I . Development to PFRTI . , (Othkr Thrust Si, in

) .,. , u * ~ 100.50K Range)

I P I i

I',I"? , 1 .. IN I. a Imire , L3 iV -I

(LI) Figure 16. XLR129-P-l Demonstrator Engine FD 24282AProgram

z • I• ' CoQualific a iitionan "zo

aCJ

ENGINE VACUUM THRUST • thouosnds of pounds

(U) Fig~ure 17. E~stimated Engi~ne Development PD 27781Cilts, Oxygen/Hydrogen Engines

(C) The engine charact~erist-ics of specific Impulse, weight:, and envelope

are based on t~he following:

* Pe~rformance at~tainable at: the Itime of preliminary flightrating test (PFRT)

* Engine inlet propeLLant conditnons:

Minimum requsred hydrogen n/ty positve section head

(NPSHi) - 60 ft

35

CONFIDENTIAL(C) The! enin c-haracteritic of spcii imule weight,_ an envelope ,_• ,i

CO,",DENTMALMinimum required oxygen net positive suction head(NPSH) a 16 ft

• Continuous throttling capability between 100% and 20% of

rated thrust

* hixture ratio range of 5 to 7 at all thrust levels

* Thrust vector control provided by mechanical gimbaling,t7 degrees

* Durability of 10 hours time between overhaul (TBO), 100reuses, 300 starts, 300 thermal cycles, and 10,000 valvecycles

0 Lightwelght, dump-cooled nozzle construction

* Performance based on the use of nozzle dump cooling forexpansion ratios greater than 35

* For high expansion ratio nozzles, radiation cooling isused aft of the lowest expansion ratio permitted by heatflux levels. (This expansion ratio varies over the para-metric range, but is approximately 200.) If radiationcooled nozzle skirts were not used, an insignificant in-crease in engine weight would result.

2. Engine' Performance

a. General

(U) Vacuum specific impulse, sea level specific impulse, altitude per-formance, and throttling performance are presented in this section forfixed and two-position nozzle engine configurations.

b. Vacuum Performance

(C) Vacuum specific impulse data are presented in Figures 18 through 20as a function of nozzle expansion ratio. Data are presented for maximumperformance (MCs), base, and minimum surface area (NSA) nozzle contours.These curves, which cover a nozzle expansion ratio range of 35 to 400,are applicable for all lightweight-nozzle (fixed and two-position)engine configurations in which dump cooling begins at an expansion ratioof 35:1. Vacuum specific impulse Ls very nearly independent of thrustlevel in the range of lOeK to 500K; this is particularly so between200K and 500K.

36

CONROFIMTAL

CONFIDENTIAL

,I M

(U) Figure 18. Vacuum Specific 1:--ipulse vs DFC 7b293No~e xpansien Racko

644, 4 4 S . ; . "

AA IW

AWAL- LEVU 2 •. %A IIJ • I

(U) Figure 19. Vacuum Specific I rpulsc vs DFC 70293Nozzle Expansion Ratio

37CONFIOF.NTAL

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P * IA4A .4t

CONFIDENTIAL

L II F'i.11re 2). Vacow:: Spec ifi c !::kjp Is, vs Dilc 70299

Nozzle Expansion Ratio

c. Al ti Ltde Pert or:'iance

i ;vncra I

I U) Al t i tude pertfor:-ianco for eng ines equIi I pped wi th I gli htwe iglit nozz les

Ls .snowi tii It.iures ZI ttirough It). Il.se data are presented as theraLi0 of specitiC i:-ipolse (at altitUde) to vacutum specific impuilIekIsalt lVa), versus altitude as a funcetiOn of nozzle expansion ratio.The a lti tide rang,Ž is from sea level to 200,000 ft (vacuumI conditions).Iheo thrust at any altitude may also be obtained by use Of these curvesbtcatse I salt'lvac a Falt lvac. LEsti,:;lted ntoZzle flow separationattitude is also shown, where applicable, for the higher expansion ratios.

'2 Two-',sition Nozzle Engine Configurations

(a, Pri:.arv Nozzle Area Rati4o - 35

(ký AlIt ,t,.,: per formahcc: ;or two-posit ion nozzle eng ine conf igiarat ionshcving a primary nozzle expansion ratio 01 35 4 p u 35) is shown in Fig-uircs 21 through 23 for MC., base, and MSA nozzle contours. These Curvesprosent altitude perfor:itance for the engines with the two-position nozzlein the extended position for high altitude operation and with it retractedfor operation at lower altitudes. The inflection points In these figuresresult fro,' the translation of the two-position nozzle, These data areusable for the entire parametric range of mixture ratios withotit si.t-nificant error. These data are inJependent of thrust level for a con-stant prirury expansion ratio. The nozzle expansion ratio range pro-vided in these curves Is from d a 50 to f - 150.

38

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CONFIDENTIAL

CONFIDENTIALlb)~ l'r i-iri Nozzlec Areai iat lo S ized ior Mininitutiý Stowied Lo4ngth

iV A.Lit tide p%' ri r: n.mc tor two- pos i tio nozzl v I cginu contifi.u rat ionst h~at provid d l illi.-awl stowe I v~Iltth (0 P ia nuil:?iLr) is Presen ted, il Fig-

(C) As 01~ the aso of tWo-poS i tionl nloZZ l% with1 U C011s tanlt prinuir y nozzlee.xpansi on ratio lt p i: I *ý tii% vat. Iat ion inl aIt iwtde pvrformanco reou It~fro' nijxtulro. ratLio 118in i %41 E ii i.:J11t . iiciq. is air addi Lionla I ci ect caused

bv thcQ variation inl thi. pri:iarv nozIlo expans ion ratio as a function ofti ruit IL icl. iiiis oiftvet is alIso .ocalvs::ia II and produc:es a totalIerror, il~tlkdincý MisLitrv ratio vtci cs , which Is no :ioro than -U.5' inioa Iovo I pert or:man.-o . Variit ions ill pr i::iary nozzl Ic xpans on ratito alsoaffecct thLv secondarv no~i.,et ranslIat ion alci .Itude ( tho a1lti tude at wh ichthe seco~ndarv nozzlec is trlalS ljLvd to its QXt%?nlded posi tion for highperfrmtance* opcirat ioni thio rcsu iltant vfoA inl mrhs tation alIti nide isa!PPfoX i:!i to v '3000 fc t . 1;ocatiso all rthe variat ions ci ted above are

no ~i ivo vStalJL , I sin:. Ic curvo. for r a t.0 is presen ted [or ctit i tdepv' *t~anofor echatlozz Ilvonto Nozzv Ic xpans ion ratios for these

cuirvos covoyr Olt, raiwo z to-i #* SU tO I 400.

1i Fihbjd Nozzle Etuti ne, C itfigkirat ion'S

(L) Alt i ide pe'rt r::zance for f ixe.d nozzl Ic ngine e'nt igurat: ions isprosented inl Fititr 27 for nozzl,. expansion ýat los of f u 35 to 4 u400.

1 t: call bc ,;evi f roit h'i s 1'itZj %ui r 1,, lower S tage aipp lIicat ions., Wih LitnZZ Icexpansion ra Lo0s groi. Lor Luali ajppl1X i:utat 1> 100: 1 are noL practical with

110...1mb 'ovatisio of !iozzi j~ low separition at low altitudes.

J. Sea 'Level P'erforma:nce

I) Gemiral

(C) To, f ac i Lirate veh ic le s iz im; and to a I I ty a d 1rec t:'malo of serlevelI peror±m'rnce, sca level specif ic ittpulne is presented in F~igures 2b~tnrtrouh 10 as a func tion of overall Io~ tLZ11 expalns ion ratio. Data areprov ided for threc nozzl Ic ' ontours for f ixed nozzle and t wo-pos it ion nozzle~

3)egioie confl4uration4.

(2) Two-I'os it ionl (4 a 35) and F ixed N'ozzl c ':ng ine Oonf l;,uratiotn5

(C) Sca level spec (fir. im;pulse for tw,). posit~ion nor.e I engiace configura-tions having a primary nozzle expansion ratio of 35 is presenred foroverallI expoins ion ratios of 30 t,1 150. Sea level specific impulse~ for

LxL'(i nozzle engine CLonfi~urat ions is prct~ented for ox sion ratios 35to. 100. F 'low separation at 'sea 'level occurs in fixed n../ lus with expan-s ion ratilos approximately'. 100 or :.reaf-or.

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W) Figure 27. Altitude Perfornmance !FC b9930

4.3

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WL) Fiutre '-8. Sea L.evel Speific lIuipuls., vs DEC 70352No)zzle Expansion: Ratio,

kt') Figur.Ž 29. Sea Level Spe'iiic Lriptlse vs DFC 70354Nozzle Expansion Ratio

CONFIDENTIAL

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CONFIDENTIAL

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(U) Fi.ture 30. Sea Level Spe..ific Impulse vs DFC 70337,Zzle IExp~isi,'in Ratio

kC) A comparis.'n Of dhCe seCa level .-peCific ih:pulse attainable with fixednozzle and two-position no:,ýIe engine configurations shows the significanti:::pr,,%vcm.rt in per.or.mance chat resuilts from the use of a two-posictionnozzle. In the case o;' two-position nozzle engine configurations for usein lower sta.e applications, the translating nozzle is in the retractedpos it ion for sea level and low altitude operation. tWith the two-pOs it ionnozzle retracted, the eng1ine operates with a low expansion ratio (fp = 35).This produces a hi-iher sea level specific im.pulse than that obtainablewith a fixed nozzle having the sa;::e overall nozzie expansion ratio. Atwo'-pos it iun nozzIC engine confiizuration with an overall nozzle expansionratio ,f 10O and a primary nozzle expansion ratio of 35 wi'[ provede asea level specific impulse that is approxim.ately 50 seconds greater thanthat attainable with a fixed nozzle for the sane conditions of nozzlecontour, mixture ratio, and vacuum thrust. Sea level thrust would beincreased in the same proportions.

e. [hrott led Perfo,.ance

(C) Throttling performrance of high pressure eni:,s is shown in Figure 31.Thiis figure presents the ratio of specific impuis_ at throttled conditionsto nominal specific impi'lse versus percent nominal thrust a- a function

of mixture ratio (r v 3.0 to r = 7.0). The effect of nozzle expansionratio on throttled performance is insignificant (less than 0.11/.) and can

be disregarded.

-5

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jow0. Nit%.

Throttled Cond itions

3. Engine Weig.ht.

a. General

(U) Engine dry weight for fixed and two-position nozzle engine configura-tion) are presented in Figure 32 through 40 for MCs, base, and 69Anozzle contours as functions of vacuum thrust and nozzle expansion ratio.

(U) The total engine dry weight includes the weight of all engine com-

ponents on the engine side of the vehicle/engine interface. Specifi-cally, it includes the following:

I. Thrust chhmber (transpiration cooled)

2. Exhaust nozzle (regenerative and dump-cooled sections)

3. Nozzle translating mechanism (where applicable)

4. Preburner assembly

5. Fuel and oxidizer turbopump assemblies

6. Transition case and gimbal

7. Engine-mounted and driven fuel and oxidizer low-speed inducers

8. Ignition system

9. Engine controls, shutoff valves, and actuators; and plumbing

10. Gimbal and actuator arm att ach.nt brackets for mechanicalthrust vector control.

Flight instrumentation and its hardware, and TVC actuator mechanismsare not included in engine weight.

46

CONFIDENTIAL

b. Two-Position Nozzle and Fixed Nozzle Engine Configurations

(C) The tot&l weight of two-position nozzln engines having a constantprimary nozzle expansion ratio, :p * 35, are presented in Figures 32through 34 for a nozzle expansion ratio range of 0 - 50 to 0 w 150.Engine weights for configurations having Itp minimum are shown inFigures 35 th.7ough 37 for nozzle expansion ratios 0 a 80 to I a 400.Fixed nozzle engine weights are presented in Figures 38 through 40.Engine weight differences between the two translating nozzle configura-tLons, ap * 35 and ep for minimum stowed lengths, are caused by varyingnozzle transla ing mechanism requirements.

(U) Relatively flat slopes for engine weight as a function of nozzleexpansion ratio are obtained with dump-cooled, lightweight nozzle engineconfigurations. This is a result, primarily, of using lightweight nozzle

construction for dump cooling beyond an area ratio of 35. As the overallnoztle expansion ratio is increased, the surface area of the regenera-Lively cooled portion of the nozzle (to I w 35 in all cases) becomessmaller because of the change in the contouring of the nozzle (which

results from the change in overall expansion ratio) and thus the nozzle

becomes lighter. Conversely, the dump-cooled portion of the nozzlebecomes larger and increases in weight as overall nozzle expansion

ratio is increased. The net result is a relatively small increase in

total engine weight with increasing expansion ratio for all thrust levels.

.. "I. I,- It p

I W10

(U) Figure 32. Total Engine Weight vs Nozzle DFC 70300

Expansion Ratio

47

- - .mamm- ----- a-

CONFIDENTIAL

s~sm

.4 * .AJ1! p S m -

I r P,,htII.. ,*Oi*I9 i.liPS4i ire, * - ).

140 ISO

R.~L i AI TIO a

(U) Figure 33. Total Engine Weight vs Nozzle DFC 70301Expansion Ratio

40"P

P0OL

WAmS. *IXM I fMg -4

"(U) Figure 34. Total Engine Weight vs Nozzle DFC 70302Expansion Ratio

48

CONFIDENIA

CONFIDENTIAL

Figr/•ure 35. Ttoal Engine We~aht vs N'OzzleExpansi on Ratio

DFC 70303

10

*"---#0

(U) F-gure 36. Total Enghe Weiht vs NozzleExpansion Ratio

DFC 70304

CONFIDEPtAL

IIIIIII I

CONFIDENTIAL

Ila

iwis._ _ _ _ _ _M9

ofUU0 *dImMile .

()Figure 237. Total Engine Weight vs Nozzle DFC 70305Expansion Ratio

SON.

:jft

WULI PIP"Sm ri,11 .#IVII

(U) Figure 38. Total Engine Weight va Nozzle DFC 70306Expansaionl Ratio

50

CONFIDENTIAL

CONFIDENTIAL

j, .. .... .0

,A * ,,i ,L

(U) Figure 39. Total Engine Weight vs Nozzle DFC 70307

Expansion Ratio

Ism

WDS

mm) l r SOIV BAII

(U) Figure 40. Total Engine Weight vi Nozzle DFC 70308Ex~pansion Ratio

51

CONFIDENTIAL

CONFIDENTIAL4. Engine Envelope

a. General

(U) Engine dimensions of overall (nozzle) exit diameter, overall length,and stowed lengths (for two-posltiu'i nozzle engine configurations) arepresented in this section in Figures 41 through 50 as a function ofnozzle expansion ratio. Length data are presented for three nozzlecontours; MC., base, and HSA. Because overall exit diameter is notsignificantly affected by nozzle contour, only a single figure is pre-sented for this parameter. Primary nozzle expansion ratiu as a functionof overall nozzle expansion ratio for minimum stowed length engine con-figurations is presented in Figures 51 through 53.

(C) Stowed engine length for two-position nozzle engine configurations,in addition to being a function of thrust level and nozzle expansionratio, Is also dependent upon the expansion ratio of the primary nozzle.For lower stage applications, a small primary nozzle expansion ratio,ep - 35, generally produces the best performance; in upper stages,larger primary expansion ratios, tp - 80, provide minimum stowed length.

(C) When the primary nozzle expansion ratio is set at a constant value(i.e., *p - 35), the primary nozzle exit plane determines the stowedlength for low overall expansion ratios. As the overall expansion ratiois increased, a point is reached where the two-position nozzle determinesthe stowed length. The inflection points in the curves of stowed lengthoccur where the exit planes of the primary and two-position nozzles arein aiignmenc.

b. Diameter

(C) Overall exit diameter is presented in Figure 41. as a function ofvacuum thrust for nozzle expansion ratios ( - 35 to e - 400. This curvemay be used for all nozzle truncations (contours).

c. Length

(1) Overall Length

(C) Engine overall length for fixed and two-position nozzle configura-tions is presented in Figures 42 through 44 for each nozzle contour asa function of nozzle expansion ratio, t a 35 to t - 400.

(2) Stowed and Minimum Stowed Length

(C) Engine stowed length for tp a 35 is presented in Figures 45 through 47.Minimum stowed lergth (ep a minimum) is presented in Figures 48 through 50.Stowed length curves are presented for each rf three nozzle contours as a

function of nozzle expansion rdtio.

(U) For minimum stowed engine length, two hardware geometry considerationsare the determining factors: (1) the nozzle translating mechanism and(2) the turbomachinery or power package. Inflections in the minimum

stowed length curves are caused by a changeover in the limiting factor.

52CONFIDENTIAL

CONFIDENTIAL

1ý~~ - M al

.. .a .. ..

104K

10%

100

l0

W91. WUS I,*1 PATIO f

(U) Figure 42. Overall Lexgth Diameterlv DFC 70274Epnosion ExpasionRto

30053

a CONFIDENTIAL

CONFIDENTIAL

-oll ..J ,

()Figure 43. Overall Length -,s Nuzzle DFC 70276Expansion Ratio

N,~,f,~ . ... ..... ___ I

Comm=1,.

QQ '00 a(

WV,: PANS[I" 141n -

(U, Fiue4.Oeal eghv ozl F 07

ExasinRai

-~ -5-

CONIDETIA

•omom

UhW~d~

V * ' 4,prU '

4.

10110

I SO

So I,,,,'

!a 0.00 kVI %1000

NOZZ0LE PMtUI •A I ATIO t

(U) Figure 45. Stowed Length vs Nozzle DFC 70366Expanssion Ratio

I• +%.Plhll llt• l..I.il A.l~Irilt. fwI 4 % ''/atu,,0 Thr,.t

, 5

U.li 0.t0

S1400 0~0,1

100 . / 000

wl200'

0o SO ---- -0i

00Zllt NP0.IOIJS•III A.TIO .4t

(U) Figure 46. Stowed Length vs N~z~zle DFC 70365Expansion Rati•-

Lpl

ICONFIDENTIAL

114nt- tt..ft . 4r.. It.." I.I.tIr ae T ,.

IIto

110 low:re

IOGK- M

.4

MOOZUt 1*0121P.1kI RA17IO -

(U) Figure 47. Stavded Length vs Nozzle DFC 70364Expansion Ratio

Ito

j M11?w 20

1.0 1140

Ito2%O

I-o~

11-06. V" I~ct. " ntl.CM

K0..fr-. . I.cI.

too2 zoo me 400

(U) Figure 48. Minimum Stowed Length vs DFC 70363Nozzle Expansion Ratio

56

CONFIDENTIAL

CONFIDENViAL

~0p IW)low

(U iueq 49. Miniu *ltowed Leghs 06

P- - owW

I NOW

*1111 1AA106%fO-

(U igur 50.yi Minimum Stowed Legt v FC706

Nozzs*f.044W~ll*del Expasio Ratio

574

"CNlb.RA

CONFIDENTIAL

S

Mt1J*I kll fOtWl h1a1 l ., f4

wuetm Mnll I~~l~Zr AIe~ UMt psi

r- o t. *4 1 4..10"t.ou

(U) Figure 51. Primary Nozzle Expansion Ratio DFC 70360

vs Nozzle Expansion Ratio

*10

* * 300 psi.

".Wo1 NoruttC

ftI e do@ Ikk.l

. 0e010 Coote'."

IIG

0A 01011100 IA00 #0IWSLL, B UIWkll LA•TlOP **

(U) Figure 52. Primary NozzlQ Expansion Ratio DFC 70358vs Nozzle Expansion Ratio

58

CONFIDENTIAL

CONFIDENTIAL

L. A•~t, -At Wn I

(U) Figure 53. Primary Nozzle Expansion Ratio DFIC 70359vs Nozzle Expansion Ratio

D. ENG1NE/VEHICLE INTERFACE DATA

L. GeneraL(U) Eng)neiVehucle interface data are presented in the tables of 7 his

section as a function of thrust level in 50K increments. These dataare based on the demonstrator engine configuration. These interfacedata include inlet condition operating region, engine inlet and TVCactuator arm attach point locations, engine inlet and power packagediameters, TVC actuator arm loads and auxiliary power available fromthe engine.

2. Engine Inlet Conditions

(C) The high pressure engines are designed to be capable of operatingover a wide range of fuel and oxidizer inlet conditions provided minimumnet positive suction head (NPSH) requirements are met. The flight en-gine inlet operating regions are the same as those for the demonstratorengine (see Figures 5 and 6) with minimum required NPSH's of 60 feetand 16 feet at the fuel and oxidizer inlets, respectively. If specialvehicle operating conditions require inlet conditions outside of thenormal engine operating regions, these should be coordinated withPratt & Whitney Aircraft to ensure engine/vehicle compatibility.

3. Engine/Vehicle Interface Locations

(U) Engine inlet and actuator arm attach point locations are presentedin Table VI. These dimensions are referenced to the engine X, Y, andZ axes shown in Figure 54. The engine fuel and oxidizer inlet flangesare in the same plane as the gimbal axis of the engine.

59

CONFIDENTIAL

CONFID.HTIAL "(U) Table Vt. Location of Engine Inlets and

Actuator Arm Attach Pointst

Vacuum Thrust (Thousands of lb)

Dimension (in.) 100 150 200 250 300 350 400 450 500

A 12 14 16 18 20 22 23 25 26

B 5 6 7 8 8 9 10 10 11

C 11 14 16 18 19 21 22 24 25

0 9 11 13 14 15 17 18 19 20

E 8 9 11 12 13 14 15 16 17

F 3 3 4 4 4 5 5 5 6

G 2 2 2 3 3 3 3 4 4

H 8 10 12 13 14 15 16 17 18

1 20 20 21 21 22 22 23 23 24

Prehirner = "Centerline Aslxr-i .

/ FuelA-- Inlet

/ �1Osidizer "-

SE--i j I I ower-Package ,-- jl

"7 j-"" I iamoeter-

T]t

-\ l \\I

G TVC ctusto,/ ,

\ ~~~~Arm Attah /.Points .E

-0 Engine I- --. Gimubal

(U) Figure 54. Engine/Vehicle Interface FD 27786ALocations

1Refer to Figure 54

60

"CONFIDENIAL(This page is Unclaossilied)

4. Inlet Line and Power Package Diameters

(U) Fuel and oxidizer Inlet line diameters at the engine/vehicln inter-face (low-speed inducer inlets) are presented in Table VII. The diametersof the inlets are determined by propellant NPSH and engine cycle require-ments.

(U) The engine power package consists of the turbomachinery, preburner,main burner injector and chamber and the associated plumbing lines; itis essentially the entire engine except for the exhaust nozzle. Thepower package maximum diameter occurs in the plane of the main turbo-pumps and preburner parallel to the gimbal axis, and generally does notexceed a diameter equivalent to that for a nozzle expansion ratio of 80.The power package diameters are presented in Table VII.

(U) Table VII. Power Package Diameter, Oxidizer

and Fuel Inlet Diameters

Vacuum Thrust (Thousands of lb)

100 150 200 250 300 350 400 450 500

Power Package 41 50 58 65 71 76 82 87 92Diameter (in.)

Oxidizer Inlet 7 9 1.0 12 13 14 15 16 17Diameter (in.)

Fuel Inlet 8 9 11 12 13 15 16 17 18Diameter (in.)

5. Gimbal. Loads

(C) Thrust vector control (TVC) for the high pressure engines is accom-plished by mechanical gimbaling. Two actuator arms attach to the engineat points 90 degrees apart. The estimated maximum gimbal loads (for eachactuator arm) are presented in Table VIII. Gimbal loads were based onthe following gimbaling requirements:

Anglh - ±7 deg

Velocity - 30 deg/sec

Rate - 30 rad/sec2

6. Auxiliary Power

(U) The high pressur', engines are designed to provide auxiliary (accessory)power for TVC and other uses. The maximum power availability is presentedin Table VIII.

61

• ~ ~ ~~ Boomi P •

i-.m

(U) Table VIII. Gimbal Loads and Auxiliary PowerAvailable from the Engine

Vacuum Thriisc (Thousands of Ib)

100 150 200 250 300 350 400 450 500

Gimbal Loa4s 16 22 28 33 39 45 51 56 62(Thousands of lb)

Auxiliary Power 43 60 80 100 117 134 151 168 185(Horsepower)

62

(This poll@ is Unclassified)

UNCLASSIFIED

REFERENCES

'1. High Pressure Oxygen-Hydrogen Rocket Engine Parametric Design Data,PWA FR-492D, 30 November 1966

*2. Applications Study for a High Performance Cryogenic Staged Combus-tion Rocket Engine, Final Report, AFRPL-TR-67-270, November 1967

63UNLSSFED _

Ulnclass ifiedSecurity Classification

DOCUMENT CONTROL DATA- R(3041t, t~ iIiticawtn of titffe findy ol sbhict And Jnde wtn Annotation et •@ftbpi

I ORIGINiATING ACTIV$ýY rc.w,P-. awthe,) U V=T I~IT. 'ZL.41PP -i rip~Pratt & Whitney Aircraft •J:J .f nDivision of United Aircraft Corporation ubt ...Florida Research and Development Center 4

3. RIPOOT TITLEAir Force Reusable Rocket Engine Program XLR129-P-1Engine Performance

4. DESCRIPTIVE NOTES (Type ot ifpart amd nclutle date.)

Special Report

S. AUTHO(•S) (LZe name. ti•el nome. Initial)

Atherton, Robert R.

G. RMEPO RI DATE 7T. TOTAL NU. OF PAGEQS 7b. NO. OF Rgrz

April 1969 70 2"41. CONACT ON GRAM? -0. to. o.,..ATOR.S REPORT NU6...I(S)

F04611-68-C-0002 PWA FR-3108& PROJECT MO.

57X 00 289N___3C. Sb1. OTHER POTMO(S) (Any otim.,an.be,. Act a,., be designed

P6341C 63408F Mi. I=

d.S669800 ' AFRPL-TR-69-19 ,-

10. AVAILAMILITY/LIMITATION NOTICES In addition to sP-ttrity requirements which must be metthis document is subject to special export controls and each transmittal toforeign governments or foreign nationals may be made only with prior approvalmf AVPT,(RPOR/STTNF0) Edwards, California 93523

11. SUPPLEMENTARY NOTES I. SPONSORING MILITARY ACTIVITY

AFFTCProcurement Division (FTMR-2)Edwards AFB, California 93523

13. A dýPIACTSThis special technical report presents information and data on the

XLR129-P-l rocket engine. Information is presented for both the demonstratorengine and flight engine versions of this rocket engine. A general descrip-tion and pertinent technical information are presented for the demonstratorengine. The demonstrator engine program schedule is also presented.Parametric design, performance, cost, and schedule data are presentedfor the flight engine. This technical repurt has been prepared for theuse of airframe manufacturers and government personnel who are conductingmission and vehicle studies."

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Air Force Reusable Rocket Engine ProgramPratt & Whitney Aircraft, FR.DCXLR129-P-1 Demonstrator EngineInstallation DrawingFlow SchematicOperating ParametersFlight Engine DescriptionNozzle Conce'ns

Parametric Engine DataI

EnvelopeEngine/Vehicle Intei.:ace Data

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