BM SA - NASA · BM SA UNCONVENTIONALNOZZLE TRADEOFF STUDY (N&SA-CR-159520) UNCONVENTIONAL NOZZLE ... Work Unit No. 11. Contract or Grant No. NA53-20109 13. Type of Report and Period

NASA CR-159520

BM SA

UNCONVENTIONALNOZZLE TRADEOFF STUDY

(N&SA-CR-159520) UNCONVENTIONAL NOZZLE

TR&DEOFF STUDY (AeroJet Liquid Rocket Co.)

311 p CSCL 21H

N79-28224

Uuclas

G3/20 29408

AEROJETLIQUID ROCKET COMPANY

Sacramento, Californi a

prepared for

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

July 1979

CONTRACT NAS 3-20109

NASA Lewis Research Center

Cleveland, Ohio

C. A. Aukerman, Project Manager

https://ntrs.nasa.gov/search.jsp?R=19790020053 2020-06-08T07:45:24+00:00Z

t. Report No.

NASA CR-159520

4, Title and Subtitle

2. Government Accession No.

Unconventional Nozzle Tradeoff Study, Final Report

7. Author(s)

C. J. 0'Brien

9. Performing Organization Name and Address

Aerojet Liquid Rocket Company

3. Recipient's Catalog No.

5. Report Date

July 19796. Performing Organization Code

8. Performing Organizahon Report No.

10. Work Unit No.

11. Contract or Grant No.

NA53-20109

13. Type of Report and Period Covered

Contractor Report, Final

14. Sponsoring Agency Code

Post Office Box 13222Sacramento, California 95813

12, Sponsoring Agency Name end Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

15. Supplementary Notes

Project Manager, C. A. Aukerman, Space Propulsion and Power Division,

i NASA-Lewis Research Center, Cleveland, Ohio

16. Abstract

I Plug cluster engine design, performance, weight, envelope, operationali characteristics, development cost, and payload capability, were evaluated' and comparisons were made with other Space Tug engine candidates using

oxygen/hydrogen propellants.

Parametric performance data were generated for existing developed or hightechnology thrust chambers clustered around a plug nozzle of very largediameter. The uncertainties in the performanc_e__p]ce_d_u_ cluster_engines with large gaps between the modules (thrust chambers) were evaluated.

The major uncertainty involves the aerodynamics of the flow from discretenozzles, and the lack of this flow to achieve the pressure ratio correspondingto the defined area ratio for a plug cluster. This uncertainty was reducedthrough a cluster design that consists of a plug contour that is formed fromthe cluster of high area ratio bell nozzles that have been scarfed. Light-weight, high area ratio, bell nozzles were achieved tlqrough the use ofAGCarb (carbon-carbon cloth) nozzle extensions.

The plug cluster engine proved to be competitive with the uprated RLIOengines, the Aerospike, and the Advanced Space Engine. High vacuum per-formance is achieved with the low pressure Plug Cluster Engine because ofthe large area ratio available with the baseline Space Tug.

17. Key Words (Suggested by Author(s)

Plug Cluster EngineSpace Tug PropulsionLiquid Rocket EngineOxygen/Hydrogen Cryogenic Engine

19. S_urity Classif. (of this report)

Unclassified -.-

tB. Distribution Statement

Unclassified - Unlimited

20. Security Classif. (of th=s page)

21. No. of Pages296Unclassified

Price"

*For sale by the National Technical Information Service, Springfield, Va. 22151

... !:.r-J:

I!!I

FOREWORD

The work described herein was performed at the Aerojet LiquidRocket Company under NASA Contract NAS3-20109 with Mr. Carl A. Aukerman,

NASA-Lewis Research Center, as Project Manager. The ALRC ProgramManager was Mr. Larry B. Bassham, the Operations Project Manager wasDr. R. J. LaBotz, and the Project Engineer was Mr. Charles J. O'Brien.

The technical period of performance for the study was froml July 1976 to 28 April 1978.

The author wishes to acknowledge the efforts of the followingALRC engineering personnel who contributed significantly to this report:

L. L. BickfordP. E. BrownK. L. Christensen

H. O. Davis (ASPC)

J. A. HeneyR. A. Hewitt

M. C. HuppertG. R. Janser

L. L. LangS. Leone

S. A. LorencE. Lueders

G. M. MeagherR. W. MichelG. H. Peirson

J. L. PieperN. M. RichardsonD. C. RousarJ. V. Smith

I also wish to thank Mr. Rudi Beichel, ALRC Senior Scientist,for his comments and assistance throughout the study effort.

iii

I!I'I"-

Section

I. Summary

II.

III.

TABLE OF CONTENTS

A. Study Objectives and Scope

B. Results and Conclusions

Introduction

A. Background

B. Purpose and Scope

C. General Requirements

D. Approach

I. Task

2. Task

3. Task

4. Task

Task

Task

Task

Task

Task

Task

I:

If:

III:

IV:

Literature Analysis

Parametric Engine Performance

Subsystem Evaluation

Preliminary Design

5. V: Plug Cluster Engine Optimization

6. VI: Plug Cluster Engine Assessment

7. VII: Lightweight Engine Structures

8. VIII: Thrust Vector Control Analysis

9. IX: Fluid Systems and Control Study

I0. X: Experimental Performance DataEvaluation

II. Task XI: Plug Cluster Engine Optimization

Literature Analysis

Objectives and Guidelines

Space Tug System Studies

I. Baseline Space Tug

2. Engine Evaluation

C. Plug and Plug Cluster Rocket Nozzle Studies

I. Plug Nozzle Performance Criteria

2. Plug Nozzle Design Criteria

H/O Thrust Chamber Technology

I. Integrated Thruster Assembly

Ao

B.

Do

Page

1

6

9

9

lO

I0

II

II

II

II

II

II

II

12

12

12

12

12

13

13

13

13

17

18

21

23

32

36

V

Section

IV.

V°

TABLE OF CONTENTS (cont.)

2. Extended Temperature Range Thruster

3. Hydrogen-Oxygen Auxiliary Propulsion Engines

E. H/O Turbopump Assembly Technology

I. APS Turbopumps

2. RLIO Turbopump Assembly

Engine Performance Methodology

A. Objectives and Guidelines

B. Module Parametric Performance Model

I. Performance Losses

2. Module Performance

C. Plug Performance Analysis

D. Plug Cluster Performance Analysis

I. Plug Cluster Design Constraints

E. Plug Cluster Performance Model I

I. Model I Calculation Procedure

2. Model I Fairing Correction

3. Model I Base Pressurization Correction

4. Model I Plug Cluster Engine Delivered Performance

F. Analysis of Experimental Plug Cluster Data(NAS 3-20104)

G. Module-Plug Cluster Performance Model II

H. Module-Plug Cluster Performance Model Ill

I. Model Ill Description

2. Model Ill Nozzle Efficiency

3. Scarfed Nozzle Performance

4. Model III Base Pressurization

5. Model Ill Plug Cluster Engine Delivered Performance

Subsystem Evaluation


B. Engine Cycle Analysis

I. Cycle Analysis Summary

Page

36

41

41

44

44

49

49

49

5O

51

52

61

66

68

77

81

81

82

88

97

98

98

I02

I03

I03

I07

ll3

ll3

ll3

126

vi

If!'I

Secti on


VI.

C. Turbomachinery Analysis

I. RLIO IIA Turbopump Assembly Analysis

2, Conceptual Turbopump Design Analysis

D. Engine Cooling Analysis

I. ITA-Type Module

2. Regeneratively Cooled Module

3. Regeneratively Cooled Plug Nozzle

E. Base Pressurization Analysis

F. Configuration Analysis

G. Parametric Weight Analysis

H. Thrust Vector Control Analysis

Preliminary Conceptual Design


B. Conceptual Design

C, Structures Analysis

D. Materials Analysis

E. Controls Analysis

I. Control System for Regeneratively Cooled Engine

2, Control System for Engine with Uncooled Plug

F. Module Design

G. Uncooled Plug Nozzle

I. AGCarb Nozzle Cycle Life

H. Uncooled Bell Nozzle Extension

I. Weight Analysis

I ,

2.

3.

,

5.

Module Weight

Plug Nozzle/Thrust Structure Weight

Module AGCarb Nozzle Extension and BaseClosure Weight

Turbopump Weight

Valve Weight

Page

129

129

135

147

147

149

152

171

174

174

176

187

187

187

194

199

201

201

2O8

211

218

220

224

225

225

226

226

226

231

vi i

Section

VII.

Vll.

IX.

X."

XI.


6. Line Weight

7. Weight Summary

Plug Cluster Engine Optimization


B. Engine Design Specification

C. Round Trip Geosynchronous Orbit Mission

D. Technology Requirements

E. Optimum Plug Cluster Engine

Plug Cluster Engine Assessment


B. Mission Exchange Factors

C. Cost Analysis

D. Life Analysis

Conclusions

Appendixes

A. Conventional Engine Operating Specifications

B. Space Tug System Study Bibliography

C. Plug and Plug-Cluster Rocket Nozzle Bibliography

D. H/O Thrust Chamber Bibliography

E. H/O Turbopump Assembly Bibliography

References

Pa e

231

231

237

237

237

242

243

247

249

249

249

256

257

259

263

263

277

283

285

287

291

viii

Table No.

I

II

Ill

IV

V

VI

VII

VIII

IX

X

XI

XII

XIII

XIV

XV

XVI

XVII

XVIII

XIX

XX

XXI

XXII

XXIII

XXIV

XXV

XXVl

XXVII

XXVIII

XXIX

LIST OF TABLES

Plug Cluster Engine Design Point

Baseline Space Tug Characteristics Summary

Baseline Space Tug Weight Breakdown

Nominal Performance, Double-Panel Aerospike EngineShowing Base Contribution

ITA Design Summary

APS Cycle Life Performance Matrix

APS Oxidizer Turbopump Performance Requirements

APS LH2 Turbopump Performance Requirements

Module Performance Parametric Ranges

Base Pressurization Data Summary

Gap Performance (CT) with Fairings

Aerodynamic Variables for Test Models

Comparison of Experimental Plug Cluster Performance

Base Pressure and Secondary Flow Effects (NAS 3-20104)

Plug Cluster Performance Model Comparisons

JANNAF Simplified Performance for Plug ClusterEngines

Plug Cluster Performance Model Comparisons

Uncertainty in Plug Cluster Engine Performance

JANNAF Simplified Performance for RLIO and ASE

Plug Cluster Engine Baseline Design Point

Cycle EXOI Preliminary Pressure Schedule

Cycle GGOI Preliminary Pressure Schedule

Cycle Analysis Summary

Cycle Analysis Summary Data

Turbopump Design Criteria

Low Speed LOX Turbopump Design Parameters

High Speed LOX Turbopump Design Point Parameters

LH2 Turbopump Design Point Parameters

Plug Heat Transfer Data

lO

14

15

28

39

43

45

46

51

64

81

89

91

97

99

108

109

llO

Ill

ll3

125

127

126

128

136

137

141

145

162

ix

Table No.

XXX

XXXI

XXXII

XXXIII

XXXIV

XXXV

XXXVl

XXXVll

XXXVIII

XXXIX

XL

XLI

XLII

XLIII

XLIV

XLV

XLVI

XLVII

XLVIII

XLIX

[

LI

LII

LIII

LIST OF TABLES (cont. }

Plug Energy Balance Calculations

Oxygen Heat Transfer Correlations Used for CooledPlug Analysis

Cooling Channel Design Calculations, Oxygen CooledPlug

Pressure Drop Estimate, Oxygen Cooled Plug

Plug Cluster Baseline Weight Summary for ParametricAnalysis

Estimated Control Hardware Characteristics

Material Selection for the Plug Cluster EngineConceptual Design

Preliminary Valve Selections

Expander Cycle Operational Sequence

Gas Generator Cycle Operational Sequence

Preliminary Valve Selection

Regen Cooled Module Life Cycle Determination

Demonstrated Experience with AGCarb

Summary of Selected AGCarb Material Properties

Materials and Fabrication Technique Trade Study

Module Weight Analysis

StructurePlug Nozzle/Thrust WeightAnalysis

AGCarb Nozzle Extension and Base Closure Weight

Typical Weight Breakdown Chart by Component forCandidate Space Tug Engines

Weight Breakdown for Lightweight Plug Cluster Engines

Plug Cluster/Bell Engine Operating Specification,Expander Cycle: Regen-Module, Model III, Pc = 20.4

Plug Cluster/Bell Engine Operating Specification,Expander Cycle: Regen/Module, Model III, Pc = 34.0

:plug Cluster/Bell Engine Operating Specification,Gas Generator Cycle: Regen-Module, Model III,Pc = 20.4Plug Cluster/Bell Engine Operating Specification, GasGenerator Cycle: Regen-Module, Model III, Pc : 34.0

Page

164

165

170

172

175

184

200

2O5

212

214

215

218

221

222

223

225

227

230

232

239

24O

241

i_| !

Table No.

LIV

LV

LVl

LVll

LVlII

LIX

LX

LXl

LXII

LXIII

LXIV

LXV

LXVI

LXVII

LXVIII

LXVIX

LXX

LXXI

LIST OF TABLES (con_.

Propellants Available for Round Trip Mission toGeosynchronous Orbit

Round Trip Plug Cluster Payloads to GeosynchronousOrbit

Plug Cluster Engine Technology Requirements

TVC Considerations and Options

Baseline Space Tug Mission Data for Engine Comparison

Mission Data for Engine Comparison

Payload Equations for Engine Comparison

Space Tug Engine Comparison (Revised)

Space Tug Engine Cost Comparison

Space Tug Engine Life comparison

Plug Cluster Engine Operating Speci fication, ExpanderCycle: Regen-Module, Model II Performance Pc =20.4 atm

Plug Cluster Engine Operating Specification, ExpanderCycle: Regen-Module, Model II Performance Pc =34 atm

Plug Cluster Engine Operating Specification, GasGenerator Cycle: Regen-Module, Model II PerformancePc : 20.4 atm

Plug Cluster Engine Operating Specification, GasGenerator Cycle: Regen-Module, Model II PerformancePc = 34 atm

Plug Cluster Engine Operating Specification, ExpanderCycle: ITA Module 16% FFC, Model II PerformancePc = 20.4 atm

Plug Cluster Engine Operating Specification, GasGenerator Cycle: ITA Module 16% FFC, Model IIPerformance Pc = 20.4 atm

Plug Cluster Engine Operating Specification, GasGenerator Cycle: ITA Module 16% FFC, Model IIPerformance Pc = 34 atm

Plug Cluster Engine Operating Specification, ExpanderCycle: Regen-Module/Uncooled Plug, Model IIPerformance, Pc = 20.4

Page

242

243

244

246

250

250

251

252

256

257

264

265

266

267

268

269

27O

271

xi

Table No.

LXXII

LXXIII

LXIV

LXXV

LIST OF TABLES (cont.)

Plug Cluster Engine Operating Specification, ExpanderCycle: Regen-Module/Uncooled Plug, Model IIPerformance, Pc = 34.0

Plug Cluster Engine Operating Specification, GasGenerator Cycle: Regen-Module/Uncooled Plug, ModelII Performance, Pc = 20.4

Plug Cluster Engine Operating Specification, GasGenerator Cycle: Regen-Module/Uncooled Plug, ModelII Performance, Pc = 34.0

Baseline Space Tug Engine Comparison

272

273

274

275

xii

I'_|I

Figure No.

1

2

3

4

5

6

7

8

9

I0

II

12

13

14

15

16

17

18

19

2O

21

22

23

24

25

26

LIST OF FIGURES

Unconventional Nozzle Tradeoff Study Program Summary

Plug Cluster Engine

Clustered Bell Nozzle Concept

Scarfed Bell/Plug Cluster Engine Concept

Space Tug Engine Evaluation

Baseline Space Tug General Arrangement and Size

Baseline Tug Engine

Advanced Engine Characteristics

Advanced Engine Evaluation

Effect of Truncating the Plug on Thrust Efficiency

Effect of Module Spacing on Propulsion System Performance

Effect of Fairings and Gap on Performance

Tilt Angle Model Performance for a Plug Cluster Nozzle

Effect of Tilt Angle on Base Pressure

Base Pressure vs Secondary Flow at Vacuum

Vacuum Thrust Coefficient Efficiency vs Secondary Flow

Effect of Radial Inward Base Bleed on Baseline Model

Performance (Design Pressure Ratio)

Prandtl-Meyer Turning Angle as a Function of Area Ratio(One-Dimensional)

Cluster Area Ratio vs Number of Modules

Plug Contour Design

Integrated Thruster Assembly is a Prime Candidate forthe Plug Cluster Engine

ITA is a Flightweight High Technology Thruster

Tested ETR Candidate Propellant Thermal Management Concepts

APS Thrust Chamber Assembly Schematics

Module Delivered Specific Impulse for a Fixed InjectorDesign Operating at a Chamber Pressure of 20.4 Atm

Module Delivered Specific Impulse for an Optimized InjectorDesign Operating at a Chamber Pressure of 20.4 Atm

Page

2

3

4

5

7

16

16

19

20

22

24

25

26

27

29

3O

31

33

34

35

37

38

40

42

53

54

xiii

LIST OF FIGURES (cont.)

Figure No.

27

28

29

3O

31

32

33

34

35

36

37

38

39

40

41

42

43

44

45

46

Module Delivered Specific Impulse for an Optimized In- 55jector Design Operating at a Chamber Pressure of 34.0 Atm

Influence of Expansion Area Ratio on Module Specific Impulse 56for a Fixed Injector Design

Influence of Expansion Area Ratio on Module Specific Impulse 57

for an Optimized Injector Design

Variation of Module Specific Impulse with Thrust Level for 58

an Optimized Injector Design

Sketch of a Plug Nozzle and Control Surface 59

Plug Nozzle Design Map 60

Plug Nozzle Two-Dimensional Thrust Coefficient 62

Features of Internal-External Expansion Axisymmetric 63Truncated Plug Nozzle Flow Fields

Base Pressurization Data Appear to Follow Nozzle Separation 65Criteria

Plug Cluster Geometry 67

Variation of the Cluster Amplification Factor with Number 69of Modules

Allowable Plug Cluster Design Conditions for Modules with 70

Zero Gap

Allowable Plu9 Cluster Design Conditions for Modules with 71

0.5 Gap (_/De)


l.O Gap (a/De)

Allowable Plu9 Cluster Design Conditions for Modules with 73

2.0 Gap (a/De)


3.0 Gap (6/De)


4.0 Gap (a/De)

Gap Efficiency Factor for Constant Plug Cluster Performance 78

Sketch of Plug Showing Computer Model Nomenclature 79

Plug Cluster Engine Performance Summary at Pc = 20.4 Atm 83

and EM : 40

xiv


Figure No.

47

48

49

5O

51

52

53

54

55

56

57

58

59

60

61

62

63

64

65

Plug Cluster Engine Performance Summary with ITA atPc --20.4 Atm

Plug Cluster Engine Performance Summary at Pc = 34

Atm and EM = 40

Plug Cluster Engine Performance Summary with ITAat Pc = 34 Atm

Plug Cluster Engine Performance Summary at Pc = 20.4

ATM and EM = lOO

Effect of Gas Properties on Required Tilt Angle

Cluster Performance as a Function of Engine AreaRatio

Pressure Tap Data Indicate the Effective Area RatioAchieved by the Flow of Gas on the Plug Nozzle

Effective Area Ratio of Plug Cluster is Less than Geo-metric Area Ratio

Mach Number Match Configuration for Model II

Scarfed Bell/Plug Cluster Engine Concept

Clustered Bell Nozzle Concept

Scarfed Nozzle Geometry

Kinetics Loss for Scarfed Nozzle

Boundary Layer Loss for Scarfed Nozzle

Cycle EXOI: Expander Topping Cycle H2-Cooled TCA,

02-Cooled Plug, Single Turbine TPA, Base Pressuri-

zation with H2

Cycle EX02: Expander Topping Cycle, H2-Cooled TCA,H2-Cooled Plug, Single Turbine TPA, Ba_e Pressuri-

zation with H2

Cycle EXD3: Expander Topping Cycle, Hp-Cooled Plug,TPA with Separate Gas Driven Turbines,-Base Pressuri-

zation with H2

Cycle EX04: Expander Topping Cycle, H2-Cooled TCA,

O;-Cooled Plug, Parallel Turbine TPA, Base Pressuri-

zation with H2

Cycle EX05: Expander Topping Cycle, H2-Cooled TCA,

H2-Cooled Plug, Parallel Turbine TPA, Base Pressuri-

zation with H2

84

85

86

87

90

92

93

95

96

I00

lOl

I04

I05

I06

ll4

115

ll6

ll7

ll8

XV


Figure No. Page

66

67

68

69

70

71

72

73

74

75

76

77

78

79

80

81

82

83

Cycle EXIIA: Expander Topping Cycle, H2-Cooled TCA, ]_902-Cooled Plug, Dual Single Turbine TPAs, Base

Pressurization with H2 (not shown)

Cycle GGOI: Gas Generator Cycle, H2-Cooled TCA, 120

02-Cooled Plug, GG Exhaust on Plug, Single TurbineTPA, Base Pressurization with Partial GG Exhaust

Cycle GG02: Gas Generator Cycle, H2-Cooled TCA, 121

H2-Cooled Plug, GG Exhaust on Plug, RLIO TPA, BasePressurization with Partial GG Exhaust

Cycle GGO3: Gas Generator Cycle, H2-Cooled TCA, 12202-Cooled Plug, GG Exhaust on Plug, Parallel TurbineTPA, Base Pressurization with Partial GG Exhaust

Cycle GG04: Gas Generator Cycle, H2-Cooled TCA, 123Hp-Cooled Plug, GG Exhaust on Plug, Parallel TurbineTPA, Base Pressurization with Partial GG Exhaust

RLIO Derivative IIA Fuel Pump Characteristics 130

RLIO Derivative IIA and lIB Oxidizer Pump Characteristics 131

Approximate RLIO Turbine Flow Parameter 132

Expander Cycle EX01 Power Balance with RL]O Turbine 733

Expander Cycle EX02 Power Balance With RLIO Turbine 134

Low Speed LOX Pump Dimensionless Performance 138Characteristics

Effect of U/C0 on Turbine Efficiency, Single Impulse 139Stage

Conceptual TPA Design, LOX Boost Pump 140

High Speed LOX and LH2 Pump Dimensionless Performance 142Characteristics

Effect of U/Cn on Turbine Efficiency, Velocity 143Compounded Stage

Conceptual TPA Design, LOX Pump 144

Conceptual TPA Design, LH2 High Speed Pump 146

ITA Wall Temperatures Based on Entrainment Fraction 148Model

xvi

:lJlI

LIST OF FIGURES. (cont.]_

Fi9ure No.

84

85

86

87

88

89

90

91

92

93

94

95

96

97

98

99

lOO

IOl

102

IO3

Film Cooling Requirements for 1200 Cycle Life atMixture Ratios from 4 to 7

Gas-Side Heat Flux Profile

Heat Transfer Computer Printout for CoFlow Regenerative

Cooling

Predicted Coolant Passage Temperatures, Downpass(CoFlow) Design

Predicted Coolant Passage Temperatures, Up-pass(Counterflow) Design

Injector End Predicted Coolant Passage Temperatures,

Up-pass Design

Predicted Coolant Pressure Drop

Comparison of Counterflow Design Temperatures atChamber Pressures of 20.4 and 34 ATM

Predicted Coolant Passage Temperatures, Up-pass(Counterflow) Design Using Nickel-200

Plug Contour for Oxygen Cooled Plug Analysis

Estimated Critical Heat Flux Characteristics of

Oxygen

Gas-Side Wall Temperature vs 02 Mass Flux, ModuleExit Plane

Gas-Side Wall Temperature vs 02 Mass Flux, DownstreamEnd of Plug

Two-Phase Pressure Drop Correlation

Plug Cluster Engine Dry Weight, Pc = 20.4 Atm

Plug Cluster Engine Dry Weight, Pc = 34 Atm

Moment Generating Capability for Hinged Engine ModuleConcept

Moment Generating Capability for Throttled EngineModule Concept

Moment Generating Capability for Hinged Panel Concept

Moment Generating Capability for Secondary InjectionConcept

150

151

i53

155

156

157

158

159

160

161

166

167

168

173

177

178

179

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181

183

xvii


Fi9ure No.

I04

I05

I06

I07

I08

I09

llO

Ill

ll2

ll3

ll4

ll5

ll6

ll7

ll8

ll9

120

121

122

123

Regen-Cooled 40:1 Module Plug Cluster EngineConfiguration EX02, Dwg. No. 1185978

Regen-Cooled lO0:l Module Plug Cluster EngineConfiguration EX02, Dwg. No. I185955

Minimum Modification ITA Plug Cluster Engine

Configuration EX02, Dwg. No. I185979

Plug Pressure Distribution

Plug Cluster Nozzle and Thrust Structure

Plug Nozzle Structure

Lightweight Module Mount Ring

Plug Cluster Engine Expander Cycle Schematic-Regen-Cooled Module and Plug

Plug Cluster Engine Gas Generator Cycle Schematic-Regen-Cooled Module and Plug

Effect of Pressure Drop on Equivalent OrificeDiameter

Plug Cluster Engine Expander Cycle - Uncooled Plug

Plug Cluster Engine Gas Generator Cycle - Uncooled Plug

Integrated Thruster Assembly (ITA), Dwg. No. 1162904

Regen-Cooled ITA, Dwg. No. I185963

Isothermal Cycle Test Data - Zirconium-Copper

Plug Cluster/Scarfed Bell Geometry (Pc = 20.4)

Plug Cluster/Scarfed Bell Geometry (Pc = 34.0)

Baseline Space Tug Engine Comparison

Plug Cluster Engine Sensitivity Summary

Plug Cluster Engine Concept Offers Many Features

Page

188

190

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202

203

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209

210

216

217

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228

229

254

255

26O

xviii

SECTION I

SUMMARY

A. STUDY OBJECTIVES AND SCOPE

The major objectives of this study program were to: (I) conduct apropulsion system analysis to assess the potential of the plug cluster engineconcept for the Space Tug baseline vehicle and nominal missions, (2) assessthe potential of utilizing an existing or high technology thrust chamber asa module for such a plug cluster application, and (3) identify the technologyrequirements for the development of a plug cluster engine.

To accomplish the objectives, the eleven task study program, summarizedon Figure I, was conducted.

Design criteria were obtained from the literature on Space Tug systems,on plug and plug cluster nozzles, on H/O thrust chambers and on H/O turbopumpassemblies.

Engine performance and envelope parametric data were established overa wide range of mixture ratios and engine geometry, using a plug clusterperformance model.

Subsystems of the engine were evaluated to determine their impact onthe design, and any limitations resulting from the utilization of the variouscycles were established.

Based upon the results of Tasks I through III and the study guidelines,three configurations and two cycles were selected to be carried into con-ceptual preliminary designs. The three configurations involved use of: (I)ITA modules, (2) minimum change ITA modules, and (3) regeneratively cooledmodules. The two cycles were the expander and the gas generator cycle. Inaddition to the cooled plug design, an uncooled carbon-carbon cloth plugdesign was evaluated.

Plug cluster engine design, performance, weight, envelope and oper-ational characteristics were evaluated for a variety of candidate clusterconfigurations (Figures 2, 3 and 4). The selected plug cluster engines werecompared with the engine candidates that were evaluated for the baseline SpaceTug. The comparison was based on mission performance, cost, life, and enginegeometry.

Upon completion of the first six tasks, an amendment was made tothe contract to address the "real world" problems of an actual engine.Lightweight engine structures were examined, with the AGCarb (carbon-carboncloth) nozzle extension providing a significant configuration improvement.

Techniques for providing thrust vector control for the plug clusterengine were evaluated, and module hinging appeared to offer the best potential.

O0

>-1.1-

E

0S.

4_

4-0OJ

GJ

N0

e--0

e"

e-0Ue"

.r=.

2

c-

c-w

4-Jt_

t,..'-

r,_

kL

3

Figure 3. Clustered Bell Nozzle Concept

4

If!I:

Figure 4. Scarfed Bell/Plug Cluster EngineConcept

The lightest weight configuration for the fluid systems, their com-ponents, and controls for a plug cluster was determined.

Analysis of the experimental cold flow data recently obtained onContract NAS 3-20104 (NASA CR-135229 "Plug Cluster Nozzle Flow Evaluation")

was made, and discrepancies in the data noted. The plug cluster engine per-formance methodology was modified to reflect the cold flow data. Engine per-formance calculated by this methodology rules out the large gap cluster con-figuration on a standard plug nozzle due to the poor aerodynamic flow condi-

tions. Optimum performance is achieved, however, through the use of a flutedplug formed from a cluster of large area ratio scarfed bell nozzles.

Throughout the entire study effort, basic data gaps and areas requiringtechnology work were identified.

B. RESULTS AND CONCLUSIONS

High vacuum performance is achieved with the low pressure plug clusterengine which makes maximum use of the large area available with the baselineSpace Tug. Low development and production costs for the engine are achievedthrough the utilization of existing developed technology. The combination ofhigh performance and low cost makes the plug cluster engine competitive withthe baseline Space Tug RLIO I_B engine and the higher pressure AdvancedSpace Engine, as shown in Figure 5.

The objectives of the program have been successfully accomplished.The fact that existing developed, long cycle life thrusters can be clusteredin various manners and numbers, allows the designer the flexibility to con-figure a large number of Orbital Transfer Vehicles (OTV) that operate atalmost any thrust level desired.

RLIO lib

MR 6

Pc (ate) 27.2

_E zosWE (Kg) 201

DE (m) 1.80

LE (m) I .4h ***

*I s (s) 460.6

n E (Is/I s ODE) 0.965

PAYLOAD(Kg)

Deploy 3740

Retrieve 1649

Round Tr|p 1001

Planetary 4845

5,000F /I

RLIO lIB

S.I. UNITS

PCE 300 PCE SOC ASE

6 6 6

20.4 34.0 136

885 8g4 400

219 194 163"*

4.32 3.36 1.07

0.61 0.94 1.28"**

464.4 466.6 469.3

0.945 Omq4g 0.966

3827 3q64

1721 1A20

1041 1098

501g 512 g

PCE 300 PCE 500

* Performance based on JANNAF simplified methodology

** Thrust/Weight ratio assumed same as for 66,723 N engine (Ref. 273

***Stowed length of deployable nozzle

4084

Igll

I150

5224

ASE

ENGLISH UNITS

RtlO lIB PCE 300 PCE 500

MR 6 6 6

Pc (psta) 400 300 500

_E 205 695 894

WE (Im) 443 482 428

D E (in.) 71 170 132

L[ (in.) 55"** 32 37

*I s (s) 460.6 464,4 466.6

nE (Is/I s ODE) 0,966 0,945 0.949

_ploy 8245 8436 8740

Retrieve 3136 3794 401 3

Round Trip 2207 2_94 2421

Planetary 10901 I 1065 II 307

io.ooo[:9.ooor

" _ 8,0001

RLIO lib PCE 300 PCE 500

AS£

6

2000

40O

404"*

42

44"**

469.3

0.966

9003

4212

2535

11518

AS[

* Performance based on JAN_F slmpli lied methodology

ee Thrust/Weight ratio assumed same as for _0,000 lbf engine (Ref. 27)

.**Stowed length of deployable nozzle

Figure 5. Space Tug Engine Evaluation

• 7

_Y f

1!1 !

SECTION II

I NTRODUCTION

A. BACKGROUND

Several analyses of the propulsion systems required for the SpaceTug vehicle have been conducted fn the past, but in each case, the studieswere conceived and conducted in a traditional fashion with primary con-sideration given to engines having conventional bell nozzles. The use ofunconventional nozzles offers a great deal of potential for high performance,long life, and flexibility in design, that had never been exploited nor evenstudied in these vehicle applications. In cases where unconventional nozzleswere considered, restrictive assumptions that were applicable only to thebell nozzles were arbitrarily imposed on the unconventional nozzles. As aresult, many options that an engine designer might have had in developingadvanced thrust chambers were ground ruled out of the studies. This restric-tive situation becomes particularly troublesome when low cost and reusabilityare required of the propulsion system in addition to high performance.

Space Tug vehicle application studies for the purpose of evaluatingcandidate propulsion systems have been based on fixed input conditions, suchas propellant combination, narrow mixture ratio range, and engine envelope,i.e., engine length and diameter. It is well known that the area ratio forconic section (bell shaped) nozzles varies in a direct relationship with nozzlelength and inversely with throat radius (E _ Ln/Rt). It is also well known thatan increase in propulsion system vacuum performance occurs primarily by anincrease in nozzle area ratio and is essentially independent of chamber pressure(Is _ _). Early candidate engines were limited in performance for a fixedlength application. There were four approaches available to achieve a higherarea ratio in this length: (I) high chamber pressure, (2) extendible nozzle,(3) multiple engines, and (4) conventional nozzle.

The approach ultimately selected to attain high area ratio was toincrease the chamber pressure to make the throat area smaller for the samenozzle length. This high chamber pressure then led to a specific set ofproblems (high unit heat flux, high wall temperatures, and small, high speed,high pressure turbomachinery) that must be solved to meet the high cycle liferequired.

What is overlooked in this approach is that the true diameter limitfor the engine installation, i.e., the vehicle diameter, has not been uti-lized by the conventional bell nozzle to arrive at a solution to the basicproblem. Unconventional nozzles, i.e., clusters of small thrusters arounda contoured plug, can utilize this dimension to arrive at engine designs whichfeature lower chamber pressures, with attendant lower heat flux, lower walltemperatures, longer fatigue life, and less critical turbomachinery.

From 1969 to 1974, the NASA sponsored a number of efforts to establishan adequate technology base for a cryogenic Attitude Control Propulsion System

for the SpaceTransportation System. The final design life goal for thethruster was50,000 cycles (pulses) and 5,000 deep thermal cycles. TheIntegrated Thruster Assembly(ITA) accumulatedthe best of the componentdesigns available andwas life cycle tested 51,005 cycles. Designs forhigher performing, regeneratively cooled thrusters were also establishedwith high cycle life capability, but not tested as extensively as the ITA.

Oneof the intriguing variations, therefore, in the application ofplug nozzles to a SpaceTug type vehicle and mission, is the possibilitythat existing developedor high technology thrust chamberscould be clusteredarounda plug nozzle of very large diameter. Thus, the primary problemsofa high pressure engine are completely avoided in exchangefor a differentset of problemssuch as clustered performance,base pressurization, andinstalled weight. Theengine designer then has a choice of problems to solveto best meet the needsof the given application, with cost comparisonsinvolving the two types of propulsion systemsalso being an important factor.

B. PURPOSE AND SCOPE

The feasibility of the clustered plug Space Tug is heavily dependenton the delivered performance and weight of the engine system, with the trade-off in performance versus the gap between module nozzle exits being a signi-ficant factor. It is the purpose of this study to conduct a propulsionsystem analysis to assess the potential of the plug cluster engine conceptfor the Space Tug baseline vehicle and nominal mission.

Plug cluster engine design, performance, weight, envelope, and opera-tional characteristics were evaluated for a variety of candidate clusterconfigurations. The selected plug cluster engines were compared with theengine candidates that were evaluated for the baseline Space Tug. The com-parison was based on mission performance, cost, life, and engine geometry.

C. GENERAL REQUIREMENTS

For purposes of this study, the engine design point for plug clusterengine evaluation was assumed to be that given in Table I, commensuratewith the baseline Space Tug requirements.

TABLE I. - PLUG CLUSTER ENGINE DESIGN POINT

Propel I ant Combi nati onMixture Ratio (nominal)Maximum Engine DiameterMaximum Engine Length(at engine gimbal, beyondbase of LOX tank)

Engine Cyclic Life(no factor of safety)Engine Thrust (nominal)

Hydrogen and OxygenO/F = 6.O447 cm (176 in.)139.7 cm (55 in.)

1200 firings

66,723 N (15,000 Ibf)

I0

I!I _!i

D. APPROACH

To accomplishthe programobjectives, a study involving eleven tech-nical tasks wasconducted. The results of the first three tasks were utilizedto select the configurations to be conceptually designedand analyzed tooptimize the plug cluster engine. Tasksconductedwere:

I. TASK I: Literature Analxsis

Significant publications pertinent to the conduction of thisstudy were reviewed and evaluated, including:

a. Space Tug system studies.

b. Plug cluster nozzle and plug nozzle experimental and analyticalstudies.

c. H/O thrust chambers of existing and high technology status.

do H/O turbopump assemblies of existing and high technologystatus.

2. TASK II: Parametric Engine Performance

A simplified plug cluster engine performance methodology wasestablished and performance maps were prepared to display the delivered spe-cific impulse in terms of the engine variables.

3. TASK III: Subsystem Evaluation

Base pressurization, engine cooling, and turbomachinery and powersubsystems analyses were conducted to determine any limitations inherent in thevarious engine cycles proposed for the plug cluster engine.

4. TASK IV: Preliminary Design

Preliminary conceptual designs of plug cluster engines were pre-pared for selected configurations and engine cycles.

5. TASK V: Plug Cluster Engine Optimization

Parametric system analyses of the plug cluster engine were conductedand tradeoffs were made in performance and engine weight to arrive at anoptimum set of engine designs. The technology requirements for such an enginewere defined.

6. TASK VI: Plu9 Cluster Engine Assessment

The plug cluster engine was compared with candidate Space Tugengines for several baseline geosynchronous and interplanetary missions.

II

7. TASK Vll: Lightweight Engine Structures

Structural techniques, designs, and materials were selected toprovide the lightest weight plug cluster engine concept for typical spaceapplications.

8. TASK VIII: Thrust Vector Control Analysis

Techniques for providing thrust vector control (TVC) for a plugcluster engine were evaluated and the best method selected.

9. TASK IX: Fluid Systems and Control Study

The lightest weight configuration for the fluid systems, theircomponents, and controls for a plug cluster engine were selected from anevaluation of several candidate systems.

I0. TASK X: Experimental Performance Data Evaluation

An analysis was conducted and an appraisal was made of the experi-mental cold flow data reported in NASA CR-135229 "Plug Cluster Nozzle FlowEvaluation". These results were compared with the performance predictions

in Tasks II, III and V. The methodology employed in Tasks II through V wasupdated and revised in order to reflect the experimentally measured effects

of gaps, fairings, tilt angle, and base pressurization.

II. TASK XI: Plug Cluster Enqine Optimization

The engine optimization obtained in Task V was revised to includethe results of the Tasks VII through IX analyses. The plug cluster assess-

ment conducted in Task Vl was revised accordingly.

12

:H| :T':

SECTION III

LITERATURE ANALYSIS

A. OBJECTIVES AND GUIDELINES

A literature analysis was conducted to provide background data on theSpace Tug system, plug cluster nozzles, H/O thrust chambers, and H/O turbo-pumps to be considered in the study. Pertinent information from the litera-ture was included in detail in the Task I Report (Unconventional Nozzle Trade-off Study - Monthly Technical Progress Report 20109-M-2, Task I - LiteratureAnalysis, Aerojet Liquid Rocket Co., Contract NAS 3-20109, September 1976).The Task I Report provided narratives on the reports containing data thatserved to allow evaluation of the plug cluster concept. The narratives includeda summary, the scope of work, results attained (pertinent figures and support-ing data), an assessment of the state-of-the-art, and the strong and weakpoints of the work. The bibliography is repeated in Appendixes B through Eof this report.

Specific information from the Task I Report that became the back-ground data for the study is summarized in this section.

B. SPACE TUG SYSTEM STUDIES

Assessment of the plug cluster engine concept as a Space Tug propul-sion system involves a multitude of factors, many of which have been pre-viously studied indepth for other Tug candidates. The system studiesinvolving the main engine propulsion have considered both storable andcryogenic propellants, interim upper stages, and full capability SpaceTugs. The literature search conducted in this study was concentrated onthe cryogenic, full capability Tug.

The envelope of the cryogenic Tug is constrained by the dimensionsof the Space Shuttle payload bay. The baseline Tug vehicle utilizes aCategory II RLIO engine with a two-position nozzle in order to conservelength. Typical engine data resulting from the study efforts indicate athrust requirement between 66,723 and 88,964 Newtons (15,000 and 20,000pounds force), and an engine mixture ratio between 5 and 6. Payload opti-mizes at the lower mixture ratio for engines with lower chamber pressure.

The selection of the RLIO engine over more advanced engines wasprimarily based upon DDT&E cost rather than the amount of payload delivered.The engine Isp increase was originally evaluated using a sensitivity of +41 kg(+90 Ib) of payload per second of Isp, and the engine weight increase wasevaluated using a sensitivity of -2.5 kg (Ib) of payload per kilogram (pound)of inert weight for the deploy mission.

I. Baseline Space Tug

The current (October 1974) NASA definition (Ref. I) of the base-line Space Tug is given in Tables II and III and in Figures 6 and 7.

13

TABLE II - BASELINE SPACE TUG CHARACTERISTICS SUMMARY (Ref. I)

VEHICLE DESCRIPTION

ENGINE Pratt & Whitney RL-IO'IIB(Retractable Nozzle)

ACTUATOR Hydraulic

APS SYSTEM 24 Hydrazine thrusters (25#)

STRUCTURE

Skirts - Graphite Epoxy/Aluminum Composite

Tanks- Aluminum Alloy/Elliptical Bulkheads

Tank Supports - Fiber Glass Struts

Thrust Structure - Fiber Glass Strut Truss

THERMAL CONTROL SYSTEM

Tank Insulation - Goldized Super floc

Active System for Fuel Cell

Heat Pipes for Other Avionics

PAYLOAD CAPABILITY TO GEOSYN-CHRONOUS ORBIT

Deploy 7926 Ibs

Retrieve 3396 Ibs

Round trip 2070 Ibs

AVIONICS SYSTEM

Antenna - Electronically steerablephased array

Platform - Strapdown

Power - Fuel Cell (2) plus Battery

Data Management - Data Bus

SC Retrieval Laser Radar

SC Deployment Inspect - TV

MAIN ENGINE PERFORMANCE

THRUST(LBS) Isp (SEC)

Full 15000 456.5

Pumped Idle 3750 434.7

Tank HeadIdle 157 377

VEHICLE CHARAC]ERISTICS

Length 30 ft

Diameter 14.67 ft

Dry Weight 5140 Ibs

Burnout Weight 5755 Ibs

First IgnitionWeight 56,779 Ibs

DeploymentAdapter &Shuttle Systems 1900 Ibs

Ground Liftoff Weight 58,679 Ibs

PAYLOAD SENSITIVITIES

DEPLOY RETRIEVALONLY

_PL-2.62 -1.38

?ws

_Pk0 0.23

_PL

-0.38 0

aPL83 Ib/sec. 59 Ib/sec.

/

14

TABLE III - BASELINE SPACE TUG WEIGHT BREAKDOWN

STRUCTURE

PROPULSION ANDMECHANICAL

THERMAL CONTROL

AVIONICS

10% GRO_FTHCONTINGENCYINCLUDING FASTENERS

TOTAL DRY WEIGHT

Weight kg_l_

895 (1,974)

611 (1,346)

200 (441)

418 (921)

212 (468)

UNUSUABLERESIDUALS

BURN-OUT WEIGHT

EXPENDABLES

PROPELLANT RESERVES

USABLE PROPELLANTS

274 (605)

FIRST IGNITION WEIGHT

ORBITER ACCOMMODATIONS(including 10% contingency)

GROUNDLIFT-OFF

2,336 (5,150)

2,610 (5,755)

248 (547)

136 (300)

22,760 (50,177)*

25,755 (56,779)

862 (1,900)

26,616 (58,679)

*Maximum propellant weight, propellant may be off-loadedto accommodate additional payload weight.

15

1.50-_

176.0DIA.

-i

174.50

LH2

TANK

/__ _ I / I_- 5S.... 110' (EXTENDED)

--360.01 (30 FT)

Figure 6. Baseline Space Tug General Arrangement and Size

-_ , . 110 _.

_IJE REFERE,;CE PL_qE ,

• L " " II

GIMBAL-,LF3P(II( "_-._W'lij ut!-Trnl_<o."o (I i

"" 55 =:I

A

Figure 7. Baseline Tug Engine

16

l)| l

Geosynchronous performance capability of the Space Tug is a func-tion of various vehicle characteristics. The partials for the deploy onlyand the retrieval only geosynchronous missions listed in Table II areexplained as follows:

PL

WS

An increase of Tug stage weight (dry weight plus unusablepropellants) by one kg (one Ib) reduces the payload that can bedeployed to geosynchronous orbit by 1.19 kg (2.62 Ib), and thatwhich can be retrieved by 0.63 kg (1.38 Ib).

PLUP

An increase of Tug usable propellant capacity by one kg (one Ib)increases the payload that can be retrieved from geosynchronousorbit by O.lO kg (0.23 Ib). In the case of deployment of amaximum weight payload _ PL/_ UP = 0 since the Tug already hasmore propellant capacity than can be utilized (i.e., propellantsmust be off-loaded to meet the Orbiter constraint of 29,484 kg[65,000 Ib] at liftoff for the Tug plus its payloads).

PL

W0

A one kilogram (one Ib) increase in weight of the equipment charge-able to the Tug but remaining in the Orbiter (such as adapterstructure and propellant fill and vent equipment) decreases theweight of payload that can be deployed to geosynchronous orbit by0.17 kg (0.38 Ib). This decrease comes about because of theOrbiter constraint for Tug plus its payload at liftoff (when theinterface weight is increased, propellant must be off-loaded tosatisfy the constraint).

3 PL

Isp

Increasing the main engine specific impulse by one second increasesthe payload that can be delivered to geosynchronous orbit by38 kg (83 Ib) and that which can be retrieved by 27 kg (59 Ib).

The baseline Space Tug'is composed of structures, propulsion andmechanical, avionics, and thermal control systems. The general arrangementand Size of the Tug systems are shown in Figure 3, and the weight breakdownis given in Table III. The thrust structure is an open fiberglass conicfrustrum truss with an aluminum gimbal block to interface with the engine.It is attached directly to the L02 tank with eight fiberglass epoxy strutsas shown in Figure 6.

The engine (RLIO Cat. liB) shown in Figure 7, is a derivative ofthe flight proven Pratt and Whitney RLIO engine. It provides a vacuumthrust of 66,723 N (15,000 Ib) and a specific impulse of 456.5 sec at a mix-ture ratio of 6.0, _ = 205:1. The life expectancy is 5 hours with 190starts. The overall stowed engine length is seen to be 140 cm (55 in.),where the gimbal point is 44 cm (17 in.) aft of the LO2 tank.

2. Engine Evaluation

A sensitivity study was conducted in Reference 2 to determine theoverall program impact when the Option 2 Category IIA RLIO main engine isreplaced with an advanced engine candidate, i.e., Category IV RLIO, Advanced

17

Space Engine (ASE), or the Aerospike (Figure 3). With the exception of the

Aerospike, the engine change effects are primarily engine related, i.e.,engine DDT&E cost, weight and specific impulse. The Aerospike engine pro-vides maximum Tug performance at an engine mixture ratio of 5.0, while theother engines maximize tug performance at an engine mixture ratio of 6.0.Therefore, a Tug using an Aerospike engine would have different tank sizes

than a Tug using the other engine candidates.

Results of this study (Figure 9) show that the Tug performanceincreases by lO to 20 percent with the use of advanced engines. For themission model used, the number of flights does not change significantly andthe fleet size does not change at all. The figure also shows that the totalprogram cost decreases with the advanced engines and the cost impact is dueprimarily to DDT&E cost (mostly due to the main engine).

C. PLUG AND PLUG CLUSTER ROCKETNOZZLE STUDIES

During the past twenty years, many investigations have been conductedin the field of unconventional rocket nozzles, and in the process, a largevolume of literature was generated. The most pertinent references on thesubject of plug and plug-cluster nozzles are listed in Appendix C.

The literature, reviewed in the Task I Report, describes experimentaland theoretical investigations of several types of plug nozzles generallyreferred to as annular-throat, discrete-throat, Aerospike, and plug clusternozzles. Inverse-plug or expansion-deflection nozzles were also discussedin the review.

In addition to plug nozzle performance in terms of thrust efficiency,specific impulse or velocity coefficient, plug wall and base pressure andheat transfer data were presented. The experimental data on thrust vectorcontrol methods applicable to plug nozzles were also reviewed. In general,a good agreement was found between the model cold-flow data and hot-flowH2/O 2 propellant test results.

The analytical methods discussed in the literature, are generallyadequate for the design and performance prediction of annular-throat isen-tropic plug nozzles, but inadequate for the analysis of nozzles whichdeviate considerably from the annular-throat configuration, such as plug-cluster nozzles utilizing bell modules. In such cases, authors of variousreports generally resort to empirical correction factors to account forshock wave interaction occurring at the module exit. These factors weredeveloped from testing of specific plug-cluster configurations and mustbe applied with caution to new plug concepts.

Most of the analytical and experimental studies were stimulated by thealtitude compensation aspect of plug nozzles, which is a desirable charac-teristic of nozzles for booster application. For this reason, the range ofmany plug variables was limited to the booster phase of rocket propulsion.The space tug vehicle operates in a vacuum at infinite pressure ratio andthe altitude compensation, which occurs at pressure ratios less than design

18

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value, does not apply. Here, the performance objective is to select the plug-cluster configuration that would produce the maximum specific impulse for aspecified nozzle length.

Review of the literature on plug and plug-cluster nozzles allows thefollowing general observations to be made concerning the advantages and dis-advantages of plug-cluster nozzles for Space Tug application:

ADVANTAGES:

(I) Plug cluster concept lends itself to modular approach, and fullutilization of available diameter.

(2) Thrust vector control can be produced by gimbaling or throttlingof individual modules or group of modules.

(3)

(4)

(5)

Design techniques for bell module design are well developed.

Plug cluster concept offers fail-operational potential for module-out or turbopump-out fail-safe modes, whereas present Tug propulsionsystems are only fail-safe.

Concept allows application of low pressure, long life propulsionsystem components.

(6) Concept leads to shorter equivalent engine length.

DISADVANTAGES:

(I) Shock wave interaction at the cluster discharge reduces nozzleperformance.

(2) Analytical methods are not available at the present time.

(3) Plug cluster engines are slightly heavier than single enginesof the same thrust level.

I. Plug Nozzle Performance Criteria

The plug and plug cluster nozzle literature indicates definitetrends in the performance of the nozzle as a function of the major designvariables. However, these trends can be misleading at the larger area ratios(_ > 80) and module gaps of this study. For example, plug engine performanceappe-ars to decrease significantly with the degree of truncation (i.e., thereduction in the ratio of plug length to isentropic plug length) as shown forthe ALRC data curve in Figure I0 (taken from Task I Report, pg. 107).

What is not indicated in the figure is that the tilt angle of theannular throat remains constant at 38 degrees, and that the loss in thrustfor a zero length plug is primarily a divergence loss. If it is assumed thata zero length plug does not turn the gas stream axially, the expected thrustefficiency would be CTlcOSO, or 0.78, which appears to be a valid extrapola-tion of the ALRC curve in Figure I0.

21

1.00

0.98

_ 0.96W

L_

_ 0.94

F-

0.92

0.90

cM e CT cos e

l.7 38° O.788

6.3 15° 0.956

4.9 18° O.927

I-I

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I I _1 I L, I I I i0 0.2 0.4 0.6 0.8

[]0

CONFIG.

ANNULAR

ANNULAR

CLUSTER

'(ZERO GAP)

I I1.0

PLUG LENGTH/ISENTROPIC LENGTH _ L/L I

Figure lO. Effect of Truncating the Plug on Thrust Efficiency

22

The loss in performance for an annular throat plug nozzle with asmaller tilt angle (Ref. 3) is shown in Figure I0 to be much less. The valuefor CTlCOSO of 0.956 is seen to be a close approximation to the experimentalefficiency, when zero flow turning is assumed.

The same trend is evident in the data for clustered modules on aplug (Ref. 4). The value for CTlCOSO is 0.927 for the assumed isentropic CTin the figure.

Another example of a literature trend involves the performanceloss due to gaps between module exits of a plug cluster engine. A typicalrepresentation is given in Figure II (Ref. 5). Experimental data (Ref. 4,pp. 11-53 and 11-84) showing the effect of fairings on gap performance aredepicted in Figure 12. Addition of the fairing is seen to improve the per-formance about 50% of the difference between the zero and one gap cases. Itis seen that CT drops significantly when the gap is increased. But this dropmay be due to the gap (loss of effective area and/or aerodynamic losses), or dueto the increase in tilt angle or change in base pressure. Figure 13 (data fromRef. 4, p. 11-39) shows the effect that can be attributed to the tilt anglewhen the plug length is held constant. Interpretation of the curves in FigureI0 requires superposition of data giving the module CT contribution (CTcosc)),the base CT contribution, and the contour CT contribution all versus the tiltangle. Such data are given in Ref. 4 but only for the baseline tilt angle.

The effect of tilt angle on base pressurization and thus CT, isdepicted on Figure 14 (Ref. 4) for zero plug length.

Base pressurization of the Aerospike annular plug engine amountsto 2.4% of the thrust as shown in Table IV (Ref. 6). Experimental data,giving the relationship between the base pressure and the base flowrate,are shown in Figure 15 (Ref. 7) for an earlier version of the engine. Figure16 (Ref. 7) depicts the nozzle thrust coefficient efficiency (CT) variationwith amount of base flow for the Aerospike. A maximum is seen to occur atabout .004 base flow.

A difference appears in this relationship when data for a plugcluster (Ref. 4) is examined (Fig. 17). The maximum now appears between1 and 2 percent (but no data points are shown between 0 and 2%). The aero-dynamic conditions are entirely different, however. Data for the plugcluster were obtained at (flowrates) pressures such that the wake might nothave closed on the plug. In vacuum, the wake will close unless the addedbase flow becomes excessive, causing flow separation.

2. Plug Nozzle DesiDn Criteria

Reference 4 describes the five geometric parameters that must bedetermined to completely define a plug cluster configuration: module arearatio (cE), number of modules (N), engine (cluster) area ratio (_E), gapdistance (S/De), and tilt angle (o). The equation relating these parametersis given as

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27

TABLE IV - NOMINAL PERFORMANCE,DOUBLF-PANEL AEROSPIKE ENGINESHOWINGBASE CONTRIBUTION

Nozzle Type

Engine Thrust, pounds

Engine Mixture Ratio

Area Ratio

Stagnation Pressure, psia

Injector Mixture Ratio

Injector Flowrate, lbm/sec

Hydrogen Injection Enthalpy, Kcal/mole

Oxygen Injection Enthalpy, Kcal/mole

ODIE Specific Impulse, Ibf-sec/Ibm

ODK Specific Impulse, Ibf-sec/Ibm

Divergence Efficiency

Boundary Layer Loss, Ibf-sec/Ibm*

Energy Release Efficiency

Base Specific Impulse**, sec.

Base Flow Ratio, Msecondary/Mprimary

Base Flowrate, Ib/sec

Base Pressure, psia

Base Thrust, Ibf

Engine Delivered Specific Impulse, sec.

Aerospike

25,000

5.5"I

200:1

I000

5.572

52.99

2.18

-I .005

498.5

497.3

0.9671

-17,93

0.995

6056.2

0.0019

0,I0

0.60

609.7

470.4

*AFBL/Minjector

** Fsecondary/Msecondary

28

i_iI !

r_

mm

i

F-:E

O

F-cEr_

&JJ

i,t

c_.

w

..JNNOz

0.0O60

0.0050

O, 0040

0.0030

0.0020 c

[

0.0010 (

m

ENGI TsOATAI• O [] MR = 5.O, c= 74.1, Pc = 1050,

Z_O NR : 3.5, E = 74.8, P : 900c

0 I I i i I0 0.01 0.02 0.03 0.04 0.05

SECONDARY-T0-PRIMARY WEIGHT FLOW RATIO, Ws/Wp

Figure 15. Base Pressure Versus Secondary Flow at Vacuum

29

l,C.)

(_}

r_z

C_)

F.-.Z

L.L.

0

_.J

N0Z

I00

98

96

94--

92--

90--

88

0

[]

[]

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HOT FIRING DATA

0 MR= 3.5

[] MR- 5.0

PREDICTED PERFORMANCE

----MR: 3.5

--MR : 5,0

I I I I I.01 .02 .03 .04 .05

WsFLOW RATIO, _-

P

Figure )6, Vacuum Thrust Coefficient Efficiency Versus Secondary Flow

30

lil_!i-

0

t',.

O

!

©

o>E-

0

0

U

E

0

.5..

a_

0

aJ

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-r-i,

^D '&N_I3IeI_I_OD X&I30"I_A

31

2EE

_M N [arctan (cos o tan 180)](Eq. I)

The choice of any three of the parameters determines the other two.

For a given number of modules, module area ratio, and gap dis-tance, the cluster area ratio and then tilt angle may be calculated. Withmodule and engine area ratios known, the tilt angle is obtained from a figuresuch as Figure 18 (Ref. 4), as the difference between the Prandtl-Meyerturning angles for the engine and module area ratios. A curve, illustratingthe relationship between cluster area ratio, number of modules, and tilt angle,for the case of 6/De = O, is shown in Figure 19 (Ref. 4). With a gap betweenthe modules, the amplification factor (_e/cM) will increase; correspondingly,the engine area ratio, as well as the tilt angle, will increase.

There are two regions that must be considered in designing thecontour of the plug (Figure 20A): (I) the expansion region of the plug, and(2) the transition region where the flows from the modules merge and mix toform an annular flow field. The method used in Ref. 4 to design the plug con-tour is as follows: (1) a single-expansion plug nozzle computer program

employing the method of characteristics is used to design a plug contour forthe desired cluster area ratio. This program provides a full-length plug

nozzle with the external expansion starting from a Mach 1 annular throat(Figure 20B). (2) A module is then positioned, as shown in Figure 20C, sothat the outer lip of the module coincides with the expansion corner of thesingle-expansion plug nozzle. Thus, the exit Mach line (corresponding to thecluster area ratio or Mach number) for both the plug cluster nozzle and thesingle-expansion plug nozzle coincide. (3) A smooth curve from the insidemodule lip is then faired into the isentropic contour.

D. H/O THRUST CHAMBERTECHNOLOGY

Hydrogen-oxygen thrust chambers that offer potential in a clustered plugconfiguration for the Space Tug application fall into two categories:(I) existing, or (2) demonstrated (high) technology status. All of the candi-date engines for the single-engine Space Tug can be correspondingly cate-gorized except for the existing RLIO that has been carried to operationalengine status.

The technology on small thrusters was recently reviewed by Gregoryand Herr (Ref. 8). Their paper covered the comprehensive program sponsoredby NASA-LeRC to provide the technology groundwork for the use of hydrogen-oxygen propellants in the Space Shuttle Attitude Control Propulsion System(ACPS) thrusters. Final reports on these projects were reviewed in Task Iof this study with the objective to independently assess the state-of-the-art of these thrusters and their components with reference to the feasi-bility of the plug cluster engine concept.

A prime candidate for the plug cluster engine is the NASA LeRC/ALRCIntegrated Thruster Assembly (ITA). Another high technology candidate is

32

1!!1 "-

$33_!_3Cl " (_) ::II_NV ONIN_IN1

!

(I)c-O

0

S,-

0

e-

0,r-

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33

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rd_

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34

_) 'OMN}:i V_IV H_,LSf'I"ID

(A)

..MACH ONE ANNULAR THROAT

/--MACH LINE CORRESPONDING TO_",_ "-.-_., MODULE EXIT MACH NUMBER_,_',,._ -'---..... /--SINGLE EXPANSION PLUG

"--.,/.. NOZZLE CONTOUR "

_'___..." "" -.. _ _ _ j_ LAST MACH LINE

(B)

PLUG LENGTH ,.j

(c)

Figure 20. Plug Contour Design

35

the Extended Temperature Range (ETR) ACPS thruster. Additional candidatesinclude regeneratively cooled thrusters.

A bibliography of pertinent reports that serve in the evaluation ofthe thrusters for the plug cluster assembly engine is given in Appendix D.

I. Integrated Thruster Assembly

The Integrated Thruster Assembly (ITA), Figures 21 and 22 (Ref. 9),is a flightweight GH2/G02 ACPS engine employing a spark initiated igniter.The nominal operating conditions are: 6672 N (1500 Ibf) thrust, 207 N/cm2(300 psia) chamber pressure, and a 4,0 mixture ratio, as given in Table V.The thruster has demonstrated a steady state specific impulse of 435 sec ata mixture ratio of 4.0 and 431 seconds at an O/F = 5.5 (Ref. 45). The ITAconsists of a premix triplet injector, a regeneratively cooled chamber, anda dump-film cooled throat and skirt; an ox rich torch type igniter andintegral exciter/spark plug; two igniter valves, and two main propellantvalves. The ITA S/N 002 was fired 42,266 times over 4200 full thermal cycles.A similar unit achieved 51,000 cycles in life testing at NASA/LeRC.

The scope of the ITA program included review of H2/02 ACPS tech-nology, design and fabrication of an optimized flightweight thruster, andtest firing to evaluate the thruster operation over a range of conditionssuch as would be encountered in a Space Shuttle application. The objectiveof the ITA program was to develop the technology for flightweight ACPSthrusters by investigating areas of unresolved technology such as: (I) chamber/injector life, (2) component interaction and optimization of a design tomeet the often conflicting requirements of steady state performance andcooling, (3) pulsing with cold propellants, (4) response time, (5) flightweight,and (6) long cycle life.

The results of the ITA program are as follows: (I) the ITA designis satisfactory, simple to operate, and has adequate life, (2) the igniteris very reliable, (3) chamber coolant part to part hydraulic characteristicshave no significant variations, (4) 51,000 pulses were demonstrated on asingle unit, (5) the predicted thermal cycle life of 65,000 cycles agreeswith measured temperature data, (6) fuel lead starts can result in damage,thus .01 to .02 sec oxidizer leads are used, (7) fuel lag shutdowns are pre-ferred, (8) the longest firing duration made with the ITA was 513 sec, and(9) the ITA weight was 6.895 kg (15.2 Ibm) exclusive of valves.

The ITA program demonstrated a lightweight, compact, high per-forming thruster which meets duty cycle and cycle life requirements. Theprimary problem area of the ITA thruster was the main propellant valves, whichstarted to leak after 20,690 pulses. The upper limit on operating pressurewas 348 N/cm2 (482 psia) due to the pressure limit of main propellant valves.Neither of these main propellant valve considerations should limit the use ofthe ITA results.

2. Extended Temperature Range Thruster

The Extended Temperature Range (ETR) Program (Ref. I0) involvedthe study of five cooling concepts (Figure 23) and the design, fabrication,

36

_! _i_

P_: 300 PSIA ....40 .... i_ ......

FIRED 51000T!ME_

Figure 21. Integrated Thruster Assembly is a Prime Candidate for the Plug

Cluster Engine

37

Aa0

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W

t_

e-I--

0

O

¢-(.3

l--

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38

TABLE V. - ITA DESIGN SUMMARY

Design Characteristics

Thrust

Chamber Pressure

Mixture Ratio

Pressure at Inlet to Valves

Fuel Flow R_te

Regen and InjectorFuel Film Coolant

Total

Oxidizer Flow Rate

Fuel Temperature

Oxidizer Temperature

Igniter Fuel Flow Rate

Core

Coolant

Total

Igniter Oxidizer Flow Rate

Igniter Core MR

Igniter Overall MR

Geometry

Throat Diameter

Exit Diameter

Chamber Contraction Ratio

Nozzle Exit Area Ratio

Chamber L*

overall Length

Overall Length (less exciter/spark plug)

Fwd End Clearance Diameter

Dimension of Cylinder Enclosing ITA

Weights (Design)

ITA (incl. Main Propellant Valves)

Main Propellant Valves

ITA (less valves)

Thrust Chamber (Incl. Insulation)

Injector

Igniter

6672 N (15001bf)207 N/cm L (300 psia)

4.0

276 N/cm 2 (400 psia)

247 g/sec (.545 Ib/sec)

65.6 g/sec (.145 ib/sec)

313 g/sec (.69 ib/sec)

1252 g/sec (2.76 ib/sec)

130oc (250OR)

208°C (376OR)

.726 g/see (.0016 ib/sec)

4.26 g/sec (.0094 Iblsec)

4.99 g/sec (.011 iblsec)

32.66 g/sec (.072 ib/sec)

45

6.55

4.88 cm (1.92 in.)

30.73 cm (12.1 in.)

5.5

40:1

43.18 cm (17 in.)

74.68 cm (29.4 in.)

61.37 cm (24.16 in.)

33.78 cm (13.3 in.)

74.68 x 36.32 cm (29.4 x 14.3 in. Dia)

14.016 kg (30.9 Ib)

7.257 kg (16.0 ib)

6.758 kg (14.9 ib)

3.933 kg (8.67 Ib)

1.887 kg (4.16 ib)

.939 kg (2.07 Ib)

Design Performance

Specific Impulse

Steady Sta_e

Pulsing @ MIB

MIB

Response (electrical slgna] to 90% thrust)

4266 N-sec/kg (435 ib=-sec/ib )

3923 N-<ec/kK (400 ib_-sec/Ib m)

222 N-sec (50 Ib-sec)

.050 sec

39

LH2

LOp. a. Gas/Liquid Injection

Upstream Valve FilmCooled Throat

b. Liquid/LiquidInjection Film CooledChamber

Figure 23. Tested ETR Candidate Propellant Thermal Management Concepts

40

and testing of two non-flightweight engine designs which are viable candi-dates for SpaceShuttle and Tugengine systems. OneETRdesign, using a24-element liquid/liquid injector, was fired 66 times as a full thruster atsea-level conditions, and lO times with a cooled chamberat altitude condi-tions, andwasdamagedafter four secondsduration. Theseconddesign wasa36-elementgas/liquid injector that was fired successfully 48 times at sea-level conditions, and 44 times with a cooled chamberat altitude conditionswith five tests of 20 secondsduration each. The operating point of bothengines is 5560N (1250 Ibf) thrust, 345N/cm2(500 psia) chamberpressure and4.5 mixture ratio with cryogenic propellants.

The ETRprogramsuccessfully demonstrateda non-flightweight36-element coaxial G/L thruster with a dumpcooled regenerative chamberand a Haynesnozzle over a chamberpressure range of 152 to 345 N/cm2(220 to 500 psia) and a mixture ratio range of 2.3 to 6.2 with fuel inlettemperature of 36 to ll6°K (64 to 208°R). A cumulative firing durationof 273 sec, including five 20 sec tests, wasmadewithout damage. Theigniter and valve capability, reliability, and durability were demonstrated.

The G/L thruster demonstrateddurations of 20 sec without damage,and a steady state performanceof 4266N-sec/kg (436 Ibf-sec/Ibm) with 18%fuel film cooling. A hydrogeninlet temperature of as low as 35.6°K (64°R)and as high as lll°K (200°R)was demonstratedin the G/L thruster. A widerange of operating conditions were tested. Combustionstability wasdemon-strated on all testing. Large amountsof thermal data were obtained.

3. Hydrogen-Oxygen Auxiliary Propulsion Engines

Technology for long life, high performing hydrogen-oxygen (H/O)rocket engines suitable for Space Shuttle auxiliary propulsion systems (APS)

were obtained in several NASA sponsored programs. Injectors, fast responsevalves, igniters, and regeneratively and film-cooled thrust chambers were

tested over a wide range of operating conditions and durations (Ref. II and12). A typical schematic of a thrust chamber that was tested is shown in

Figure 24.

The scope of the H/O APS programs included the screening ofcandidate cooling methods during analysis and design studies, and the

fabrication and testing of the selected designs. Design criteria and per-formance summaries are indicated for these designs in Table VI.

E. H/O TURBOPUMP ASSEMBLY TECHNOLOGY

The plug cluster engine concept is dependent upon the turbo-machinery subsystem design, performance and weight. Since the perform-ance of a conventional space engine is essentially insensitive to thelevel of thrust chamber pressure, pump discharge pressures can be low,

and consequently, the turbopump weight, which is then a small percentageof the total engine weight, is low.

For the plug cluster engine, pump weight optimization will dependupon the number of turbopump assemblies (TPAs) selected to feed the

41

|

J o

|

|

¢,_°Fw

E

c_

_r_

E¢0

cJ

It]:3

.C:Im

QJS-

or--l,

42

.... TABLE VI - APS CYCLE LIFE PERFORMANCE MATRIX (Ref. II)

PHASE I CHASE;_R DE_q]CNS AND CO.NI)ITT bN._

AMBIENT TEY}'EPATI;!_E .ROPLLL_U,TS

P , psia (N/cm 2)C

MR

Regen Chamber, 10% FFC

I , ibf-sec/Ibm

s (N-s/kg)

Nf

NfT

Film Cooled Chamber, 20% FFC

I ibf-sec/ibms' (N-s/kg)

Nf

Nf T

I00 (69)

44q

(4400) ,

3.5xi073.5xi04

455

3xiO _

2xlo 3

448

(4390)

9xlO 3

300 (207)

452

(4429)

2xI

443

(4341)106 _

9×i0 J

yl

3xI0_

2x10 J

430

(4214)

9xlO 3

500 (345)

455(4459)1.2xi0 _

447

4380)

PHASE II CHAMBER DESIGNS AND CONDITIONS

COLD PROPELLANTS

Regen Chamber, 9% FFC

Is , ibf-sec/Ibm(N-s/kg)

Nf

Nf T

Film Cooled Chamber

15%FFC

Is, ibf-sec/ibm(N-s/kg)

Nf

Nf T

20% FFC

Is , ibf-sec/Ibm(N-s/kg)

Nf

Nf T

106 _

2x105

lO_6>IO b

444

(4351)106 .

l.SxlO 4

441

(4322)106

105 -

438

(4292)

i iO@10 3

1

442

(4332)106 .

1.5xlO 4

438

(4292).

4xlO 5

105

435

(4263)106

105

Nf ffiThermal cyclic llfe for pulses of 200 Ibf-sec or less.

Nf * Thermal cyclic life for full thermal cycles."T

431

(4224)

106. .

l. SxlO 4

428

(4104)105

105

425

(4165_8xlO _105

442

(43_2)

7.5xlq;

4xi07

4XIN"

2xiO"

Firings >Z_50 Ib-sec totai i-ru!_e.

All designs provide 106 pulse capability for 50 ib-sec bit impulse.

43

thrusters (modules). Mission reliability, as well as the number of enginerestarts per mission, and the effect on chilldown propellant requirementsalso depend strongly upon the number of TPAs in the feed system.

An important factor in the selection of the number of TPAs is thegeometric size of the critical components such as bearings and seals.Miniaturization of these components must be avoided.

The turbopump selection further impacts the system payload capabilitythrough the suction pressure requirements, and the classical tradeoffof tank weight versus NPSH and chilldown propellant flow must be made.

A bibliography of reports on turbopump technology pertinent to theplug cluster engine Space Tug application is given in Appendix E. It isapparent that the TPA technology is of sufficient status to allow enginesystem and design studies to be conducted in a realistic manner.

I. APS Turbogumm__s

Small, high-performance L02 (Table VII) and LH2 (Table VIII) turbo-

pump assembly configurations were fabricated with each unit consisting of pump,turbine gas generator, and appropriate controls (Ref. 13). Development test-

ing was conducted on each type to demonstrate performance, durability,transient characteristics, and heat transfer under simulated altitude condi-

tions. Following successful completion of the development effort, two L02

turbopump units and one LH2 turbopump unit were acceptance tested. A weldfailure in the turbine manifold of one LH2 turbopump unit prevented its

acceptance. The test results on the L02 turbopump assembly correlated wellwith predicted performance, while the LH2 turbopump test results showed lowerthan anticipated developed head at the design point and in the high flow range

of operation. The lower developed head is attributed to higher than anti-cipated pump flow passage resistance from effects typical of small multi-

stage pumps. The results of this program have established a sound technologybase for future development of small, high performance turbopumps and gas

generators.

Assessment of the state-of-the-art of these turbopump configura-

tions shows the breadboard designs are somewhat heavier than desired for use

on an engine. Further design refinements would likely be required for

adaptation to engine installations.

EDM and casting methods were extensively used in fabrication.

Although some difficulties were encountered, the processes were evidentlyquite successful. The art of fabricating small turbomachinery componentswill undoubtedly develop further as their use is increased.

2. RLIO Turbopump Assembly

Reference 14 examined selected RLIO derived candidate engines for

the cryogenic Space Tug to define detailed engine system performance, mechan-ical and operational characteristics. A critical element evaluation estab-

44

i'i I-T

TABLE Vll - APS OXIDIZER TURBOPUMP PERFORMANCE REQUIREMENTS (Ref. 13)

Pump:

Turbine:

Turbopump:

Flow, m3/sec (gpm)

Inlet Pressure N/m2 (psia)

Developed Pressure, N/m2 (psid)

Inlet Temperature, K (R)

Energy Source

Exhaust Pressure N/m2 (psia)

Life, tbo, hrs

Operating Cycles

Start Time, sec

"ON" Time

"OFF" Time

Useful Life

Seal Leakage

Maximum Surface

Temperature

Turbine to Pump Heat Flow

6. 309 x 10-4 (I00)

137,895- 344,738 (20-50)

1.103 x 107 (1600)

92.8 - 103.9 (167-187)

O21H2

24.317 (35)

lO

lO,O00

1.5

2 sec to 600 sec

5 sec to 24 hrs

l0 years

Minimized

589 K (1060 R)

<52,752 Joule/hr(50 Btu/hr nonoperative)

<158,256 Joule/hr(150 Btu/hr operative)

45

TABLEVIII - APSLH2 TURBOPUMPPERFORMANCEREQUIREMENTS( Ref. 13)

Pump:

Turbine:

Turbopump:

F1ow

Flow

Developed Pressure

Inlet Pressure

Inlet Temperature

Energy Source

Exhaust Pressure

Life, tbo

Operating Cycles

"ON" Time

"OFF" Time

Start Time

Turbine to Pump Heat Flow

2.01 kg/s (4.5 lb/sec)

0.02902 m3/sec (460 gpm)

1.103 x lO7 N/m 2 (1600 psia)

124,106 - 344,738 N/m2 (18 - 50 psia)

20.8 - 25 K (37.5 - 45 R)

02/H2

2413.7 N/m2 (35 psia)

lO hrs

lO,O00

2 sec (minimum)

5 sec to 24 hrs

1.5 sec

158,256 Joule/hr(50 Btu/hr Static)

52,752 Joule/hr(150 Btu/hr Operating)

46

11_I_

lished the feasibility of various engine features such as tank headidle,pumpedidle, and autogenoustank pressurization and two-phasepumping. Thetank headidle and pumpedidle modeare attractive as a meansof achievingpumpchilldown with minimumloss of total impulse. The two-phasepumpingcapability relates to minimizing the weight of gas pressurants.

Four engines were investigated with chamberpressures from 400 to870 psia. The turbopumpassemblyconfigured for two-phasepumpingrequireda larger diameter inducer for the LH2pump,and a low speedboost pumpforthe LOXsystem.

47

SECTIONIV

ENGINE PERFORMANCE METHODOLOGY


The program objective was to prepare performance maps for the plugcluster engine concept, displaying the delivered specific impulse in termsof the following variables:

Combustion chamber pressure

Engine area ratioEngine diameterNumber of clustered modules (or thrust per module)Module area ratioModule mixture ratio

Engine mixture ratioPlug nozzle base pressure (or base flow rate)

The program requirement was to utilize simplified engine performance

methodology. Test data correlations for base pressurization and fairingcorrections were incorporated into the model. Nominal conditions for the

plug cluster engine were those of the baseline Space Tug as given in Table I.

A Task Report (Unconventional Nozzle Tradeoff Study - Monthly TechnicalProgress Report 20109-M-4, Task II - Parametric Engine Performance, AerojetLiquid Rocket Company, Sacramento, California, Contract NAS 3-20109,November 1976) was issued summarizing the data generated.

Upon receipt of the experimental cold flow data from Contract NAS3-20104, the engine performance methodology was revised. Performance maps

reflecting the experimentally measured effects of gaps, fairings, tilt angle,and base pressurization were generated. Calculations for the baseline caseindicated only a small performance improvement over that obtainable from alow area ratio (_ = 40) module. These data led to a reevaluation of the

plug cluster design and to the formulation of a design and consistent per-formance methodology.

B. MODULE PARAMETRIC PERFORMANCE MODEL

The approaches taken to define the performance of the plug clusterengine involve the establishment of the individual thruster (module) per-formance as well as the performance contribution from the plug nozzle exten-sion. Module performance is discussed in this section.

Module parametric performance analysis was accomplished using a

computer model constructed to meet the study's specific requirements. Itwas built upon the procedures specified by the JANNAF Liquid Rocket Per-formance Subcommittee (Ref. 15) and was a modification of a computer modelformulated for another engine study (Ref. 16). The JANNAF Subcommittee has

recommended two performance analysis methods. The standard procedure which

49

utilizes the best available analytical procedure is primarily used forsingle point performanceanalysis of existing engine systems. The secondmethodis a simplified procedurewhich utilizes design chart data and lowercost computerprograms. It is designed for the parametric analysis of enginesystemsandwas ideally suited to this study. Thesimplified method, there-fore, wasutilized.

Theprogramcalculates delivered moduleperformanceand the moduleenvelope as a function of engine thrust (F), chamberpressure (Pc), arearatio (EPS), mixture ratio (O/F), film cooling level, nozzle length (%Bell),and injector type. To accomplishthis wide-range, parametric analysis witha minimumcost, the JANNAFprocedureshavebeen expandedto include: (1) ODEand ODKIsp and C* data tabulations as a function of O/F, P F, and EPS,(2) injector design limits, and (3) envelope design data. _elivered moduleperformanceand envelopeare determined for any set of design and operatingconditions through the evaluation of the one-dimensional equ#librium (ODE)specific impulse and the appropriate performancelosses. The moduleenvelope is determined from the calculated performancelevel and the nozzledesign and chamberlength requirements andspecific operating conditions.A brief description of the methodsused to evaluate the aboveparametersfo I lows.

I. Performance Losses

al One-Dimensional Equilibrium (ODE) and One-DimensionalKinetic (ODK) Performance

The ODE and ODK Isp and C* are included in block data formin a subroutine. The data were calculated using the JANNAF approved ODK/

ODE computer program. A parametric evaluation of the ODE and ODK Isp andC* over a wide range of nozzle expansion ratios, O/F ratios, and chamberpressures was accomplished and its results are included in the evaluation

program. The ODE Isp is included in the computer printout under the headingISPT.

b. Divergence Loss

The nozzle divergence loss (% DL) is evaluated for Rao(Bell) nozzles using design charts similar to those presented in AppendixA of Ref. 17. Data from these charts are contained in block data format

in a subroutine which supplies the nozzle divergence efficiency and nozzle

length for a specified nozzle area ratio and % Bell. The divergence effi-ciency as a function of length and area ratio is determined from a method-of-characteristics computer program using the design technique developed

by Rao.

c. Boundary Layer Performance Loss

The boundary layer performance loss (% BLL) is evaluated

using the Design Charts presented in Appendix B of Reference 17. The DesignChart data are included in block data format in a boundary layer loss sub-

50

IllI!

routine of the computer program. Inputs to the subroutine include thenozzle area ratio and throat radius, chamber pressure, gamma (1.20), nozzleexit angle, CSTAR, and wall temperature ratio.

d. Fuel Film Cooling Loss

The fuel film cooling loss (% FCL) is calculated using thethermal exchange stream tube model from the ITA program (Ref. 9). Thismodel resulted in fair correlations of the ITA film cooling loss. The filmcooling loss was also evaluated as part of the Plug Cluster Module Demon-stration Program (Contract NAS 3-20107), and these results were incorporatedinto the Parametric Analysis Program that was used for engine optimizationin Task V.

e. Energy Release Loss

Two options were included in the program. The first optionassumes a fixed injector design (i.e., ITA) is utilized at all operatingconditions. The energy release loss (% ERL) is based on the empirical energyrelease performance loss determined from the ITA program and extended usingthe Gas/Gas mixing model developed under Contract NAS 3-14379 (Ref. 18).This fixed injector design results in a larger energy release loss withincreasing mixture ratio and decreasing propellant temperature.

The second option included in the Parametric Computer Pro-gram allows for development of a new injector design for each operatingcondition. In this case, it was assumed that the new injector could bedeveloped to produce an energy release efficiency of 99% which is comparableto the ITA design at nominal operating conditions (O/F = 4, Pc = 207, F =6672, ambient propellants). The 99% ERE option was utilized in all of theengine optimization studies as it represents the more realistic approach.

2. Module Performance

Performance and module geometric parameters for a fixed ITA typeinjector and a new injector design (ERE = 99%), fully developed for eachoperating point over the specified range of design and operating conditions,are shown in the following tabulation:

TABLE IX. - MODULEPERFORMANCEPARAMETRIC RANGES

Propellants: (I); Hydrogen (T = 139°K), Oxygen (T = 208°K) *Injector Design: (2); Fixed (ITA Type) and Variable (ERE Constant)Thrust Level: (4); 2224, 6672, 13345, 22241N (500-5000 Ibf)Chamber Pressure: (2); 20.41 and 34.02 ATMMixture Ratio: (5); 4, 5, 5.5, 6, 7Area Ratio: (4); 40, I00, 150, 200Film Cooling: (4); O, 15, 20, 25%Nozzle Length: (I); 75.5% BellTotal Cases = 1280

*ITA operating temperatures - the propellants for the Space Tug engine, however,are stored as liquids at their normal boiling point. The performance calcula-tions included in Figures 25-30 show the trends in performance, but are from6 to I0 seconds higher in ODE specific impulse due to the use of ITA conditions.

51

The parametric performanceanalysis results were tabulated in theTask II Report for the fixed injector (ITA) and optimized (ERE= 99%)injec-tor designs respectively. Included in each table are the estimated deliveredperformance,performancelosses, and engine envelope dimensionsfor 1280specific design points which cover the range of parametric conditions includedin this study. The calculated delivered specific impulse is summarizedas afunction of mixture ratio and fuel film cooling in Figures 25-27. Figure 25showsthe variation of delivered Isp with mixture ratio for the fixed injec-tor design. Figures 26 and 27 contain similar plots of delivered specificimpulse for the optimized injector design at chamberpressures of 20.4 and34 atm, respectively. In all cases, maximumspecific impulse is obtainedat the low mixture ratio (4.0) point for a constant value of fuel film cool-ing. Thein_Nuenceof mixture ratio is less, however, for the optimizedinjector design since its energy release loss is unaffected by the operatingmixture ratio.

Theeffect of area ratio on modulespecific impulse at an assumedconstant 20%FFCis illustrated in Figures 28 and 29 for the fixed and opti-mized injector designs, respectively. Modulespecific impulse increasesapproximately 10-20 sec over the range of area ratio included in this study(40-200). The effect of module thrust level on specific impulse is showninFigure 30 for an optimized injector design. With the assumptionof a con-stant energy release efficiency, specific impulse increases slightly(approximately I%) with increasing thrust level becauseof reducedkineticand boundarylayer performancelosses.

C. PLUGPERFORMANCEANALYSIS

Plug cluster engine performancehas beenshownto reach, as a limit,the performanceof an annular throat plug nozzle. To a first approximation,then, the performancecontribution of the plug nozzle portion can be estimatedby comparingplug cluster and annular plug nozzle performancedata. Therefore,the performanceof the annular throat plug is discussed in this section.

Plug nozzle contour and performancewasanalyzed by meansof a com-puter programbasedon theory developed in Reference19. This well knowntheory is valid for an annular throat truncated plug nozzle expandingfromMach1 to a desired Machnumber,ME, at the plug exit. The sketch of theplug nozzle control surface is shownin Figure 31.

In performing parametric calculations, the MachnumberMEand initialflow angle oE are chosenand the base pressure is assumedzero. (OEand MEare design parameterswhich determine the plug area ratio and truncated length.)The results are plotted as a function of plug area ratio (_E) and non-dimensional plug length (L/RE) with Machnumberas a parameteras showninFigure 32. It can be seen that the effect of truncation of an ideal plug oflength LI is to reduce the Machnumberleaving the plug. At a given plugarea ratio, the truncation can be defined as follows:

L/L1% = (L/RE/LI/RE) I00 (Eq. 2)

52

!!| ill

FIXEDINJECTORDESIGN(ITA)

Pc = 20.4 ATMArea Ratio = 40:1

F = 6672N (1500 Ibf)75%Bell Nozzle

460 -

440

"O"

v

420

E

u.r,-

"_ 400,,,-,,

E

U4"o

380

360

0FFc

?o25

I I I I I4 5 6 7 8

Mixture Ratio

Figure 25. Module Delivered Specific Impulse for a Fixed Injector DesignOperating at a Chamber Pressure of 20.4 ATM

53

OPTIMIZEDINJECTORDESIGN(ERE= 99%)

Pc = 20.4 ATMArea'Ratio = 40:1

F = 6672N (1500 Ibf)75%Bell Nozzle

460

440

v

_- 420

E

{.},r--W-

"G400

E

380

0FFC

!

360 , , I I I I ,i3 4 5 6 7 8

Mixture Ratio

Figure 26. Module Delivered Specific Impulse for an Optimized Injector DesignOperating at a Chamber Pressure of 20.4 ATM.

54

OPTIMIZED INJECTOR DESIGN (ERE = 99%)

Pc = 34.0 ATM

Area Ratio = 40:I

F = 6672 N (1500 Ibf)

75% Bell Nozzle

460 -

440

"C

420,-gE

u.r-

_ 400

E

380

360

FFC

I I I I I4 5 6 7 8

Mixture Ratio

Figure 27. Module Delivered Specific Impulse for an Optimized Injector Design

Operating at a Chamber Pressure of 34.0 ATM.

55

FIXED INJECTOR DESIGN (ITA)

Pc : 20.4 ATM

% FFC = 2O

F = 5672 N (1500 Ibf)

75% Bell Nozzle

460 - O/F

44O

420

400

E=U

380

_4

6

_7

360 I I I I40 80 120 160 200

Area Ratio, E

Figure 28. Influence of Expansion Area Ratio on Module Specific Impulse for aFixed Injector Design.

56

OPTIMIZEDINJECTORDESIGN(ERE= 99%)Pc = 20.4 ATM%FFC= 20

F : 6672N (1500 Ibf)75%Bell Nozzle

460

440

m-,-

_. 420

400

E

380

- O/F

360 I I I I

40 80 120 160 200

Area Ratio, E

Figure 29. Influence of Expansion Area Ratio on Module Specific Impulse for

an Optimized Injector Design.

57

OPTIMIZED INJECTOR DESIGN (ERE = 99%)

Pc = 20.4 ATM

% FFC = 20

O/F = 6.0

75% Bell Nozzle

460

440 -

IlJ

v

.2420 -

E

U.r,=,

q--

.g400 -

E

U

380 -

Figure 30.

I

I

_A_rea Ratio2OO

150

I00

4O

I ,, I I I L I4000 8000 12,000 16,000 20,000 24,000

I , I, I I I(I000) (2000) (3000) (4000) (5000)

Thrust, F, NClbf)

Variation of Module Specific Impulse With Thrust Level for anOptimized Injector Design

58

An

/

4

II

II

I/,,,

X

U

4-

_..

OtJ

"I[3

(IJ

NNOZ

tO

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U

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59

33 'O.L:I_I _a,_v 6n[d

60

The plug nomenclature is indicated in Figure 31. The ratio L/L I is commonlyreferred to as the isentropic plug length percentage. The isentropic pluglength is really a misnomer as all plug calculations are isentropic. Inactuality, an isentropic plug is a plug in which the flow field is axialor effectively one-dimensional. All truncated plugs have flow fields whichare not purely axial and thus contain divergence losses.

The calculated plug nozzle vacuum thrust coefficient can also beplotted as a function of plug length (L/RE), area ratio (E E) and degree oftruncation (L/LI) as shown in Figure 33. This figure illustrates the effectof area ratio and truncation on achievable vacuum thrust coefficient assumingan isentropic expansion along the optimized plug contour from an annularthroat.

The plug nozzle performance will be higher than indicated in Figure 33due to a finite base pressure acting on the plug base. In vacuum, or atdesign pressure ratio, the wake behind the base will be closed (see Figure 34)and the base pressure, without base injection, will be a function of theexpansion process along the plug wall. Experimentally obtained base pres-sures for both low and high area ratio plug nozzles follow the trend shownin Table X and Figure 35.

A relatively consistent correlation between PB and PE is evidentfrom Figure 35, where the base pressure obtainable is defined as that in whichthe nozzle separation criteria holds. Base pressures above the nozzle separa-tion criteria (PB/PE > 3.6, Ref. 42) cause an enlargement of the wake on theplug base, and thereby reduce the effective area ratio obtainable with a plugnozzle.

The data on plug nozzle base pressurization do not lead to a correla-tion between the base flow rate required to maintain a base pressure relation-ship, such as PB : 2.5 PE. In some cases, no bleed flow was required. Inother cases, percentage flows as high as one percent were required. Testingin facilities with finite volumes sometimes leads to conditions whereinwake closure is not achieved, unless a vacuum is pulled on the base regionat the start of the experiment to "snap" close the wake. This phenomenacould very well explain some of the variation in test data concerning basepressurization of truncated plug nozzles. The assumption was made, therefore,to utilize a bleed flow of 0.2% (see Table IV) of the engine flow, which isconsistent with the latest test data on the Aerospike where the wake is closed(Ref. 6).

D. PLUG CLUSTER PERFORMANCEANALYSIS

The selection of a performance model for the plug cluster engineproved to be the most difficult task in the study. This was due to thefact that the accepted, somewhat empirical, approach was based on utilizinga computer program for an annular plug with all external expansion. Experi-mental data indicated that the performance for a cluster of modules (internalexpansion sections with discrete throats) could be very closely approximatedby the annular plug model, providing the gap between the modules was close to

61

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65

zero. At the start of the study, no method was available that could adequatelyrepresent a plug cluster where the gap was much greater than zero.

Several approaches were taken before a satisfactory plug cluster engineperformance model was formulated. The initial approach relied heavily ondefining the module performance accurately, and included the contributionprovided by the plug. The plug contribution amounted to an area ratio increasefactor and a base pressurization thrust component. An empirical performanceimprovement was added to account for the use of fairings in the gap betweenthe" modules.

The second approach incorporated the experimental results from ContractNAS 3-20104 into a revision of the initial model.

The final approach is based entirely on the JANNAF simplified pro-cedures for bell nozzles. The method provides a straightforward estimate ofthe performance for plug cluster engines. The method is possible because theengine is envisioned to be formed from a cluster of high area ratio bellnozzles with zero gap (cf. Figure 4). A baseplate is provided, and the bellsare scarfed. The plug cluster engine, therefore, resembles a cluster ofmodules, with large gaps, placed on a fluted plug. Since the aerodynamicsof the flow from each module is identical to that for a scarfed bell nozzle,the JANNAF simplified procedures provide an accurate representation of theperformance, providing the base contribution is included.

The progressive effort in defining the performance models was bene-ficial in obtaining an understanding of the weaknesses in some of the designapproaches, and this understanding led to the formulation of the optimum plugcluster engine design configuration. Therefore, the rationale for each modelwill be summarized in the following sections to document the resulting per-formance associated with each design philosophy.

I. Plug Cluster Design Constraints

The nomenclature used to describe the plug cluster system is shownin Figure 36. Geometric constraints for the cluster have been identified(Ref. 4) in an equation which relates the area ratio amplification factor(_E/EM) to the number of modules, the module spacing, and the module tiltangle:

c E

EM 1 [ (I + 6/D e) cos C)]___N sin [tan -I (cos o tan )7 + cos o I(Eq. I)

where:

cluster area ratio = 4 (RE2/Dt 2 N)

module area ratio = (De/D t)

N = number of modules

66

la,J

I

q-0

_ta

,p- -I_

°r-

.%

,Ic

t

Z

I-ClJ

67

De =

Dt =

RE :

gap between adjacent modules

module exit diameter

module throat diameter

cluster exit radius

A solution of the equation for a constant tilt angle of I0 ° isshown in Figure 37 to illustrate the cluster geometric constraint. As shown,the cluster amplification factor increases with both the number of modulesand the gap between modules. Physically, for a fixed module size, the clusterradius (BE) increases as the number of modules increase so that the exitarea (REz) increases at a greater rate than the cluster throat area (NAt).Increasing the module tilt angle will decrease the amplification factor onlyslightly because of the reduced exit radius due to the module tilt.

A cluster of Bell nozzle modules around a truncated plug generatesa complex three-dimensional non-isentropic flow field, which is presently toodifficult to describe with a simple model (cf. Figure 34). However, someuseful insight into the flow problem can be realized by assuming that thegas expands isentropically from the Mach number at the module exit to theMach number at the cluster exit. Under this assumption, the tilt angle (c))is equal to the difference in the corresponding Prandtl angles.

o = v E - v M(Eq. 3)

where:

vE = Prandtl-Meyer angle of plug exit Mach number ME

I ]I/2 _ )I/2= (y__)I/2 tan-I (y-Iy_TT)(ME2 - I) tan -I (ME2 1 (Eq. 3a)

vM = Prandtl-Meyer angle of module exit Mach number Me

= (Y+l)I/2y-I tan-I ( ) (Me2 - 1 - tan (Me - I) (Eq. 3b)

The allowable plug cluster configurations are thus defined by com-bining the geometric and Prandtl-Meyer constraints [Eqs. (I) and (3)]. Theseresults are shown in Figures 38, 39, 40, 41, 42, and 43, respectively, formdoule gaps (6/D E) of 0.0, 0.5, 1.0, 2.0, 3.0, and 4.0.

E. PLUG CLUSTER PERFORMANCEMODEL I

The initial plug cluster parametric model of this study approximatedthe performance of the plug cluster system by the performance of the moduleplus the added thrust generated on the exposed plug and plug base. Thisapproach appears to be justified since the sum to the thrus_ generated bythe modules comprises 95% or more of the total plug cluster engine thrust.(The result is valid for the Space Tug engine cluster system which utilizes

68

Ii| i|i

MODULE TILT ANGLE, B = lO°

20

GAP BE_EE: MODULES I /

>

_ 10

0 ,5

0

lO 15 20

NUMBER OF MODULES, N

Figure 37. Variation of the Cluster Amplification Factor With Number of Modules.

69

y=l.2

50O

_S/De = 0

_M-- 250 2OO

j /40O

300

._ 100

200

4OI00

o

5 I0 15 20


Figure 38. Allowable Plug Cluster Design Conditions for Modules with Zero Gap.

70

i!_111

y=l.2

(S/De" = O. 5

l.I.l

500

400

b)

F-

,:Ci,i

U'}

3OO

2OO

I00

_M = 150 I00

I

0 = I0°

14°o

5 I0 15 2O


6O

4O

2O

Figure 39. Allowable Plug Cluster Design Conditions for Modules with 0.5 GAP

(6/De).

71

CLUSTER AREA RATIO vs NUMBEROF MUDULES

72

8OO

700

600

500

d

400

W

.-J

300

200

I00

Fi gure 40.

c = 160M 120 I00

-y,= 1.2

6/De = 1.0

8O

6O

0 = 14° 18 22° 26°

4O

2O

0 5 I0 15 20

NUMBEROF MODULES, N

Allowable Plug Cluster Design Conditions for Modules with 1.0 GAP

(_ID e)

Ill 'I

m,*L_I'--

.-If,..)

CLUSTER AREA RATIO vs NUMBEROF MODULES

cM = 60

800

700 --

6OO

500

6/D e 2.0

Y/

400 2O

/

Io 32° 36°O:

i

0 0 5 lO 15 20

300

2O0

100

NUMBER OF MOOULES, N

Figure 41. Allowable Plug Cluster Design Conditions for Modules with

2.0 Gap (6/De)

73

74

80O

700

600

5OO

w_J

d_- 400

e_

m 300

..J(..)

200

100

Figure 42.

CLUSTER AREA RATIO vs NUMBER OF MODULES

_M = 40 30

y= 1.2

_IDe_ = 3.0

38°

34°

42 °

46 °

20

10

5 lO 15 2O

NUMBER OF MODULES

Allowable Plug Cluster Design Conditions for Modules With 3.0

GAP (_/De)

UII

CLUSTER AREA RATIO VS NUMBEROF MODULES

800

7OO

6OO

LaJCO

- 500I--,4I--"

,,- 400I'--

..J

_00

20O

100

_M,_40 30 20

//

/

O = 34_8 ° 42o460 50 °

5 I0 15 20


Figure 43. Allowable Plug Cluster Design Conditions for Modules with 4.0

GAP (a/D e )

75

relatively high area ratio modules (40-200) and operates in a vacuum environ-ment. For booster systems, the module performance is about 91% of the totalengine thrust.) Thus, the module performance contribution has a one to twoorder of magnitude greater effect on the total engine performance than the

exposed plug contribution. In other words, a I% error in the module thrustcontribution will result in approximately a I% error in the plug clusterengine performance, while a I0% error in the exposed plug thrust contribu-tion will result in only a 0.5% or less error in the plug cluster engineperformance.

Based on this criteria, the following assumptions were made in

developing Plug Cluster Model I:

(1) Module performance is based on the JANNAF Simplified PerformanceEvaluation Procedure and thus includes the effects of operating conditions

(O/F, Pc, propellant temperature), expansion kinetics, boundary layer losses,fuel film cooling losses, and incomplete energy release performance losses.The exposed plug thrust contribution can be estimated using the plug designCF curve (Figure 33) and a base CF contribution from an empirical correlation.The CF curves are based on isentropic, constant gamma (perfect gas, frozen

flow) flow conditions and are used only to ratio the total engine performanceto the module performance. The throat of a module was assumed to coincide withthe throat of the annular plug.

(2) For an isentropic plug (i.e., L/LI = I00%) the module tilt angleis determined from the Prandtl-Meyer angle difference for an expansion from

the module exit condition to the plug exit condition [Eq. (3)]. The module

tilt angle is assumed to decrease linearly with the plug length as theexposed plug length is reduced from the isentropic value so that the tilt angleis zero when the exposed plug length is zero (i.e., there is no module tiltfor a zero percent plug).

(3) No correction is made to the exposed plug thrust contribution forthe effect of discreet module throats as opposed to the annular continuous

throat configuration assumed in the calculation of the plug performancecurves. This assumption is best for a large number of modules and smallmodule gaps and worst for a small number of modules and large module gaps.The discreet throat effect must be determined from experimental data.

These assumptions lead to a reduced cluster efficiency with increasing

module gap which is on the same order as the values reported in the literature.This effect is shown in the following derivation:

FC F Plug with Gap ._

CT with Gap LCFI Plu'g"with Gap-J

CT without Gap =F-CF Plug without Gap_

J(Eq. 4)

76

I!!I _

CT with GapCT without Gap

FCF with Gap I F CFI without Gapl

: m ]cFwit,outG p]LcFI

where it is assumed that the delivered CF ratio (CF with Gap/CF withoutGap) is approximately 1.0 and the ideal (CFI) ratio is a function of theplug area ratio increase with increasing module gap.

A special case for the comparison of gap effects can be obtained byengine-out operation of a module cluster without any module gap (i.e.,constant module tilt angle, module area ratio, gap 6/DE = 1.0 with everyother module out). Under this condition, the plug area ratio will increasebecause of the reduced module throat area which will have the effect ofreducing the plug cluster performance efficiency and total thrust withoutmaterially changing the delivered specific impulse. Thus, for all otherdesign parameters constant except the module gap, the plug area ratio can bedefined as follows:

Ewith Gap : Ewithout Gap (I.0 + 6/D E) (Eq. 5)

and,CFI at c = cE without Gap

CFI at _ = (_E without Gap)(l + a/DE)(Eq. 6)

The efficiency ratio calculated in this manner is shown in Figure 44.Note that the efficiency ratio increases as the engine area ratio increasesbecause of the diminishing effect of area ratio on performance. The generaltrend in the curve can be compared with the trend in Figures II and 12 for aI0 and 20 percent plug of 15 to 30 area ratio.

I. Model I Calculation Procedure

The Plug Cluster Computer Program Model I is set up to calculateperformance based on a module arranged in a specific cluster configuration.The input consists of a desired gap, module, engine area ratio and numberof modules for an isentropic plug as defined in Figures 38 through 43. Asthe plug is truncated, the number of modules and gap are held constant whilethe engine area ratio and tilt angle vary to accommodate the geometricrequirements imposed by Equation I. Physically, this has the effect ofmoving the cluster on the plug (by the amount Re sin o) as the plug isshortened (Figure 45).

The working plug forward boundary is at L0 where:

= + Rt) sin oL0 LM cos e - (Re (Eq. 7)

The degree of truncation of the plug is defined in terms of theisentropicplug length (L/LI). Plug truncation lengths are chosen in the program based

77

NOTE: Assumes CF Gap = CF Without Gap

1.00

0.99r_

C_

O

,r,- "_ 0.98F--

E0

u 0.97i,

"G,r,-

4- 0.96i,i

¢:I

0.95

0.94

CT I at c = CE without gapwith ag_a]z__ = CFI E

LCT without gap CFI at c E CE without gap (l + 6/D E

I I I I I0.2 0.4 0.6 0.8 1.0

EngineArea Ratio

E

E

500

200

100

4O

2O

Gap Between Modules, _/D e

Figure 44. GAP Efficiency Factor for Constant Plug Cluster Performance.

78

c_J

OD

l,l v

_J

(D

°i.==

u_

-IJ

I.--=I

_J

--I

.=_

ffJ

3

Z

"00

S-

¢J

O

e"or,,

0r"

0

e-U

Ill

79

on a percentage of isentropic lengths (5, I0, 15, 20, 25, 30).

fol 1ows :The plug-cluster vacuum thrust coefficient is expressed as

[ccrE , cT0j + Cr( /DE: ol(Eq. 8)

where:

ACFB

ACFF

CFE = Plug-cluster thrust coefficient

CFM = Module thrust coefficient calculated by means ofJANNAF Simplified Analysis

CFpL : Thrust coefficient of a plug of length L/L I

CFp0 = Thrust coefficient of a plug of length Lo/L I

= Increase in thrust coefficient due to base pressure

= Increase in thrust coefficient due to fairings

CF= Ratio of thrust coefficients with gap [assumed

CF (6/D e = O) equal to 1.0 for this analysis as per Eq. (6)]

The value of CFM is obtained from the module parametric model and is equal to:

CFM = FM/Pc At (Eq. 9)

The values of CFPL, CFpo, and the isentropic plug length are obtained byinterpolation of Figure 33. The contributions from the fairings and basepressurization are obtained from empirical relationships to be described.Finally, the thrust, total mass flowrate and specific impulse of the plugcluster system are calculated as follows:

FE = CFE Pc N At (Eq. I0)

mE = mM M + mBase (Eq. II)

IsPE= FE/mE (Eq. 12)

8O

2. Model I Fairing Correction

The performance methodology indicates that the calculated engineefficiency (Figure 44) reflects the same trend as the experimental data(Figure II) in showing a loss as a function of module gap. Therefore, itappeared reasonable to apply a fairing correction (ACFF) to Eq. (8) toobtain the predicted delivered performance for the plug cluster engine.

Data for two different gaps (6/De = 0.185 and 1.0) from Ref. (4)(p. II-53) are given in Table XI (see Figure 12). Utilizing these data, an

TABLE XI. - GAP PERFORMANCE (CT) WITH FAIRINGS (Ref. 4)

Configuration a/D E : 0 6/D E = 0.185 6/D E = 1.0

Plug Length = 10%No Fairings 0.966Straight Fairings

0.950 0.9140.954 0.940

equation (curve fit) can be written of the form

CT (without gap) - CT (with gap)

CTF X + CT (with gap) (Eq. 13)

where:ACFF = [CTF- CT (with gap_ CFI(Eq. 14)

and where X is equal to 4 and 2, respectively, for 0.185 gap and 1.0 gap.That is, the fairing correction becomes larger as the gap is increased.

Since no data were available on fairings for gaps greater than one,Eq. (13), with X equal to 2, was utilized for all of the gap calculations

(6/De = l to 4).

3. Model I Base Pressurization Correction

The base pressure can be 2.5 to 3.6 times the static pressure ofthe exhaust gas for the fully expanded plug as shown in Table X. This pres-sure is recognized to be the standard separation criteria (Pe > 0.4 P ambient)for DeLaval nozzles. The achievement of such a base pressure may require afinite mass flow into the base, the amount being presently determined fromexperiment. Base pressurization of the Aerospike (Ref. 6) annular plugengine amounts to 2.4% of the thrust as previously shown in Table IV. Uti-lization of these data for the 200:1 area ratio Aerospike plug allows thedevelopment of the equation.

Wbase = K Pbase Abase (Eq. 15)

81

where the constant K : 1.731 (1.695 x 10-4 ) was derived (Ref. 6) from Wbase =0.045 kg/s (0.I0 Ib/s), Pbase = 0.041 atm (0.6 psia) and Abase -- 0.634m 2(983.1 in.2).

Since the nozzle wake is closed, Eq. (15) would be expected to holdforsmall changes in the base area and/or tilt angle. It has been assumedto be valid for the larger area plugs of this study, but requires verification.

4. Model I Plug Cluster Engine Delivered Performance

Parametric performance data are given in Figures 46 through 50 forModel I engines. The figures include the module losses computed by JANNAFprocedures, where the nomenclature is: ODE - one dimensional equilibrium,KL - kinetics loss, DL - divergence loss, BLL - boundary layer loss, andERL - energy release loss. These losses are indicated by a bar giving atotal loss of about 20 seconds in Figure 46. Because the plug nozzle isdesigned to turn the module exhaust, the module DL term is assumed zeroin the plug cluster performance calculations. The true loss may be betweenthat of zero and the module loss shown, giving an uncertainty band equiva-lent in thickness to the DL band shown for the module. The plug length wasmaintained essentially constant to be more representative of a practicalapplication.

The lower performance line shown represents the engine performancefor just the modules and truncated plug. Improvements provided by fairingsand base pressurization increase the cluster performance as shown. Forexample, expansion (Figure 46) of the EM = 40 modules on an EE = 72 plug isseen to provide a 5% (23 second) improvement in engine performance. Themethod of combining the module and plug nozzle performance appears to beoverly optimistic for the zero gap configuration, as the indicated lossesare less than the module loss bar. At large gaps, the delivered performanceappears correct, as the difference between the ODE line and the deliveredline is equal to or greater than the module loss bar.

The base correction for the zero gap point in Figure 46 amountsto 1.5 seconds, or 0.3,%. The maximum base pressurization correction shownat E _ 400 amounts to about seven seconds, or 1.5%.

No fairing correction is taken for a gap of zero, but for positivemodule gaps, the previously presented method was utilized. In Figure 46,the maximum fairing correction amounts to about nine seconds (2%) atE _ 400.

The gap = 3 point of Figure 46 shows engine losses considerablygreater than those from the module alone. These differences can be attributedto gap and truncation terms in addition to the conventional losses.

In Figure 47, the (16%) film cooled module engine performance lossesare described. The film cooling performance loss (2.7%) is seen to have amajor impact on the engine performance.

82

_,_ -T _.-

6o'}

I

u_

r_E

"G

ct_

490

48O

470

46O

450

440

43O

Regeneratively Cooled ModuleMR = 5.5 N = I0

ODE

Module

Gap 6/De 1

Del i vEred Performance

Base CorYection

ModuleLosses

// /

. KL

444, ERL

///

BLL/

/.--44DL

,',,,\K

5O

Deli vered

I ,0

Fairing Correction

2515

Plug (No Correction) lo=_LL I

Plug LengthEssentiallyConstant

I l i J 1 J I200 400 600 800

Plug Cluster Area Ratio

Figure 46. Plug Cluster Engine Performance Summary at Pc : 20.4 ATM and EM = 40

83

49O

480

470

46O

6_J

I

450%E

or--4-

_ 440CI.

430

42O

410 [

Fi gure 47.

Film Cooled Module _'3

M = 40 MR = 5.5 N OD=IE_IO 2

Module _f_'Fuel Film Coolinghap = i

/,I

Module I Delivered Performance

lL°s sesj_z"/].,._ ODEI _

- - . ' / _i/ralrlng borrecl:lon ji

BLL 50/// 25

_-_ 15L__

I0 = % El\\\\ Plug (No Correction)\\\\

Plug LengthFCL Essentially

_ _ Constant

\\\N \ \

/DL/ / /]" _ _ Delivered

I I I i I i I i I0 200 400 600 800


Plug Cluster Engine Performance Summary with ITA at Pc= 20.4 ATM.

84

!TTTIT-

490Regereratively Cooled Module

MR = 5.5 N = I0'3

48O

47O(]J

I

U'}

_ 460

(_},iP-,

a.

m

i

ModuleLosses

• |s! ,|_L. ,

450

BLL '

440

2L 2i."x\\\

430

Module

Gap 6/D e

ODE

ODE

Deli vered

Plug (no correction) 1 L

I0 %IT


I l I , I i I i I0 200 400 600 800


Figure 48. Plug Cluster Engine Performance Summary at Pc : 34 ATM and _M : 40

85

U

!

(3.

E

(.I

q.-

"Gr_

cz)

480

47O

46O

490 --

ModuleLosses

ODE

/''I

450ERL [

/ / /1/

BLL I

440

FCL I

430 ' ....

k/ DL ,4////i

420

Film Cooled Module

EM : 40 MR=5.5 N=IO

ModuleGap 6/Dc

5O

Del i vered

16% Fuel Film Cooling

Del i

Base Correction

Fairing Correction

215

Plug (No Correction) L

I0 %L-II



Figure 49. Plug Cluster Engine Performance Summary with ITA at Pc = 34 ATM.

86

I!! :I

Regeneratively CooledModuleMR = 5.5 N = I0

U

I

%E

u

"G

490

480

470 --

460 --

450 -

440

Module Gap

f/De :

//7" KL

ERL

//////BLL

//////

DL

ODE

o

L

30 Plug (No Correction)15 12 : % FL--


Delivered

I , I , I , I j I0 200 400 600 800

Plug Cluster Ratio Area

Figure 50. Plug Cluster Engine Performance Sun_nary at P = 20.4 ATM and _M100. c :

87

Figures 48 and 49 depict the proceding two cases at the higherchamber pressure of 34.0 atm (500 psia). A corresponding case for EM =I00 modules, is given in Figure 50.

Examination of the data presented in Figures 46 and 50 for Model I,shows that the clustered plug performance improvement for a cM = 40 and aCM = lO0 module is 7.8% and 5.7%, respectively, at the large area ratios ofthis study.

F. ANALYSIS OF EXPERIMENTAL PLUG CLUSTER DATA (NAS 3-20104)

Module-plug cluster performance Model II, presented in the next sec-tion, incorporates the results of the cold flow experiments conducted underContract NAS 3-20104 (Ref. 46). Prior to the use of these data, however, itwas a requirement of this contract to make an appraisal of the data and tocompare the results with those available in the literature. This sectiondocuments the work performed to analyze the test data.

Tests were performed on twelve different cluster configurations shownin Table XII. Freon or air were used as the test medium to determine theeffects of gap, module area ratio, cluster area ratio, fairings, fences,tilt angle, and base pressurization on cluster performance. Validity ofthe air data is questionable because condensation shocks could have had aneffect on the results. In addition, the applicability of the air data isquestionable because the tilt angles and module-match points were designedbased on the use of Freon. Figure 51 shows that the change in the ratioof specific heat capacities has a strong influence on these design parameters.

A comparison of the experimental data from Contract NAS 3-20104 withthat given in Reference 4 is shown in Table XIII. It is seen that there isgeneral agreement with the engine performance based on the efficiency (nls)of the engine for zero gap cases, but that either the base or module per-formance derived from the data do not agree. For example, the ContractNAS 3-20104 results in the table indicate a negative recovery of the tiltangle (cosine 0) loss, whereas the Contract NAS 8-11023 results indicate somerecovery even for a zero length plug. It is not expected that the largedifference in area ratio between the cited cases should have any effect forzero gap cases.

For a fixed number of n_dules and module area ratio, the test datain Figure 52 show that increasing the cluster area results in a decreasein performance and efficiency. This decrease in performance and efficiencyis due to the mismatch of aerodynamic flow fields for discrete bell nozzlesexhausting onto an annular plug contour, as indicated in Figure 53. The lossin performance is conventionally reported as due to the increase in gapassociated with the increase in cluster area ratio.

The trend in the data is what would be expected for a plug clusterconfiguration based on an annular plug nozzle, where the contour has notbeen optimized for discrete internal expansion sections (bell nozzles) witha gap between the nozzles. Typical static pressure data are shown on a

88

I!! :!

TABLE XII - AERODYNAMICVARIABLES FOR TEST MODELS

Config.

A-I]

A-IO

D-IO

E-IO

B-IO

C-IO

A-7

A-l]

A-IO

D-IO

E-IO

B-IO

C-IO

A-7

C C

5OO

40O

400

400

400

400

200

5OO

4OO

40O

400

400

400

2OO

_M N 6/DE gT EXPTL OT Ec/CM EXPTL__ Eq.(3)

(FREON)

Based on y = 1.15

40 12 1.96 27.9 31.9 12.48

40 12 1.62 26.93 29.4 lO.Ol

40 20 1.06 25.93 29.4 9.85

40 5 2.88 26.93 29.4 10.60

80 12 .77 17.2 19.5 5.04

200 12 .Ol 6.63 8.0 2.00

40 12 .77 19.23 21.5 5.02

Based on y = 1.4 (AIR)

40 12 1,96 27,9 19.9 12.48

40 12 1.62 25.93 18.6 lO.Ol

40 20 1.06 25.93 18.6 9.85

40 5 2.88 25.93 18.6 I0.60

80 12 .77 17.2 11.9 5.04

200 12 .Ol 6.63 4,6 2.00

40 12 .77 19.23 14.0 5.02

12,36

9.92

9.80

I0.34

5.0l

l.99

4.99

12.67

I0.15

9.94

II.05

5.08

2,00

5.06

89

BT= _E - _M Eq. (3)

:_ y+l_-l- tan'l_ y-Iy+l (M2 -l) -tan'l;M 2 -lEq (3a,

13o[Y=l.l

120F

llO I

_ lO0 I

9O

80 I I I I I

0 lO0 200 300 400 500

ENGINE AREA RATIO c = AE/AT

Figure 51. Effect of Gas Properties on Required Tilt Angle.

90

Ii!i!i

TABLE XIII- COrIPARISON OF EXPERIMENTALPLUG CLUSTER PERFORMANCE

PARAMETER

NAS 8-11023*

N 24 24

SM 4.85 4.85

_E 15.0 15.00 18 18

PLUG LENGTH 0 9.4

10DE (MODULE) 63.72 63.72

!ODE (ENGINE) 69.66 69.66W (ENGINE) 3.48 3.48

IDEL (ENGINE) 66.45 67.29

qlS (ENGINE) 0.954 0.966

IDEk (BASE) 3.99 3.43

PB/PE 2.93 2.5

% BASE I s 6.0 5.I

WBASE 0 0

IDEL (MODULE)***: 62.47 63.86

hiS (MODULE) 0.980 1.002

IODE (MODULE) . cos 0 60.60 60.60

% cos 0 loss recovered**** 59.9 I04.5

12

7.7

15.0

I0

0

66 62

69 66

3 48

67 15

0 964

1 25

1 48

1 9

0

65.90

O. 989

65.61

28.7

NAS 3-20104**

12

195.6

386.5

6.63

0

94.78

95.65

0.574

91.46

0.956

0.47

1.60

1.7

0

90.99

O. 960

94.14

<0

* Reference 4: Base area estimated from photographs of hardware; flow rateassumed from sample case given in Appendix A.

** Test 35.01; Configuration C-lO; Air Media

*** IDEL (ENGINE)- IDEL (BASE) = IDEL (MODULE)

**** [IDE L (MODULE) 10DE (MODULE) x cos e]/[loD E (MODULE) - IODE MODULE) x cos e]

91

__)

>-f_)Z

c_)k-4IJ-b--LAJ

F---C/)

-rF--

LJJ

ZL_J

O.gl8 -

0.914 -

0.910 _

O.906 -

O.902 -

0.898 _

0.894

CONTRACT NAS 3-20104 (REFERENCE 46)

TEST CE Y/De BTNO. CONFIG. Isp

49.02 A-7 200 0.77 19° 63.39

28.01 A-lOc 400 1.62 26° 63.07

46.01 A-ll 500 1.96 28° 62.29

_M = 40N=I2

I I I200 300 400 500

ENGINE AREA RATIO tE = AE/AT

CT

0.919

0.897

0.895

Figure 52. Cluster Performance as a Function of Engine Area Ratio.

92

lili!!

N

I I I I I I I I

HIgN31%0£

H19N37%0

I I I I I I

(1J

QJ,r-

U

0

{.) r..-_J N4.- N4-- OLI..IZ

,,.,.(U4-I (}Je:l c"L,I._

"_ c-e- Oi-.-i

COft_

..IJ(._

r',4-O

_ e..-

,,4

l,

93

schematic of the plug cluster in Figure 53. The two values listed refer topressures measured on the plug nozzle wall between two modules (gap of 1.6module exit diameters) and on the center line of one module exit. We would

expect the static pressure values to be lower as the gas expands on the

plug nozzle, and to approximate the isentropic relationship. Correspondingly,we would expect to be able to calculate the effective area ratio using theisentropic relations (with suitable two dimensional corrections).

The calculated aerodynamic area ratios based on the pressure measure-ments are shown on Figure 53 for O, 15, and 30 percent plugs, and also onFigure 54. It is seen that the gas expands from the module area ratio of

40 to an equivalent area ratio of 84 at the plug exit. Now, by definition,this plug cluster configuration has an area ratio of 400. However, sincewe are examining the gas flow data on the surface of the plug, we must comparethe data with the geometric plug flow area ratio. This area ratio is deter-

mined by taking the defined engine area, subtracting the area occupied bythe plug, and dividing this value by the sum of the module throat areas.Three such area ratios are indicated on Figure 53.

It is seen, therefore, that less than one-third (84/303) of the avail-

able geometric plug flow area ratio was actually achieved with the 30% plugtest configuration of Figure 53. This result is consistent with the one-third (40/I19) value expected for a zero length plug, the larger plug flowarea being the result of the gap between the modules. We thus have a grossdiscontinuity in area ratio at the match point.

The results from Contract NAS 3-20104 may be interpreted by examiningthe geometric flow area for a zero gap version of Figure 53. Despite the

fact that an even number of modules cannot be added to change this con-figuration to zero gap, the assumption can be verified by examining a cluster

with a 6/De = l gap, for example. The addition of more modules increasesthe sum of the throat areas, and thus decreases the geometric plug flowarea ratio to that shown in Figure 54 at zero gap. It is seen that the

aerodynamic area ratio based on the pressure measurements more closely fitsthe zero gap geometric plug flow area ratio. Thus it can be concluded that

the stream tubes emanating from the nozzles (modules) apparently followessentially the same aerodynamic path regardless of the area available. Theperformance at large gaps (for the tested configuration) is thus about thesame as at low gaps (area ratios).

The test results from Contract NAS 3-20104 indicate a serious flaw

in the design criteria of high area ratio plug clusters based on methodologies

developed from low area ratio testing. The low area ratio, low gap,

methodology stipulates a one-dimensional matching of the module Mach numberwith that of an annular plug Mach number, as shown in Figure 55. But this

approach becomes unsatisfactory at gaps much greater than zero, because wehave a three dimensional problem. Geometric remedies, such as the addition

of fairings between the modules, offer a partial solution for large gaps,as shown in Contract NAS 3-20104 and also indicated in Figure 12.

As shown in Table XlV the effective specific impulse of the base

injection flow (Isp Base) is from 6 to 13% less than the specific impulse

94

il;_I

I I0L_

"F-

Z ::3w

4-

Z,,, 0

W

n:lQJL

Q;

0

°l,,,-IJ-

95

!0--4

0

L0

g-r--

4.J

.i,m

e-o

e-u

E

Z

-r-

L_

°_...

96

TABLE XIV. BASE PRESSUREAND SECONDARYEFFECTS (NAS 3-20104)

% PBTest Config. FVAC _ ISPE FBase Total ISPBase % Plug

29.01 A-IOb 59.72 0 64.93 .2317 0.4 .014 - 3029.02 A-IOb 60.30 .0107 64.86 .8303 1.4 .051 61.08 30

28.01 A-lOc 58.94 0 63.84 .8445 1.4 .035 - 1528.03 A-lOc 59.59 .0109 63.63 1.4685 2.5 .061 61.78 15

30.01 A-lOd 56.33 0 61.17 2.6872 4.8 .062 - 030.02 A-lOd 56.93 .0107 61.05 3.2050 5.6 .074 52.91 0

Fvac I AFBase FB = PBABISPE = Wp (I + Ws/Wp) SPBASE - Us

generated by the primary flow. The overall specific impulse with secondaryinjection for Contract NAS 3-20104 testing was always less than the Isp with-out base injection. This result differs from that found in Reference 4,where the efficiency of the zero gap plug cluster was slightly better orequal to that for no base flow up to a secondary flow of about two percent.It also differs from that found for the annular throat (Aerospike) configura-tion, where the secondary flow specific impulse amounted to 6056 seconds(Table IV) for flows as low as 0.2 percent. The base pressurization resultsfrom Contract NAS 3-20104 correspond to those obtained for below design pointtesting of plug nozzles, wherethewake is not closed. The results indicatethe possibility that the flow did not provide a closed wake, makingsecondary injection not as effective.

The results from the testing on Contract NAS 3-20104 represent selectedpoint design plug cluster configurations, as resources did not allow asystematic investigation of the many variables. Nevertheless, the data andtheir comparison with related test data from the literature, indicate thatlow performance will be obtained with large gap cluster configurations ofbell nozzles on annular plug designs. The results conclusively define theproblem as being one of assuring an aerodynamic flow match between the belland the plug. Solution of the problem leads to unconventional plug con-tours that resemble a fluted plug, and provide much higher performance thanconventional type clusters.

G. MODULE-PLUGCLUSTER PERFOrmANCEMODEL II

The initial computer model (Model I) represented an engine configurationin which the cluster throat location coincided with the equivalent area ratioannular plug throat location. In order to provide a better approximation ofthe test data from Contract NAS 3-20104, Model I was revised to incorporatea Mach number match point. The configuration model is that shown in Figure 55.Uncorrected plug performance is calculated in a similar manner to that forModel I, except that CFLO is now evaluated at the point of Mach number match.

97

Theuncorrected plug performanceis multiplied by a gap efficiencyfactor (see Figure 44), which is defined as the ratio of the uncorrectedcluster efficiency to the efficiency of a zero gap configuration at thegiven cluster radius (REc).

Correction for base pressurization wasprovided by a correlationderived from Contract NAS3-20104data utilizing the effective base pres-sure PBas a function of the plug exit pressure PE.

Application of Model II to a cluster configuration used in ContractNAS3-20104resulted in predicted performanceof CF= 1.927 comparedto themeasuredCF= 1.928 - 1.933. Plug cluster engine performancecomputedforthe baseline case is shownin Table XVfor both Models I and Iio Cold flowtest information are included in the table for comparison.

It is seen from Table XVthat Model II, basedon Contract NAS3-20104test data correlations, predicts that a plug cluster engine will be only 91%efficient. Model II is not considered to be an accurate representation of aplug cluster engine, becauseof the approximations that have beenutilized.Themodel, however, doesprovide engine performanceconsistent with thecold flow results for Contract NAS3-20104configurations. The modelwillrequire r6vision to predict the performanceof optimumcluster designs.

H. MODULE-PLUGCLUSTERPERFORMANCEMODELIII

The problemof achieving highly efficient aerodynamicflow for theplug cluster engine conceptwas solved by joining high area ratio, partiallyscarfed, bell nozzles in the mannershownin Figure 56. Discussions con-cerning the performanceof the scarfed-bell or fluted-plug cluster engineconcept are presented in this section.

Since the JANNAFsimplified performancemethodologyis well establishedfor full bell nozzles, Model III includes the analysis of the configurationshownin Figure 57. Theperformanceof the scarfed-bell plug cluster engineis expected to very closely approachthat for the clustered bell conceptshownin Figure 57.

In order to makea direct comparison(in Section VIII of this report)betweenthe plug cluster engine and candidate SpaceTugengines, such as theRLIOand the AdvancedSpaceEngine (ASE), this section also includes thecalculated performancefor these engines.

I. Model III Description

The JANNAF simplified methodology, as utilized in this study,reduces to the equation

aFBLISPdelivere d = IsPoDE (nKl N nNOZ hERE _--)

98

i_ :11

TABLE XV - PLUG CLUSTER PERFORMANCE MODEL COMPARISONS

Model I Model II

Vacuum Thrust, KN (Klb)

Chamber Pressure, atm (psia)

Vacuum Specific Impulse, s

Mixture Ratio (O/F)

Engine Area Ratio (AE/At)

Module Area Ratio (Ae/At)

Number of Modules

Module Gap (6/De)

Engine Diameter, cm (in)

Plug Base Diameter, cm (in)

Engine Length, cm (in)

Percent (L/LI) Plug

Tilt Angle, deg.

Base Flow Ratio %

Uncorrected Is, s

Base Correction als, s

Gap/Fairing Correction aIs, s

Delivered Is, s

Engine Efficiency nIS

72.5 (16.3) 68.7 (15.5)

20.4 (300) 20.4 (300)

467.4 443.8

5.44 5.5

458 458

40 40

lO lO

2 2.05

320 (126) 320 (126)

218 (86) 173 (68)

86 (34) 91 (36)

15 20

3.5 27.7

0.2 0

449.5 454.6

+ 8.9 + 0.6

+ 9.0 - If.4

467.4 443.8

O.962 O.914

NAS 3-20104

Test 46.01

Config. A-ll

lw

I0.7 (157)

mB

493

40

12

1.96

nD

15

27.9

0

0.895

99

D.

c-O

¢-.-r-

c--LU

IZn

r=-

L

°e,,-

100

_ I

O.

¢.)_.-0

N0

r....

Q,I

l-

or.-LI-

lOl

where IsPODE is the one dimensional equilibrium specific impulse, nKIN isthe kinetic efficiency (IsPODK/IsPODE), nNOZ is the nozzle or divergenceefficiency, hERE is the energy release efficinecy, _FBL is the boundarylayer thrust decrement, and F is the nominal engine thrust.

The values of IsPODE and IsPODK are generated by the JANNAF ODE-ODK-TDK computer program, and nNOZ and AFBL are found using the charts andmethodology from Reference 17. The value of hERE = 0.995 was used in theperformance calculations for all of the engines.

Because the plug cluster engine (PCE) is not strictly a bellnozzle configuration, its performance required additional calculations asfollows:

° Determine the IsPdelivered for scarfed nozzles.

Calculate exit pressure (PE) corresponding to overall PCEarea ratio.

Determine base pressure (PB) as a function of PE (normallyassumed to be 2.5 x PE but can be as high as 3.6 x PE)-

o Determine base area (AB) of PCE.

Calculate thrust loss (_FoT) due to module tilt angle (OT);equals Fm (I - coseT).

Calculate delivered thrust of PCE; FpCE = N Fm + PB AB -N aFoT, where N is the number of modules in the cluster.

Calculate the delivered specific impulse; IS_del = FPCE/WENGINE, where the flow rate to the engine, WENGINE, mayinclude a base bleed contribution.

The rationale and assumptions used in the calculations for thePCE are described in the following.

2. MODEL III Nozzle Efficiency

Nozzle divergence efficiency is obtained in the standard mannerfrom the following equation

l + cosnNOZ : 2

where e is the nozzle divergence angle.

In the case of the plug cluster engine, there is some questionregarding the use of the module nozzle efficiency, since the module istilted toward the axis of the engine. For the case when the tilt angleequals the nozzle divergence angle, the flow from the outer portion of thenozzle is aligned with the axis of the engine. The flow from the inner

102

portion of the nozzle is turned aerodynamically in forming the wakeonthe base of the plug. Since the methodof calculation includes a tiltangle (cos c)) loss, inclusion of the modulenozzle divergence loss doesnot appear to be warranted. Since there is someuncertainty, however, thePCEdelivered performancewill be presentedwith andwithout this loss.

3. Scarfed Nozzle Performance

The scarfed bell/plug cluster nozzle concept (Figure 56) can beenvisioned to be a fluted plug nozzle with both internal and externalexpansion components. As such, the performance would very closely approach

that for the full bell cluster shown in Figure 57. Were the scarfed bell/plug cluster nozzle to operate as a cluster of scarfed nozzles with a smallamount of base thrust contribution, the performance would be less. In order

to present this degree of uncertainty, PCE calculations were made assumingthe module thrust contribution to be only that from a scarfed nozzle.

The first step in determining the performance of a scarfed nozzleis to determine the area ratio (_eff) of an equivalent unscarfed nozzle.

In this manner it is assumed that the delivered performance of a scarfednozzle corresponds to the area ratio at the intersection of the scarfingplane and the lengthwise nozzle axis. This method of scarfing is shownin Figure 58, along with the method chosen for the PCE design.

The analytical expression for ceff is obtained by assuming a15-degree conical nozzle and by specifying the two area ratios (El and E2)between which the nozzle is scarfed.

4_IE 2eff =

(¢_I + ¢_2 )2

For the two scarfed nozzles considered, the expression yields:

When Nozzle is Scarfed From Eeff

El = 40 to E2 = 500 97

El = I00 to _2 = 500 191

Using these values of _eff, the kinetics and boundary layer losses wereobtained from Figures 59 and 60, respectively.

As seen in Figure 58, the PCE scarfed nozzle is not as severelyscarfed as the one utilized in this analysis. However, the more conserva-

tive _eff was utilized for this analysis.

4. Model Ill Base Pressurization

A relatively consistent correlation between PB and PE was pre-viously cited in Table X and Figure 35 (Section IV,C), where the base pres-sure obtainable is defined as that in which the nozzle separation criteria

103

i,i

NNOZ

i,ii,

U

WUm,

A

v

i,i

NN

Z

L_

U

:b-

Z.=C

v

5.-

NQ

Z

"O

LO

.r'-i,

104

iT| ! _-

Z

-j

1.000

.990

.980

.970

LOX/LH2

0 Pc = 300 Rt = 0.9

[] Pc = 500 Rt = 0.7

JANNAF ODK METHODOLOGY

MR = 5.0

MR = 4.0MR = 7.0

i I l i40 I00 500 I000

Figure 59. Kinetics Loss for Scarfed Nozzle.

105

II II

4"1 _ 1.1.1/'_ r,," (__

zI,I

C_ C_ ILl

II IIkid

¢sd _ (0 _ I.L

"" " on z

° _ <::]0._

I I J I 1 I ! I I

(387) _3_,V7 _,_JVGNflO8--IV

00

0

o

v

I.i.I

m,,.

0

W

C)i,-4l"--

":Cl,.l,J

,,:C

,5NN0

I-

(.}V'}

f,-0

W-

f#710,.=I

G.I

.--I

C

0_3

_3

Cn

106

holds. Base pressures above the nozzle separation criteria (PB/PE > 3.6)

cause an enlargement of the wake on the plug base, and thereby reduce theeffective area ratio obtainable with a plug nozzle.

The assumption was made, as cited in Section IV,C to utilize

bleed flows of 0.2% of the engine flow.

5. Model Ill Plug Cluster Enqine Delivered Performance

The plug cluster engine performance calculated by the outlined

methodology is summarized in Table XVI. Calculated values are given forchamber pressures of 20.4 atm (300 psia) and 34.0 atm (500 psia), and formixture ratios of 5 and 6. Three types of nozzles are assumed: (1) per-formance equivalent to a full bell nozzle, (2) performance equivalent toa scarfed bell at _ = 40, and (3) performance equivalent to a scarfed bellat E = lO0.

Table XVII gives a comparison of the three performance models.

In order to estimate the possible uncertainty in the calculatedvalues of performance, a base case was selected, and the assumptions weremodified to determine their effect on performance. The result of theuncertainty analysis is summarized in Table XVIII.

The lower limit in performance is achieved by a cluster of bell

nozzles with zero tilt angle. For this case, it is assumed that the base pres-sure is equal to PE, and that there is zero base bleed. The nozzle efficiencyis now 0.994 as the divergence loss of the bell nozzle must be taken into

account, The resultant performance is found to be 460.7 seconds (nls = 0.943).A possible upper limit for the cluster is found to be 471.7 seconds (nls =0.965). Zero base bleed and the maximum base pressure consistent with nozzle

separation criteria is assumed. Also a smaller boundary layer loss is assumed,which is consistent with more rigorous calculations. A loss of 0.6 second in

specific impulse is required for a gas generator cycle plug cluster engine.

In order to further evaluate the validity of the performanceprediction for the plug cluster engine, calculations were made for theRLIO and ASE using the same JANNAF simplified procedures. The results of

these calculations are given in Table XIX. Comparisons with the perform-ance values presently accepted for these engines are shown.

I07

o° .

§ _-

o°

LK')

O_

co § _o°ii

W

r-. _.

V1ILl

;E o_

"' 0 § _o_

(/)

"J _ _-_" •

.J _

E:) U ,_

IJ-

0 ILl _ "_

i,

I..iJ

..a _ •

(x-) '5"

LL

I

8 " ,.ox _ _g "7,_ °_ 'o _ _

0

108

_. _ . _. ?_

°!. _ o "= _ ,-: _°' _ _ _ . . _._ _ ; ,.,.; o _ ,3 _ °

. _ _ ,.,., o oi. _

o"=. _. m m _, ,.o

_. o° o .=. oo _ _ o • _ _ '_ _ _.

.

o c3

Ii! I

TABLE XVII - PLUG CLUSTER PERFORMANCE MODEL COMPARISONS

Model l Model II Model III

Vacuum Thrust, KN (Klb)

Chamber Pressure, atm (psia)

Vacuum Specific Impulse, s

Mixture Ratio (O/F)

Engine Area Ratio (AE/At)

Module Area Ratio (Ae/At)

Number of Modules

Module Gap (6/De)

Engine Diameter, cm (in)

Plug Base Diameter, cm (in)

Engine Length, cm (in)

Percent (L/LI) Plug

Tilt Angle, deg.

Base Flow Ratio %

Uncorrected Is, s

Base Correction aI , ss

Gap/Fairing Correction als, s

Delivered Is, s

Engine Efficiency his

+

+

72.5 (16.3) 68.7 (15.5)

20.4 (300) 20.4 (300)

467.4 443.8

5.44 5.5

458 458

40 40

lO lO

2 2.05

320 (126) 320 (126)

218 (86) 173 (68)

86 (34) 91 (36)

15 20

3.5 27.7

0.2 0

449.5 454.6

8.9 + 0.6

9.0 - II.4

467.4 443.8

0.962 0.914

67.1 (15.1)

20.4 (300)

463.9

5.5

895

500

lO

0

433 (170)

246 (97)

82 (32)

0

5.3

0.2

459.3

+ 4.6

0

463.9

0.947

lOg

TABLEXVIII. - UNCERTAINTYIN PLUGCLUSTERENGINEPERFORMANCE

Expander Cycle Base Case:

MR = 6.0 EM = 500 N = I0

Pc = 300 _E = 895

Delivered IspVariable Performance

nNOZ : 1.0 t_B : .065 (0.2%)

PB = 2.5 PE eT = 5"3°

_FBL = 47.9

Difference From

Base Case Isp

Base Case

nNOZ = 0.994

WB : 0

WB = 0.33 (1%)

PB = 3.6 PE

PB = l.5 PE

eT=O

WB = 0

PB = PE

nNOZ = 0.994

_FBL = 38

Gas Generator Cycle

I

464.4

461.5 -2.9

465.3 +0.9

460.6 -3.8

467.l +2.7

463.9 -0.5

460.7 -3.7

467.5 +3.1

463.8 -0.6

The JANNAF simplified procedure is seen to give conservative per-

formance prediction for the ASE, and to give correct performance predictionfor the RLIO, providing a more conservative nozzle efficiency and otherlosses are utilized in the calculation. It is anticipated, therefore, that

the preceding methodology for the plug cluster engine will provide a reasonablyaccurate assessment of the performance potential.

110

T_.|!

TABLE XIX. JANNAF SIMPLIFIED PERFORMANCE FOR RLIO AND ASE

Engine RLI0 RLI0Description IIB IIB ASE ASE

Cycle Exp. Exp. SC SC

Pc 400 400 2000 2000

MR 5 6 5 6

200 200 400 400

% Bell 75 75 90 90

F nom (Ibf) 15K 15K 20K 20K

Ispode 477.3 477.2 484.0 485.8

_ere 0.995 0.995 0.995 0.995

nkin 0.998 0.998 0.996 0.996

-nnoz 0.996 0.992 0.994 0.994

aFBL 300.92 299.0 380.35 380.35

noverall 0.969 0.965 0.966 0.966

Isp delivered 462.5 460.6 467.6 469.3

(Ref. 26) (Ref. 27)456.2* 473.0

*Includes dump cooled nozzle loss and nNOZ = 0.982

Ill

SECTION V

SUBSYSTEMEVALUATION


Analyses were conducted for the four major subsystems of the plugcluster engine to determine the configurations, operating conditions, andweights that must be considered for the complete engine system analysis.The subsystems analyzed are:

Base PressurizationEngine Cooling (Thruster Module and Base Region)Turbomachinery and PowerThrust Vector Control

The extent of the subsystem analysis was carried out only to determinethe effect on engine performance limitations imposed on engine design, andto define the geometry for the subsequent weight estimates.

The design point for the plug cluster engine evaluation was assumedto be that given in Table I, commensurate with the baseline Space Tug re-quirements. Upon completion of the analysis, the selected configurationsand associated rationale were reviewed with the NASA LeRC Project Managerto select the specific configurations to be carried into the conceptualdesign phase.

B. ENGINE CYCLE ANALYSIS

Candidate cycles that were evaluated include expander topping andgas generator cycles (Figures 61 to 70). Parallel turbine arrangementsand single turbine arrangements with a direct-drive fuel pump and a gear-driven oxidizer pump were compared. The cycle analysis was conductedutilizing a preliminary version of the 66.7 kN (15,000 pound) thrust plugcluster engine at the baseline design point (Table XX). Conclusions derivedfor the design point are essentially applicable for the thrust levels (between44.5 kN and 111.2 kN [I0,000 and 25,000 pounds force]) under consideration andfor chamber pressures to 34 atm (500 psia).

TABLE XX. PLUG CLUSTER ENGINE BASELINE DESIGN POINT

Vacuum ThrustChamber PressureMixture RatioEngine Area RatioNumber of Modules

66,723 N (15,000 Ibf)20.4 atm (300 psia)

6_400

10

The expander topping cycle is basically a closed cycle because theturbine flow can be included in the main chamber flow. A small portion(about 0.2%) is directed through the nozzle base to maximize the base pressurethrust contribution. The gas generator cycle is an open cycle. The turbine

113

; jr'w.' ._

\ \

i,

Figure 61. Cycle EXOI: Expander Topping Cycle, H2-Cooled TCA, 02-CooledPlug, Single Turbine TPA, Base Pressurlzation with H2

114

I!I I

Q;

0

io,J°

Od

_=I'-- °w--

Q) t'-'-- 00 *r-0"_

! N

-r" S.-

.N _,)

Z_&,,

I'---S,-

_ 0r,-

',' I---

• i r,-,,,,

0 e-

&a-

"r--h

115

Figure 63. Cycle EX03: Expander Topping Cycle, H2-Cooled TCA, TPA withSeparate Gas Driven Turbine, Base Pressurization with H2

116

Figure 64. Cycle EX04: Expander Topping Cycle, H_-Cooled TCA, O_-Cooled

Plug, Parallel Turbine TPA, Base Pressbrization with I_2

117

I

Figure 65. Cycle EX05: Expander Topping Cycle, H_-Cooled TCA, H2-cooled Plug,Parallel Turbine TPA, Base Pressurizatlon with H2.

lib

1!!!-

Figure 66 Cycle EXlIA: Expander Topping Cycle, H2-Cooled TCA, 02-CooledPlug, Dual Single Turbine TPAs, Base Pressurization withH2 (Not Shown)

119

Y

V

I GB

Figure 67. Cycle GGOI: Gas GeneratorGG Exhaust on Plug, Single

Partial GG Exhaust.

Cycle, Hp-Cooled TCA, 02-Cooled Plug,Turbine TPA, Base Pressurization With

120

I!!I_

!

I

-tT

I

Figure 68. Cycle GG02: Gas Generator Cycle, H2-Cooled TCA, H2-Cooled Plug,GG Exhaust on Plug, RLIO TPA, Base Pressurization with PartialExhaust.

GG

121

GG

Figure 69. Cycle GG03: GasGenerator Cycle, H2-CooledTCA,02-CooledPlug,GGExhauston Plug, Parallel Turbine TPA,BasePressurization WithPartial GGExhaust.

122

GG

7-

Figure 70. Cycle GG04: Gas Generator Cycle, H2 Cooled TCA, H2-Cooled Plug,GG Exhaust on Plug, Parallel Turbine TPA, Base Pressurization WithPartial GG Exhaust

123

exhaust flow is greater than the optimumbase flow required for base pressuriza-tion with the result that the majority of the flow is dumpedfrom the engine,producing a low thrust contribution. Since the chamberpressure is relativelylow for the plug cluster engine, the turbine flow rate, which is determinedby the turbomachineryrequirements, is relatively small, such that the thrustloss is only on the order of 0.6%. A portion of this loss can be regained bydumpingthe gases on the plug in the gaps betweenthe moduleexits.

In the expandercycle, turbine power is derived from passing mostof the hot hydrogen(and hot oxygenin somevariations) from a coolingjacket through low pressure ratio turbines. A control reserve can be ob-tained when5 to 20%of the available turbine flow bypassesthe turbine. Thecombinedhydrogenflow, minus a small amount(0.2%) providing base pressuriza-tion, is injected into the main combustionchamber.

Several types of expander cycles were examined. Cycle EXOI, depictedin Figure 61, represents a plug cluster engine utilizing an RLIOtype turbopumpassembly(TPA). Note that hydrogenis used to cool the modules, and oxygenthe plug and base. A discussion concerning modification of this pressureschedule to best utilize an existing RLIOTPAis presented in the next section(Section V.C.). The pumpdischarge pressures are 34.0 and 39.5 atm (500 and580 psia) for the oxygenand hydrogenpumps,respectively.

Cycle EX02,shownin Figure 62, is identical to that in Figure 61except that hydrogenreplaces oxygenas the plug coolant. The pumpdischargepressures are 26.8 and 43.2 atm (394 and 635 psia) for the oxygenand hydrogenpumps. This cycle wasone of those selected for conceptual design study. Thepressure schedule for a baseline engine, utilized in the preliminary cycleevaluations, is given in Table XXI.

Cycle EX03(Figure 63) with oxygenand hydrogenpumpdischarge pressuresof 40.5 and 35.7 atm (595 and 525 psia) utilizes both hot hydrogenand hotoxygen-driven turbines. The feasibility of obtaining sufficient heat input tothe oxygenat the baseline pressure conditions, and at a short plug length,is marginal, as discussed in Section V.D. on engine cooling. This cycle, andthose depicted in Figures 64 and 65 offer the potential of lighter weightturbomachinery. Cycle EX05(Figure 65) wasselected for further design analysis.

An expandercycle configuration, utilizing two TPAs, is shownin Figure66. This cycle provides an approachto a fail-operational modeas opposedtoa fail-safe failure modedesignated for the single engine baseline SpaceTug.EachTPAof Cycle EXIIA delivers propellant to one-half of the modules. The

ropulsion system, therefore, has the capability of operating at full thrustall modulesfiring), at 50%thrust (one'half of the modulesfiring), or atin-between thrust levels, dependinguponthe throttle capability of the finaldesign. Failure of component(s)in oneTPA-fed subsystemwould allow aminimumof 50%thrust capability for the SpaceTug to return to a servicestation for repair. Since a weight penalty would be associated with CycleEXIIA, it is included here only to showa further potential that a modulecluster can offer.

124

Ill !

TABLE XXI - CYCLEEX02PRELIMINARYPRESSURESCHEDULE

Pressure_ atm (psia)

Main Pump Discharge

AP Line (2%)

AP Plug Coolant Jacket

Plug Coolant Jacket Outlet

AP Line (1%)

_P Fuel Shutoff Valve (1%)

Valve Outlet

Orifice Outlet

Coolant Jacket Inlet

AP Coolant Jacket

Coolant Jacket Outlet

AP Line (1%)

Turbine Inlet

AP Turbine (Total to Static)

Turbine Outlet

AP Line (1%)

Shutoff Valve Outlet (1%)

Orifice Outlet

Main Injector Inlet

AP Injector (I0%)

Chamber Pressure

02

26.8 (394)

O.B(8)

--u

--m

0.3(4)

26.0 (382)

23.8 (350)

23.1 (340)

2.7 (40)

20.4 (300)

H2

43.2 (635)

0.9 (13)

3.4 (50)

38.9 (572)

0.3 (5)0.3 (5)

38.2 (561)

37.7 (554)

36.5 (537)

6.7 (98)

29.6 (435)

0.3 (4)

29.3 (431)

4.9 (72)

24.4 (359)

0.3 (4)

24.0 (352)

23.7 (348)

22.9 (337)

2.5 (37)

125

In the gas generator cycles shownin Figures 67-70, a small amountof propellant (about I% of the engine flowrate) is burned_n a gas generatorto powerhigh pressure ratio turbines. Themixture ratio is selected togive a gas temperature of about 922°K (1660°R) (MR= 0.7 to 0.9, dependinguponthe heat input to the propellants in the cooling circuit). In the casesevaluated, the required turbine flowrate exceedsthe optimumbase pressuriza-tion flowrate (0.2,_) by a factor of five.

Cycle GGOI,depicted in Figure 67, represents a plug cluster engineutilizing an RLIOtype TPA. The oxygenand hydrogenpumpdischarge pressuresare 33.9 and 27.7 atm (498 and 407 psia), respectively.

In Cycle GG02,shownin Figure 68 hydrogenis usedas the coolantfor both modulesand plug, with oxygenand hydrogenpumpdischarge pressuresof 27.1 and 31.8 atm (398 and 467 psia). This cycle wasselected for concep-tual design analysis. A Typical pressure schedule for this cycle is given inTable XXII (next page).

In Cycles GG03and GG04,Figures 69 and 70, a lighter weight parallelturbine arrangementis utilized. Cycle GG04wasone of the all hydrogen-cooled cycles selected for further study.

I. Cycle Analysis Summary

The results of the cycle analysis are presented in Table XXIII and

XXIV. There appear to be no limitations in the power balance for the regen-eratively cooled plug and modules, even at chamber pressures to 34 atm. However,no expander cycle power balance was possible for an ITA module with an expansion

ratio of _M = lO0. This was due to the lack of sufficient LH coolant tempera-ture rise Tn the shortened plug and the chamber portion of th_ module.

TABLE XXIII. CYCLE ANALYSIS SUMMARY

Cjcle EX02Regen-Cooled ITA (16% FFC)

GG04

Regen-Cooled ITA (16% FFC)

Pc atm 20.4 20.4 20.4 20.4

EM 40 40 40 40Is sec 467.4 454.9 466.8 454.4Als sec 0 0 0.6 (0.13%) 0.5 (0.10%)WGG 0 0 0.38 0.36

Pc atm 34.0

EM 40 NotIs sec 471.I CalculatedAIs sec 0

WGG 0

34.0 34.0

40 40470.I 457.7

1.0 (0.21%) 0.8 (0.17%)0.57 0.57

Pc atm 20.4 20.4

_M I00 I00Is sec 469.2 NoAls sec 0 Power

WGG 0 Balance

20.4lO0

Not468.5

0.7 (0.15%) Calculated

0.40

126

IT!IT

TABLEXXII - CYCLEGG02PRELIMINARYPRESSURESCHEDULE

Pressure, atm (psia) TCA

02 H2

Main Pump Discharge 27.1 (398) 31.8 (467)

_P Line (2%) 0.5 (8) 0.6 (9)

AP GG Valve (I0%) ....

Valve Outlet ....

GG Inlet ....

AP GG (I0%) ....

Turbine Inlet ....

aP Turbine (Total to Static) ....

Turbine Outlet ....

AP Line (20%) ....

Plug Dump ....

_P Shutoff Valve (2%) 0.5 (8) --

Valve Outlet 26.0 (382) --

Orifice Outlet 23.8 (350) --

Plug Coolant Jacket Inlet -- 31.2 (458)

aP Plug Jacket -- 3.4 (50)

Plug Coolant Jacket Outlet -- 27.8 (408)

AP Line (1%) -- 0.3 (4)

Fuel Shutoff Valve Inlet -- 27.5 (404)

AP Shutoff Valve (1%) -- 0.3 (4)

Valve Outlet -- 27.2 (400)

Coolant Jacket Inlet -- 26.7 (392)

AP Module Coolant Jacket -- 2.7 (40)

Coolant Jacket Outlet -- 24.0 (352)

Orifice Outlet -- 23.7 (348)

Main Injector Inlet 23.1 (340) 22.9 (337)

AP Injector (I0%) 2.7 (40) 2.5 (37)

Chamber Pressure 20.4 (300)

GG

02

27.1 (398)

0.5 (8)

2.7 (39)

23.9(3Sl)

23.2(341)

2.3 (34)

H2

31.8 (467)

0.6(9)

3.1 (46)

28.0 (412)

25.7 (377)

4.8 (70)

20.9 (307)

17.9 (263)

3.0 (44)

0.6 (9)

2.4 (35)

127

I

CO

0

0

gx

128

iii | _-

C. TURBOMACHINERYANALYSIS

The turbomachinery selection and design studies were conducted bygiving consideration to the following system capabilities:

Idle mode engine firing to provide thrust for propellantsettling in the tanks and thermal conditioning (chill)of the propellant feed system.

° Two-phase flow pumping capability in both oxidizer and fuel pumps.

These capabilities are essentially the same as those provided by theRLIO derivative IIA configuration. The existing RLIO TPA, however, has a 5-hour life limit.

I, RLIO IIA Turbopump Assembly Analysis

The fuel and oxidizer pump performance curves used forRLIO derivative IIA turbopump performance predictions are shown in Figures71 and 72. The turbine flow parameter is shown in Figure 73.

These curves were constructed utilizing the information in Reference 14Appendices. Application of this pump to the various plug cluster enginecycles requires certain modifications, which are outlined in the following.

° Cycle EXOI

A modified RLIO derivative IIA turbopump was selectedfor this engine. The modification consists of a 10% reduction in fuel pumphead coefficient accomplished by impeller trimming. No changes are anticipatedin the turbine, the low speed or the high speed oxidizer pump. Cycle powerbalance was achieved with a turbine speed of 27,850 rpm and a turbine bypassflow of 39% (Fig. 74).

° Cycle EX02

As with Cycle EXOI, a modified RLIO derivative IIAturbopump was selected. The turbopump modification consists Of a 19% reductionin oxidizer pump head accomplished by trimming the oxidizer impeller. Cyclebalance was achieved with a turbine speed of 27,500 rpm and a turbine bypassflow of 40% (Fig. 75). As an alternate, the turbine could be modified by in-creasing its flow area. This modification would reduce turbine speed and turbinebypass flow. In addition, less trimming of the oxidizer pump would be required.

° Cycle EXIIA

The turbomachinery for this cycle either combines scaledversions of the turbopumps selected for Cycle EXOI, or represents a plug clusterof double the thrust level. The dual turbopumps provide a redundancywith reduced thrust capability in the event of one turbopump failure

° Cycle GGOI

This cycle utilizes a modified RLIO derivative IIA turbo-pump. A redesigned turbine is required to accommodate the hot combustion pro-ducts from the gas generator. The fuel pump impellers are trimmed (similar toEXOI) or redesigned to provide the required propellant pressures. Heat shieldsmay be required to avoid excessive heat flux into the gearbox.

129

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LL

Turbine Inlet Temp - 263°K (474°R)Fuel Pump Speed - 27,850 rpm(Fixed by LOX Pump Head)

2,5 -

2.0 -

1.5 -

(P

I=

l.O -

0.5

0 , I35

|

5OO

Figure 74.

TurbineBy-PassF1ow 39%

Fl owLimits

C) From Preliminary Power Balance

[Z] Untrimmed Fuel Pump

Trimmed Fuel Pump

I I I 1

40 45 50 55

_uel _Pump Di_c_rg_ Pcessu_e, ATM

I I

600 700

Fuel Pump Discharge Pressure, psia

Expander Cycle EXOI Power Balance With RLIO Turbine.

I

800

133

M

Turbine Inlet Temp - 278°K (500°R)

Fuel Pump Speed - VariedLOX Pump Trimmed to Required Head

U

&/3

r_

3t_r_

0

&J-

2r_

}--

1

2.5

2.0

_I .5 -

OP

l.O -E

S-

0.5

0 i35

500

Figure 75.

I _ Required for Power

Turbine

By-PassFlow 42%

_ne Flow Limit

0- Preliminary Power Balance

NF = 24,900 rpm

40

- Power Balance 10% LOX Pump Trim

I27,500

, ,I45 50

Fuel Pump Discharge PPessure, ATM

I I

600 700

Fuel Pump Discharge Pressure,psia

Expander Cycle EX02 Power Balance With RLIO Turbine.

J55

I

800

134

IllI

° Cycle GG02

This cycle utilizes a modified RLIOderivativeIIA turbopump. As with Cycle GGOI,a redesignedhot gas turbine is required.A heat shield maybe required to reduce the heat flux to the gearbox.

2. Conceptual Turbopump Design Analysis

Section III,E, established the state-of-the-art of turbopumps re-quired for this application. Utilizing these data and the NASA guidelines fromReferences 21 and 22, pumps were conceptually designed to meet the baselineSpace Tug requirements including a lO-hour life expectancy. Table XXV summarizesthe design criteria.

The low speed LOX pump consists of a high head inducer drivenby a five-stage hydraulic turbine, with the turbine drive fluid being taken

from the discharge of the high speed LOX pump. The design parameters arelisted in Table XXVI. The nondimensional pump performance map is shown inFigure 76, and the hydraulic turbine efficiency performance is presented inFigure 77. The turbopump cross-section is shown in Figure 78.

It is anticipated that the low speed pump will be cooled toliquid oxygen temperature prior to full speed operation.

The high speed LOX turbopump consists of a full shrouded singlestage centrifugal pump driven by a velocity compounded gas turbine such asshown in Figure 70. The shaft is supported by a spring-loaded angular contactball bearing cooled by liquid oxygen. The design parameters are listed in

Table XXVII. The nondimensional pump head-flow and efficiency performanceis shown in Figure 79 and the turbine efficiency performance is shown inFigure 80. It is anticipated that the pump will be cooled to liquid oxygentemperature prior to full speed operation. Tank head idle-mode and pumpidle-mode are a logical sequence. The turbopump cross-section is shown inFigure 81.

The liquid hydrogen turbopump consists of a fully shroudedsingle stage centrifugal pump driven by a velocity compounded gas turbine.The pump impeller has an inducer stage designed to provide a zero NPSH pumpingcapability. The shaft is supported by two sets of spring-loaded angular con-

tact ball bearings. The rotor axial thrust is supported by a self-compensatingthrust balance incorporated in the impeller back shroud.

The turbopump design parameters are listed in Table XXVIII. Thenondimensional pump head-flow and efficiency characteristics are shown in

Figure 79. The drive turbine efficiency is shown in Figure 80. The pump cross-section is shown in Figure 82.

Application of the conceptual turbopump designs to the variousplug cluster engine cycles are outlined in the following:

° Cycle EX03

For this cycle, low speed pumps with a hydraulic turbine

drive are used for both the oxidizer and fuel to provide two-phase flow pumpingcapability. The high speed turbopumps incorporate single-stage centrifugal

135

TABLEXXV- TURBOPUMPDESIGNCRITERIA

GENERAL PUMP REQUIREMENTS

I. Pump Zero Tank NPSH2. TPA Life Expectancy lO hours - 1200 Engine Starts

PUMPS

150 (I000) <Ns >600 (4000)

Ns _ 225 (1500) Preferred

N v_q N - rpmNs : (H) 3/4

Q - m3/min (gpm)

H - m (ft)

TURBINES

Full Admission

AaN2 <258 x lO9 (40 x lO9) - (Blade Stress Consideration) -

c.m2 x rpm2 (in 2 x rpm 2) (Ref. 21)

INDUCERS

Cm jz_NPSH) 2g ,C

Design Limits

NASA LIMITS

Cm

@ =_-_t > 0.06

I Fluid CLOX 2.3

LH2 I.3

S = 3,000 (20,000) High Speed LOX

= 15,000 (lO0,O00) High Speed LH2

= 4,500 (30,000) - Low Speed LOX

_ N ¢-q--

(NPSH)314

N- rpm

Q - m'3/min (gpm)

NPSH - m (ft)

SHAFTING

Ncl > 1.5 NDesign Ncl - Lowest Shaft Critical Speed

o (DN)3_ 1,361 atm; D - mm, N - rpm(Watt) N2 = 4.8 x lO5' o<

[(Horsepower) N2 o (DN)3 , o< 20,000 psi]= 5.26 x IOW

D- mm, N - rpm

BEARINGS NASA LIMITS

Fluid DN Limit

LOX 1.3 x lO6

LH2 2 x lO6

(Ref. 22)

136

IIII

TABLEXXVI. LOWSPEEDLOXTURBOPUMPDESIGNPARAMETERS

PUMP- High HeadInducerFlow, m Kg/s (Ib/sec) = 14.1 (31)

R" NPressure ise, _-_ , (psi) = 0.14 x lO6 (20)

Flow, Q sm-_3(gpm_ = 0.0122 (Ig3)

Head Rise, m (ft) = 12.2 (40)

Specific Speed = 3500

Design Speed, rpm = 4000m

Tip Speed, _ (ft/sec) = 18.1 (59.4)

Head Coefficient = 0.37

Tip Diameter, cm (in) = 8.64 (3.4)

Efficiency, n = 0.68

NPSH, m (ft) = 0.68 (2.25)

Suction Specific Speed = 30,000

Inlet Flow Coefficient = 0.13

TURBINE - Five-Stage Hydraulic

Flow, _ Kg/s (Ib/sec) = 2.77 (6.1)N

Pressure _rop_ (psi) = 1.54 x 106 (223)m

Flow, Q _-- (gpm) = 0.0024 (38.6)

Head, m (ft) = 135 (445)

Pitch Line Velocity m/s (ft/sec) = 9.3 (30.5)

Pitch Diameter cm (in) = 4.39 (I.73)

Blade Height, cm (in) = 0.254 (O.l)


137

1.4

Noncavi tatin g

1.0

0

00.4 0.6 0.8 1.0 1.2 1.4 1.6

0

Figure 76. Low Speed LOX Pump Dimensionless Performance Characteristics.

138

Iii ! _:

0.9

0.8

0.7

0.6-

t-

"G

_0.5-q.L_

O.

O. --

0,20.3

f

Figure 77.

• J

I _ I II

0,4 0.5 0.6U - Blade SpeedC-o Ideal Nozzle Velocity"

Effect of U/Co on Turbine Efficiency, Single Impulse Stage.

139

v

_J

QO

0 I!0

II °_

U

r.-C_

v

E

I !

eaE

o0

X0

I'--

%-I

0

C0

0

-r'-I.I-

140

Ill T

TABLE XXVII - HIGH SPEED LOX TURBOPUMP DESIGN POINT PARAMETERS

PUMP - Single Stage Centrifugal (Fully Shrouded)

Flow, m Kg/s N(Ib/sec) = 16.83 (37.1)

Pressure Rise, _-_ (psi) = 3.1 x I06 (450)m3Flow, Q__ (gpm) = 0.0146 (231)1S

Head Rise m, (ft) = 274.3 (900)

Specific Speed, Ns = 1,589

Design Speed, rpm = 17,176

m (ft/sec) = 75.3 (247)Tip Speed

Head Coefficient, _ = 0.475

Efficiency, q = 0.68

Tip Diameter, cm (in) = 8.38 (3.3)

INDUCER

NPSH, m (ft) = 9.37 (31)

Suction Specific Speed, S = 20,000

Tip Diameter cm (in) = 5.00 (I.97)

Tip Speed m/s (ft/sec) = 45.1 (148)

Inlet Flow Coefficient = 0.18

m3Flow, Q_- (gpm) = 0.0146 (231)

TURBINE

Velocity Compounded - Full Admission

Pitchline Blade Speed, Um, m/s (ft/sec) = 137 (450)

Pitch Diameter, cm (in) = 15.24 (6)N

= 0.325 x lO6 (47.2)Inlet Pressure _ (psi)Inlet TemperatUre °K (°R) = 922 (1660)

Exit Pressure (static) N (psi) = 0.I09 x 106 (15.8)

Ideal Nozzle Velocity, _o m/s (ft/sec) = 1936 (6351)U-mm = 0.071CoEstimated Efficiency, n = 0.32

Flow Kg/sec, (Ib/sec) = O.ll (0.245)

Exit Annular Area Aa, cm2 (in2) = I02 (15.8)

AaN 2, cm2 x rpm 2 (in 2 x rpm2) = 30.6 x I09 (4.60 x lO9)

BEARINGS

Bore m (in) = 25 (0.984)

DN mm x rpm = 0.429 x lO6

SHAFT CRITICAL SPEEDS

Not Determined

IIncludes flow to hydraulic turbine drive for low speed pump.

141

O

El E

"c)

H

1,2 o !I

1,0

0,8

0,6 L1 ,,

C_) = Design Value0

0,4

0,2 --

00,4

(_ = Design ValueO

Noncavitating

I I I I I I0,6 0,8 1,0 1,2 1,4 1,6

Figure 79. High Speed L0X and LH2 Pump Dimensionless Performance Characteristics.

142

i'.| I

0.7

0.6

0.5

_0.4

"G,f,-

W-4-

"'0.3

0.2

O.l

I i L i L

0 0.05 0.I 0.15 0.2 0.25

U Blade SpeedCo Ideal Nozzle Velocity

I

0.3

Figure 80. Effect of U/Co of Turbine Efficiency, Velocity Compounded Stage.

143

1I

r-,

_Q

v

r =-

vl;..

_N

O _r- _

U_ -.r-

aJ'_

,---__%---

0I,_ I (_J _

C_I I

• _ E m

X0

.._

4_

_J

_Jt-

OiJ

CO

_n

144

T',!I

TABLE XXVIII - LH2 TURBOPUMP DESIGN POIIITPAPAMETERS

PUMP - Single Stage Centrifugal (Fully Shrouded)

Flow - _ Kg/sec (Ib/sec) : 2.45 (5.4)

Pressure Rise, _ (psi) = 2.67 x lO6 (391)

Flow - m3 (gp_)m_= 0.0355 (563)sec

Head Rise, m (ft) = 3992.9 (13,100)

Specific Speed, Ns = 150O

Design Speed, N = 77_400 rpm

Tip Speed m/s ft/sec = 295 (967)

Head Coefficient _ = 0.45


Tip Diameter, DT, cm, (in) = 7.37 (2.9)

INDUCER

NPSH m (ft) = 15.7 (52)

Suction Specific Speed = 95,000

Tip Diameter cm (in) = 5.84 (2.3)

Tip Speed m/s (ft/sec) = 237 (777)

Inlet Flow Coefficient, ¢ = 0.065

TURBINE

Velocity Compounded - Full Admission

Pitch Line Blade Speed, Um m/s (ft/sec) = 427 (1400)

Pitch Diameter cm (in) = I0.5 (4.14)N

Inlet Pressure m-2 (psi) = 6.55 x lO5 (95)

Inlet Temperature, K (°R) = 922 (1660)

Exit Pressure (static), _ (psi) = 0.I09 x lO5 (15.B)

Ideal Nozzle Velocity, CO m/s (ft/sec) = 2,219 (7,281)

u-m : o.192CoEstimated Efficiency : 0.62

Flow, Kg/sec (Ib/sec) = 0.0837 (0.185)

Exit Annular Area, Aa, cm2 (in 2) = 33.55 (5.2)

AaN2, cm2 x rpm 2 (in 2 x rpm2) = 210 x 109 (31 x 109 )

BEARINGS

Bore mm (in) : 20 (0.787)

DN mm x rpm : 1.55 x lO6

SHAFT CRITICAL SPEEDS

Not Determined

145

im

0

o

ii

.o

i._r-

E

1.1-I, c_

I.I_ "r"

v

II

e-

_ v

,--_%,_

v • I

I _"_,J _/'_

_ _ _ ""-• E4-_

_ C_ l- Q,I ¢o

_ I .t_ .i _

(_, i-- _ -r-

i'm

1"

01,-

-r.-I

e-

l'--

U¢--

0

(._

CO

,p-

146

!!! :I" !

pumps and two-stage turbines. At the design point, the turbine bypass flowis no greater than 20 percent. A bypass around the plug cooling passages,

rather than the turbine, could be used if the heat flux from the plug shouldprove inadequate to heat all the oxidizer flow.

° Cycle EX04

The turbomachinery for this cycle is similar to that used for

Cycle EX03 except that both turbines are provided with gaseous hydrogen.An oxygen seal package similar to that used in the RLIO is required between

the oxidizer pump and its drive turbine.

° Cycle GG03

The turbopumps utilized for this cycle are single-stage, cen-trifugal, high speed pumps, and hydraulic, turbine driven, low speed pumps.

An oxygen seal package is required for the high speed oxidizer pump. The hotgas turbine is a velocity compounded or Curtis stage.

D. ENGINE COOLING ANALYSIS

I. ITA-Type Module

Cooling analyses conducted early in the program on film-cooledskirts for both 40:I and 200:I nozzles (Figure 83) led to recommended designpoints of 21.5% FFC (fuel film cooling) and 24% FFC, respectively, for a

2560°R wall with cycle life of 1200. A reexamination of the film coolingrequirements of the 40:I nozzle was made using the results obtained onContract NAS3-20107 (Plug Cluster Module Demonstration Program).

The effect of increasing the module area ratio from 40 to 200

on module film cooling requirements was evaluated using the entrainmentfraction model. There is some uncertainty as to how to apply the entrainmentfraction, k, data obtained with a 40:I nozzle to a 200:I design so resultsfor two representative assumptions were obtained as shown in Figure 83. The"k vs x" model assumes that the axial entrainment fraction distribution in

the 200:I nozzle is the same as in the 40:I nozzle. This assumption yieldsmaximum nozzle wall temperature about equal to the 40:I nozzle prediction,thereby indicating that the film cooling requirements are about the same.

The "k vs x/L" model assumes that the entrainment fraction

correlates with non-dimensional axial position rather than the absolutevalue of axial position. This assumption yields high nozzle temperature for

a 200:I design which means that higher film coolant flow rates are required.

It is believed that the "k vs x/L" model is most likely torepresent the entrainment fraction distribution in a 200:I nozzle because

the entrainment is strongly influenced by wall curvature which tends to

correlate with x/L. Therefore, it is believed that a 2 - 3% FFC percentageincrease would be required if the module area were increased from 40:I to 200:I.

147

E Module

40/I

_ 200/120011

30001

5000 -

2500

4000 -

2ooo

L

3000 -'L_J

15ooGJ

"_ o2000

_000

n_

aJ(3.E(1J

I000.- 500%

0 0

Figure 83.

NozzleEntrainment

Fraction

APS DatakvsXk vs X/L

r- L200 "-

Max Temp Range

ITA DesignO/F = 5.5

Pc : 20.4 ATM (300 psia)

139°K (250°R) H2208°K (375°R) 02

(K vs S/L)

Throat Temp-/_'. _"

1422°K (2560°R)Maximum uesirableNozzle Temperature

Max. Temp.40/I Nozzle

(K vs X)

I0

I I I I

15 20 25 30% 'FFC

ITA Wall Temperatures Based on Entrainment Fraction Model.

148

;II! I;

The effect of overall module mixture ratio on film cooling

requirements and the parametric relationship between film-cooling-per-formance loss, mixture ratio, and fuel film cooling percent are shown in

Figure 84. These results were obtained for the ITA engine configuration butthe APS film coolant injection sleeve design was assumed because it is more

efficient (less coolant mixing). The analysis was performed using the HOCOOLcomputer program and the post-test entrainment fraction model. The designcriteria used for defining the film coolant requirements are: cyclic life of

1200 cycles (throat limit - includes safety factor of 4), I% maximum creepin lO hour (nozzle limit), 2560°R maximum nozzle temperature, and 1660°R film

coolant injection sleeve temperature (copper material).

2. Re_enerativel X Cooled Module

Preliminary regenerative cooling analysis of a module witha 40:I nozzle area ratio was conducted using the following ground rules:

(1) chamber and nozzle are entirely fuel cooled with no film cooling,

2) chamber pressure is 20.4 atm (300 psia) or 34.0 atm (500 psia),3) mixture ratio is 5.5, and (4) total cycle life is 1200 cycles. Azirconium copper chamber with rectangular coolant passages, similar to the

ITA design was used.

Gas-side boundary conditions were based on data generated

in Ref. 18, in which heat fluxes were calculated from gas-side thermocoupleresponses by means of a two-dimensional SINDA model. Test hardware wascomparable to the ITA design, i.e., premix injector, identical chamber con-tour. The present analysis was based on the reactive gas-side model and a

reference temperature equal to the mean of the wall and the recovery tempera-

ture. The correlating factor, Ca, was adjusted to make the predicted fluxprofile agree with the data of R_f. 18, as shown on Figure 85.

Coolant side heat transfer was based on the Hess and Kunz

correlation with a constant correlating factor, CL, of 0.0208. The walltemperature distribution in the coolant correlation was based on the bulk

temperature over the land and external wall and on the centerline walltemperature over the internal wall. As described in Ref. 23, this formulationmatched the results of two-dimensional SINDA analyses reasonably well.

Channel geometry was not varied extensively. The sixty channelITA design, with a channel width of 0.152 cm (0.060 in) was used as a start- _iing point, and when found satisfactory, the channel depth was varied to obtainthe change in wall temperature with channel depth. In the expansion section,both constant channel width and constant land width configurations were investi-gated. Both are satisfactory, but the constant channel width design will beexcessively heavy, while the constant land width design leaves a large span

across the coolant passage. Although not modeled, a bifurcation to double thenumber of channels will reduce both the weight and the span.

149

40 -

6

U90-_ 30 -

Uf-

04-

20 -

E• r =

OO

(..)

E

•_ lO

<1

0 ,5

150

Design Criteria:

NfT- -> 1200 cycles

I0 hr life (1% creep)

1422°K (2100°F) max. nozzle temp.g22°K (1200°F) max. Cu sleeve temp.

Is = 398 sec

Is Values for:4O/l Nozzle

Amb. Temp. 02/H2

= 435 sec

Is : 449 sec

Figure 84.

lO 15 20 25 3O

% of Fuel Film Cooling

Note: Results based on

Ta_kra_XPnCtMDFracti°n M°del / / / /

= 40/I /

Acmble;_. Te_p_4 (300 psia) // /

I . I , ... I - I .. IlO 15 20 25 30

% of Fuel Film Cooling

Film Cooling Requirements for 1200 Cycle Life at Mixture Ratiosfrom 4 to 7.

15"

lO -

U_p

c-

mr_

r--b-

5 -

-r

25 -

20 -

15

IJ-

5 -

0

0 Gas-Gas Program (Ref. 18)

Predicted

n

00

0

5 10 15

CM• I . .

5

Axial Distance (in.)

I ,I20 25

I

I0

F_gure 85. Gas-Side Heat Flux Profile.

151

Figure 86 shows a partial computer program output for a0.508 cm (0.20 in) channel depth coflow design in which the channel width isconstant between the injector and a point 8.89 cm (3.5 in) past the throat,and the land width is constant at 0.508 cm (0.20 in) thereafter.

The results of the analysis are summarized in Figures 87-92

for the 20.4 atm (300 psia) chamber pressure case.

Figure 87 presents the predicted gas-side wall temperature andtemperature drop from gas-side to back-side at the maximum temperature point --1.27 cm (0.5 in) forward of the throat -- as a function of channel depth for

the coflow design.

Figure 88 shows similar data at the same location in the counter-

flow design. Temperatures and gradients are comparable to those of the co-flow design. However, at the injector and the gas-side, temperatures areconsiderably higher than near the throat, as shown in Figure 89, although

the gradients are much reduced.

The predicted coolant pressure drop for both the coflow andcounterflow design is shown in Figure 90. The loss coefficients at the entranceand exit are taken to be 0.5 and l.O respectively, with the friction drop based

on a surface roughness of 0.000163 cm (0.000064 in).

The corresponding data for a module operating at a 34 atm (500

psia) chamber pressure is compared in Figure 91 at the injector (forward) endandnear the throat.

From the standpoint of thermal considerations, there is an ample

margin on wall temperature, pressure drop, and flow velocity for the regenera-tively cooled zirconium copper module with a gas-side wall thickness of 0.152cm (0.06 in.). Thus, the mechanical design (see Section VI,F) can concentrate

on integration of the module most effectively into the entire system, minimiz-ing module weight, cost and fabrication complexity.

Nickel and stainless steel chambers were also considered on a

preliminary design basis. Figure 92 shows the maximum wall temperatures pre-dicted for Nickel-200 as a function of channel depth, for a counterflow 60

channel design. The pressure drops are low and a reduction in wall tempera-ture appears feasible within a reasonable pressure schedule.

The stainless steel design produced wall temperatures in excess

of 1422°K (2100°F) above the throat for a channel depth of 0.406 cm (0.160 in);it is not apparent that reasonable temperatures can be achieved without an

extensive design effort.

3. Regeneratively Cooled Plug Nozzle

The geometry of the plug cluster engine analyzed was assumed toconsist of ten 40:I area ratio modules distributed around a 400:I area ratio

plug. The configuration is sun_arized in Figure 93 and Table XXlX.

152

2

.i

g:z

,11

u_ .

=

c

-u

..r

g

o.

o°

.J

>

p-

(.0

u_

>-

2 _r.<

zq

t

o

J_

I!

z

ii

g ;J _

Li z ,*_

%

_ _u

g

2

,e

_J2

:z

4

< l

r "

.I

_ ,.5 ._

= ."

a%

m

_ ._ ..-

- _. : -_:

.................... ..%-_,.,._,,,,.,,

N-

O

e-

%0

(..)

(1)

-r-

qJc-

O

IJ..

o(..)

0

0

4Jc-

O

ID%-

e'-

I--

-r

G)

:,.,.,. t ._.... / _-': !... •.,

153

2

,<

C

.,:-1.

_..,,.aT

#5¢.,

M

r'o _ --

b

z

_c

,<c

ac

.T

q

:I r

,2ii

z _

z

_r

r ..;

i!

°

t

:g-

z

" rT

!

:i _ ___c

1

• 2

. T

.3.

_ _l

r

...,°,0,.,°,,..°,,°,_,,,

,,,o.,.,,o.,,,,.o,...o_.

_ _._ _._ _°_ _ _

_=..__._:..._ ....

. . , , , , • , • . , • , • • • . , • • • • • • • •c_

i1111

_ _'_ _ ?__ ....

• , • • o • , • . • • • • • • • • . • • • • • ,

• • , , . • , • , . • • , • . • . . • o • • • ..

0

C_v

c"

ri-

O

o

S-

e-_J

_C

o

O

4-.)

O

e-

"7-

l,.-

_J

e'-

I---

"1-

t

154

_I I

(°F)

I000

900

800

700

600

500

l) Coflow

2) 60 Channel Design3) Channel Width 0.152 cm (0.060 in)

4) Data for Max Temperature (I.27 cm[0.5 in] Forward of Throat)

5) Mat'l: Zirc-Copper

6) Pc = 20.4 atm (300 psia)

Tint

_J$-

4-}

I.

_J

_JF--

700 -

600 -

500 ,,I I I I I0.3 0.4 0.5 0.6 0.7

cm

T_wg2

TBS

I0.8

Figure 87.

k I I0.I 0.2 0.3

Channel Depth (in)

Predicted Coolant Passage Temperatures, Down Pass (CoFlow) Design.

155

(°F)

lO00

900

800 -

700 -

600 -

500 -

- (°K)

80O

700

QJ

t_

r_E

600

-50O

l) Counterflow2) 60 Channel Design3) Channel Width 0.152 cm (0.060 in)4) Location 1.27 cm (0.5 in) Forward

of Throat


6) Pc = 20.4 atm (300 psia)

Twg3

iTint

__ 3 (870)

___Twg3

(760)

Twg_(671)

___Twg2_Tint

(514)

Twg2

___J

TBS

t J .....I, I I I

0.3 0.4 0.5 0.6 0.7 0.8cm

Figure 88.

[ J JO.l 0.2 0.3

Channel Depth (in)

Predicted Coolant Passage Temperatures, Up-Pass (Counterflow) Design;-

156

I!|-

(OF1000

90C

80C-

700-

60C

500-

400 "

300!

Figure 89.

(°K)800

700

600

GJ

m:

I.-

5OO

400

Twg3

l) Counterflow

2) 60 Channel Design

3) Channel Width 0.152 cm (0.060 in)4) Location: Forward End


6) Pc = 20.4 atm (300 psia) Tint

TW TwTnI I I I ,,i

0.3 0.4 0.5 0.6 0.7cm

Twg 2

TBS

I0.8

I, , ,,,_ , I , I0.I 0.2 0.3

Channel Depth (in)

Injector End Predicted Coolant Passage Temperatures, Up-Pass Design.

157

(psia)

50 -

40 -

3O

2O

I0

(atm)

4, --

- 2.0-

0

_1.0-

_ 0

I) 60 Channel Design2) Channel Width 0.152 cm

(0.060 in)

3) Pc = 20.4 atm (300 psia)

Uppass

Downpass

I I I I ,I I

0.3 0.4 0.5 0.6 0.7 0.8cm

IO.l

Figure 90.

I ! I111

0.2Channel Depth (in)

Predicted coolant Pressure Drop.

I0.3

158

If!TT:

Counterflow

ChannelDepth0.51 cm (0.2 in)Pc = 20.4 atm (300 psia) Pc = 34 atm (500 psia)AP: 2.04 atm (30 psi) AP 5.58 atm (82 psi)AT : 223°K (402°F) AT : 194°K(350°F)

778 ib I 800°K(940)| _(980° F)

At Injector

819 850

(lOl5 __.__](1070)

581 _ 579 536 _ 535

(585) (583) (505) (503)

1.27 cm (0.5 in) Forward of Throat

713

(823t

381I

(226)

794 825

736 (970 _ (I025)

______j(865)

-_382 389 --_" 389

(227) (241) (240)

Figure 91. Comparison of Counterflow Design Temperatures at Chamber Pressuresof 20.4 and 34 ATM.

159

(°F)

200O -

1500 -

lO00 -

Figure 92.

l) Counterflow

2) 60 Channel Design3) Channel Width: 0.152 cm 0.060 in4) Location 1.27 cm(0.5 in) Forward of Throat5) Mat'l: Nickel -2006) Gas Side Wall Thk: 0.076 cm (0.030 in)7) AP 2.72 atm (40 psia)

8) Pc = 20.4 atm (300 psia)

(°K)

1300 -

1200 _

llO0 -

I000 -

E

900 -

Twg2 - TBS -_ J

Twg3-Tint_

f \_ Twg3

Twg3I

Tint

i I I I I I0.3 0.4 0.5 0.6 0.7 0.8

Twg2

TBS

Channel Depth (cm)

I I IO.l 0.2 0.3

Channel Depth (in)

Predicted Coolant Passage Temperatures, Up-Pass (Counterflow) Design

Using Nickel - 200.

160

1111 [-

/

II

O

I

1113

•u[ 'sn_p_l 6ntd '_1

L I0

16:

LI. I_ 0 u4_ 0 _ _ 0 0

•,- Or'- u3 ,'-- tO C,.I _0 l'-,"00CL E I",. oO U'_ CD U'> tO C_

k-

m_,,t

ILl

13-

IX

X

c_F--

00_"

II

.._

0

If

0

N

0_ _ ,-- 0

r-- 0 0 0 0 0_I o_I

162

A heat flux distribution for the plug nozzle was estimated

using a Bartz-type equation for heat transfer coefficient (Ref. 24), andCornell data for the base heat flux distribution (Ref. 25), as indicated

on Table XXIX. The heat fluxes estimated range from 703 kW/mz (0.43 Btu/in_

sec) at the module exit to 85 kW/mZ (.052 Btu/inz sec) at the base outer radius.These heat fluxes were calculated using the mass flux adjacent to the plug

wall along the centerline of a module and circumferential heat flux variations wereneglected. Data relative to the variation in mass flux between modules was notavailable; however, it is conceivable that a lower mass flux, normally con-ducive to a lower heat flux, will exist along the centerline between modules.On the other hand, it appears that shock phenomena will tend to increase the

plug wall heat flux between modules. Due to these uncertainties, the accuracyof the estimated heat flux is probably on the order of + 50%. Experimental

plug heat flux data are needed to determine the extent of circumferentialvariations and to verify the heat flux magnitude along the module centerline.The Ref. 25 heat flux data are for a plug cluster engine with zero gap and

are therefore not entirely applicable.

A radiation cooled plug was also considered. The wall tempera-tures estimated for a radiation cooled plug are listed in Table XXlX.

These temperatures range from I067-1867°K (1460-2900°F).

Table XXX presents the results of plug energy balance calcula-

tions which yielded coolant outlet temperature for two plug cluster engine con-

figurations: l) a 40:I area ratio module and a 400:I area ratio plug, and 2)a 200:I area ratio module and a 400:I area ratio plug. Oxygen and hydrogen

coolants were considered. For the oxygen cooled case, the entire plug can becooled with liquid oxygen if the module area ratio is 200:I, but a two-phase

cooling system is required if the module area ratio is 40:I. If all of the hydro-gen is utilized as a plug coolant, the outlet temperature would range from 40-

106°K (80-190°R) depending on the module area ratio.

The feasibility of oxygen cooling of the plug cluster engine

depicted inFigure 93 (40:I module, 400:I plug) was investigated. The corre-lations used to evaluate the oxygen heat transfer coefficient were obtainedfrom Refs. 26-29 and are summarized in Table XXXI. The estimated critical heat

flux for subcooled and two-phase oxygen is also plotted in Figure 94. Thesecritical heat flux estimates are a crucial factor in evaluating the stainless

steel coolant channel design and need to be verified experimentally.

° Counter Flow vs Parallel Flow

The initial analysis objective was to determine the bestinlet location for the oxygen since it enters the cooling passages as a liquidand exits as a gas. It was found that the counterflow arrangement indicated inFigure 93 is best. This is demonstrated on Figures 95 and 96 which show plug

163

TABLE XXX - PLUG ENERGY BALANCE CALCULATIONS

AhZQplug KJ/kg

_rn _p]u_ kW _Btu/sec) Coolant (Btu/lb) Tin °K(°R)

40 400 3278 (3109) all 02 237 (102) 92 (165)

40 400 all H2 1302 (560) 22 (40)

40 400 75% H2 1732 (745) 22 (40)

40 400 50% H2 2603 (1120) 22 (40)

40 400 25% H2 5207 (2240) 22 (40)

200 400 1034 (981) all 02 75 (32.2) 92 (165)

200 400 all H2 411 (177) 22 (40)

200 400 75% H2 549 (236) 22 (40)

200 400 50% H2 823 (354) 22 (40)

200 400 25% H2 1646 (708) 22 (40)

Tout°K(°R)

157 (283) (G)

I04 (188)

132 (237)

185 (333)

354 (638)

132 (238) (L)

46 (83)

53 (95)71 (127)

127 (229)

(I) Pin02 = 40.8 atm (600 psia), hin : -129 KJ/Kg (-55.4 Btu/Ib),w = 13.83 Kg/s (30.48 Ib/sec)

PinH2 = 43.2 arm (635 psia), hin = -188 KJ/Kg (-81Btu/Ib),

= 2.51 Kg/s (5.54 Ib/sec)

(2) Pout02 = 34 atm (500 psia), PoutH2 = 36.4 atm (535 psia)

Major Assumptions: I. Plug Twall - 533°K (500°F).

2. No film cooling effects on plug.

3. Heat flux proportional to pV at the plug

wall along module _ to the 0.8 power.

4. No circumferential heat flux variation on plug.

5. Oxygen Tsat = 147°K (265 °R) at 37.4 atm (550 psia).

164

!!! 'I

TABLEXXXI OXYGENHEATTRANSFERCORRELATIONSUSEDFORTHEOXYGENCOOLEDPLUGANALYSIS

Oxygen Heat Transfer Correlations

I. Subcooled Liquid 02 Heat Transfer

a. Forced Convection

ALRC 02 Correlation (Ref. 26) evaluated at typical Tbulk

and Twall

p = 34 arm (500 psia), Th = lll°K (200°R), Tw = 139°K (250°R)

hd-O5/(pV)-95= 2.30xlO-2_(l.O9xlO "4)

h - kW/m2-°K (Btu/in2 sec °F)

d - cm (in)

pV - Kg/m2-s (lb/sec ft2)

b. Burnout or critical heat flux (nucleate-to-film-boiling transition)

based on N204 data and per correlation (Ref. 27):

CBo = A + B VATsub, kW/m2 (Btu/in2 sec)

V ATsub

_m_K/s (ft °F/sec). A B<I097 (2000) 981 (0.6) 1.85 (.00062)>1097 (2000) 2451 (1.5) 0.507 (.00017)

c. Nucleate Boiling: TwL = Tsat + ATsH, (ATsH = 283°K or 50°F)

2. Gas 02 Heat Transfer

Approximation of ALRC correlation (Ref. 26) prediction for

P : 34 atm (500 psia), Tb = 153°K (275°R)

hd'05 = 1.82xlO-2(8.65xlO -5) [0.7 (Tw/Tb)-'B](pv).g5

3. Two-Phase 02 Heat Transfer

h - kW/m2-°I (Btu/in2 sec°F)

d - cm (in)

pV - Kg/m2-s (Ib/sec ft2)

Tw, Tb - oK (°R)

Film boiling based on Giarratano and Smith Correlation (Ref. 28)

x, quality h2ph/hgas

0 - .Of 0.25

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30 l.g3/pV

H' = Hf + .724 aHfg [(I - _ ) ] (Btu/Ib)

c. Nucleate Boiling: Twl = Tsa t + _Tsh, (ATsH = 283°K or 50°F)

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wall temperature at the module exit and at the downstream end of the plug asa function of coolant oxygen mass flux and the coolant state. At the module

exit (Figure 95) acceptable wall temperature can be maintained by liquid oxygencooling with a nucleate boiling of a forced convection mechanism but excessive

Mach number (>0.5) are required for gas. The results plotted on Figure 96shows that gas cooling is feasible at the downstream end of the plug as a walltemperature of about 811 °K (lO00°F) can be maintained with a coolant Machnumber less than 0.3.

° Coolant Channel Design

Preliminary design calculations for sizing the cooling chan-

nels of an oxygen cooled (counter flow) plug are summarized in Table XXXIi.A 92°K (165°R) liquid oxygen inlet temperature was assumed.

At the coolant inlet (z = 29.2 cm or ll.5 in), the oxygen is

a subcooled liquid and it is desirable to avoid film boiling. Consequently,the coolant velocity is governed by critical heat flux consideration. A

0.6 m/s (2 ft/sec) velocity is sufficient to yield _n adequate burnout safetyfactor (I.5). The required mass flux is 0.07 Kg/cm_s (l.O Ib/in_ sec).

At the next analysis station (z = 50.8 cm or 20 in), the heat

flux is lower and the oxygen is still significantly subcooled. As a result,the required velocity is lower 0.37 m/s (I.2 ft/sec).

At the third analysis point (z = 76.2 cm or 30 in), the oxygen

bulk temperature has reached the saturation temperature (147°K at p =37.4 atm, or 265°R at p = 550 psia assumed) and bulk boiling is beginning tooccur.

As long as the oxygen remains slightly subcooled, nucleateboiling can be easily maintained. However, after a certain amount of bulk

boiling occurs, it is difficult to maintain nucleate boiling on the coolantchannel walls.

For the fourth analysis point, z = lOl.6 cm (40 in), thecoolant is 48% vapor and the estimated critical heat flux characteristic

(Figure 94) indicates that reduced mass fluxes are required to maintainnucleate boiling and avoid film boiling. This is necessary to avoid annularflow where the liquid does not touch the wall. The approach is indicated byoption (a) for the z = lOl.6 cm analysis point where the channel area has beenincreased by a factor of 5. Options (b) and (c) are film boiling designs inwhich the wall is cooled to a 811-i367°K (lO00-2000°F) temperature by de-creasing the channel flow area (by a factor of 28-52) so that annular flow doesoccur but the gas velocity adjacent to the wall is sufficient to provide the re-quired cooling.

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The results obtained for analysis points 3, 4 and 5 showthereare two choices for the coolant channel design in the two-phaseregion: (I)a design which includes a two-phasenucleate boiling region, and (2) a design whichdoes not. The design which does not include two-phasenucleate boiling is con-sidered most practical for fabrication purposes. A design which does include anucleate boiling region would have a lowest pressure drop but would be extremelydifficult to design and fabricate since it would be necessary to first increasethe flow area by a factor of_five (decrease the massflux from 0.035 to 0.007kg/cmL s or 0.5 to 0.I Ib/in L sec), and then de_reaseit by a factor_of at least70 (increase massflux from 0.007 to 0.49 kg/cmL s or 0.I to 7 Ib/in _ sec).

Cooling of the plug base region is relatively straight-forwardsince it involves only gas-forced convection heat transfer.

° Pressure Drop Estimate

The pressure drop in an oxygencooled plug wasestimated byassumingthat the coolant channelswould be designed for film boiling in thetwo-phaseregion. The coolant channel geometryand pressure drop calcula-tions are summarizedin Table XXXlII.

Theestimated loss wasover 47.6 atm (700 psi) which is so largethat it probably rules out oxygenplug cooling as a practical concept. Mostof the pressure drop is estimated for the two phaseregion wherethe estimatedfriction loss is 39.8 atm (585 psia). The two-phaseflow APwasestimated usingthe Ref. 30 water data as indicated in Figure 97.

E. BASE PRESSURIZATION ANALYSIS

The early literature, summarized in Section III.C.I, Figures 15 and 16and Table IV, indicates that only a small relative base flow rate (about 0.2%)is required for a large area ratio plug nozzle operating in vacuum conditionswhere the wake is closed aft of the plug base. Analysis of these data forvacuum operation reveals that the base pressure, corresponding to the 0.2%flow, is 2.5 times the static pressure of the exhaust gas on the edge of theexpansion section of the plug. This value is recognized to be the standardseparation criteria (P_ > 0.4 Pamb_o-_) for DeLaval nozzles. (Also given asPe _ 0.28 Pambient f°r'hTgh area r_Y_ nozzles in Reference 42.)

Achievement of the optimum base pressure may or may not require afinite mass flow into the base, the amount presently being determined from ex-periment (See Table X and Figure 35). Since the amount of flow should be de-pendent upon both the diameter of the base and the pressure level, an equation(Eq. 17, Section IV,E.3) was formulated for the parametric analysis usingAerospike (Ref. 6) data.

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Wbase= K PbaseABase (Eq. !5)

In solving this equation for the flow rate, it is assumedthat the flowpressure is given by

Pbase= 2.5 Pe

wherePe is the ODEpressure at the engine area ratio.

The performanceimprovementobtained by a base pressurization correctionusing Eq. 15 is seento be between0.3 to 2.1% l to lO secondsspecific impulseas shownin Figures 46-50 of Section IV. This improvementseemsreasonablewhencomparedwith the 2.4%thrust increase due to base pressurization of the Aero-spike (see Section IV,E.3.)

The schematicsfor the various engine cycles utilizing base pressuriza-tion are given in Figures 61-70 (Section V.B.) In all cases, the performanceimprovementwassufficient to justify the additional weight (2 to 5 Ibs)required to achieve pressurization (see Section VI.)

Fm CONFIGURATION ANALYSIS

Plug cluster engine configuration layouts were prepared for candidatecycles utilizing modules with area ratios of 40,100 and 200:I. These layouts(Section VI.B), in conjunction with the parametric weight analysis (Section V.G.),and the parametric engine performance (Section IV.E.), led to the selection by

the NASA Project Manager of the configurations:

o ITA Module (_M = 40), a/De = 2

° Minimum Change ITA, _/De = 2

° Regeneratively Cooled Modules (_M = lO0), 6/De = l

The cycles selected for these configurations were the expander cycle (EX02) withan RLIO turbopump assembly and a gas generator cycle with a state-of-the-art

technology turbopump design. Both cycles were to utilize an H2-cooled plug.

G. PARAMETRIC WEIGHT ANALYSIS

For purposes of the parametric weight study, the plug cluster engine wasassumed to be composed of a combination of the following components:

° Regeneratively Cooled Combustion Chamber (WCC)

° Regeneratively Cooled Thrust Chamber Nozzle (WTCN)

° Thrust Chamber Nozzle Extension (WNOZ)

° Main Injector (WINJ)

174

° Ignition System (WIGN)

o Main Turbopump (with Gear Box) (WTPA)

° Main Turbopump (Parallel Turbines) (WTPA)

° Valves and Actuators (WV)

° Propellant/Gas Lines (WL)

o Gas Generator (WGG)

° Miscellaneous (Electrical Harness, Instrumentation, Brackets, EngineMount, Gimbal (WMISC)

° Plug Nozzle (WPN)

The engine dry weights do not include:

o Gimbal Actuators and Actuation System

° Engine Controller

° Pre-Valves

° Tank Pressurant Heat Exchangers and Associated Equipment

o Contingency (a total contingency is normally included in the vehicleweight statement)

Baseline engine weight statements were established for the expanderand gas generator cycle engines by comparing like components with the RLIOlIB, the single- and double-panel Aerospike, and the Advanced Space Engine.

These baseline weights were revised during the program to conform to thepreliminary conceptual design layouts. The initial component weights utilizedin the parametric analysis, are given in Table XXXIV (see Section VI for re-vised weights and a detailed breakdown by component).

TABLE XXXIV. PLUG CLUSTER BASELINE WEIGHT SUMMARY FOR PARAMETRIC ANALYSIS

Component

Baseline Weight kg (Ib)

Expander Cycle Gas Generator Cycle

WCC (per module) 2.12WTCN (per module) 3.2WNOZ (per module) 1.6

WINJ (per module) 1.88WIGN 9.9

WTPA (Gear Box) 31.8WTPA (Parallel Turbines)21.3WV I0.4

WL 17.4WGGWMISC 27.4WPN 38.8

Comments

(4.68) 2.12 (4.68) All modules(7.0) 3.2 (7.0) Regen Module Only(3.6) 1.6 (3.6) ITA Module Only(4.14) 1.88 (4.14) All Modules(21.9) ll.9 (26.3) All Chambers

(70.0) 31.8 (70.0) RLIO TPA(47.0) 21.3 (47.0) All Cycles(22.9) I0.4 (22.9) Will Vary for GG(38.3) 17.4 (38.3) Will Vary for GG

2.5 (5.6) GG Cycle Only

(60.5) 27.4 (60.5) All Cycles(85.5) 38.8 (85.5) All Cycles

175

With the baseline engine weight established, engine componentweightscaling relationships were derived as a function of thrust, chamberpressure,and nozzle area ratio. Thesescaling relationships were used to calculatethe weights over the parametric ranges of interest. The equations, whichwere established through geometryconsiderations and empirical data fits ofhistorical data (References23, 31, 32), were modified to obtain the best fitfor a variety of engine types (References5, 9, 14, 23, 32-34).

Theresults of the parametric weight analysis are presented in Figures98 and 99 for chamberpressure of 20.4 and 34 atm respectively. It is seenthat the engine weight for a modulearea ratio of 200 becomesexcessive. Forthis reason, configurations with 200:I moduleswere not selected for furtherstudy. The inclusion of AGCarbnozzle extensions later in this study, however,showedthat area ratios as high as 500:I could be utilized.

H. THRUSTVECTORCONTROLANALYSIS

Preliminary analytical evaluation of four basic thrust vector control(TVC)concepts for the plug cluster rocket engine wasaccomplished. Thefour concepts are gimbaling, throttling or engine out, hinged panels, andsecondaryinjection.

The initial evaluation involved an assessmentof the momentgeneratingcapability for all the concepts. The required TVCmomentgenerating capabilityis identical to that momentwhich would be generated by a 66.7 KN(15,000 Ibf)thrust engine operating at a gimbaled angle of 4 degreesor 21,280 joules(188,342 inch-pounds). The analysis and test information contained in Prattand WhitneyAircraft Report PWAFR-lOl3, Reference4, formed the basis for thisportion of the study. The information wasmanipulated to yield the lateralforce, the axial force, and the momentproducing displacement of the axialforce for the following TVCconcepts:

o Gimbaling - The only mode of operation considered is the so-calledhinged motion of a modular engine in a plane which intersects the plug nozzlecenterline. The corresponding moment generating capability is shown in FigurelO0 for I, 2, and 3 hinged modular engines. The required moment of 21,280 joules

(188,342 inch-pounds) can be achieved by this TVC scheme with one module at ahinge angle of 52 degrees.

° Throttling or Engine Out - The differential throttling of modules

will result in the moment generating capability shown in Figure lOl for 2, 3,and 5 throttled modular engines. The required moment can be achieved by

throttling 3 engines to approximately 12% nominal thrust.

o Hinged Panels - The hinging of a panel or flap consisting of a 60degree sector of the plug surface located at the upstream end of the plugnozzle will result in an estimated longitudinal force of 71,g81 N (16, 182

pounds) for a panel or flap hinge angle of 18 degrees. The assumed relationshipbetween moment generating capability and hinge angle is shown in Figure 102.Note that the desired moment can be achieved with an estimated hinge angle of

34 degrees.

176

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E

E

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200

I00

40

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Figure 98. Plug Cluster Engine Dry Weight, Pc = 20.4 ATM.

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Figure 99. Plug Cluster Engine Dry Weight, Pc - 34 ATM.

178

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PLUG LENGTH = 15% OF ISENTRORIC

MOOLILETHRUST - (1500 LBF) 6672 NTILT kl_E = 17oNUMBEROF MODLES - 10TOTAL ENGINE THRUST - (16302 LBF) 72.51 KN

LRe- (63.12 IN.) 160 cm(151.2 IN.) 384 cm

200,000

160,000

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-66.7 KN THI_T

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A 2 ENGINES Nll_

[] I ENGINE HINGED

10 20 301 1 I l

40 50 60 70

HINGED ANGLE, DEGREES

I I I80 90 I00

HINGED ANGLE

VEHICLE

C.G.¢ VEHICLE

_MOMENTS = FLL + FAd

Figure lO0.T _

Moment Generating Capability for Hinged Engine Model Concept.

179

240,000200,000

i 80,00040,000

U5

0

30,000 _

m

2S, 000

20,000

I0,000

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100

PLUG LENGTH = 15% OF ISENTROPIC

MODULE THI_UST= (1500 LBF) 6672 NTILT ANGLE = 170

NUMBER OF NODULES - 10

TOTAL ENGINE THI_UST- (16,302 LBF) 72.51 KN

Re = (63.12 IN) 160 cmL = (151.2 IN) 3.84 cm


-66.7KN THRUST

gO

5 ENGINESTHROTTLED

3 Era61NESTHROTTLED

2 ENGIWESTHROTTLED

I I I _i I I I I80 70 60 50 40 30 20 10 0

PERCENT THRUST OF THI_TTLED ENGINES

THROTTLED ENGINE

VEHICLE

E.G.VEHICLE

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Figure lOl. Moment Generating Capability for Throttled Engine Module Concept.

18O

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TOTAL ENGINE THRUST = (16302 LBF) 72.51N

d/Re = .012, FL/FA- .035, FA - (16182 LBF)

71.98 N @ 18o HINGE ANGLE

R = (63.12") 160 cme

L = (151.2") 384 cm

m

2O ,000

15,000

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_ 5,000


-66.7 KN THRUST

ONLYAVAILABLEDATAPOINT- /

//

_ ofl I I I 1 I ! I 10 4 8 12 16 20 24 28 32 36

HINGE ANGLE, DEGREES

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C.G. _ 'L

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VEHICLE

Figure 102. Moment Generating Capability for Hinged Panel Concept.

181

o Secondary Injection - The injection of fluids into the supersonicprimary stream produces a shock wave which causes an asymmetrical pressure dis-tribution on the plug nozzle wall. The resulting laterial force and axialforce displacement characteristics were evaluated for both gas slot injectionat the end of the plug and for gas injection at the outside diameter (OD) of50 percent of the modules. The moment generating capability for slot injection

is shown in Figure I03. Only slot injection yields the required moment andthis occurs at a relative weight flow of 4.5 percent. Relative weight flow isdefined as the ratio of secondary injection flow to total primary flow convertedto percent.

The next phase of the preliminary evaluation of the four basic TVC con-

cepts involved an estimation of the control hardware characteristics necessaryto achieve the equivalent of the following gimbaled single engine requirements.

Gimbal angle : 4 degrees

Rate = 4 deg/sec

FrequencyResponse : flat to 5 hertz

Table XXXV contains a tabulation of the estimated control hardware

characteristics. The actuation system estimated weights for hinged moduleswere found to be similar for both hydraulic (including pump) and electromechan-ical systems using historical data from past programs. The throttling, engineout, and secondary injection systems will require flow control valves which,

in turn, must be operated by an actuation device. The estimated pressure drop

and weight flow requirements were converted to a fluid K requirement [KW ._eWeightflow/(pressure drop x specific gravity)I/2]. An array o_ historical date -

garding the use of LOX and LHR valves and actuation devices on past engineprograms was likewise arrangea as a function of KW and provided the basis forthe estimation of both the weight and envelope dimensions. Electromechanical

actuation was presumed for the valves in this study based upon past actuationtrade studies.

A Summary evaluation of the TVC schemes proceeds as follows:

° Gimbaling - A minimum of four hinged modular engines are requiredto achieve pitch and yaw control with a total actuation system weight ofapproximately 36.3 Kg (80 pounds). The required maximum module hinge angle isapproximately 53 degrees. There are questions involving the mechanics ofhinging the modules through a large angle that remain to be answered.

° Hinged Panels - A minimum of four hinged panels are required to achievepitch and yaw control with a total hydraulic actuation system weight of approxi-mately 83.5 Kg (184 pounds). The required maximum panel hinge angle of 34degrees is based upon very little data and more information is necessary todetermine the effects of panel shape, size, location and hinge angle. A 34degree panel hinge angle raises questions concerning the erosion of the panel

while deflecting the combustion gases.

182

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0

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240,000

>. ,., 200,000ac --J•:c r-ae-_ N

160,000

-J ,.,1

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_ "' 80,000

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30,000 L

25,000

_q 20,000

I0,000

5,000

PLUG LENGTH = 15% OF ISENTROPICINJECTION SLOW DIMENSION : 400 CIRCULAR ARC

TOTAL ENGINE THRUST = (16302 LBF) 72.51KN

GAS INJECTION

Re = (63.12") 160 cm

L = (151.2") 384 cm

4o GIMBAL MOMENT

(15,000 LBF) -66.7 KN

THRUST

0 I I I0 I 2 3 4 5 6 7 8

RELATIVE WEIGHT FLOW, PERCENT

45o

VEHICLE I_

]E]MOMENTS = FLL + FAd

INJECTION SLOT

_VEHICLE

Figure I03. Moment Generating Capability for Secondary Injection Concept.

183

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° Throttling or Engine Out - A minimum of three deeply throttled engine

modules are required to achieve the required moment in a pitch or yaw plane.As a result, all ten engine modules must either be operated in a throttledmode or in an on-off mode to achieve the full pitch and yaw control. In either

case, there will be two modules which will be required to respond to both apitch and a yaw conmand. This overlap of control for two modules is not aproblem if engine out or on-off control is used. In addition, engine out

operation eliminates the potential combustion stability problem associated withdeep throttling of engine modules. The total weight for the ten flow controlvalve/actuator combinations is approximately 150 Kg (330 pounds). The averagepower required to obtain adequate valve transient response is estimated to be75 watts per module.

° Secondary Injection - A minimum of four gas injection slots locatedat the plug base are required to achieve pitch and yaw control. The total

weight for the four flow control valve/actuator combinations is approximately69 Kg (152 pounds). The average power required to obtain adequate valve

transient response is estimated to be 50 watts per slot. Past tests at ALRCon the Minuteman secondary injection system indicate that the generated sideforces are directly responsive to changes in injectant flow rate for fre-quencies up to 20 hertz. This concept raises questions concerning the weight

and complexity associated with the hardware necessary to deliver the injectantto the flow control valve.

In conclusion, it appears from this preliminary analysis that hingingengine modules to achieve the required pitch and yaw control moments would bethe most desirable concept from the standpoint of axial force capability, weight,and reliability. If weight reductions of 20% are made in the near future through

the use of composite materials, the hinged module approach still appears themost promising.

185

I ,

SECTION VI

PRELIMINARY CONCEPTUALDESIGN

m. OBJECTIVES AND GUIDELINES

Preliminary conceptual designs of selected plug cluster engine systemswere generated based on the information developed in Tasks I-III for filmand regeneratively cooled systems. The detail of the designs allowed theapproximation of complete engine weights from which perturbations or trade-offs could be conducted to optimize the plug cluster engine.

Tradeoff and sensitivity factors between subsystem operating points,plug cluster engine geometry, plug cluster engine performance, and installedengine weight, were established for the nominal configurations of the plugcluster engine.

An engine component list was prepared and compared with those forcandidate engines in past Space Tug studies to assure that similar componentsand requirements were included in the weight statement. A common frame ofreference was thus established for the weight of the plug cluster engine.

Consideration of AGCarb, carbon-carbon cloth, lightweight structuresled to modification of portions of the conceptual designs. An AGCarbuncooled plug nozzle was investigated in depth. A cluster of large arearatio scarfed bell nozzles, with AGCarb nozzle extensions from E : 40 toc : 500, was also investigated.

B, CONCEPTUAL DESIGN

Conceptual design layouts were made for the expander and gas generatorcycle configurations. Typical layouts are shown in Figures 104-106. Theselayouts contain individual valving for the igniters and two additional mainpropellant valves. This number of control elements is more than considerednecessary for the minimum valve configuration. The selection of the moreconservative system is based on the preliminary controls analysis presentedin Section VI,E.

The RLIO turbopump assembly is shown for the expander cycle configura-tions (Figures 104 and 105), and a parallel turbine state-of-the-art TPA isshown for the gas generator cycle configuration (Figure 106).

Four different modules are utilized in the conceptual designs:

l_I Integrated Thruster Assembly (ITA), (2)Minimum Modification ITA, andRegeneratively cooled ITA with both a 40:I and a lO0:l module area ratio.These are discussed in Section VI.F.

Three types of fairings are shown in the figures: (I) straightfairings, which historically have shown the highest performance, (2) contouredfairings, which appear to add excessive weight, and (3) scarfed nozzles, wherethe uncut portion of the nozzle becomes the contoured fairing.

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Detailed discussions concerning the structure, materials and controlsfor the conceptual design configurations are given in the following sections.

C° STRUCTURES ANALYSIS

Stress analysis calculations were performed in support of the structuraldesign of the plug cluster engine configurations. Results of the moduledesign analysis are reported separately in Section VI.F. The various com-

ponents were designed to provide minimum weight by comparing the known loadsand stresses to stainless steel allowable strength values and critical bucklingloads. Critical stress modes, such as buckling, tension, and bending wereidentified for each component, and the following safety factors were utilized:

Safety Factor on Yield = l.l - 1.25Safety Factor on Ultimate = 1.4Safety Factor on Buckling = 1.25 - 1.4

Design criteria used in the calculations are:

Module Thrust =

Engine Thrust =Accelerati on =

Plug Nozzle Temperature =Life =Pressure Profile =

6672 N (15DO Ibf)68,058 N (15,300 Ibf)

0.2 g533°K (500°F)1200 cycles

(given in Figure I07)

The configuration with labeled structural components is illustrated in FigureI08.

An arrangement of brazed tubes and circumferential stiffeners was

found to be the lightest weight structure for the regeneratively cooled plugnozzle. This arrangement, shown in the sketches of Figure I09, utilizesfive equally spaced stiffeners for an assumed uniform external pressure loadof 0.04 atm (0.6 psi). If it is assumed that all of the radial load is

carried by the stiffeners, the total radial load per stiffener is 7784 N(1750 Ib), and the load per unit length is 947 N/m (5.4 Ib/in). For a

stiffener of cross section 3.81 cm x 5.08 cm (1.5 in x 2 in), the required

thickness for buckling stability is 0.025 cm (O.Ol in). The bending stressin the tube with a 0.23 m (9 in) span between support is 592 atm (8690 psi).The buckling margin of safety is 0.2 for the plug wall, where the margin ofsafety is defined as

M.S. - allowable stressapplied stress x safety factor -l

The lightweight module mount ring shown in Figure llO was designed basedon a required buckling load of 5940 N/m (33.9 Ib/in). The buckling margin of

safety of 0.5 and the bending margin of safety of 1.8 are obtained for thisstructure.

194

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The plug torus was designed to serve two functions: (I) distribute

coolant to the plug wall, and (2) serve as the structural member of attachingthe plug to the thrust structure. The minimum wall thickness for a marginof safety of zero is 0.058 cm. A wall thickness of 0.064 cm (0.025 in) wasselected.

The thrust struts (rod braces) were sized for 950 Ibf compression ineach 1.46 m (57.6 in) long strut. The strut wall thickness was selected as0.064 cm (0.25 in) for a zero margin of safety.

Both aluminum and phenolic impregnated fiberglass cloth honeycombstructures were analyzed for the base closure structure. The margin ofsafety proved to be large, allowing the use of commercially available thick-nesses.

D. MATERIALS ANALYSIS

The selection of materials for the plug cluster engine conceptualdesign (Figures 2 and I04 are typical) was based on propellant compatibility,required mechanical properties, and fabricability, with the primary emphasisbeing placed on compatibility. A listing of the material selected for eachengine component _ given in Table XXXVI.

The requirements of long life, low maintenance, postfire condensationand storage in coastal environments dictate the selection of materials with

high resistance to pitting, crevice and stress corrosion. Design effects suchas galvanic couples and theinfluence of fabrication, particularly on stressCorrosion cracking susceptibility must be considered.

Hydrogen incompatibility is manifested in metals by a loss of toughnessboth with decreasing temperature and hydrogen absorption. The low operatingtemperatures and pressures of the engine allow the use of austenitic stainless

steels which are both highly resistant to embrittlement by hydrogen absorptionand possess excellent toughness over the range of service temperatures. Theuse of the susceptible nickel base alloys will be limited to the possible useof an electroformed nickel close-out of the module zirconium copper chamberliner. Limited data indicate that as-deposited electroformed nickel is sus-ceptible to hydrogen embrittlement; however, sufficient ductility is retainedin the weaker, annealed condition to al:lowits use. The remaining selected

materials, i.e., copper and copper alloys, aluminum alloys and titanium alloys(under lO0°F) are highly resistant to hydrogen embrittlement.

Oxygen incompatibility is manifested in metal either by loss of tough-

ness at lower temperatures, reduction of fatigue life or catastrophic oxida-tion. With the exception of titanium alloys, the alloys anticipated for

hydrogen service will also be used in oxygen. These materials possess ex-cellent cryogenic toughness, and their ignition temperatures in oxygen are well

above their respective service temperatures. Ignition is not a problem exceptwhere aluminum alloys would be subjected to high energy inputs or where organic

199

TABLEXXXVI- MATERIALSELECTIONFORTHEPLUGCLUSTERENGINECONCEPTUALDESIGN

Component Material

Module Chamber/Regenerati ve

Plug NozzlePlug Aft End Closure

Plug Wall StiffenerPlug Support Rods

Engine Support RingThrust Struts

LOX Boost Pump and LH2Low Speed Pump

HousingsTurbine Nozzles

Impeller and Turbine RotorsBearingsShaft

LH2 High Speed Pump

Turbine HousingTurbine RotorsTurbine Nozzles

Pump HousingImpellerBearingsShaft

LOX Pump

Turbine HousingTurbine RotorsTurbine Nozzles

Pump Housing

ImpellerBearingsShaft

Zirconium Copper Liner EF Nickel Close-out

CRES 347 or Carbon-Carbon CompositeAluminum or Fiberglass Phenolic Honeycomb

CRES 347CRES 347CRES 347CRES 347

A356 Aluminum

6061T-6 Aluminum7075 T-73 AluminumCRES 440

CRES A-286

CRES 347 CastA-286CRES 347 Cast5AI-2.5Sn ELI Titanium

5AI-2.5Sn ELI TitaniumCRES 440CA-286

CRES 347 CastA-286CRES 347 CastA356 Aluminum7075 T-73 Aluminum

CRES 440CCRES-A- 286

20O

contaminantscould ignite and provide a secondarysource of energy to ignitethe metals.

All selected non-metallic materials will be limited to those whichare acceptable in accordancewith MSFC-SPEC-IOIand I06.

A fiber reinforced graphite composite is a candidate material for theplug nozzle. This material's chemical compatibility with combustiongases(water vapor and hydrogen) is excellent. Its calculated regression rate,due to reaction with water vapor, approacheszero at temperatures below2500°Fand is less than 2 mils/hr at 3000°F. Material regression due to reaction withhydrogenwasmeasuredat 4 mils for a ten hour test period at 3000°Fand 4psia.

E° CONTROLS ANALYSIS

The controls analysis was conducted in two parts: (1) for the fully

regeneratively cooled (modules and plug nozzle) engine and (2) for the engineutilizing an uncooled plug nozzle. The analysis of the regeneratively cooled

engine (Figure I04) is summarized in the first section. The second sectionoutlines the results of the study of the engine with an uncooled plug nozzle(Figure 2) where a minimum weight control system was devised.

I. Control System for Regeneratively Cooled Engine

Engine cycle schematics were prepared to correspond to theregeneratively cooled conceptual design configurations. The schematics shownin Figures Ill and ll2 are expander and gas generator (GG) cycles. Minordifferences occur in the cooling circuits for different modules (e.g., ITAand minimum modification ITA).

A preliminary evaluation of the valves and controls required for thetwo engine cycle concepts shown in Figures Ill and ll2 was performed to pro-vide a degree of confidence that the defined system schematics could control

the engine. This evaluation was performed in a very broad manner and did notinclude any formal analysis of system transients. The basic approach used wasto examine the original schematics for both the expander and GG cycles, identifypotential problem areas, attempt to minimize the control problems by adding,deleting or relocating controls and then to examine the revised system withregard to preliminary definition of controls and control modes. With this

approach, the resultant schematics should be representative of what will berequired; however, final definition will require programmed analyses to evaluatethe varied transient conditions that may be encountered.

For both concepts, the engine start would begin with tank head opera-tion through a cooldown phase followed by a pumped idle mode and then full

thrust operation. Although the basic approach is quite simple and has beenused successfully for other cryogenic, pump fed engines, the use of lO engine

modules presents some additional considerations regarding location and sequenc-ing of controls.

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The expander cycle, shown in Figure III, uses GH2 to drive a turbinewhich is coupled to both pumps by a gearbox. The oxidizer circuit has a tankshutoff valve, a single main oxidizer shutoff valve, a check valve (to control

oxidizer tank pressurization gas), and small shutoff valves located at eachmodule to control igniter flow. The fuel circuit has valves comparable in

function to those in the oxidizer circuit plus a bypass valve that serves to

control the flow through the turbine after the H2 has passed through the coolantpassages. Auxiliary sensing devices and an electronic controller will be re-quired to properly accommodate the various conditions under which the enginemust start and shutdown.

The GG cycle shown in Figure ll2 uses hot gas to drive parallelturbines, each of which is coupled to a pump. The valves in both the fueland oxidizer circuits are comparable to those defined for the expandercycle. In addition to the common valves, a fuel and oxidizer GG valve are

required. Also, the bypass valve is relocated and functions as a throttlevalve to control hot gas flow to the oxidizer turbine.

Based upon the examination of the systems, Table XXXVII was preparedto show a preliminary definition of the required valves.

For each valve defined, viable options exist dependent upon moredefinitive performance requirements. One major variable is the allowablepressure drop for the valve. The pressure drop could be a driver in selec-

tion of the type of valve, particularly if system weight were critical. Thecurves of Figure ll3 show the effect of pressure drop on equivalent orifice

diameter for a valve flowing liquid oxygen and a valve flowing GH2 with flowconditions typical of those required for the main propellant shutoff valves.

As is shown, the change in orifice diameter is very significant below about 20psi. Since weight is a function of valve size, the final system pressureschedule could have a very definitive effect on the weight of the requiredcontrols. This also affects the type of valve since valves with differentshutoff elements have different size requirements to provide the same equiva-lent orifice flow.

Although the basic system schematics are thought to be practical as

depicted, there are questions that cannot be completely resolved by thelimited analysis performed to date. Several pf these questions are analyzedwith regard to potential problems and possible options to resolve the pro-blems.

Start Transient

With various sensing elements, signals and an electroniccontroller, a desired engine start should be attainable for a given setof conditions; however, there is some concern as to whether the same logic

can be applied for all conditions. The effects of variations such as fullvs empty lines, hot vs cold regenerative cooling section, single phase vstwo phase propellants and temperature soakback into valves and turbopumps

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are aspects requiring additional consideration. A morecomplexcontrol approachmaybe required to assure a smoothstart under the varying conditions that couldexist on restarts. Changesthat maybe required are useof moremodulatingor step position controls and start sequencevariations that would be selectedas a function of several monitored parameters.

Mixture Ratio Control

The concernabout mixture ratio (MR)control is relatedprimarily to the start and shutdowntransients. It is a concern becauseonlyone valve controls flow to all the modules. The needfor a flow balanceddis-tribution systemto the modulesis apparent.

Evenwith valves at each module, as is donefor theigniter circuits, MRexcursions during the transients could be rathersevere. TheMRrange would be influenced by propellant conditions, drivingpressure and the sizing of the igniter valves or flow orifices. It seemsreasonable to assumethat with a moredetailed analysis, this potential problemcould be accommodatedby proper orificing or a modulating control in one circuit.

ModuleInteraction Effects

With the modulesclustered around the nozzle, starttiming and interaction effects are a concern. The thrust generated by a modulewith just the igniter portion operating would be so low that no problemwouldresult from a start variation. As main modulethrust comesup throughidle mode,a variation from side to side could induce a turning momentto thevehicle. Anymomentscould be corrected by an attitude control systemor gimbal capability; however,here again the needfor a balanced flow anddistribution network is emphasizedto minimize the potential effect.

Another aspect of interaction relates to the commonmain control valve andmultiple feed lines. Any significant pressure pertur-bation in a modulechambercould reflect back into the feed system. Thefluctuations at one modulecould then effect other moduleswith varioustime lags. Dependentuponline lengths and propellant properties, anypressure disturbances maybe either amplified or attenuated. Thepotentialfor this effect could be reducedby makingthe systemstiffer, i.e., havinghigher injector pressuredrops, and controlling starts to limit Pc spikes.

Line Cooldown

Under the tank headand pumpedidle modestart the lineswill be chilled. Multi-position main shutoff valves and bypass bleed orificesare required for this operation.

Propellant Utilization

Propellant utilization in the GG cycle could be achievedquite readily by a special control signal to the throttle valve. On the expander

207

cycle, precise control could not be readily achieved with the valves depictedon the schematic. Somedegree of compensationin one direction only couldbe accomplishedby the turbine bypassvalve; however, the control and sensinglogic required is believed to be complex. A simpler approachmaybe to makeone of the main propellant shutoff valves capable of mudulation.

Tank Pressurization

Theschematics showsimple check valves to control thepropellant flow back to the tanks for autogenouspressurization. Thefeasibility of this approachis somewhatquestionable considering possi-ble pumpdischarge pressure variations, check valve crack and reseat

. accuracy, desired range of tank pressure and the early mission conditionswhere ullage will be small. An acceptable alternative would be to makethese valves a pressure differential sensing unbalancedpoppet arrangement.This approachcould be usedwith the valve size being comparableto a conven-tional spring loaded check valve.

Thrust Throttling

Although a throttling requirement is not currentlyimposed,a throttling capability could offer an attractive option to someother vehicle control requirements. The GGcycle could readily accommodatethrottling by makingthe GGvalves modulating rather than on-off. The ex-pander cycle would probably require the main shutoff valves to havea modulat-ing capability. This would imposea larger penalty than the GGcycle sincethe valves involved are muchlarger

Other options could be used for a steppedthrustcapability rather than true throttling over a specified range. In additionto control, as described aboveusing multiposition valves instead of fullmodulating valves, an approachof modulecontrol would be feasible. By addingmain propellant control valves to groups of modules, groups of 2, 3, or 4modulescould be shutoff or started to changethrust. A similar approachmight be used, with different modulegroupings, to provide maneuveringmomentswithout requiring engine gimbaling or use of auxiliary control thrusters.

Noneof these areas of concernappear to be overwhelming.However,rather extensive systemanalyses would be required to assure thatthe proper control parametersand control logic are used to provide the desiredperformancecharacteristics over the full range of operating and restart condi-tions.

I. Control System for Engine With Uncooled Plug

Minimum weight control system schematics were formulated for

both expander cycle and gas generator cycle engines with an uncooled plugnozzle (Figure 2) These are given in Figures ll4 and llS.

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Primary control of the expander cycle (Figure 114) is attained

by use of modulating or multi-position valves in the GH2 turbine drive circuit.Secondary control, for mixture ratio (MR) and propellant utilization (PU) is

achieved by a modulating valve in the oxidizer circuit downstream of the pump.

A preliminary definition of a start and shutdown sequenceof operations for this expander cycle is given in Table XXXVIII. The

sequence requires use of a controller havingcomputational and logic capabilities(i.e., microprocessor) rather than a controller having only timers and signalsequencing capabilities.

Primary control of the gas generator cycle (Figure ll5) is

achieved by modulating GG valves. Secondary control for MR and PU is obtainedby the throttle valve which controls hot gas flow to the oxidizer pump turbine.

A sequence of operations for start and shutdown of theGG cycle is shown in Table XXXIX. The comments relative to the requiredcontroller as discussed for the expander cycle also apply to the GG cycle.

Component weight estimates for the major control componentsare listed in Table XL. The total weight for the minimum valve expandercycle is 23.8 Kg (52.5 Ibs), while the corresponding weight for the GG cycleis 20.6 Kg (45.5 Ibs).

F. MODULE DESIGN

Four different modules are utilized in the conceptual designs: (1)Integrated Thrust Assembly (ITA) shown in Figure ll6 and described in SectionIII.D.I, (2) Minimum Modification ITA, and (3) Regeneratively Cooled ITA shown

in Figure ll7 for both a 40:I and a lO0:l module area ratio.

The minimum modification ITA utilized a regen cooled nozzle extensiondownstream of the regen-film cooled throat section. The fully regen module,

Figure ll7, requires no film cooling, and therefore, represents a majordeparture from the basic ITA design.

The ITA design has been shown to possess the capability of over1200 cycles operation at a mixture ratio of 5.5 (Reference 45) Analysis toestimate the life cycle capability of the regeneratively cooled module designis as follows:

Design criteria for the structures analysis are:

Coolant Channel Pressure = 38.1 and 66.7 atm (560 and 980 psia).Chamber Pressure = 20.4 and 34.0 atm (300 and 500 psia).

Coolant Channel Temperatures (given in Figures 87, 88, and 91)Design goal : 1200 thermal cycles for a lO hour duration.Safety Factor on Yield = l.l.Safety Factor on Ultimate = 1.4.

Chamber Material = Zirconium Copper.

211

TABLE XXXVlII - UNCOOLEDPLUG EXPANDERCYCLE OPERATIONS SEQUENCE

Command Element Response Element Action

Tank shutoff valvesStart signal

Differential pressureswitch

Ignition monitor

Controller

Temperature sensor

Spark exciter and oxid-izer igniter solenoidvalve

Controller

Spark exciter

Bypass valve

Oxidizer line Checkvalve

pressure

Pc transducer

ControlIer

Pc transducer

Controller

Bypass valve, thrustcontrol valve

Thrust control valve

Both shutoff valves open; fueland oxidizer start flowing.

Spark exciter energized; solenoidvalve opens; fuel and oxidizerflow in igniter is ignited.

Controller samples all modulesto confirm burning in each igniter.

Spark exciter is de-energized;igniter burn continues; GH_ flowsinto chamber thru main injector

and combusts; system cooldowncontinues.

Pump housing temperature sensorsreach the set temperature; control-

ler energizes bypass valve to movefrom full open to intermediateopen position; controller maintainslock-out on MR and thrust controlloops; turbine rotates; pumpsstart pumping fuel and oxidizer.

Pump discharge pressure increasesto open oxidizer line checkvalve;oxidizer flows into main chamber

and ignites with fuel.

Controller samples all chamberpressures and confirms all haveachieved pre-determined pressure.

Controller commands bypass valveto move to steady state positionand activates the thrust control

loop.

With thrust control loop activated,the low Pc signal causes the thrustcontrol valve to move toward the

closed position; valve closingforces rated flow thru the turbine;

Pc overshoot controlled by pre-

programmed valve travel rate.

212

TABLE XXXVIII (cont.)

Command Element Response Element

Pc transducer Controller

Controller MR valve, oxidizer

igniter valve

i

Action

Controller samples all thrustersto confirm full thrust.

Upon confirmation of proper Pc

the controller de-energizes theoxidizer igniter valves and

activates the MR control loop;steady state operation established;thrust controlled by Pc transducer

acting on the thrust control valve;MR controlled by PU tank signalacting on MR valve.

Shutdown involves simultaneous programmed functions, which are executed

by the controller. The shutdown signal results in deactivation of control loops,

bypass and thrust control valves open, tank shutoff valves close and the MR

valve goes to the nominal position. Upon confirmation of PC decay, the purge

valves are opened to clear oxidizer and fuel bleeds out the injector and base

bleed port.

213

TABLE XXXIX - UNCOOLED PLUG GAS GENERATOR CYCLE OPERATIONAL SEQUENCE

Command Element Response Element Action

Start signal

Differential

pressure switch

Ignition monitor

Controller

Temperaturesensor

Ox line

pressure

Pc transducer

Controller

Pc transducer

Pc transducer

Controller

Tank shutoff valves,throttle valve

Spark exciters, ox.

igniter valves

Controller

Spark exciter

GG valves

Ox checkvalves

Controller

GG valves

GG valves

Controller

Throttle valve, Ox

igniter valves

Shutoff valves open; throttle valve

goes to nominal position; fuel andoxidizer start flowing.

Spark exciter energized; solenoidvalve opens; fuel and oxidizer flow

and are ignited at thrusters and GG.

Controller samples all modules & GG

to confirm burning in each igniter.

Spark exciter is de-energized; igniter

burn continues; GH2 flows thru maininjector and combusts; system chilldowncontinues.

Pump housing temperature sensors reachthe set temperature; controller com-mands GG valves to about 50% openposition while maintaining control

loop lock-out; turbine and pumpsrotate; fuel and oxidizer pressure rise.

Pump discharge pressure increases toopen the oxidizer line checkvalves;oxidizer flows into the main chambers

and ignites with the fuel.

Controller samples all chamber pressuresand confirms all have achieved pre-determined pressure.

Controller removes thrust control loop

lock-out; MR control loop remains lockedout.

Low Pc signal causes GG valves to move

to the full open position; flow thruturbines goes to rated flow and thrustrises to full thrust level.

Controller samples all thrusters toconfirm full thrust.

Controller activates the MR control loopwhich lets the throttle valve respondto tank PU signals; oxidizer igniter

valves are de-energized; steady statethrust established and controlled byPc signals to GG valves.

Shutdown involves simultaneous pre-programmed functions which are

executed by the controller. The shutdown signal results in deactivation of

control loops, the throttle valve is commanded to a pre-determined position,

the GG valves are closed at a controlled rate and the tank shutoff valves

close in response to a timer signal. The controller samples Pc decay

and initiates an oxidizer purge to clear oxidizer lines. Fuel bleeds out

thru the main chamber and the hot gas out the base bleed and plug wall

ports.

214

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The estimate of thermal strain in the coolant channel was made usingthe equation:

_T = K.e.AT

where K = 2 is a factor based on detailed finite element analyses of similarstructures, _ = coefficient of expansion at the average wall temperature,

Tave, and AT = gas side temperature Twg 3 minus cold side temperature Tint.

The total strain in the coolant channel is given by the sum of thethermal and pressure (bending stress) strains. Since the pressure strainin this case is negligible, the allowable cycles can be read directly fromFigure ll8. The calculations for life determination are summarized inTable XLI. It is seen that in all cases the cycle life for the regeneratively

cooled module is greater than the required 1200 cycles.

TABLE XLI. REGEN COOLED MODULE LIFE CYCLE DETERMINATION

Chamber Channel Twg 3Pressure Depth

atm(psia) cm (in) °K (°F)

20.4(300) .38 (.15) 701 (802)', .64(.25)761(910)

20.4(300) .38 (.15) 678 (760)" .51 (.20) 713 (824)

AT

°K (°F)

536 (504)625 (665)

541 (S14)587 (596)

" .64 (.25) 739 (870) 628 (671)

34.0(500) .51 (.20) 794 (970) 661 (729)

Life

ET Flow C_xcles

.011 Coflow 3400

.014 " 2000

.010 Counterflow 3700

.012 " 2700

.014 " 1900

.015 " 1700

Although creep life determination was not included in this study, it

is apparent that there is adequate life for the low magnitude stress conditionsthat exist (cf. Reference 45).

G, UNCOOLED PLUG NOZZLE

Preliminary calculations were made for an uncooled plug nozzle con-figuration using graphite and carbon technology for materials of construction.AGCarb lOIK, a low modulus graphite composite which can be fabricated infree standing structures, was chosen as a typical candidate material. Ithas been used to launch communication satellites. (SVM-7 is the Aerojet

Solid Propulsion Company designation for Apogee Kick Motor used to orbit theRCA SATCOM, U. S. Domestic Communications Satellite.) This material is fullycharacterized and its properties are well understood. AGCarb 5451, anothercandidate, is made from a higher modulus, higher density version of AGCarb lOl

which provides improved erosion resistance.

218

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Demonstrated experience with AGCarb is noted in Table XLII. Summaryof selected AGCarb material properties is given in Table XLIII and a typicaldensity of graphite composite structures varies from 1.45 to 2.20 g/cc, thehigher value being for pyrolytic graphite.

Plug dimensions assumed for the nozzle calculations are 2.84 m (112 in)diameter by 2.16 m (85 in) diameter by 1,28 m (50 in) length with the geometrybeing approximated by a frustum of a cone. If the density of the AGCarb istaken as 1.45 g/cc, then the nozzle weight is given as 18 Kg (40 Ib) for a wallthickness 0.14 cm (0.055 in), which corresponds to the weight of 45 Kg (99 Ib)of the regeneratively cooled tubular structure described in Section VI.C. (andTable XLVI).

A structural analysis was performed to determine the required wallthickness for a plug nozzle with the pressure distribution shown in Figure 107.A wall thickness of 0.130 cm (0.051 in.) is acceptable for the plug from afabrication point of view and this thickness was selected for analysis. Buck-ling is the critical failure mode, and was, therefore, utilized for deter-mining the number and placement of required circumferential ring stiffeners.Four stiffeners are required with K-408 AGCarb and three with K-550D AGCarb.The circumferential ring stiffeners have square cross sections with dimen-sions of 2.54, 2.29, 2.03 and 1.91 cm (I, 0.9, 0.8 and 0.75 in), respectively.

A tapered panel section with the thickness varying from 1.02 to 0.25 cm(0.4 to 0.I in) was found to be structurally suitable for the nozzle fairing.Carbon-carbon bonding techniques ensure adequate bonding strength at thefairing-nozzle interface.

I. AGCarb Nozzle Cycle Life

The life of the carbon-carbon cloth (AGCarb) plug nozzlewas evaluated using an erosion rate expression of Heddon and Loewe (Reference43). It is a Hinshelwood-type equation and is based on experimental workwith a nuclear reactor grade graphite (density = 1.76 g/cc, surface area : 7.8 m2/g

at temperatures of 1213 - 1330°K (1724 - 1886°F) and H20 partial pressures of3.4 x 10-4 to 1.02 x 10-3 atm (0.005 to 0.015 psia). The equation is:

220

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TABLE XLIV - MATERIALS AND FABRICATION TECHNIQUE TRADE STUDY

Candidate

Materlal

AGCarb

Ablative

Bulk Graphite

Pyrolytic Graphite

(throat Insert only)

Reinforcement

Precursor "

Rayon, Continuous

Pan

Pitch

Fabric Weave

Plain (square)

Harness Satin

Selected ReJected Comments

Good thermal stability, low erosion, gas compatible, flight tested,

free standing.

Too heavy; duration limited

Thermal shock resistance not demonstrated.

Can be used if necessary wlth AGCarb chamber shell. PG washerpacks are used on MX and C-4 hlgh pressure solid rocket.

Selected for graphite yarn. Demonstrated on SVM6 and SVM7.

Higher cost, not demonstrated, lower interlamlnar shear, fabrication

loss greater.

New precursor, not demonstrated, feb. technicques not proven

low reliability.

Demonstrated, flexible fabric, intermediate strength, best

interlemlnar shear.

Tends to delaminate.

Fabrication - Reinforcement

2D

3D

Orientation

Rosette

Shingle

Tape Wrap

Angle Layup

Orthogonal

Cylindrical

Matrix

Resin Pitch

Low Pressure

High Pressure

Chemical Vapor

__epositio_ Carbon

CVD Resin Pitch

coatings

PG or SIC/PG

X

X

x

(Throat insert)

x

x

X

x

X

x

X

Demonstrated low cost fabrication techniques.

More costly, not demonstrated, primary 2D alternate.

Low axial compression and tensile, not demonstrated as free

standing, low cost fabrication.

Low cost, method to achieve high density.

Costly, structural advantages not needed, demonstrated on reentry systems.

Free standing, excellent mechanical properties; not demonstrated,costly, long process time.

Demonstrated, low cost, most fabrication experience.

Costly, high density not needed in chamber.

Not'demonstrated, costly, best 2D interlamlnar shear, lower fibercontent.

More costly than resin/pltch, not demonstrated in flight, someimprovement in shear over straight resin pitch, primary alternate.

Firing time too long for developed and demonstrated coating technology.

Multlple starts requirement not demonstrated.

223

where

rm= Kl CH20

l + K2 . CH2

Kl = 5 x lO12 e-68'O00/RT

K2 = 6.7 x lO-5 e14'500/RT

Mole)T:T--

cc/g-s

cc/Mole

CH20 = H20 concentration g/cc

CH2 = H2 concentration Mole/cc

T = temperature °K

R = gas constant 1.9872 cal/Mole - °K

This equation was checked with data from Lewis, Floyd and Cowlard

(Reference 44), who investigated various carbons (pyrolytic graphite, vitreouscarbon and erosion- resistant synthetic graphite) at pressures of l to 3

atmospheres and surface temperatures of 1500 to 3000°K. The erosion rate

calculated for plug cluster conditions (Pc = 20.4 atm [300 psia] and T =I067-1875°K [1460 - 2915°F] - see Table XXIX)

Area Ratio

on Plu9

Erosion Rate cm/lO hr (mil/lO hr)

S_nthetic Graphite Pyrolytic Graphite

40 0.16 (63) .005 (2)

458 3 x lO-5 (.010) 8 x lO-7 (0.0003)

Examination of the table indicates that pyrolyzed graphite nozzles

are capable of meeting the lO-hour life requirement with ease, while syntheticgraphite nozzles would erode somewhat at the module-plug interface (_M = _0).The AGCarb nozzle will exhibit properties between those for synthetic graphite

and pyrolytic graphite shown in the table. Should erosion be a problem at themodule interface, a coating of pyrolytic graphite or metal carbide could be

applied.

H. UNCOOLEDBELL NOZZLE EXTENSION

The successful application of AGCarb carbon-carbon cloth composite

materials for the plug nozzle structure prompted the investigation of thesematerials for uncooled nozzle extensions for the modules.

224

!!!

Thebasic moduleis regeneratively cooled to an area ratio of E = 40.TheAGCarbnozzle extension is attached at this point to extend the area ratioto E = 500. The surface temperatureat the attach point is 1867° (2900°F) Asindicated previously for the plug, the cycle life of the AGCarbis greaterthan lO hours for the environemntal conditions of this study.

Fabrication of the full or scarfed bell nozzle extension is withinthe state-of-the-art for this size nozzle (79 to I02 cm[31 to 40 in] exitdiameter). Assemblyof the cluster and installation of the base closure isalso readily accomplished. The scarfed bell nozzle assemblyforms a flutedplug with ideal aerodynamiccontour, as opposedto the assemblyof the samenumberof _ = 40 moduleson an annular plug.

I. WEIGHTANALYSIS

The baseline plug cluster weights for the expanderand gas generatorcycles were established by careful analysis of existing componentweights,scaling equations, and layout drawings (Figures I04 and I06). Revisions tothe componentweights were madeto incorporate materials and design changes.

I. Module Weight

The existing ITA module weight breakdown (cf. Tables V and XXXIV)

was examined and a 15 percent weight reduction was realized by assuming thata welded joint would replace line flanges. Elimination of the oxidizerflange (PNl162901-1), the fuel flange (PNl162901-2), the fuel inlet line

(PNl162906-1), and the oxidizer inlet line (PNI162885-I) resulted in a weightsavings of 0.52 kg (l.14 Ib) chargeable to the module. The heavy injector

head and flange were modified to reduce the weight by 0.23 kg (0.5 Ib), andsolid state circuitry was utilized to reduce the ignition system weight from0.99 kg (2.19 Ib) to 0.77 kg (I.69 Ib). This weight reduction is reflected

in the plug cluster engine baseline module weight given in Table XLV.

The nozzle extension weight for the regeneratively cooled modulewas estimated utilizing the design data from Ref. 35 (p. 705).

TABLE XLV. MODULE WEIGHT ANALYSIS

Module

Baseline BaselineITA,gjr_ ITA Module Regen Module

Component Kg (lb) Kg (lb) Kg (lb)

Igniter 0.99 (2.19) 0.77 (l.69) 0.77 (l.69)

Nozzle Extension 1.64 (3.61) 1.64 (3.61) 3.18 (7.00)Chamber 1.56 (3.43) 1.56 (3.43) 1.56 (3.43)Chamber Line/Torus/Flange 0.56 (I.24) 0.05 (O.lO) 0.05 (O.lO)Injector Assembly 1.88 (4.14) 1.65 (3.64) 1.65 (3.64)

6.63 (14.6]) 5.66 (12.47) 7.20 (15.86)

225

2. Plu9 Nozzle/Thrust Structure Weight

The weight of the plug nozzle plus fairings and thrust mountis given in Table XLVI.

A similar weight analysis was performed for the uncooled AGCarb

plug nozzle and associated thrust structure, The weights are also summarizedin Table XLVI.

Differences in the weights of the common con_)onents for theregen and uncooled plugs are found in the table. The major difference lies inthe thrust structure assumed for the uncooled plug which is 31.4 kg (16.4 +

15.0) comparied to 22.1 kg (I0.2 + ll.9) for the cooled plug nozzle. Thisdifference indicates the uncertainty in the selection of structure for the

two preliminary designs.

3. Module AGCarb Nozzle Extension and Base Closure Weight

The AGCarb nozzle extension (E = 40 to • = 500) weight wascalculated for modules operating at both 20.4 and 34.0 atm (300 and 500 psia)chamber pressures. Geometry data for the individual module and the cluster

configuration are given in Figures ll9 and 120.

The procedure for computing the weights was to: (1) calcu-late surface area (As), (2) assume _ wall thickness (t = 0.127 [0.050 in]) and

density (p = 1.45 g/cc [0.052 Ib/inJ]), and (3) calculate the weight from W = AStp. Tapered (0.127 to 0.064 cm thickness) nozzle weights were also calculated.

For the case of scarfed nozzles the surface area reduction

due to scarfing was assemed to be 40% the corresponding surface area reductionfor a 15° conical nozzle.

The resultant nozzle and base closure weights for the clusterengines formed by high area ratio bell nozzles are given in Table XLVII.

4. Turbopump Weight

The RLIO turbopump weight (35.9 Kg Or 79.1 Ib) from Reference

14 was utilized in determining the baseline weight of 31.8 Kg (70 Ib) for theplug cluster engine. A redesign of this pump according to 1977 state-of-the-art would show a marked reduction in weight.

A parallel turbine turbopump assembly based on current state-

of-the-art (Figures 78, 81 and 82) is expected to weigh only 21.3 kilograms(47 pounds).

226

I'|!

TABLEXLVI - PLUGNOZZLE/THRUSTSTRUCTUREWEIGHTANALYSIS

Con_)onent

Thrust Ring

Plug Wall

Base Closure

Struts/Plates

Fairings

RegenerativelyCooled

Plu9 Nozzle

kg (Ib)

10.2 (22.5)

44.9 (99.0)

4.6 (lO.l)

11.9(26.Z)

14.7(32.4)

Uncooled AGCarb

Plug Nozzle

kg (Ib)

16.4(36.2)

18.6(41.1)

5.6(12.3)

15.0(33.0)

16.1(35.5)

Total 86.3 (190.2) 71.7 (158.I)

227

i

(14 ")

366 cm

(176")

447 cm

I '

") 164 cm_ _ RT - (.9") 2.3 cm

__-_ DEXIT = (40.2") 102 cm-- _ OT : _ LCH_ER -_(17") 43 cm

_ 5.330 L TOTAL = (81.6")207cm

/Base

Db = (96.7") 246 cm

Ab = (7358.2 IN2) 47472 cm2

(E = 894.8

Pc = (300 psia) 20.4 atm /Pc35,500(MR: 5) /27,500 (MR = 6 /

P'E (.0085 PSIG)(MR = 5)/

5.78 x lO-4atm /

(.0109 PSIG) /

(MR = 6) /

7.42_

__/ (68.1") 173 cm(DISTANCE TO SINGLE BELL NOZZLE TYPEENGINE (RL-IO AR ASE) GIMBAL BLOCK)

_-- LT = (32.4")_D_ (22.6")-

82.3 cm 57.4 cm

k

DIA- (170.3")433 cm

(123.1")312.7 cm

(BASELINE SPACE TUG)

Figure 119. Plug Cluster Engine Geometry (Pc = 20.4)

Lv

228

Ill I

(176")

447 cm

F m

\NOZ (,, . ) .b c

Rt = (.7) (Pc = 5_) 1.78 cmDEXIT = (31.3") 79.5L' = (17") 43 cmLTOTAL = (67.22)") 170.7_90% BELL

Db = 75.2") 191 cm

Ab = (4446.1 IN2) 28684 cm2

(E = 894.8J

Pc = (500 PSIA) 34.0 atm /

_P-_c= 35,000 (MR= 5) I

l

rE 28,000 (MR = 6) /

(144") :'PE = (.0143 PSIA) /

366 cm (MR = 5) A /

9.73 x 10-_ atm //

(.0179 PSIA)

(MR = 6)1.218 x

10-3_///

jr _ z (68.1") 173 cm ----_---_(DISTANCE TO SINGLE BELL NOZZLE

TYPE ENGINE (RL-IO OR ASE)

_- LT = (37.1")_94.2 cm

jBase

DIA = (132.4"]

336 cm

D

(123.1") 312.7 cm

(BASELINE SPACE TUG)

Figure 120. Plug Cluster Engine Geometry (Pc = 34.0)

229

TABLE XLVII - AGCarb NOZZLE EXTENSION AND BASE CLOSUREWEIGHTFOR PCE FORMEDFROMBELL NOZZLES

COMPONENT

Nozzle Extension*(I0 Modules)E : 40 to E = 500

Pc = 20.4 atm

67.1 (148)

BASELINE WEIGHTKg (Ib)

Pc = 34.0 atm

40.9 (90.1)

Scarfed Nozzle **(_ 40 to _ = 500) 4o.o (88.1) 24.4 (53.9)

Base Closure 8.7 (lg.l) 5.3 (11.6)

Weight EffectivePlug (Unscarred)

75.8 (167.1) 46.1 (101.7)

Weight EffectivePlug (Scarfed)

48.6 (107.2) 29.7 (65.5)

*Tapered Nozzle 75% of Weight Shown**Tapered Nozzle 80% of Weight Shown

230

III i

5. Valve Weight

The baseline valve weight of 24.1 kilograms (53.2 pounds)

was established for the selected cycle (Figure Ill) and its required numberof valves by comparison with the weights of corresponding valves of the

candidate Tug engines given in Table XLVIII.

A revised valve weight statement was prepared (see Section VI.E.,Table XL) for the uncooled plug and clustered bell systems. The valve list-ing is included in Table XLIX.

6. Line Weight

The line weights were determined from the engine layout drawings(Figures I04 and I06). Minimum wall thicknesses calculated using a safetyfactor of 1.5 were a factor of two to ten lower than the wall thickness valuesutilized.

The total line weight for the baseline expander cycle engineis 16.15 kilograms (35.6 pounds).

Line weights were reevaluated for the uncooled plug and

clustered bell systems. The expander cycle and gas generator line weightsamounted to 12.7 and 12.3 Kg (28.0 and 27.2 Ib), respectively, showing a20% reduction in weight. The revised line weight breakdown is given inTable XLIX.

7. Weight Summary

The baseline plug cluster engine weight, corresponding tothe regeneratively cooled plug nozzle designs, in Figures I04 and I06 aresummarized by component in Table XLVIII. Controls, connecting and miscell-aneous hardware weight, consistent with that for the candidate Space Tugengines, are included in the table.

A similar weight breakdown is given in Table XLIX for theuncooled plug cluster engine, the clustered bell engine, the scarfed bell/fluted plug cluster engine at two chamber pressures, and the scarfed bell/fluted plug engine utilizing a gas generator cycle. Note that the GG

cycle reduces engine weight by about ll.3 Kg (25 Ib), and that operationat the higher chamber pressure (34.0 atm [500 psia]) reduces engine weightby 24.5 Kg (54 Ib).

The minimum plug cluster engine weight appears to be about181 Kg (400 Ib) for the higher pressure engine utilizing a GG cycle. Ingeneral, however, the plug cluster engines weigh more than the candidate

Space Tug engines listed in Table XLVIII. This might be expected due to thegeometrical configuration of the plug cluster. Every effort has been made

231

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233

TABLE XLIX - WEIGHT BREAKDOWNFOR LIGHTWEIGHT PLUG CLUSTER ENGINES

ENGINE:

Thrust (N)

Chamber Pressure (atm)

Mixture Ratio

Engine Area Ratio

Engine Diameter (cm)

EquivaTent Engfne Length (cm}

% Plug Nozzle

n

Isv (seconds)

Injector Assembly

Thruster Chamber & Primary Nozzle

Thrust Mount & Gimbal Assy.

Nozzle Extension

Plug Nozzle & Fairings

Base Closure

Turbopumps & Mounts

Gas Generator or Preburner

Ignition System

Lines: Total Weight

Ox Lines

Fuel Lines

Hot Gas L_nes

Line Supports

Valves: Total Weight

Oxidizer Inlet Shutoff Valve

Fuel Inlet Shutoff Valve

Oxidizer MR & PU Control Valve

Oxidizer Injector Check Valve (lO)

Turbine Bypass Valve

Thrust Control Valve

Tank Pressurizing Valves

Solenoid Valves (3)

Igniter Valves [Oxidizer)

Purge System Check Valves

GG Inlet Control Valve (Bipropellant)

GG Ox Throttle Valve

GG Igniter Valve (Oxidizer)

Controls, Connecting & Misc Hdwr

TOTAL ENGINE WIEGHT (Kg)

MODEL 11PLUG CLUSTER

UNCOOLED PLUG

EXPANDER CYCLE

68,950

20.4

5.5

458

320

85.g

15

0.914

443.8

16.5

47.8

31.4

O

34.7

5.6

31.8

0

7.7

12.7

4.9

5.6

0

2.3

MODEL IllCLUSTERED

BELL (CM=500)EXPANDER CYCLE

67,230

20,4

5.5

895

433

82.3

0

O, 946

463.9

16.5

• 47.8

16.4

67.1

0

8.7

31,8

0

7.7

12.7

4.9

5.6

O

2.3

MODEL IllPLUG CLUSTER

(SCARFED BELL)EXPANDER CYCLE

67,230

20.4

5.5

895

433

82.3

O. 946

463.9

16.5

47.8

16.4

40.0

0

8.7

31 .B

0

7.7

1.27

4.9

5.6

O

2.3

MODEL II]

PLUG CLUSTER

[SCARFED BELL)¢_ASGENERATOR CYCLE)

67.141

20.4

5.5

895

433

82.3

• 0.944

463.3

16.5

47.8

16,4

40.0

0

8.7

21.3

2.5

7.7

12.3

4.9

2.0

2.0

2.3

MODEL Ill

PLUG CLUSTER

(SCARFED BELL%[EXPANDER CYCLE)

67,230

34.0

5.5

895

336

g4.2

o.g50

465.9

12.8

43.0

16.4

24.4

O

5.3

24.2

3.4

7,7

II .l

4.4

2.8

1.9

2.0

28.3 28.3 28.3 25.2 37.8

3.g

3.9

3.6

4.1

2.7

2.1

0.5

3.4

3.6

0.7

B.8

225.3

3.9

3.9

3.6

4.1

2.7

2.1

0.5

3.4

3.6

0.7

3.g

3.9

4.1

0.5

3.4

3.6

0.7

2.8

2.2

0.2

8.8

207.2

3.9

3.9

3.6

4.1

2.7

2.1

0.5

3.4

3,6

0.7

8.8 B.8

218.7245,8

5.1

5.1

4.9

5.4

3.6

2,B

0.6

4.5

4.9

o.g

8.0

Ig4.1

234

I!! !I

TABLE XLIX (cont.)

ENGINES:

Thrust (lbf)

Chamber Pressure (psia)

Mixture Ratio

Engine Area Ratio

Engine Diameter (in.)

Equivalent Engine Length (in.)

Plug Nozzle

q

Isv (seconds)

Injector Assembly

Thruster Chamber & Primar_Nozzle

Thrust Mount & Gimbal Assy.

Nozzle Extension

Plug Nozzle & Fairings

Base Closure

Turbopumps & Mounts

Gas Generator or Preburner

Ignition System

Lines: Total Weight

MODEL II

PLUG CLUSTERUNCOOLED PLUG

EXPANDER CYCLE

15,500

300

5.5

458

125.9

33.B

15

D,914

443.8

36.4

I05.4

69.2

0

76.6

12.3

70.

O

16.9

28.0

MODEL Ill

CLUSTERED

BELL (_M=500)EXPANDER CYCLE

15.114

300

5.5

895

170.3

32.4

0

O. 946

463.9

36.4

105.4

36.2

148.

0

19.1

70.

0

16.9

28.0

MODEL Ill

PLUG CLUSTER

(SCARFED BELL)EXPANDER CYCLE

15,114

300

5.5

895

170.3

32.4

0.946

463.9

36.4

105.4

36.2

88. l

0

19.1

70.

0

16.9

28.0

MODEL Ill

PLUG CLUSTER

(SCARFED BELL)GASGENERATORCYCLE

15,094

300

5.5

895

170.3

32.4

0,944

463.3

36.4

105..4

36.2

BB. 1

0

19.I

47.

5.6

16.9

MODEL III

PLUG CLUSTER

(SCARFED BELL)

{EXPANDER CYCLE)

15,114

500

5.5

8gS

132.4

37.1

0.g50

465.9

28.2

94.8

36.2

i3.9

0

I1.6

53,4

7.5

16.9

27.2 24.6

Ox Lines

Fuel Lines

Hot Gas Lines

Line Supports

Valves: Total Weight-

Oxidizer Inlet Shutoff Valve

Fuel Inlet Shutoff Valve

Oxidizer Injector Check Valve (10)

Turbine Bypass Valve

Thrust Control Valve

Tank Pressurizing Valves

Solenoid Valves (3)

Igniter Valves (Oxidizer)

Purge System Check Valves

GG Inlet Control Valve (Bipropellant)

GG Ox Throttle Valve

GG Igniter Valves (Oxidizer)

Controls, Connecting & Misc Hdwr

TOTAL ENGINE WEIGHT (lbm)

10.7

12.3

0

5.

10.7

12,3

0

5.

10.7

12.3

0

5.

62.5 62.5 62.5

8.5

8.5

9.o

5.9

1.6

1.0

7,5

8.0

1.5

19.5

8.5

8.5

9.0

5.9

4.6

1.0

7.5

8.0

1.5

496.8

8.5

8.5

9.0

5.9

4.6

1.0

7.5

8.0

1.5

19.5 19.5

542. 482. l

10.8

6.9

4.5

5.0

9.8

6,2

4.1

4.5

55.5 23.3

8.5

8.5

g.o

1.0

7.5

8.0

1.5

6.1

4.9

0.5

11.3

11.3

12.0

7.9

6.1

1,3

10.0

10.7

2.0

19.5 17.6

456. g 428.

235

in this study to use similar engine state-of-the-art technology (turbopumps,injectors, combustion chamber, etc.) to provide an equivalent comparison of

engines. Advantage was taken of the unique configuration of the plug clusterto evaluate the effect of lightweight uncooled nozzles.

236

SECTION VII

PLUG CLUSTER ENGINE OPTIMIZATION


Parametric system analyses were conducted to optimize the plugcluster engine concept for a Space Tug round trip to geosynchronous orbitmission. The engine design point for the optimization was the baselinedesign established in the preliminary design effort. It is consistentwith the guidelines listed in Table I.

The payload capability is included using the exchange factors avail-able from Space Tug system studies. Subsystem limitations imposed on theengine design or operation were evaluated.

Points of the study where additional technology will improve thefeasibility of the plug cluster concept as a building block approachfor future applications to advanced space vehicles were summarized.

B. ENGINE DESIGN SPECIFICATION

Specifications for the conventional engine design configurations aregiven in Appendix A Tables LXIV through LXX. The specifications are forengines utilizing expander and gas generator cycles, regeneratively-cooledand film-cooled modules, module area ratios of 40, and engine operatingpressures of 20.4 and 34 atm. Geometric and performance data given in thetables were derived from the performance Model I presented in Section IV,and the performance was revised to reflect the Model I! results.

In addition to these specifications, similar data are given in theAppendix Tables LXXI through LXXIV for the uncooled plug configurations.

Specifications for the recommended (optimized) engine configurations,the plug cluster/scarfed bell engines, are given in Tables L through LIII.Performance model III was utilized to generate these data.

Some of the effects that can be noticed by examination of the tablesare: (I) an increase in chamber pressure leads to an increase in engineperformance, and a decrease in engine diameter; (2) the gas generator cycleshows a small (about 0.1%) decrease in engine performance at Pc of 20.4 atm,a lower pump discharge pressure and a higher turbine operating temperaturethan a corresponding expander cycle; (3) the fuel film cooled ITA moduleleads to a significant decrease in specific impulse compared to a corres-ponding regeneratively cooled module; (4) the conventional plug clusterengine design (Tables LXIV through LXX) does not realize the high area ratioperformance potential; (5) the plug cluster/scarfed bell, Tables L throughLIII (PCE) engine design achieves the high area ratio performance potentialthrough optimization of the aerodynamic flow contour of the plug nozzle.

237

TABLE L -PLUG CLUSTER/SCARFED BELL ENGINE OPERATING SPECIFICATIONEXPANDER CYCLE: REGEN-MODULE

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow RateEngine Area Ratio (AE/AT)Module Area Ratio (Ae/At)Number of Modu}esModule Gap (a/D e)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Eixt DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket _PCoolant Inlet TemperatureCoolant Exit Temperature

MODEL IIIP = 20.4

C

SI UNITS

71.97 kN20.41 atm

463.9 s5.44

15.82 kg/s13.36 kg/s

2.46 kg/s8955OO

700

433 cm246 cm

82.3 cm8.59 cm4.72 cm

102 cm16.51 cm164 cm207 cm

2.46 kg/s2.04 atm

22 K246 K

ALTERNATE UNITS

16,180 Ibf300 psia

34.88 Ibm/s29.46 Ibm/s

5.42 Ibm/s

170 in96.7 in32.4 in3.38 in1.86 in

40.2 in6.5 in

64.6 in81.6 in

5.42 Ibm/s30 psia40 R

442 R

(5.50 TCA)(34.82 TCA)

TURBINES

Inlet PressureInlet TemperatureGas Flow Rate

Specific Heat RatioMolecular Weight

Shaft HorsepowerPercent Bypass

32.7 atm

246 K

1.38 kg/s1.40

2.02 g/mol290 kW44

480 psia442

3.06 Ibm/s

390 hp

MAIN PUMPS

Oxidizer Pump Flow Rate

Oxidizer Pump Discharge PressureFuel Pump Flow Rate

Fuel Pump Discharge Pressure

13.36 kg/s27.2 atm

2.46 kg/s36.7 atm

29.46 Ibm/s

400 psia5.42 Ibm/s

540 psia

238

ItI I I

TABLE LI - PLUG CLUSTER/SCARFED BELL ENGINE OPERATING SPECIFICATIONSEXPANDERCYCLE: REGEN-MODULE

MODEL IIIP = 34.0

C

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow RateEngine Area Ratio !AE/AT)Module Area Ratio (Ae/At)Number of ModulesModule Gap (a/De)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Eixt DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket APCoolant Inlet TemperatureCoolant Exit Temperature

SI UNITS ALTERNATE UNITS

71.63 kN 16,100 Ibf34.02 atm 500 psia

465.9 s5.44

15.68 kg/s 34.56 Ibm/s13.24 kg/s 29.19 Ibm/s

2.44 kg/s 5.37 Ibm/s8945OOI0

0336 cm 132.4 in191 cm 75.2 in

94.2 cm 37.1 in6.63 cm 2.61 in3.66 cm 1.44 in

79.5 cm 31.3 in16.51 cm 6.5 in

127.6 cm 50.2 in171 cm 67.2 in

2.44 kg/s 5.37 Ibm/s5.58 atm 82 psia

23 K 42 R218 K 392 R

(5.50 TCA)(34.50 TCA)

TURBINES

Inlet PressureInlet TemperatureGas Flow RateSpecific Heat RatioMolecular WeightShaft HorsepowerPercent Bypass

55.0 atm 808 psia218 K 392 R

2.23 Kg/s 4.92 Ibm/s1.402.02 g/mol

506 kW 679 hp8

MAIN PUMPS

Oxidizer Pump FlowRateOxidizer Pump Discharge PressureFuel Pump Discharge Pressure

13.24 kg/s 29.19 Ibm/s43.5 atm 640 psia66.8 atm 982 psia

239

TABLELII - PLUGCLUSTER/SCARFEDBELLENGINEOPERATINGSPECIFICATION

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow Rate

Engine Area Ratio (AE/AT)Module Area Ratio (Ae/At)Number of Modules

Module Gap (a/De)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat Diameter

Module Exit Diameter

Module Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow Rate

Coolant Jacket AP

Coolant Inlet Temperature

Coolant Exit Temperature

GAS GENERATOR CYCLE: REGEN-MODULE

MODEL IllP = 20.4c


72.54 16,310 Ibf

20.41 atm 300 psia463.3

5.3315.97 kg/s 34.20 lbm/s13.44 kg/s 29.64 lbm/s2.52 kg/s 5.56 lbm/s

895500lO0

433 cm 170 in246 cm 96.782.3 cm 32.48.59 cm 3.38 in4.72 cm 1.86 in

I02 cm 40.2

16.51 cm 6.5 in164 cm 64.6

207 cm 81.6

2.43 kg/s 5.36 Ibm/s2.04 atm 30 psia22 K 40 R

246 K 442 R

(5.50 TCA)(34.82 TCA)(29.46 TCA)(5.36 TCA)

TURBINES


6.53 atm 96 psia922 K 1,660 R

0.17 kg/s 0.38 ibm/s1.36

3.8 g/mol189 kW 254 hp

O

MAIN PUMPS

Oxidizer Pump Flow RateOxidizer Pump Discharge PressureFuel Pump Flow Rate

Fuel Pump Discharge Pressure

13.44 kg/s 29.64 Ibm/s27.22 atm 400 psia2.52 kg/s 5.56 Ibm/s

27.2 atm 400 psia

GAS GENERATOR

Chamber Pressure

Combustion TemperatureMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow Rate

6.80 atm lO0 psia922 K 1,660 R0.9

0.17 kg/s 0.38 Ibm/s0.08 kg/s 0.18 lbm/s0.09 kg/s 0.20 Ibm/s

240

Ill I _

TABLELIII - PLUGCLUSTER/SCARFEDBELLENGINEOPERATINGSPECIFICATIONGASGENERATORCYCLE:REGEN-MODULE

MODELIIIP = 34.0

C

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RaCeOxidizer Flow RateFuel Flow RateEngine Area Ratio (AE/AT)Module Area Ratio (Ae/A t)Number of ModulesModule Gap (a/De)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Exit DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket APCoolant Inlet TemperatureCoolant Exit Temperature


72.54 kN 16,310 Ibf34.02 atm 500 psia

465.05.25 (5.5O)

15.91 kg/s 35.07 Ibm/s (34.50 TCA)13.36 kg/s 29.46 Ibm/s (29.19 TCA)

2.54 kg/s 5.61 Ibm/s (5.31 TCA)8945O0I0

0336 cm 132.4 in191 cm 75.2 in94.2 cm 37.'I in

6.63 cm 2.61 in3.66 cm 1.44 in

79.5 cm 31.3 in16.51 6.5 in

127.6 cm 50.2 in171 cm 67.2 in

2.41 kg/s 5.31 Ibm/s5.58 atm 82 psia

23 K 42 R218 K 392 R

TURBINES


6.46 atm 95 psia922 K 1,660 K

0.26 kg/s 0.57 Ibm/s1.363.8 g/mol

316 kW 424 hp0

MAIN PUMPS

Oxidizer Pump Flow RateOxidizer Pump Discharge PressureFuel Pump Flow RateFuel Pump Discharge Pressure

13.36 kg/s 29.46 I bm/s43.55 atm 640 psia

2.54 kg/s 5.61 Ibm/s47.6 atm 700 psia

GAS GENERATOR

Chamber PressureCombustion TemperatureMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow Rate

6.80 atm I00 psia922 K 1,660 R

0.90.26 kg/s 0.57 Ibm/s0.12 kg/s 0.27 Ibm/s0.14 kg/s 0.30 Ibm/s

241

The engine length given in the tables is equivalent to the engine

length reported for the baseline Space Tug candidate engines, which ismeasured from the gimbal point at the aft end of the LOX tank. Sincethe modules of the plug cluster engine are clustered around the LOX tankforward of its aft end (Figures If9 and 120), the actual engine length has

no constant reference point, but varies with the engine area ratio, modulearea ratio and chamber pressure. In order that a direct comparison could bemade between the various types of propulsion systems, the engine length wasdetermined from the centerline of the LOX tank, and its equivalent length

from the gimbal point of the baseline Tug was determined.

The longer engines shown for the higher pressure (34 atm) systems are

the result of selecting the higher performing (20% LI) plug. At an equalpercent plug (15% LI), the higher pressure engine is shorter.

C. ROUND TRIP GEOSYNCHRONOUS ORBIT MISSION

The baseline Space Tug round trip payload (WPL) to geosynchronousorbit is given in Reference 36 as 939 Kg (2070 Ib), with a velocity incre-

ment (Av) budget of 8,680 m/s (28,478 ft/sec). The useable main enginepropellants amount to 22,629 Kg (49,889 Ib), and the burnout weight (WBo)

is 2617 Kg (5770 Ib) when the AP_ propellant is ignored, The voIume of

the LH2 and LOX tanks is 52.39 mj (1850 ft_) and 18.12m° (640 fts), respectively.For engine mixture ratios other than 6, propellant off-loading must take place

as given in Table LIV.

TABLE LIV. PROPELLANTS AVAILABLE FOR ROUND TRIP MISSION TO GEOSYNCHRONOUSORBIT

PropellantMixture LH2 L02 OffloadedRatio

Kg :_-i-b) K9 _Tb) Kg (Ib)4.0 3,233 (7,127) 12,931 (28,508) 6,465 (14,254) LO25.0 3,233 (7,127) 16,164 (35,635) 3,232 (7,127) LO2

5.5 3,233 (7,127) 17,780 (39,199) 1,616 (3,563) LO2

6.0 3,233 (7,127) 19,396 (42,762) 0 (0)

7.0 2,771 (6,109) 19,396 (42,762) 462 (1,018) LH2

Solution of Equation 24 gives WI, the ignition weight (24,720 Kg

or 54,499 Ibm), where

WIAv = g Is In + (Eq. 24)

Wpl_ WBO

?¢?

g is the constant (9.807 m/s 2 or 32.2 ft/sec2), Is is the RLIO specific

impulse (456.5 sec), and Av, WpL and WRn are as given above. The equationcan be rearranged and then solved for _e payload capability of other engines

when the off-loaded propellant (Wpm) is accounted for and the appropriatespecific impulse and engine weightV_WE ) are utilized.

WpL = Wl - WpOL - (WBo - WE [RLIO] + WE) (Eq. 25)

eav/gls

The RLIO engine weight (WE [RLIO]) shown in Eq. 25 is 201 Kg (443 Ibm).

The round trip payload to geosynchronous orbit for the optimized plug

cluster engine (_M = 500) is given in Table LV. It is seen that off-loadingpropellant from the baseline mixture ratio (MR=6) Space Tug design pointreduces the capability of the plug cluster engine. Therefore, the maximum

payload is achieved at an MR of 6.

TABLE LV. ROUND TRIP PLUG CLUSTER (PCE) PAYLOADS TO GEOSYNCHRONOUS ORBIT

PC MR Is WE WpL _WpL/_Is

atm (psia) -- Kg-l-(-_-.m-T- Kg (Ibm) Kg/s {'Ibm/sec)

20.4 (300) 5 463.4 219 (482) 547 (1205) 13 (29)

" 5.5 463.9 219 (482) 793 (1749) 14 (31)" 6 464.4 219 (482) I041 (2294) 15 (33)

34.0 (500) 5 465.2 194 (428) 595 (1311) 13 (29),, 5.5 465.9 194 (428) 846 (1865) 14 (31),' 6 466.6 194 (428) I098 (2421) 15 (33)

D. TECHNOLOGY REQUIREMENTS

The plug cluster concept appears to offer a unique building blockapproach for advanced space vehicles. The status of technology to developsuch an engine is very favorable. The ITA-type modules have been demonstratedto deliver long life (greater than 1200 cycles at a mixture ratio of 5.5). An

existing RLIO turbopump could be used with a 5-hour life, or developed turbo-pump technology could be applied to a new design. Existing AGCarb carbon-

carbon cloth nozzle technology is available for the high area ratio bellnozzle extensions. There is an inherent low cost associated with the

utilization of off-the-shelf technology in the development of a plug clusterengine.

The feasibility of the plug cluster concept is based upon certainassumptions and preliminary conceptual designs. Points of the study where

additional technology will improve the feasibility of the concept aresummarized in Table LVI. Thrust vector control considerations are listed

in greater detail in Table LVII.

243

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I!!i!

E, OPTIMUM PLUG CLUSTER ENGINE

Two plug cluster/scarfed bell engines were selected for comparisonwith the Space Tug candidate engines. Both engines utilized regeneratively

cooled modules (EM : 500) and the expander cycle. The plug cluster engineoperating at a chamber pressure of 20.4 atm (300 psia) is designated PCE300 and its counterpart at higher chamber pressure is PCE 500 and is describedin Table XLIX.

247

II| I _

SECTIONVlll

PLUG CLUSTER ENGINE ASSESSMENT


The results of the plug cluster engine study were compared with resultsfrom previous Space Tug and Orbit-to-Orbit engine studies involving highpressure engines using bell nozzles and engines using annular plug (Aero-spike) nozzles. The comparison was directed toward making a common groundof reference between the studies with regard to the assumptions and complete-ness. An assessment of the plug cluster concept was made as a result ofthese comparisons.

Specific missions that were considered for the comparison are:

(I) Round trip to geosynchronous orbit.(2) Placement of payload into geosynchronous orbit.(3) Retrieval of payload from geosynchronous orbit.(4) Placement of payload into planetary or escape trajectory.

B. MISSION EXCHANGEFACTORS

Payloads and payload sensitivities for Space Tug missions aregiven in References I, 2 and 37. These data, however, were derived fromvehicle designs during various stages Of the Space Tug studies, and there-fore, are not entirely consistent. For example, studies to determine theoptimum mixture ratio for the various engine candidates included vehicleredesign to accommodate the different propellant tank volumes required.

In order to p_ovide a common ground of reference for the enginecomparison, the ideal velocity (Av) budget (Reference 36) for each missionwas utilized, and the payload calculated for the baseline Space Tug in themanner previously described for the round trip mission (cf. Section VlI.C.)The mission data are summarized in Table LVIII.

To simplify the analysis, the APS (auxiliary propulsion system)contribution to the ideal velocity was ignored, and ignition weightswere computed using the appropriate form of Equation 24 (Section VlI.C.)The resultant data are given in Table LIX. Equations were developed foreach mission as shown in Table LX. The nomenclature for the equations isgiven in Section VII.C and the baseline Space Tug data in that section andin Tables LVIII, LIX and LX.

Results of the calculations are given in Appendix A (Table LXXV) formodel I plug cluster engines compared to early estimates for the candidateSpace Tug engines. Mission exchange factors (_PL/_Is and _PL/_WE) are alsoincluded in the table.

249

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250

Iii I

Since the compilation of the data in Appendix A (Table LXXV), additional datahave become available on each of the listed engines: (1) further developmentof the ASE has led to more realistic weight estimates for the engine, (2)high area ratio tests have been conducted using the RLIO, (3) test data have

shown that the conventional configuration for the plug cluster engine (modelI) does not achieve the high performance predicted, and (4) a redesign ofthe plug cluster concept indicates that the high area ratio performance

potential of this system can be achieved.

An attempt was made to compare the Model III candidate engines using

the latest empirical data, and also to compare them on the same analyticalbasis (i.e., JANNAF Simplified Methodology). A summary of the enginecomparison using the JANNAF simplified methodology for each engine is given

in Table LXI, and the overall summary depicted in Figure 121.

The results of this comparison show that the plug cluster engine conceptderived from a cluster of scarfed bell nozzles offers a competitive payloadwhen compared to the previously studied Space Tug engines. By adoptinga zero gap configuration and by utilizing the available vehicle diameter,

the tradeoff in engine weight with performance becomes favorable, as shownin Figure 122.

TABLE LX. PAYLOAD EQUATIONS FOR ENGINE COMPARISON

Mission

WI _ WpoLm

Round Trip WpL eAv/gls (WBo + AWE)

WI _ WpOL AVin/gls

Deploy WpL eAVout/gls e (WBo + AWE)

(Eq. 25)

(Eq. 26)

Av n/glse i WI - WpoL

(WBo+AWE) - AVout/gls

Retrieve WpL = AVin/gls e (Eq. 27)l e

WI - WpoL AVin/gls

Interplanetary WpL - AVout/gls - WKS - e (WBo+AWE) (Eq. 28)e

Where: AWE = WE-WE(RLIO), WKS = Kick Stage Weight + etc. (3984 Kg or

8783 Ibm)

251

n "_ •

nin

1

n,x

v

7"

O

O

t.u

ILl

I

X

,._1

i,n

._1

I"-'

Z

I---

I'--

Z

1,--.4

OZ

ii

N

i

"" D

logo

w

j_

z

o

252

I!! I_

r---

0

ot'-

0

I--

_..I e,,.-

"1-

.../

, _--

0

t_ _ _ o

ii

I

g

d

oo0o

o

g

g

O4

0

"W

253

S.l. UNITSRLIO lIB PCE 300 PC[ 5_G

MR 6 6 6

PC (atm) ?7.2 20,4 34,0

_E 205 895 894

W E (Kg) 201 219 194

D E (m) l.RO 4.32 3.35

L E (m) 1,40 *** O.RI 0,94

*I s Is) 460,6 464.4 466._

r E {Is/I s ODE) 0.q65 0.945 0.949

PAYLOAD _KJl)

Deploy 3740 3_27 3q64

Retrieve 1649 1721 IR20

Round Trip 10F)l I041 109R

Planetary 4_45 5019 517q

4,500

4,000

RLIO lIB PCE 300 PCE 500


** Thrust/Welght ratio assumed same as for _G,7_3 tl enqine {Rrf 27)

***SLowed length of deployable nozzle

ASE

6

1 36

400

183"*

1.07

1.28"**

469.3

0.966

4084

1911

I150

5224

ASE

ENGLISH UNITS

RLIO lIB PC[ 300 PCE 500

)4R 6 6 6

Pc (psla) 400 300 500

{E 205 895 894

WE {lbm) 443 482 428

DE (in.) 71 170 13?

LE (in.) 55*** 32 37

*I s Is) 460.6 464.4 466.6

nE (Is/I s ODE) 0.965 0.945 0.949

PAYLOAD {I_)

Deploy $1245 8436 R740

Retrieve 3135 3794 401 3

Round Trip 2207 2294 7421

Planetary I0901 11065 II 307

9"000 F8,000|

RLIO lIB PCE 300 PCE 500

ASE

6

2000

400

404**

42

44e**

469.3

0.966

9003

4212

2535

11518

ASE


** Thrust/We|ght ratio assumed same as for 20,000 Ibf engine (Ref. 27)

***Stowed length of deployable nozzle

254

Figure 121. Space Tug Engine Evaluation

II

0

III

w

I i I , I , I i0 0 0 0

C_

0

0

OD

I

"l-

ILl

t._Z

ZLIJ

rO

or'-

t-

OJt""

=r-

ILl

U1.

r--

r--

LL

5(]N033S - 3S7ndWl _I_I33dS 03_3A173{3 3NIgN3

255

C. COST ANALYSIS

Engine development cost has been a major selection criteria indiscussions concerning candidate Space Tug and Orbit Transfer Vehicles (OTV).In fact, the RLIO uprated engines were reconTnended by both MDAC and Convair(References 2 and 37) in their Space Tug studies primarily on the basis ofDDT&E.

Cost data for the RLIO versions, the Aerospike, and the AdvancedSpace Engine are given in Table LXlI. The estimates were taken from theSpace Tug studies (References 2 and 37). The cost data for the plugcluster engines given in the table were estimated in several ways summarizedin the following. The initial cost estimate of $0.4M to $0.7M for a plugcluster engine was made based upon conceptual design layouts. The DDT&Ecost estimate for a 5-year development program was obtained by plotting theDDT&E costs shown in the table versus chamber pressure. While this plotshowed some scatter due to the widely divergent engine designs, it didreflect a trend in development cost with chamber pressure, giving somecredence to the selected PCE cost.

TABLE LXII. SPACE TUG ENGINE COST COMPARISON

................. (1973 Dollars) .............

DDT&E _ Engine Maintenance_ $M/Year

RLIO IIA 13 0.7 0.22RLIO liB 50 0.8 0.22RLIO IV 119 0.9 0.23A/S 140 I.I 0.17ASE 154 1.0 0.15PCE300 52 0.4 0.16PCE500 60 0.4 0.16

A determination was made of the number of equivalent engines re-quired during the development program. Comparison with the Space TugStorable Engine Study (Reference 16) showed that from 17 to 22 equivalentengines were required, and that the DDT&E cost was between $41.5M to $70M,depending upon the cycle chosen. Based on the previously cited costs perPCE300 fabrication, the manufacturing (plus procuren_nt)

Cost of the DDT&E effort is between $6.8M and $15.4M or from 13 to 30percent of the total cost. Comparison of these percentages with the44 percent obtained from the on going OME program indicates that the manu-facturing/procurement cost is low for this size program. A single enginecost of $1M would bring the plug cluster development cost percentage inline with that for OME. This higher figure, however, does not seem reason-able when the module high technology status is considered.

The DDT&E cost of $52M to $60M is, therefore, seen to represent areasonable value when compared with values generated for the other OTVcandidate engines. The figure also appears to be consistent with previousALRC estimates for the development of similar space engines.

256

!_T

D. LIFE ANALYSIS

Typical engine life projections given for the Space Tug candidate

engines are listed in Table LXIII. All of the engines except the RLIO lIBhave identical life requirements if it is assumed that the Aerospike and theASE cycle life is 300 times a safety factor of four (1200 cycles). The

Aerospike, however, is also projected to have some scheduled maintenanceand refurbishment after 60 cycles or 2 hours of operation, with a total

engine service life of 50 hours or 1500 cycles (Reference 6).

TABLE LXIII. SPACE TUG ENGINE LIFE COMPARISON

Baseline Tug Aerospike ASELife RLIO lIB (Ref. l) (Ref. 6) (Ref. 39) PCE300/500

Hours 5 lO lO 5"-I0

Cycles 190 300 300 1200

*RLIO Turbopump

Critical components that dictate the minimum life between over-

hauls are the injector, thrust chamber, bearings, seals, and the igniter.Initial analysis of the low pressure plug cluster engine components indi-cate that lifetimes greater than those listed in Table LXIII are state-of-the-art. The plug cluster engine, therefore, should surpass the life

capability of the high pressure engines and will be superior to the RLIO

lIB, which utilizes fifteen-year-old technology.

257

Ill I :+

SECTIONIX

CONCLUSIONS

Themajor conclusions (Figure 123) resulting from the UnconventionalNozzle Tradeoff Study are:

PLUG CLUSTER FEATURES

° COMPETITIVE PAYLOAD

° INCREASED PAYLOAD LENGTH

° DESIGN FLEXIBILITY

o LONG LIFE

° EXlSTING TURBOMACHINERY

° DEMONSTRATEDMODULES

o EXISTING NOZZLE TECHNOLOGY

o LOW COST

o LOW PC ENGINE SYSTEM

The major feature of the PCE is that it is capable of delivering acompetitive payload. While verification of the performance is needed, theperformance methodology developed in this program indicates only a smalluncertainty.

Another feature of the PCE is the allowance of increased payload lengthdue to the shorter engine length.

The PCE offers considerable design flexibility, since the capabilityto increase or decrease the number of modules (and thrust) is inherent in thecluster concept. Fail operation features can be provided by the cluster con-figuration that are not possible with single engine configurations.

Long life has been demonstrated for the ITA modules. While lifeverification of the fully regeneratively cooled module is required, sufficientdata have been accumulated to indicate the soundness of the approach.

Another feature of the PCE is that an existing turbopump assemblycould be utilized. Likewise, well developed turbopump technology could beapplied.

Existing AGCarb carbon-carbon cloth nozzle technology is available.Verification of the greater than l-hour predicted life for this nozzle isneeded.

259

260

i,

e-

¢0

0

4_

I.)e-

O

e-

• r"-

e-"L.ul

4-)

-r=

An important feature of the PCE is the inherent low cost associatedwith utilizing off-the-shelf technology. Low cost is also inherent in theoperation of low pressure systems which comprise the PCE. While a cost

analysis should be conducted to verify this favorable feature of the PCE,there is little uncertainty involved in predicting low cost operation for sucha Space Tug system.

This study indicates that the performance of the PCE is competitiveto other Space Tug candidate engines.

261

_i_I_

I _ _,

r _

SECTION X

APPENDIXES

A, CONVENTIONALENGINE OPERATING SPECIFICATION

These Engine Operating Specifications are Shown on Tables LXlV ThroughLXXV

263

TABLE LXIV - PLUG CLUSTER ENGINE OPERATING SPECIFICATIONEXPANDERCYCLE: REGEN-MODULE

MODEL II PERFORMANCEP = 20.4

C

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow RateEngine Area Ratio (AE/A T)Module Area Ratio (Ae/A t)Number of ModulesModule Gap (_/D e)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Exit DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket _PCoolant Inlet TemperatureCoolant Exit Temperature

SI UNITS

68.72 kN20.41 atm

443.85.44

15.82 kg/s13.36 kg/s2.46 kg/s

4584OI0

2319.8 cm217.9 cm85 9 cm8 59 cm4 72 cm

29 87 cm16 51 cm35 46 cm60 33 cm2 46 kg/s4.42 arm

22 K376 K

ALTERNATE UNITS

15,451 Ibf300 psia

(5.50 TCA)34.88 Ibm/s (34.82 TCA)29.46 Ibm/s5.42 Ibm/s

125.9 in85.8 in33,8 in3.38 in1 86 in

II 76 in6 5 in

13 96 in23 75 in

5 42 I bm/s65 psia40 R

677 R

TURBINES


31.6 arm376 K

l .05 kg/sl .402.02 g/tool

290 kW57

464 psi a677 R

2.32 Ibm/s

390 hp

MAIN PUMPS

Oxidizer Pump Flow Rate 13.36 kg/sOxidizer Pump Discharge Pressure 27.2 atmFuel Pump Flow Rate 2.46 kg/sFuel Pump Discharge Pressure 36.7 atm

29,46 Ibm/s400 psia

5.42 Ibm/s540 psia

264

TABLELXV- PLUGCLUSTERENGINEOPERATINGSPECIFICATIONEXPANDERCYCLE:REGEN-MODULE

MODELII PERFORMANCEP = 34.0

C

PARAI,IEI[P,

ENGINE

Vacuum, ThrustChar,her PressureVacu_;m Specific ]m!,ulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow RateEnglne Area Ratio (AE/A T)Iff;ch!le Are_, Ratio (Ae/i;L)Nu_her of _,k_dulesModulo Gop ((./De)Engine DiameterPlug Base DiameterEn3ine I.engti_Moduie Chamber Iliam_tcr14odt(le Thrc_t Dia:,,_,terModule Exit DiameterModule Cl,amber l.engLh14odule Noz:,]e Len::]_.hModule i._ngtl_Coolant Jacket il_.; R_tb_Coolaut ,]c.cl:et ,d_Cuol ant ] p] et Tem_J.:_._cI:qr::Coo! an_: Lxi _ Te;,lf,,:r;_i.u;',,

Sl UIIITS ALTERNATE UNITS

68.77 15,460 Ibf34.02 atm 500 psia

447.3 s5.44

15.68 kg/s 34.56 Ibm/s13.24 kg/s 29.19 Ibm/s

2.44 kg/s 5.37 Ibm/s458

40I02

247.4 cm 97.4 in168.1 cm 66.2 in106.2 cm 41.8 in

6.63 cm 2.61 in3.66 cm 1.44 in

23.11 cm 9.10 in16.51 cm 6.5 in27.41 cm 10.79 in52.27 cm 20.58 in2.44 kg/s 5.37 Ibm/s

10.3 atm 152 psia23 K 42 R

269 K 485 R

TUR_I HFS

Jnlet PressureI n l -,t rc;,:;,era tureGas [-IG: Ral.cSpecific lleat. R;:_ti,:,Molecular Wci(,hlShaft l!ors epowerPercent Bypass

55.3 atm 812 psia269 K 485 R

2.24 Kg/s 4.94 Ibm/s1.402.02 g/mol

520 kW 697 hp8

14t_IN PUi,IPS

Oxidizer I"u;1;l; F.o;v Ri:_:;.' 13.24 kg/sOxidizer Pu,_:p Discila_g_-: i'ressure 43.5 atmFuel Pump Flo,,; Ratc 2.44 kg/sFuel Pump Discharge Pressure 66.8 atm

29.19 Ibm/s640 psia

5.37 Ibm/s982 psia

(5.50 TCA)(34.50 TCA)

265

TABLELXVI - PLUGCLUSTERENGINEOPERATINGSPECIFICATIONGASGENERATORCYCLE:REGEN-MODULE

MODELII PERFORIVtANCEP = 20.4

C

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flaw RateFuel Flow RateEngine Area Ratio !AE/A T)Module Area Ratio (Ae/A t)Number of ModulesModule Gap (6/D e)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Exit DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket APCoolant Inlet TemperatureCoolant Exit Temperature

TURBINES


MAIN PUMPS

Oxidizer Pump Flow Rate


69.27 kN 15,573 Ibf20.41 atm 300 psia

443.25.33

15.97 kg/s 35.20 Ibm/s13.44 kg/s 29.64 Ibm/s

2.52 kg/s 5.56 Ibm/s458

40I0

2319.8 cm 125.9 in217.9 cm 85.8 in

85.9 cm 33.8 in8.59 cm 3.38 in4.72 cm 1.86 in

29.87 cm 11.76 in16.51 cm 6.5 in35.46 cm 13.96 in60.33 cm 23.75 in

2.43 kg/s 5.36 Ibm/s4.42 arm 65 psia

22 K 40 R376 K 677 R

6.53 atm 96 psia922 K 1,660 R

0.17 kg/s 0.38 Ibm/s1.363.8 g/mol

189 kW 254 hp0

13.44 kg/s 29.64 Ibm/sOxidizer Pump Discharge Pressure 27.22 armFuel Pump Flow RateFuel Pump Discharge Pressure

C_ASGENERATOR

Chamber PressureCombustion TemperatureMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow Rate

2.52 kg/s27.90 arm

6.80 atm922 K

0.90.17 kg/s0.08 kg/s0.09 kg/s

400 psia5.56 Ibm/s

410 psia

I00 psia1,660 R

0.38 Ibm/s0.18 Ibm/s0.20 Ibm/s

(5.50 TCA)(34.82 TCA)(29.46 TCA)(5.36 TCA)

266

..... _ o.

TABLE LXVII - PLUG CLUSTER ENGINE OPERATING SPECIFICATIONGAS GENERATORCYCLE: REGEN-MODULE

MODEL II PERFORMANCE

Pc : 34.0

PARAMETER

ENGINE

V:,cuum Thrust

Va c!_Ulil "'" "":_iJC'C1fi c Im',>ulseMixi:rre P,aCio ((_/i-}Total Fl or: [',ateOxidizer Fiow R,teFuel Flow Rate

Engine Area Ratio !AE/,_I)Module Area Ral;io _Ae/l',_)N,,irbcr of r,lodulcsl,ic_dule Gap (5/P e)Engine DiameterPlug Base I?iameLe ,_Engine I e;JgthMedule Chamber Pi._met_rF.odule Thro_,t Dianete'Modcle F',!it Di_'meLcrModule Chamhcr lengthMo(',u!eNc,zzle Lc'n]chModule LengthC-,ol_:u'Jacl<et Il _v1P.c,te

Cc,"]ant ,]_cb't /,PCoolar, t ]uleL Trmn_rotureCf_olar, t Exi t Tel,p__l e,t':r_

l UF,"?,! NES

Inlet P.,',.,":.sure

Iul(:i Tc:mperat!;,e...... FiSpecific Heat RatioMol ecul ar Wei _h',:Shaft !lo,'s epovr, rPercent Bypass

MA}I_ P[!!,iPS

Oxidizer P_:m;,Flov F'L_!.e


69.63 kN 15,650 psia34.02 atm 500 psia

446.45.25

15.91 kg/s 35.07 Ibm/s13.36 kg/s 29.46 Ibm/s2.54 kg/s 5.61 Ibm/s

45840lO

2247.4 cm 97.4 in168.1 cm 66.2 inI06.2 cm 41.8 in

6.63 cm 2.61 in3.66 cm 1.44 in

23.11 cm 9.10 in16.51 6.5 in27.41 cm I0.79 in52.27 cm 20.58 in

2.41 kg/s 5.31 Ibm/s10.3 atm 152 psia23 K 42 R

269 K 485 R

6.46 atm 95 psia922 K 1,660 K

0.26 kg/s 0.57 Ibm/s1.363.8 g/mol

316 kW 424 hp0

13.36 kg/sOxidizer Pu,np Discloser, I,.-.'Prcs.;ure 43.55 atm

2.54 kg/s49.47 arm

6.80 atm922 K

0.9O. 26 kg/sO. 12 kg/sO. 14 kg/s

Fuel Pu,,l'FIo'.' .kaLe:Fuel Pump Disc_;,_r!jc P;es_;u,'e

GAS GENFI,_AI0_

CNa IIlb(-"r Pre_ sure,

Combustion Temper )tureMixture Ratio (0/7)Total Flow RatcOxidizer glow RateFuel Flow !J,ate

29.46 Ibm/s640 psia

5.61 Ibm/s727 psia

lO0 psia1,660 R


(5.50 TCA)(34.50 TCA)(29.19 TCA)

(5.31TCA)

267

TABLE LXVIII -

PARAMEIrp,

ENGINE

Vacuum Ti_rustChan_ber PressureVacutm, Specific li_lJulse14i×_ur'e Ratio (O/F)[oral lIn;¢ RateOxidizer Fio,v RateFuel Plm; RateEngine Area Ratio (AE/A T)_.Ic_dule Area R_._.io ,,'c, :,,_Ulil[)er of Modules

Modiile Gap (_S/De)[ngine l)iameterPlug Base DialreterEntitle LengthI,iodule Chamber [)ia!_ieter

_ i hrou_- lal ,.. Buy

1!o(t:_".e ' -"L.X It OiT.tiT, L"L(::"r u t ,blc_dule c, a!_t-i l.'_ntjth

_bdu]e Nozzle Lengthl,ladu] e LengthCoolant Ja(:kot. Fie,.., i'.ater'oola:,,t. J_,.kc [ ,,._')t']OO ] [lilt: T r}'l (7_. "tef/]l)C :",L,_.tJ!_O

CQ,c: 1 <;_i L ix [ t Te_r,:,e ra l:tl'¢o

PLUG CLUSTER ENGINE OPERATING SPECIFICATIONEXPANDER CYCLE: ITA MODULE 16% FFC


C

_I !Jf_]_ AL.TE_iAC_ U!41TS...............................

68.88 kN 15.480 Ibf20.41 atm 300 psia

431.9 s5.44

16.26 kg/s 35,85 Ibm/s13.73 kg/s 30,28 Ibm/s2.53 kg/s 5,57 Ibm/s

4584OI02

319,8 cm 125,9 in217.9 cm 85.8 in85.9 cm 33.8 in

8.59 cm 3,38 in4.72 cm 1,86 in


2.53 kg/s 5.57 Ibm/s4.08 atm 60 psia

22 K 40 R104 K 188 R

(5.50 TCA)(35.79 TCA)

(5.51TCA)

TURB! I<ES

I nl et PrcsstmeInlet Temperatc, reGas Fie;.., RateSpecific !4e_._ i,',_tleMolecul .Jr _,.bl ,V!:tShaft Flor-_cpo_,_(,rPercent By p__ss

39.74 atm 584 psia104 K 188 R

2.53 kg/s 5,57 ]bm/s1.402.02 g/mol

353 kW 474 hp0

bIA! li PU_,2S

Oxidizer Pump FIJ;,._ P,ateO×idizer Pure!; lJischdige Pi-e%,_i-e

Fuel Pu:_lp Flow Rai,nFucl Ptml[, l)is,.:har!!,' l';(;f;qtlrc:

13.73 kg/s 30.28 Ibm/s27.22 arm 400 psia2.53 kg/s 5.57 Ibm/s

44.64 atm 656 psia

268

TABLELXIX - PLUGCLUSTERENGINEOPERATINGSPECIFICATIONGASGENERATORCYCLE: ITA MODULE16%FFC

MODELII PERFORMANCEP = 20.4

C

PARAMETER

ENGINE

V_:cuum Tl_rustCha,6er Pre%sureVacuum Specific l_g;:_!-,eMixture Ratio {O/F)To&_, Flow RateOxidizer Flow RateFuel Flow Raie

,,,.tlo (AFII,-F)Fr_gine Area _, "Hodule Area R,_Cio (_,,.ir 1)r_umber of. V,odttles

Modale Gap (_/De)Engine DiameterPlug Base Dial,'eter[ngil;e l.em.ithModule Chamber Diam,.:terMo'Jule Thro_.t DiameterigoduJe Exi_ Di_mel.er_,_,)duleCi_amber Le_q1',,

l,_o(iule No;.zle Lct+gtl-,Me;d,lle l_eng!:h

Coola_t Jacl.et Flou i_ateCoolum_t ,]aclcct A[;CoolanL !nIc,t Tempu_aLureCool_,nt F;,ii- -lenli-,or<,ti_re

TUI:[',i_!FS

]nlel Pr;;s _;tire'_]',:t l'O'.'il)C;r('_!11','"Gas Fl(;__:,Le

Speclfic ;',cnLP,ali,,Ho lecul Jr {h.>! g!itSh;:,ft Ho:'L;opo;;erl'ercenT Bypass

MA_N l_!.il,4!_S

Oxidizer l>t,:np FIe'.., Tb;i.tOxi di __er P!.:;nV lJi _ _I,;. :'ge Prt:'.;s ureFuel Pur,,p Flow Rat_,Fuel f>ui!li; i)ischarg:-: Pi:-ssure

GAS GFi.IERAIOR

Chamber P;'ps g_IreCoi,_Ju;; t i o_; Tem?eratureMixture Ratio (0/i)Total Flow RateOxidizer Fiow Rat..._Fuel I:lov: i_c..,{:r-

Sl UNITS

69.38 kN20.41 atm

431.45.34

16.40 kg/s13.81 kg/s

2.59 kg/s4584OI0

2319.8 cm217.9 cm

85.9 cm8.59 cm4 72 cm

29 87 cm16 51 cm35 46 cm60 33 cm

2 50 kg/s4 08 atm

22 K104 K

6.46 atm922 K

O. 16 kg/sI. 363.8 g/mol

189 kW0

13.81 kg/s27.22 atm

2.59 kg/s27.56 atm

6.80 atm922 K

0.9O. 16 kg/s0.08 kg/s0.09 kg/s

ALTERNATE UNITS

15,596 Ibf

300 psia

36.15 Ibm/s30.45 Ibm/s5.70 Ibm/s

125.9 in85.8 in33.8 in

3.38 in

1.86 in

11.76 in6.50 in13.95 in23.75 in

5.51 Ibm/s

60 psia40 R188 R

95 psia1,660 R

0.36 Ibm/s

254 hp

30.45 Ibm/s400 psia

5.70 Ibm/s405 psia

I00 psia1,660 R


(5.50 TCA)

(35.79 TCA)(30.28 TCA)(5.51TCA)

269

TABLE LXX - PLUG CLUSTER ENGINE OPERATING SPECIFICATION

PARAt.IETER

ELrGIN[

Vacuum Thru_,tChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flo_ Rate

Engine Area Ratio (AE/AT)Module Area Ratio (Ae/At)Number of Fk)du-!esModule Gap (..i ;e,Engine Diame,:erPlrg Base Diar,_.terEngine LengthMoHule Chamber Did:mete;Module Throat Dia!_:eterModule Exit DiameterModule Chamber LeJ,,'TLhModule Nozzl.f [.e;_sthil,;dul e Length

rCoolant Jacket rl,.v R_LeCoolant Jackal ,;PCoo!ani: Inlet-'-_ ........Coolant Exit l_iq,erature

TURBIIIEL,

In!at PressureI,,l e t Temper,,L,:_-__Gas Flow R_,_.mSpecific I!e,.t '_111oMolecular 1.1_.i w."uShaft liors e_;o;,,_,rPercenl Bypass

MAI _{ PU_;PS

Oxidi;r.r ;'w:T: Flo;.: Rai:eOxidizer Pu;.,2 Dischc, rge PressureFuel Pump Flov. RaLeFuel Pu_np Discharge Pressu_e

GAS GFNERAIOR

Chamber PressureCombus t i on Ic:;_pe ra t ureMixture Ratie (O/F)Total Flo,._ RaT,_Oxidii:er F_,.,v.,I_c,teFuel [ lo','_R_:Lc

GAS GENERATORCYCLE: ITA MODULE 16% FFC


C

SI UNITS AL'[FRNA_E ,_,t_,"

69.66 kN 15,660 Ibf34.02 atm 500 psia

434.65.26

16.34 kg/s 36.03 Ibm/s13.73 kg/s 30.27 Ibm/s

2.61 kg/s 5.76 Ibm/s4584OI02

247.4 cm 97.4 in168.1 cm 66.2 in106.2 cm 41.8 in

6.63 cm 2.61 in3.66 cm 1.44 in


2.48 kg/s 5.46 Ibm/s9.53 atm 140 psia

23 K 42 R97 K 174 R

6.46 atm 95 psia922 K 1,660 R

0.26 kg/s 0.57 Ibm/s1.363.8 g/mol

316 kW 424 hp0

13.73 kg/s 30.27 Ibm/s43.55 atm 640 psia

2.61 kg/s 5.76 Ibm/s48.65 atm 715 psia

6.80 atm lO0 psia922 K l ,660 R

0.90.26 kg/s 0.57 Ibm/s0.12 kg/s 0.27 Ibm/s0.14 kg/s 0.30 Ibm/s

(5.50 TCA)(35.46 TCA)(30.00 TCA)

(5.46 TCA)

27O

I!| I

TABLELXXI - PLUG CLUSTER ENGINE OPERATING SPECIFICATIONEXPANDERCYCLE: REGEN-MODULE/UNCOOLEDPLUG


C

PARAMETER

ENGINE


Engine Area Ratio (AE/ATIModule Area Ratio (Ae/A t)Number of Modules

Module Gap (_/D e)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Exit DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket APCoolant Inlet TemperatureCoolant Exit Temperature

SI UNITS

68.72 kN20 41 atm

443 85 44

15 82 kg/s13 36 kg/s2 46 kg/s

45840I02

319 8 cm217 9 cm85 9 cm

8 59 cm4 72 cm

29 87 cm16 51 cm35 46 cm60.33 cm2.46 kg/s2.04 atm

22 K246 K

ALTERNATE UNITS

15,451 Ibf300 psia

34.88 ]bm/s29.46 Ibm/s5.42 Ibm/s

(5.50 TCA)(34.82 TCA)

125.9 in85.8 in33.8 in3.38 in1.86 in

11.76 in6.5 in

13.96 in23.75 in

5.42 Ibm/s30 psia4O R

442 R

TURBINES

Inlet Pressureinlet TemperatureGas Flow RateSpecific Heat RatioMolecular WeightShaft HorsepowerPercent Bypass

32.7 atm246 K1.38 kg/s1.402.02 Q/mol

290 kW44

480 psia442 R

3.76 Ibm/s

390 hp

MAIN PUMPS

Oxidizer Pump Flow RateOxidizer Pump Discharge PressureFuel Pump Flow RateFuel Pump Discharqe Pressure

13.36 kg/s27.2 atm

2.46 kg/s36.7 atm

29.46 I bm/s400 psia

5.42 Ibm/s540 psia

271

TABLELXXII - PLUGCLUSTERENGINEOPERATINGSPECIFICATIONEXPANDERCYCLE:REGEN-MODULE/UNCOOLE[)PLUG

MODELII PERFORMANCEPc = 34.0

PARAMETER

ENGINE

Vacuum ThrustChamber PressureVacuum Specific ImpulseMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow RateEngine Area Ratio (AE/A T)Module Area Ratio (Ae/A t)Number of ModulesModule Gap (a/De)Engine DiameterPlug Base DiameterEngine LengthModule Chamber DiameterModule Throat DiameterModule Exit DiameterModule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket APCoolant Inlet TemperatureCoolant Exit Temperature

TURBINES


MAIN PUMPS

Oxidizer Pump Flow RateOxidizer Pump Discharge PressureFuel Pump Flow RateFuel Pump Discharge Pressure


68.77 kN 15,460 Ibf34.02 arm 500 psia

447.35.44

15.68 kg/s 34.56 Ibm/s13.24 kg/s 29.19 Ibm/s

2.44 kg/s 5.37 Ibm/s458

40lO

2247.4 cm 97.4 in168.1 cm 66.2 in106.2 cm 41.8 in

6.63 cm 2.61 in3.66 cm 1.44 in

23.11 cm 9.10 in16.51 cm 6.5 in27.41 cm 10.79 in52.27 cm 20.58 in2.44 kg/s 5.37 Ibm/s5.58 atm 82 psia

23 K 42 R218 K 392 R

54.98 atm 808 psia218 K 392 R

2.23 Kg/s 4.92 Ibm/s1.402.02 g/mol

506 kW 679 hp15

13.24 kg/s 29.19 Ibm/s43.5 atm 640 psia

2.44 kg/s 5.37 Ibm/s64.62 atm 950 psia

(5.50 TCA)(34.50 TCA)

272

TABLE LXXIII - PLUG CLUSTER ENGINE OPERATING SPECIFICATION

GAS GENERATOR CYCL[: REGEN-MODULE/UNCOOLED PLUGMODEL II PERFORMANCE

Pc = 20.4

PARAMETER

ENGINE

Vacuum ThrustChan_)e r Pressure

Vacuum Specific h,I>ulseMixture Ratio (O/F)Total Flow Rate

Oxidizer Flow RateFuel Flow Rate

Engine Area Ratio (AI/AT)Module Area Ratio (Ae/At)Nun_er of Modules

Module Gap (_S/De)Engine Dian_ter

Plug Base Diameter

Engine LengthModule Chamber Diameter

Module Throat DiameLerModule Exit DiaiIwDter

Module Chanl)er Length

Module Nozzle Length

Module LengthCoolant Jacket Flow IRateCoolant Jacket AP

Coolant Inlet Temperature

Cool ant Exit Temperature

TURBINES

Inlet Pressure

Inlet TemperatureGas Flow Rate

Specific lleat Ratio

Molecular Weight

Shaft llorsepower

Percent Bypass

MAIN PUMPS

Oxidizer Pump Flow Rdte

Oxidizer Pump Dischar(ie PressureFuel Pump Flow Rate

Fuel Pump [)ischarge Pressure

C_S G[_NERATOR

Chamber Pressure

Con_us t i on Tempera Lure

Mixture Ratio (O/F)Total Flow Rate

Oxidizer flow Rat(,Fuel Flow Rate

SI UNITS

69.27 kN

20.41 atm443.2

5.33

15.97 kg/s13.44 kg/s

2.52 kg/s45}_

4010

2310. g cm217.() cm

g5.9 (:m_l. 59 ('m

4 72 cm29 }_7 cm16 51 cm35 46 cm60 33 cm

2 43 kg/s2.04 atm

22 K246 K

ALTERNATE UNITS

15,573 Ibf

300 psia

(5,50 TCA)35.20 Ibm/s (34.82 TCA)29.64 Ibm/s (29.46 TCA)

5.56 lbm/s (5.36 TCA)

125.9 in85._I in33. H in

3.38 ill1.86 in

11.76 in6.5 in

13.96 in23.75 in

5.36 lbm/s30 psia_0 R

442 R

6.53 atm 96 psia922 K 1,660 R

0.17 kg/g 0.38 Ibm/s1.36

3.g g/H_I

189 kW 254 hp0

13.44 k(I/S 20.64 1hm/s

27.22 atm 40(} psia2.52 kq/s 5.56 Ibm/s

27.22 arm 400 psia

6.}]0 atm I00 psia()22 K 1,660 R

0.()

0.17 kH/', ().3H lhm/s

0,08 kg/s (].IH lbmls

0.09 kg/s 0.20 ll)m/s

273

TABLE LXXIV - PLUG CLUSTER ENGINE OPERATING SPECIFICATION

PARAMETER

ENGINE


Engine Area R_tio (A_/AT)Nodule Area Ratio (Ae/At)Number of ModulesModule Gap (6/D e)Engine DiameterPlug Base DiameterEngine Length_bdule Chamber DiameterModule Throat DiameterNodule Exit DiameterNodule Chamber LengthModule Nozzle LengthModule LengthCoolant Jacket Flow RateCoolant Jacket &PCoolan._ Inlet TemperatureCoolant Exit Temp_=rature

TURBINES

Inlet PressureInlet TemperatureGas Flow RateSpecific Heat RatioMolecular UeightShaft HorsepowerPercent Bypass

MAIN PUt.IPS

Oxidizer Pump Flo,v RateOxidizer Pump Dischar!le PressureFuel Pump Flow RateFuel Pump Discharge Pressure

GAS GEIIERATOR

Chamber PressureCombustion TempeFatureMixture Ratio (O/F)Total Flow RateOxidizer Flow RateFuel Flow Rate

GAS GENERATORCYCLE: REGEN-MODULE/UNCOOLEDPLUGMODEL II PERFORMANCE

Pc = 34.0


69.63 kN 15,650 Ibf34.02 atm 500 psia

446.45.25

15.91 kg/s 35.07 Ibm/s•13.36 kg/s 29.46 Ibm/s

2.54 kg/s 5.61 Ibm/s458

4O]02

247.4 cm 97.4 in168.1 cm 66.2 in

I06.2 cm 41.8 in

6.63 cm 2.61 in

3.66 cm 1.44 in23.11 cm 9.10 in16.51 6.5 in

27.41 cm 10.79 in

52.27 cm 20.58 in

2.41 kg/s 5.31 Ibm/s

5.58 atm 82 psia23 K 42 R

218 K 392 R

6.46 atm 95 psia922 K 1,660 K

0.26 kg/s 0.57 Ibm/s1.363.8 9/mo1

316 kW 424 hp0

13.36 kg/s 29.46 Ibm/s43.55 arm (40 psia

2.54 kg/s 5.61 Ibm/s47.63 atm 700 psia

6.80 arm iO0 psia922 K 1,(60 R

0.9

0.26 kg/s 0.57 Ibm/s

0.12 kg/s 0.27 Ibm/s0.14 kg/s 0.30 Ibm/s

274

(5.50 TCA)(34.50 TCA)(29.19 TCA)(5.3] TCA)

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275

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i

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_g i i?

276

B. SPACE TUG SYSTEM STUDY BIBLIOGRAPHY

I. "Space Tug Systems Study (Cryogenic)," Vol. I: Final Overview Briefing

Manual, McDonnell-Douglas Astronautics Co., Contract NAS8-29677, NASA-CR-120459, MDC G5031, Jan 1974.

2. Ibid, Vol. 2: Technical Compendium Final Report, NASA-CR-120460, MDCG5029.

3. Ibid, Vol. 3: Executive Summary Final Report NASA-CR-120461, MDC G5030.

4. Ibid, Data Dump Overview, MDC G4813, Sep 1973,

5. Ibid, September Data Dump, Vol. l, Summary, Program Option l, Sep 1973.

6. Ibid, September Data Dump, Vol. 2, Summary, Program Option 2.

7. Ibid, September Data Dump, Vol. 3, Summary, Program Option 3.

8. Ibid, Concept Selection Overview, MDC G4770, Jul 1973.

9. Ibid, First Review Overview, MDC G4729, May 29, 1973.

|0. "Space Tug Systems Study (Cryogenic)," Final Report, Vol. I Final

Briefing, General Dynamics, Convair Aerospace Division, ContractNAS8-29676, CASD-NAS73-033, Jan 1974.

II. Ibid, Vol. II - Compendium.

12. Ibid, Vol. Ill - Executive Summary.

13_ Ibid, First Review, May 29, 1973.

14. Ibid, Data Dump, Vol. I - Program l Summary, Sep 18, 1973.

15. Ibid, Vol. II - Program 2 Summary.

16. Ibid, Vol. III- Program 3 Summary.

17. "Space Tug Systems Study (Storable)," Final Report, Vol. l, Overview

Presentation, Martin Marietta Corp., Contract NAS8-29675, MCR-73-314,Jan 1974.

18. Ibid, Vol. 2, Compendium.

19. Ibid, Vol. 3, Executive Sumn_ry.

20. Ibid, Requirements Assessment Presentation, MCR-73-82, II Apr 1973

277

21. Ibid, First Review Presentation, Propulsion Addendum, 30 May 1973.

22. Ibid, Program Concept Evaluation, MCR 73-193, 19 Jul 1973.

23. Ibid, Propulsion Addendum, 19 Jul 1973.

24. Ibid, Vol. l, Summary, Program Option l, Selected Option Data Dump,

MCR 73-235, Sep 1973.

25. Ibid, Vol. 2, Summary, Program Option 2

26. Ibid, Vol. 3, Summary, Program Option 3

27. "Storable Space Tug Systems Study", Final Presentation, Vol. I, Grumman

Aerospace Corp., Contract NAS8-29674, 300 RP-73-013 Final Report,31 Jan 1974

28. Ibid, Final Report, Vol. 2, 300 RP-73-015

29. Executive Summary, Vol. 3, 300 RP-73-014.

30. Ibid, Requirements Assessment Meeting B80-RP-95-73-O02, 12 Apr 1973.

31. "Space Tug First Review," Propulsion Panel Meeting, Grumman AerospaceCorp., Contract NAS8-29674, 31 May 1973.

32. "Storable Space Tug Systems Study," Concept Selection Meeting, GrummanAerospace Corp., Contract NAS8-29674, B80-RP-95-73-O04, 20 July 1973.

33. Ibid, Propulsion Panel Meeting, 21Ju] 1973.

34. Ibid, Data Dump, Vol. I - Summary, Program Option I, 300 RP-73-004,19 Sep 1973.

35. Ibid, Vol. II - Summary, Program Option II, 300 RP-73-005.

36. Ibid, Vol. Ill - Summary, Program Option Ill, 300 RP-73-006.

37. "Baseline Tug Definition Document," Rev. A, Preliminary Design Office,Program Development, NASA Marshal1 Space Flight Center, 26 Jun 1972.

38. "Space Tug Point Design Study," Vol. I, Management and Technical Summary,

McDonnell Douglas Astronautics Co., Contract NAS7-101, MDC G2818, Feb 1972.

39. Ibid, Vol. If, Design and Analysis.

40. "Space Tug Point Design Study, Final Report, Vol. I, Summary, NorthAmerican Rockwell Space Division, Contract NAS 7-200, SD72-SA-O032,II Feb 1972.

278

V!I

41.

42.

43.

44.

45.

46.

47.

48.

49.

50.

51.

52.

55.

56.

57.

58.

Ibid, Vol. II, Operations, Performance, and Requirements.

Ibid, Vol. III, Design Definition.

Ibid, Vol. IV, Program Requirements.

Ibid, Vol. V, Cost Analysis

"MSC In-House Storable Propellant Space Tug Study Report Draft,"Payloads Engineering Office, Future Programs Division, NASA MannedSpacecraft Center, Jan 1973.

"Addendum I, Space Tug System Study Data Package, Propulsion SystemsControl Data," NASA/Marshall Space Flight Center, 17 April 1973.

"Baseline Space Tug System Requirements and Guidelines," NASA/Marshall

Space Flight Center, 15 Jul 1974 MSFC 68M00039-I

"Baseline Space Tug Configuration Definition," NASA/Marshall SpaceFlight Center, MSFC 68M00039-2, 15 July 1974.

"Baseline Space Tug Flight Operations," NASA/Marshall Space FlightCenter, MSFC 68F_00039-3, 15 Jul 1974.

"Baseline Space Tug Ground Operations, Verification, Analysis and Processing,"NASA/Marshall Space Flight Center, MSFC 68M00039-4, 15 July 1974.

"Baseline Space Tug Executive Summary," NASA/Marshall Space Flight Center,MSFC 68M00039-5, II Oct 1974.

"Orbit-To-Orbit Shuttle (Chemical) Feasibility Study," Vol. 1 - Management

Summary, McDonnel Douglas Astronautics Co., Contract F04701-71-C-0173,SAMSO-TR-71-221-1, Oct 1971.

Ibid, VoI. 2 - Technical Sunwnary, SAMSO-TR-71-221-2.

Ibid, Vol. 3, Book 1 - Missions, Operations and Requirements, SAMSO-TR-71-221-3, Book 1

Ibid, VoI. 3, Book 2 - Missions, Operations & Requirements, Appendix,SAMSO-TR-71-221-3, Book 2.

Ibid, Vol. 4, Book 1 - Design and Integration, SAMSO-TR-71-221-4, BookI.

Ibid, Vol. 4', Book 2 - Design and Integration - Appendices, SAMSO-TR-71-221-4, Book 2.

Ibid, VoI. 5, Book 1 - Subsystenm: Structures and Propulsion, SAMSO-TR-71-221-5, Book I.

279

59. Ibid, Vol. 5, Book2 - Subsystems:Avionics, SAMSO-TR-71-221-5,Book2.

60. Ibid, Vol. 5, Book3 - Avionics, Appendix, SAMSO-TR-71-221-5,Book3.

61. Ibid, Vol. 6 - ProgramPlans and Costs, SAMSO-TR-71-221-6.

62. "ExtendedApplications Study of AMOOSAeromaneuveringOrbit to OrbitShuttle and AMRSAeromaneuveringe-_-_en-trySystem,Orientation and StudyPlan Review,_heed Missiles & SpaceCo., Huntsville, Contract NAS8-31997,LMSC-HRECSDD496798,Apt 1976.

63. "SpaceTug Storable EngineStudy," Vol. I: Executive Summary,AerojetLiquid RocketCo., Contract NAS8-29806, UPD 396; DR-MA-04, 31 Jan 1974.

64. Ibid, VoI. II: Engine Design and Performance Characteristics, DPD 396;DR-MA-03.

65. Ibid, Vol. III: Engine/Tug Interface and Operational Requirements, DPD396; DR-MA-03.

66. Ibid, Vol. IV: Project Planning Data, DPD 396; DR-MA-03.

67. Ibid, Final Program Review, DPD 396; DR-MA-02-9, 22 Jan 1974.

68. "Advanced Space Engine Preliminary Design Final Report," Pratt & WhitneyAircraft Div., Contract NAS3-16750, NASA CR-121237, FR-5654, Dec 1973.

69. "Advanced Space Engine Preliminary Design, Rocketdyne Div. Rockwell Intl.,Contract NAS3-16751, NASA CR-121236, R-9269, Oct 1973.

70. "O_/He Advanced Maneuvering Propulsion Technology Program Engine SystemStbdi_s Final Report," Vol. I: Aerospike Engine Configuration Design andAnalysis, Rocketdyne Div., No. American Rockwell Corp., Contract F04611-67-C-0116, AFRPL-TR-72-4, Dec 1971.

71. Ibid, Vol. II: 25,000-Pound-Thrust Bell Engine Configuration Design andAnalysis.

72. Ibid, Vol. III: lO,O00-Pound-Thrust Bell Engine Configuration Design andAnalysis, Jan 1972.

73. "Design Study of RLIO Derivatives," Final Report Vol. I Program Summary,Pratt & Whitney Aircraft Div., Contract NAS8-28989, FR-6011Vol. I, 15Dec 1973.

74. Ibid, Vol. II Engine Design Characteristics.

75. Ibid, Vol. II Appendices.

280

I!I !

76. Ibid, Vol. III, Part 1 Preliminary interface Control Document.

77. Ibid, Vol. III, Part 2 Operational and Flight Support Plan.

78. Ibid, Vol. IV,DevelopmentPlans and ProgramCosts.

79. "Application of Bell 8096Engine for SpaceTug Propulsion," Final Report,VolumeI - Executive Summary,Bell AerospaceDiv., Contract NAS8-29765,DPD394; DR-MA03,Report 8096-910513,Feb 1974.

80. Ibid, Vol. II - EngineDesignand PerformanceAnalysis.

81. Ibid, Vol. III - Engine/Tug Interface and Operational Requirements.

B2. Ibid, Vo1. IV - Project Planning Data.

83. Ibid, Vol. V - Costs.

84. "Pre-PhaseA Study for an Analysis of a ReusableSpaceTug," Vol. I -ManagementSummary,North AmericanRockwell Corp., Contract NAS9-10925,NASA-CR-II4937,SD-71-292-I-VoI. I, 22 Mar 1971.

85. Ibid, Vol. II Technical Summary,NASA-CR-II4936,SD-71-292-2-VoI. II.

86. Ibid, Vol. III - Mission and Operations Analysis, NASA-CR-II4935,SD-71-292-3-Voi. III.

87. Ibid, Vo1. IV - Spacecraft Conceptsand SystemsDesign, NASA-CR--I14938,SD-71-292-4TVol.IV.

88. Ibid, Vol. V - Subsystems,NASA-CR-II4939,SD-71-292-5-VoI. V.

89. Ibid, Vol. VI - Planning Documents,NASA-CR-II4940,SD-71-292-6-VoI.VI.

90. "Performanceof RecoverableSingle and Multiple SpaceTugsfor MissionsBeyondEarth Escape," NASA/LeRC,NASA-TM-X-68136,Nov1972. (Zinmlerman,A.V.)

91. "EuropeanSpaceTug Pre-PhaseA Study," Appendix3 Propellant Controland Propulsion, HawkerSiddeley DynamicsLtd., ELDO-CTR-17/7/44,HSD-TP-7227-APP-3,Jan 1971.

92. "IUS/TugOperations and Mission Support Study," Vol. III SpaceTug Operations,International BusinessMachinesCorp., Contract NAS8-31009,NASA-CR-143855,IBM-75W-OOO72-VoI.III, May1975.

93. "DODUpperStage/Shuttle SystemPreliminary RequirementsStudy," Vol. ISummary,North AmericanRockwell Corp., Contract F04701-72-C-0305,SAMSO-TR-72-195-Voi. I, SD-72-SA-OII5-1-VoI. I, Aug1972.

94. Ibid, Vol. II OOSVehicle and Shuttle Interface Analysis, SD-72-SA-OII5-2.

95. Ibid, Vo1. II, AppendixI DODUpperStage Functional Analssis and SystemRequiren_nts, SD-72-SA-OII5-3.

281

98.

99.

I00.

I01.

102.

Ibid, Vol. II, Appendix II ReferenceShuttle Definition, SD-72-SA-OII5-4.

"A Study of Compatibility of a Cryogenic UpperStagewith SpaceShuttle,"General Dynamics/Convair,Contract NAS3-14389,NASA-CR-120897,GDCA-BNZ71-020-7, Apr 1972.

"Orbit-To-Orbit Shuttle EngineDesignStudy," Final Report Book1Parametric Cycle Study, Aerojet Liquid RocketCo., Contract F04611-71-C-0040,AFRPLTR-72-45BookI, May1972.

Ibid, Book2 25KLB EngineDesign.

Ibid, Book3 25KLB EngineMaintenance,DevelopmentPlans, Cost Estimatesand 1OKLB EngineDesign.

Ibid, Book4 Appendices.

"Study of Liquid Oxygen/LiquidHydrogenAuxiliary Propulsion Systemsforthe SpaceTug," Final Report, SpaceDiv., Rockwell International Corp.,Contract NAS3-18913.NASACR-134790,SD75-SA-0043,15 Jun 1975.

282

li! !

C. PLUGANDPLUG-CLUSTERROCKETNOZZLEBIBLIOGRAPHY

°

,

.

°

,

°

°

°

I0.

II.

12.

"Results of a Program for Theoretical and Experimental Evaluation of thePlug Nozzle Concept", Aerojet-General Corporation (V. H. Ransom,S. A. Lorenc, J. J. Williams, T. L. DeYoung), Technical Memorandum No.159 SRP, 8 Mar. 1961.

"Results of Plug-Nozzle Cold-Flow Experiments", Aerojet-GeneralCorporation (J. J. Williams), Technical Memorandum 160 SRP, 20 Apr. 1961.

"Mode] Tests of Several Rocket Exhaust Nozzle Configurations IncludingThrust Vector Control Devices", Fluidyne Engineering Corporation(R. G. Brasket and C. W. Landgraf), Report on F.E.C. Project 0160,Nov. 1960.

"A Second-Order Numerical Method of Characteristics for Three-DimensionalSupersonic Flow", Volume I. Theoretical Development and Results, JetPropulsion Center, Purdue University (V. H. Ransom, J. E. Hoffman, andH. D. Thompson), Lafayette, Indiana 47907, Technical Report AFAPL-TR-69-98,Oct. 1969.

"Performance of Annular Plug and Expansion-Deflection Nozzles IncludingExternal Flow Effects at Transonic Mach Numbers", Lewis Research Center(R. A. Wasko), Cleveland, Ohio, NASA TN D-4462, Apr. 1968.

"Design and Analysis of Altitude-Compensating Rocket Nozzles", Aerojet-General Corporation (C. A. Bryce, T. L. DeYoung, J. C. Eschweiler, andR. L. Strickler), Report No. 8548-SR-I, Dec. 1964.

"Study of High Effective Area Ratio Nozzles for Space Craft Engines",Aerojet-General Corporation Report No. NAS7-136-F, Vol. 1 and 2, ContractNAS7-136, Jun. ]964.

"J-2T Aerodynamic Spike Thrust Chamber Study, Phase Zero Final Report",Advanced Projects Department Rocketdyne Engineering, Report No. R-5905,Contract No. NAS8-19, 15 May 1965.

"Mu]tichamber Plug-Nozzle Systems Analysis", Rocketdyne (M. S. Bensky),Report No. R-6494, Contract No. NAS8-20237, May 1966.

"Advanced Engine Aerospike Experimental Program Thrust Chamber Performance"(Volume 2, Final Report), Rocketdyne Engineering, Report No. R-7760,Contract NAS8_20237, May 1966.

"Plug Nozzle Study" - Final Report, Pratt and Whitney Aircraft, Report No.FR-IO13, Contract NAS8-II023, 8 Sep. 1964.

"An Experimenta] Hot Rocket Model Investigation of a Plug Cluster NozzlePropulsion System. Part I: Base Thermal and Pressure Environnmnt for aModule Chamber Pressure of 300 psia and Simulated Altitudes to 150,000Feet", Cornell Aeronautical Lab. (R. J. Sergeant), Cal No. HM-2024-Y-5 (I),Contract No. NAS8-20027, Sep. 1967.

._ _/_ 283

13.

14.

15.

16.

17.

18.

"Trade-Off Between Different Possible Engine Concepts Using LOX/LH2 --

Low Thrust High Pressure Engines for European Launch Vehicle With Con-ventional Bell Nozzles or Aerospike Nozzles, Europa 3, Stage 2,"Cryorocket, MBB and SEP, NASA DCAF E003091, ELDO-CTR-17/3/191, 30 Mar 1973.

"Space Tug Engine Study, Phase A -- Ranking of Different Engine Designs,ELDO Launch Vehicle Studies," Cryorocket, Neuilly (France), NASA DCAF

E003091, ELDO-CTR-17/3/174, 21 Aug 1972.

"A Liquid Ox_vgen-Liquid Hydrogen Rocket Engine for Satellite LaunchingVehicle Upper Stages -- Providing for De Havilland Third Stage of Multi-Engines Installation in Second Stage," Rolls-Royce Ltd. (England), NASADCAF E003091, Mar 1973.

"Zero NPSP Pumping Tests with an Operational Hydrogen Turbopump," Aerojet-

General Corp., Nuclear Rocket Operations (W. E. Campbell), in 9th Liq.

Prop. Symp., Vol. II, pp. 169-81, Sep 1967.

"Mark Ill Turbopump Assembly Performance Test Results," Aerojet-General

Corp., NASA-CR-79824, RN-S-OIIO, SNP-I, Sep 1964.

"Operation Handbook, Partial Model Turbopump P/N P9230792," Aerojet-GeneralCorp. (C. J. Reynolds and S. R. Andrus), Contract DAAG05-69-0453, AGCRMR-O127, Dec 1969.

284

D. H/OTHRUSTCHAMBERBIBLIOGRAPHY

l °

.

°

.

.

.

.

t

.

I0.

II.

12.

13.

"Hydrogen-Oz_vgen Space Shuttle ACPS Thruster Technology Review," NASA/LeRC(J. W. Gregory and P. N. Herr), AIAA Paper 72-1158, 29 Nov 1972.

"Integrated Thruster Asse_ly Program," Aerojet Liquid Rocket Co., ContractNAS3-15850, NASA CR-134509, Nov 1973.

"Extended Temperature Range ACS Thruster Investigation," Aerojet LiquidRocket Co. (A. L. Blubaugh and L. Schoenman), Contract NAS3-16775, NASACR-134655, 1 Aug 1974.

"Hydrogen-Oxygen Auxiliary Propulsion for the Space Shuttle," Volume I:High Pressure Thrusters, Aerojet Liquid Rocket Co., Contract NAS3-14354,NASA CR-120895, 30 Jan 1973.

"Hydrogen-Oxygen APS Engines," Volume I: High Pressure Thruster, Rocket-dyne Div. of NAR (R°D. Paster), Contract NAS3-14352, NASA CR-120805,Feb. 1973.

"High Pressure Reverse Flow APS Engine," Bell Aerospace Co. (J. M. Senneff),Contract NAS3-14353, NASA CR-120881, Nov 1972.

"Investigation of Gaseous Propellant Combustion and Associated Injector/Chamber Design Guidelines," Aerojet Liquid Rocket Co. (D. F. Calhoon,J. I. Ito, and D. L. Kors), Contract NAS3-14379, NASA CR-121234, 31 Jul 1973.

"Combustion Effects on Film Cooling," Aerojet Liquid Rocket Co. (D. C. Rousarand R. L. Ewen), Contract NAS3-17813, NASA CR- , 1976.

"Test Results with a Regeneratively Cooled LOX/LH2 Thrust Chamber for theAriane 3rd Stage Propulsion System," Messerschmitt-Boelkow-Blohm GMBH(H. Dederra), AIAA Paper 75-I189, Sep 1975.

"Development of Pulsed Hydrogen/Oxygen Attitude-Control Engines,"Messerschmitt-Boelkow-Blohm GMBH (G. Pulkert), DGLR Paper 72-077,Oct 1972.

"Investigation of Thrusters for Cryogenic Reaction Control Systems," TRWSystems Group (R. J. Johnson), Contract NAS3-I1227, NASA-CR-72608,Mar 1970.

"Space Shuttle High Pressure Auxiliary Propulsion Subsystem Definition,"TRW Systems Group (R. A. Benson, et al), Contract NAS9-11013, NASA-CR-115162, Mar 1.971.

"Ignition Devices for ACPS," Aerojet Liquid Rocket Co. (S. D. Rosenberg),Contract NAS3-14348, in MSFC Proc. - Space Transportation System Pro-pulsion Technol. Conf., Vol. 2, pp. 613-64, 28 Apt 1971.

285

14.

15.

16.

"Space Shuttle ACPS Shutoff Valve," Rocketdyne (G. M. Smith), ContractNAS3-14350, in MSFC Proc. - ibid, Vol. 2, pp. 563-612, 28 Apr 1971

"Space Shuttle ACPS Shutoff Valve," Marquardt Corp. (H. Wichmann),Contract NAS3-14349, in MSFC Proc. ibid, Vol. 2, pp. 501-62,

28 Apr 1971.

"Space Tug Engine Performance Verification Program," Vol. I, ProgramSummary, Pratt & Whitney Aircraft, Contract NASS-31008, FR-7509,

14 May 1976.

17. Ibid, Vol. II, Final Report.

286

E. H/O TURBOPUMPASSEMBLY BIBLIOGRAPHY

,

Q

"Turbopumps for Cryogenic Upper-Stage Engines," Rocketdyne Div.(A. T. Zachary, A. Csomor, and L. L. Tignac), Contract NAS8-27794,R-9367, Dec 1973.

"Design Study of RL-IO Derivatives," Vol. II, Engine Design Character-istics, Pratt & Whitney Aircraft, Contract NAS8-28989, FR-6011,15 Dec 1973.

3. Ibid, Vol. II Appendices.

4. "Turbopump Systems for Liquid Rocket Engines," NASA Space Vehicle DesignCriteria (Chemical Propulsion), NASA SP-8107, Aug 1974.

°

°

,

°

.

I0.

II.

12.

"An Experin_ntal Investigation of Two-Phase Liquid Oxygen Pumping," NASA/MSFC (L. A. Gross), NASA-TN-D-7451, Nov 1973.

"Study and Technological Work on Some Aspects of the U.S. Space ShuttleAuxiliary Propulsion System," Cryorocket, Neuilly (France), Rept-D-9/71-MK/DH,Aug 1971.

"Study on a Combined OMS/ACPS Propulsion System," Messerschmitt-Boelkow-Blohm GMBHand Cryorocket, MBB-UR-82-71, Aug 1971.

"Preliminary Design Study of Space Shuttle Auxiliary Propulsion System,"Societe Europeenne De Propulsion and Cryorocket (France), SEP-NT-3054/71-M,Aug 1971.

"Experimental Findings from Zero-Tank Net Positive Suction Head Operationof the J-2 Hydrogen Pump, NASA/MSFC (H. P. Stinson and R. J. Strickland),NASA-TN-D-6824, Aug 1972.

"Study of Influence of the Variation of the Parameters on the Functioningof the 200KN HP Engine -- Without Considering Return of Part of Pump Flowto Tank, Europa 3, Stage 2," Societe Europeenne De Propulsion (P. Bordier),NASA DCAF E003091, SEP-TL/LH-3049/71M, 29 Jun 1971.

"Studies of the Parameter Variations of the 200KN HP Engine -- ConsideringReturn of Part of Pump Flow to Tank, Europa 3, Stage 2," Societe EuropeenneDe Propulsion (P. Bordier), NASA DCAF E003091, SEP-TL/LH-3050/71M,29 Jun 1971.

"European Space Tug Engine Study Pre-Phase A General Studies," Cryorocketand Messerschmitt-Boelkow-Blohm GMBHand Soeiete Europeenne De Propulsion,NASA DCAF E003091, ELDO-CTR-17/3/IOI, 29 Jan 1971.

287

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15.

16.

17.

18.

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22.

"The Effects of BaseBleed on Plug Nozzles," Aeronautical ResearchCouncil,London,ARC-R&M-3466RAE-TR-65043,1967.

"A Report on the Constructive Design and on the Weight Estimation of theSkin Structure of the Plug-Nozzle of the SecondStageof the SpaceTrans-portation System,Neptun," TechnischeUniv., Berlin (J. Allesch),TUB-IR-3/1968,Mar. 1968.

"Analytical and Experimental Study of Axisymmetric Truncated Plug NozzleFlow Fields," Notre DameUniv. (T. J. Mueller, et al), Contract NAS8-25601,NASA-CR-123941,Sep. 1972.

"Comparisonof Performanceof a Double-CorneredPlug Nozzle with a Con-ventional Convergent-DivergentRocketNozzle," ArmyMissile Command,RedstoneArsenal (A. T. Stokes), RK-TR-72-17,Jun. 1972.

"Three DimensionalThrust ChamberLife Prediction," BoeingAerospaceCo.,Contract NAS3-19717,NASA-CR-134979,D180-19309-I, Mar. 1976.

"Fortran Programfor Plug Nozzle Design," NASA/MSFC(C. Lee and D. Thompson),Contract NASA-TM-X-53019,Jul. 1964.

"Truncated Plug Nozzle ComputerProgramDocumentation,"Vol. I, Notre DameUniv. (C. R. Hall), Contract NAS8-25601,NASA-CR-132965,UNDAS-PD-6OI-I-VoI-I,Jul. 1971.

Ibid, Vol. II, NASA-CR-132966.

"Plug-Nozzle Program," General Electric Co. (A. R. Graham),Contract NAS5-445,NASA-CR-135802,DM-61-154,Oct. 1961.

"Analytical and Experimental Investigations of High ExpansionArea RatioNozzles," Vol. I, Analytical and ModelCold-Flow Studies, Rocketdyne(D. W. Hege), Contract AF04(611)-8509,RPL-TDR-64-3-VoI.I, Dec. 1964.

23. "An Analytical and Experimental Study of NonuniformPlug Nozzle Flow Fields,"Notre DameUniv. (C. R. Hall and T. J. Mueller), NSR-15-O04-029,AIAAPaper71-41, Jan. 1971.

24. "Design of MaximumThrust Plug Nozzleswith Variable Inlet Geometry,"ColoradoState Univ. and Purdue Univ. (G. R. Johnson, H. D. Thompson, andJ. D. Hoffman), Computers and Fluids, Vol. 2, pp. 173-90, Aug. 1974.

25.

26.

"The Mach Disc in Truncated Plug Nozzle Flows," ARO, Inc. and Notre DameUniv. (T. V. Giel and T. J. Mueller), Contract _IAS8-25601, AIAA Paper75-886, Jun. 1975.

"The Computation of Steady Exhaust Nozzle Flows by a Time Dependent Method,"AVCO Corp. (K. P. Sarathy and R. Bozzola), AIAA Paper 76-151, Jan. 1976.

288

111]

27. "Fortran Program for Design of Pure External and Internal-ExternalExpansion Plug Nozzles," Brown Engineering Co. (C. C. Lee), ContractNAS8-2485, NASA-CR-51300, R-41, Mar. 1963.

28. "Fortran Computer Program to Design Plug Nozzle By Using Rao MaximumThrust Theory," Brown Engineering Co. (C. C. Lee), Contract NAS8-5289,NASA-CR-51301, Jul. 1963.

29. "Plug Nozzle Propulsion Efficiency," NASA/LaRC (C. E. Mercer and

L. B. Salters), NASA-TN-D-18O4, May 1963.

30. "NASA Plug Nozzle Handbook," General Electric Co. (A. R. Graham),

NAS9-3748, NASA-CR-II7OI8, Aug. 1965.

31. "An Experimental Model Investigation of a Plug Cluster Nozzle PropulsionSystem," Cornell Aeronautical Lab (K. C. Hendershot and R. J. Sergeant),Contract NAS8-823,, NAS8-20027, in APL Eighth Liq. Prop. Symp., Vol. I,pp. 293-334, Oct. 1966.

32. "Cold Flow Thrust and Pressure Investigations of a Plug Cluster NozzleUsing Short-Duration Techniques," Cornell Aeronautical Lab (R. K. Ellisonand K. C. Hendershot), Contract NAS8-823, NASA-CR-68956, CAL-HM-1510-Y-13,Apr. 1965.

33. "Base Pressure and Heating of Truncated Plug Rocket Exhaust Nozzle,"

Martin Co. (R. C. Schmidt), in APL Bull. of the 6th Liq. Prop. Symp.,Vol. I, pp. 537-53, Aug. 1964.

34. "Base Pressure Effects on the Thrust Performance of Unconventional Rocket

Nozzles," Douglas Aircraft Co. (G. W. Burge and D. W. Kendle), in APL Bull.of the 5th Liq. Prop. Symp., Vol. II, pp. 730-68, Dec. 1963.

35. "Study for Evaluation of Plug Multichamber Configuration," Phase I Pratt& Whitney Aircraft, Contract NAS8-I1436, FR-1415, 29 Oct 1965.

36. Ibid, Phase II, FR-2400, 31 Oct 1967.

289

fil!

SECTIONXl

REFERENCES

l °

"

,

.

°

.

°

_°

.

I0.

II.

12.

Baseline Space Tug Executive Summary, NASA/Marshall Space Flight Center,MSFC 68M00039-5, II October 1974.

Space Tug Systems Study (Cryogenic), September Data Dump, McDonnell-Douglas Astronautics Co., Contract NAS8-29677, Septen_)er 1973.

Wasko, R. A°, Performance of Annular Plug and Expansion-DeflectionNozzles Including External Flow Effects at Transonic Mach Numbers,NASA TN D-4462, April 1968.

Booz, D. E. and Schilling, M.T., Plug Cluster Nozzle Study, Pratt andWhitney Aircraft, Florida, PWA FR-IOI3. Contract NAS 8-11023,8 September 1964.

Bensky, M.S. Multichamber Plug-Nozzle Systems Analysis, Rocketdyne,Canoga Park. California, R-6494, Contract NAS 8-20237, May 1966.

Diem, H. G., Huang, D. H. and Wheeler, D.B., 02/H2 Advanced Maneuvering

Propulsion Technology Program Engine System Studies Final Report, Volume 1

Aerospike Engine Configuration Design and Analysis, Rocketdyne, CanogaPark, California, AFRPLLTR-72-4, Contract F04611-67-C-0116, December 1971

Monteath, E. B., Advanced Engine Aerospike Experimental Program ThrustChamber Performance, Volume 2, Final Report, Rocketdyne, Canoga Park,California_ R-7760, Contract NAS 8-20349, July 1969.

Gregory, J. W. and Herr, P.N., Hydrogen-Oxygen Space Shuttle ACPS Thruster

TechnolocLv Review, NASA/Lewis Research Center, AIAA Paper 72-I158,29 November" 1972.

LaBotz, R. J. and Blubaugh, A. L., Integrated Thruster Assembly Program,Aerojet Liquid Rocket Company, Sacramento, California, NASA CR-134509,Contract NAS 3-15850, November 1973.

Blubaugh, A. L. and Schoenman, L., Extended Temperature Rangs ACPSThruster Investigation, Aerojet Liquid Rocket Company, Sacramento,

California, NASA CR-134655, Contract NAS 3-16775, 1 August 1974.

Schoenman, L. and LaBotz, R. J., Hydrogen-Oxygen Auxiliary Propulsion forthe Space Shuttle, Volume I: High Pressure Thrusters, Aerojet LiquidRocket Company, Sacramento, California, NASA CR-120895, ContractNAS 3-14354, 30 January 1973.

Paster, R. D., Hydrogen-Oxygen APS Engines, Volume I_ High PressureThruster, Rocketdyne, Canoqa Park, California, NASA CR-120805, ContractNAS 3-14352, February 1973.

291

REFERENCES(Contd)

13.

14.

15.

16.

17.

18.

19.

20.

21.

22.

23.

Zachary, A.T., Csomor, A. and Tignac, L.L., Turbopumps for CryogenicUpper-Stage Engines, Rocketdyne, Canoga Park, California, R-9367,Contract NAS8-27794, December 1973.

Adams, A., Mayes, T.C. and Walter, J.D., Design Study of RLIDDerivatives, Final Report, Volume II, Engine Design Characteristics,Pratt & Whitney Aircraft, West Palm Beach, Florida, FR-6011, ContractNAS8-28989, 15 December 1973.

JANNAF Rocket Engine Performance Prediction and Evaluation Manual,Johns Hopkins University, Applied Physics Laboratory, Silver Spring,Maryland, CPIA Publication 246, April 1975.

williams, c.w., and Mellish, J.A., Space Tug Storable Engine Study,Volume II; Engine Design and Performance Characteristics, AerojetLiquid Rocket Company, Sacramento, California, DPD 396; DR-MA-03,Contract NAS8-29806, 31 January 1974.

Pieper, J. L. for ICRPG Performance Standardization Working Group,Aerojet General Corporation, Sacramento, California, CPIA PublicationNo. 178, 30 September 1968.

Calhoon, D.F., Ito, J.l. and Kors, D.L. Investigation of GaseousPropellant Combustion and Associated Injector/Chamber Design Guidelines,Aerojet Liquid Rocket Company, Sacramento, California, NASA-CR-121234,Contract NAS 3-14379, 31 July 1973.

Rao, G.V.R., Spike Nozzle Contour for Optimum Thrust, AFBMD/STL Symposiumon Advances in Ballistic Missile & Space Technology, 4th, Los Angeles,California, August 24-27, 1959, Planetary & Space Science, Volume 4,pp.92-101, Pergamon Press, 1961.

Johnson, G.R., Thompson, H.B. and Hoffman, J.D., Design of MaximumThrust Plug Nozzles with Variable Inlet Geometry, Vol. I, TheoreticalDevelopment and Results, Jet Propulsion Center, Purdue University,Lafayette, Indiana, AFRPL-TR-70-75, October 1970.

Liquid Rocket Engine Turbines, NASA Space Vehicle Design Criteria(Chemical Propulsion)_ NASA SP-SIIO, January 1974.

Turbopump Systems For Liquid Rocket Engines, NASA Space VehicleDesign Criteria (Chamical Propulsion), NASA SP-8107, August 1974.

Luscher, W.P. and Mellish, J. A._ Advanced High Pressure Engine Stud_for Mixed-Mode Vehicle Applications, Aerojet Liquid Rocket Company,NASA CR-135141_ Contract NAS 3-19727, January 1977.

292

REFERENCES(Contd)

24.

25.

26.

27.

28.

29.

0.

31.

32.

33.

34.

Rousar, D.C. and Ewen, R.L., Combustion Effects on Film Cooling,Aerojet Liquid Rocket Company, NASA CR-135052, Contract NAS 3-17813.February, 1977.

Sergeant, R.L., An Experimental Hot Model Investigation of a Plug ClusterNozzle Propulsion System, Part I: Base Thermal and Pressure Environmentfor a Module Chamber Pressure of 300 psia and Simulated Altitudes to150,000 Feet, Cornell Aeronautical Laboratory, Inc., CAL No. HM-2045-Y-5 (I),Figure 33, September 1967.

Rousar, D. and Miller, F., Cooling With Supercritical Oxygen, AerojetLiquid Rocket Company, AIAA Paper 75-1248, 29 September 1975.

Van Huff, NoE. and Rousar, D.C., Ultimate Heat Flux Limits of StorablePropellants, Aerojet General Corporation, 8th Liquid Propulsion Symposium,Volume II, 7 November 1966.

Giarratano, P. J., and Smith, R.V., Comparative Study of Forced ConventionBoiling Heat Transfer Correlations for Cryogenic Fluids, Advances inCryogenic Engineering, Vol. II, 492-506 (Equation 4), August 1965.

Tong, L.S., Boiling Heat Transfer and Two-Phase Flow, John Wiley & Sons,Inc., New York.

Davidson, W.F., Studies of Heat Transmission Through Boiler Tubing atPressures from 500-3300 Pounds, Transactions ASME, 1948.

Mellish, J.A., Space Tug Storable Engine Study, Aerojet liquid RocketCompany, Contract NAS 8-29806, DPD 396, 31 January 1974.

Luscher, W.P., Orbit-to-Orbit Shuttle Engine Design Study, AerojetLiquid Rocket Company, Contract F04611-71-C-0040, AFRPL TR-72-45,May 1972.

Diem, H.G., Burry, R.V. and Evans, S.A., 02/H2 Advanced ManeuveringPropulsion Technology Program, Engine System Studies Final Report,Volume II 25,000-Pound-Thrust Bell Engine Configuration Design andAnalysis, Rocketdyne, Canoga Park, California, AFRPL TR-72-4, ContractF04611-67-C-0116, December 1971.

Diem, H.G., Burry, R.V. and Evans, S.A., 02/H2 Advanced Maneuvering

Propulsion Technology Program, Engine System Studies Final Report,Volume Ill lO,O00-Pound-Thrust Bell Engine Configuration Design andAnalysis, Rocketdyne, AFRPL TR-72-4, Contract F04611-67-C-0116,January 1972.

293

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36.

37.

38.

39.

40.

43.

44.

45.

46.

REFERENCES (cont.)

Luscher, W. P., Orbit-to-Orbit Shuttle Engine Design Study, Book 3,Aerojet Liquid Rocket Company, AFRPL TR-72-45, Contract F04611-71-C-0040, May 1972.

Baseline Space Tug Configuration Definition, NASA/Marshall SpaceFlight Center, MSFC 68M00039-2, 15 July 1974.

Space Tug Systems Study (Cryogenic), Final Report, General Dynamics,Convair Aerospace Division, CASD-NAS73-033, NAS8-29676, January 1974.

Addendum I, Space Tug System Study Data Package, Propulsion SystemsControl Data, NASA/Marshall Space Flight Center, 17 April 1973.

Advanced Space Engine Requirements and Operating Conditions, NASA/Lewis Research Center, 1976.

Shubert, W. C., Thompson, P. S. Mayes, T. C. and Brown, J. R., Space

Tug Engine Performance Verification Program, Pratt & Whitney AircraftReport FR-7509, Contract NAS 8-31008, 14 May 1976.

Advanced Space Engine, Rockwell International Document LC332-977D.

Handbook of Supersonic Aerodynamics, Section 17, Ducts, Nozzles andDiffusers, NAVWEPSReport 1488 (Volume 6), January 1964, p. 368.

Heddon, K. and Loewe, A., Carbon, Volume 5, No. 4, 339-53 (1967); NASAMicrofiche N69-34048.

Lewis, J. C. A., Floyd, I. J. and Cowlard, F. C., Laboratory Investi-gation of Carbon-Gas Reactions of Relevance to Rocket-Nozzle Erosion,Paper No. 29, AC4_RDConf. Proc. 52 (1970); NASA Microfiche N70-24576.

Rousar, D. C., Plug Cluster Module Demonstration, Aerojet Liquid RocketCompany, NASA CR-135385, Contract NAS 3-20107, July 1978.

Kirby, F. M. Plug Cluster Nozzle Flow Evaluation, Final Report, Rocket-dyne Division, NASA CR-135229, Contract NAS 3-20104, to be published.

294

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