nasa · sequence of operation, hydrogen cool ed , gas generator cycle alternate mode 1 engine start...

NASA CR - 135141

(HIS&-CS-135141) bDV&NC?a HIGH PEdSSURB N77-23194 ESGXNS STUDY FOR N I X E C - U O D R VEHICLE I P P L I C B T X O N S Final Beport ( A e r o j e t L i q u i d socket Co.) 3 1 9 y HC A14/HF 901 CSCL 21H Unclas

G3/20 29365

N A S A \

ADVANCED HIGH PRESSURE ENGINE STUDY

FOR MIXED-MODE VEHICLE APPLICATIONS

F ina l Report

by

W. P. Luscher and 3 . A. M e l l i s h

AEROJET t IQU I D ROCKET COMPANY

Prepared f o r

NATIONAL AERONAUTICS AND SPACE ADNINISTRATION

NASA Lewis Research Center

Contract NAS 3-1 9727

https://ntrs.nasa.gov/search.jsp?R=19770016250 2020-03-31T18:00:16+00:00Z

- - 1. Report No. 2. Gavernnient Access~on No. 3. R e c ~ p ~ e n t ' s Catalog No.

NASA CR- 1351 41 4. T ~ l l e and Sublltle 15. Rcporl Date

Advanced High Pressure Engine Study f o r Mixed-Mode Vehic le Appl icat ions, F i na l Report

I 1 8. Perlorm~ng Organtzallon Report No.

W. P. Luscher and J. A. Me l l i sh I 9. Pedorm~ng Organ~zalton Name and Address

Aero je t L i q u i d Rocket Company - Post O f f i c e Box 13222 Sacramento, Ca l i f o rn i a 95813 NAS3-19727

I

15. Supplementary Notes

qency Name and Address

Nat ional Aeronautics and Space Admin is t ra t ion Washington, D. C . 20546

P ro j ec t Manager, D. D . Scheer , Chemi ca l Energy D i v i s ion , NASA-Lewis Research Center, Cleveland, Ohio

13. Type 01 Report and Per~od Covered

Contractor Reporr;, F i na l - 14. Swnsor~ng Agency Code

16. High pressure l i q u i d rocke t engine design, performance, weight, envelope and operat iona l c h a r a c t e r i s t i c s were evaluated f o r a v a r i e t y o f candidate engines f o r use i n mixed-mode, s ing le-s tage- to-orb i t appl i ca t i ons .

P rope l lan t proper ty and performance data were obtained f o r candidate Mode 1 f ue l s which included; RP-1, RJ-5, hydrazine, monomethyl -hydrazine, and methane. The common ox id i ze r was l i q u i d oxygen. Oxygen, the candidate Mode 1 f ue l s , and hydrogen were evaluated as t h r u s t chanber coo lants . Oxygen, methane and hydrogen were found t o be the n m t v i a b l e coo l i ng candidates. Water, 1 i t h i um and sodium-potassi um were a lso evaluated as auxi 1 i a r y coo lan t systems. Water proved t o be the bes t of these bu t the system was heavier than those systems which cooled w i t h t he engine prope l 1 ants.

Engine weight and envelope parametr ic data were estab l i shed f o r candidate Mode 1, Mode 2 and dua l - fue l engines. Del ivered engine performance data was a l so ca l cu l a ted f o r a l l candidate Mode 1 and dua l - fue l engine.

Engines c a r r i e d i n t o p r e l iminary design were; ( 1 ) LOX cooled, LOXIRP-1 Mode 1, staged combustior cyc le , (2) hydrogen cooled, LOXIRP-1 Mode 1, gas- generator cycle, and (3) dua l - fue l LOX cooled, LOXIRP-1 Mode 1/LH2 Mode 2, staged combustion cyc le .

---- --- 17. Key Words (Sgggested by Authcr(s\

Single-Stage-to-Orb i t M i xed-Mode Propuls ion Unc lass i f i ed - Un l im i ted High Pressure L i q u i d Rock c Engine Mode 1 and Mode 2 Dual -Fuel

- - 19. Secur~ty C l a s s ~ f . (ot lhls report) 20. Sec rr~ty Classlf . (of this page) -/%-NO. of Pages 122. Price'

Unc lass i f i ed I- Unc lass i f i ed 3 04

*For sa le by the Nat iona l Technical In format ion Service, Sp r i ng f i e l d , V i r g i n i a 22151

FOREWORD

The work described here in was performed a t the Aero je t L iqu id Rocket Company under NASA Contract NAS3-19727 w i t h M r . Dean D, Scheer, NASA-Lewis Research Center, as Pro jec t Manager. The ALRC Program Manager was M r . Werner P. Luscher and the Pro jec t Engineer was M r . Joseph A. Me11 ish .

The technical per iod performance f o r the study was from 30 June 1975 t o 30 June 1976.

The authors wish t o acknowledge the e f f o r t s of the fo l low ing ALRC engineering personnel who contr ibuted s i gni f i- cant ly t~ t h i s repor t :

R. E. Anderson L. L. B ick fo rd A. T. Caffo R. L. Ewen J. W. Hidahl M. C. Huppert G. R. Janser J. E. J e l l i s o n C. J. O'Brien J. W. Salmon D. H. Saltzman M. Short K. Y . Wong

We a lso wish t o thank M r . Rudi Beichel , ALRC Senior Sc ien t is t , f o r h i s comments and assistance throughout the study e f f o r t .

TABLE OF CONTENTS

Sect ion

I

I I

111

Page

1 SUMMARY

A, Study Object ives and Scope B. Results and Conclusions

INTRODUCTION

A. Background B. Purpose and Scope C. General Requirements D. Approach

TASK I - PROPELLANT PROPERTIES AND PERFORMANCE

A. Object ives and Guidel ines B . Physical and Thermodynamic Property Data C. Operational Considerations D. Combustion Gas Property Data E. Main Chamber Theoret ica l Performance Data

TASK I 1 - COOLANT EVALUATION

A. Object ives and Gui del i nes B . TCA Geometry and Mix ture Rat io Se lec t ion C. Low Cycle Fat igue Analysis D. Thermal Analysis

TASK I 1 1 - CYCLE EVALUATION

A. Object ives and Guidel i nes B. Task I 1 1 Del ivered Performance C. Engine Cycle Power Balances

TASK I V - ENGINE WEIGHT AND ENVELOPE

A. Object ive and Guidel ines B. Base1 i ne Engine Weight Statements C. Parametric Weight Data D. Parametric Envelope Data

V I I TASK V - AUXILIARY COOLANT FEASIBILITY

A. Object ives and Gui del i nes B. Auxi 1 i a r y Coolant Proper t ies

Section

I IP

V I I I -..

TABLE OF CONTENTS (cont. )

C. Mater ia ls Compat ib i l i ty D. Chamber Cool ing E. Engine Data

TASK V I - ENGINE PRELIMINARY DESIGN

A. Objectives and Guide1 ines 5. Ejos t Pump Dr ive Select ion C. Boost Pump NPSH Ef fec ts D. Engine Design and Operating Spec i f i ca t ions

Model 1 LOX/RP-1 Baseline Design

a. Conf igurat ion and Nominal Operating Conditions

b. Engine Operation and Control c. S t a r t and Shutdown Data d. Design and Off-Design Engine Performance e. Engine Mass Propert ies Data

Dual -Fuel Engine Design

a. Conf igurat ion and Nominal Operating Condit ions

b. Engine Operation and Control c. S t a r t and Shutdown Data d. Design and Off-Desi gn Engine Performance e. Engine Mass Propert ies Data

A l te rna te Mode 1 Engine Design

a. Conf igurat ion and Nominal Operating Condit ions

b. Engine Operation and Control c. S t a r t and Shutdown Data d. Design and Off-Design Engine Performance e. Engine Mass Proper t ies Data

E. Turbomachinery Design F. Thrust Chamber Design

1. Performance 2. Thermal Analysis

G. Main I n j e c t o r Design H. Preburner Design I. Valve Select ion and Siz ing J. Mater ia l Select ion

Paqe

1 54 156 1 58

163

163 166 170 171

178

1 78

178 186 186 186

191

191

193 203 2 03 2 03

21 0

21 0

21 0 220 220 220

224 247

247 257

265 2 68 282 287

TABLE OF CONTENTS (cont.)

Section

I X CONCLUSIONS AND RECOMMENDATIONS

A. Concl us i ons B . Recomnendations for Technology Work

REFERENCES

Page

292

LIST OF TABLES

'1'7bl e No. - Page

I

I I

I 1 1

I V

V

V I

V I I

V I I I

I X

X

X I X I I

XI11

X I V

xv xv I X V I I

X V I I I

X I X

X X

XX I

X X I I

X X I I I

X X I V

xxv X X V I

X X V I I

X X V I I I

Engine Pre l iminary Design Data Sumnary

Para1 1 e l Burn Mixed-Mode Vehicle D e f i n i t i o n

Series Burn Mixed-Mode Vehicle D e f i n i t i o n

Basel ine Engine D e f i n i t i o n s

Proper t ies o f Candidate Propel 1 ants

Vapor Pressure o f RJ-5

Densi ty o f RJ-5

V iscos i t y o f RJ-5

Heat Capacity and Thermal Conduc t i v i t y o f RJ-5

Extrapolated Heat Capaci t ies o f Methane

Extrapolated Densi t ies o f Methane

V iscos i t y o f Methane

Thermal Conduct iv i ty o f Methane

Propel 1 an t Cost Sumnary

LOX/RJ-5 Main Chamber Gas Property Data

LOX/RP-1 Main Chamber Gas Property Data

LOX/MMH Main Chamber Gas Property Data

LOX/N2H4 Main Chamber Gas Property Data

LOX/LH2 Main Chamber Gas Property Data

LOX/CH4 Main Chamber Gas Property Data

LOX/RJ-5 Oxidizer-Rich Preburner Gas Proper ty Data S umna ry

LOX/RP-1 Oxidizer-Rich Preburner Gas Property Data Sumnary

LOX/MMH Oxidizer-Rich Preburner Gas Property Data Sumnary

LOX/N2H4 Oxid i zer-Ri ch Preburner Gas Property Data Sumnary

LOX/LH2 Oxidizer-Rich Preburner Gas Property Data Sumnary

LOX/CH4 Oxi d i zer-Ri ch Preburner Gas Property Data Sumnary

LOX/MMH Fuel -Rich Preburner Gas Property Data Summary

LOX/N2H4 Fuel-Rich Preburner Gas Property Data Summary

LIST OF TABLES (cont.)

Table No.

X X I X

XXX

X X X I

X X X I I X X X I I I

X X X I V

xxxv X X X V I

xxxv I I X X Y V I I I X X X I X

XL

XL I

XLII

X L I I I

XLIV

XLV

XLV I

XLVI I

XLVI I 1

XLIX

L

L I

L I I

L I I I

LIV

LV

LOX/LH2 Fuel-Rich Preburner Gas Property Data S u m r y

LOXIRJ-5 %el-Rich Preburner Gas Property Data Sumnary

LOX/RP-1 Fuel-Rich Preburner Gas Property Data Sumnary

LOX/CH4 Fuel-Rich Preburnsr Gas Property Data S u m r y

Cool an t Eva1 uat ion Study Guide1 ines

Task I 1 Mode 1 Engine Geometry and Mixture Ratio Select ion Sumnary

Task I 1 Chamber Design L imi ts

Number o f Cool ant Channel s i n Combustion Chamber

Task I 1 Coolant Evaluation Sumnary

Maximum Bearing DN Values

Face Contact Seal Maximum PV, FV and PFV Factors

Task I11 Turbine I n l e t Temperature Selections

Task I11 Cycle Evaluation Summary

Task I 1 1 Candidate Mode 1 Engine Comparisons

Task I V - Engine Weight De f i n i t i ons

Task I V - Base1 i ne Engine Weight Statements

Candidate Mode 1 Engine Envelope Parametrics

Candidate Mode 1 Engine Envelope Parametri cs

Hydrogen Cooled , Gas Generator Cycle Engi ne Envelope Parametrics

Hydrogen Cooled , Gas Generator Cycl e Engi ne Envelope Parametri cs

Dual -Fuel Engine Envelope Parametri L;

Mode 2 Engine Envelope Parametri cs

Properties o f Candidate Auxi 1 i ary Cool ants

Density of L iqu id NaK (56)

Viscosi ty o f L iqu id NaK (56)

Summary o f Task V Chamber Designs

Task V Aux i l i a r y Cooled Engine Sumnary

v i i i

I

Table No.

LV I

L'JII

LVI I I

LIX

LX

LXI

LXI I

LXI I I LXIV

LXV

LXVI

LXVI I

LXVI I I

LXIX

L XX

LXXI

LXXI I

LXXI I I

LXXIV

LXXV

LXXVI

LXXVI I

LIST OF TABLES (cont.

Mode 1 LOX/RP-1 Engi ne Operati ng Speci f i cations, Nominal MR=2.9

Mode 1 LOXIRP-1 Base1 i ne Engi ne Pressure Schedule, Nominal MR=2.9

Sequence o f Operations, Base1 i ne LOXIRP-1

Mode 1 LOXIRP-1 Base1 ine S t a r t and Shutdown Transient Data Summary

Mode 1 LOXIRP-1 Baseline Engine Design and Off-Design MR Performance

Mode 1 LOXIRP-1 Base1 i ne Engi ne Wei ght Statement

Dual -Fuel Nominal Engine Operating Speci f icat ions

Dual -Fuel , Mode 1 Nomi nal Engine Pressure Schedule

Dual -Fuel , Mode 2 Nominal Engine Pressure Schedule

Sequence o f Operation Dual -Fuel Engine

Dual-Fuel Engine S t a r t and Shutdown Transient Data Summary

Dual-Fuel, Mode 1 Design and Off-Design MR Performance

Dual-Fuel, Mode 2 Design and Off-Design MR Performance

Dual -Fuel Engine Weight Statement

A1 ternate Mode 1 Hydrogen Cooled, Gas Generator Cycle Engine Operating Speci f icat ions

A1 ternate Mode 1 Hydrogen Cooled, Gas Generator Cycle Engine Pressure Schedule

Sequence o f Operation, Hydrogen Cool ed , Gas Generator Cycle

Al ternate Mode 1 Engine S t a r t and Shutdown Transient Data Summary

A1 ternate Mode 1 Hydrogen Cooled, Gas Generator Cycle Engine Design and Off-Design Performance

A1 ternate Mode 1 Hydrogen Cooled, Gas Generator Cycl e Engine Weight Statement

Base1 i ne Mode 1 and Dual -Fuel Engine Low Speed LOX TPA Design Point

Base1 i ne Mode 1 and Dual -Fuel Engine High Speed LOX TPA Design Point

Page

182

185

187

188

'189

190

196

201

202

204

206

207

208

209

21 4

21 8

21 9

221

222

22 3

226

228

LIST OF TABLES (cont .)

Table No. Page

233 LXXVI I I Base1 i ne Mcde 1 and Dual -Fuel Engi ne Low Speed RP-1 ' TPA Design Point

LXXIX Basel ine Mode 1 and Dual-Fuel Engine High Speed RP-1 TPA Design Point

LXXX Dual-Fuel Engine Low Speed LH2 TPA Design Point

LXXXI Dual-Fuel Engine High Speed LH2 TPA Design Point

LXXXI I Alternate Mode 1 Engine High speed LOX TPA Design Point

LXXXI I I A1 ternate Mode 1 Engine High Speed RP-1 TPA Design Point

LXXXIV A1 ternate Mode 1 Engine Low Speed LH2 TPA Design Point

LXXXV A1 ternate Mode 1 Engine High Speed LH2 TPA Design Point

LXXXVI Dual -Fuel Nozzl e Contour Eva1 uatia.7

LXXXVI I Mode 1 Base1 i ne Del i vered Speci f i c Impul se

LXXXVIII Mode 1 Gas Generator Cycle Delivered Speci f ic Impulse

LXXXIX Mode 1 Dual -Fuel Del i vered Speci f ic Impulse

XC Mode 2 Dual -Fuel Del i vered Speci f ic Impul se

XC I Cool ant Jacket Sunmiary

X C I I Mode 1 LOXIRP-1 Base1 ine and Dual- Fuel Engine Main I n j e c t o r Design Data

X C I I I A1 ternate Mode 1 Gas Generator Cycle Engine Main I n j e c t o r Design Data

X C I V Gxidizer-Rich Preburner Desi grl Data Summary, LOX/RP-1 Base1 i ne

X C J Fuel -Rich Preburner Design Data Sumnary , LOX/RP-1 Easel i ne

XCV I 02/H2 Oxi d i zer-Ri ch Preburner Design Data S u m r y , Dual -Fuel

X C V I I H7/C2 Fuel -Rich Preburner Design Data Summary, ~ t a l -Fuel

X C V I I I A1 ternate Mode 1 Co-Axial Gas Generator Design Data Summary

Table No.

X C V I X

C

C I

C I I

C I I I

C I V

C v cv I

LIST OF TABLES (cont. )

Valve Sizing: Mode 1 LOX Cooled Base1 i n e

Val ve S i z i ng : Hydrogen Cool ed , Gas Generator Cycl e

Valve S i i i ng : Dual -Fuel Engine

Hot Gas Check Valve Weight and Envelope

Estimated Power Requirements f o r Electromechanical Valve Actuation Mater ia ls Select ion

Conclusions

Techno1 ogy Recommendations

LIST OF FIGURES

Figure No.

Advanced Hf gh Pressure Engine Study Program Summary

Propel I an t Cost Pro ject ions

LOX/RP-1 Experimental Preburner Charac te r i s t i cs

Emp i r i ca l l y Derived LOX/RP-1 Fuel -Rich Gas Proper t ies

LOXIRP-1 ODE and Frozen Performance Comparison

LOX/RJ-5 ODE Vacuum Performance Versus O/F

LOXIRP-1 ODE iacuum Performance Versus O/F

LOXIMMH ODE Vacuum Performance Versus O/F

LOX/N2H4 ODE Vacuum Performance Versus OIF

LOX/LH2 ODE Vacuum Performance Versus O/F

LOX/CH4 ODE Vacuum Performance Versus O/F

LOXIRJ-5 ODE Performance Versus Area Rat io

LOXIRP-1 ODE Performance Versus Area Rat io

LOX/WH ODE Performance Versus Area Rat io

LOY/N2H4 ODE Performance Versus Area Rat io

LOX/LH2 ODE Performance Versus Area Rat io

LOX/CH4 ODE Performance Versus Area Rat io

Tensi le Proper t ies (Z i r con i um Copper)

Tens i le S t ress -S t ra in

Creep-Rupture and Low Cycle Fatigue

C0nduc.i.i v i t y and Expansion

Shear Coaxial Element Performance

E f f e c t o f Chamber Contract ion Rat io on Energy Release E f f i c i e n c y

E f f e c t o f Chamber Contract ion Rat io on Combustion Pressure Loss

Task I I LOX/RJ-5 and LOXIRP-1 Del i vered Vacuum Performance vs Mix ture Rat io

Task I I LOXIMMH and LOX/N2H4 Del ivered Vacuum Performance vs Mix ture Rat io

Page

1

2 9

5 1

56

58

59

6 0

61

62

63

64

65

66

67

68

69

70

7 3

74

7 5

76

78

78

78

80

81

Figure No.

2 7

LIST OF FIGURES (cont . )

Page - Task I 1 LOXlCH4 Del ivered Vacuum Performarse vs 82 Mix tu re Ra t io

Low Cycle Fat igue L i f e Requirement 8 5

Oxygen Heat Transfer C o e f f i c i e n t Va r i a t i on With Wall 9 3 Temperature

Coolant Pressure Drop With Oxygen Cool ing 95

Coolant Pressure Drop With Hydrogen Cool i n g 9 6

Coolant Pressure Drop With Hydrazine Compounds 98

Cool an t Pressure Drop W i t h Methane Cool i ng 39

Task I 1 1 Mode 1 Oxygen Cooled Engine Cycle Schematic 103

Task I I I Mode 1 Fuel Cooled Engine Cycle Schematic 1 04

Task I 1 1 Para1 l e l Burn, Hydrogen Cooled Mode 1 Englne 105 Cycl e Schema t f c

Task I 1 1 Mode 1 Hydrogen Cooled Gas Generator Cycle 136 Schematic

Task I 1 1 Dual-Fuel, Oxygen Cooled Engine Cycle Schematic 107

LOXIRJ-5 Staged Combustion Del ivered Performnce vs 113 Area Rat io

LOXIRP-1 Staged Combustion Del i vered Performance vs 114 Area Rat io

LOXIMMH Staged Combustion Del ivered Performance vs : I 5 Area Rat io

LOXIN2H4 Staged Combustion Del i vered Performance vs 11 6 Area Rat io

02/H2 Staged Combustion Del ivered Performance vs 117 Area Rat io

LOX/CH4 Staged Combustion Del i vered Performance vs 118 Area Rat io

Maximum Pump Discharge Pressure Requirements fo r 120 Candidate Mode 1 Engines

Mode 1 Staged Combustion Cycle Engine Weiqht Parametrics 128

E f f e c t o f Area Rat io Upon Mode 1 Staged Combustion Cycle 129 Engine Weight

LIST OF FIGURES (cont.)

Figure No.

Mode 1 Hydrogen Cooled Gas Generator Cycle Engine Weight Parametrics

E f Fect o f Area Ratio Upon Mode 1 Hydrogen Cooled Gas Generator Cycle Engine Weight

Dual-Fuel Engine Weight Parametrics

Mode 1 LOX/CHq Engine Weight Parametrics

Mode 2 LOX/LH2 Engine Weight Parametrics

E f fec t o f Area Ratio Upon Mode 2 LOX/LH2 Engine Weight

Mode 1 Auxi l i a r y Cooled Engine Cycle Schematic

High Speed Pump Configurat ion Concept f o r Tap-Off o f Hydraulic Turbine Drive F l u i d

Head Coef f i c ien t vs Speci f ic Speed

E f fec t o f NPSH on Low Speed RP-1 Turbopump Weight

E f fec t o f NPSH on Low Speed LOX Turbopump Weight

E f fec t o f NPSH on Low Spec ! LH2 Turbopump Weight

E f fec t of NPSH on Low Speed LOX & RP-1 Pump Impel ler Diameters

E f fec t o f NPSH on Low Speed LH2 Pump Impel ler Diameters

Base1 ine LOX/RP-1 Mode 1 Engine Cycle Schematic

Mode 1 LOX/RP-1 Base1 ine Engine Assembly (Top View)

Mode 1 LOX/RP-1 Base1 ine Engine Assembly (Side View)

Dual -Fuel , Oxygen Cooled Engine Cycl e Schematic

Dual -Fuel Engine Assembly (Top View)

Dual-Fuel Engine Assembly (Side View)

Mode 1 Hydrogen Cooled, Gas Generator Cycle Schematic

Mode 1 Hydrogen Cooled, Gas Generator Cycle Engine Assembly (Top View)

Mode 1 Hydrogen Cooled, Gas Generator Cycle Engine Assembly (Side View)

Base1 i ne Mode 1, Dual -Fuel and A1 te rna te Mode 1 Low Speed LOX TPA

Base1 ine Mode 1 and Dual-Fuel High Speed LOX TPA

Page

130

132

133

134

136

137

147

169

172

173

174

175

176

177

179

180

181

192

1 94

195

21 1

21 2

21 3

225

227

LIST OF FIGURES (cont.)

Figure Wo . ,Basel i n e Mode 1 , Dual -Fuel and A1 ternate Mode 1 Low Speed RP-1 TPA

Base1 ine Mode 1 and Dual-Fuel High Speed RP-1 TPA

Dual -Fuel Low Speed LH2 TPA

Dual -Fuel High Speed LH2 TPA

Al ternate Mode 1 High Speed LOX TPA

A1 ternate Mode 1 High Speed RP-1 TPA

Al ternate Mode 1 Low Speed LH2 TPA

A1 ternate Mode 1 High Speed LH2 TPA

ZrCu S lo t ted Chamber L i f e L imi ts

ZrCu S lo t ted Chamber Pressure and Temperature Wall Thickness L im i t s

Mode 1 LOXIRP-1 Baseline and Dual-Fuel Thrust Chamber I n j e c t o r

Mode 1 LOXIRP-1 Baseline Backup Main I n j e c t o r

A1 ternate Mode 1 Gas Generator Cycle Engine Thrust Chamber I n j e c t o r

Mode 1 LOXIRP-1 Base1 i ne Oxidizer-Rich Preburner

Mode 1 LOXIRP-1 Base1 ine Fuel -Rich Preburner

021H2 Oxidi zer-Rich Preburner, Dual -Fuel

H2/02 Fuel -Rich Preburner , Dual -Fuel

A1 ternate Mode 1 Co-Axial Gas Generator

Page

2 30

231

236

2 38 242

244

246 24 9

258

259

266

269

270

272

273

276

277

280

SECTION I SUMMARY

A. STUDY OBJECTIVES AND SCOPE

The major object ives o f t h i s study program were t o provide parametric data, pre l iminary designs and I d e n t i f y technology requirements o f advanced high pressure engines f o r use i n mixed-mode single-stage-to-orbi t vehicles.

The two sets o f l i q u i d rocket engines used as the propulsion f o r the mixed-mode vehic le concept are ca l l ed Mode 1 and Mode 2. Mode 1 engines use re1 a t i v e l y low performance, high densi ty propel 1 ants wh i le the Mode 2 engines use high performance, low density oxygenlhydrogen propel 1 ants. A t h i r d . erigine type, ca l l ed the dual-fuel , uses both high density fuel and low density hydrogen sequent ia l ly i n a common t h r u s t chamber. This study was p r imar i l y concerned w i th the Mode 1 and dual - fue l engines.

To acxompl i sh the object ives, the s i x task study program sumnarized on Figure 1 was conducted.

Task I -- Propel 1 ant Propert ies and Performance

Cool ant Evaluation

Task I11

Cycle Evaluati on--3

Task I V 0

Englne Weight and 4 Envel ope

Task V

Review

Baseline Fuel Select ion

Candidate Mode 1 Engine Select ion

Task V I

Prel iminary Engine Designs

Baseline Mode 1

Dual -Fuel

A1 ternate Mode 1

Auxi 1 i ary Cool ant Feasi b i 1 i t y

TECHNOLOGY REQUIREMENTS

Figure 1. Advanced High Pressure Engine Study Program Summary

High pressure engine design, performance, wei ght , envelope and oper- a t i ona l cha rac te r i s t i c s were evaluated f o r a v a r i e t y o f candidate engines.

Propel1 an t proper ty , performance and c h a r a c t e r i s t i c data were obtained f o candidate Mode 1 f ue l s which included; kerosene (RP-1) , A SHELLDYNE-H (RJ-5) , hydrazi ne (N2H4) , monomethyly-hydrazi ne (MMH) , and methane ( C H ~ ) . L i q u i d oqlgen was the comncn ox id i ze r .

Oxygen, the candidate Mode 1 fue ls , and hydrogen were evaluated as t h r u s t chamber coolants f o r the Mode 1 engines. The dual - f ue l engine was oxygen cooled and the basel ine Mode 2 engine was a hydrogen cooled con- f i g u r a t i o n s i m i l a r t o the Space Shu t t l e Main Engine (SSME) . Water, l i t h i um, and sodium potassium (NaK 562) were a lso evaluated as a u x i l i a r y systems f o r coo l ing the chamber.

Engine weight and envelope parametr ic data were es tab1 i shed f o r the candidate Mode 1 engine, the oxygenlhydrogen Mode 2 engine and the dual - f u e l engine. Eel i ve red engine s p e c i f i c impulse data were a l so ca lcu la ted a t the most p r a c t i c a l t h r u s t chamber pressure f o r a l l candidate Mode 1 and the dual - f u e l engines.

Based upon the r e s u l t s o f Task I through V and the study gu ide l ines, three en ines were c a r r i e d i n t o a p r e l iminary design phase. These engines were; ( 1 7 base1 i n e Mode 1 , LOX cooled , LOXIRP-1, staged combusti on cyc le , (2) dual - fue l , LOX cooled, LOXIRP-1 (Mode 1) and LH2 (Mode 2) , staged combustion cycle, and (3) a1 te rna te Mode 1, hydrogen cooled, LOXIRP-1, gas generator cyc le .

Throughout the e n t i r e study e f f o r t , bas ic data saps m d areas r e q u i r - - . i n g technology work were i d e n t i f i e d .

B. RESULTS AWD CONCLUSIONS

Oxygen, methane and hydrogen were found t o be t he most v iab dates f o r cool i n g the Mode 1 engine. RP-1 and RJ-5 are un feas ib le

l e cand f o r

coc l i n g a t chamber pressures above 136 atmospheres (2,000 psis) because o

proved t o be the best a u x i l i a r y coolant bu t the system us ing one of the main p rope l lan ts as a coolant.

The Task I through V review resu l t ed i n the f o l

" Change the basel ine f u e l from RJ-5 t o RP-1.

f coking and gumming i n t he coolant channel s. ' For 1 ong-term' use; N2H4 and MMH a re incompat ible w i t h z irconium copper, ZrCu, which was t h e spec i f i ed t h r u s t chamber mater ia l f o r the study e f f o r t . I n add i t ion , t he f u e l - r i c h combustion products of LOXIMMH and LOXlN2Hq c rea te a t u rb i ne design 1 i f e problem because t he combust^ on temperatures are i n excess o f 1 OOO°F (1800°R). Water

was heavier than those

lowing major decis ions:

' ' ie lec t the hydrogen cooled, gas generator cyc l e engine as t h e idat:? lode 1 f o r p re l im ina ry design.

"Include methane (CHq) i n the f i r s t f o u r study tasks.

Current RJ-5 p rope l l an t costs are 4.41 $/Kg (2.00 $ / ~ b ) . This cos t appeared t o be p r o b i b i t i v e and RP-1 was se lected as the base l ine Mode 1 f u e l foe the p re l im ina ry engine designs. However, the h i ghe r dens i t y o f RJ-S s t i l l makes t h i s f u e l a t t r a c t i v e f o r use i f the cos t goes down when purchased i n l a rge quan t i t i e s .

T b cu r ren t cost o f CH4 i s s i m i l a r t o t h a t o f RP-1, 0 . 1 ~ ? $,'Kg (0.36 $/L, l) and o f f e r s about 10 secs. greater s p e c i f i c impulse f o r an engine we i f h t penal ty o f approximately 218 Kg (480 Lb). Because of t he l a t e e n t r y o f t ~ i e CHd i n t o the study, a p re l im ina ry design o f a LOX/CH engine could d n o t be incorporated w i t h i n the remaining study t ime span. owever, because of the po ten t i a1 bene f i t s p rev ious ly mentioned, the LOX/CH4 engine should be s tud ied f u r t h e r i n a p re l im ina ry design phase.

The hydrogen cooled, gas generator cyc l e engine was se lec ted as the candid ~ t e Node 1 engine becauze p r e l iminary t r a d e - o f f s tud ies showed potent7 a1 i inprovements i n e i t h e r veh i c l e gross l i f t - o f f weight o r d r y weight when compared t o the basel ine.

The engine design data summary f o r t he th ree engines c a r r i e d i n t o the pre:irrt inary design phase i s shown on Table I.

Sup;)orti ng research and techno1 ogy programs are recommended t o f i 11 bas ic data gaps o r prov ide c r i t i c a l i n fo rmat ion p r i o r t o en te r i ng a h igh pressure, Mode 1 o r dua l - fue l engine development. These programs are requi red t o v e r i f y p rope l l an t and combustion gas p roper ty data a t h i gh pressure, prov ide heat t r a n s f e r in fo rmat ion on l i q u i d oxygen as a coolant, ob ta in engirle and component l i f e data and t o v e r i f y engine performance a t h igh t h r u s t cllamber pressure.

TABL

E I. -

ENG

INE

PREL

IMIN

ARY

DES

IGN

DAT

A SU

MM

ARY

I B

ase1

ine

I

Mod

e 1

Mai

n P

rop

el 1

an

ts

Oxi

diz

er

Fue

l

Coo

lant

'

Cyc

le

Thr

ust,

N

(Lb

) se

a-~

eve

l Va

cuum

b~

ixe

d e = 40 N

ozzl

e

Oxy

gen

RP-

1

O~

ge

n

Sta

ged

Com

bus t

i on

2 .

~o

XI

O~

~

(607

,000

) 2

.92

6~

10

(6

57.7

00)

Spe

ci fi c

Im

puls

e,

sec.

S

ea-L

evel

Va

cuum

T

hru

st C

ham

ber

Pre

ssur

e,

Atm

. (p

sia

)

Eng

ine

Mix

ture

Rat

io,

O/F

Noz

zle

Are

a R

ati

o,

c

LOX

Flo

w R

ate,

K

g/se

c (l

b/s

ec)

RP-

I F

low

Rat

e,

Kg/

sec

(Ib

/se

c)

H2

Flo

w R

ate,

K

g/se

c (l

b/s

ec)

Eng

ine

Dry

Wei

ght,

Kg

(lb

)

Eng

ine

Leng

th,

CM (

in.)

Max

imum

Dia

met

er,

CM (

in.)

ENG

INE

CON

A1 te

rna

te

Mod

e 1

323.

6 35

0.6

272

(400

0)

2.9

40

632

-7 i 13

94.8

)

218.

2 (4

81 -0

) --

21 1

2 (4

657)

277

(103

)

224

(88

jd

'~x

ten

ible

Noz

zle,

R

etra

cted

/Dep

loye

d

d~ow

erhe

ad

eMod

e 2

Noz

zle

Ex

it D

ia.

Oxy

gen

RP-

1

Hyd

roge

n ---

Gas

G

ener

a to

r

2 .7

0xlo

6 (6

07,0

00)

2 .9

27xl

o6

(658

,000

)

323.

5 35

0.7

IGU

RAT

ION

Dua

l -F

uel

--

"!?&

-L

I Mode 2

Oxy

gen

RP-

1 O

xyge

n LH

Z

Oxy

gen

Sta

ged

Com

bus t

i on

Ow

gen

Sta

ged

Com

bus t

i on

SECTION I I INTRODUCTION

A, BACKGROUND

The NASA i s cur ren t ly conducting studies o f advanced recoverable vehic le concepts fo r the post 1990 time per iod i n order t o i d e n t i f y technology needs and provide guidance i n agency planning. One o f the concepts under study i s a v e r t i c a l takeoff, hor izonta l landing (VTOHL) single-stage-to-orbi t (SSTO) vehic le which u t i l i z e s mixed-mode ascent propulsi on. M i xed-mode propulsion involves the sequenti a1 o r para1 l e l use o f high densi ty impulse propel lants and h igh spec i f i c impulse propel- lan ts i n a s ing le stage t o increase vehic le performance and reduce vehic le weight. The concept has been described i n the 1 i tera ture (e.g., "Reusable One-Stage-to-Orbit Shutt les: Brightening Prospects" by Robert Salkeld and Rudi Beichel , Astronautics and Aeronautics, June, 1973) and theo- r e t i c a l l y provides the performance necessary f o r a VTOHL-SSTO Space Trans- p o r t a t i on Vehicle.

The propulsion system f o r the veh ic le concepts under study consists o f two sets of engines (herein ca l l ed Mode 1 and Mode 2 engines) which u t i l i z e oxygen as a comnon ox id izer . Mode 1 engines burn a h igh bu lk density f u e l during l i f t - o f f and ea r l y ascent t o minimize the performance penalty associated w i t h car ry ing the weight o f Mode 1 f u e l tanks t o o r b i t , Mode 2 engines burn hydrogen and complete the t r a j e c t o r y t o o r b i t . I n some concepts the Mode 1 and Mode 2 engines are burned sequent ia l ly (ser ies burn concept) wh i le other concepts have both the Mode 1 and Mode 2 engines i n operat ion dur ing l i f t - o f f and ea r l y ascent ( p a r a l l e l burn concept). An engine va r ia t i on introduced i n the previously c i t e d reference in tegrates hydrogen turbopump system hardware w i t h the basic oxygen-hi gh bulk densi ty f ue l engine thus permi t t ing sequential burn o f both fue l s w i t h the comnon ox id izer i n an engine which shares common components such as the t h r u s t chanber (dual- fuel engine). I n a l l o f the concepts, the engines must be 1 ightweight and operate a t high chamber pressure t o permit high expansion area r a t i o w i t h i n the confine o f the VTOHL vehic le base area.

B. PURPOSE AND SCOPE

The feasi b i 14 t y o f the mi xed-mode single-stage-to-orbi t vehic le i s heavi l y dependent on the del i vered performance o f the engine systems . I t was the purpose o f t h i s e f f o r t t o a n a l y t i c a l l y determine the performance o f candidate engine systems and t o i d e n t i f y technology needs i n the propul s i on area.

Parametric studies and engine pre l iminary designs were conducted based upon cu r ren t l y achievable component performance 1 eve1 s and cu r ren t l y ava i lab le mater ials.

C. GENERAL REQUIREMENTS

For purposes o f t h i s study, a p a r a l l e l burn vehic le w i th the requirements and operat ing condit ions given i n Table I 1 and a ser ies b u m vehic le w i t h the requirements and operat ing condit ions given i n Table I 1 1 were assumed as basel ine vehicle concepts. Basic requirements f o r the basel ine Mode 1, Mode 2, and dual- fuel engines, consistent w i th the basel ine vehicles, are given i n Table I V .

The i n i t i a l basel ine Mode 1 engine f o r both the p a r a l l e l bum vehic le and the series burn vehic le was a f i x e d 40:1 area r a t i o 90% b e l l nozzle, staged-combustion turb ine d r i ve cyc le engine u t i l i z i n g RJ-S/oxygen as propel 1 ants. The t h r u s t chamber i s regenerat i ve ly cooled w i t h oxygen,

The base1 i ne Mode 2 engine f o r the para1 l e l burn veh ic le i s a staged combustion cyc le hydrogen-oxygen engine w i t h a f i x e d 40:l area r a t i o 90% be:l nozzle f o r sea l eve l operat ion and an extendible 200:l area r a t i o 90% b e l l nozzle f o r vacuum operation. The chamber i s regenerat ive ly cooled w i t h hydrogen.

The basel ine dual- fuel engine f o r the ser ies burn vehic le i s a basic Mode 1 engine which incorporates those design features and components necessary t o provide capabi 1 i t y f o r sequenti a1 i n - f l i ght operation as a hydrogen-oxygen Mode 2 engine w i th a 200:l area r a t i o 90% b e l l nozzle. The chamber i s regenerat ively cooled w i th oxygen dur ing both operational modes.

D. APPROACH

To accomplish the program object ives , a study i nvo l v ing s i x technical tasks was conducted. The resu l t s o f the f i r s t f i v e tasks were reviewed t o se lec t a basel ine fue l , update the basel ine engine requirements and performance, and t o se lec t a Mode 1 engine concept, i n add i t ion t o the basel ine Mode 1 and dual- fuel , f o r p re l iminary design. Tasks conducted were:

1. TASK I: PROPELLANT PROPERTIES AND PERFORMANCE

Property data and charac ter is t i cs o f various candidate propel - lan ts were obtained and establ ished as a r e s u l t o f t h i s task.

Oxygen i s assumed as the common ox id i ze r f o r the Mode 1 engines and the hydrogen-fu l e d Mode 2 angines. Candidate Mode 1 fue l s are kerosene k (!?O-1), SHEELDYNE-H (RJ-5), hydrazine, monomethyl-hydrazine (MMH) and methane (CH4) .

The data includes:

a. Physical and thermodynamic property data f o r the candidate fuels .

b. Log is t i cs and safety aspects such as pro jected avai l a b i 1 i t y and cost, hand1 ing, chemical and thermal s tab i li t y , materi a1 compat ib i l i t y , corros i v i t y and t o x i c i ty .

TABLE 11. - PARALLEL BURN MIXED-WDE VEHICLE DEFINITION

Gross Weight: 1,448,400 kg (3,193,200 l b .)

Propul s i on Sys tern

Mode 1

No. of Engines Vacuum Speci f i c Impul se , Sec. Vacuum Level Thrust per Engine, N ( l b ) Sea Level Thrust per Engine, N ( l b ) Ideal Velocity , m/sec ( f t l sec ) Fuel Density, kg/m3 ( l b / f t 3 ) Oxidizer Densi t y , k g / d ( l b l f t 3 ) M i xture Rati o Area Ratio

Mode 2

No. o f Engines Vacuum Speci f i c Impul se , Sec . Vacuum Level Thrust per Engine, N ( l b ) Sea Level Thrust per Engine, N ( Ib) Ideal Velocity , m/sec ( f t /sec) Fuel Densi t y , k g / d (1 b / f t3) Oxidizer Densi t y , kg/m3 ( 1 b / f t3 ) Mixture Ratio Area Rati o (Sea Level /Vacuum)

TABLE 111. - SERIES BURN MIXED-MODE VEHICLE DEFINITION (DUAL-FUEL ENGINE)

Gross Weight: 1,598,500 kg (3,524,000 1b .)

Propulsion Sys tern

Mode 1

No. o f Engines Vacuum Specif ic Impulse , sec . Vacuum Level Thrust per Engine, N ( lb ) Sea Level Thrust per Engine, N ( lb ) Ideal Velocity , m/lec (f t / jec) Fuel Density, kg/m ( l b / f t ) Oxidizer Density , kg/m3 (1 b/f t3) Mixture Ratio Area Ratio

Mode 2

No. o f Engines Vacuum Speci f i c Impulse , sec . Vacuum Level Thrust per Engine, N (1 b) Ideal Velocity, m/ ec ( f t l sec ) 3 Fuel Density, kg/m ( l b l f t 3 ) Oxidizer Densi ty, k g / d ( l b l f t 3 ) Mixture Ratio Area Ratio (vacuum)

Pro

pel 1

ants

:

Oxi

diz

er

Fue

l M

ixtu

re R

ati

o,

(O/F

) Se

a Le

vel

Thr

ust,

N (I

b)

Vacu

um T

hrus

t,

N (l

b)

Vacu

um I

mpu

lse,

Se

c.

Dri

ve C

ycle

N

ozzl

e Ty

pe

Noz

zle

Exp

ansi

on R

ati

o

Pro

pe

llan

t In

let

Tem

p.:

Oxi

diz

er

OK(O

R F

uel,

OK

(O

R)

NPSH

@ E

ngin

e In

let:

Oxi

d<ze

r, m

(ft

) F

uel,

m (

ft)

Cha

nber

Pre

ssur

e,

Atm

. (p

sia

) S

ervi

ce L

ife

Bet

wee

n O

verh

auls

Gim

al

hg

le (

Squ

are

Pa

tte

rn),

D

egre

es

Pa

rall

el

Bur

n S

eri

er B

urn.

D

ual -F

uel

Mod

e 1

I M

ode

2 I

me

1

345.

6 S

tage

d Co

mh.

90%

Be

ll

40

90.3

(1

62.5

) 30

0 (5

40)

272

(4,0

00)

250

Cyc

les

or

20

Hou

rs

Run

Tim

e t

10

Hyd

roge

n 7

.O

466.

5 S

tage

d Co

nb.

90%

Be

ll

40

SL/

200

Vac

204

(3,0

00)

250

Cyc

les

or

20 H

ours

Ru

n Ti

me

f 7

345.

6 S

tage

d C

d.

90

% B

ell

40

90.3

(1

62.5

) 30

0 (5

40)

14.0

46

27

2 (4

,000

) 25

0 C

ycle

s o

r 20

Hou

rs

~u

n

Tim

I M

ode

2

Oxy

gen

Hyd

roge

n 7

-0

;:29x

106

(515

,250

46

6.5

Sta

ged Cd.

90%

Be

ll

200

204

(3,0

00)

250

Cyc

les

or

20 H

ours

R

un T

ime

c , Main chamber and prebumer propel l a n t conbination conbus ti on gas thermdynarni c and transport property parametric data.

d. Main chamber theoretical rocket performance parametric data.

2. TASK 11: COOLANT EVALUATION

The re l a t i ve coolant capabi 1 i ty of candidate Mode 1 fuels, owgen and hydrogen were determined. Concepts considered are:

Propel lant Combination Cool ant

RJ-SIOxygen O V 9en RJ-5/0~tgen RJ-5 RJ -5/Oxygen Hydrogen RP-1 /Oxygen RP -1 Hydrazi ne/Owgen Hydrazi ne WIOwgen MMH CH4lOwgen CH4

The re l a t i ve mer i t o f the various coolants i s established on the basis o f the attainable thrust chanber pressure, as re f lec ted i n the coolant pressure drop, wi th considerati on of potenti a1 propel 1 ant property problems and l imi tat ions.

3. TASK 111: CYCLE EVALUATION

Engine cycle pressures, temperatures and del i vered performance f o r various candidate Mode 1 engine concepts were established.' Concepts are:

Main Chamber In jec tor Propel l a n t Conbination Cool ant Description

RJ -5lOxygen RJ-5/0xygen RJ-510xygen

RJ-5lOxygen

RP- 1 /Oxygen Hydra r i ne/Oxygen MMH/O>(ygen CHqIOwgen

Oxygen RJ-5 Hydrogen

Hydrogen

RP-1 Hydrazi ne MMH CH4

Gas-Gas , Staged Conbusti on Cycle Gas-Gas , Staged Conbusti on Cycle Gas-Gas , Staged Combusti on Cycle, Paral le l Bum H 102 Liquid-Liquid, s Generator (H2/02) Cycle

L Gas-Gas, Staged Conburtion Cycle Gas-Gas, Staged Conbustion Cycle &-Gas, Staged Conbustion Cycle Gas-Gas , Stdged Combus ti on Cycle

The candidate Mode 1 concepts were studied over a pract ica l , achievable chamber pressure range.

4. TASK I V : ENGINE WEIGHT AND ENVELOPE

Weight and envelope estimates f o r candidate Mode 1 engine concepts and the baseline Mode 2 engine concept were established over ranges o f thrust,

thrust chamber pressure and nozzle area rat ios.

5. TASK V: AUXILIARY COOLANT FEASIBILITY

The re1 a t i ve mer i t o f auxi 1 i ary cool ants which include water, l i thium, and sodium-potassium (NaK 56%) f o r an RJ-5!0xygen Mode 1 engine concept were determined, The re l a t i ve merit of the aux i l i a ry coolants was established on the basis o f cool ant capabi 1 i ty , auxi 1 i a r y system wei ght and complexi t y , and propel1 ant property consi deratlons .

6. TASK V I : ENGINE PRELIMINARY DESIGN

Preliminary designs and associated data were prepared f o r thre, candidate engine concepts and t he i r major com onents . Engines carr ied i n t o the design phase were:(l) baseline Mode 1, (2 dual-fuel, and (3) candidate Mode 1 resu l t ing from the study,

P

SECTION I 1 1

TASK I - PROPELLANT PROPERTIES AND PERFORMANCE

A. OBJECTIVES AND GUIDELINES

The object ives o f t h i s task were t o provide propel lant and combustion gas property data, operat ional character is t i c s , and theore t ica l performance for the various candidate propel lants and prope l lan t combinations considered i n t h i s study. To accompl i s h these object ives, 1 i tera ture surveys and analyses were conductea. Much o f the p r o p e l i A property data i s r rdd i l y avai lab le i n the 1 i tera ture and the best references are c i t e d herein. Ana ly t i ca l l y derived data o r r e l a t i v e l y new information, such as the ~ a t a on RJ-5, i s presented i n t h i s repor t .

Theoretical performance and combusti on gas property data were cal cu- l a ted fo r the fo l lowing parametric ranges.

Preburners (Fuel 6nd Oxidi zer-Ri ch)

Chamber Pressure: 136 t o 476 atm. (2000 t o 10000 psia) Mixture Ratio: def ined by the combustion tem erature range o f

700' t o 1367'K (800 t o 2000°Ff

Thrust Chamber

Chamber Pressure: 68 t o 340 atm. (1000 t o 5000 ps ia) Area Ratio: 1 t o 400 M i x ture Ratio: corresponding t o s to i ch i ometric f 5 ~ %

Propel 1 an t Stoichiometr i c Combination o/F

02/RJ-5 3.13 02/RP-1 3.42 02/MMH 1 ..'4

1 .oo 7.94 4 .OO

B. PHYSICAL AND THEROOYNAMIC PROPERTY DATA

A sumnary o f propel lant property data f o r the candidate fue l s and ovgen i s shown on Table V. Each o f the various propel lants are discussed i n the paragraphs which fol low and appropriate references c i ted .

TABL

E V.

- P

ROPE

RTIE

S O

F CA

hDID

ATE

PRO

PELL

ANTS

OXY

gcn

Hyd

roge

n RP

-1

RJ-

5 IC

H

Hyd

razl

ne

hth

4n

e

O2

"2

(a2)

12.3

7 C

14H

18.1

5 C

H6W

2 t'

F4

="

4 31

.998

8 2.

0159

4 17

3.51

51

186.

45

46.C

7c37

32

.045

28

' .

16.0

43 -

54.3

72

13.8

35

224.

8 26

8.2'

22

0.78

27

4.68

90

.68

(-361

.81%

) (-

434.

767)

(4

5)

(23)

1

(-62

.27)

(3

r.75

) (-

296.

4)

I I

I

I

I

I

I

Bo

ilin

g P

oint

. OK

90.1

88

20.2

68

492.

6

51 7-

547

360.

80

387.

4 11

1.64

(O

F)

(-29

7.34

6)

(-42

3.19

7)

(1.4

27)

(470

-525

) (1

89.7

7)

(237

.6)

(-258

.7)

Cri

tic

al

Tcl

pcr

atu

n,

OK

154.

581

32 .9

76

679

780

585

653

190.

6 (o

r]

(-19

1.43

3)

(-400

.31

3)

(763

) (9

44.3

3)

(593

) (7

16)

(-116

.7)

Vapo

r Pr

essu

re

at

298.

15°K

, kN

/d

-- --

1.8

.MB

6.

595

1.89

2 --

(at

77.F

. p

sia)

--

-- (.

26)

(.00

70)

(0.9

57)

(0.2

74)

-- . I D

ensi

ty.

liq

uid

s

t 29

8.15

OK

, kg

m3

I 1140

.8~

I 70

. 78b

I 10

67.4

I 87

0.2

1100

3.7

I 42

2. 6

b (a

t 77

OF.

lb

/ft

) (7

1.23

) (4

.419

) la

m (4

9.94

) (6

6.63

6)

(54.

325)

(6

2.65

9)

(26.a)

I I

I --

Har

t C

apac

ity,

li

qu

id

at

298.

15'K

. J/

g-O

K (

1.6

96

~

1 9

.d

I

1.98

1

1.28

(a

t 77

OF,

Ll

tu/lb

-OF)

(.

405)

(2

.316

) ( .

474)

( ,

307)

( .

7N1)

(0

.835

) 3.

078

3.

d

I Vls

cosl

ty,

liq

uid

a

t 29

8.15

°K.

a/m

2 . 1 9

wb

.0

13Zb

1

1.53

1 2

2.76

0.

775

I I 0.

913

I 0.

11ss

b (a

t 77

.F.

IbJf

t-re

c)

(1.3

16~1

0-4)

(~

387x

10-5

) (1

.04~

10-3

) (1

.530

~10-

2)

(5.2

1~10

'4)

(6.1

35

~1

~~

) (7

.76

~1

6)

Th

em

1 C

ondu

ctiv

ity,

lij.

at

298.

15.K

. M

/m-"K

.1

515~

.0

98gb

.I

37

.I617

.2

48

.4W

.1

99

(a

t 77

.F.

Btu

/ft-

su-O

F)

(2.4

33~1

0'5)

(1

.589

~10

'~)

(2.2

~10-

5)

(2. 6

0x10

'5)

(3.9

8~10

'5)

(7.8

6~

1~

5)

(3.1

0)

Hcl

t ~

f F

om

tio

n.

liq

uid

a

t 29

8.15

°K,

kcal

/ml

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1. Properties o f Oxygen

Detailed properties o f oxygen are avai lable i n a number o f sources and a l l the preferred ones are d i r e c t l y traceable t o the Cryogenics Div is ion of the National Bureau o f Standards. The best sources o f oxygen data are Ref's. 1, 2, and 3. For the thermal conductivi ty data, i n the v i c i n i t y o f the c r i t i c a l point where anomalous behavior occurs, Table V I o f Ref. 4 was used.

2. Properties o f Hydrogen

The detai led properties o f parahydrogen are avai lable i n a number o f source documents but are generally a l l traceable t o the Cryo- genics Division o f the National Bureau o f Standards. I n t h i s study program, Ref. 5 was used as the primary data source.

3. Properties o f RP-1

The properties o f RP-1 used, were taken from Ref. 6. These properties are dominantly obtained from Ref. 7 which i s an al ternate data source.

4. Properties o f Monomethyl hydrazi ne (MMH)

The primary source o f p r o ~ e r t i e s data on MH i s Ref. 8. Ref- erences 9 and 10 are supplementary data sources.

5. Properties o f Hydrazine

The primary source o f properties data on N2H4 i s also Ref. 8. Reference 11 i s a supplementary data source.

6. Propert i er o f RJ- 5 (she1 1 dyne-~R)

The most comprehenisve compilation o f properties data on RJ-5 has been prepared by the Fuels Branch, A i r Force Aero Propulsion Laboratory and i s the primary data source, Ref. 12. However, because o f data gaps and inconsistencies i n the avai lable information, the various recomnended propert ies o f RJ-5 are discussed.

a. Empirical Formula, Molecular Weight and Heat o f Formation

Various heats o f formation and empirical formula were derived based upon 1 i terature searches, comnunications w i th the Sun O i l Company and communications wi th NASAILeRC. The data i s sumnarized herein t o show the sens i t i v i t y t o the avai lable information. The l a tes t data (Sept. 1975) obtained from the A i r Force Aero Propulsion Laboratory a t Wright-Patterson AFB by the Lewis Research Center was used i n t h i s study. This data i s based upon the average of analyses on three batches o f RJ-5.

Source _L_

Chemical Composition 1 somer X W t . - -

I L I terature Search (Ref. 13) Clp Hla 85.3 C14 H20 10.3 C14 H22 4.4

I1 Communication w i t h Sun O i l C14 H18 C14 H20

I11 Communication w i th NASA/LeRC C14 H18 92.45 ( Recomnended) C14 H20 7.55

Based upon the compositions, the fo l lowing can be derived f o r each case:

Mean Molecular Empi ri cal Source Weight Fomul a

I 186.703 C14 H18.4 I I 186.593 C14 H18.29

I I I (Reconmended) 186,450 C14 H18.15

The net heats o f combustion and formation obtained f o r each o f the calculat ions are:

Net Heat o f Combustion Source - k -b)

I 1829 .065 I I 1829.650

I11 (Recomnended) 1854.250 17,901 )

Heat o f Formation* k cal/q-mole

-19.4 -15.6 +13.0

The e f f e c t o f a l l ca lculated heats o f formation and empir ical formula on theoret ica l spec i f i c impulse i s only approximately 3 t o 4 secs. A pos i t i ve heat o f formation value resu l t s i n higher spec i f i c impulse.

A l l o f the c i t e d calculat ions aFpear t o be approximately v a l i d f o r themchemi c a l l y character iz ing RJ-5. Previous values were based on an extremely small data base. Therefore, the l a t e s t values obtained by NASAILewis from the Aero Propulsion Lab were recomnended. Further carbon and hydrogen analyses should be performed on various avai lable l o t s o f RJ-5 and heats o f combustion determined on each o f those l o t s . With t h i s extended data base, a more r e l i a b l e empir ical formula and heat o f formation could be established as we l l as the v a r i a b i l i t y and precis ion of the values.

b , Freezing Point

RJ-5 does no t e x h i b i t a sharp freezing p o i n t bu t ra the r complex supercooling and Tract ional crys t a l l i zat ion behavior. RJ-5 under-

9 oes essent i a1 l y complete c rys ta l 1 i zat ion a t temperatures lower than 231 'K -45°F)' followed by p a r t i a l me l t ing on warming t o 250°K (-10°F), r e s o l i d i -

f i c a t i o n on hold ing a t t h i s temperature and remel t ing on warming t o 253°K (-5OF) iv i th the l a s t c r ys ta l d isappear i~ ig a t 268°K (+23"F) (Ref. 13). Assuming t h a t 268°K (+23"F) i s the h ighest temperature a t which an s o l i d phase can e x i s t i n equ i l ib r ium with the 1 i q u i d phase, 268°K (+23"F f i s recomnended as the "threshold f reez ing point," a value o f most importance i n rocketry.

c. B o i l i n g Point

Since RJ-5 consists o f a mixture o f compounds, i t exh ib i t s a b o i l i n g range ra ther than a s p e c i f i c b o i l i n g po in t . The range o f 517 t o 547°K (470 t o 525OF) (Ref. 12) was recomnended f o r use.

d. C r i t i c a l Temperature

The c r i t i c a l temperatures given i n Ref. 12, 9 4 V F and 504OC, are inconsis tent w i t h each other ( i .e., 945OF = 507.z°C). Although nb t specif ied, these values are probably estimates. Analyses were performed which resu l ted i n an estimated value o f 780°K (506.85OC o r 944.33OF). These values are consistent, c lose t o those i n the reference, and hence, were used. The inaccuracy i n t h i s value i s estimdted t o be f 15°K.

e. C r i t i c a l Volume and C r i t i c a l Density

No values f o r c r i t i c a l volume o r densi ty o f RJ-5 were found i n the 1 i terature. Therefore, these values were estimated through ana ly t i ca l techniques. The estimated values used i n t h i s study are:

C r i t i c a l Volume = 622.4 c d / g Mol C r i t i c a l Densi ty = 0.300 glcm3

The probable inaccuracies i n these estimates are f5%.

f. C r i t i c a l Pressure and Compressibi l i ty

The c r i t i c a l pressure values o f Ref. 12 were a lso incon- s i s t e n t and p r e s u ~ b l y are estimated values. Analyses were again performed t o a r r i v e a t a consistent se t o f values. This resu l ted i n the fo l low ing estimated values:

C r i t i c a l Pressure = 25.2 atm (370 psia) C r i t i c a l Compressibi l i ty (Zc) = 0 .2448

g. Vapor Pressure

The vapor pressure o f RJ-5 has been measured by A t l a n t i c Research (Ref. 14) a t 373, 398, 423, and 350°K w i t h dupl icate vapor pressure measurements taken a t each temperature. A least-squares curve f i t o f the e igh t data points, wherein the pressure measurements using the Bourdon tuce gage are given twice the weight o f those using a mercury manometer, y ie lds the fo l lowing equation:

Log P(psia) = 4.817675 - 3741.925 !I? '459.67 + OF'

This equation was used f o r the temperature range o f 373 t o 448°K (212-347OF).

For higher temperatures, the vapor pressure a t 448°K (347"F), the normal average b o i l i n g point, and the c r i t i c a l po in t were used t o define the three constants f o r an Anotoine-type vapor pressure equation which i s given below for the temperature range 448 < T < 780°K (347 c T < 944.33OF).

- - - - Log P(psia) = 4.751597 - 2496.043

m . 8 6 3 + T (2)

The above equations were used i n t h i s program t o def ine the vapor pressure o f RJ-5. Values calculated from these equations are shown on Table V I .

h. Density

The density o f RJ-5 (anbient pressure) has been measured by A t l a n t i c Research (Ref. 14) over the temperature range of 0 t o 150°C (32°F - 302°F). The n ine values reported were curve- f i t t o g ive the fol lowing equivalent equation:

~ ( g / m l ) = 1.08716 - .00079288(°C) f o r 0 - < T - < 150°C (3) p ( l b/f t3) = 68.75385 - .02750753("F) f o r 32 5 T - c 302°F (3a)

p = density

For densi t ies i n the compressed l i q u i d s ta te and a t temperatures substant ia l l y above 150°C, the density can be estimated by the f o l lowf ng equation:

p ( l b l f t 3 ) = 19.931 (pR) where pR = reduced densi ty (densi t y l c r i t i c a l density)

(4)

Values o f pR were obtained from Table 3-6 o f Ref. 15.

Density values calculated from the equations above are summarized i n Table V I I .

TABLE V I . - VAPOR PRESSURE OF RJ-5

Temperature O K

255.6 31 1 . 1 366.7 422.2 477.8 533.3 588.9 644.4 700.0 755.6

Temperature O F

0 100 200 300 400 500 600 700 800 900

TABLE VII. - DENSITY OF RJ-5

B n s i ty Sat. Lip. At 340 Atm. At 680 a h .

Engl i sh Units

Density, lb(m)/ft3 Sat. Liq. A t 5000 Psia At 10000 Psia

i, viscos i t y

The v i scos i t y o f RJ-5 has been measured by A t l a n t i c Research (Ref, 14) a t low shear rates w i t h Cannon-Fenske viscometers i n the tempera- tu re range o f 250 t o 448OK (-10.3 t o 347OF).

The rheology of RJ-5 was a lso invest igated i n d e t a i l by A t l a n t i c Research (Ref. 14) w i th an extrus ion rheometer i n the low tempera- tu re range, 219 t o 296OK (-65 t o 72OF); where RJ-5 exh ib i t s a t r a n s i t i o n from Newtonian t o pseudoplastic behavior as temperature decreases,

The v i scos i t y o f RJ-5 has been estimated by Shel l Develop- ment Company (Ref. 16) over a range o f condit ions using the well-known Jossi-Stiel-Thodos cor re la t ion o f residual v i scos i t y and a c r i t i c a l proper t ies funct ion w i t h reduced density.

The v i scos i t y used was a smoothed combination o f ; (1) the experimental v i scos i t y data (Ref. 14) , (2) a graphical extrapolat ion o f t h a t data, and (3) the estimated v i scos i t y data a t the higher temperatures (Ref. 16). The data are sumiwized on Table V I I I .

j . Heat Capaci ty and Thermal Conducti v i t y

The heat capacity and thermal conduct iv i ty o f RJ-5 apparently have not been experimental1 deter~~i ined. Estiiilated values are reported by At1 a n t i c Research (Ref. 14 1 which are s t t r i butable t o She1 1 Development (Ref. 16). The accuracy of these data are q u i t e uncertain bu t were used i n t h i s program for lack of any other data. These data are summarized i n Table I X .

7. Properties o f Methane (CH4)

a. T r ip le Point

The t r i p l e po in t values o f methane were obtained from a recent NBS compilation (Ref. 17) and are shown below,

and are:

20

T r i p l e Point Values o f Methane

Temperature: 90.680 OK(1 63$24'R, -296.446OF) Pressure : 11743.57 N/m (1.70326 psia, 0.11590 atm) Dens i t y :

Liquid: 451.562 k g / d 28.1901 lb / f t 3 ) Vapor: 0 .251 5326 kg/m 4 (0.01 570268 I b / f t 3 )

b. C r i t i c a l Point

The c r i t i c a l po in t values o f methane obtained from Ref. 17

TABLE V I I I . - VISCOSITY OF RJ-5

Viscosity ~ - s e c / m ~ (1 bmlf t -secl

TABLE I X . - HEAT CAPACITY AND THERMAL CONDUCTIVITY OF RJ-5

Temperature " K (OF) -

Thermal Conducti v i ty W/w0K (BTU/ft-sec-OF)

. I66 2.67

. I61 2.58

. I56 2.50 lo-5

. I47 2.36

. I36 2.19 1 0 ' ~

. I25 2.00

. I11 1.78

C r i t i c a l Point Values o f Methane

Temperature : 1906555"K(342 99g0R, -1 16.671 OF) Pressure: 4.598825 MI4 m i (667.003 psia, 45.3869 atm) Density : 160.43 kg/ (10.015 l b / f t 3 ) Compressi b i 1 i ty: 0.29027

A c. P-V-T and Derived Thermodynamic Properties

The pressure-volume-temperature data and deri.!ed thermodynamic propert ies o f methane ( i n te rna l energy, enthal py , entropy, spec i f i c heats a t constant pressure and constant volume, and speed o f sound) are avai lable i n Ref. 17 f o r the e o f pressures from the t r i p l e po in t pressure t o 680 atm (10,000 and f o r the range o f temperatures from the mel t ing l i n e t o 500°K (440°F

Values o f density and heat capacity a t constant pressure, Cp, are no t read i l y avai lable above 500°K (440°F) and a t high pressures, 136 t o 680 atm (2,000 t o 10,000 psia) . Because values i n t h i s pressure range and t o temperatures near 922°K (1200°F) were necessary t o conduct heat t rans fer analyses, such values were estimated. These extrapolated values are given i n Tables X and X I and are intended t o supplement the data avt i i lab le i n Ref. 17. The heat capacit ies were estimated based upon the idea l gas heat capacit ies given i n Ref. 18 and heat capacity correct ions f o r the e f f e c t o f pressure given by Gambi 11 (Ref. 19) .

Densities were extended i n t o the higher temperature and pressure region u t i 1 i z i n g estimated compressib i l i ty factors from L dersen's generalized tables (Ref. 20) and then adjust ing those estimated va f ues upward or downward s l i g h t l y so t h a t the resu l t i ng estimated f l u i d densi t ies agreed wi th those given i n Ref. 17 a t 500°K (900°R) f o r each pressure being considered.

d. Viscosity

The v i scos i t y of many substances can be c lose ly approximated by the fo l lowing expression:

n = no + t o (5)

where Q~ = the gas v i scos i t y a t low pressure and i s a funct ion o f temperature only

AQ = Tl-QG, the so-cal led res idual v i scos i t y which i s a funct ion of densi t y only, as a f i r s t good approximation (except near the c r i t i c a l po in t )

Recommended values of q o from Ref, 21 are given i n Table X I I . These re fe r - ence values are based on an analysis of 27 sets o f experimental data.

TABLE X. - EXTRAPOLATED HEAT CAPACITIES OF METHANE

Cp, J l g - O K

136 atm 272 atm 408 atm 544 atm 680 atm

Engl i sh Uni t s

Temp. , O R 2000 psi a 4000 psia 6000 psia 8000 psia 10,000 psia

TABLE X I . - EXTRAPOLATED DENSITIES OF METHANE

Temperature Dens 4 ty , k9/m3 O K 136 atm 272 atm 408 atm 544 atm 680 atm

600 700 800 900

1 000

Temp, O R

1080 1 260 1440 1620 1800

Engl i sh Units

Density, 1 b / f t3 2000 psia 4000 psia f!!%WL 8000 psia. 10,000 psia

' ? y - - - t'? ~ g ~ $ i s s r s r a ~ s ~ g ; ; $ g ~ ~ g $ C C C C C C C -

- C I A I ~ ~ ~ ~ f D ~ ~ ~ g # j f ~ # ~ g ~ $ ~ $ ~ ~ f C C C C C

Y Y I Y Y Y Y Y Y I Y Y Y Y Y Y Y Y I - 1 1

Values o f A n versus densi ty are presented graph ica l l y i n Ref. 22. D i f f i c u l t y i n reading the graph accurately l e d t o the decision t o recor re la te the most comprehensive ex erimental data. This was accomplished using values o f no (Ref. 21 ! and experimental values of n (Ref. 22 and 23). The correspondi i~g values of density were determined from Ref. 17. A good graphical co r re la t i on o f A q versus density was then obtained. The o r i g i n a l values o f r, were obtained a t temperature: from 103 t o 273OK (-274' t o 32OF) and a t pressures up t o 340 atm (5000 ps la) . More than 100 data points were u t i l i z e d i n the cor re la t ion . The r e s u l t i n g values of A n versus density are 1 i s t e d i n Table XII.

e. Thermal Conducti v i t y

The thermal conduct iv i ty o f methane has been corre lated i n a manner q im i l a r t o t h a t used i n co r re la t i ng the v i scos i t y data, i ,e.,

where ko = the gas thermal conduct iv i ty of low pressure which i s a funct ion c f temperature only

~k = k-ko, the so-cal led res idual thermal conduct iv i ty which i s p r i m a r i l y a funct ion o f density, as a f i r s t good a proximation (except near the c r i t i c a l p o i n t ! .

Values o f ko were obtained from Ref. 24 and are given i n Table XIII. These values are based on an analy; ! s o f eleven sets o f experimental data by the authors o f Ref. 24.

Values o f ~k versus densi ty are based on a graphical cor- r e l a t i o n which was developed i n t h i s study. The ~ k ' s , as def ined by Eq. ( 6 ) , were obtained using values of ko (Ref. 24) a d values o f k p r imar i l y from Ref. 25. The corresponding val ues of dens i t y were determined from Ref. 17. The o r i g i n a l values of k were obtained a t tempwatures from 99 t o 235OK (-281-to -37OF) and a t pressures up t o SO0 atm (7360 ps ia) . Forty-f ive dzta points were u t i l i z e d i n the cor re la t ion . Enhanced thermal conduct iv i t ies i n the v i c i n i t y o f the c r i t i c a l po in t have been neglected i n the co r re la t i on because of i nsu f f i c i en t data and because expected operat ing pressures are an t ic ipa ted t o be very subs tant ia l l y greater than the c r i t i c a l pressure.

The r e s u l t i n g values o f ~k versus densi ty are l i s t e d i n Table XIII.

C. OPERATIONAL CONSIDERATIONS

Operational considerations i ncl ude those propel 1 ant properties or characteristics that have a significant impact upon the re1 iabil ity and cost of the engine and its potential impact on the environment such as; (1) cost and availability, (2) safety and ground handling, (3) chemical and/or thermal stabil i ty , (4) corrosivi tylmaterial s compati bi 1 i ty, and (5) environmental effects.

Three factors that strongly affect the cost of the candidate propellants are the cost of; (1 ) petroleum and/or natural gas, (2) energy to produce, and (3) added operating and/or capital costs arising from new environmental and/or occupational health requirements . Recent changes of significant magnitude in some of these cost factors and uncertainty as to the magnitude of further changes in the near-term makes propellant cost projections very difficult. However, an attempt to estimate propellant "hard" cost was made in this study. Two methods of approach were tried. These are:

O Extrapolation of historical data " Use of cost escalation indices

The historical propel 1 ant cost data on oxygen, hydrogen, RP-1, N2H4 -and MMH were provided by the NASA/LeRC Project Manager as costs to the government for mid 1966 and 1975. The RJ-5 cost estimate is based upon telephone communications between ALRC and Air Force Aero Propul sion L? boratory, Wright-Patterson Air Force Base personnel . The base point nistorical cost for methane was obtained from NASAIMSFC by ALRC to support the "Phase A/B Study for a Pressure-Fed Engine on a Reuseable Space Shuttle Booster," Contract NAS8-28217. These data, obtained in October 1971, indicated that methane and propane were expected to cost about 0.055$/kg (0.025 $11 b) while RP-1 was approximately 0.066 $/kg (0.03 $/lb). For purposes of this study, these costs were considered to be equal. Extrapolation of this data to 1976 results in a cost of approximately 0.132 $/kg (0.06 $/lb) which agrees with the data shown in Ref. 26.

The historical data and estimates were plotted on semi- logarithmic graph paper and linearly extrapolated to 1990 as shown in Figure 2. It i: -eal ized that there have been step changes in some of the propellant costs over the past years. However, the assumption is that a straight 1 ine averages out these steps. The historical data on hydrogen indicates that the cost has remained constant. An article in the July 14, 1975 issue of Aviation Week and Space Technology (Page 21 ) indicates that the price of hydrogen will grow to 3.31 $/kg ($1.50/lb) over the remainder of a 12-year NASA contract with Air Products to 1987.

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The escalat ion indices used was o b t a i w d from the Bureau t , Labor Standards as o f 1 October 1975. A r a t e o f 9%/year i s cu r ren t l y being used by procurement people a t ALRC.

The h i s t o r i c a l data ex t rspo la t ion and the escalated costs are w i t h i n reasonable agreement as shown by the f igure. The data i s summarized on Table X I V . The projected cost f o r 1990 shown on t h i s t a b l e i s based upon the escalat ion fac tor .

Both the extrapolat ion and escalated cost pro ject ions assume t h a t current manufacturing methods w i 11 be used t o produce the propel 1 ants.

2. Safety and Ground Handling

The safety and g r o u ~ d handl i ng charac ter is t i cs o f the candidate propel lants are conveniently summarized i n d e t a i l i n CPIA Publ i - cat ion No. 194 (Ref. 27). This repo r t i s a standard data source f o r such information.

Because 02/H was def ined as the only Mode 2 engine propel - l a n t combination and 021 g J-5 as the basel ine Mode 1 engine propel lant combination f o r t h i s program, the yanking o f candidate propel lants w i t h respect t o safety and handling cb ,~ rac te r i s t i cs reduces t o the ranking of RP-1, N2H4, MMH, and CHq versus RJ-5 as a1 te rna t i ve Mode 1 fuels . RP-1 and CHq compared w i t h RJ-5 tend t o be ranked s l i g h t l y below RJ-5 from the safety and handl i ng standpoint p r i m a r i l y because o f t h e i r higher vol a t i l i t y (flamnabi 1 i t y and t o x i c hazards). However, experience w i t h RP-1 would ind ica te t h a t a l l three can be handled safely. Hydrazine and MMH present a s i g n i f i c a n t l y greater cqfety and handl i n g problem than RJ-5, RP-1 o r CHq f o r a number o f reasons: (1 ) higher v o l a t i l i t y , (2) very tox ic , (3) h igh ly flammable, (4) higher r e a c t i v i t y w i t h a i r and mater ia ls o f con- s t ruct ion, and (5) t h e i r a b i l i t y t o decompose rap id l y and exothermical ly from ce r ta in thermal o r c a t a l y t i c s t imu l i .

Oxygen, the ox id i ze r f o r both Mode 1 and Mode 2 engines, has several safety and handl i n g cha rac te r i s t i cs important t o note: (1 ) i t s cryogenic nature, and (2) i t s strong tendency t o reac t ox ida t i ve l y w i t h many mater ia ls . By i t s e l f , oxygen i s a very s table mater ia l , but many organic contaminants i n a dense oxygen phase can y i e l d explosive mixtures which can be i n i t i a t e d by a va r ie t y o f s t i m u l i such as shock, f r i c t f o n , etc. Gaseous oxygen contaminated by gaseous organics, o r f i n e l y d iv ided sol i d organics, o r metal, can s i m i l a r l y form explosive mater ia ls t h a t are sensi- t i v e t o thermal s t imu l i . Oxygen, a t high temperatures, w i l l cause almost any organic mater ia l t o burn and r e a d i l y causes a f a s t surface ox ida t ion o f metals. Most metals w i l l a c t i v e l y burn i n oxygen a t s u f f i c i e n t l y high temperaturelpressure condit ions .

Propel lant

LH2

L02

MMH

TABLE X I V . - PROPELLANT COST SUmARY

(a)~ased upon escalation factor.

Hydrogen, the f u e l f o r the Mode 2 engine, a1 so exh ib i t s several safety and handling charac ter is t i cs worthy of special consideration: (1) i t s very low density and extreme v o l a t i l i t y which are a t t r i b u t a b l e t o i t s very low molecular weight and b o i l i n g point , (2) i t s high f l a m a b i l i t i n a i r and other oxidizers, (3) i t s very high heat o f combustion, and (47 i t s adverse e f f e c t on ce r ta in metals due t o hydrogen enbr i t t lement phenomenon.

A1 though each propel 1 ant exh ib i t s many unique charac ter is t i cs , a l l have been sa fe ly handled i n past rocket engine programs. Therefore, safety i s no t a s ing le major propel lant se lec t ion c r i t e r i a .

3. Chemical and/or Thermal Stabi 1 i t y

Oxygen and hydrogen are both thermal ly s tab le bu t thcy are h igh l y reac t ive i n ox idat ion and reduct ion react ions , respect ive ly . As previously mentioned, oxygen w i l l form explosive mixtures w i t h a wide var ie ty o f both organic and inorganic mater ia ls having fue l value. These mixtures w i l l react exothermical ly when exposed t o ce r ta in thermal o r shock st imul i . The violence o f the react ion w i 11 depend upon the n ix tu re , the degrec o f confinement, and the i n t e r f a c i a l area o f the oxygenlmaterial mixture. Hydrogen behaves qui t e simi 1 a r l y except i t s p o t e n t i a l l y dangerous mixtures form w i th mater ia ls having ox ida t ive capabi li t i e s (e .g., a i r and oxygen. )

The hydrocarbon fuels , RJ-5, RP-1, and CHq are r e l a t i v e l y s tab le bu t they do have tendencies t o gum, crack, o r coke a t elevated temperatures. These tendencies are increased when contaminated by water and/or oxygen. Because these fue ls are flammable, t h e i r mixtures w i th ox id izers such as a i r o r oxygen are subject t o the burning o r explosive chemical react ions t yp i ca l o f heavy hydrocarbons.

The gumning, cracking, and coking charac ter is t i cs o f RP-1 and KJ-5 are o f p a r t i c u l a r concern when they are used as regenerative coolants, o r i n f ue l - r i c h gas generators (pr,iburners. ) Comprehensive comparati ve thermal s t a b i l i t y t es t i ng of these two fue l s by the same t e s t procedures i n the same t e s t equipment are no t known t o have been conducted, so t h e i r r e l a t i v e thermal s t a b i l i t i e s are d i f f i c u l t t o assess. It would appear, however, t h a t RJ-5 has the po tent ia l f o r the b e t t e r thermal s t a b i l i t y because i t i s a synthet ic f u e l and can be p u r i f i e d t o remove unstable components t o almost any degree desirable (only a m i 1 i t a r y spec i f i ca t ion d ra f t present ly ex is ts f o r i t ) . RP-1 i s a spec ia l l y p u r i f i e d kerosene, bu t nonetheless, may contain ce r ta in spec i f ied maximum quanti t i e s o f " impur i t ies" (as ex is ten t and po ten t i a l gums, su l fu r , mercaptan , aromatics, and o lef ins) which are known t o have a considerable in f luence on thermal s t a b i l i t y .

On the basis o f thermal s t a b i l i t y tes ts by the JFTOT (Je t Fuel Thermal Oxidation Test, ASTM D 3241-73T) t e s t method on RJ-5 and J e t A-1 fue ls (Ref. 28) , and the close s im i l a r i t i e s between J e t A-1 and RP-1, RJ-5 i s thermal ly s table ( i n terms o f deposit formation on heated surfaces) t o a temperature of approximately 561°K (550°F) wh i le RP-1 i s probably s table t o

about 547 t o 561°K (525550°F) using s i m i l a r c r i t e r i a . From th i s , i t can be estimated t h a t l i qu id -s ide wa l l tempeyatures i n excess of 589°K (600°F) w i l l lead t o the f o u l i n g o f the coolant flow passages, p a r t i c u l a r l y w i th repeated engine cycles .

Methane i s thermal ly s tab le a t temperatures up t o approximately 978°K (1300°F), Ref. 29. Therefore, coolant passage f o u l i n g would no t be expected w i t h t h i s propel 1 ant.

The use 9 f RP-1, RJ-5, o r CH4 i n fue l - r i ch preburners can present a thermal s t a b i l i t y probl ;m manifested as coking. Coking of heated hydrocarbon fue l was studied under the T i tan I (LR91-AJ-3) engine development contract because the f u e l was used as the regenerative coolant, and a f u e l - r i c h gas generator was u t i l i z e d , Ref. 30. JPL conducted tes ts by f low ing j e t propulsion fuel through a heated tube and exhausting the heated vapors through a known o r i f i c e . It was noted t h a t a t approximately 994°K (1330°F). coking suddenly became so pronounced t h a t the tes ts could n o t be continued due t o complete choking of the tube. On the basis o f these t e s t s , i t was recomnended t h a t any j e t propulsion fue l - r i ch gas generator be run a t temperatures below 978°K (1300°F), the time a t which any f u e l may be above t h i s temperature be minimized, and any l oca l temperatures over 978°K (1300°F) be e l iminated by uni formly mixing the propel lants dur ing the i n j e c t i o n process. S imi la r tes ts conducted by Aerojet, us in RP-1 fue l , showed t h a t 1 i t t l e coke formed a t a temperature of 922°K (I~OOOFQ; a l i g h t amount formed a t 978°K (1300°F), and a heavy amount formed a t 1033°K (1400°F).

On the basis o f these t e s t r e s u l t s and experience w i t h hydrocarbon fue ls i n At1 as, T i tan I, and the F-1 engines , i t would appear t h a t ccking problems w i th e i t h e r RP-1 o r RJ-5 i n fue l - r i ch preburners present no major d i f f i c u l t y provided the gas temperature i s con t ro l l ed below 922°K (1230°F). This assessment should, however, be experimental ly v e r i f i e d i n the case o f RJ-5. For study purposes, a f u e l - r i c h gas tu rb ine i n l e t tempera- tu re o f 867°K (1100°F) was selected.

The candidate fue l s , hydrazi ne and monomethyl hydrazine , have unique thermal s tab i 1 i t y charac ter is t i cs which can have important bearings on t h e i r su i t ab i 1 i ty f o r use i n high-pressure, staged-colnbustion-cycle engqnes. Both o f these fuels e x h i b i t ce r ta in thresholds o f s t a b i l i t y and these thresholds are dependent upon the materials o f construct ion and the nature of the s t imu l i t h a t imparts energy t o the f u e l .

The thermal s tabi 11 t y o f MMH and N2H4 and N H4-MMH (90110) was the subject o f 3 comprehensive experimental evaluat ion tRef. 31). The materials of construct ion, the physical s ta te o f the fue l , and the nature of the sens i t i z i ng stimulus have important inf luences on the threshold s e n s i t i v i t y temperatures of these fue ls . The l iqu id/vapor phase f o r MMH undergoes i n i ti a1 exothermi c decomposi ti on a t an average temperature of 499°K (438°F) i n the ten metals studied bu t var ia t ions i n temperature of nearly 233°K (+ 6 0 " ~ ) occur because o f the nature o f the various metals. S im i l a r l y , the average threshold temperature f o r explosive decom~osi t ion

i s 547OK (524OF) w i t h metal-caused va r i a t i ons o f about -214 t o 283OK (-75 t o +50°F). The presence o f MMH on l y i n the vapor phases decreases t h2 threshold sensi t i v i t y temperature about 289 t o 306OK (60 t o 90°F). A comparison of the s e n s i t i v i t y of mixed l i qu id / vapo r phases o f MMH o r N2H4 t o moderate hea t ing ra tes versus a combination o f hea t ing fo l lowed by ad iaba t i c compression of the vapor phase, shows t h a t MMH vapor i s n o t appreciably sensi ti ve t o compression whi l e N2H4 vapor i s general l y very sens i t i ve t o compression and i s f u r t h e r sens i t i zed by some metals. I n view of t he marked d i f f e rence i n the s e n s i t i v i t y o f MMH and NzH4 vapors t o compression, i t i s impor tant t o note t h a t even a smal l add1 t i o n o f MMH t o N2H4 ( i .e., 10% wt) g r e a t l y reduces the s e n s i t i v i t y o f N2H4 t o compression. The e f fec t i veness o f ammonia (NH3) i n reducing N2H4 s e n s i t i v i t y i s a l so estab l ished, a1 though l e s s succ inc t l y documented.

The use o f MMH as a coolant wi 11 necess i ta te designs t h a t insure the bu l k temperature o f MMtl does n o t exceed about 478OK (400°F). The use o f hydrazine imposes s i m i l a r thermal , l i m i t s under normal con- d i t i o n s , b u t a much more severe l i m i t , about 367OK (200°F), if simul- taneous exposure t o vapor phase ad iaba t i c compression occurs such as dur ing s t a r t o r shutdown t rans ien ts . The f a c t t h a t a copper a l l o y i s used t o form a p a r t of t he coolant passage, and t he e f f e c t o f copper a l l oys on the thermal s t a b i l i t y o f MMH and N2H4 has n o t been establ ished, po in t s up the need f o r f u r t h e r experimental work.

4. Materi a1 s Compati b i 1 i t y l c o r r o s i v i t y

a. Oxygen

Oxygen i n c o m p a t i b i l i t y i s manifested e i t h e r by l o s s o f toughness a t low temperatures, reduc t ion o f f a t i g u e 1 i fe , o r ca tas t roph ic ox ida t ion .

I n the case o f most common s t r u c t u r a l a l l oys , the ox ida t i ve a t tack becomes appreciable on l y a t h i gh temperature because the reac t i on i s i n h i b i t e d by the format ion o f a p r o t e c t i v e ox ide 1 ayer, I n the case o f organic mater ia ls , the choice o f acceptable mater ia ls i s very l i m i t e d , as i s the app l i ca t i on of the mater ia ls t o var ious engine uses. While the above i nd i ca tes a s i g n i f i c a n t poss ib le problem i n us ing oxygen, the long- s tanding experience w i t h oxygen i n a v a r i e t y o f vehic les (e.g., A t las , T i t an I, Saturn, etc.) has lead t o a w e l l es tab l i shed group o f mate r ia ls and corresponding engine uses t h a t permi t r e1 i able appl i c a t i on o f oxygen. Several repor ts prov ide good compi la t ions o f mate r ia ls ' acceptable f o r use i n oxygen serv ice (Ref. 32, 33, and 34.) T i t a n i um and i t s a l l o y s and aluminum and i t s a l l o y s are incompat ible from an ox ida t i on s tandpoint when they are sub jec t t o h igh energy inputs .

The aluminum, austeni t i c i r on , copper, n i c k e l and coba l t base a1 loys a1 1 posess adequate toughness f o r app l i ca t ions t o 94°K (-290°F) . The ef fect o f oxygen and water vapor as found i n oxygen combustion products on fa t igue l i f e has been inves t iga ted (Ref. 35). The s t r a i n u t i l i z e d i n

these tes ts (Ref. 35) sugyests t h a t both h igh cycle and low cycle fat igue are de le ter ious ly affected, L imi ted f rac tu re toughness tes ts have been performed i n oxygen. Inconel 718 d i d no t e x h i b i t environmental l y enhanced subcri t i c a l f law growth when tested i n 68 atm (1000 ps i ) oxygen a t room temerature (Ref. 36). Considerable t e s t i n g remains t o be done t o es tab l ish the effects of gaseous oxygen on f rac tu re toughness and h igh and low cyc le fat igue of candidate mater ia ls .

b. Hydrogen

One o f the most important mater ia ls considerat ion i n the use o f l i q u i d hydrogen fue l i s embritt lement. The s u s c e p t i b i l i t y o f metals t o embri t t lement by h igh p u r i t y , high pressure, 272 t o 510 atm (4000-7500 psi ) gaseous hydrogen has been extensively invest igated f o r candidate a1 loys fo r Space Shut t le main engine (SSME) components (Ref. 37 and 38) .

Hydrogen incompati b i 1 i t y i s m n i fes ted by loss o f toughness both w i t h decreasing temperature and by hydrogen absorption. Embri t t lement by hydrogen absorption i s th? most severe a t room temperature except f o r those mater ia ls which are af fected by hydrogen react ions w i t h i n the metals such as hydride formati on, hydrocarbon gas formati on through the reduct ion of carbides o r water vapor formation through the reduct ion o f oxides.

Mater ia ls such as Inconel 718, a mater ial which i s suscepti - b l e t o hydrogen embri t t lement, are used extensively i n hydrogen fuel systems, However, they are l i m i t e d t o low temperature appl icat ions where enbr i t t l e - ment e f fec ts are minimal . These appl i ca t ions inc lude engine components which are a t ambient as we l l as c h i l l e d temperatures a t engine s t a r t ; and hence, the former experience t r a n s i t i o n from an e m b r i t t l i n g t o a non- e ~ r i t t l i n g condit ion. A s i m i l a r condi t i o n ex i s t s i n the ho t gas system where n icke l base a l loys are heated from low temperature t o elevated temperatures i n an enbri ttli ng environment dur ing engine s t a r t , The sui ta - b i li ty o f these materials i n these appl icat ions depends upon surface chi ll - i n g o f the mater ia l p r i o r t o reaching design stresses o r rap id t r a n s i t i o n t o the nonenbri ttl i n g elevated operat ing temperature, The t o t a l time t h a t the mater ia l w i l l be subjected t o the e m b r i t t l i n g temperature range w i l l be extremely s h o r t from the standpoint o f the app l ica t ion of f rac tu re mechanics t o the design.

The inf luence o f hydrogen i n a fue l - r i ch h o t gas system composed o f the hydrogen-oxygen combus ti on products i s uncertain. The e f f e c t of hydrogen p u r i t y has been invest igated (Ref. 39). This reference shows tha t the in t roduc t ion o f 0.6% 02 completely arrested hydrogen i n - duced crack growth i n an a l l o y s tee l . A s i m i l a r e f f e c t i s reported fo r the in t roduc t ion of moisture i n t o hydrogen. However, these resu l t s w e inconsis tent w i t h the data o f Ref. 37 which shows increased e n b r i t t l i n g effects w i t h the in t roduc t ion of water vapor. I n view o f these incon- s i stencies, hydrogen embri t t lcment must be considered from the standpoint o f pure gas e f fec ts u n t i l more d e f i n i t i v e information becomes avai lab le. The current design o f the ho t gas system o f the SSME c a l l s f o r the pro- t ec t i on of suscept ible mater ia ls w i t h p la t ings and weld overlays o f unaffected mater ia ls t o avoid embritt lement.

35

c. RP-1, RJ-5, and CH4

The hydrocarbon fuels, RJ-5, RP-1 and CH4 do not present any s i g n i f i c a n t co r ros i v i t y lmater ia ls compa t ib i l i t y problems. They e x h i b i t excel lent corcpatibi 1 i t y t o a va r i e t y of s teels , s ta in less steels , aluminum al loys, and t i tan ium a l loys and t o a v a r i e t y o f elastomeric mater ia ls such as V i ton, n i t r i t e , and f l uo ros i 1 i cone rubber compounds. However, a carbur i - zat ion po tent ia l ex i s t s i n t h e i r f ue l -ri ch combustion products, Carburi - z a t i on should no t occur below approximately 1000°K (1 350°F) (Ref. 40) . Where temperature poses a carbur izat ion problem, n icke l base a1 loys can be employed t o reduce carburi z a t i oq .

d. MMH and Hydrazine

Metals i n contact w i th the hydrazine fue l can be subject t o s t ress corrosion cracking and can catalyze decomposition. Stress corrosion i s caused by the react ion of the metal w i t h impur i t ies i n the fuel t o produce hydrogen on the metal surface.

The t i tan ium and aluminum a1 loys are unaffected wh i le the nrartensi t i c (CRES 410 and 4130 a l l o y s tee l ) and the n i cke l base a l l o y (Incone: 718) d isp lay crack growth a f t e r varying incubat ion periods (Ref. 41). The incubat ion period f o r crack growth i n MMH i s ten times t h a t of hydrazine. Although Inconel 715 i s suscept ible t o stress corrosion i n hydrazine, the incubat ion period f o r crack g r ~ w t h i s SO0 hours; a time we l l beyond the current requi rernents o f 1 i q u i d rocket enqlnes . Addit ional t e s t i n g i n hydrazine and MMH a t elevated temperatures i s required t o es tab l ish compa t ib i l i t y 1 i m i ts . A1 though maximum embri t t lement e f fec ts f o r several mater ials occurs a t room temperature, the e f f e c t o f incresed hydrogen a c t i v i t y no t h igher temperature could fu r the r in f luence stress corrosion cracking and must be determi ned .

The compatibi 1 i t y o f uncoated z i rconi urn copper (ZRC ) w i t h N2H4 and MMH i s questionab i e a t elevated propel l a n t temperatures, 3 3 Y 0 ~ (140°F), f o r long term use.

5. Environmental Ef fects.

The use of oxygen and hydrogen i n rocket engines w i l l have no s ign i f i can t adverse e f f e c t on the environment b?cduse both are nornlal components i n the a i r and t h e i r only s table react ion product i s water.

The use o f hydrocarbon fue ls such as RJ-5, RP-1, and CH4 w i l l produce minimal environmental e f fec t s , the effects being general ly simi l .ar t o those of j e t a i r c r a f t .

The use of the hydrazine fuels , N?H4 o r MMH, add two add i t iona l points t h a t may present environmental constra ints : (1) these fue ls are very tox i c and some unburned fue l w i l l l o c a l l y enter the atmosphere o r the ear th 's surface and ( 2 ) the react ion products o f oxygen and N2H4 o r MMH w i 11 y i e l d some ni t rogen compounds o f environmental concern (n i t rogen oxides a l d incompletely burned n i t rogen compounds such as amnonia and amines.)

D. COMBUSTION GAS PROPERTY DATA

The TRAt4 72 computer program described i n Ref. 42, was used t o calcu- l a t e theore t ica l one-dimensional equi 1 i b r i um and frozen main chanber and preburner gas propertv data. These data are sumnarized herein.

1. Main Chamber

Main chanber gas stagnation temperature, cha rac te r i s t i c exhaust ve loc i ty , molecular weight, thermal conduct iv i ty , v iscos i t y , s p e c i f i c heat, spec i f i c heat r a t i o , and D i t tus-Boel t e r f ac to r data were parametrical l y ca lcu lated f o r s i x study propel lant combinations (LOX/RJ-5, LOX/RP-1 , LOX/CH4, LOXIN2H4, LOX/MMH and LOXILH2). The data were calculated a t chanber pressures of 68, 136, 272 and 408 atm (1000, 2000, 4000, and 6000 psia.) The mixture r a t i o s ranged from 0.5 t o 1.5 the sto ich iometr ic value. A s u f f i c i e n t number of mixture r a t i o points fo r each prope l lan t combination were run t o permit accurate i n te rpo la t i on of data when stored i n the computer data rout ine. The computer program used (Ref. 42) i s s i m i l a r t o the JANP!AF One Dimensional Equi 1 i b r i um (ODE) model bu t has been extended t o inc lude trans- p o r t property ca lculat ions . The data were calculated f o r the fo l l ow ing mixture r a t i o values:

Propel 1 ant Combination Mixture Ratio, O/F

The proper t ies are sumnarized on Tables X V through XX f o r each o f the propel lant combinations and a representat ive s e t of mixture r a t i o values covering the t o t a l range analyzed. The data show t h a t the c h a d e r pressure in f luence i s small and increases w i t h increasing mixture r a t i o .

2. Preburners

The conbusti on gas proper t ies were a1 so calculated f o r f ue l - r i ch and ox id i ze r - r i ch preburner operat ion w i t h a1 1 s i x study prope l lan t combina- t ions. These data were developed over a chamber pressure range of 136 t o 680 atm. (2000 t o 10,OSO psia) and mixture r a t i o ranges corresponding t o combustion gas temperatures between a t l e a s t 700 t o 1367°K (126C0R t o 2460°R). It should be noted t h a t f o r the; hydrat ine based fuels , (hydrazine and monomethyl -hydrazine) the existence o f a monopropel l a n t flame makes operat ion below a ce r ta in minimum flame temperature impossible. The monopropella,nt flame temperatwe: for N2H and MMH a t 408 atm (6000 ps ia ) are approxlntately d 1000 and 1333°K (1800°R an 240i.1°R), respect ive ly . I n addi t ion, the e f f e c t o f pressure ( f o r the range invest igated) upon the proper t ies o f the ox id izer - r i c h combustion products i s i n s i g n i f i c a n t and can be neglected.

Cha

mbe

r P

res

su

re

1

.. I

Mi x

ture

R

atio

-

9L

L

1.7

2.8

3.7

5.1

1.7

2.8

3.7

5.1

1.7

2.8

3.7

5.1

1.7

2.8

3.7

5.1

TABL

E X

VI.

- L

OXI

RP-

1 M

AIN

CHAM

BER

GAS PROPERTY D

ATA

Ch

arac

teri

stic

V

elo

city

m

/sec

if

t/se

c)

Rat

io o

f S

pec

ific

Hea

ts

Froz

en

Mot

ecu

l ar

Wei

ght

18.9

24.0

26.2

28.4

18.9

24.2

26.5

28.6

18.9

24.5

26.7

28.9

18.9

24.6

26.9

29.0

Cha

mbe

r P

ress

ure

Am. -

Ips

ia)

Mix

ture

R

ati

o

O/F

0.8

1.4

2.0

2.6

0.8

1.4

2.0

2.6

0.8

1.4

2.0

2.6

0.8

1.4

2 -0

2.6

. .

.

, .

*

,-

.

. ,

TABL

E X

VII

. -

LOX/

MM

H M

AIN

CHA

MBE

R rja

S

PRO

PERT

Y SU

mA

RY

Com

bust

ion

Ch

ara

cte

rist

ic

Vel

oci

t ,ii

ET

dGg

Ra

tio

of

Sp

eci

fic

Hea

ts

Qui

1 i bri

urn

Fro

zen

Mo

lecu

lar

Wei

qht

16.5

20.8

23.3

24.9

16.5

20.9

23.4

25.1

16.6

21.1

23.6

25.3

16.6

21.2

23.8

25.4

Chamber

Pre

ssur

e A

tm. - (

psi

a)

Mi x

ture

R

atio

O

iL

TABL

E XV

III.

-

LOX/

N2H4

MR

lN C

HAMB

ER

GAS

PROZ

ERTY

SU

MMAR

Y

Com

bust

i on

Ch

arac

teri

stic

R

atio

of

--

-

--

.

. . -

Tern

era

ture

V

el o

ci t

Sp

ecif

ic H

eats

+

m

e E

qu

ilb

riu

n

-

Froz

en

Wol

ecul

ar

Wei

qh

t

16.0

18.7

20.0

22.1

16.0

18.8

20.1

22.2

16.0

18.8

20.3

22.3

16.0

18.9

20.3

22.4

Cha

mbe

r P

ress

ure

Kim.

(ps

iaj

Mix

tu -e

R

ati

~

O/F

2.0

2.8

3.6

4.5

6.0

2.0

2.8

3.6

4.5

6.0

2.0

2.8

3.6

4.5

6.0

2.0

2.8

3. G

4.5

6.0 .

.

TABL

E XX

. -

LOX/

CH

4 M

AIN

CHA

MBE

R

Com

bus t

i on

Tem

pera

ture

,

OK

-

(OR

)

2557

(4

602)

3391

(6

103)

3569

(6

425)

3534

(6

361)

3379

(6

082)

2562

(4

611)

3451

(6

211)

3666

(6

598)

3627

(6

528)

3453

(6

215)

2564

(4

616)

3504

(6

307)

3762

(6

771)

3721

(6

697)

3526

(6

346)

2566

(4

618)

3532

(6

357)

3818

(6

872)

3774

(6

794)

35

67

(642

0)

Ch

ara

cte

rist

ic

GAS

PROP

ERTY

S

UM

RY

Ra

tio

of

Sp

eci

fic

Hea

ts

Equ

i 1 i b

ri un

Fro

zen

Mo

lecu

lar

Wei

ght

16.0

19.7

22.2

24.0

26.1

16.0

19.8

22.4

24.2

26.3

16.0

19.9

22.6

24.5

26.5

16.0

20.0

22.7

24.6

26.6

The preburner gas roper t ies were a lso calculated using the previously referenced (TRAN 72 ! computer program. The preburner t ransport property ca l cu l a t i on were 1 i m i ted t o the frozen condi t ion because the low gas stagnation temperatures l i m i t thermal disassociat ion o f the combustion products,

Propert ies were calculated f o r nominal i n l e t temperatures o f 298°K (77OF) f o r RJ-5, RP-1, N H4 and MMH and the normal boi 1 i n g p o i n t (NBP) f o r LOX, LH2 and Cd4. I$ a propel lant i s used t o cool the conburtion chanher, the e f f e c t o f p rope l lan t preheating was accounted :or by ad jus t ing the heat o f formation i n the computer program. That i s , the fue l heat of formation was increased f o r fue l - r i ch preburners when the fue l o r ox id i ze r i s used t o cool the chamber and the LOX heat o f formation i s increased f o r ox id izer - r i ch preburners when LOX i s used as a coolant, Because o f the propor t iona l ly small f ue l mass i n jec ted i n the ox id i ze r - r i ch preburners , the e f f e c t o f f ue l preheating was neglected.

Propel lant preheating can be shown as enthalpy increase above the nominal propel 1 an t i n l e t condit ions . The range o f cool ant enthal py increases evaluated was varied f c r each prope l lan t t o account f o r the avai lab le f low r a t e di f ferences.

The ODE equi 1 i brium combustion temperature, cha rac te r i s t i c ve loc i ty , r a t i o o f spec i f i c heats and molecular weight data are sumnarized on Tables X X I through X X I X a t a chamber pressure of 408 atm (6000 ps ia) f o r a l l preburners except the fue l - r i ch LOXIRJ-5, LOXIRP-1 and LOXlCH4 preburners. Chamber pressures of approximately 408 t o 544 atm (6000 t o 8000 ps ia) are t yp i ca l o f the operat ing requirements r e s u l t i n g from t h i s study.

The f u e l - r i c h LOX/hydrocarbon preburner data has been adjusted from the ODE equ i l ib r ium values t o r e f l e c t the experimentally observed nonequi 1 i b r i um performance o f these mixtures . This nonequi 1 i b r i um performance has been empir ical l y v e r i f i e d by many researchers, inc lud ing data reported during the T i tan I engine development program (Ref. 43) .

F i gure 3 compares experimental f ue l -ri ch performance (Ref. 43) t o ODE performance f o r LOXIRP-1 a t a chamber pressure o f 37.4 atm (550 psia). These data were used t o develop the fo l lowing formulas f o r ODE data adjustment.

const. 0

'*exp = C*ODE x n p const. b

TABL

E X

XI.

-

LOX

RJ-

5 O

XID

IZE

R-R

ICH

PRE

BURN

ER

GAS

PRO

PERT

Y DA

TA

SUM

MAR

Y

Pre

ssur

e =

408

atm

(60

00 p

sia

) C

ham

ber

Coo

lant

: LO

X

Mix

ture

R

ati

o

2 5

40

60

2 5

40

60

25

40

60

Pro

pel 1

an

t P

reh

ea

tin

g

(Cha

nqe

in

Com

bust

ion

Tem

pera

ture

OK

(O

R)

Ch

ara

cte

ris

tic

V

elo

city

, m

lsec

(f

t/s

ec

)

Ra

tio

of

Sp

ec

ific

H

eats

1.27

5

1.31

1.32

5

1.27

5

1.31

1 .j

25

1.27

5

1.31

1.32

5

TABL

E X

XI I. -

I.OX/

RP-

1 O

XID

IZE

R-R

ICH

PR

EBUR

NER

GAS

PRO

PERT

Y DA

TA

SUM

MAR

Y

Pre

ssur

e =

408

atm

(60

00 p

sia

) C

ham

ber

Coo

l an

t :

RP-

1

Mo

lecu

lar

Wei

ght

32.3

32.2

32.1

32.3

32.2

32.1

32.3

32.2

32.1

I P

rop

ell

an

t P

reh

ea

tin

a

I I

I (c

hang

e i;;

I

Com

bust

ion

I C

ha

rac

teri

sti

c

I R

ati

o o

f I

I::

IP

(305

0)

1.27

31

.8

I (2

600)

(1

675)

I ::

: (2

405)

I 1

.302

1 3

1.9

Mix

ture

R

ati

o

-~n

tha

l py)

Cal

. (B

tu/l

b)

Tem

pera

ture

OK

(O

R)

Ve

loci

ty,

m/ s

ec

(ftl

se

c)

Spe

c i f i c

Hea

ts

Mo1

ecu

l ar

Wei

ght

Mix

ture

R

ati

o

TABL

E X

XII

I. -

LOX/

Mt4H

O

XID

IZER

-RIC

H

PREB

URNE

R G

AS P

ROPE

RTY

DATA

SU

MM

ARY

Pre

ssur

e =

408

atm

(60

00 p

sia

) C

ham

ber

Coo

lant

: MM

H

TABL

E XX

IV.

- LO

X/N

2H4

OX

IDIZ

ER

-RIC

H

PREB

URNE

R G

AS P

ROPE

RTY

DATA

SU

MM

ARY

Pro

pel 1

an

t P

reh

ea

tin

g

(Cha

nge

in

En t

ha 1

py )

Cal

. (B

tu/l

b)

Pre

ssur

e =

408

atm

(60

00 p

sia

) C

ham

ber

Coo

lant

: N2

H4

Com

bust

ion

Tem

pera

ture

K

(OR

)

Mo

lecu

lar

Wei

ght

29.3

30.1

30.5

Mix

ture

R

ati

o

Ch

ara

cte

ris

tic

V

el o

ci t

y ,

m/ s

ec

(ftl

se

c)

Pro

pel 1

an

t P

reh

ea

tin

g

(Cha

nge

in

En

tha

l py)

C

al .

(Btu

/l b

)

Ra

tio

of

Sp

ec

ific

H

eats

M

ole

cula

r W

eigh

t

Com

bust

ion

Tem

pera

ture

OK

(OR

)

Ch

ara

cte

ris

tic

V

elo

city

, m

/sec

( f

tls

ec

)

Ra

tio

of

Sp

ec

ific

H

eats

m o o N m o

n n n h h F 0 0 h C D ( 3 C O r J m - ~ Y V V

VI L

n 0 aJ c u m -

CO 0 h - F

h h nnn - 0 a m 0 * ' a m m * I n r g I n r g h * * * * d U V U V Y

I -

I n n n n n n h h h

m m a o c o m o a m 0 - 0 0 c o b - m o m

e lnut m w r . U Y U N @ J c u V V W W U - ' - ' %

OC W z O C * 3e w"% OZE a=>

v, I v e z!z

1 n A w > Dl- L O C

W *n I 0 cue z n \ Xv,

3s H H u > X X

W A m e l-

e e m m u m u e m m m m m m m m m m . . . . . . . . . P P P F P - 7 - 7

n n n n n n n n n h a m C O b a O m 0 F - m a h N N m m m W e h N O 0 3 m m c o

e U e W W Y U V V U V V

h h h n-n n n n

P ( U m m Y N F e a r n h ~ m

0 - m O - m o w e - P C F F F F F F

aJ L 0 3 C,C,

2 5L

- m o F ~ O F V , O 0 0 - 0 0 - 0 0 - . . . . . . d d d 0 0 0 o o o

TABL

E X

XIX

. -

LOX/

LH9

FUEL

-RIC

H

PREB

URNE

R GA

S- P

R~P

ER

TY DA

TA

SUM

MAR

Y

Pre

ssur

e =

408

atm

(60

00 p

sia

) C

ham

ber

Coo

lant

: LH

2

Mix

ture

R

ati

o

0.5

1 .o

1.

5

0.5

1 .o

1.

5

0.5

1 .o

1.

5

Pro

pel 1

an

t P

reh

ea

tin

g

(Cha

nge

in

Com

bust

ion

Tem

pera

ture

O

K

f OR)

Ch

ara

cte

rist

ic

Ve

loci

ty,

m/s

ec

(ft/

se

c)

Ra

tio

of

Spe

c? C

i c

Hea

ts

1.39

1.36

1.32

1.39

1.35

1.31

1.39

1.35

1.31

Mo

lecu

lar

Wei

qht

3.04

4.03

5.04

3.04

4.03

5.04

3.04

4.03

5.04

Mixture Rat io , O/F

Figure 3. LOXIRP-1 Experimental Preburner Characteri s l i cs

- L L (2000) 0 u

Y 3

I)

0,

5 (1600)-- C, cC1 L

E QJ I- I= (1200) 0 *r C1 ln 3 n E U

(800) -,

-r

0.20 0.24 0.28 0.32 0.36 0.40 0.44 -

Mixture Ratio, O/F

400 -. -! 1300

1200

] loo- -

1000 -- --goo.-

800 - 700 - -

Pressure = 37.4 atm (550 psia)

*-

--

-

I 1 I

I 1 I

1 I

where:

To (or CI = cor rec ted exper imenta l ly based conbustion eXP exP temperature o r c h a r a c t e r i s t i c ve l oc i ty value

) = ODE combustion temperature o r c h a r a c t e r i s t i c (or '*ODE v e l o c i t y value OODE

rl To

= experimental e f f i c i e n c y f a c t o r def ined a t any equivalence r a t i o by d i v i d i n g the experimental

const. 0 To ( o r C*) value by t he ODE To ( o r C*) value

Equ.i,valence r a t i o = S t o i c h i o n e t r i c O/F d i v i ded by design O/F.

has be speci f ana ly t des i gn

cannot

E f f i c i e n c y f ac to r s were developed versus equivalence r a t i o f rom the data o f F igure 3 and used t o p r e d i c t To and C* values a t h igher chamber pressures. The LOX/RP-1 f ac to r s were a lso assumed t o be v a l i d f o r both the LOXIRJ-5 and LOX/CH4 p rope l l an t combinations. This data i s presented on Tables X X X , X X X I and X X X I I . The emp i r i ca l f u e l - r i c h gas generator p roper ty data, which

en der ived from t e s t r esu l t s , a l so r e s u l t s i n a much lower r a t i o of i c heats and a h igher molecular weight than a re p red ic ted by the ODE i c a l model. For example, the a n a l y t i c a l and experimental data a t a p o i n t o f 408 atm (6000 ps i a ) a re compared below:

LOX/RP-1 Experimental Theoret i ca l

Mix ture Ra t i o 0.22 0.22 Combustion Temp., "K ( O R ) 867 (1560) 1172 (2110) Molecular Weight 29.83 24.2 Ra t i o o f Spec i f i c Heats, Y 1.095 1.172

The experimental data d i f f e r s because the a n a l y t i c a l models accurate ly p r e d i c t the composit ion of the exhaust products, The

experimental data shown on Figure 4 has been e m p i r i c a l l y der ived from the r e s u l t s o f T i t a n I and A t l as type ecgine gas generator t e s t i n g i n a pressure range from 27.2 t o 68 atm (400 t o 1,000 ps i a ) . However, component designers fami 1 i a r w i t h the past work f e e l t h a t the exper imenta l ly d e r i ~ e d Y i s con- se rva t i ve . Because no experimental data e x i s t s a t h i gh pressure, the low pressure data has been used t o ad jus t the t h e r o e t i c a l p roper ty data pre- d i c t i o n s . The experimental l y de r i ved molecular weight was n o t assumed t o vary w i t h pressure and the spec i f i c heat a t constant pressure was c a l - cu la ted from the p e r f e c t gas law.

Mix

ture

R

ati

o

0.10

0.30

0.50

0.10

0.30

0.50

0.10

0.30

0.50

TABL

E XX

X. - L

OX

/RJ-

5 FU

EL-

RIC

H

PREB

URNE

R G

AS P

ROPE

RTY

DATA

S

UW

RY

Pre

ssur

e =

408

atm

(60

00 p

sia

j C

ham

ber

Coo

lant

: R

J-5

Pro

pel 1

an

t P

reh

ea

tin

g

(Cha

nge

in

I Com

bust

ion

-~n

tha

l py

) C

al .

(Btu

/l b

)

Ch

ara

cte

ris

tic

V

elo

city

, m

/sec

(f

t/s

ec

) _

Ra

tio

of

- S

pe

cif

ic

Hea

ts

1-0

6

1 .I

35

1 .I

85

1.06

1 .I3

5

1 .I

85

1.06

1.13

5

1 .I

85

Mol

ecu

l ar

Wei

qht

42.7

33.8

23.0

42.7

33.8

23.0

42.7

33.8

23.0

a L 0 a UC, xci

CZ I:

? m u , ? m ~ ) - C * I V ) . o o o d d d d d d

, Ratio of Specific Heats

N N N N N W W MW, Molecular Weight C-' 0 N P Q, 03 0 N P OI %

0 W

T, Combustion Temp., OK --I

E. MAIN CHAMBER THEORETICAL PERFORMANCE DATA

One-dimensional equi 1 i b r i um (ODE) and frozen (ODF) sea-level and vacuum spec i f i c impulse were calculated over the same chamber pressure an4 mixture r a t i o ranges as the main chamber combustion gas property data. Performance was calculated f o r expansion area r a t i o s ranging from 1 : 1 t o 400:l. ODF performance i s between f i v e and e igh t percent below ODE performance f o r a11 prope l lan t combinations. ODF performance would y i e l d erroneous spec i f i c impulse values and nonoptimum TCA mixture r a t i o s if u t i l i z e d . A t yp i ca l comparison of ODE and ODF performance i s shown on Figure 5 f o r LOX/RP-1. Therefore, the ODE performance was used t o conduct a l l analyses i n t h i s study e f f o r t .

For each study propel 1 ant combination, ODE vacuum spec i f i c impulse was i n i t i a l l y p l o t t e d versus mixture r a t i o f o r various values of chamber pressure. A t h igh t h r u s t chamber pressure, 272 atm (4000 psia), the optimum sea-level performance occurs a t nozzle area r a t i o o f approximately 40:1, Therefore, t h i s was the spec i f ied basel ine area r a t i o i n the study. ODE vacuum spec i f i c impulse vs mixture r a t i o for a l l s i x propel lant combinations i s presented on Figures 6 through 11.

Sea-level and vacuum ODE performance i s presented as a func t ion o f nozzle area ;-atio f o r approximately optimum mixture r a t i o s f o r the Mode 1 propel lants and a t an O/F = 7.0 f o r LOX/LH2 on Figures 12 through 17.

The propel l a n t temperature used i n t h i s performance eval uat ion were:

" Oxygen, NBP: 90.Z01( (162.4"R) " KP-1: 298°K (537"R) " RJ-5: 298°K (537"R) O MMH : 298OK (537"R) " Hydrazine: 298°K (537"R) O Hydrogen, NBP: 20.3"K (36.5"R) " Methane, NBP: 112°K (201.6OR)

Cham

ber

Pre

ssur

e =

272

atm

(40

00 p

sia)

No

zzle

Are

a R

atio

, E

=

40:1

1

I I

I I I

I I

I I

I I 1

I 1 .

I I I I

6 1.

8 2.

0 2.

2 2.

4 2.

6 2.

8 3.

0 3.

2 3.

4 3.

6 3.

8 M

ixtu

re R

atio

Fig

ure

5.

LOX/

RP-1

OD

E an

d Fr

ozen

Per

form

ance

Com

pari

son

PC =

Cha

mbe

r P

ress

ure

E

= N

ozz

le A

rea

Ra

tio

I

I

I

I

I I

t I

4 m

2.0

2.4

2.8

3.2

3.6

4.0

4.

4 4

.8

5.2

5.6

6.0

M

ixtu

re R

atio

, O

/F

atm

(p

sia)

408

(600

0)

272

(400

0)

136

(200

0)

68

(100

0)

Fig

ure

11.

LO

X/CH

4 OD

E Va

cuum

Per

form

ance

Ver

sus

O/F

Vacuum

.-A

PC, atm

1 2 4 6 10 20 40 60 100 200 400 Nozzle Expansion Area Rat io , Ac/AT

Figure 12. LOXIRJ-5 ODE Performance Versus Area Rat io

O/F = 2.9:1 Vacuum

Sea-Level

I I I L

1 1 I I ;-

6 10 2 0 40 60 100 2 00 400 Nozz;e Expansion Area Rat io , A,/At

Figure 13. LOX/RP-1 ODE Performance Versus Area Rat io

Vacuum --

Sea-Level

P c atm ( p s i a )

200 ! 1 L I I I

I 1 1 I I 1 I

1 2 4 6 1b 20 200 400 Nozzle Expansion Area Rat!o, AE/AT

Figure 14. L O X / W ODE Performance Versus Area Rat io

Sea-Level

2 00 I I

1 2 4 6 lb 20 40 60 -"d2b0 1 0 400 Area Rat io , AE/AT

Figure 15. LOX/NZH4 ODE Performance Versus Area Rat io

2641 I 6 . I I 1 1 . I l 1 1

,-.-- I +- -I

2 2 0 40 60 80 100 ' 499 Nozzle Expansion Area Rat io - AE/AT

Figure 16. LOX/LH2 ODE Performance Versus Area Ratio

69

atm (ps ia )

Sea Level I 1 I

1 I I I -

1 2 4 6 1 0 2G 4'9 60 160 2 00 40C Nozzle Expansion Area Rat io , AE/AT

Figure 17. LOX/CHq ODE Per fw .~ . .~ l :ce Versus Area Rat io

SECTION I V

TASK I I - COOLANT EVALUATION


The object ives o f t h i s task were to : (1) def ine the t h r u s t chamber geometry and mixture r a t i o f o r the various candidate Mode 1 prope l lan t cwb ina t ions and engine cycles, and (2) t o es tab l ish the r e l a t i v e cool ing capabi 1 i ty o f the candidate Mode 1 fuels, oxygen and hydrogen.

The fo l lowing propel 1 an t combinations and cool ants were considered:

Propel 1 ants Cool ant Cycle

LOX/RJ-5 OW gen Staged Combustion LOXIRJ-5 Hydrogen Staged Combustion LOXIRJ-5 Hydrogen Gas Generator LOXIRJ-5 RJ-5 Staged Combustion LOXIRP-1 RP-1 Staged Combustion LOXIHydrazi ne Hydrazi ne Staged Colnbustion LOXIMMH MMH s ta e Combustion LOXlCH4 CH4 Stale! Combust i on

The two hydrogen cool ing studies d i f f e r e d only i n the length o f the co!nh!stion chamber; the gas genzrator cyc le required 22.9 cm (9 inches) addi t ional length t o accommodate h igh densi ty propel lant i n jec t i on .

Parametric studies over the chamber pressure range from 136 atm t o 340 atm (2000 ps ia t o 5000 psia) were required, w i th ihe re1ati:e mer i t o f thd various coolants based on the a t ta inab le chamber pressure as determined by pressure drop requirements. A serv ice l i f e o f 250 cycles was specif ied; design c r i t e r i a and coolant evaluat ion a lso considered 1 imi ta t ions such as coking o f the hydrocarbon fue l s and c a t a l y t i c decomposition o f MMH and hydrazi ne.

Addit ional Task I 1 guidel ines provided by NASA/LeRC are given i n Table X X X I I I and Figures 18 through 21. Rectangular channel construct ion was spec i f ied i n the high heat f l u x p a r t o f the char&er using a zirconium- copper a1 1 oy . Table X X X I I I provides channel dimension and wa l l thickness l i m i t s plus i n l e t pressures and temperatures. Figures 18 through 21 show the zirconium-copper proper t ies used i n the study.

B. TCA GEOMETRY AND MIATURE RATIO SELECTION

The TCA geometry and mixture r a t i o were i n i t i a l l y selected f o r the Mode 1 LOA '?J-5 base1 i n e engine. The approach and resu l t s are discussed herein .

TABLE X X X I I I . - COOLANT EVALUATION STUDY GUIDELINES

Cool ant I n l e t

Propel 1 an t Combination Coolant

Avai lab le Temp. Coolant - " K (OR)

RJ-5 Dxygen Oxygen Total Flow 111 (200)

RJ-5IOxygen RJ-5 Total Flow 311 (560)

RJ-5/Oxygen Hydrogen (Minimize) 61 (111))

RP-1 /Oxygen RP-1 Total Flow 311 (560)

Hydrazi ne/Oxygen Hydrazi ne Total Flow 311 (560)

MMH/Oxygen MMH Total Flow 31 i (560)

CH4/Oxygen CH4 Total Flow 144 (260)

THRUST CHAMBER

" Regenerati vel y Cooled

O High Heat Flux Port ion o f Chamber Shal l be o f Nontubular Construction w i t h the Following Dimensional L imi ts :

Minimum S l o t Width = 0.762 mn (0.03 i n . )

Maximum S l o t Depth/Width = 4 t o I*

Minimum Land Thickness = 0.762 mm (0.03 i n . )

Minimum Wall Thickness = 0.635 mn (0.025 i n . )

Mater ial (Nontubular Port ion) : Copper A1 l o y (Zirconium Copper) Conforming t o the propert ies given Figures 18, 19, 20 and 21

O Service Free L i fe : 250 Cycles Times a Safety Factor of 4 O Possible Benef i t o f Carbon Deposit ion on Hot Gas

Side Wall s h a l l be Neglected

*Applied Herein from the I n j e c t o r End t o Area Ratio 1.5:l.

k, STRENGTH COEFFICIENT, MN/~' (ksi )

Reduction of Area, % RA

E, MODULUS OF ELASTICITY, GN/~' (lo3 ksi )

I-, Strain Hardening Exponent

Creep-Rupture

-- Rupture, hrs

1 1 00

Range: 2.2"K t o 866.7"K (-456°F t o 1100°F)

Low-Cycle Fatigue

I- 1 I

100 1000 10,000 Cycles to Failure

Figure 20. Creep-Rupture and Low Cycle Fatigue

--

I I

I I

1 I

I 33

.1

255.

4 47

7.8

700.

0 92

2.2

(-4

00

) (0

) (4

00

) (8

00)

(1 23

0)

TEM

PERA

TURE

O

K (O

F)

Figu

re 2

1.

Cond

ucti

vity

and

Exp

ansi

on

The Mode 1 LOXIRJ-5 engine u t i l i z e s a staged combustion cyc le comprised o f p a r a l l e l f u e l - r i c h and ox id i ze r - r i ch preburners and a gaslgas i n jec ted primary th rus t chamber. Simp1 i f i e d gaslgas mixing analyses were conducted w i t h an ALRC developed gaslgas mixing model (Ref. 44). I n j e c t o r energy release e f f i c i ency (ERE) was evaluated as a func t ion o f chamber length ( L ' ), chamber pressure (PC), chamber contract ion r a t i o (%), mixture r a t i o , and i n j e c t o r element type and pressure drop. The ERE goal used i n the study i s 98%.

The analysis was i n i t i a t e d by se lec t ing an i n i t i a l design po in t and evaluat ing i n j e c t o r ERE as a funct ion o f chamber length f o r three i n j e c t o r element types. The elements evaluated were a shear coaxial , a fue l -ox id izer - fue l (F-0-7) external impinging t r i p l e t , and a premix t r i p l e t . I n j e c t o r element s ize ( i .e., t h r u s t per element), o f course, a lso af fects performance. The chamber length study was conducted for a constant t h rus t per element o f 2233N (502 l b f ) which r e s u l t s i n 1309 elements a t a vacuum th rus t l e v e l o f 2.92 MN (656.400 l b f ) . This element s ize was selected based on ALRC SSME an3 M-1 design experience.

Figure 22 shows ERE versus chamber length f o r the shear coaxial element and a1 so notes the assumed i n i t i a l design condit ions. Minimum coaxial performance i s predicted f o r the equal pressure drops i n the fuel and ox id i ze r i n j e c t i o n o r i f i ces . This occurs because the f u e l - r i c h and ox id i ze r - r i ch gases have near equal dens i t ies resul t i n g i n near equal i n j e c t i o n ve loc i t i es which 1 i m i t s tu rbu len t shear mixing. Engine pressure schedule c r i t e r i a a1 so d i c t a t e equal pressure drops. Therefore, t h i s case was selected f o r analysis. Figure 22 indicates a maximum chamber length requirement o f 20.3 t o 22.9 cm (8 t o 9 inches) t o achieve a 98% ERE goal.

S imi la r anal2,es were conducted f o r the F-0-F t r i p l e t and premix t r i p l e t elemenxs. The F-0-F t r i p 1 e t chamber length requirement i s approximately 15.2 cm ( s i x inches). The premix element impinges two rectangular fue l s l o t s normal t o the c i r c u l a r ox id izer element below the i n j e c t o r face plane. The fue l and ox id izer then mix i n the o r i f i c e cup f o r 2 t o 3 cup diameters before being in jected. The premix element shows high performance requ i r i ng only 8.9 t o 14 cm (3.5 t o 5.5 inches) t o reach the study performance goal.

The r e s u l t s ind ica te t h a t f o r a given performance ieve l , the longest and hence, heaviest chamber i s required by the shear coaxial element. However, t h i s element has the advantage o f producing excel l e n t chamber compa t ib i l i t y and i s wel l modeled empi r ica l l y . Thi: would r e s u l t i n r e l a t i v e l y low DDT&E cost. The t r i p l e t and premix elements requ i re shor ter and 1 i gh te r chambers t o achieve the performance goal but are a1 so higher r i s k . Both can producd heat t rans fer problems; the external impinging t r i p l e t can produce chamber streaking and the premix produces r e l a t i v e l y high i n j e c t o r face heat f luxes. The premix presents an addi t ional problem f o r the Mode 1 engine; the 811°K (lOOO°F) ox ia i ze r - r i ch and f u e l - r i c h gases are near ly hypergol i c , according t o p re l iminary analysis. Any combust ion i n the face mixing cup would r e s u l t i n element development problems. An add i t iona l considerat ion i s t h a t f o r the l a rge Mode 1 engine th rus t l eve l ,

Energy Release Eff. %

Combustion Pressure Loss % of Plenum Pressure

-I

-1 VI 0 cn I

a 1

1 - Energy Release Eff. % 4

the shear coax ia l element would by f a r be the eas ies t element t c incorpora te i n t o a non-complex man i fo ld /pa t te rn design. For the reasons mentioned, the shear coaxi a1 element and associated chamber leng th was selected.

Figure 22 shows t h a t the requ i red chamber leng th t o achieve an energy re lease e f f i c i m c y o f 98% i s approximately 9" . The chamber cons is ts o f a c y l i n d r i c a l ser t i o n and a 30" convergent sect ion. P re l im inary 1a.vouts o f the chamber using proper r a d i i o f curvature i n the corners and a t the t h r o a t r e s u l t s i n a s l i g h t chamber leng th increase t o 24.7 cm (9.74") . This small increase i n o v e r a l l 5ength i s p re fe r red t o reduc ing the c y l i n d r i c a l sec t ion length any fu r ther . Hence, f o r the parametr ic study, the chamber was s p l i t i n t o a constant leng th c y l i n d r i c a l sec t ion and a va r i ab l e length convergent sec t ion which i s p ropor t iona l t o the t h r o a t rad ius . I n t h i s fashion, the convergent angle remai ns approximately constant .

The in f luences of m ix tu re r a t i o , chamber con t rac t i on r a t i o , and chamber pressure on ERE were a l so determined f o r a f i x e d chamber length. M ix tu re r a t i o does no t a f f e c t ERE s i g n i f i c a n t l y from 2.0 t o 3.0:1 ( t k an t i c i pa ted design range). F igure 23 shows t h a t ERE increases as chamber con t rac t i on r a t i o ( € c ) decreases. The s e l e c t i o n o f t he design chamber con- t r a c t i o n r a t i o was tempered w i t h the knowledge t h a t t he Rayleigh l i n e combus t ion pressure l o s s increases sharp ly w i t h decreasing con t rac t i on r a t i o as shown on F igure 24. A design con t rac t i on r a t i o value o f 2.5:1 was se lected t o m i r~ im ize tne combustion pressure l o s s and chamber weight and t o a t t a i n near maximum performance.

The TCA geometry se l ec t i on process sumnarized here in f o r t he LOX/ RJ-5 Mode 1 base1 i n e was repeated f o r the o the r Task I 1 propel l a n t com- b ina t ions . The shear coax ia l element was used f o r c11 enyine performance s tud ies except t he hydrogen-cooled gas generator cyc i e which u t i l i z e s l i q u i d l l i q u i d TCA p rope l l an t i n j e c t i o n . In jec to r , f o r t h i s engine cyc l e must have the a b i l i t y t o produce f i n e l y atomized p rope l l an t sprays t o maximize perfomance. Th is r e s u l t s i n an increased ~hamber l eng th requ i re - ment.

M ix tu re r a t i o se lec t ions were made on the bas is o f maximum de l i ve red performance. For the Task I I analyses, de l i vered per fomance was assumed t o be 97% (98% ERC p lus 1% f o r o ther losses) of t he t h e o r e t i c a l vacuum value. Th's was substant ia ted i n l a t e r d e t a i l e d analyses.

The mix tu re r a t i o se lec t ions f o r the var ious p rop2 i l an t combinations a re based upon the de l i ve red performance ( I s ) curvzs shown on Figures 25, 26 and 27. Both engine and veh i c l e design cons iderat ions f a v m opera t ing the LOX cooled engine t o t h e " r i g h t " o f the peak s ince t h i s r e s u l t s i n both more LOX f o r coo l i ng and a h igher bu lk dens i t y . From an engine design viewpoint, i t i s more des i r ab le t o operate t he f u e l cooled engines t o t h e " l e f t " o f the peak t o increase t he f ue l a v a i l a b l e f o r coo l ing . This, however, r e s u l t s i r reduced p rope l l an t bu lk dens i ty . I n o rder t o avo id b i as i ng t he cool i n g study i n f avo r o f e i t h e r the LOX o r f u e l cooled systems, m ix tu re r a t i o s were se lected a t the maximum s p e c i f i c impul se magni tudes.

The TCA geometry c r i t e r i a and mix tu re r a t i o se lec t ions a re summarized on Table X X X I V .

C. LOW CYCLE FATIGUE ANALYSIS

Low cyc l e t h r u s t chamber thermal f a t i g u e analyses were conducted i n con junct ion w i t h the coo lant heat t r a n s f e r eva lua t ion t o e s t a b l i s h t he chamber pressure upper l i m i t , i f any, created by t h e chamber 1 i f e requ i re - ment.

Past experience has shown t h a t t he t h r o a t reg ion i s t he most c r i t i c a l area from a f a t i g u e l i f e standpoint . Therefore, analyses t o determine the maximum a l lowable w a l l temperature f o r a design chamber 1 i f e of 1000 cyc les (250 cyc l2s t imes a sa fe ty f a c t o r o f f o u r ) were concentrated a t a sec t ion through t he chamber t h roa t . The p rope r t i es of chamber mate r ia l , Z R C ~ (z i rcon ium copper), a re shown on Figures 13 through 21 .

The i n i t i a l ana lys is was conducted f o r t he LOX/RJ-5 basel ir,e engine a t a chamber pressure o f 272 atm (4000 ps i a ) . An approximate channel cross- sec t ion was se lected and t he pressures and temperature estimated. The s t ress and s t r a i n were ca lcu la ted along w i t h estimate:. o f creep damage and cyc l e l i f e . The assumed channel geometry was then mod i f ied u n t i l a s a t i s f a c t o r y design was achceved. The thermal g rad ien ts were def ined by t he heat t r ans fe r ana lys is and a plane s t r a i n f i n i t e element model was c o n s t r u ~ l e d . Maximum e f f e c t i v e s t r a i n s were determined f o r t b m n a l grad ients and compared t o t he cyc l e 1 i f e data f o r ZRC~. Frorn these con~.:risons, the expected cyc l e l i f e was estab l ished. Th is procedure was repeated f o r the o ther chamber pressure;, 136-340 atm (2000-5000 ps:,) and the o ther study p rope l l an t combinatiuns.

The r e s u l t s of t he low cyc l e f a t i g u e analyses a re summarized on F igure 28 f o r t he base l ine LOX/RJ-5 engine a t 272 auv (4000 ps i a ) . This f i g u r e shows t he maximum permi s s i b l e temperature d i f Ference between the hot gds s i de wa' 1 sur face and the p rope l l an t bu lk temperature ( e s s e n t i a l l y the backside wa l l temperature) f o r a v a r i a t i o n i n c y c l e l i t e . For a design cyc l e l i f e requirement o f 1090 cycles, the maximum rT i s 867°K (llOO°F). I n order t o con ta in t he h igh pressuye i w " i v e d , the s t r u c t u r a l c loseout would have t o be approximately 0.762 cm (0.30 i n . ) t h i c k . The recomnended c loseout ma te r i a l i s n i c k e l .

Fur ther calculation^ a t a chamber pressure o f 346 a t n ~ (5W0 ps ia ) revealed t h a t t he rna;;imum e f f e c t i v e s t r a i n was on l y m in ima l l y a f fec ted . Therefore, F igure 28 was considered t 3 be v a l i d f o r t he chamb?r pressure parametr ic analyses. The 1 i m i t i n g temperature d i f f e r e n t i a l c r i t e r i a $!as a l so found t o be v a l i d f o r the o ther p rope l l an t combinations c o n s i k r e d pa rame t r i ca l l y i n t he Task ? T t hen ia l ana lys is .

Pro

pel 1

an

t C

ombi

natio

n

LOX/

RJ-

5

LOX/

RP-

1

LOX/M

MH

LOX/

N2H

4

LOX/

CH4

LOX/

RJ-

5

TABL

E XX

XIV.

- T

ASK

I1 M

ODE

1 EN

GIN

E GE

OMET

RY

AND

MIX

T'JR

E R

ATIO

SEL

ECTI

ON

SUMM

ARY

Eng

i ne

Cyc

l e

Stg

. Co

mb.

Stg

. Co

mb.

Stg

. Co

mb.

Stg

. Co

mb.

Stg

. Co

mb.

Gas

Gen

.

Cha

mbe

r C

ontr

acti

on

O/F

R

ati

o

2.7

2.5

2.9

2.5

1.6

2.5

1 .O

2.5

3.5

2.5

2.7

2.5

Ma1 n

P

rope

l la

nt

Inje

cti

on

Gas

-Gas

Gas

-Gas

Gas

-Gas

Gas

-Gas

Gas

-Gas

Liq

uid

-Liq

uid

Ls

= C

ham

ber

Leng

th =

Cyl

ind

rica

l Le

ngth

+ C

onic

al S

ecti

on L

engt

h

L' cm

-.

( Inc

hes 1

7.62

t1.3

03

Rt

(3t1

.303

Rt

)

7.62

t1.3

03

Rt

(3t1

.303

R

t)

7.62

+1.5

,3

Rt

(3t1

.303

Rt

)

7.62

+1.3

03

Rt

(3t1

.303

R

t)

7.62

+1.3

03

Rt

( 3+1

.303

Rt

)

30.5

t1.3

03

Rt

(12t

1.30

3 R

t)

Temperature Differential (AT), OF

Temperature Differential (AT), OK

A d A d 4 Q, Q) 0 rU P Q,

0 8 o

0 0 0 0

0 0 0 0 0

i D. THERMAL ANALYSIS / / i

Detai led chamber designs were developed f o r the f o l lowing combinations I D

o f coolant and chamber pressure: : I ! . I

Cool ant Pressure

Atm. - O W W 136, 204, 272, 340 (2000, 3000, 4000, 5000) Hydrogen (S . C . Cycle 136, 238, 340 2000 3500, 5000) Hydrogen (6.6. t y c l e i 238 3500

I 136, 238, 340 2000, 3500, 5000 Hydrazi ne MMH 136, 238, 340 2000, 3500, 5000 Methane 136, 272, 340 2000, 4000, 5000 r i i

i - The chamber geometry and operating mixture ra t ios were establ i shed as

described previously i n t h i s section. ! !

Results f o r the hydrogen-cooled gas generator cycle a t chamber pressures . . other than 3500 psia were obtained by scal ing the 3500 psia data based on the corresponding staged combusti on cycle resul ts . i j i i

: i A1 1 designs are based on straddle-mi 11 machining wi th a constant land

width o f 1.02 mn (0.040 i n .) Based on channel optimization studies f o r oxygen cooling, the 4:1 channel depthlwidth l i m i t o f the guidelines was applied i n 1 I

! 1 the throat region. Therefore, applying the thermal design c r i t e r i a f o r each concept t o the throat region determined the number o f channels, and applying them elsewhere determined the local channel depth. The design c r i t e r i a which i 1

i I contro l led the various conce~ts are sumnarized i n Table XXXV. Plane s t r a i n L J

i

f i n i t e element analyses by the indicated cycle l i f e c r i te r ion , could be sa t i s f ied by 1 i m i t i n g the loca l dif ference between the maximum gas-side wall temperature and the essent ia l ly uniform external wal l temperature t o 867OK

t i ] &

(1100°F). The 58g°K (600°F) coking l i m i t on coolant-side wall temperature prevented the development o f pract ica l designs wi th RJ-5 or RP-1 cooling.

I

TABLE XXXV. - TASK I1 CHAMBER DESIGN LIMITS

Cool an t C r i t e r i a

Oxygen and CH4 ATW 2 867.K (1 100°F) Cycle L i fe

Hydrogen hTw T,< 867'K (1 100°F) Cycle L i f e

< 867OK (1100°F) Strength Tws - RJ-5, RP-1 < 589°K (600°F) Coking Twc - w, N2H4 < 644OK (700°F) Coolant C r i t i c a l Temperature Twc -

ATw = Wall temperature d i f f e r e n t i a l

Tws = Gas-side w a l l temperature

Twc = Coolant s ide wa l l temperature

1, Chamber Wall Construction

A1 though tubes could be used t o save weight i n the low heat f l u x p a r t o f the nozzle, rectangular channel cons t ruc t i on was assumed throughout t o s imp l i f y the Task 11 analyses. A design was selected which i s p rac t i ca l t o fabr ica te y e t provides the f low area va r ia t i on desired f o r e f f i c i e n t cool - i n g w i t h i n the channel depthlwidth cons t ra in t o f 4: 1 . Straddle-mi 11 machining, which y i e l d s a constant land width, i s used except i n the a f t end where constant width channels are proposed. The nunber of channels i s doubled when the channel width becomes too large from a s t ruc tu ra l standpoint. For t h i s task, t h i s p o i n t was selected a t an area r a t i o o f 7.6:l. Straddle m i l l i n g would be terminated when the channel width again becomes too la rge ,'or e l se a t the p o i n t where s t radd le m i l l i n g cannot form the e n t i r e channel. However, t o s imp l i f y t h i s analyses, a constant land width was used over the e n t i r e length.

A land width o f 1.02 mm (0.040 in . ) was selected f o r a l l designs along w i t h the minimum specif ied gas-side wa l l thickness o f 0.635 mn (0.025 in.) per the study guidel ines and an external wa l l thickness of 1.52 mn (0 .O6O i n .) Channel opt imizat ion studies w i t h oxygen cool ing ind ica ted there was a r e l a t i v e l y small advantage i n going t o the minimum spec i f ied land width o f 0.762 mm (0.030 i n , ) , St ruc tura l analyses i nd i ca te a much th i cke r external wa l l i s required, but i t i s 1 i ke ly t h i s wa l l would be made o f a lower con- d u c t i v i ty mater ia l such as n icke l . I n t h a t case, the conductance of the external wa l l might be less than assumed herein; however, t h i s change would have no e f f e c t on the designs w i t h oxygen, hydrazine, methane, and MMH cool ing and only a small e f f e c t on the hydrogen-cooled designs.

The channel geometry parameters which remained t o be determined f o r each design were the numbei rrf channels and the channel depth ax ia l p r o f i l e . With the land width f i x e d and the channel depth l i m i t e d t o four times the channel width, the maximum l o c a l coolant flow area was s e t by the nunber o f channels. Channel opt imizat ion studies w i t h oxygen cool ing indicated t h a t i t was desi rable t o design a t the channel depthlwidth l i m i t o f four. However, t h i s could be accomplished a t only one ax ia l posi t ion. A t other locat ions i t was necessary t o s a t i s f y the thermal design c r i t e r i a w i t h lower depth/width r a t i o s o r t o overcool , i .e. , n o t reach the appl icable wa l l temperature l i m i t s . I n order t o avoid overcool ing i n h igh f l u x regions, the number of channels i n each design was se t by s a t i s f y i n g the design c r i t e r i a a t the th roa t w i t h a channel depth/width r a t i o close t o four . Lower r a t i o s then resu l ted from s a t i s f y i n g the design c r i t e r i a between the th roa t and the i n j e c t o r end. The resu l tan t nunber o f channels f o r each design i s given i n Table X X X V I .

I n a l l designs the coolant enters a t area r a t i o 1.5:1, f lows through the th roa t region t o the i n j e c t o r end and then through the nozzle from 1.5: 1 t o 40:l. Straight-through f low paths from each end were i n - vest igated f o r oxygen cool ing b u t requi red much higher pressure drops, A l l pressure aro calculat ions were based on a surface roughness of ,002 mn (80 m i c r o i nches . !

Cool ant

Oxygen

TABLE X X X V I . - NUMBER OF COOLANT CHANNELS I N COMBUSTION CHAMBER*

Chamber Pressure, aim (psia)

136 (2000) 204 (3000) 238 (3500) 272 (4000) 340 (5000)

250 260 260 260

Hydrogen (S.C. Cycle) (Para1 1 el Burn) 450

Hydrogen (6.6. Cycle)

Hydrazi ne

MMH

Methane

q h e number o f channels i s doubled i n the nozzle a t area r a t i o 7.6:l

2 . Methods o f Analysis

Design data presented herein were generated w i th a r e enerative- 2 cool ing program simi 1 ar t o the computer program suppl i ed t o NASA- ewis under Contrdct NAS3-17813 (Ref, 45). Two-dimensional conduction e f fec ts and the spat i a1 va r ia t i on o f the coolant heat t rans fer coef f ic ient were simulated i n t h i s program. A program opt ion represents the not wal l , the land and t h a t p a r t o f the external wa l l adjacent t o the channel as f i n s , That p a r t o f the external wa l l adjacent t o the land i s assumed t o be isothermal. The land f i n can be s p l i t i n t o two segments w i th d i f f e r e n t coolant heat t rans fer coef f i c ien ts . One coef f i c ien t i s appl ied t o the hot wa l l and the f i r s t p a r t o f the land f i n s , The other c o e f f i c i e n t i s appl ied t o the res i o f the land f i n and the external wa l l f i n . The in te r face between segments o f the land f i n corresponds t o a spec i f ied coolant-side wa l l temperature.

A l i m i t e d nu&er o f two-dimensional mode network analyses were performed a t the maximum heat f l ux l oca t ion near the th roat . These studies accomplished the fol lowing:

r Provided deta i led temperature d i s t r i b u t i o n s f o r the cycle l i f e analysis.

a Provided the basis f o r determining the computer model simul a t i on parameters f o r oxygen methane and hydrogen cool i n g .

r Established the optimum channel geometry f o r a f i x e d coolant f 1 ow area w i th oxygen cool i ng .

r Defined loca l coolant ve loc i t y requirements f o r RJ-5 and RP-1 cooling.

r iktermined the accuracy o f the computer model f o r MMH and hydrazine cooling,

Gas side heat t rans fer was handled i n the fo l lowing manner. A two-dimensional nozzle expansion analysis and a TRAN 72 computer program (Ref. 42) ca lcu la t ion were used t o determine per t inent parameters a t the ed e o f the wal l boundary layer for LOXIRJ-5. Parameters establ ished were; (13 the r a t i o of two-dimensional t o one-dimensional m s s ve loc i t ies , (2 ) the r a t i o o f s t a t i c t o stagnation tewperatures and, (3) the r a t i o of adiabat ic wal l t o stagnation temperatures. A recovery fac to r e.qual t o the 113 power o f the Prandtl number was used i n t h i s analysis. A l l r a t i o s determined were assumed t o apply t o the other propel 1 ant combinations under invest iga- t ion . A l l combustion product propert ies were evaluated using the TRAN 72 computer program,

Faximum heat f l u x occurs s l i g h t l y upstream o f the throat . Heat transfer from the combust!on products t o the chamber wal l was ca l - culated as:

-0.2 -0.6 0 = 0.026 C P u Ref

g f e Prf C (Taw-Twg) p f

(9)

where: T

Tf = 0.5 (Taw-Twg) (12)

The coe f f f c ien t C accounts f o r f low accelerat ion ef fects. Nomenclature i s as fol lows: 9

0 English Let te rs

gas-side heat t rans fer co r re la t i on c o e f f i c i e n t

c spec i f i c heat; rp i s an integrated average between the coolant bu lk temperature and the wal l temperature

D Local chamber diameter

Pr Prandtl number

Re Reynolds number

T temperature

u Axia l ve loc i t y

0 Greek Let te rs

1.1

P

0

a Subscr

aw

e

f

w g

v i scos i t y

density

gas-side heat f l u x

i p t s

adiabat ic wa l l

f rees t ream

f i l m temperature

gas-side wa l l surface

3. Oxygen Cool i ng

A recent ly developed ALRC co r re la t i on (Ref. 46) f o r supercr i t i c a l oxygen was used. This co r re la t i on y i e l d s heat t rans fe r coe f f i c i en ts which are sens i t i ve t o bul k tsmperature, wal l temperature and pressure.

Oxygen proper t ies f o r temperatures below 333OK (600°R) were taken from NBS data (Refs. 1 and 3), and extrapolat ions o f t h i s data. Russian data, (Ref. 47), f o r density, spec i f i c heat and enthalpy were used f o r temperatures above 333OK (600°R). Transport proper t ies f o r t h i s temperature range were taken from ALRC predict ions f o r pressures up t o 340 atm (5000 psia). A tab le o f these values i s included i n Ref. 46. Extrapolat ion t o higher pressures was based on the dense gas cor re la t ions o f Ref. 1 w i t h the d i l u t e gas cont r ibu t ion i n fe r red from the proper t ies a t 340 atm (5000 psia).

A channel opt imizat ion study was conducted ea r l y i n the Task I 1 e f f o r t t o def ine the channel geometry which minimizes the l oca l gas-side wa l l temperature f o r a f i x e d coolant v e l o c i t . This study assumed a loca l t h roa t s t a t i c pressure o f 204 atm (3000 s i a , a bul k temperature o f 153OK n (27E0R) and a t o t a l f low area of 35.9 cm (5.56 in.2). These assumptions g ive an oxygen v e l o c i t y o f 198 m/sec (650 f t l s e c ) and a heat t r a n f e r c o e f f i c i e n t which var ies w i t h wal l temperature as shown i n Figure 29. Because the heat t rans fe r c o e f f i c i e n t i s much higher a t low wal l temperatures, the land i s a very e f f e c t i v e f i n and maximum wal l temperatures occurred a t the channel center l ine. Two-dimensional network analyses w i t h a hot wal l thickness o f 0.762 mn (0.030 in . ) were used f o r t h i s study. The optimum conf igurat ion i s t h a t w i t h minimum channel width, which i s def ined by the land width and channel depthlwidth constraints. The r e s u l t s showed tha t the channel depth a f fec ts the maximum wal l temperature much less than channel width, so the advantage i n reducing the iand width comes p r i - mar i l y from the channel width reduct ion allowed. Use o f a 1.02 mn (0.040 in . ) land i n the analysis, instead o f the 7.62 mn (0.030 i n . ) minimum guide1 ine, r e s u l t s i n a 278OK (40°F) higher optimum wal l temperature.

Determination o f the simulat ion parameters i n the s i m p l i f i e d Ref. 45 computer model was a lso based on the coolant condit ions assumed above, but w i t h a somewhat lower coolant v e l o c i t y and the hot wa l l thickness reduced t o 0.635 mn (0.025 i n . ). Due t o wa l l temperature dependence shown on Figure 29, the coolant heat t rans fe r c o e f f i c i e n t var ies by more than fac to r o r two. It was found t h a t the max.imuin temperature o f the two dimensional analyses i s matched by the s i m p l i f i e d analysis by: (1) using a coef f i c ien t on the hot wal l which i s 11 percent greater than t h a t ca l - culated from the center1 i n e wal l temperature and (2) evaluat ing the c o e f f i - c i e n t f o r the external wal l and the e n t i r e land w i t h a wal l temperature equal t o the bulk temperature.

Channel depths from the i n j e c t o r end through the th roa t were determined by the cyc le l i f e c r i t e r i o n and var ied from about 6.35 mn (0-25

i n . ) a t the i n j e c t o r Face t o 7.62 mn (0.30 i n . ) a t the th roa t t o 25.4 mn (1.0 i n . ) a t the e x i t . Note t h a t w i t h the s t radd le-mi l l design selected

Pre

ssur

e =

204

atm

(30

00 p

sia)

B

ulk

Tem

pera

ture

= 153OK (

27S

0R)

Vel

oci

ty =

198

Wse

c (6

50 f

tls

ec

)

I

I . 1

I

1

8

I I

I

I

150

200

250

300

350

400

450

Wall

Tem

pera

ture

, OK

4 I I

I

I I

8

a I

I

I i

200

300

400

500

600

700

800

900

Wall

Tem

pera

ture

, O

R

Fi g

um 2

9.

Oxygen H

eat

Tra

nsf

er C

oef

fici

ent

Var

iati

on

wi t

h M

a1 1

Tem

pera

ture

herein, the channel depth does not have t o change s ign i f icant ly from the in jec to r through the throat.

Figure 30 shows the required oxygen pressure drop as a functior, o f chamber pressure.

4. Hydrogen Ccol ing

Two hydrogen cool i ng concepts were investigated : the para1 1 e l burn, staged-combustion c c l e and the gas generator cycle. The l a t t e r requires a 22.9 cm (9 in. f longer combustion chamber t o accomnodate high density l i q u i d propel lant in ject ion. Coolant heat t ransfer coef f ic ients were based on the Hess and Kunz cor re la t ion (Ref. 48). and the f low path was the same as i n the oxygen-cooled design. The simp1 i f i e d computer model provided excel 1 ent simul at ion o f two-dimensional conduction analyses using a coolant coef f ic ient var ia t ion s imi lar t o that f o r oxygen, but wi th the hot wall coef f ic ient equal t o (rather than 11 percent greater than) that f o r the center1 ine wall temperature.

Local wall temperature d i f f e ren t i a l s were again 1 imited t o 867°K (1100°F) t o sa t i s fy the cycle l i f e c r i te r ion . I n addition, loca l maximum wall temperatures were l im i ted t o 867OK (llOO°F) i n view o f the s ign i f icant st ructural property degradation above t h i s value. This consideration res t r i c t s the coolant bulk temperature r i s e i n order t o cool the nozzle without excessive pressure drop, and thus sets the minimum coolant f low rate.

Detai 1 ed design studies were conducted wi th hydrogen f low rates o f 6.35 and 9.07 kglsec (14 and 20 lb/sec) f o r the para l le l burn, staged combustion and gas generator cycles, respectively. These f low rates were chosen from prel iminary heat t ransfer and cycle balance analyses. Coolant flow rates d id not vary wi th chamber pressure, since the t o ta l heat load var ia t ion i s small. Resultant required pressure drops are p lo t ted i n F i ure 31. A staged combustion cycle f low ra te i s 7.26 kglsec (1 6 1 blsec ! would give bul k temperature r ises comparable t o the gas generator cycle values and s ign i f i can t l y reduce nozzl'e pressure losses a t low chamber pressures. However, since pressure drops are so low i n general wi th hydrogen cooling, studies o f pressure drop vs. f low ra te f o r t h i s cycle were not undertaken since they would not mater ia l ly a f f ec t the study resul t s .

5, RP-1 and RJ-5 Cooling

Cooling w i th RJ-5 or RP-1 was found t o be impractical based on 1 i m i t i n ~ coolant-side wall temperatures t o 589°K (600°F) i n order t o prevent s ign i f icant coking. Two-dimensional conduction network analyses i n the throat region determined the maximum cool ant ve loc i t ies required f o r a chamber pressure of 136 atm (2000 p?fa) based on the Hines corre la t ion (Ref. 49), RP-1 i s the bet ter coc*.ant, but s t i l l required a

LOX/RJ-5 PROPELLANTS O/F = 2.7

100 200 300 400 500 Chamber Pressure, atm

I 1

1 1 a I a 1 1 I I

4000 6000 .-I

2000 10,000 Chanrber Pressure, psf a

Figure 30. Coolant Pressure Drop w l th Oxygen Cool lng

31. Coolant

LOXIRJ-5 Propellants

Gas Genera t o r /

Staged Combus t i o n = 6.35 kglsec

"2 (14 lb lsec)

100 200 300 400 Chamber Pressure, Atm.

1 I 2000 4000 600C

Chamber Pressure, psia

Pressure Drop with Hydrogen Cooling

velocity o f 175 m/sec (575 ft/sec). The resul tant dynamic head and f r i c t i o n losses preclude maintaining pos i t ive s ta t i c pressures.

Design studies f o r the hydrazine and H4H systems were based on 1 imi t ing the coolant-side w a l l temperature t o 6 4 4 O K (700°F). Super- c r i t i c a i M tests reported i n Ref. 51) indicate s ign i f i can t mluct ions i n heat transfer coefficient f o r war1 temperatures cbove 6M°K (700°F). These and other tests wi th storable supercr i t ical f l u i d s indicate tha t standard bulk temperature correlations, e.g., the Hines correlat ion used herein, are applicable u n t i l the wall temperature reaches the c r i t i c a l tenperature. Since no data are avai lable f o r supercri t i c a l ilydrazine, the l i m i t i n g wall temperature was set j u s t below the c r i t i c a l temperature, 653OK (716OF). Resulting ou t le t temperatures f o r MH are well below the decmposi t i o n temperatures observed wi th most wall material s. However, the hydrazine ou t le t temperatures are a t the threshold sens i t i v i t y temperatures observed wi th l iquid/vapor samples subjected t o adiabatic compression. It should be noted that adding a small amount of WI t o hydrazine increases t h i s thresh01 d temperature s ign i f icant ly . Figure 32 shows the pressure drops required f o r chamber cool ing w i th H I and hydrazine. Hydrazine pressure drops are reasonable, but H4H pressure drops are high. ZrCu was assuned as the chamber material throughout these studies i n accordance wi th the guidelines. However, materials investigations show that uncoated ZrCu i s incompatible wi th both MMH and N2H4 a t temperatures above 330°K (140°F) f o r long term use and may not be feasible f o r even short term use.

7. Met haw Cool ing

The supercr i t ical oxygen correlat ion (Ref. 46) developed a t ALRC i n 1975 was used t o predict the heat t ransfer coeff ic ients f o r the super- c r i t i c a l methane coolant. I n addition, a CLF-5 supercr i t ical correlat ion which i s also described i n Ref. 46 was used as a cross reference. The CLF-5 correlat ion does not contain a pressure factor which reduces the heat transfer coeff ic ient a t pressures above 212 atm (3120 psia) as the oxygen correlat ion does. Thus, i t resulted i n higher coeff icients and lower cooling requirements. The oxygen correlat ion was chosen because no supercri t i c a l methane correlat ion could be located, and the oxygen data approached the same pressure range as the methane coolant pressures studied.

The cool ant f low path and channel construc~ion assumptions were ident ical t o those previously described, and resulted i n the selection o f 270 channels f o r the combustion chamber.

The required coolant jacket pressure drop as a function of chamber pressure using methane as a coolant i s shown on Figure 33.

Figure 32.

100 200 300 40\. Chamber Pressure, A t m

I I

1 I

1 I I I

2000 3000 4000 5000 6000 Chamber Pressure, psi a

Cool ant Pressure Drop with Hydrazi ne Compounds

LOXICH4 Propel 1 ants

O/F = 3.5

1 1

1 I

I v

I I

I ' 200 300 400 500 Chamber Pressure, Atmospheres

. I

1 I I I I 1 I ( I I

2000 4000 6000 10,000 Chamber Pressure, psia

Figure 33. Coolant Pressure Drop with Methane Coollng

8. Sumnary o f Results and Conclusions

Coolant jacket pressure drops, coolant bul k temperature r ises and ou t le t temperatures resu l t ing from t h i s study are sumnarized on Table X X X V I I.

Based upon t9e resu l ts of t h i s task, the fo l lowing conclusions were reached :

O Hydrocarbon fue ls such as RJ-5 and RP-1 are not pract ica l coolants f o r high pressure appl icat ion due t o the low wall temperature which must be maintained t o prevent s ign i f i can t coking.

Oxygen cooling i s feasib le f o r chamber pressures up t o a t least 272 atm (4000 psia).

O Hydrogen cool ing requires the smallest pressure drop and i s pract ical f o r chamber pressures beyond 340 atm (5000 psia). Flow rates as low as 7.26 kgjsec (16 1 bjsec) f o r the staged combustion cycle and 9.07 kgjsec (20 lb jsec) f o r the gas generator cycle are su f f i c ien t .

O Cooling wi th MMH requires the highest pressure drop, but i s feasible providing tha t the copper can be coated wi th a compatible material which does not s ign i f i can t l y a1 t e r the conductivi ty o f the chamber wall s.

" Cooling wi th hydrazine i s feasible from a pressure drop standpoint f o r chamber pressure beyond 340 atm (5000 psia) but coolant bulk temperatures are close t o the adiabatic compression sensi t i v i ty thresh01 d. Addition o f a small amount o f MMH would provide a much more stab1 e mixture and should re ta in much of the heat t ransfer advantage o f hydrazine.

O Use o f methane as a regenerative coolant i s feasib le f o r thrust chamber pressures beyond 340 atm (5000 psia).

SECTION V

TASK I11 CYCLE EVALUATIOPI


The objectives of t h i s task were t o determine engine cycle pressures, temperatures and del ivered performance f o r the candidate Mode 1 engines. Each candidate was evaluated over a chamber pressure range o f 136 t o 340 atm (2000 t o 5000 psia) o r a t the maximum value corresponding t o a cycle power balance, th rus t chamber cool ing o r propel l a n t property 1 im i t .

Candidate engine concepts considered were :

Propellant Combination

RJ-5/0wgen RJ -5IO~ygen R J-5/0wgen (Para1 1 e i Burn) RJ-5/0yygen (Gas Gen Cycle) RP-1 /Oxygen Hydrazi ne/OXygen M W h ~ g e n CH4/Oxygen RJ-5 & LH2/0wgen

Cool ant

Hydrogen Hydrogen RP-1 Hydrazi ne MMH w 4 Oxygen

Reference Schematic

Figure 34 F i gure 35 Figure 36 Figure 37 Figure 35 Figure 35 Figure 35 Figure 35 Figure 38

The engine cycle power balances w e y perfwm,d a t a sea-level. Tus o f 2.70 MN (607,000 lb), w i th the engine mixture r a t ~ o s and coolant ~ a c e pressure drops as determined i n Task 11. Engine performance data were evaluated for a combus ti on ef f ic iency of 98%. Addi t iona l study guide1 ines are presented on Tables X X X V I I I , X X X I X and XL and below.

e System Pressure Losses @PIP upstream)

In jectors: - L iqu id - 15% min. - Gas - 8 % m i n .

Val ves : - Shutoff - 1% - L iqu id Control - 5% min. - Gas Control 10% min.

a Boost Pump Drive Requirements no t Considered

@ Main Pump Suction Specif ic Speed = 20,000

FUEL

Fna

L- O

XYG

EN

- CO!!

BUST

I OH

PR

ODUC

T5

Fig

ure

35.

Ti

. ;k

I1

1 M

ode

1 Fu

el C

oole

d E

ngin

e C

ycle

Sch

emat

ic

TABLE XL, - TASK I1 I TURBINE INLET TEMPERATURE SELECTIONS

Turbi ne Propel 1 ant Turbine !;let mp. Combi nation Drive Gas Cr i te r ia

O2/RJ-5 Ox-Rich 922 (1660) L i f e i he l -Rich 867 (1560) Co k i ng

02/RP-1 OX-Rich

Fuel -Rich L i f e

W i n g

0 2 / w Ox-Ri ch 922 (1660) L i f e Fuel -Rich 922 (1660)(') L i f e

O2lN2"4 OX-Ri ch Fuel -Rich

L i f e L i f e

021H2 Fuel -Rich & Ox-Rich 922 (1660) L i fe

02/CH4 Ox-Rich 922 (1663) L i f e

Fuel -Rich 867 (1560) Coking

l~ reburner Temp. = 1 383°K (2490°R) (2)~reburner Tmp. = 1 O28OK (1 850°R)

The turbine i n l e t temperature c r i t e r i a l i s t e d on Table XL were selected t o meet the l i f e requirements o f the engfne and t o p roh ib i t turbine coking problems i n the case o f the fue l - r ich hydrocarbon preburners. Experimental data a t lower pressures has shown tha t coking i s ins ign i f icant a t a temperature o f 867°K (llOO°F). For the fuel- r ich LOX/M and LOX/NzH4 combinations, the mnopropel 1 ant tenperature exceeds pract ical turbine i n l e t tenperature l i m i t s f o r long l i f e . Therefore, f o r these cycles, i t was assumed that a heat exchange device can be used t o reduce the temperature o f the fue l -ri ch prebumer conbus ti on gases.

B. TASK I I I DELIVERED PERFORMANCE

Delivered perfcrmance was calculated f o r a1 1 the engines over a &amber pressure range o f 136 t o 340 atm (2000 t o 5000 psia) . Delivered vacuum and sea-1 evel performance was calculated as fo1 1 ows :

ISD vac - "sp ~osses (13) Del

vacuum delivered impulse

where :

Isp vac = Del

- I:,vac - ODE

ODE vacuum impulse

Real engine performance losses including chemical kinetics, boundary 1 ayer, nozzle divergence, and in jec to r energy release

= Sea-level del ivered impulse

= Sea-1 evel nozzle ex1 t force performance decrement

The real engine ,\erformance losses considered i n the analyses are sunnarized below:

A I = Kinetics loss accounting f o r incomplete chemical re- S P ~ ~ combination i n the chamber convergent and expansion

sections A 1 = Boundary layer loss accounting f o r nozzle f r i c t i o n

s P ~ ~ A I = LOSS accounting f o r nozzle heat transfer. This loss i s

S P ~ ~ zero f o r a1 1 the TCA fuel o r ozid izer cooled engines since heat i s recycled t o the preburners. The loss i s pos i t ive for the hydrogen cooled Mode 1 concepts

A I = Divergence loss accounting for non-axi a1 nozzle exi t S P ~ ~ moment urn

AI = Injector energy release loss assumed to be 2% of ODE S P ~ ~ ~ performance which i s consistent w i t h the Task I1

geometry selection to achieve this value. A I = Gas generator cycle loss. GG flow assumed to be dumped

s P ~ ~ i n TCA nozzle cycle a t area rat io consistent w i t h turbine s t a t i c ex i t pressure

ODE performance i s documented in Section I11 for a l l the Mode 1 engine propellant combinations . The nozzle chemical kinetics loss was calculated with the One-Dimensional Kinetic (ODK) option of the JANNAF TDK reference program (Ref. 51 ) . The chamber boundary 1 ayer friction loss was calculated with an ALRC computerized formulation of the JANNAF boundary layer chart technique described i n Ref. 52. The nozzle divergence loss was calculated from a Rao nozzle design program. The program predicted a divergence efficiency of 99.8% for a contoured 90% Bell length, area rat io equal to 40:l nozzle.

The resul ting del i vered performance, for the staged conbustion cycle engine concepts, i s presented on Figures 39 through 44. The data are shown for the mixture ratios selected i n Task 11.

The para1 1 el burn, hydrogen cooled staged combus ti on engine concept has a 1.2 sec lower delivered perfomance than the baseline LOX/RJ-5 engine. T h i s occurs because the heat i n p u t into the hydrogen coolant i s transferred to the Mode 2 engine preburners. The 1.2 secs performance loss can be subtracted as a constant value from the performance of Figure 39.

The hydrogen cooled,gas generator cycle Mode 1 engine concept has one additional performance loss cowpawd to the staged combustion concepts. The gas generator flowrate i s dumped in the nozzle downstream of the throat and thus , delivers something less than 40:l expansion perfomance. The specific imoulse of tho 22s generator flow was calculated assuming the gases wil l 'be expanded i n the nozzle from the turbine exit s t a t i c pressure to the nozzle 40:l wall ex i t pressure. This pressure was obtained from a perfect gas method of characteristics solution for a 40:1, 90 percent Bell contoured nozzle. The gas generator cycle perfomance loss was calculated by subtracting the sumned TCA and gas generator flow mass weighted per- formances from ODE performance for the fuel and oxidizer a t nominal mixture rat io (O/F = 2.7 fnr LO /RJ-5). The calculated gas generator performance losses are 0.5, 1 .O, 1 .g and 2.4 seconds for TCA chamber pressures of 136, 204, 272, and 340 atm (2000, 3000, 4000, and 5000 psia) , respectively.

The gas generator performance loss plus the 1.2 sec heat loss must be subtracted from Figure 39 to obtain the delivered perfomance for the hydrogen cooled, gas generator cycle engi ne concept.

O/F = 2.7:l 90% Bel l Nozzle

I,, Vacuun

260 I I I

I I

I I

I I I

10 20 30 40 50 60 Expansion Area Ratio, Pi/AT

F i gure 39. LOX/RJ-5 Staged Combustion Del i vered Performance vs Area Ratio

O/F = 2.9 90% Bel l Nozzle

Chamber Pressure atm (psia) -

Area Ratio

Figure 40. LOXIRP-1 Staged Combustion D e l i ~ e r e d Performance vs Area Ratio

11 4

O/F = 1.6 90% Bell Nozzle


U QJ V) .. A

e ,w 340 -- Y

QJ V) F 3

B u F 320 3- L a > .r P QJ 0

300 --

280 --

1 I

. I I

I I I

20 30 40 50 60 Area Ratio

Figure 41 . LOXIMMH Staged Combustion Del i vered Performance vs Area Ratio

115

.O 1 Nozzle


Vacuum

Arza Ratio

Figure 42. LOX/N2H4 Staged Combustion Del i vered Performance vs Area Ratio

116

O/F = 7.0 90% Bell Nozzle

Chamber Pressure

Sea-Level

Chamber Pressure atm (psia) - 470 -

460-9

450.-

440 - - 430--

420 --

4101- $ 2 400,. n V)

n Y

390-- VI

r- 3 P E 380-- n

0 a L 0, > 370-- .c

E 360 - - 350,-

340--

330-p

320,.

310.-

1 I 1

1 I

I I

I I

20 40 60 1 00 2b0 300 Area Ratio

Figure 43. 02/H2 Stdged Combustion Del ivered Performance vs Area Ratio 117

O/F = 3.5 90% Bel l Nozzle PC

atm (psia) 340 (5000) 272 (4000)

I Vacuum s P

204 (3000) 136 (2000)

0 -340 (5000)

1 2 7 2 (4000)

204 (3000)

I Sea-Level s P

136 (2000)

30 40 Expansion Area Ratio, AE/AT

Flgure 44. LOX/CH4 Staged Combustion Del ivered Performance vs Area Ratio

118

C. ENGINE CYCLE POWER BALANCES

Engine pump dischar e pressure versus chatrber pressure relat ionships were evaluated f o r the var 1 ous candidate coolants and engine cycles. Cycles using RJ-5 and RP-1 as coolants were not evaluated because the coolant evaluation study resul ts showed tha t operation i n a reusable engine would be l im i ted t o pressures belgw 136 atm (2000 psia).

The resul t s of the cool ant eval ua t i on studies , cool ant jacket AP and flow rate, were used t o conduct these analyses i n conjunction wi th the sys tem pressure drop c r i t e r i a .

The cycle evaluation assumes tha t the boost pumps w i 11 produce su f f i c i en t discharge ressure t o i n t a i n a main pump suct i pec i f i c speed o f 3000 rpm x 9 1 4 x min -178 (20,000 rpm x pm x ft-3145. The 9 boost pump dr ive requirements were not considered n the power balances,

Parametric pump performance curves f o r head cae f f i c ien t vs speci f i c speed and pump ef f ic iency vs impeller diameter were used t o calculate the pump eff iciencies. I n general, pump ef f ic ienc ies i n the range o f 75 t o 80% were obtained depending upon the size o f the impeller. The performance of the RJ-5 pumping system was lowered by 2 percentage points t o account for the extremely high v iscosi ty of the propellant.

Design point turbine e f f ic ienc ies used i n conducting the power balance calculat ions for the staged combustion cycle engines are as follows:

LOX Rich Turbines - 80% Candidate Fuel - R i ch Turbines - 74% Hydrogen-Ri ch Turbines - 81%

For the low flow, high pressure r a t i o turbines o f the gas generator cycle engine, turbine e f f ic ienc ies o f 603 were assumed. This e f f ic iency corresponds t o tha t obtainable wi th a two-stage turbine a t a veloci ty r a t i o o f 0.2. Eff iciency, i n t h i s case, i s not c r i t i c a l i n establ ishing the main propel lant pump discharge pressure and does not mater ia l ly a f fec t the power balance.

Maximum pump discharge pressure requirements f o r the candidates are presented on Figure 45. The maximum feasible operating pressure f o r the staged combustion cycle engines i s considered t o be 80% o f the chanber pressure a t which the curve i s asymptotic. For example, the baseline LOX cooled engine asymptote i s approximately 354 atm (5200 s ia) which resu l ts i n a maximum recommended pressure o f 283 atm (4160 psia ! with the 20% margin.

Pert inent conclusions derived from the cycle evaluations along wi th operational considerations a!*e sunmarized or. Table XLI. It should be noted that the major i ty of the propel lant combinations (par t i cu la r l y RJ-5 and RP-1) are viable candidates if oxygen, rather than the fuel, i s used t o cool the combus ti on chanber .

A summary o f engine data a t a thrust chamber pressure o f 272 atm (4000 psia) i s shown on Table XLII.

10 I I

I I I

1 I

100 200 300 400 Thrust Chamber Pressure, Atm.

Fuel & Ox. Pumps,

Stg. Cycle

LOX Cooled Baseline, Stg. Cycle

Stg. Cycle

Cool an t

Fuel & Ox. Pumps, Hz Cooled Gas Gen. Cycle

1 I I

I 1

o 1000 2000 3000 4000 5000 6obo Thrust Chamber Pressure, ps ia

Figure 45. Maximum Pump Discharge Pressure Requirements f o r Candidate Mode 1 Engines

TABL

E X

LI-.

- TA

SK

111

CYCL

E EV

ALU

ATIO

N S

UW

RY

Coo

l an

t

02

"2

H2 w

N2H

4

RJ-

5

RP-

1

CH

4

Max

P,, L

-

Cyc

l e -

Atm

(p

sia

) R

emar

ks

S .C

. 28

5 (4

200)

' V

iab

le C

andi

date

Par

a1 1 e

l B

urn

G.G.

S.C.

S.C.

S .C

.

S .C

.

S.C.

340

(50

00

)~

Via

ble

Can

dida

te

>340

(>

5000

) V

iab

le C

andi

date

306

(450

0)'

Ma

teri

al

co

mp

ati

bil

ity

qu

est

ion

ab

le

Fue

l -R

ich

gas

com

bust

ion

tem

p.

prob

lem

>340

(>

5000

) M

ate

ria

l c

mp

ati

bi l i ty

qu

est

ion

ab

le,

Exp

losi

ve d

ecom

posi

ti o

n o

f N

2H4

vapo

rs,

Fu

el-

rich

gas

com

bust

ion

tem

p.

prob

lem

.

<I3

6

(<20

00) b

Cok

ing

in c

hann

els

<I3

6

(<2

00

0)~

C

okin

g in

cha

nnel

s

272

(40

00

)~

Via

ble

Can

dida

te

a~

as

ed

upon

20%

pow

er m

argi

n

b~

rop

el la

nt

pro

pe

rty

1 id

ted

TABL

E X

LII

. - T

ASK

I11

CAN

DID

ATE

MOD

E I

ENG

INE

COM

PARI

SONS

Th

rust

. m

(lb

) S.

L.

Vac .

Is.

Sec

S.L.

Vac .

Iota

1 F

low

Rat

e.

kg/s

ec

(Ws

w)

Mix

ture

Lt

io

h.

Flo

w R

ate.

kg

/sff

(l

bls

ec

)

Fue

l F

low

Rat

e.

kgls

ec

(lb

lse

c)

LH

Flo

w R

ate.

kg

/sec

(1

2/se

c)

Coolant

Co

ola

nt

Jx

ke

t AP

. at

m (

ps

i)

too

lan

t E

xit

Tem

p.

OK

(^

R)

Ox.

Pm

y D

isch

arge

Pre

ssur

e.

atm

(p

sia

)

Fue

l P

mp

Ois

char

ge P

ress

ure.

at

m (

psi

a)

LH2

Pu

p O

fsch

arge

Pre

ssur

e.

a-

psi^

)

LOX

Tur

bopm

p S

peed

. RP

n

Fu

el

Tur

bopc

np S

peed

. RP

M

LH2

Turb

opun

p S

peed

. RP

M

LOX

Pln

p H

orse

pow

er

Fue

l P

mp

Ho

rse

po

we

r

LH2

Pm

p t

brse

pow

er

Cha

mbe

r P

ress

ure

= 2

72 a

tm (

4000

psi

a)

No

zzle

Are

a R

ati

o =

40:

l

Sta

ged

Coc

lbus

tion

LOV

RJ-

5 B

ase

line

,

2.70

(6

07.0

00)

2.92

(6

56.4

00)

321.

3

347.

4

(188

9.2)

2.7

(137

8.6)

(510

.6)

-- LOX

(1 56

0)

(332

)

(920

0)

(760

0)

- -

15.1

00

21.6

00

- - 57

.400

19.6

00

- -

Sta

ged

Con

bu,ti

on

LOX

/mI

2.70

(6

07.0

00)

2.92

(6

56.9

00)

Sta

ged

Co

llbu

stio

n

LOxJ

N*

- 2.

70

(607

.000

)

2.93

(6

57.6

00)

LOX

/RJ-

5 P

ara

lle

l B

urn

Sta

ged

Co

du

stio

n

LO

x/W

2.70

(6

07.0

00)

A. OBJE

SECTION V I

TASK I V - ENGINE YEIGHT AND ENVELOPE

CTIVE AND GUIDELINES

The object ive o f t h i s task was t o provide parametric engine weight and envelope data f o r the candidate Mode 1 engines, the dual-fuel engine and the base1 inc Mode 2 engine.

The parametric ranges considered were:

Parameter Mode 1 Enq i nes

Mode 2 Engines

Thrust, MN (Lb) 1 .78 t o 4.0 Sea-Level 2.2"o 4.45 Vac. (400K t o 900K Sea Level ) (500K t o 1 OOOK Vac. )

Chamber Pressure, Atm. 136 t o 340 136 t o 340 (psia) (2000) t o (5000) (2000) t o (5000)

Nozzle Area Ratio 10:l t o 60:l 100:l t o 300:l

Parameter Dual -Fuel

).lode 1 - Mode f

Thrust, MN (Lb) 2.7 t o 4.0 Sea-Level 2.16 t o 3.44 Vac. (607K t o 900K Sea-Level) (486K t o 774K Vac.)

Chamber Pressure, Atm. 182 t o 363 136 t o 272 (psfa) (2670 t o 5330) (2000 t o 4000)

Nozzle Area Ratio 40:l 100:l t o 300:l

B. GASEL1 NE ENGINE WEIGHT STATEMENTS

For purpose of the parametric study, i t was necessary t o establ ish the elements o f engine weight t o be included i n the scaling study and t o establ ish baseline engine weight statements. Table XLI I I l i s t s the engine components included i n the parametric analyses. Those items not included are also l is ted.

Engine weight statements are shown on Table XLIV. Included are the three engines on which preliminary designs were completed i n Task V I , the Mode 2 engine, and the LOXlCH4 methane cooled candidate Mode 1 engine.

I n the Task I V effort, i t was necessary t o i n i t i a l l y generate weights and parametric data for use i n the evaluation and select ion o f candidate Mode 1 engines f o r prel iminary design. These data were based on

TABLE XLIII , - TASK I V - ENGINE WEIGHT DEFINITIONS

I nc 1 uded Not Included - Regenerati vely Cooled Combus t i o n Gimbal Actuators and Actuation Chamber Sys tem

Regenerati vely Cooled Thrust Chamber Engi ne Control 1 c r Fixed Nozzle

Thrust Chamber Norzl e Extension Pre-Val ves (Mode 2)

NozzleExtension Deployment System Tank Pressurant Heat Exchangers (Mode 2) and Associated Equipment

Main In jec tor

Main Turbopunps

Boost Pumps

Preburners (or Gas Generator)

Propellant Valves and Actuation

Gimbal

Hot Gas Manifold ( i f required)

Prnpell ant Lines

Ign i t i on System

Miscel laneous (E lect r ica l Harness, Instrumentation, Brackets, Auxi 1 i a r y Lines and Controls)

Contingency (a t o ta l contingency i s normally included i n the t6ehicle weight statement)

TABL

E X

LIV

. - T

ASK

IV -

BA

SE

LIN

E E

NG

INE

WEI

GH

T ST

ATEM

ENTS

Com

pone

nt

Gim

bal

Mai

n In

jec

tor

and

Ma

nif

old

Mai

n C

ham

ber

Fix

ed

No

zzle

Ext

en

dib

le N

ozzl

e

Ext

. N

ozz

le D

eplo

y.

Sys

tem

Pre

bu

rne

rs

Val

ves

and

Act

ua

tio

n

Boo

st P

unps

Mai

n Pu

mps

Lin

es

Mi s

cel ?

ane

ous

To

tal

Dry

Vac

T

hru

st/W

t

LOX/

RP-

1

Mod

e li

l)

96

(21

1)

385

(84

8)

124

(27

4)

1

(24

9)

- - - - 18

1 (3

98

)

91

(200

)

159

(351

)

528

(116

4)

209

(46

0)

227

(50

2)

--

2113

(4

657)

141

Mod

e 2

(2)

Wei

ght

Dua

l -F

uel

(1

96

(21

1)

385

(84

8)

124

(274

)

89

(19

7)

592

(130

6)

292

(64

4)

411

(90

5)

305

(67

2)

209

(46

1)

986

(217

4)

449

(98

9)

246

(54

2)

--

4184

(9

223)

71

(LO

X/R

P-1

)

56

(LO

XIL

H2)

(')B

ased

up

on c

alc

ula

ted

we

igh

ts fro

m

pre

lim

ina

ry d

esi

gn

la

yo

uts

("B

ased

up

on s

ca

l i ng

of

SSM

E

(3)B

ased

up

on s

ca

lin

g o

f in

-hou

se

IRID

co

nce

ptu

al

de

sig

n a

nd

his

tori

ca

l d

ata

kg

(lb

) LO

XlC

H4

Mode

95

(20

9)

348

(767

)

127

(28

1)

249

(550

) - - - - 12

1 (2

67

)

268

(59

0)

151

(33

3)

585

(128

9)

206

(45

4)

227

(50

2)

--

2377

(5

242:

125

LOX/

RP-1

H

z

Coo

led

G.G.

c

yc

le('

)

95

(211

)

349

(76

9)

157

(347

92

(20

3) - - - -

9

(20

) G.

G.

133

(29

3)

167

( 368

3 94

(8

69

160

(35

3)

227 -

( 502

1

1784

(3

935)

167

scal ing o f h i s t o r i c a l weights of s i m i l a r components and/or estimates obtained from conceptual designs. A f ixed 90% b e l l nozzle was assumed f o r the Mode 1 engines and a f i x e d 90% b e l l nozzle t o an area r a t i o o f 40:l and an extendible 90% be1 1 nozzle beyond was assumed f o r the Mode 2 and dual- f ue l engines. Upon completion o f the pre l iminary design e f f o r t i n Task V I , the weight statements and parametric data were updated f o r the three pre- l im ina ry design engines t o r e f l e c t weights and dimensions as calculated from the p re l iminary design 1 ayouts.

Mode 2 engine baseline weight was obtained by modifying the SSME engine ta rge t weight (Ref. 53) f o r the study guide1 ines and assumptions, and then scal ing t o the required t h r u s t leve l . The SSME weight was modif ied by: (1 ) e l iminat ing the contingency and miscel 1 aneous components excluded i n t h i s study, (2) ca l cu la t i ng the SSME nozzle weight per u n i t surface area and ad jus t ing the f i xed nozzle weight f o r a 40:1, 90% b e l l and, (3) adding a nozzle extension and deployment system weight which i s based upon the ALRC SSME design (Ref. 54).

I t should be noted t h a t adjustments i n ind iv idua l components were not made t o account f o r the mixture r a t i o change from 6 t o 7. Past studies of hydrogen cooled, LOXILH2 engines have shown t h a t var ia t ions i n component weights w i t h n i x t u r e r a t i o tend t o compensate so t h a t the resu l t i ng t o t a l engine weight i s essen t i a l l y constant.

The base1 i ne weight f o r the LOX/CH4 methane cooled engine i s based on the f ~ i t i a l Task I V weight statements, i .e. scaled h i s t o r i c a l weights and estimates from conceptual designs. Because of the low densi ty o f methane, the baseline weight includes a scale fac to r which accounts f o r the volumetric flow r a t e difference between CHq and RP-I engine components.

C. PARAMETRIC WEIGHT DATA

With the base1 i ne engine weight establ ished, engine component weight scal ing re la t ionsh ips were derived as funct ions o f thrust , t h r u s t chamber pressure and nozzle area r a t i o . These scal i n g re la t ionsh ips were used t o ca lcu la te the weights over the parametric ranges o f i n te res t . The scal i ng equations were establ ished througi geometry considerations and empirical data f i t s o f h i s t o r i c a l data. These techniques have proven t o be sa t is fac tory i n past parametric studies o f t h i s nature such as, the OOS Engine Study (Contract F04611-71 -C-0040), the Space Tug Storable Engine Study i -on t rac t NAS8-29806), and the parametric analyses conducted f o r the ear ly Phase B, Study (Contract NAS8-26188).

1 . Mode 1 Staged Combus ti on Cycle

Figure 46, shows the Mode 1 staged combustion cyc le engine weight as a funct ion o f t h r u s t chanber pressure f o r a nozzle area r a t i o of 40 : 1 . Weight va r i a t i ons w i t h t h rus t are a1 so shown on the f i g u r e .

The data on Figure 46 were calculated f o r the Mode 1 LOX cooled basel ine engine b u t are also w i t h i n the ca lcu la t ion accuracy f o r the p rac t i ca l candidate Mode 1 fue l cooled staged conbustion cycles except f o r LOX/CH4. Prac t ica l chanber pressure l i m i t s f o r each o f the cycles were shown on Table XLI o f the previous section,

Minimum engine weight occurs between 190 and 231 atm (2800 and 3400 ys ia) depending upon the t h r u s t l eve l . The lower th rus t l eve l has a corresponding higher chamber pressure. However, a chamber pressure increase t o 272 atm (4000 ps ia) resu l t s i n on ly a modest engine weight increase.

It can a lso be establ ished from the data shown on Figure 46 t h a t the engine t h r u s t t o weight r a t i o decreases as the engine t h r u s t l eve l increases. This would i nd i ca te t h a t i t may be desirable t o c l u s t e r more small engines ( i .e., f ou r 2.70 MN (607,000 l b ) t h r u s t engines ra the r than three 3.6 MN (809,000 l b th rus t ) engines.

The e f f e c t of area r a t i o upon the Mode 1 staged combustion cyc le engine weights i s shown on Figure 47. The data are p l o t t e d f o r the baseline t h r u s t chamber pressure o f 272 atm (4300 psia) . The engines are heavier a t an area r a t i o of 10:l than a t 20:l because a 10:l nozzle a t 272 atm (4000 ps ia) i s we l l underexpanded and low performing a t sea-level. This resu l t s i n a large th roa t s ize and correspondingly la rge surface areas f o r the combustion chamber and nozzle.

The trends shown on Figure 47 are s i m i l a r f o r the other engines considered i n t h i s study.

2. Hydroge~ Cooled, Gas Generator Cycle

The hydrogen cooled, gas generator cycle engine weight as a function o f chamber pressure and t h r u s t l eve l i s shown on Figure 48. This engine i s schematically depicted on F i g w e 37 o f the previous sect ion. The trends w i th chamber pressure and t h r u s t are s i m i l a r t o those discussed f o r the Mode 1 s ta l2d corhustion cycles. However, the chamber pressure r e s u l t i n g i n minimum engine weight i s h igher f o r t h i s cycle.

This engine weighs approximately 318 kg (700 l b ) less than the basel ine LOX cooled, LOXIRP-1 engine a t a t h r u s t o f 2.7 MN (607K l b ) and chamber pressure o f 272 atm (4000 ps ia) .

.P El, E W

El, Y

.I - C, .c El, .C

. aJ C .r El, E W

Sea-Level Thrust, MN (Ib)

2,000 1 I I I I I I I I I 1 I 4 5 0 1 00 150 200 250 300 350

Chamber Pressure, A t m .

1 on'o I I I

2000 3000 4000 5060 Thrust Chamber Pressure, ps ia

Figure 46. Mode 1 Staged Combustion Cycle Engine Weight Parametrics

Chamber Pressure = 272 Atm (4000 psia)

Sea-Level Thrust , MN (k lb )

2.70 (607 ) - eel i n e

2b 4 6 6b Nozzle Area Ratio

Figure 47. E f f e c t of Area Ratio bpon Mode 1 Staged Combustion Cycle Engine Weight

129

Sea-Level Thrust , MN ( I b )

L. BASELINE

I I I I I t I

l o o 1 50 2d0 250 300 350 Chamber Pressure, Atm .

1

I I I

1 1

I I t

1000 2000 3000 4000 5000 Thrust Chamber Pressure, psi a

Figure 48. Mode 1 Hydrogen Cooled Gas Generator Cycle Engine Weight Parametrics

The e f f e c t o f area r a t i o upon the hydrogen cooled gas generator cyc le engine weight i s shown on Figure 49. The data are p l o t t e d f o r a th rus t chamber pressure o f 272 atm (4000 ps ia) . As w i t h the Mode 1 staged combus- t i o n cyc le engines, the GG cyc le engines are heavier a t an area o f 10:l than a t 20:l because o f the nozzle underexpansion and low performance a t sea l eve l which r e s u l t s i n a la rge th roa t s ize and correspondingly la rge surface areas f o r the combustion chamber and nozzle.

3. Dual-Fuel Engine

The dual -fs.el engine concept i s shown schemt ica l l y on Figure 38. Oxygen i s used t o cool the engine i n both modes o f operation. Be- cause the chamber cool ing problem i s more severe f o r the Mode 2 LOX/LH2 operation, the Mode 2 chamber pressure and t h r u s t i s lower than Mode 1. The parametric weight data f o r t h i s engine i s shown on Figure 50 as a funct ion o f the extendible nozzle area r a t i o and Mode 1 t h r ~ s t leve l . The data are shown f o r the recommended operat ing chamber pressures o f 272 atm (4000 ps ia) f o r Mode 1 and 204 atm (3000 ps ia) f o r Mode 2.

For a given Mode 1 sea-level th rus t , the Mode 2 vacuum th rus t w i 11 vary w i th the nozzle area r a t i o because the th roa t area of the dual - f u e l combustion chamber i s f ixed. The t h r u s t var ies as fol lows:

Mode 1 Nozzle Mode 2 Sea-Level Thrust , Are a Vacuum Thrust, M (Klb) Rat io MN (Klb)

4. Mode 1 LOXICH4 Engine

The Mode 1 LOXICH4, methane cooled engine weight parametrics are shown on Figure 51. The data are presented as a funct ion o f t h r u s t chanber pressure and t h r u s t l eve l for the basel ine nozzle area r a t i o o f 40:l.

Because of the lower densi ty o f methane, the base po in t f ue l components are heavier. This resu l t s i n lower pressures f o r minimum engine weight. However, the weight penalty f o r operat ing a t a th rus t chamber pressure of 272 atm (4000 ps ia) remains modest.

Chamber pressure = 272 A t m (4000 psia)

SEA-LEVEL THRUST, MN (KLB)

BASELINE

0 2 0 40 60 NOZZLE AREA RATIO

Figure 49. E f fec t o f Area Ratio Upon Mode 1 Hydrogen Cooled Gas Generator Cycle Engine Weight

MODE 1 PC = 272 atm (4000 psia) MODE 2 PC = 204 atm (3000 psia)

MODE 1 AREA RATIO = 40

MODE 1 Fsl = 4.0 MN (900,000 1 b)

! i

MODE 1 Fsl = 3.34 MN (750,000 l b )

1 b0 150 2-00 2-50 300 EXTENDIBLE NOZZLE AREA RATIO

Figure 50. Dual -Fuel Engine Weight Parametrics

Sea-Level Thrust ,

L Base1 ine

b 1 I I I 1 I I

I I

50 100 150 200 2 50 300 350 Chamber Pressure, psi a

I I

1 I

I I

1 1

I

1 000 2000 3000 4000 50b0 Thrust Chamber Pressure, psia

Figure 51. Mode 1 LOX/CH4 Engine Weight Parametrics

5. Mode 2 LOX/LH2 Engine

The LOX/LH2 Mode 2 engine descr ipt ion fo l lows:

Scaled SSME

Engine Cycle: Staged combustion w i t h dual f u e l - r i c h preburners . Propel lant State a t Thrust Chamber I n j e c t o r I n l e t : Fuel -ri ch gas and 1 i quid oxygen.

Combusti on Chamber: Regenerati ve ly cooled w i t h hydrogen.

Nozzle: Fixed b e l l t o E = 40 f o r sea-level operation; extendible nozzle f o r vacuum operat i on.

The weight data f o r t h i s engine i s shown on Figure 52 f x a basel ine nozzle area r a t i o o f 200:l. The data are shown as a funct ion o f t h r u s t chamber pressure and th rus t leve l . The minimum engine weight occurs a t a th rus t chamber pressure o f approximately 136 atm (2000 psia). A t a base1 i n e t h r u s t of 3.32 MN (746K I b ) , the weight increase between 136 atm and 204 atm (2000 t o 3000 ps ia) i s only 227 kg (500 I b ) ,

From the data shown on t h i s f igure , i t can be establ ished t h a t the engine t h r u s t t o weight r a t i o decreases as the engine th rus t l eve l i s increased .

The e f f e c t o f area r a t i o upon the engine weight parametrics i s shown on Figure 53. The data are presented f o r the basel ine operatfng pressure of 204 atm (3000 ps ia) .

D. PARAMETRI C ENVELOPE DATA

The elements o f engine length include:

Gimbal

I n j e c t o r and ho t gas manifold

Fixed Nozzle

Nozzle Extension (Mode 2)

Scal i ng equations based upon geometric considerations were formulated as functions of th rus t , t h r u s t chamber pressure and area r a t i o . The diameter and length paranetr ics f o r the Hode 1 and dual- fuel engines were calculated using the envelopes establ ished dur ing the Task V I pre l iminary design e f f o r t

Chamber Pressure, atma I 1

I I

1 1 I 1000 2000 3000 4obo 5000

Fixed Nozzle t o E = 40 Extendible Nozzle t o E = 200 Vacuum Thrust, MN ( l b )

Thrust Chamber Pressure, psia

Figure 52. Mode 2 LOX/LH2 Englne Weight Paramtr ics

8500 -

8000

7500.-

7000

6500-.

- 4.45 (1,000,000)

"

--

as a base. Thus, the parametrics assume a s i m i l a r engine packaging arrangement and the same percentage be1 1 nozzle lengths.

The diameter parametrics include an est imation o f the powerhead diameter (pump envelope) t o es tab l ish whether the nozzle e x i t o r t h i s envelope i s greater. For the Mode 2 engines, the powerhead diameter must be establ ished t o determine i f the nozzle a t the spl I i w i l l fit over the pump i n order t o reduce the stowed nozzle engine length.

The envelope parametrics do no t inc lude an a1 lowance f o r glmball "g.

1. Candidate Mode 1 Engines

The data shown i n t h i s sect ion arc v a l i d f o r a l l Mode 1 staged combusti on engine cycle candidates, i nc l udi ng LOXICHq when modi f i e d as noted i n the tables. This i s t r u e because the maximum v a r i a t i o n i n t h r u s t c o e f f i - c i en t between a l l M~de 1 propel lant combinations considered i n t h i s study i s 3%. This has on ly a 1-112% inf luence upon the engine l i n e a r dimensions which i s we1 1 w i t h i n pre l iminary design ca lcu la t ion accuracy.

The overa l l e f f e c t a f t h r u s t and chamber pressure upon the Mode 1 engine dimensions i s summarized on Tab12 XLV a t the design po in t area r a t i o o f 40:l. The e f f e c t o f area r a t i o upon the dimensifins i s summarized on T :~ le XLVI a t the design po in t t h r u s t chamber press,:.e o f 272 atm (4000 ps ia) . Small var ia t ions i n powerhead diameter ( less than 1 cm) resu l t i ng from engine f lowrate var ia t ions f o r the d i f f e r e n t area r a t i o s a t a given th rus t leve l and chamber pressure were neglected.

2. Hydrogen Cooled, Gas Generator Cycle Engine

She Mode 1 hydrogen cooled, gas generator cycle engine envelope data i s sumnarized on Tables XLVII and XLVIII. The engine length i s d i f - fe ren t from the staged combustion cyc le data because the chamber length required t o achieve a given energy release e f f i c i e n c y i s longer f o r 1 iqu id - 1 i q u i d propel lant i n jec t i on . The powerhead diameter and nozzle e x i t outside diameters a lso vary from the Mode 1 engine data prev iously presented because o f d i f ferences i n engine packaging and the manifolding for the hydrogen coolant a t the nozzle e x i t .

3. Dual-Fuel Engine

The dual- fuel eogine envelope parametrics are shown on Table XLIX f o r the maximum recomendea operat ing pressure o f 272 atm (4000 ps ia) i n Mode 1 and 204 atm 1;000 ps ia) for Mode 2. The data are presented for the nozzle extension i n the deployed and stowed posi t ions. The Mods 1 nozzle area r a t i o was f i x e d a t 40:l f o r t h i s study. Because the resu l t - i ng powerhead d i a ~ e t e r s were laymger than the nozzle e x i t diameter a t 40:1, the nozzle w i l l not stow over the turbomachinery. Therefore, i t

I D d N M O h r - ( V N

n - n n n n n n nnnh

nnnn n n n n hnnn L QIhQIh h p - 0 3 W d N C Q aJ V ) ( V O Z - * N O Q I I n m r . F F -

w - F C C F P - F ?

w w u W U Y V W Y W U E Q

nnnn nnnn n n n n 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 8 5 _og,o ,o _ o g s z N O *

n n n n m tn o m e h a 3

n n n n O C n N G c o o n

h h h

O N * QI OO h

h h h

03 h m m m ( U F C C Y Y W

h h h

N m m h e m F C C W U Y

h h h

0 0 0 0 0 0 0 0 0 m b V) Y Y Y

h h h

0 0 0 0 0 0

TABL

E XL

IX

. -

DU

AL-F

UEL

EN

GIN

E EN

VELO

PE P

ARAM

ETRI

CS

Mod

e 1

PC

= 27

2 at

m (

4600

psi

a)

Mod

e 2

PC

= 20

4 at

m (

3000

psi

a)

Mod

e 1

Are

a R

ati

o =

40

Mod

e 1

E

xte

nd

ible

S

tow

ed

Ext

ende

d E

xte

nd

ible

S

ea- L

evel

N

ozzl

e

Noz

zle

Eng

ine

Noz

zle

Eng

ine

Noz

zle

Pow

rrhe

ad

Th

rust

A

rea

Len

th

Leng

th

Ex

it I.D.

Dia

met

er

MN

(K l

bj

R

ati

o -

(in

. f

cm -

Jin

.) -

cm

(in

. )

2.70

I 3.

34 1

4.0 1

NOTE

:

Par

amet

rics

are

bas

ed o

n th

e T

ask

VI

pre

lim

ina

ry d

esig

n w

hich

co

nsi

ste

d o

f a

75%

bel

l n

ozz

le

to 4

0:l

(fix

ed

) co

mbi

ned

wit

h a

110

% b

e1 1

no

zzle

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119

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was assdmed i n the parametrics t h a t the nozzle could be re t rac ted t o the r e l a t i v e pos i t i on establ ished f o r the Task V I dual- fuel engine pre l iminary design.

4. Mode 2 Engine

The Mode 2 engine consists o f the same elements o f length as the Mode 1 engine plus the add i t ion o f an extendib le b e l l nozzle. For a1 1 cases, the ove ra l l nozzle ( f ixed p lus extension) was assumed t o be a 90% b e l l t runcated a t the f ixed 40:l area r a t i o .

Lengths o f other components are based upon scal ing o f values t yp i ca l of the SSME (Ref. 53). These are:

O Length from gimbal center t o i n j e c t o r face = 20"

" Chamber length = 14"

Engine envelope data, as a func t ion o f extendible nozzle area r a t i o and th rus t leve l , are presented on Table L f o r the extendib le nozzle i n both the extended and stowed posit ions. These data are summarized f o r the base1 i n e chamber pressure o f 204 atm (3000 psia). For a l l cases, the ca l - culated pump envelope exceeds the e x i t diameter o f the 40:l f i x e d nozzle. Therefore, the nozzle extension cannot be re t rac ted over the turbomachinery and associated components.! It has been assumed t h a t a l l o f these components are packaged above the th roa t plane and t h a t the nozzle extension can be re t rac ted t o t h i s po in t . This r e s u l t s i n the f i x e d nozzle po r t i on being greater than the nozzle extension a t c = 100:l but smaller than the nozzle extension a t 200:l and 300:l. Therefore, the stowed length i s equal to :

" Length o f gimbal, i n j e c t o r and hot gas manifold, chamber and the f i x e d 40:l nozzle a t E = 100

O Length o f gimbal, i n j e c t o r and hot gas manifold, chamber and the extendible nozzle a t E = 200 and 300.

Based upon the geometry and scal i n g equations, the area r a t i o a t which the nozzle extension and f i x e d nozzle lengths are exac t ly equal i s 136:l. Hence, stowed length below t h i s area r a t i o i s governed by the f i x e d nozzle.

The e x i t diameter o f the extendib le nozzle exceeds the pump envelope over the e n t i r e range o f var iables. Therefore, t h i s nozzle e x i t diameter defines the maximum engine diameter f o r the Mode 2 engines.

V) *P

C tJ 'r .c 3 n E 0 .r U) E aJ C, X aJ V) 1 C P

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SECTION V I I

TASK V - AUXILIARY COOLANT FEASIBILITY

A, OBJECTIVES AND GUIDELINES

The object ives of t h i s study were to: (1) determine the r e l a t i v e mer i t o f a u x i l i a r y cooled Mode 1 engine concepts, (2) provide cycle power balance data f o r the maximum recommended chamber pressure, and (3) e s t i - mate engine weight and envelope data.

The engine cyc le schematic analyzed i s shown on Figure 54. The base1 i ne propel 1 ants were spec i f ied as LOX/RJ-5. L iqu id oxygen, ra the r than RJ-5 i s used i n the heat exchanger t o avoid the gumning and coking problems which severely l i m i t e d the operat ing pressure of the RJ-5 cooled chamber i n Task I I.

Addit ional study guidel ines are as fo l lows :

Baseline Sea-Level Thrust = 2.70 MN (607K 1 b ) Engine Mixture Rat io = 2.7 ChanSler Guidel ines - Same as Task I 1 Cycle Evaluation Guidel ines - Same as Task I I I Auxi 1 i a r Cool ants - Water, Lithium, and Sodi um-Potassium (NaK 5 6 d Cool ant Jacket I n l e t Temperatures :

Water - 311°K (560°R) L i th ium - 456°K (820°R) NaK - 311 (560°R)

L iqu id Oxygen Heat Exchanger I n l e t Temperature - l l l ° K (200°R)

B. AUXILIARY COOLANT PROPERTIES

Analyses and l i t e r a t u r e surveys were completed and the physical and thermodynamic property data f o r the candidate auxi 1 i a r y cool ants accumulated . A summary o f t h i s data i s shown on Table L I .

The primary source o f data f o r water i s the ASME Steam Tables. Pro- per t ies f o r L i th ium and NaK (56%) are discussed i n the paragraphs which fo l low:

1 , Thermophysi cal Properties o f L i th ium

The primary source of data on l i t h i u m i s an Aerojet-General Nucleonics repo r t (Ref. 55) , This reference i s c i t e d as the AGN repo r t i n t h i s discussion.

a. Density

Water - Formul a "2* Molecular Weight 18.01534

Freezing Point, OK 273.16

(OF) (32 .O) Bo i l i ng Point, OK 373.16

(OF). (212) C r i t i c a l Temperature, OK 647.31

(OF) (705.47) C r i t i c a l Pressure, '

M N / ~ Z

C r i t i c a l Density, kg/m3

Vapor Pressure, L iqu id a t 298.15K, 3.166 k~/m2)

( a t 77OF, ps ia) (0.4592)

TABLE L I . - PROPERTIES OF CANDIDATE AUXILIARY COOLANTS

Density, L iqu id

a t 298.15K, kglm 997.1

( a t 77OF, 1 b / f t 3 ) (62.247)

Heat Capacity, l i q u i d

a t 298.15K, J1g.K 4.177 ( a t 77OF, Btu/lb°F) (0.9983)

Viscosi ty , L iqu id a t 298.15K, mN.S/m 2 .89O4

Sodium-Potassium (NaK 56%)

Na,572 K.428

29.886

$278.16

( 4 1 ~ 1 0 8 6

( ~ 1 494) ~ 2 5 4 2 2 40**

( 4 1 16 2 72**)

-30.9**

( a t 77OF, lb,,,/ft-sec) (.5983 x (5.596 x

Thermal Conduct iv i ty , L iqu id

a t 298.15K. W1m.K ,6095 ~ 2 2 . 5

( a t 77OF, Btu/ ( 9 . 7 8 3 ~ 1 0 - 5 ) (53.61 X I O - ~ ) f t .sec°F)

L i t h i um

L i

6.939

453.69 (356.95)

%I61 5

(52448)

31 73 t 400

(5252 2 720)

*Values a t the f reez ing po in t . **Psuedocri t i c a l value calcu lated by Kay's r u l e .

(1) So l i d L i th ium

The densi ty o f sol i d 1 i thium has been obtained over the temperature range o f -269°C t o the me1 t i n g po in t (180.54OC). The data developed by three invest igators were combined by AGN t o develop a co r re la t i ng equation covering the e n t i r e temperature range. The recomnznded equation i s given below:

p(s) = 0.560 + 9.243 x I O - ~ T (1 5) 3 where p(s) i s i n g/cm and T i s i n degrees Celsius. The standard dev ia t ion

o f the f i t i s 0.026 g/cm3. The confidence range a t the 95% confidence l eve l i s + 0.048 g / c d near the mel t ing point , 2 0.028 g / c d a t O°C, and 5 0.051 g/ci? a t -273.15"C.

(2) L iqu id L i th ium

The densi ty o f 1 i q u i d 1 i th ium has been measured over the temperature range from the me1 t i n g po in t (180.54OC) t o 1125OC. The data developed by four invest igators are i n very good agreement and were combined by AGN t o develop a co r re la t i ng equation covering the e n t i r e temperature range. The recomnended equation i s given below:

I I 3 where p 1 i s i n g/cm and T i s i n degrees Celsius. The standard dev ia t ion

of the i i s 0.0018 g/cm3. The confidence a t the 95% confidence l eve l i s + 0.0040 g/cm3 near the mel t ing point, 0.0032 g / c d a t midrange, and + 0.0012 - - g / c d i s the highest temperature range.

b. Speci f ic Heat

(1) So l id L i th ium

Data on the spec i f i c heat o f sol i d l i t h i u m have been c o l l ected by the Purdue Un ivers i ty Thermophysical Propert ies Center and corre lated t o the f o l lowing recommended equation f o r the temperature range o f 200°K t o the mel t ing po in t (453.6g°K).

where C P ( ~ i s i n cal/g.OK and T i s i n O K . The standard dev ia t ion o f the 4 f i t i s 0.0 21 cal/g.OK.

(2) L iqu id Li thium

The enthal py o f 1 i q u i d 1 i thium has been measured w i t h reference t o the i c e po in t over various temperature ranges from near the mel t ing po in t (180.54OC) t o 1226.8OC. The data developed by three i n v e s t i - gators are i n good agreement and were combined by AGN t o develop a co r re la t i ng equation ( t h i r d order polynomial) covering the e n t i r e temperature range. The recomnended enthal py equation i s given below:

where ( H ~ - H ~ ~ ~ ~ ~ ) (n) i s i n ca l /g and T i s i n degrees Celsius.

An equation for the heat capacity ( s p e c i f i c heat) o f l i q u i d 1 i t h i um was derived from the above equation by d l f f e r e n t i a - ti on t o give the f o l lowing recommended equation :

where C ~ ( 9 "

i s i n cal/g.OC and T i s i n degrees Celsius. - . .

c. Thermal ~ o n d u c t i v i ty

The ava i lab le data on the thermal conduct iv i ty of s o l i d and l i q u i d l i t h i u m has been c r i t i c a l l y analyzed a t the Themnophysical Properties Research Center a t Purdue Un ivers i ty . The recomnended values developed from t h a t analysis are presented on Page 360 o f Ref, 56.

d. Viscosi ty

The v i scos i t y o f l i q u i d l i t h i u m has been measured over the temperature range from the me1 t i n g p o i n t (180.54OC) t o 1 181°C. The data developed by s i x inves t iga tors are i n f a i r agreement and were combined by AGN t o develop a co r re la t i ng equation covering the e n t i r e temperature range. The recommended equation i s :

where ~ ( n ) i s i n m i 11 ipo ise and temperature i s i n degrees Kelvin. The r e l a t i v e standard deviat ion o f the f i t i s 9.1%. The confidence range a t the 95% confidence leve l i s + 0.25 m i l l i p o i s e near the mel t ing point , * 0.11 m i l l i p o i s e a t mid-range, and + 0.10 m i l l i p o i s e i n the highest range invest igated.

e. Vapor Pressure

The vapor pressure o f 1 i q u i d 1 i thium has been measured over the temperature range of 935 t o 1608OC. The data developed by f i v e invest igators are i n f a i r agreement and were combined by AGN t o develop the co r re la t i ng equation covering the e n t i r e temperature range. The recommended equati on i s :

where P a ) i s i n atmospheres and T i s i n degrees Kelvin. The r e l a t i v e standar 6 deviat ion of the f i t of the vapor pressure data i s 3.3%. The confidence range a t the 95% confidenc l eve l i s i 1.2 x 10-4 atm f o r the lowest temperature range, k 8.2 x 10-$ atm i n the mid-range and t 4.5 x 10-3 atm i n the highest temperature range.

The normal b o i l i n g po in t ca lculated from the above co r re la t i ng equation i s 1342.17OC + 0.12°C a t the 95% confidence l e v r ;

2. Themphysical Properties o f NaK (56)

a. Density o f L iqu id NaK (56)

The densi ty o f l i q u i d NaK (56% w t K) i s calculated from the fo l lowing equation (Ref. 57) :

Using data f o r ~k and PNa from Ref. 57 the densi t y o f NaK (56) has been calculated accordingly. These data are s u m r i z e d i n Table L I I .

The data i n the tab le are c lose ly f i t t e d by the fo l low- i n g 1 inear equation :

' ~ a ~ ( 5 6 ) = 0.8997-2.376 x I O - ~ T (Tln°C) (23) where o i s i n g/cm3

b. Viscosi ty o f L iqu id NaK(56)

The v i scos i t y of l i q u i d NaK (56% wt K) i s calculated from the fo l lowing equation (Ref, 57) :

Using dats f o r vk and PNa from Ref. 57 the v i scos i t y o f NaK(56) has been calculated accordingly. These data are s u m r i z e d i n Table L I I I .

I n the temperature region up t o 400°C (673.16OK) the v iscos i ty data o f NaK(56) given i n the tab le are c lose ly f i t t e d by the f o l 1 owing equation:

632.01056) (25) ~ ( c e n t i p o i s e ) = (0.102079-2.0344 x IO-~T) exp (0,10438 + ,-

where T i s i n O K and I 673.16OK

c. Thermal Conductivity o f NaK(56)

The thernal conduct iv i ty o f a NaK very c lose ly approxi- mating NaK(56) has been pub1 ished by Ewing , et . a1 . (Ref. 58) f o r the temperature range o f 200 t o 500°C. Ewing's data have been extended t o higher and lower temperatures based on data f o r NaK(78) given i n Ref. 57. The racomnended thermal conduct iv i ty f o r NaK(56) can thus, be oxpressed by the fol lowing equation: .

Temp., O C

5 0

100

1 50

TABLE L I I . - DENSITY OF L I Q U I D NaK(56)

Dens i t y , c ~ / c r ~

(0.887)

0.8750

Temp. , O C Density , CJ/CIII~ -- 400 0.8052

Temp., O C

SO

TABLE LIII. - VISCOSITY OF LIQUID NaK(56)

Viscosity, Cent i poi se Temp., O C

400

450

500

5 50

600

65C

700

Vi scosi t y , Centi poise

-7 2 K(W/cm.OC) = 0.220 + 2.07 x I O - ~ T - Z . ~ x 10 T (26)

where T i s i n O C

This equation represents Ewing's data w i t h i n an average dev ia t i on o f 0.4% and maximum dev ia t i on of 1.4%.

d. Spec i f i c Heat of NaK(56)

The s p e c i f i c heat o f l i q u i d NaK a l l oys can be c a ~ i z ! a t e d by the fo l lowing equat ion according t o (Ref. 57) :

where XNa and Xk are the weight f r a c t i o n s o f Na and K, respec t i ve ly , and Cp(Na) and Cp( ) are the speci f i c heats o f Nz and K, respect ive ! y.

Based upon t9e heat capaci ty equations f o r Na and K given i n (Ref. 57), the heat capaci ty equat ion f o r l i q u i d NaK(56) i s evaluated t o >o as fo l lows :

-7 2 ' ~ ( N a ~ - 5 6 ) = 0.26325-1.1017 x I O - ~ T + 1 .I0026 x 10 T (28

where Cp i s i n cal/g."C and T i s i n "C

e. Vapor Pressure of L i q u i d NaK(56)

The vapor pressure o f NaK(56) can be ca lcu la ted from t h a t o f i t s const i tuents assuming the v a l i d i t y o f Raou l t ' s law (Ref. 57). Based upon vapor pressure co r re l a t i ons fo r the cons t i tuen ts (Ref, 57) , the vapor pressure of N a ~ ( 5 6 ) was ca lcu la ted a t 100°C increments f rom 100°C t o 1200°C. Those ca lcu la ted values were then cu rve - f i t t e d t o y i e l d the f o l l o w i n g vapor pressure expression:

where T i s i n OK

The average dev ia t i on between the o r i g i n a l ca lcu la ted values and those c a l - cu la ted from the above equation i s 2.44%. The maximum dev ia t i on i s about 6% b u t occurs a t a temperature where the vapor pressure i s very low (%4 x 10-6 atm a t 200°C). The ca lcu la ted normal b o i l i n g p o i n t i s 1086.5"K which i s i n very good agreement w i t h experimental values.

C. MATERIALS COMPATIBILITY

1. Water

Corrosion data most app l i cab le t o the advanced h igh pressure engine water coo l ing system i s t h a t determined f o r the s t r u c t u r a l mate r ia ls i n h igh pressure, water cooled nuc lear reactors . Most o f t h i s t e s t i n g has

been conducted on steels , s ta in less s tee ls and n i cke l base a l loys i n auto- claves and r e c i r c u l a t i n g loop systems a t temperatures t o 617OK (650°F). It was noted t h a t the r e c i r c u l a t i n g systems requi red f i l t e r s f o r the removal o f insoluble corrosion products and t h a t water p u r i t y i s essent ia l f o r adequate resistance of the aus ten i t i c s ta in less s tee ls t o s t ress corrosion cracking (Ref. 59). Both hydroxide and ch lo r ide contaminants w i l l s tress crack these materials. The st ress corrosion cracking threshold temperature i s 394OK (250°F) f o r 347 and 316 s ta in less steels i n water con- t a i n i n g 875 ppm NaCl (Ref. 60). The s u s c e p t i b i l i t y o f the aus ten i t i c s ta in less steels, i nc lud ing Incoloy 800, t o s t ress corrosion cracking a t 478°K (400°F) i n water conta in ing 10 ppm NaCl and 12-17 ppm 02 i s shown by Ref. 61. Nickel-chromium a l l oys must be u t i l i z e d i f water p u r i t y cannot be assured.

O f the high conduct iv i ty mater ia l candidates f o r the ccol i n g system, i .e., aluminum a l loys , coppers, and n ickel , the l a t t e r two metals are considered sa t i s fac to ry f o r serv icz i n high temperature water although stress corrosion data are no t general ly ava i lab le and should be determined. Aluminum a1 loys have low st rength a t h igh operat ing temperatures, and are suscept ible t o p i t t i n g corrosion. The corrosion ra tes f o r copper a t 373OK (212OF) and n icke l a t 42Z°K (300°F) are reported a t less than 2 m i l 3 per year (MPY) i n neutra l pH water (Ref. 62). Tests conducted w i t h copper base a l loys, namely adri-a1 t y brass and cupronickels a t 478OK (400°F) i n neutra l pH water containing less than 10 ppm 02 indicated n e g l i g i b l e corrosion ra tes (Ref. 62). However, over 90% o f the metal i n the corrosion product was released t o the system.

Titanium a l l oys and 17-4PH HI025 s tee l are recomnended f o r pump components where h igh fa t i gue strengths and c a v i t a t i o n resistance are required. The t i tan ium a1 loys are preferred fo r a 1 ightweight design and fo r high resistance t o corrosion fa t i gue i n the event o f water contamination.

Galvanic corrosion e f fec ts should be minimal i n the coo l ing system provided aluminum a l l oys are no t u t i l i z e d as mater ia ls o f con- s t ruc t ion . Galvanic corrosion may be enhanced by water impuri t i e s other than 02 through t h e i r a c t i v a t i n g o f the passive surfaces of the s ta in less steels and the n icke l -chromium a l loys .

2. NaK

Two types o f corrosion can occur i n NaK coolant c i r c u i t s ; i .e., mass t rans fer and in te rgranu lar at tack. The f o m r corrosion mechanism involves the so lu t i on o f the container mater ia l i n the ho t por t ion o f the sys tem and deposi ti on i n the cool po r t i on *and occurs when exposure times are s u f f i c i e n t t o reach saturated solut ions i n the ho t por t ion and supersaturated solut ions i n the cool por t ions. I n t e r - granul a r attack i nvo l ves the formati on o f corrosion products w i t h i n the gra in boundaries which embri t t l e and/or weaken the materi a1 . Mass t ransfer corrosion can r e s u l t i n r e s t r i c t i o n s o r plugging o f the coolant c i r c u i t .

The a u s t e n i t i c s t a i n l ess s tee l s and the n i c k e l base a l l o y s are s a t i s f a c t o r y f o r long t ime serv ice t o 978°K (1300'F) (Ref. 63). OFHC copper i s acceptable f o r l ong t ime se rv i ce a t (700°F) and 867°K (1100°F) f o r sho r t t ime serv ice (Ref. 64). N icke l i s acceptable as a h i gh conduct i - v i ty mater ia l i n sodium (Ref. 65). T i tan ium i s r a t e d as acceptable f o r long term use :o 867°K (1100°F) and f o r sho r t t ime use t o 1144°K (1600°F) (Ref. 64).

Data on the e f f e c t o f NaK on the low cyc le f a t i g u e l i f e o f t he engineer ing a l l o y s i s n o t genera l l y ava i lab ie . Na does n o t impa i r the low cyc le fa t igue l i f e of 316 s ta i n l ess s tee l (Ref. 63). Tests conducted on austeni t i c s t a i n l ess s tee l s and n i c k e l base a1 loys t o determine t h e i r s u s c e p t i b i l i t y t o long term c a v i t a t i o n damage i n 1089°K (1500°F) Na ind ica ted t h a t the mater ia ls were s a t i s f a c t o r y w i t h the n i c k e l base a1 loys sus ta in ing the l e a s t c a v i t a t i o n damage (Ref. 66).

3. L i t h i um

Metals i n molten li thium are sub jec t t o the same type o f a t t ack experienced i n Na and NaK. The austeni t i c s t a i n l ess s tee l s e x h i b i t severe s o l u t i o n corrosion, i n t e rg ranu la r a t tack and mass t r a n s f e r above 918°K (1300°F) (Ref. 63) and appear t o be s u i t a b l e f o r 1 i m l t e d serv ice a t 867°K (1100°F) (Ref. 64). The n ick21 base a1 loys are more sever ly at tacked and are n o t recomnended f o r 1 i t h i um serv ice. Stress corros ion t e s t s conducted on a n u h e r o f a l l o y s i n l i t h i u m a t 589°K (600°F) and 756°K (960°F) showed t h a t t i t a n i u m and the a u s t e n i t i c s t a i n l ess s tee ls are h i g h l y r e s i s t a n t t o s t ress corros ion wh i le the n i c k e l base a l l o y s are h i g h l y suscept ib le (Ref. 67). Copper and i t s a l l o y s are l i s t e d as possessing poor res is tance t o l i t h i u m a t 575°K (575"F), on ly 378°K (220°F) above the me l t p o i n t o f l i t h i u m (Ref. 64). Copper i s r e a d i l y d isso lved i n l i t h i u m as evidenced by the copper-1 i thium phase diagram. The use o f copper o r i t s a1 loys i n molten l i t h i u m w i 11 be 1 i m i t ed t o sho r t t ime serv ice, and corros ion r; and mass t r a n s f e r cha rac te r i s t i c s must be determined t o es tab l i sh the ti ;?d temperature 1 i m i t s . D. CHAMBER COOLING

Chaher thermal design analyses s tud ies were conducted f o r the th ree candidate a u x i l i a r y coolants. The chamber pressure range considered was 272 t o 340 atm (4000 t o 5000 ps ia ) . The Task I 1 r e s u l t s i nd i ca ted t h a t chamber pressures o f a t l e a s t 272 atm (4000 ps ia ) could be achieved us ing one o f the main engine p rope l lan ts as cool ants . Therefore, i f the a u x i l i a r y cooled cyc le i s t o be compet i t ive, h igh chamber pressure must be achieved because of the added sys tern we! yh t and complexity .

The chamber designs d i f f e r e d from those o f the Task I 1 s tud ies i n two ways. I n order t o prov ide s impler designs f o r Task V, the coolant was assumed t o flow i n one pass from the i n j e c t o r end t o the e x i t a t area r a t i o 40: 1. The number of cool an t channels upstream of area r a t i o 7.6: 1 was f i x e d a t 300 f o r a l l Task V designs i n order t o p rove ., chanriel widths greater than about 1.52 mrn (0.060 i n . ) As a r e s u l t , a l l channel depth lwidth r a t i o s

i n the high heat f l u x regions are we l l below the 4:1 l i m i t spec i f ied i n the Task I 1 guidelines.

Design c r i t e r i a and heat t rans fer cor re la t ions used f o r the auxi 1 i a r y cool ants are described below.

1 . Water Cool i ng

An i n l e t pressure o f 190 atm (2800 ps ia) was selected i n order t o provide subcri t i c a l operat ion w i th high sa tura t ion temperatures, wi thout ge t t i ng so close t o the c r i t i c a l pressure t h a t the heat t rans fe r cor re la t ions used might be i nva l i d . Operati on a t supercr i t i c a l pressures provides no advantages wi thout considering i n d e t a i l the reduced heat t rans fer coef- f i c i en ts and u l t imate heat f l u x l i m i t s which occur when the coolant-side wa l l temperature exceeds the c r i t i c a l temperature; such considerations were beyond the scope o f the present study. I n order t o maintain nucleate bo i l i ng , l oca l coolant ve loc i t i es i n the combustion chamber were determined such t h a t maximum coolant heat f luxes d i d no t exceed 80 percent o f the bus nout heat f l u x . The burnout co r re la t i on used herein was taken from Ref. 68. With nucleate boi 1 i n g contro l 1 i n g the coolant-side wal l temperature t o a maximum o f 275°K (35°F) above the l oca l saturat ion temperature, gas-si de wa l l temperatures were below 867°K (1100°F f o r a l l chanb r pressures considered. The D i t tus- Boel ter co r re la t i on (Tab 1 e I o f Ref. 457 w i t h bulk coo;ant proper t ies was used f o r forced convection heat t rans fer t o water, i.e., when the coolant- s ide wal l temperature was less than the sa tura t ion temperature.

A water f low r a t e o f 59 kg/sec (130 l b l sec ) was selected based on prov id ing an e x i t subcooling o f a t l e a s t 272°K (30°F) w i t h 10 percent f low mald is t r ibu t ion . A water i n l e t temperature o f 311°K (100°F) was assumed.

2. L iqu id Metal Cooling

A review of the l i q u i d metal heat t rans fer l i t e r a t u r e ind icated t h a t the most popular ly accepted co r re la t i on i s o f the form:

where:

Nu = Nusselt number Re = Reynolds number Pr = Prandtl number b = Bulk rnin = Minimum

w i th Numi varying from 4.3 t o 7.0. The present studies are based on a Numin o f B.n as recommended i n Ref. 69 for a number o f 1 i q u i d metals, inc uding dium-potassium a: loys. No experimental studies o f 1 i th ium heat t rans fer were found.

Gas-side wal l temperatures were 1 imi ted t o 867°K (llOO°F), as i n Task 11, due t o the s i g n i f i c a n t reduct ion i n zirconium-copper a l l o y mechan-

ica! p roper t ies a t h igher temperatures. I n order t o meet t h i s l i m i t a t area r a t i o 40:l w i t h small pressure gradients, out1 e t bu l k temperatures were 1 im i t ed t o 756°K (900°F). Based on i n l e t temperatures o f 456OK (360°F) f o r l i t h i u m and 311°K (100°F) f o r NaK, the r e s u l t i n g f l o w ra tes were 54 and 154 kg/sec (120 and 340 I b/sec ) respec t i ve ly .

3. Results and Conclusions

Channel depth p r o f i l e s and r e s u l t a n t pressure drops were generated based on the above f l ow rates, heat t r a n s f e r co r re l a t i ons and design c r i t e r i a . Pressure drop, coolant bu l k temperature r i s e and maximum wa l l temperature r e s u l t s are shown on Table LIV.

A l l th ree a u x i l i a r y coolants were found t o be acceptable from a chamber coo l ing s tandpoint f o r chamber pressures as h igh as 340 atm (5000 p s i ) . Table LIV shows t h a t water and NaK pressure drops are s i m i l a r and are i n the 20 t o 48 atrn (300-700 p s i ) range f o r the chamber Fressure range studied. L i t h i um pressure drops are i n the 7 t o 20 atm (100-300 p s i ) range. Note t h a t t hz nuc leate b o i l i n g i n the water cooled designs maintains the maximum w a l l temperature below the 1 i m i t imposed on the l i q u i d metal cooled designs. Although the l i t h i u m r e - qu i res the lowest f l ow r a t e and pressure drop, i t s temperature must be maintained above 454°K (357°F) t c prevent f reez ing. I n add i t ion , i t should a lso be emphasized t h a t l i t h i u m i s incompat ible w i t h copper and i t s a1 loys a t temperatures above 575°K (575°F) f o r long term use.

E. ENGINE DATA

I n add i t i on t o the coolant jacke t p resswe drop data and cyc le eva luat ion gu ide l ines, i t was necessary t o conduct p re l im ina ry heat exchanger analyses t o evaluate the pressure drop o f t h i s component p r i o r t o conducting the engine power balances . The A P estimates re - s u l t i n g from t h i s analyses are as f o l l ows :

HEAT EXCHANGER. AP . . -

AUX . LOX, AUXT 6 0 5 ~ ~ ~ COOLANT .-- atm ( p s i ) atm ( p s i )

Water 15.6 230) t 102 (1500) L i t h i um 5.4 80) 2.0 (30) NaK 14.3 (210) 5.4 (80)

The pressure drop o f each coolant through the heat exchanger was ca l - cu la ted by determining the design v e l o c i t y necessary t o avo id f reez ing the coolant. LOX, a t a 367°K (200°R) i n l e t temperature, i s the o ther heat exchangsr f l u i d .

For each o f the auxi 1 i a r y cool ants, engine cyc le power balances were conducted which resu l t ed i n the f o l l o w i n g main pump discharge pressure requi rements .

TABLE LIV. - SUMMARY OF TASK V CHAMBER DESIGNS

Cool ant

Flow Rate, kg/sec ( I blsec)

Bulk Temperature Rise, OK, (OF)

PC = 272 atm (4000 ps ia)

PC = 340 atm (5000 psia)

Out le t Temperature, OK (OF)

PC = 272 atm (4000 ps ia)

PC = 340 atm (5000 psia)

Pressure Drop, atm (ps i )

PC = 272 atm (4000 psia)

PC = 306 atm (4500 psia)

PC = 340 atm (5000 ps ia)

Maximum Wall Temperature, OK (OF)

PC = 272 atm (4000 psia)

PC = 340 atm (5000 psia)

Water L i th ium

Pump Discharge Pressure, atrn ( ps i a ) Aux.

Cool an t - LOX - RJ -5 Cool an t

Water, PC = 272 atm (4,000 ps i a ) 497 (7,300) 478 (7,030) 145 (2126)

PC = 340 atm (5,000 ps i a ) 748 ( 1 1,000) 702 (10,320) 165 (2420 j

NaK, PC = 272 atm (4,000 ps i a ) 495 (7,280) 478 (7,030) 37 (545)

PC = 340 atm (5,000 ps i a ) 762 (11,200) 710 (10,440) 67 (993)

Li th ium, PC = 272 atm (4,000 ps i a ) 483 (7,100) 475 (6,980) 19.7 (290)

PC = 340 atm (5,000 ps i a ) 714 (10,500) 692 (10,170) 34.1 (502)

With the discharge pressure requirements determined, engine weight and envelope est imates werc made. For the same LOX pump discharge pressure as the base l ine LOX cooled, Mode 1 engine, the a u x i l i a r y cooled cyc les can achieve a h igher t h r u s t chamber pressure ( i .e., 313 atm (4600 ps i a ) vs 272 atm (4000 p s i a ) ) . However, f o r comparative purpose t o the o ther Mode 1 engines considered i n Task 111, a chamber pressure o f 272 atm (4000 ps i a ) was se lected f o r the data summary.

Table LV summarizes p e r t i n e n t auxi 1 i ary cooled engine cyc l e data. A t a t h r u s t chamber pressure o f 272 atm (4000 ps ia) , the engine performance i s i d e n t i c a l t o the base l ine engine, The maximum d i f f e rence i n a u x i l i a r y cooled system weights i s about 136 kg (300 1bs .) The auxi 1 i a r y cooled system weight inc ludes the f o l l o w i n g components:

' Pump and Turbine ' Heat Exchanger

Accumulator ' Lines ' Coolant

Because the main pump discharge pressures are reduced compared t o the basel ine, t he bas ic LOXIRJ-5 components weigh 2058 kg (4537 l b s . ) . However the weight o f the a u x i l i a r y LJmponents r e s u l t i n increased engine weight compared t o the basel ine as shown below:

Weight, kg ( lbs . ) Basel i ne

Cool an t LOX - Water L i t h i um - NaK

Basic Engine 2130 (4696) 2058 (4537) 2058 (4537) 2058 (4537) Aux. Cool an t Sys tem - -- - - _ C

310 (683) - 277 (610) 177 (390)

To ta l 2130 (4696) 2335 (5147) 2235 (4927) 2368 (5220)

TABLE LV. - TASK V AUXILIARY COOLED ENGINE SUMMARY

Engine Sea-Level Thrust, MN (1 b)

Engine Speci f ic Impul se, sec

Sea-'-eve1

Vacuum

Chamber Pressure, atm (psia)

Nozzle area r a t i o

Engine Length, cm ( in . )

Nozzle E x i t Dia, cm ( in. )

Pump Envelope Dia, cm ( i n . )

Engine Weight, kg ( l b )

Water

L i t h i l m

Na K

)Most v iab le candidate.

This data shows t h a t the weight d i f fe rences between a u x i l i a r y cooled systems are minor and the NaK system i s the heav iest because f l ow r a t e requirements are the highest. This r e s u l t s i n a greater wet weight.

Although the l i t h i u m cooled system r e s u l t s i n the l i g h t e s t engine weights, operat ional problems would seem t o rec lude i t s use. For example, the me l t l ng p o i n t o f l i t h i u m i s 454OK (3570FY. Th is means t h a t upon engine shutdown, the l i t h i u m w i l l s o l i d i f y i n the system unless i t i s heated. This, o f course, would create a severe power d r a i n on the veh ic le . Because o f handl ing and sa fe ty considerat ions, i t i s advisable t o mainta in a c losed envi ronrrent w i t h both 1 i thium and NaK. Therefore, dumping o f these cool ants upon engine shutdown i s n o t f eas ib l e , Water can be dumped t o avo id i t s f reez ing. This can r e s u l t i n some e u i v a l e n t weight saving s ince the weight o f the water, approximately 72.6 kg 160 I b s ) , need n o t be c a r r i e d through- ou t the e n t i r e ? shi c l e t r a j e c t o r y .

9 Because ~f the hand1 ing , operat ional and c o r r o s i v i ty problems

associated w i t h us ing l i t h i u m and NaK and weight d i f fe rences are r e l a t i v e l y mi nor, water i s obv ious ly the super io r auxi 1 i ary cool an t .

SECTION V I I I

TASK V I - ENGINE PRELIMINARY DESIGN

Based on the resu l t s of the parametric studies, concepts were selected t o be invest igated and character a pre l iminary engine design e f f o r t . I n addi t ion, RP-1, was selected as the basel ine f u e l f o r a l l three engines

The three selected engine concepts are:

1 . Base1 i ne LOX/RP-1 Mode 1 Engine

three engine ized f u r t h e r i n ra the r than RJ-5,

2. Dual-Fuel Engine 3. A1 ternate Mode 1 Hydrogen Cooled, Gas Gererator Cycle Engine


The object ives o f t h i s task were t o provide engine pre l iminary designs and data f o r the three candidate engine concepts. The designs include pre l iminary layouts o f the engine a s s e h l i e s and the major engine components. The data includes the performance, mass property, envelope, and operat ional charac ter is t i cs o f the engines and t h e i r components.

The guide1 ines used i n t h i s pre l iminary design study are:

' Enqi ne Performance

Based upon JANNAF Methodology

Engine System Pressure Losses

Same as Task 111

Thrust Chamber

Same as Task I 1 (Table XXXIII).

' High Pressure Pumps

Inducers and/or impel lers u t i l i z e d i n the high pressure pumps s h a l l be designed f o r operat i on above i n c i p i e n t cav i ta t ion .

Impel ler burs t speed s h a l l be a t l eas t 20% above the maximum operat ing speed.

Impel ler e f f e c t i v e st ress a t 5% above the maximum operat ing speed s h a l l no t exceed the allowable .2% y i e l d s t ress. (Does not apply t o areas i n which l oca l y i e l d i n g i s permitted.)

Low Pressure Pumps

I n l e t Flow Coe f f i c i en t : .06 (Minimum) I n l e t Flow Maximum Ve loc i t y :

WOE 1 FUELS, = p3F Turbines

Blade r o o t steady-state s t r ess s h a l l n o t exceed the al lowable 1% 50 hour creep s t ress .

Stress s t a t e a t the blade r o o t as def ined by the steady- s t a t e s t ress and an assumed v i b r a t o r y s t ress equal t o the gas bending s t ress s h a l l be w i t h i n the a l lowable s t ress range diagram o r modi f i ed Goodman d i agram.

No blade na tu ra l frequencies w i t h i n ' 152 of known sources o f e x c i t a t i o n a t steady-state operat ing speeds.

Disk b u r s t speed s h a l l be a t l e a s t 20% above the maxilnum operat ing speed.

Disk maximum e f f e c t i v e s t ress a t 5% above the maximum operat ing speed s h a l l n o t exceed the a1 lowable .2% y i e l d s t ress . (Does n o t apply t o areas i n which l o c a l y i e l d i n g i s permit ted.)

Bearinqs

Turbopump designs s h a l l u t i l i z e r o l l i n g element bear ings.

Maximum DN: Same as Task 11; (Table XXXVIII)

Seals - Turbopump dcs i gns s h a l l u t i 1 i ze conventional type seals . Face Contact Seal Maximum PV, FV, and PfV Factors: Same

as Task I 1 1 (Table XXXIX).

General

Components which are subject t o a low cyc lc f a t i gue mode of f a i l u r e sha l l be designed f o r a minimum o f 250 cycles times a safety f a c t o r o f 4.

Components which are subject t o a f rac tu re mode o f f a i l w e sha l l be designed f o r a minimum o f 250 cycles times a sa fe ty f a c t o r of 4.

Components which are subject t o a h igh cyc le fa t i gue mode o f f a j l u r e sha l l be designed w i t h i n the allowable st ress range diagram (based on the material endurance 1 im i t .) I f st ress range mater ia l property data are n o t avai lab le, modif ied Gcodman diagrams constructed as shown below s h a l l be st! l i zed .

Allowable Alternating Stress Line

.I

.I a L

'. Y

1:1 Ratio V)

01 C .... \ i L

0 Y C

-\ <

Hp;rn Stress h u

' Ft" Lower' of F t ~ 1.6

or - 1 . I

Fa = Haterial Endurance Limi t . Fty= Material Y i e l d S t rmgth (.P offset ) Ftu= Material Ultimate Strength

Ef fec t i ve stress sha l l be based on the Mises-Hencky constant energy of d i s t o r t i on theory.

Unless otherwise noted under component ground ru les speci- f i e d herein, the fo l lowing minimum fac tors o f safety s h a l l be u t i l i z e d :

Factor of Safety ( .2% y i e l d ) = 1.1 x L i m i t Load Factor o f Safety (Ul t imate) = 1.4 x L i m i t Load L i m i t Load: The maximum pred ic ted load o r pressure a t the most c r i t i c a l opera t ing condi ti ?n .

Components su5 jec t t o pressure loading s h a l l be designed t o the f o? lowing minimum proo f and b u r s t pressures:

Proof Pressure = 1.2 x L i m i t Pressure Burs t Pressure = 1.5 x L i m i t Pressure

B. BOOST PUMP DRIVE SELECTION

Low Speed Pump u r i v e Yethod

I n order t o minimi zs turbomachinery weight, a low speed boost pump i s used t o permi t the main o r h igh speed turbopump t o operate a t i t s optimum speed. The purpose o f t h i s study was t o determine the ~ o s t a p p r o p i a te d r i v e method f o r the low speed boost pump. This inc luded trade-offs t o determine the e f f e c t o f a l t e r n a t e boost pump d r i v e methods and design p o i n t n e t p o s i t i v e suc t ion head l e v e l s on engine weight, com- p l e x i t y , performance, s t a r t t r a n s i e n t cha rac te r i s t i c s and boost pump i n l e t diameter .

Tb? boost pump d r i v e methods considered are:

a. Gear Dr ive b. HydraulSc Turbine 2 r tve c. Gas Turbine D r i .e

The hydrau l i c t u rb i ne i s powered by f l u i d tapped o f f from an i n te r s tage s t a t i o n of the h igh speed pimp.

The gas t u rb i ne i s powerxd by h o t gas from the preburner. I n order t o avoid complex seal problems, gas from the f u e l - r i c h preburner i s used i n the low speed f u e l pumps and gas from the oxidizer-r i ! .h p r@- burner t o d r i v e the low s ~ e e d LOX pump. No other gas supply was con- s idered i n the study.

Only s i ng le reduct ion gear t r a i n con f igura t ions were considered i n the gear t r a i n study.

The engine boost pump trade study was i n i t i a t e d p r i o r t o the select ion o f RP-1 as the baseline fue l . Therefore, much o f the e f f o r t was concentrated on the o r i g i n a l LOX/RJ-5 base1 i ne engine. However, the resu l t s and select ions made are no t a f fec ted by the fue l change t o RP-1.

a. Cycle and Engine Performance Considerations

As ant ic ipated, the low speed purq power requi re- l e n t i s qu; te small compared t o the high speed pump power requirement. This i s i 11 u s t r ~ t e d below f o r the o r i g i n a l basel ine LOXIRJ-5 e n g i t , ~ :

FLUID I M ( f t ) 1 SPEED SPEED I I

*Estimated low speed .mp head r i s e required f o r a, gh speed pusp suct ion spec i f i c speed o f 4.14 RPMX (M3/sec) .?jxM-3/4 (20,000 RPM x G P M ~ / ~ x Ft-3/4)

LOW SPEED HEAD RAT1 0

I I #PSH

The energy ex t rac t ion from the cycle required t o d r i v e the boost pumps i s d ' rec t ly depeqdent on the e f f i c i e n c y o f the d r i ve system. Estimated d r i ve ef f ic iencies are :

Pump Head Requi rement , M ( f t )

LOW * HIGH '

DRIVE METHOD .- - DRIVE EFFICIENCY

Gears 0 -98 Hydraulic Turbine 0.50 - 0.65 Gas Turbine 0.50 - 0.70

A1 thcugh the gear d r i v e i s the most e f f i c i e n t , the e f f i c i ency of the turbine dr ives are high enough t o warrant f u r t h e r con- s iderat ion. A range o f e f f i c i e n c i r s i s ind icated f o r the turb ine dr ives t o i l ' u s t r d t e the e f f e c t o f the turb ine type on e f f i c iency . For the gas turbines, turb ine t i o speed and the number o f stages are the p r i n c i p l e var.ia5:es. Due t o the low speed o f the boost pump, optimum t i p speods requi re r e l a t i dely large diameter turbines and consequently r e s u l t i n high weight.

For the hydraul ic turb ine dr ive, the e f f i c iency indicate^ i s the ~ r o d u c t o f the turb ine e f f i c i ency and the e f f i c i ency ~f the pump supplying the working f l u i d , As w i th the gas turbine, the e f f i c iency achievable i s dependent primari5, on t i p speed and number of stages.

Because the boost pump d r i ve system power requi rement Ss smi l l , differences between the boost pump d r i ve methods on the engine cyc:e balance and performance i s neg l i q ib le f o r any reasonably

achievable d r i v e e f f i c i ency . Therefore, the choice o f a boost pump ,dr ive method must be based on other considerations.

b. S t a r t Transient and System Complexity

A gas tu rb ine dr ive, i n which the hot gas i s b led from the preburner, would undoubtedly requ i re hot gas valves f o r f low cont ro l dur ing the engine s t a r t t rans ient . As present ly conceived, a hydrau l i c t u rb ine d r i ve would no t requ i re any flow cont ro l device other than a t r i m o r i f i c e i n the tu rb ine supp:y l i n e t o assure a spec i f i c operat ing speed. A gear d r i v e i s idea l i n the sense t h a t speed cont ro l i s d e f i n i t e . The gear dr ive, however, requi res t h a t the low speed and high speed pump are attached. This r e s t r i c t s the pump packaging and mounting arrangements w i t h the t h r u s t chamber and preburner assemblies.

A1 though s i g n i f i c a n t development problems would no t be an t ic ipa ted w i t h gears operat ing i n RJ-5 (o r RP-1) and possib ly LH2, the use o f gears i n LOX would undoubtedly requ i re comprehensive design studies and experimental development .

c. Engine Weight Considerations

The r e l a t i v e weights o f t hs low speed LOX pump d r i v e system which are based upon comparisons o f p re l im inary sizes t o h i s t o r i c a l data are:

DRIVE METHOD DRIVE WEIGHT, kg ( 1 b l

Gear 13.6 (30) L iqu id Turbine 36.3 (80) Gas Turbine* 544 (1200)

*Drive gas from preburner

The gear d r i v e i s c l e a r l y the l i g h t e s t i n weight, but fo r the LOX syc n a develqment program would be requi red t o va l i da te the concept feas i b i i i t y .

The gas tu rb ine d r i v e i s unacceptably heavy. The weigh^ i s h iqh because o f the very hiqh tu rb ine i n l e t wessures. 340 t o 408 atm (5000-to 6000 p s i ) , and h igh i n l e t temperatureYm867 t o 9 2 2 " ~ (1560 t o 1 660°R).

d. Low Speed Pump Drive Method Select ion

I n order t o maintain f l e x i b i l i t y i n the turbomachinery packag g arrangement, whi le s t i l l achieving a good low speed pump d r i v e e f f i c i i :y, the hydraul ic tu rb ine d r i v e was selected f o r a l l p rope l lan ts . Figure i l l u s t r a t e s the manner i n which f l u i d i s tapped from the h igh pressurt 3ump't.- d r i v e the hydraul i c turb ine.

C. BOOST PUMP NPSH EFFECTS

Low speed hydraul i c t u rb ine dr iven boost pump p re l iminary designs which were generated fo r each of the th ree engine concepts are discussed i n Section V I I I .E. For each o f the boost pumps, analyses were conducted t o determine the effect of varying design p o i n t NPSH on boost pump operat ing character- i s t i c s , weight, and diameters. The fo l low ing ne t p o s i t i v e suctior, head l e v e l s were invest igated:

NPSH, M ( f t )

LOX 0.61 (2) 2.4 (8) 4.9 (16)

LH2 1 0 1 (33) 39.6 (130) 78.6 (258)

RP-1 13.7 (45) 16.2 (53) 19.8 (65)

1. Boost Pump Design Point Condit ions

Design p o i n t suct ion s p e c i f i c speeds were selected t o minimize the 1 i kelyhood o f encountering c a v i t a t i o n induced f low and pressure o s c i l l a - t ions. Magnitudes selected f o r the pre l im inary design e f f o r t were as f o l 1 ows :

Propel 1 ant

Suction Spec i f i c Speed (1

RPM (M', set)' RPM ( G P M ) ' ~ ~ M3/4 ~ t ~ ' ~

RP- 1 5.2 LOX 6.2 LH2 (Dual -Fuel ) 8.9 LH2 (A1 ternate Mode 1 ) 6.0

(1) These u n i t s w i l l be used throughout t h i s sect ion wi thout h r t h e r i d e n t i f i c a t i o n .

Head r i s e requirements o f the boost pumps were determined by specifying high speed pump operat ing suct ion s p e c i f i c speeds o f 4.14 (20,000) along w i t h a1 lowance f o r duct ing pressure losses. This resu i ted i n the fo l low ing magnitudes:

Propel 1 ant Boost Punp Head, AH

M ( F t . ) RP-1 112.8 370 LOX 99.1 326 LH2 (Dual -Fuel ) 61 3 1800 LH2 (A1 ternate Mode 1 ) 324 953

B o o ~ t pump i n l e t diameters were establ ished w i t h i n the c r i t e r i a f o r i n l e t ve loc i t y as a func t ion o f NPSH spec i f ied i n the guidelines, i.e.,

Propel 1 ant

RP- 1

LOX

Max, I n l e t Veloci ty, Cm

Cm

Hydraul ic tu rb ine f low was selected t o be 20 percent o f the low speed pump flow. The head o f the h igh speed pump inducer (and tu rb ine head) was selected comnensurate w i t h hydraul i c tu rb ine e f f i c i ency , 1 i ne losses, and t r i m o r i f i c e requirements. This was estimated t o be approximately e igh t times the low speed pump head.

2. Analysis and Results-

I n conducting the analysis, flow, head r i se , i n l e t f low c o e f f i - c ient , i n l e t h u b t i p r a t i o , and suct ion s p e c i f i c speed were held constant a t the design po in t magnitudes. Pump speed was determined from the suct ion spec i f i c speed re la t i onsh ip as a func t ion o f NPSH. The i n l e t diameter was then calculated f o r the design po in t i n l e t f low c o e f f i c i e n t and hub-t ip r a t i o . The impel ler discharge t i p speed was calculated as a func t ion o f head r i s e by es tab l ish ing the head c o e f f i c i e n t as a func t ion o f stage spec i f i c speed (Figure 56). The mean impel ler o u t l e t diameter was then calculated from the t i p speedlpump speed re la t ionsh ip . For the constant f low, constant head boost pump condit ions, weight was estimated t o vary w i t h the cube o f i n l e t diameter.

Figures 57, 58, and 59 show the e f f e c t s o f NPSH on the low speed turbopump weights and f igures 60 and 61 show the e f f e c t s on i n l e t and o u t l e t diameters. The NPSH range f o r each pump corresponds t o dn i n l e t pressure rang!? s f approximately 1.07 atm (15.7 ps ia ) t o 1.54 atm (22.7 psia). As indicated i n the f igures, s i g n i f i c a n t increases i n LOX and LH2 boost pump weight and impel ler diameters occur as NPSH i s reduced. Because o f the lower vapor pressure o f RP-1, the NPSH i s r e l a t i v e l y high even a t the lo,-er i n l e t pressures and the change i n RP-1 pump weight and diameter i s much small er.

D. ENGINE DESIGN AND OPERATING SPECIFICATIONS

This sect ion describes the engine system charac ter is t i cs o f a l l three engine concepts selected f o r p re l i m i nary design e f f o r t s .

HEAD - 77.4 m (370 FT)

SUCTION SPECIFIC SPEED - 5.2 (25,000)

16 18 20 2 2 NPSH - m

Figure 57. E f f e c t o f NPSH on Low Speed RP-1 Turbopump Weight

NPSH - FT 0 4 8 12 16 2 0

FLOW - 0.554 m3/sec (8785 GPM)

HEAD - 99.1 m (326 FT)

SUCTION SPECIFIC SPEED - 6.2 (30,000)

NPSH - m

Figure 58. Effect of NPSH on Low Speed LOX Turbopump Weight

NPSH - FT

NPSH - FT 0 1 00 200 - 300

I I I I

ALTERNATE MODE 1 LOW SPEED V)

LH2 TPA (H2/02 GG CYCLE) - -50 2 0 2 0 ., Y FLOW - .I29 m3/sec (2040 GPM) I

I .-40

5 HEAD - 324 m (950 FT) g 2 W

SUCTION SPECIFIC SPEED - 6.0 (29,000) -- 30 S x n

n 10.- aE .-20 2 3

0 n m

Y s P:

-- 10 3 I-

I 1 I I I I I I , 0 0 20 40 60 80

160

cn 120

Y

I

k z 80.-

n J n 0 m & 3

40-

0

Figure 59. E f f e c t of NPSH on Low Speed LH2 Turbopump Weight

I I I I 1

.- DUAL-FUEL LOW SPEED LH2 TPA

FLOW - 0.898 m3/sec (14240 GPM)

HEAD - 613 m (1800 FT) -- SUCTION SPECIFIC SPEED - 8.9 (43,000)

I I I I I I I I 20 40 66 80

NPSH - m

NPSH - FT 0 4 8 12 16 2 0

LOW SPEED LOX PUMP

\ FLOW-. 554m3/sec (8785 GPH)

HEAD - 99.1 m (326 FT) 5 0

SUCTION SPECIFIC SPEED 6.2 (30,000)

\ MEAN DIAMETER I

INLET

1 2 3 4 5 6

NPSH - m NPSH - FT

2o

l6

s. 0

I 12 w W I- W

3 8 . . a

4

14 15 16 17 18 19 2 0

NPSH - in Figure 60. E f f e c t o f NPSH on Low Speed LOX & RP-1 Pump Impel ler Diameters

176

- -

L--- INLET -. MEAN DIAMETER

.- OUTLET

LOW SPEED RP-1 PUMP

FLOW - .272 m3/sec (4310 GPM) HEAD - 112.8 m (370 FT) --

., SUCTION SPECIFIC SPEED - 5.2 (25,000) ..

I 1 1 I I I !.

1: 7

6 I U r

- 5 - m P:

- . 4 F r - 3 "0

2

1

0

NPSH - FT

100 290 3CO /

I I I I 1 0

.

-

I I I ! I I 1 I - 0 0 20 40 60 80

NPSH - m

\ DUAL - FUEL LOW SPEED LH2 PUMP

FLOW - 0.898 m3/sec (14,240 GPM)

\ HEAD - 613m (1800 FT)

\ SUCTIOY SPECIFIC SPEED - 8.9 (43,000)

MEAN DIAMETER

L TIP DIAMETER INLET

NPSH - FT 0 100 200 390

--6 V, W x 0

m.4 E I

E C

20 .- 16

5 '2 I

w W

8

I I I I 1

.. TIP DIAMETER

INLET .- --

ALTERNATE MOrE 1 LOW SPEED

u n

0 - LH2 PUMP (H2/02 GG CYCLE) MEAN DIAMETER "2 3 Y

OUTLET L I , FLOW - . I29 m3 /sec (2040 GPM) " HEAD - 324 m (950 FT)

SUCTION SPECIFIC SPEED - 6.0 (29,000) I 1 I I I I ! I

60 60 0 2 0 40 NPSH - m

1. Mode 1 LOX/RP-1 Base1 i n e Design

a. Conf igura t ion and Nominal Operating Cond

The engine cyc le f o r the Mode 1 Base1 i n e engine design i s a dual preburner staged-combustion cyc le w i t h a LOX cooled t h r u s t chamber. The engine schematic i s shown on F igure 62 i n c l u d i n g the engine con t ro l requirements which b a s i c a l l y cons i s t d f main f u e l ~ n d o x i d i z e r valves. Cycle balance i s obtained by o r i f i c e s i n the f u e l - r i c h preburner propel- l a n t l i nes .

The p r e l i m i r a r y assembi'y drawing o f t he basel ine engine i s shown on Figures 63 and 64. The main fea tu re i s t h e sidemounted TPA's and boost pump assembl i e s Hot gas duc t ing i s minimized through com- p l e t e l y in tegra ted TPA's which a re mounted t o t he engine by means o f t he ho t gas cross-over ducts. The TCA i s t o t a l l y regenera t i ve ly cooled w i t h LOX. The TCA design incorporates a s l o t t e d zirconium copper chamber t o a nozzle area r a t i o ot 15: l . A two pass Inconel 718 tube bundle i s used from heye t o the 40:1 nozzle e x i t area r a t i o . The coolant i s i n t r o - duced a t an area r a t i o o f l . 5 : l and f lows up t o the plane o f the i n j e c t o r face. The coolant i s co1 l ec ted and in t roduced a t the area r a t i o of 1.5 t o cool the nozzle. The coolant e x i t s from t h ? tube bundle a t an area r a t i o o f 15: l and i s then introduced i n t o t he preburners.

The design d e t a i l s f o r t he principal components a re presented i n appropr ia te paragraphs o f t h i s sect ion.

The gimbal 1 ed envelope was evaluated f o r a 1 O0 s u a r e pa t t e rn and i s a l so shown on F igure 63. The r e s u l t s i nd i ca ted a 240.2 cm (94.6 i n . ) square requirement.

The Mode 1 base1 i n e engine nominal opera t ing spec i f i c - cat ions are shown on Table LVI and an engine pressurk schedule f o r nominal operat ion i s shown on Table LVI I .

b. Engine Operation and Control

Each o f the engitze cyc les were examined t o e s t a b l i s h the main con t ro l requirements. P a r t i c u l a r a t t e n t i o n was g iven t o the s t a r t and shutdown sequences. The LOXIRP-1 modes o f operat ion a re pat terned a f t e r F-1 (Ref. 70) and T i t a n I (Ref. 71 ) engine experience. The pr imary concern w i t h these p rope l lan ts i s t o keep contaminants ou t o f the LOX mani- f o l ds . Therefore, a l l LOX/RP-1 combustors a re s t a r t c d o x i d i z e r - r i c h t o reduce the chance o f RP-1 en te r ing the o x i d i z e r c i r c u i t s .

Fig

ure

62.

B

asel

ine

LOX/

RP-1

M

ode

1 E

r~g

ine C

ycle

Sch

emat

ic

TABLE LVI. - MODE 1 LOX/RP-? ENGINE OPERATING SPECIFICATIONS

Nominal MR = 2.9

Enqi ne

Sea-Level Thrust, MN ( l b ) Vacuum Thrust, MN (1 b)

Sea-Level Speci f i c Impul se , sec

Vacu~ m Speci f i c Impul se, sec Total Flow Rate, kg/sec (1 blsec)

Mixture Ratio Oxidizer Flow Rate, kglsec (1 blsec)

Fuel Flow Rate, kglsec (1 blsec)

Thrust Chamber

Sea-Level Thrust, MN (1 b) 2.70

Vacuum Thrust , MN (1 b) 2.926

Sea-Level Specif ic Impulse, sec 323.5

Vacuum Specific Impulse, sec 350.6

Chamber pressure, atm (psia) 272.2 Nozzle Area Ratio 40

Mixture Ratio 2.9 Throat Diameter, cm (in.) 26.31 4

Nozzle Ex i t Diameter, cm (in.) 166.421

Coolant Jacket Flow Rate (LOX), kg/sec ( lb lsec) 632.7

Coolant Jacket AP, atm (psi) 126.9

Coolant I n l e t Temp., O K (OR) 111.1

Coolant Ex i t Temp., O K (OR) 202.8

In jec tor Ox.-Rich Gas Flow Rate, kg/sec (1 blsec) In jec tor Fuel-Rich Gas Flow Rate, kglsec (1 b/sec)

Chamber Length, an ( i r; Chamber Diameter, cm (in.)

TABLE LVI (cant . ) Preburners

Chamber Pre ssure, atm (psi a Combustion Temperature, O K ( OR) Mixture Ratio Oxidizer Flow Rate, kg/sec (1 b/sec) Fuel Flow Rate, kg/sec (1 b/sec)

Turbi nes

I n l e t Pressure, atm (psia) I n l e t Temperature, O K (OR)

Total Gas Flow Rate, kg/sec (Wsec ) Gas Properties

C , speci f ic heat @ constant &ssure, d/kg O K (Btu/ l b OR)

. y , Ratio o f specific heats Shaft Horsepower mhp (HP) Efficiency, X

Pressure Ratio ( to ta l to s ta t i c )

Main Pumps

Total Outlet Flow Rate, kg/sec (1 b/sec) Volumetric Flow ate, n3/sec (gPd NPSH, m (ft.)

LOX/RP-1 RP-1/LOX Ox.-Rich Fuel -Rich

440.8 (6,479) 439 (6,452) 922 (1,660) 867 (1,560) 45.0 0.22 587.5 (1,295.3) 45.1 (99.5)

LOX PW RP-1 PUW

440.8 (6,479) 439 (6,452) 922 (1,660) 867 (1,560)

600.6 (1,324.1) 250.3 (551.7)

LOX RP-1

632.7 (1,394.8) 218.2 (481 .O)

TABLE LVI (cont. )

Main Pumps (cont.) LOX RP-1

Discharge Pressure, atm ( psia) 632.8

Nwnber o f Stages 2

Speci f i c Speed, (RPM) (m3/sec)l/Z/ O. j 1 (m)3/4 Total Head Rise, m ( f t ) 5660

Efficiency, % 82

Low Speed TPA

Pumps

NPSH, m (ft) 4.88

I n l e t Flow Rate, kg/sec 632.7 (1 b/sec)

Outlet Flow Rate, kg/sec 759.2 (1 b/sec)

Discharge Pressure, atm (psia) 12.4

Hydraul i c Turbine

I n l e t Pressure, atm (psia) 92.9

Outlet Pressure, atm (psia) 12.4 Flow Rate, kglse: (lb/sec) 126.6

Speed, rpm 2560

Pre

burn

er

Pro

pel l

an

t

Nom

inal

MR

= 2.

9

OX-

Ri c

h

LOX

RP-

1

Fue

l -R

ich

Pre

ssur

e,

atm

(p

sia

)

Mai

n Pu

mp

Dis

char

ge

ki

n S

hu

toff

Val

ve I

nle

t AP

Sh

uto

ff V

alve

V

alve

Ou

tle

t AP

Lin

e

Coo

lant

Jac

ket

Inle

t AP

Coo

lant

Jac

ket

Coo

lant

Jac

ket

Ou

tle

t AP

Lin

e

Pre

burn

er C

ontr

ol I

nle

t AP

Con

trol

Pre

burn

er I

n1

et

493.

4 (7

,252

) 54

2.9

(7,9

79)

487.

8 (7

,169

) 51

6.5

(7,5

91)

+ AP

Pre

burn

er

52 -6

(7

73)

102.

1 (1

,500

) 48

.9

(717

) 77

.5

(1,1

39)

t

Tu

rbi n

e In

1 e

t 44

0.8

(6,4

79)

439

(6,4

52)

4

AP

Tur

bine

(T

ota

l to

Sta

tic)

14

5 14

3.2

(2,1

04)

+

(2,1

31)

Mai

n In

jec

tor

Inle

t (T

ota

l)

295.

8

(4,3

48)

' 29

5.8

(4,3

48)

hP

Inje

cto

r 23

.7

(348

) 23

.7

(348

) Ch

alRbe

r P

ress

ure

272.

2 (4

,000

) 27

2.2

(4,0

00)

The engine s t a r t and shutdown sequence i s presented on Table LVII I . The sequence of operations l i s t e d are assumed t o occur a f t e r the vehicle prevalves have been opened and the cryogenic components have been c h i l l e d down t o the main engine shutof f valves.

c. Star t and Shutdown Data

The engine s t a r t and shutdown transients were also estimated. The staged combustion cycle engines were assumed t o be c h i l l e d down and bled-in t o the main chamber valves p r i o r t o receipt o f the f i r e signal. These engines were then assumed t o be star ted under tank head. The t ransient estimates are based upon the analyt ical model ing o f s im i la r engine configurations on the ALRC Liquid Engine Transient Simulation (LETS) model.

The s t a r t and shutdown transient data i s sumnarized '

on Table LVIX. Star t t o 90% o f rated thrust and shutdown down t o 5% o f rated thrust are generally specif ied values t o establ ish t ransient times.

d. Design and Off-Design Engine Performance

The design and off-design engine performance a t the design thrust level f o r + 10% MR excursions are presented on Table LX. The nominal engine operatingmixture r a t i o i s 2.9. A requirement f o r the engine t o operate over the en t i re mixture r a t i o range i n f l i g h t was not i den t i f i ed during the course o f t h i s study. Therefore, f o r t h i s analysis, i t was assumed tha t the preburner flow o r i f i c e sizes are changed t o operate a t the various mixture ra t ios . Continuous i n f l i gh t operation over the en t i re MR range would resu l t i n higher engine pressure drop requirements.

e. Engine Mass Properties Data

The engine mass properties data, consist ing o f the engine weight breakdown, gimbal 1 ed moments o f i n e r t i a and center o f grav i ty 1 ocation , were computed from the prel i m i nary engine and component layout drawings.

The weight breakdown f o r the base1 ine engine i s shown on Tab1 e LX1.

The center o f grav i ty i n the axia l d i rec t ion was calculated t o be 62 cm (24.4 inches) from head-end o f the gimbal.

TABLE LVI I I. - SEQUENCE OF OPERATIONS BASELINE LOX/RP-1

Start - 1 . Purge Oxidizer Lines and Manifolds i n Fuel and Ox.

Rich Preburners . 2. Energf ze Spark Igni ters . 3. Open Main Ox. Valve (#I).* 4. Open Igni ter Valves on Fuel and Ox. Rich

Preburners . 5. Open Main (RP-1) Fuel Valve ( 1 2 ) .

Shutdown

1. Close MainOx. Valve (11). 2. I n i t i a t e Ox. Purge. 3. Close Main (RP-1) Fuel Valve (12). 4. Close Ign i ter Valves on Fuel and Ox. Rich

Preburners. 5. Cutoff Ign i ter Spark Energy.

*Note: Numbers refer t o valves on Figure 62.

TABLE LIX. - MODE 1 LOX/RP-1 BASELINE START AND SHUTDOWN TRANSIENT DATA SUMMARY

Star t t o 90% F

Time, sec 2.52

Total S ta r t Impulse, kg-sec (1 b-sec) 69,000 (1 52,000)

LOX Consumption, kg (1 b) 129.3 (285)

RP-1 Consumption, kg ( l b ) 120.2 (265)

Shutdown t o 5% F

Time, sec 0.50 Total Shutdown Impulse, kg-sec 80,800 (1 78,200) (1 b-sec) LOX Consumption, kg (1 b) 162.4 ( 358)

RP-1 Consumption, kg (1 b) 73.0 (161

TABL

E LX

. -

Ble

Sea

-Lev

el

Thr

ust,

MN

(lb

) V

acum

Thr

ust,

MN

(I

b)

WDE

1

LOX/

RP-

1 BA

SELI

NE

ENG

INE

DESI

GN

AND

OFF

-DES

IGN

MR

PE

RFOR

MAN

CE

; M

ixtu

re R

ati

o

t

2.61

2.

90

3.19

1

Sea

-Lev

el

Sp

eci

fic

Impu

lse,

se

c 32

1.7

323.

6 32

1.1

Vac

uun

Sp

eci

fic

Impu

lse,

se

c 34

9.1

350.

6 34

7.7

To

tal

Flo

w R

ate,

kg

lsec

(lb

lse

c)

855.

9 (1

886.

9)

850.

9 (1

875.

8)

875.

5 (1

890.

4)

237.

1 (5

22.7

) 21

8.2

(481

.0)

204.

7 (4

51.2

) 1

Fue

l Flow R

ate,

kg

/sec

(l

bls

ec)

61

8.8

(136

4.2)

63

2.7

(139

4.8)

65

2.8

(143

9.2)

O

xid

ize

r F

low

Rat

e,

kg/s

ec

(lb

/se

c)

C

277.

5 (4

,078

) 27

2.2

(4,0

00)

270.

3 (3

,973

) 4

Cha

mbe

r P

ress

ure,

at

m (

psi

a)

7

TABL

E LX

I. - W

DE

1 LO

X/R

P-1

BASE

LIN

E EN

GIN

E W

EIGHT

STA

TEM

ENT

Com

pone

nt

Gia

bal

Mai

n In

jec

tor

Cop

per

Cha

Rde

r an

d N

ozzl

e (E

=

14.7

) Tu

be B

undl

e N

ozzl

e (E

=

14.7

to

40)

F

uel

Ric

h P

rebu

rner

O

xid

ize

r R

ich

Pre

bu

mr

Fue

l V

alve

s an

d A

ctu

atio

n

Oxi

diz

er

Val

ves

and

Act

ua

tion

La

w Sp

eed

LOX

TPA

Law

Spe

ed R

P-1

TPA

Hig

h S

peed

LOX

TP

A H

igh

Spee

d R

P-1

TPA

Hot

6as

Ma

nifo

ld

Low

Pm

ssur

e Li

nes

Hig

h P

ress

ure

Line

s Ig

nit

ion

Sys

tem

W

sce

l 1 an

eous

TO

TAL

The center o f gravity and gimballed moments o f i ne r t i a fo r the engine are sumr i zed below.

Axis - X Y Z

Center o f Gravity

X - X Y - Y z - z 2 2 2 Gimballed Iner t ia kg-m2 (slug-ft2) kg-m2 (slug-ft ) kg-m (slug-ft )

The coordinate system i s defined on Figures 63 and 64.

2, Dual-Fuel Engine Design,

a. Configuration and Nominal Operating Conditions

This engine concept i s peculiar t o mixed mode propulsion systems since the Mode 1 and Mode 2 propellants u t i l i r e a comnon thrust chamber and LOX feed system. A schematic (without boost pumps) i s shown on Figure 65. The engine uses a LOX cooled thrust chamber, gaseous propellant main in ject ion and dual preburners i n both operating modes. The LOX TPA has two preburners; one f o r each mode o f operation. Mode 1 operation u t i l i z e s a nozzle area r a t i o o f 40:l and Mode 2 a 200:l retractable nozzle extension.

The TCA design uses a slotted zirconium copper chamber to a nozzle area r a t i o o f 15:l. An Inconel 718, two pass, tube bundle i s used from 15:1 t o an area r a t i o o f 40:l. The coolant flow path i s the same as the base1 ine design. The nozzle extension i s regeneratfvely cooled with LOX from the 40:1 point t o an area r a t i o o f 132:l and raJia- t!on cooled from here t o the exit. Columbium i s used as the material throughout the radiation cooled section.

The schematic ident i f ies the control valves required for engine operation i n both Mode 1 and Mode 2. This engine concept i s exclusively used fo r a series burn application and the Mode 2 feed system has to be isolated during Mode 1 operation. This affected the number and location o f the valves.

OXYG

EN

FUEL

(R

P-1)

= HYD

ROGE

N =

COneU

STIrn

PRO

bUer

S

Figu

re 6

5.

Dua

l -Fu

el , Ox

ygen C

oole

d E

ngin

e C

ycle

Sch

emat

ic

O O

rifi

ce

X V

alve

Q ve

nt

The preliminary layout o f the dual - fuel engine i s shown on Figures 66 and 67. The engine featurcs f ixed boost pumps f o r a1 1 three propellants clustered around the engine gimbal center, The TPA's are side mounted wi th t he i r axis perpendicular t o the thrust axis i n order t o obtain favorable center o f gravi ty location. The preliminary layout indicates the s p l i t and stored length o f the c dendible nozzle.

The engine envelope data i s sumnari zed below.

Engine Length, cm ( i n . ) Fixed Nozzle Extendible Nozzle Retracted Extend1 b l e Nozzle Deployed

Nozzle E x i t Diameter, cm (in.)

The gimbal led envelope was evaluated f o r a 10' square pattern i n Mode 1 and a 7' square pattern I n Mode 2. The resu l t - i ng square dfmensions are:

ENVELOPE GIMBAL ANGLE CM (in.)

Fixed Nozzle 10" Retracted Nozzle l o 0 Deployed Nozz 1 e 7e

The engine operating conditions f o r the engine nominal design point are presented i n Table LXII f o r both the Mode 1 and Mode 2 engine operation.

The engine pressure schedules f o r Mode 1 and Mode 2 nominal engine operation are presented i n Table LX I I i and Table LXIV.


The study of the dual -fuel engine r e s d ted !n the iden t i f i ca t ion o f addit ional controls f o r not only operational purposes but t o p roh ib i t the flow o f hot gases through inact ive components. For example, when the ovgen-r ich preburner i s operating wi th LOXIRP-1, valves 7 and 9 (Figure 65) must be closed t o keep the LOXIRP-1 preburner exhaust products from backing through the LOX and LH2 propel lant feed system. The conbusti on products o f both fue l -ri ch preburners (RP-1 /LOX and H2/02) are exhausted i n t o a comnon main i n j ec to r manifold. Therefore, there i s an open flow path throu the inact ive system when the other i s operating, Closing o f f t h i s flow w f t h preburner va',ves does not solve the problem

I H

IGH

SPE

ED

I A

H,

TPA

' H

IGH

SPE

ED

/RP-

I TP

A

Fig

ure

66.

D

ual-F

uel

Eng

ine

Ass

embl

y (T

op V

iew

)

nx. D

IA.

--392 m

/

(154

.5

in.)

/ -

- - -

(121

.00

in.)

I

Fi g

ure

67.

Dua

l -Fu

el

Eng

ine

Ass

embl

y (S

ide

Vie

w)

Engine

TABLE LXII. - WAL-FUEL NOMINAL ENGINE OPERATING SPECIFICATIONS

Sea-Level Thrust, MN (1 b) Vacuum Thrust, MN (1 b) Sea-Level Specif ic Impul se , sec Vacuum Specif ic Impul se, sec Total Flow Rate, kg/sec (1 b/sec) Mixture Ratio Oxidizer Flow Rate, kglsec (1 b/sec) Fuel Flow Rate, kg/sec (1 b/sec)

Thrust Chamber

Sea-Level Thrust, MN ( l b ) Vacuum Thrust, hN (1 b) Sea-Level Specif ic Impul se , sec

Vacuum Specific Impul se , sec Chamber Pressure, atm (psia) Nozzle Area Ratio Mixture Ratio Throat Diameter, cm (in.) Nozzle Ex i t Diameter, cm (in.) Coolant Jacket Flow Rate (LOX), kg/sec (1 b/sec) Coolant Jacket AP, atm (ps i ) Coolant I n l e t Temp., O K ( O R )

Coolant Ex i t Temp., O K ( O R )

In jec tor Ox.-Rich Gas Flow Rate, kg/sec (1 b/sec)

LOX/RP-1 Mode 1

LOX/LH2 Mode 2

a*7% o f the LOX flow bypasses the coolant jacket i n Mode 1 operation and suppl ies the fuel - r i ch preburner.

TABLE LXII (cont .)

LOX/RP-1 LOX/LH2 Thrust Chamber (cont . ) Mode 1 Mode 2

In jector Fuel-Rich Gas Flow Rate, 250.8 (552.9) 114.2 (251.8) kg/sec (1 blsec) Chamber Length, cm (in.) Chamber Diameter, cm (in.)

Preburners

Chamber Pressure, atm (psia) Combustion Temperature, OK ( O R )

Mixture Ratio

Oxidizer Flow Rate, kg/sec (Wsec) Fuel Flow Rate, kglsec (1 blsec)

Turbines

I n l e t Pressure, atm (psia)

I n l e t Temperature, OK (OR)

Total Gas Flow Rate, kg/sec (1 blsed)

Gas Properties Cp, Specific Heat @ Constant Pressure, J/ kg OK (Btu l l b OR)

y, Ratio o f Specific Heats

LOX/RP-1 RP-1 /LOX Ox. -Rich Fuel -Rich

440.8 (6,479) (6 971 4)

922 (1,660) 867 (1,560) 45.0 0.22

588.8 (1,298.1) 45.2 (99.7)a

LOX Pump RP-1 Pump

Shaft Horsepower, mHp (HP) 59,826 Efficiency, % 80 Speed, rpm 15,100 Pressure Ratio (Total t o 1 ,490 Stat ic)

'coolant jacket bypass flow.

TABLE L X I I (cont. )

Main pun^ -- Total Out let Flow Rate, kglsec (IbJsec)

3 Volumetric Flow Rate, m /sec (gpm) NPSH, m ( f t ) S u c t i m Spec , f i c S ed, (RPM)(d/s)1/ 2/(m)F4 Speed, rpm Discharge Pressure, atm (psia)

Number o f Stages

Speci f ic Speed, ( R P M ) ( ~ ~ / s ) ~ / ~ I (m) 314 Total Head Rise, m (ft)

Efficient), %

LOW SPEED TPJ

Pumps NPSH, m (ft) I n l e t Flow Rate, kglsec (1 blsec)

Out le t Flow Rate, kglsec ( I blsec) Discharge Pressure, atm (psia)

Hydraul i c Turb I lie I n l e t Pressure, atm (psia) Outlct. Pressure, atm (ps.ia) Flow Rate, kglsec ( l t l s e c )

Speed, rpm

LOX RP- 1

634.0 (1,397.8) 218.6 (482.0)

LOX RP-1

i .J

j I t . j

U I I I I i L 1 / 1 ! -

i..l i 1 iJ L.

[]

[ j

ill n n n

Preburners

TABLE LXII (cont. )

Dual -Fuel , Mode 2 Nominal

LOW LH2 OX.-Ric Fuel -Rich

Chamber Pressure, atm (psia) 301 .1 (4,425) 31 5.2 (4,632) Combustion Temperature, OK ( O R ) 922 (1,660) 922 (1,660) Mixture Ratio 110 0.9 Oxidizer Flow Rate, kg/sec (Wsec ) Fuel Flow Rate, kg/sec (1 b/s=)

Turbines LOX Pump LH2 Punp

I n l e t Pressure, atm (psia) 301 .1 (4,425) 31 5.2 (4,632) I n l e t Temperature, O K ( O R ) 922 (1,660) 922 (1,660) Total Gas Flow Rate, kglsec (1 b/sec) 394.8 (870.3) 114.2 (251.8)

Gas Properties Cp, Specif ic Heat 8 Constant Pressure, J/kg OK (Btu/ 1159 (0.277) 8242 (1.97) l b OR)

y, Ratio o f Specif ic Heats 1.312 1 .357 Shaft Horsepower, mHp (HP) 32,347 (31,900) 58,305 (57,500)

Efficiency, % 77 81

Speed, rpn 12,600 (12,600) 33,200 (33,200) Pressure Ratio (Total t o Sta t ic ) 1.37 1.27

Main pumps LOX LH2

Total Outlet Flow Rate, 445.3 (981.8) 63.6 (140.3) kg/sec (1 blsec) Volumetric Flow Rate, m3/sec (gpn)

0.392 (6,210) 0.903 (14,310)

NPSH, m (ft) 67.4 (221) 518.2 (1,700) Suction Specif ic Speed, [RPM) (m31sec) 1 /2/ (m) 3/4

TABLE LXII (cont. )

Main Pumps (cont .) LOX L H ~

Speed, rpm 12,600 33,200 Discharge Pressure, atm ( psia) 469.5 (6,900) 381 .O (5,600)

Number o f Stages 2 3

Specific Speed, (RPM) (m3/sec)ll2/ 0.269 (1 ,300)a 0.21 7 (1,050)a (m) 314 Total Head Rise, m (ft) 4197.1 (13,770) 54,102 (177,500)

79 80 Efficiency, %

Low Speed TPA

Pumps NPSH, m ( f t )

I n l e t Flow Rate, kg/sec (1 blsec)

Outlet Flow Rate, kg/sec (1 blsec)

LOX LH2

Discharge Pressure, atm ( psia) 9.19

Hydraul i c Turbine I n l e t Pressure, atm (psia) 65.73

Outlet Pressure, atm (psia) 9.19 Flow Rate, kglsec ( lb lsec) 89.1

Speed, rpn 21 50

TABL

E L

XII

I . - Y

AL-

FUE

L,

MODE

1 N

OM

INAL

ENG

INE

PRES

SURE

SCH

EDUL

E

Pre

burn

er

Pro

pel 1

an

t P

ress

ure,

at

m (

psi

a)

Mai

n Pu

mp

Dis

char

ge

Mai

n S

hu

toff

Val

ve I

nle

t AP

Shu

tcf f V

alve

Val

ve O

utl

et

AP

Lin

e

Coo

lant

Jac

ket

Inle

t AP

C

oola

nt d

acke

t C

oola

nt J

acke

t O

utl

et

aP L

ine

Pre

burn

er C

on

tro

l In

let

AP C

ontr

ol

Pre

burn

er I

n1

et

aP

Pre

burn

er

Tur

bine

In

1 e

t AP

Tur

bine

(T

ota

l to

Sta

tic)

C

heck

Val

ve I

nle

t AP

Che

ck V

alve

M

ain

Inje

cto

r In

let

aP

Inje

cto

r n

~

Ch-r

Pre

ssur

e 0,

(1)

TCA

By-

Pas

s Fl

ow.

Ox-

Ric

h LO

X R

P-1

632.

8 (9

,300

) 55

1.1

(8,1

00)

632.

8 (9

,300

55

1.1

(8,1

00)

6.33

(9

3)

5.51

(8

1)

626.

4 (9

,207

) 54

5.6

(8,0

19)

2.72

(4

0)

2.72

(4

0)

623.

7 (9

,167

) --

132.

0 (1

,940

) - -

491.

7 (7

,227

) - -

2.72

(4

0)

-- 48

9.0

(7,1

87)

542.

9 (7

,979

)

4.90

(7

2)

5.44

(8

0)

484.

1 (7

,115

) 53

7.4

(7,8

99)

43.3

(6

36)

96.6

(1

,420

) 44

0.8

(6,4

79)

145.

0 (2

,131

)

Fue

l -Ri c

h L

OX

(~

) R

P-1

A n h

8 8 _ Z % % f I t iz U!G sc . P

m V ) V ) m * V) m m QD Y W Y Y Y Y Y Y Y

- - I . . .

* m o m h h (3 0

nn A nn-

C, Q) r c W

E , 4 r C I r) > V) .r '4- 0 'I-

3 % 2 n V)

o f leakage through the turbopump shaf t and up the suction l i ne , Turbopump shaft seals and vents o r hot gas check valves and vents are required. The hot gas check valve has been p re l im inar i l y selected as the most reasonable approach.

The LOX/RP-1 combustors are again sequenced t o s t a r t oxi d i ze r - r i ch . The LOX/LH2 combus t o r sequences are based upon t es t experience wi th both fue l and ox id i zer-r ich modes o f operation, The fue l - r i c h LOX/LH2 conbus tors are star ted fue l -ri ch and the oxi d i zer-r i ch combustors are star ted oxidizer-rich. Experience has shown tha t i t i s possible t o s t a r t fuel -ri ch LOX/LH2 combustors oxid izer-r ich but attenpts , both planned and inadvertent, t o s t a r t ox id i zer - r i ch conWctors fuel - r i ch resulted i n destruction o f the hardware (Ref. 72 and 73).

The cryogenic components, both LOX and LH2, f o r t h i s dual-fuel engine are assumed t o be ch i l l ed down on the launch pad.

The s t a r t and shutdwn sequence for the dual-fuel engine i s shown i n Table LXV.

c. S ta r t and Shutdown Oiita

The dual-fuel engine was also assumed t o be ch i l l ed down and bled-in t o the main chamber valves p r i o r t o rece ip t o f the f i r e signal and star ted under tank head. The t ransient estimates are based upon the analyt ical modeling o f s im i la r engine configurations such as, the ARES and ALRC SSME design .

The s t a r t and shutdown data i s sumnarized on Table LXVI .

d . Design and O f f -Desi gn Engine Performance

Engine cycle balances and performance were evaluated a t the design thrust leve l over a mixture r a t i o range encompassing the design po in t mixutre r a t i o ? 10%. These data are sumnarized on Tables LXVII and LXVIII f o r Mode 1 and 2 operation, respectively,


The engine mass properties data were computed from the preliminary engine and component 1 ayout drawings .

The weight breakdwn f o r the dual-fuel engine i s shown on Table LXIX.

The center o f gravi ty i n the axi a1 d l r e c t i on was calculated t o be 199 cm (78.2 in,) from the head-end o f the gimbal w i t h the nozzle extensi on deployed.

TABLE LXV. - SEQUENCE OF OPERATION DUAL-FUEL ENGINE

Mode 1 - Star t -

1 . Purge Oxidizer Lines and ~ a n i fo ld, 2. Energize Spark Ign i ters . 3. Open Main Ox. Valve (dl).* 4. Open I gn i t e r Valves on Fuel and Ox. Rich Preburners. 5. Open Ox, Valves on Fuel and Ox. Rich Preburners (#2 and Y3). 6. Open Main Fuel (RP-1) Valve (W4). 7. Open Ox. Rich Preburner Fuel (RP-1) Valve (#5).

Shutdown

1. Close Ox. Rich Preburner Fuel (RP-1) Valve (#5).

2. Close Main Ox. Valve (d l ) , 3. I n i t i a t e Ox. Purge. 4. Close Main Fuel (RP-1) Valve (14). 5. Close I gn i t e r Valves on Fuel and Ox. Rich Preburners . 6. Cutoff I gn i t e r Spark Energy. 7. Close Fuel and Ox. Rich Preburner Ox. Valves (#2 and 13).

Mode 2

Star t - 1 . Energize Spark Ign i ters . 2. Open I gn i t e r Valves on Fuel and Ox. Rich Preburners. 3. Open Main Fuel Val ve (16) .* 4. Open Main Ox. Valve (# I ) . 5. Open Ox. Val ves on Fuel and Ox. Rich Preburners (#7 and #8). 6. Open Ox. Rich Preburner Fuel -Valve (#9).

*Numbers re fe r to valves on Figure 65.

TABLE LXV (cont.)

Mode 2 (cont.)

Shutdown

1 . Close Igniter Valves. 2. Cutoff Igniter Spark Energy. 3. Close Ox. Rich Preburner Fuel Valve (19). 4. Close MainOx. Valve # I . 5. In i t ia te Ox. Purge. 6. Close Main Fuel Valve. 7. Close Ox. Valves on Fuel and Ox. Rich Preburners (17 and 18).

TABLE LXVI. - DUAL-FUEL ENGINE START AND SHUTDOWN TRANS I ENT DATA SUMMARY

Made 1 Mode 2 Start t o 90% F

Tlme, sec 2.52 2.52 Total Start Impulse, kg-sec 69.000 (1 52.000) 58,500 (1 29,000) (1 b-sec) LH2 Consumption, kg (1 b) - - 35.4 (78) LOX Consumption, kg ( lb ) 129.3 (285) 93.0 (205) RP-1 Consumption, kg ( lb) 120.2 (265) - -

Shutdown t o 5% F Time, sec 0.50 0.50 Total Shutdown Impulse , 80,800 (178,200) 63,400 (139,700) kg-sec (1 b-sec) LH2 Consumption, kg (1 b) - - 23.1 (51 LOX Consunptlon, kg (1 b) 162.9 (358) 113.9 (251) RP-1 Consunption, kg ( lb) 73.0 (616) o -

TABL

E L

XV

II.

- DU

AL-F

UEL,

W

OE

1 DE

SIG

N AN

D O

FF-D

ESIG

N MR

PER

FORM

ANCE

Eng

ine

Sea

-Lev

el

Thru

st, 14(

(1 b)

Vacu

un T

hrus

t,

HI4

(lb

) S

ea-L

evel

S

ped

f i c

Impu

l se ,

sec

Vacu

un S

pec

ific

Im

puls

e, s

ec

To

tal

Flow

Rat

e,

kg/s

ec

(lb

/sec

) Fu

el F

low

Rat

e,

kg/s

ec

(lb

/sec

) O

xid

izer

Flo

w R

ate,

kg

/sec

(

Ws

d

Cham

ber

Pre

ssur

e,

atm

(p

sia)

Mix

ture

Rat

io

2.61

2.

90

3.19

TAB

'-€ L

XV

III .

- D

JAL-

FUEL

, MO

DE 2

DES

IGN

AND

OFF

-DES

IGN

MR

PER

FORM

ANCE

E ' 20

0

Va

cw

Thr

ust,

MN

(1

b)

2.29

(5

15,2

50)

2.29

(5

1 5,2

50)

2.29

(5

15,2

50)

Vac

uun

Sp

eci

fic

Impu

l se ,

sec

461.

6 45

9.2

455.

4 T

ota

l F

low

Rat

e,

kgls

ec

(lb

/se

c)

506.

3 (1

116.

2)

508.

9 (1

122.

1)

513.

2 (1

131.

4)

Fue

l F

low

Rat

e,

kgls

ec

(1 b/

sec)

69

.4

(152

.9)

63.6

(1

40.3

) 59

.0

(130

.0)

Oxi

diz

er

Flo

w R

ate,

kg

/sec

46

3.9

(963

.3)

445.

3 (9

81.8

) 45

4.2

(100

1.4)

(I bls

ec)

C

ham

ber

Pre

ssur

e,

atm

(p

sia

) 20

5.0

(3,0

14)

204.

1 (3

,000

) 20

1 .5

(2

,962

)

Mix

ture

Ra

tio

6.

3 i

.O

7.1

Component

Gimbal Main In jec tor Copper Chamber and Nozzle (E = 14.7) Tube Bundle Nozzle (E = 14.7 to 40) Extendible Nozzle* Extendible Nozzle Deployment System Fuel -Rich Preburners

O2lH2 02/RP-1

Oxidi zer-Ri ch Preburners

O2jH2 02/RP-1

Fuel Valves and Actuation Oxidizer Valves and Actuation Hot Gas Check Valves Low Speed LOX TPA Low Speed RP-1 TPA Low Speed LH2 TPA High Speed LOX TPA High Speed RP-1 P A

High Speed LH2 TPA Hot Gas Manifold Low Pressure Lines High Pressure Lines I gn i t i on System M i scel 1 aneous

TOTAL

*Tube bundle t o E = 132 and col umbiun t o the ex i t .

TABLE LXIX. - DUAL-FUEL ENGINE WEIGHT STATEMENT

Weiqht kg ( lb )

The center o f grav i ty and gimballed moments o f i n e r t i a for the engine are:

AXIS X . Y 7 . . - - "

Center of Gravity, cm ( i n .) - Nozzle Retracted 176 169.21 6.35 (-2.51 Nozzle Deployed 199 78.2 6.35 (-2.5

Y-Y - z-z - GI h a l l ed Iner t ia , kg-m2 (s 1 Nozzle Retracted 12,300) 16,500 12,160 Nozzle Deployed 18,960) 25,500 I 18,800 1


3. Alternate Mode 1 Engine Design

a. Configuration and Nominal Operating Conditions

The gimbal l ed envelope was evaluated f o r a 10' square pat tern and i s also shown on Figure 69. The square dimension i s 264.6 cm (104.2 in.).


Based upon the resu l t s o f Task I through V, the hydrogen cooled, gas generator cycle en ine was selected as the a1 ternate Mode 1 candidate. The cycle schematic 9 without boost pumps) i s shown i n Figure 68. This cycle uses LH2 t o cool the combustion chamber and nozzle and a fue l - r i ch LOX/LH2 gas generator t o d r i ve the t l l rbines o f the main propel lant pumps. LOX/RP-1 are in jec ted as 1 iquids i n the main in jec tor . The engine cycle features the hydrogen turbine i n series w i th the LOX and RP-1 turbines which operate i n pa ra l le l .

The prel iminary engine layouts are shown i n Figures 69 and 70.

The TCA design uses a s lo t ted zirconium copper chamber t o a nozzle area r a t i o o f 25:l. A s ingle pass A-286 tube bundle i s used from 25:1 t o the nozzle e x i t (area r a t i o o f 42.7:l). The coolant f low path i s shown on Figure 70.

The nominal engine operating conditions and pressure schedule i s presented on Tables LXX and LXXI, respectively.

The engine control requirements are pr imar i ly governed by the engine s t a r t and shutdown sequences. The sequence o f operation for the LOXIRP-1 components i s again patterned a f t e r the F-1 . The s t a r t and shutdown sequences are described on Tab1 e LXXII .

TABLE LXX. - ALTERNATE WDE 1 HYDROGEN COOLED, GAS GENERATOR CYCLE ENGINE OPERATING SPECIFICATIONS

Nominal MR = 2.9

Sea-Level Thrust, MN (1 b)

Vacuum Thrust, MN (1 b) Sea-Level Speci f ic Impulse, sec

Vacuum Speci f ic Impul se , sec

Total Flow Rate, kglsec (1 blsec)

Mixture Ratio (wLOX/wRp_l ) LOX Flow Rate, kglsec (1 blsec)

RP-1 Flow Rate, kglsec (1 b/sec)

Hydrogen Flow Rate, kg/sec (1 blsec)

Thrust Chamber

Sea-Level Thrust, MN (1 b)

Vacuum Thrust, MN ( l b )

Sea-Level Speci f i c Impul se , sec

Vacuum Specif ic Impulse, Sec

Chamber Pressure, atm (psia)

Nozzle Area Ratio

Mixture Ratio (ULOX/WRP-l )

Throat Diameter, cm ( in. ) Nozzle E x i t Diameter, cm ( in. )

Coolant Jacket Flow Rate (LH2), kglsec (1 blsec) Coolant Jacket AP, atm (ps i )

Coolant I n l e t Temp., O K (OR)

Cool ant E x i t Temp., O K (OR)

LOX Flow Rate, kglsec (1 blsec)

RP-1 F l ~ w Rate, kglsec ( Ib lsec)

Chamber length, cm ( in.)

Chamber Diameter, cm ( in . )

TABLE LXX (cont. )

Turbine Exhaust Performance

Sea-Level Thrust, N (1 b) Vacuum Thrust, N (1 b) Sea-Level Specific Impulse, Sec Vacuum Specific Impulse, sec Gas Flow Rate, kglsec (Iblsec)

Gas Generator

Chamber Pressure, atm ( psia) Combustion Temperature, O K (OR) Mixture Ratio LOX Flow Rate, kglsec (lblsec) Hydrogen Flow Rate, kglsec (Iblsec)

h)

--I

' Q

,

Tu

rbi n

es

TABL

E LX

X (c

on

t. )

LOX

Inle

t P

ress

ure,

at

m (

psi

a)

174.

9 (2

,571

)

Inle

t T

empe

ratu

re,

OK

(OR

) 85

9.4

(1,5

47)

To

tal

Gas

F

low

Rat

e,

kgls

ec

(1 b

lse

c)

9.66

(2

1.3)

Gas

P

rop

ert

i es

C ,

Sp

eci

fic

Hea

t @

Con

stan

t P

ress

ure,

~

j%

~

OK

(B

~U

II~

O

R)

Y,

Ra

tio

of

Sp

eci

fic

Hea

ts

Sh

aft

Hor

sepo

wer

, m

Hp

(HP

)

Eff

icie

ncy

, %

Spe

ed,

rpn

Pre

ssur

e R

ati

o (

To

tal

to S

tati

c)

Tu

rbin

e E

xit

Sta

tic

Pre

ssur

e,

atm

(p

sia

)

Tu

rbin

e E

xi t

To

tal

Tem

p. , O

K (O

R)

Ma i n

Pum

ps

To

tal

Ou

tle

t F

low

Rat

e,

kgls

ec

(lb

lse

c)

626.

1 (1

,380

.4)

Vo

lum

etr

ic F

low

Rat

e,

m31

sec

(gpm

) 0.

553

(8,7

30)

NPSH

, m

(ft

.)

96.3

(3

16)

Su

ctio

n S

peci

f i c

Spe

ed,

(RP

M) (m

31se

c) lI2/ 4

.14

(20,

000)

~

(m) 3

14

Spe

ed,

rpn

Dis

char

ge P

ress

ure,

at

m (

psi

a)

Nun

ber

of

Sta

ges

1

Sp

eci

fic

Spe

ed,

(RP

M) (

m3

/~e

c)~

l~/

(m)

314

RP- 1

174.

9 (2

,571

)

859.

4 (1

,547

)

4.85

(1

0.7)

Mai

n Pu

mps

(c

on

t .)

To

tal

Hea

d R

ise,

m

(ft

)

Eff

icie

nc

y,

%

Low

Spee

d TP

A

Punp

s

NPSH

, m

(ft

)

Inle

t Fl

ow R

ate,

kg

/sec

(l

b/s

ec)

Ou

tle

t Fl

ow R

ate,

kg

jsec

(l

b/s

ec)

Oi s

char

ge P

ress

ure

, at

m (

ps

ia)

Hyd

raul

ic T

urb

ine

Inle

t P

ress

ure

, at

m (

psi

a)

Ou

tle

t P

ress

ure

, at

m (

ps

ia)

Flow

Rat

e,

kg/s

ec

(lb

/sec

)

TABL

E LX

X (c

on

t.)

LOX

RP-1

TABLE LXXI. - ALTERNATE MODE 1 HYDROGEN COOLED, GAS GENERATOR CYCLE ENGINE PRESSURE SCHEDULE

Nominal MR = 2.90

Thrust Chamber Flows

Propel 1 ant Pressure, atm ( ~ s i a )

Main Pump Discharge AP Line Shutoff Valve I n l e t AP Shutoff Valve Main In jec to r I n l e t AP In jec to r Chamber Pressure

Propel 1 ant Pressure, atm ( ~ s i a )

Main Pump Discharge

Shutoff Valve I n l e t

AP Shutoff Yalve

Valve Out let AP Line Coolant Jacket I n l e t AP Coolant Jacket Coolant Jacket Outlet

AP Line G.G. Control I n l e t

AP Gas Generator G.C. I n jec to r I n l e t AP Gas Generator

Turbi ne In1 e t

LOX

347.0 (5,100)

3.33 (49) 343.7 (5,051)

3.47 (51) 340.2 (5,000) 51 .O (750)

289.2 (4,250)

Gas Genera t o r Flows

LOX

347.0 (5,100) --

RP- 1

347.0 (5,100)

3.33 (49) 343.7 (5,051)

3.47 (51) 340.2 (5,000)

51 .O (750) 289.2 (4,250)

TABLE LXXII. - SEQUENCE OF OPERATION HYDROGEN COOLED, GAS GENERATOR CYCLE

Star t 7

Purge Gas Generator and Thryst Chamber Oxidizer Lines and Manifold. Energi r e Spark Ign i ters . Open Gas Generator Ign i t e r Valves . Open Main LH2 Valve ( # I ) .* Open Gas Generator Ox. Valve (12). Open Thrust Chamber Ox. Valve (53). Open Thrust Chamber I gn i t e r Valves . Open Thrust Chamber RP-1 Valve (14).

Shutdown

Cutoff Gas Generator Spark Energy. Close Gas Generator Ign i te r Valves . Close Ox. Gas Generator Valve (12). Close Ox. Thrust Chamber Val ve (#3) . I n i t i a t e Ox. Purge. Close Main LH2 Valve. Close Thrust Chamber RP-1 Valve (14). Close Thrust Chamber Ign i t e r Val ves . Cutoff Thrust Chamber Spark Energy.

*Numbers re fe r t o the valves on Figure 68.

c. S ta r t and Shutdown Data

The engine i s assumed t o be bled-in and the hydrogen and oxygen components c h i l l e d down t o .ne main propel lant shutof f valves p r i o r t o rece ip t o f the s t a r t signal. The thrust, t o t a l impulse, and propel- 1 ant consumptions during t ransient gperation are suimarized i n Table LXXIII from 90% o f rated thrus t t o shutdown a t 5% o f rated thrust . Transient estimates are modeled a f t e r the F-! engine s t a r t and shutdown.

d. Design and Off-Design Engine Perfnrmance

Engine performance a t the design th rus t leve l over a mixture r a t i o range encompassing the design po in t mixture r a t i o 2 105 i s sumnarized on Table LXXIV.

A requirement f o r t h i s engine t o operate over t h i s range i n f l i g h t has not as yet, been ident i f ied . Therefore, the power balances assume that the engine o r i f i c e s i res are changed t o operate a t the various mixture ra t ios . The operating points are representative o f those t ha t would be required w i th a variable control valve although some penalty would be incurred a t the points where no o r i f i c e i s required.


The mass propert ies dated f o r the a1 ternate Mode 1 engine were calculated from the pre l i m i nary engine and component 1 ayout drawings.

The weight breakdown f o r t h i s engine i s shown on Table LXXV.

The center o f r a v i t y i n the 2xia l d i rec t ion was calculated t o be 93.7 cm (36.9 in . ! from the head-end o f the gimbal.

The center o f g rav i t y and gimbal 1 ed moments o f iner?3a f o r the engine are:

Center o f Gravity, cm (in.) 93.7 (36.9) -1.78 (-0.7) -3.8 (-1.5)

2 Gimballed Iner t ia , kg-m (s lug-f t2) 431 (318) 1301 (960) 1313 (969)


TABLE LXXII I. - ALTERNATE MODE 1 ENGINE START AND SHUT DOWN TRANS I ENT DATA SUmARY

Star t t o 90% F Time, sec 0.80 Total Star t Impulse, kg-sec (1 b-sec) 69,030 (1 52.200) LH2 Consunption, kg ( Ib) 52.6 (116) LOX Consunption, kg ( lb ) 166.5 (367)

RP-1 Consunption, kg ( l b ) 3.2 (7'1

Shutdown t o 5% F Time, sec 0.50 Total Shutdown Impulse, kg-sec (1 b-sec) 80,800 (178,200) LH2 Consuption, kg ( I b) 3.2 (7)

LOX Consumption, kg (1 b) 160.1 (353)

RP-1 Consunption, kg (1 b) 72.1 (159)

Eng

i ne

TABL

E LX

XIV

. - A

LTER

NATE

MOD

E 1

HYDR

OGEN

COO

LED

GAS

GENE

RATO

R CY

CLE

ENG

INE

DES

IGN

AND

OFF

-DES

IGN

MR

PE

RFOR

MAN

CE

Sea

-Lev

el

Thr

ust,

MN

(1

b)

Vacu

um T

hrus

t,

MN

(1 b)

Sea

-Lev

el

S~

ec

i f i c

Impu

l se ,

sec

Vacu

um S

peci

f i c

Impu

lse ,

sec

To

tal

Flo

w R

ate,

kg

lse

c (1

b/se

c)

LOX

Flo

w R

ate,

kg

lse

c (1

bls

ec)

RP-

1 F

low

Rat

e,

kgls

ec

(1 b/

sec)

Hyd

roqe

n F

low

Rat

e,

kg/s

ec

(1 b

lse

c)

Th

rust

Cha

mbe

r

Sea

-Lev

el

Thr

ust,

MN

(l

b)

Vacu

um T

hrus

t,

MN

(1 b

)

Sea

-Lev

el

Spe

ci f i c

Impu

l se ,

sec

Vacu

um S

pe

cifi

c Im

puls

e , se

c

Cha

mbe

r pr

essu

re,

atm

(p

sis)

Ena

ine

Mix

ture

Ra

tio

TABLE LXXV. - ALTERNATE MODE 1 HYDROGEN COOLED, GAS GENERATOR CYCLE ENGINE WEIGHT STATEMENT

1.j Component

Gimbal 95.7

Main In jec tor 348.8 Copper Chamber and Nozzle ( e = 25) 157.4

Tube Bundle Nozzle (e = 25 t o 42.7) 92.1

Fuel -Rich Gas Generator 9.1

Fuel Valves and Actuation

Oxidizer Valves and Actuation

Low Speed LOX TPA

Low Speed RP-1 TPA

Low Speed LH2 TPA High Speed LOX TPA

High Speed RP-1 TPA

High Speed LH2 TPA

Low Pressure Lines

High Pressure Lines I gn i t i on Systen

M i sce l l aneous

TOTAL

E. TURBOMACHINERY DESIGN

Preliminary designs f o r the high speed and low speed turbopumps were established f o r the three candidate engine designs,

1 . Basel ine Mode 1 2. Dual-Fuel 3. Alternate Mode 1

The turbomachinery was designed i n general accordance wi th the engine cycle balance data presented i n Section V I I I D . Some modifications t o the data would be required f o r the next design i tera t ion.

1 . Base1 ine Mode 1 Engine

The low speed LOX turbopump f o r the baseline Mode 1 engine i s shown on Figure 71. This pump uses a sc ro l l type discharge t o accomnodate the engine assembly. The shaft i s supported by two spring loaded angular contact ba l l bearings. A t the design point, the estimated pump ef f ic iency i s 85%. The hydraulic turbine i s driven by f low from the high speed pump. Low speed pump and turbine design data are shown on Table LXXV I.

The high speed LOX pump i s shown i n Figure 72. Signif icant design features include:

" High head inducer w i th tap-off f low f o r the low speed pump turbine.

Two centr i fugal pump stages.

O Axial thrust reacted by self-compensating thrust balancer.

Each bearing package consists o f two spring loaded aiigular contact ba l l bearings, one located between inducer and f i r s t stage impeller, and the second between the l a s t impel 1 er and turbine.

O Single stage turbine; configured for zero reaction a t the ro to r hub wi th f ree vortex-velocity d i s t r i bu t i on from hub t o t i p .

The pump and turbine design parameters are l i s t e d i n Table LXXVII. The pump performance a t the design point i s estimated t o be 82%. The turbine was configured as a s ingle stage wi th f ree vortex ve loc i ty diagrams (approximately 10% reaction a t the mean diameter). Based on the information i n Reference 74 and the design r a t i o o f mean blade speed t o nozzle spouting velocity, the turbine e f f ic iency a t the design point, i s estimated t o be 80% based on the t o t a l to s t a t i c pressure ra t io .

The RP-1 turbopumps designed f o r the baseline Mode 1 engine are shown on Figures 73 and 74.

-67.

3 cm---\

(26.

50

in.)

Figu

re 7

1 .

Bas

e1 i ne

Mod

e 1 , D

ual -

Fuel

an

d A1

tern

ate

Mod

e 1

Low

Spze

d LO

X TP

A

FOLD

TABLE LXXVI. - BASELINE MODE 1 AND DUAL-FUEL ENGINE LOW SPEED LOX TPA DESIGN POINT

Pump Turbine

Speed, rpn FIOW, m3/sec (gm)

Head, m ( f t )

Power, MP (hp) NPSH, m ( f t )

Suction Sped f i c Speed

Specific Speed

No. o f Stages

Efficiency, %

2560

(8,785) 0.111 (1,755)

(326) 762.0 (2,500) (1,000) 1014 (1,000)

(16) - - (30,000) -- (3,127) 0.177 (857) (Stage)

4 81

FLOW

TO

OUT

LET

FLOW

l

CENT

RIFU

GAL

\

I 1

'

(19.

10

in.)

4e

.5 cm

I --

<

PREB

URNE

R HS

G

78.2

Cm-

(3

0.80

in

.)

Fig

ure

72.

B

ase1

ine

Mod

e 1

and

Dua

l-Fue

l H

igh

Spee

d LO

X TP

A

TABLE LXXVII. - BASELINE MODE 1 AND DUAL-FUEL ENGINE HIGH SPEED LOX TPA DESIGN POINT

15,000

0.665 (1 0,540)

0.554 (8,785)

Inducer Head, m ( f t ) 823.0 ( 2,700)

Stage Head, m (ft) 2418.6 (7,935)

Overall Head, m ( f t ) 5660.1 (18,570)

No. o f Stages 2 + Inducer

8 Power, mHp (hp) 60,434 ( 59,600) $

K Inducer Suction 4.14 (20,000) Speci f i c Speed

NPSH, m ( f t )

Inducer Specif ic Speed

Stage Speci f i c Speed

Overall E f f i c i ency , %

- . , > / j ) i . . . .

j 7

Twb i ne

No. o f Stages

Speed, rpm Flow, kg/sec (1 b/sec) I n l e t Pressure, atm (psia) Pressure Ratio

r: I n l e t Temperature, O K ( O R )

Nozzle Outlet Velocity , m/sec ( ft/sec) Blades Speed (mean), m/sec ( f t /sec)

i I Efficiency, %

. i i 1 Rotor Tip Dia, cm (in.) Hub Dia, cm (in.)

2 2 Annular Area, Aa, cm ( in . )

i Aa~2 (Annular Area x Speed : t, Squared)

TABLE LXXVI I (cont . )

&u % % Q fa mv,

The design parameters f o r the low speed RP-1 pump are l i s t e d i n Table LXXVIII. The low speed pump incorporates a s c r o l l discharge design t o accommodate the engine i n s t a l l a t i o n . The estimated pump performance a t the design po in t i s 83%. The tu rb ine vector diagrams were selected f o r a r a t i o o f mean blade speed t o nozzle discharge v e l o c i t y of approximately 0.5 (near optimum f o r an impulse turb ine) . The hydraul ic tu rb ine i s d r iven by propel- l a n t f low from the main pump.

The high speed fue l turbopump i s shown i n Figure 74. The con- f i g u r a t i o n i s bas i ca l l y the same as the LOX turbopump and incorporates essen t i a l l y the same design features. The pump and tu rb ine design parameters are 1 i s t e d i n Table LXXIX. The estimated pump performance a t the design po in t i s 82%. The tu rb ine was configured f o r f r e e vortex blading w i t h approximately 10 percent reac t ion a t the r o t o r mean diameter. The tu rb ine e f f i c i e n c y i s estimated from Ref. 74 t o be a t l e a s t 75 percent. The b lading cent r i fuga l s t ress w i l l be q u i t e conservative based on A ~ N Z (see Fig. 30, Page 45 o f Ref. 74).

2. Dual-Fuel Engine

The LOX and RP-1 turbopumps used f o r the dual- fuel engine are :he same as those designed f o r the baseline Mode 1 engine.

The low speed LH2 pump f o r use on the dual- fuel engine i s shown on Figure 75. The conf igurat ion consists o f a high head inducer stage dr iven by a four stage hydraul ic turbine. The design po in t parameters are 1 i s t e d i n Table LXXX. The pump performance a t the design p o i n t i s 85%. The r o t a t i ~ g assembly i s supported by spr ing loaded angular contact bearings as indicated i;? Figure 75. The ax ia l t h rus t on the bearings i s kept low by placement o f the l aby r in th seal between the pump and turb ine.

The high speed LH2 turbopump conf igura t ion i s shown i n Figure 76. As w i t h the PP-1 and LOX high speed pumps, the inducer stage supplies the f l u i d t o d r i v e the hydraul ic turb ine. A three-stage cen t r i f uga l pump was necessary t o maintain the stage spec i f i c speed above 1000. The tu rb ine end bearing package i s outboard o f the turb ine t o avoid e i t h e r h igh shaf t s t ress o r an excessive value o f bearing DN.

The pump and turb ine design parameters are l i s t e d i n Table LXXXI. A1 though the turbopump conf igurat ion s a t i s f i e s a1 1 the s t i pu la ted design c r i t e r i a ; f u r the r design refinements would be desi rable t o reduce weight. Further studies would be required t o explore po ten t i a l shaf t c r i t i c a l speed and v i b r a t i o n problems.

3. Al ternate Mode 1 Engine

The a1 te rna te Mode 1 hydrogen cool ed, gas generator cyc le engine has lower pump discharge requirements than the other candidates as discussed

TABLE, LXXVSSI. - BASELINE MODE 1 AND DUAL-FUEL ENGINE LOW SPEED RP-1 TPA DESIGN POINT

Speed, rFJm FIOW, m3/tec (gpn) Head, m ( f t ) Power, mHp (hp) NPSH , m ( f t ) Suction Spec1 f i c Speed Spec i f i c Speed Stages Efficiency, %

871 5 (4310) 0.054 (862) (370) 871.7 (2860)

( 385 390 (385) (65) (25 ,ooo) (6700) 0.308 (1490) (Stage)

3 78

TABLE LXXIX. - BASELINE MODE 1 AND DUAL FUEL ENGINE HIGH SPEED RP-1 TPA DESIGN POINT

Pump Speed, rpm I n l e t flow, m3/sec (gpm) ~ x i t flow, m3/sec (gpn) Inducer Head, m ( f t ) Stage Head, m (ft) Overall head, m ( ft)

Number o f stages

Power, mHp (hp) Inducer Suc+ on Specific speed

NPSH, m (ft) Inducer Speci f i c Speed Stage Specif ic Speed Overall Efficiency, %

(5,172) (4.310) (3,154) (9,928) (23,010)

2 + Inducer (26,200)

TABLE LXXIX (cont. )

Turbine

'f - 1! (I.

1'3. o f Stages

Speed, rpm

Flow, kg/sec ( lb lsec) I n l e t Pressure, atm (psia) Pressure Ratio I n l e t Temperature, O K ( O R )

Nozzle Out le t Velocity, m/sec ( f t l s e c ) .

Blade S eed (mean), m/sec ( f t /sec ! Efficiency, %

Rator T ip Dia, cm ( in, ) Hub Dia, cm ( in. )

2 Annular Area, c,n2 ( i n . ) Aa N~ (Annular Area x Speed Squared)

TABLE LXXX. - DUAL-FUEL ENGINE LOW SPEED LH2 TPA DESIGN POINT

Speed, rpn FI OW, m3/s)/sec (gpnj Head, m ( f t ) Power, mhp (hp) NPSH Suction Specific Speed Specific Speed Stages Efficiency, %

Pump Turbine

11,420

(1 4,240) 0.180 (2,848)

(1,800) 41 45 (1 3,600) (538) 545.5 (538) (loo) - 0

(43,100) - o

(4,930) 0.283 (1,369) (Stage) 4 78

106.

7 cm

-

(42.

00

in.)

63.2

cm-

w

rTH

RU

ST

BAL

ANCE

R W

TW

O ST

AGE

(24.

90

in-)

TO

UTLE

T FL

OW

/ TH

REE

STAG

E \-

- I

/ /

Fig

ure

76.

D

ual -

Fuel

H

igh

Spe

ed L

H2 T

PA

TABLE LXXXI. - DUAL-FUEL ENGINE HIGH SPEED LH2 TPA DESIGN POINT

Punp

Speed, rpm I n l e t flow, m3/sec (gpa)

~ x i t FIOW, m3/sec (gpn)

Inducer Head, m (ft) Stage Head, m (ft)

Overall Head, m (ft) No. o f Stages

Power, ~ H P (hp) Inducer Suction Speci f i c Speed

NPSH, m ( f t )


Stage Speci f ic Speed

Overall Efficiency, %

(1 7,090)

(1 4,240)

(1 5,000)

(55,133) (1 80,400)

3 + Inducer

( 59,000) (20,000

( 100)

(3,183)

(1,095)

Turbine

No. o f Stages

Speed, rpm

F l ow, kg/sec

I n l e t Pressure, Atm (psia)

Pressure Ratio

I n l e t Temperature, O K (OR)

TABLE LXXXI (cont . )

F i r s t Stage Nozzle Outlet Velocity, mlsec ( f t l sec ) 762.0 (2500)

Rotor Blade Speed (mean), m/sec ( f t l sec ) 439.2 (1 441 )

Overall Efficiency, % 81

Second Stage Rotor Tip Dia., cm ( in. ) 28.96 (11.4)

Hub Dia., ( in. ) 21.84 (8.6) 2 Annular Area, Aa, cm2 (in. ) 283.9 (44.0)

2 AaN (Annular Area X Speed Squared) 309.1 x l o 9 (47.9 lo9)

i n Section VC. However, two more pumps are requi red f o r the LH2 flow. Because o f the reduced pump discharge pressure requirement, the main LOX and RP-1 pumps can operate a t higher speeds than the base1 i n e Mode 1 engine pumps. Consequently, the required NPSH f o r the h igh speed pump inducers i s greater. The Mode 1 engine low speed LOX and RP-1 pumps meet these requirements w i t h a s l i g h t increase i n speed. The LH2 turbopumps are essen t i a l l y scaled down versions o f the dual- fuel engine LH2 turbopumps.

The low speed LOX turbopump selected f o r the a l te rna te Mode 1 engine (Figure 71 ) i s the same as t h a t selected f o r the baseline Mode 1 and dual - f ue l engines. However, the operat ing speed i s increased from 2,560 rpm t o 2,650 rpm. The small increase i n ro ta t i ona l speed does no t s i g n i f i c a n t l y affect the low speed pump NPSH requirement and the design po in t performance i s reduced by on ly approximately 0.1%.

The h igh speed LOX pump conf igurat ion i s shown on Figure 77 and the design parameters are l i s t e d i n Table LXXXII. The turbopump con- s i s t s o f a high head inducer, a s ing le stage cen t r i f uga l pump and a v e l o c i t y compounded tu rb ine (Cur t i s Stage). The r o t o r i s supported by two bearing packages, each o f which consists o f two spr ing loaded angular contact b a l l bearings. Rotor t h r u s t i s balanced by a balance piston. F l u i d i s tapped- o f f o f the h igh head inducer t o supply the hydraul ic tu rb ine which dr ives the low speed LOX pump. A pre l iminary design f o r the hot gas seal between the pump and the f u e l - r i c h tu rb ine i s a lso shown on the f igure . The pump con- f i g u r a t i o n shown has not been optimized from the standpoint o f c r i t i c a l speeds and r o t o r dynamics a t design speed. Further refinements i n the conf igura t ion would be required t o achieve an optimum r e l i a b l e design. The weight, however, would no t change appreciably from the weight o f the conf igurat ion shown.

The low speed RP-1 turbopump selected f o r the a l t e rna te Mode 1 engine i s i den t i ca l w i t h t h a t selected f o r the Mode 1 and dual- fuel engines (Figure 73). I t s operat ing speed, however, was increased from 8715 rpm t o approximately 8800 rpm t o increase the head from 112M (366 fee t ) t o 123M (405 feet) . The small increase i n speed does not s i g n i f i c a n t l y a f f e c t the NPSH requirement o r the pump performance.

The high speed (RP-1) pump conf igura t ion i s shown on Figure 78, and the design parameters are 1 i s t e d i n Table LXXXIII. The turbopump con- s i s t s o f a high head inducer, a s ing le stage cen t r i f uga l pump, and a ve loc i t y compounded tu rb ine (Cur t i s Stage). The r o t o r i s supported by two bearing packages each o f which consists of two spring loaded angular contact b a l l bearings. Rotor t h r u s t i s balanced by a hydrostat ic balance piston. As w i t h the LOX TPA for the a1 te rna te Mode 1 engine, r o t o r dynamics studies and c r i t i c a l speed estimates should be made i n the next design i t e r a t i o n .

The low speed LH2 turbopump i s shown i n Figure 79, and i s essen- t i a l l y a scaled version o f the low speed pump configured f o r the dual-fuel engine. The pump and hydraul ic tu rb ine design parameters are l i s t e d i n

PUMP

OUT

LET

9.9

CM

TURB

INE

INLE

T FR

OM

H.S

. LH

TP

A 6.

1 CM

f2.

4 IN

.)

DIA

\

FLOW

15.2

CM

(6.0

0 IN

.)

DI A

I --.A

(36.

95

IN.)

Fig

ure

77.

A

lte

rna

te M

ode

1 H

igh

Spe

ed L

OX T

PA

Pump

TABLE LXXXII. - ALTERNATE MODE 1 ENGINE HIGH SPEED LOX TPA DESIGN POINT

1 1 ; Speed, RPM I I n l e t Flow, m3/sec (gpm)

. , j / o u t l e t FIOW, m3/sec (gpm) , .

Inducer Head, m ( f t )

! 1 : Stage Head, m ( f t ) 1 .

Overall Head, m ( f t ) ! . I No. o f Stages I Power, mHp (hp)

Turbine

Inducer Suction Specif ic I

I NPSH, m ( f t )

Inducer Specif'c Speed i Centrifugal Stage Specific Speed

Overall Eff ic iency, %

Number o f Stages

. . Stage Type

Speed, rpm

Flow, kg/sec (1 blsec)

I n l e t Pressure, atm (psia)

Pressure Ratio (Total t o Sta t ic )

I n l e t Temp., O K (OR)

Nozzle Vel oci t y (Ideal ) , mlsec (f t lsec)

Mean Blade Speed, mlsec ( f t l sec )

Overall Ef f ic iency , %

Second Rotor Tip Dia, cm ( in . ) Hub Dia, cm ( in. )

2 2 Annular Area, cm ( in . ) A ~ N ~ *Annular Area x Speed Squared)

(1 0,476)

(8,730)

(2,900)

(7,130) (1 0,030) 1 + Inducer

(33,300)

(20,000)

( 360 (4,170)

(1,939)

Vel oc i t y Compounded (Curt is Stage)

TABLE LXXXIII. - ALTERNATE MODE 1 ENGINE HIGH SPEED RP-1 TPA ,IESIGN POINT

Pump

Speed RPM 3 In le t , Flow m /sec (gpm)

Outlet FIOW, m3/sec (gpn) Inducer Head, m ( f t )

Centrifugal Stage Head, m ( f t )

Overall Head, m ( f t )

No. o f Stages

Power, mHp (hp)

Inducer Suction Specific Speed

NPSH, m ( f t j Inducer Speci f i c Speed

Centrifugal Stage Specif ic Speed

Overall Efficiency, %

Turbi ne

Number o f Stages

Stage Type

Speed, RPM

Flow, kglsec (1 blsec)

I n l e t Pressure, ata (psia)

Pressure Ratio (Total t o S ta t i c ) I n l e t Temp., O K (OR)

Nozzle Velocity (Ideal ), m/sec ( f t l sec ) Mean Blade Speed, m/sec ( f t l sec )

Overall Eff ic iency, %

Second Rotor Tip Dia, cm (in.)

Hub Dia, cm ( in. )

Annular Area, an2 (in.2)

A ~ N ~ (Annular Area x Speed Squared)

(5,136)

(4,280)

(4,000) (! 0,270) (1 4,270)

1 + Inducer

(16,800) (20,000)

(450)

(4,260) (1,900)

Velocity Compounded (Curt is Stage)

INDU

CER

12.3

-cm

(4.8

4 in

.)

DIA . I

PUM

P O

UTLE

T 7.

1 an

(2.7

8 in

.)

FLOW

FRO

M H

.S.

4 ST

AGE

TUR

BIN

E

OUT

LET

TUR

BIN

E

Fig

ure

79.

A7 te

rnat

e M

ode

1 Lo

n Sp

eed

LH2 TPA

Table LXXXIV. The pump e x i t s c ro l l configurat ion was selected f o r ease o f mounting on the engine.

The high speed LH2 turbopump configurat ion i s show^ i n Figure 80, and the design parameters are 1 i s t ed i n Table LXXXV. The configurat ion i s a scaled version of the high speed LH2 pump configured f o r the dual-fuel engine . F. THRUST CHAMBER DESIGN ANALYSES

1 . Performance

The object ive o f the performance analysis was t o pred ic t the del ivered spec i f ic impulse performance o f the candidate engines using JANNAF methodology. As par t o f t h i s analysis, i t was necessary t o establ ish an "optimum" compromise nozzle contour f o r the two modes o f operation of the dual-fuel engine.

a. Dual -Fuel Nozzle Contour Analysis

The object ive of the dual-fuel contour analysis was t o determine re l a t i ve performance t radeof fs and t o select a nozzle f o r the dual- fuel engine which o f fe rs the best performance i n both modes o f operation. I n the analysis, 40:1 and 200:l expansion r a t i o nozzles o f various lengths were generated w i th a RAO optimum nozzle contour pro ram. For combinations o f these contours, the TDK computer program (Ref. 51 3 was u t i l ized t o evaluate the performance f o r various Mode 1 and Mode 2 engine configurations.

The Mode 1 performance was evaluated w i th a LOX/RP-1 propel- l an t combination a t a mixture r a t i o o f 2.9 w i th gamna equal t o 1.129. Mode 2 performance i s based on a LOX/LH2 nozzle combination a t a mixture r a t i o of 7.0 w i th gamma equal t o 1.210.

A less than optimum Mode 2 contour resu l ts from combining an optimum 40:l Mode 1 contour w i th an optimum 200:l contour. The opposite i s also true. I n general, 3 performance t radeof f ex is ts between the various Mode 1 and Mode 2 nozzle combinations. The resu l ts o f the study are sumnarized on Tab1 e LXXXVI .

Case 1 i l l u s t r a t e s the r e l a t i v e performance values for the shortest possible optimum 200:l nozzle contour (73.2% be l l based on RAO c r i t e r i a ) combined w i th an optimum 40:l contour o f su f f i c i en t length (83.7% b e l l ) t o provide a 90% be l l nozzle for the Mode 2 engine configuration. Case 2 i l l u s t r a t e s the r e l a t i ve performance values f o r the shortest possible optimum 40:l nozzle contour (57.9% be l l based on RAO c r i t e r i a ) combined w i th an optimum 200: 1 contour o f su f f i c i en t length (85.4% be l l ) t o provide a 90% be l l nozzle f o r the Mode 2 engine configuration. Case 1 resu l ts i n the best performance a t E = 40:l a t the expense o f the performance a t E = 200:l. Con- versely, Case 2 gives the best performance a t E = 200:l a t the expense o f

TABLE LXXXIV, - ALTERNATE MODE 1 ENGINE LOW SPEED LH2 TPA DESIGN POINT

Pump Turbine

Speed, RPM 18,790 18,790

FIOW, m31sec (gpm) 0.129 (2,040) 0.026 (4081

Head, m ( f t ) 289.6 (950) 2438.4 (8,000)

Power, mHp (hp) 41 .G (41 ) 4 i .6 (41

NPSH, m ( f t ) 27.43 (90) - - Suction Specif ic Speed 6.003 (29,000) -- Speci f i c Speed 1.027 (4,960) 0.263 (1,270) (Stage)

No. of Stages 1 Efficiency, % 85

TABLE LXXXV. - ALTERNATE MODE 1 ENGINE HIGH SPEED LH2 TPA DESIGN POINT

Pump

Speed, RPM 3 I n l e t Flow, m /sec (gpm) 3 Out le t Flow, m /sec (gpm)

Inducer Head, m ( f t )

Stage Head, m ( f t )

Overall Head, m (ft)

No. o f Stages

Power, d i p (hp)

Inducer Suction Speci f ic Speed

NPSH, m ( f t )


Stage Specif ic Speed

Overall E f f i c iency , %

Turbine

No. o f Stages

Speed, RPM

Flow, kg/sec ( Ib lsec)

I n l e t pressure, atm (psia)

Pressure Ratio (Total t o S ta t i c ) 1.43

I n l e t Temp., O K (OR) 922.2

Nozzle Velocity, m/sec { f t l sec) 898.6

Mean Blade Speed, m/sec ( f t /sec 404.5

Eff ic iency, % 7 2

Last Stage Rotor T ip Dia, cm ( in. ) 12.62

Hub Dia, cm ( in. ) 9.45 2 Annular Area, Aa, cm2 ( i n . ) 54.84

N~ Aa 271 l o 9

TABLE LXYXVI . - DUAL-FUEL NOZZLE CONTOUR EVALUATION

Case - 1 Mode 1

Mode 2 (Overal l )

2 Mode 1

Mode 2

3 Mode 1 Mode 2

4 Mode 1 Mode 2

% Bel l Length .

*Compared t o Optimum 80% Mode 1 and 90% Mode 2 Be1 1 s .

Performance Change, %*

-Om2 ' Design -0.68

the performance a t E = 40:l ( w i t h i n the design cons t ra in t o f a 90% b e l l leng th f o r t he E = 200:l nozzle). Note t h a t i n e i t h e r case, the Mode 2 nozzle r e s u l t s i n a performance loss from the s ing le Mode 2 base1 i ne nozzle and as the Mode 2 nozzle performance loss i s minimized, t he Mode 1 performance loss i s s i g n i f i c a n t l y increased.

As a resu l t , longer Mode 2 nozzle contours were invest igated i n order t o improve the Mode 2 performance wi thout changing the Mode 1 performance.

Case 3 i l l u s t r a t e s the r e l a t i v e performance f o r an 80% be1 1 Mode 1 nozzle combined w i t h an optimum 110% be1 1 nozzle f o r the Mode 2 engine. I n t h i s arrangement, a 0.85% s p e c i f i c impulse performance loss occurs i n the Mode 2 engine w i t h no loss f o r tbe Mode 1 engine. The Mode 2 ove ra l l nozzle leng th f o r t h i s concept i s equiva lent t o a 121% b e l l nozzle.

The best ove ra l l engine performance appears t o be obtained by s a c r i f i c i n g some perfomance f o r bozh the Mode 1 and Mode 2 conf igurat ions as i l l u s t r a t e d i n Case 4. I n t h i b case, an optimum 75% b e l l nozzle (Mode 1 ) i s combined w i t h an optimum 110% b e l l , whicn r e s u l t s i n a Mode 2 engine w i t h an ove ra l l 119% b e l l nozzle length. This combination e f f e c t s a 0.21% and 0.68% drop i n Mode 1 and Mode 2 vacuum s p e c i f i c impulse, respect ive ly , compared w i t h the base1 ine 80% b e l l Mode 1 and 90% b e l l Mode 2 nozzles. This case was selected f o r design. The nozzle i s heavier but the performance loss i s reduced t o 0.7 sec i n Mode 2.

b. Del i vered Performance

Tables LXXXVII through XC present de l i vered I sp a t nominal and - + 10% o f nominal mixture r a t i o f o r various engf ne candidates.

The del ivered performance va? ues were ca lcu lated using the JANNAF simp1 i f i e d ca l cu la t i on technique described i n Section 3 o f CPIA pub1 i- cat ion 246. The ODK (One Dimensional K ine t i c ) computer program was used t o deternii ne the theore t ica l vacuum spec i f i c impul se and k i n e t i c performance loss, fo r the respect ive propel l a n t combinations and mixture r a t i o s . Del i vered I sp was obtained by cor rec t ing the ODK performance f o r the fo l low ing losses: (1 ) i n j e c t o r energy release e f f i c i ency , (2) nozzle divergence e f f i c i ency , (3) nozzle and chamber boundary l aye r loss, and (4) where appl icable, a sea- l e v e l I sp correct ion. For the gas generator cycle, an add i t iona l ca lcu lated 1.5 second loss was subtracted as a r e s u l t o f the gas-generator f low being dumped downstream o f the throat . To make up f o r t h i s loss such t h a t the del ivered vacuum and sea-1 eve1 del i vered s p e c i f i c impul se are near ly equiva- l e n t t o the baseline performance, the gas generator nozzle expansion r a t i o and chamber pressure were increased t o 42.7:1 and 289 atm (4250 ps ia) , respec- t i v e l y ,

An i n j e c t o r energy release e f f i c i e n c y o f 0.985 was used. This i s based on a 24.8 cm (9.75 in . ) chamber length f c r the base1 i ne Mode 1 and

TABL

E LX

XX

VII.

-

MOD

E 1

BASE

LIN

E D

ELIV

ERED

SPE

C1"C

IM

PULS

E

E

= 40

:l

80%

Be

ll L

engt

h

Mix

ture

Ra

tio

(O

/F)

ODK

Vacu

um I

(sec

) s P

E

nerg

y R

elea

se E

ffic

ien

cy

Div

erge

nce

Eff

icie

nc

y

Co

rre

cte

d I

(se

c)

SP

Bou

ndar

y La

yer

Loss

(s

ec)

PC =

272

atm

(40

00 p

sia

)

Pro

pel 1

an

t LO

X/R

P-1

(Nom

i na

l )

2.61

2.

90

3.19

I De

live

red

Vac

uum

IS

p

(sec

) 34

9.1

350.

6 34

7.7

1 Se

a Le

vel I

Co

rre

ctio

n (

sec)

s P

27

.4

27.0

26

.6

L~

eli

ve

red

Sea

Leve

l I

(sec

) I

s P

321

.7

323.

6 32

1.1

I

TABL

E LX

XX

VII

I. - M

ODE

1 GA

S GE

NERA

TOR

CYCL

E D

ELIV

ERED

SP

EC

IFIC

IM

PULS

E

= 28

9 a t

m (

4250

ps

i )

80%

Be1

1 L

engt

h P

rope

l 1 a

nt

= LO

XIR

P-1

Mix

tuv

Ra

tio

(O

/F)

(Nom

inal

) 2.

61

2.90

3.

19

ODK

Vac

uwn I

(se

c)

360.

0 36

1.6

358.

7 s P

E

nerg

y re

lea

se e

ffic

ien

cy

Div

erge

nce

eff

icie

nc

y

Co

rre

cte

d I

(sec

) SP

B

ound

ary

lay

er

loss

(se

c)

Gas

G

ener

ator

lo

ss

(se

c)

Del

i ve

red

Vacu

um I

(sec

) I

349.

1 35

0.7

347.

8 SP

Sea

lev

el I

Co

rre

ctio

n (

sec)

27

.5

27.2

26

.8

SP

De

live

red

Sea

Le

vel I

(sec

) 32

1.6

323.

5 32

1 .O

SP

TABL

E LX

XX

IX.

- MOD

E 1

DU

AL-F

UEL

D

ELIV

ERED

SP

EC

IFIC

IM

PULS

E

E

= 40

:1

75%

Be1

1 L

engt

h

= 27

2 at

m (

4000

psi

a)

Pro

pel 1

an

t LO

X/R

P-1

(Nom

i na

l )

Mix

ture

Ra

tio

(O

/F)

2.61

2.

90

3.19

ODK

Vacu

um IS

p

(se

c)

358.

8 36

0.3

357.

3

Ene

rgy

rele

ase

eff

icie

nc

y

Di v

erge

nce

eff

icie

nc

y

Co

rre

cte

d I

(sec

) 35

0.4

351.

9 34

9.0

SP

Bou

ndar

y la

ye

r lo

ss (

sec)

2

.O

2.0

2 .O

a

De

live

red

Vac

uum

I

(se

c)

348.

4 34

9.9

347.

0 SP

Sea

Leve

l I

Co

rre

ctio

n (

sec)

27

.4

27.0

26

.6

s P

De

live

red

Sea

Le

vel IS

P

(se

c)

321

.O

322.

9 32

0.4

TABL

E XC

. -

MODE

2

DU

AL.

LFU

EL D

ELIV

ERED

SP

EC

IFIC

IM

PULS

E

Mix

ture

Ra

tio

(O

/F)

ODK

Vacu

um I

(se

c)

s P

E

= 2

00:l

PC =

20

4 at

m (

3000

ps i

a)

119%

Be

ll L

en

gth

P

rop

el 1

an

t LO

X/L

H2

(See

E

ncl

osu

re 4

)

Ene

rgy

rele

ase

eff

icie

nc

y

Div

erge

nce

eff

icie

nc

y

Co

rre

cte

d I

(se

c)

s P

Bou

ndar

y L

aye

r Lo

ss

(se

c)

I De

live

red

Vac

uum

I

(se

c)

SP

461.

6 45

9.2

455.

4 I

dual-fuel engines and a 72.8 cm (18.50) length f o r the a1 ternate Mode 1 engines. The nozzle divergence e f f i c i ency i s based on the value computed w i t h the TDK program (Ref. 51) using the ideal gas opt ion.

The chamber boundary 1 ayer f r i c t i o n 1 oss was calculated w i t h an ALRC computerized formulat ion o f the JANNAF boundary l aye r char t technique (Ref. 52). Sea-level performance was cal cul ated by subtract ing the ambient pressure nozzle e x i t force from the del ivered vacuum performance.

2. Thermal Analysis

Heat t rans fe r and hydraul ic analyses were conducted t o formulate regenerat ively cooled th rus t chamber designs f o r the three candidate engine concepts. The analyses were conducted using the Task V I performance governed nozzle contour resu l t s o f the previous paragraphs, updated cyc le l i f e governed thermal 1 i m i ts , and chamber curvature enhancement o f the 1 i q u i d s ide heat t rans fe r coe f f i c i en t .

The resu l t s from the cyc le l i f e c r i t e r i o n and wa l l s t rength l i m i t s are presented on Figures 81 and 82. Figure 81 presents two curves showing the a1 lowable temperature d i f f e r e n t i a l AT, from the maximum gas-side wal l temperature (Tw ) t o the average external wal l temperature (TBs) as a func- t i o n o f the l a t ? e r temperature f o r the bar re l sect ion o f the chamber and f o r the remainder of the chamber and nozzle. Figure 82 presents the al lowable channel width t o wal l thickness r a t i o as a func t ion o f the chamber gas-side wal l temperature f o r three pressure d i f f e r e n t i a l values across the gas-side wal l . This curve i s appl icable t o zirconium-copper s l o t t e d chambers which are proposed f o r use i n a l l three engine concepts.

The curvature enhancement e f fec ts were t reated i n the same manner as tha t used i n Reference 75. This cor rec t ion f o r the e f f e c t s o f curvature on f r i c t i o n coe f f i c i en ts i s a lso appl i cab le t o l oca l heat t rans fe r coeff i - c ients. For the purposes o f t h i s analysis, on ly the heat t rans fer coef f i - c ien ts o f the gas side l i q u i d wal l was corrected, the other wal ls of the passage were exempted from curvature e f fec ts .

The coolant heat t rans fer co r re la t i on used f o r the baseline and the dual-fuel engine concepts where LOX cool ing i s employed was the same as used i n the Task I 1 analysis, wi'h curvature enhancement e f fec ts added. The use o f the superc r i t i ca l oxygen co r re la t i on developed a t ALRC has been pre- v ious ly described i n Section I V D along w i t h the co r re la t i on used f o r the hydrogen cooled gas generator cycl e concept. Thi s co r re l a t i on, developed f o r s u p e r c i r i t i c a l hydrogen by Hess and Kunz (Ref. 48) i n 1964, was a lso modif ied t o include %rvature enhancement e f fec ts .

Gas s ide heat t rans fer cor re la t ions were the same as those used i n Task 11.

Allowable Temperature D i f f e r e n t i a l (AT) = TWG - TBS

For Various Bdckup Cyl i nder Wall Temperatures (TRq)

T~~

Throat, Convergent & Divergent Region tg = .762 cm (.3 in . )

= 867°K (1100°F) 800

h

Y 0 u

I- a Barre l

7 ~ ; n Region tg = 1.524 cm (.6 in.)

Safe Zones i .e . Temperatures

Pred ic t > 1000 Cycles t o F a i l u r e

\ I 1

I I

1 I

I I

1 I I .

150 200 2 50 300 350 400 TBS ( O K ) 700

F i g u r ~ 81. ZrCu S lo t t ed Chamber L i f e L im i t s

- - I 1

I I I I I I

I 1

I I

- 300 -200 -1 00 0 100 200 300 Backup Cyl inder Wall Temperature, TBS (OF)

A l l three of the engine chamber designs generated u t i l i z e d a s i m i l a r coolant f l ow pattern. I n a l l three designs the coolant was i n i t i a l l y i n t r o - duced i n t o a s l o t t e d zirconium-copper chamber j u s t downstream from the th roa t a t an area r a t i o o f 1.5:l and flowed through the th roa t region up t o the plane o f the i n j e c t o r face. From here the coolant i s gathered, brought ex te rna l l y back t o j u s t downstream o f the en t ry po in t a t 1.5:l and f lows on down through the nozzle. This basic f low path was chosen because i t r e s u l t s i n the minimum cool ant pressure drop. More conventional ( in jec to r - th roa t -nozz le ) f l ow paths were analyzed, but resu l ted i n higher pressure drops by 10 t o 25%. The pressure drop i s minimized by the chosen f low path p r i m a r i l y because the coolant passes through the highest heat f l u x ( t h roa t ) reg ion wh i le a t the lowest bulk temperature. This has two advantageous e f fec ts ; i t al lows the tmpera - t u r e d i f f e r e n t i a l between the external wa l l and the hot wa l l t o be ma, imized i n the most c r i t i c a l cool ing region, and the lower bulk temperature r e s u l t s i n 2 higher 02 heat t r ans fe r coe f f i c i en t . This r e s u l t s i n lower coolant v e l o c i t y requirements and thus 1 ower pressure drop.

A l l three engine chamber designs were analyzed f o r both f u l l s l o t t e d chamber conf igurat ions and f o r tube bundles t o be attached a t optimized area ra t i os . The three primary tube mater ia ls considered were: Zr-Cu, A-286, and Inconel 718. S t ruc tu ra l analys is were used t o determine the required tube wal l thickness as a func t ion of the tube radius fo r the three mater ia l s mentioned a t t h e i r cyc le 1 i f e temperature 1 i m i t s . Zr-Cu was e l iminated because it requires a high wa l l th ickness t o radius r a t i o . Inconel 718 was chosen f o r use on the LOX cooled designs becuase i t has a higher maximum temperature l i m i t and lower thickness t o radius r a t i o . Because o f i t s i ncompa t i b i l i t y w i t h hydrogen, Inconel was no t used f o r the gas genera- t o r cyc le design, snd thus A-286 tubes were used.

The r e s u l t s o f the heat t r ans fe r analyses are summarized on Table X C I .

a. LOXJRP-1, Mode 1 Base1 i ne Engine, LOX Cooled

The approved base1 i ne Mode 1 engine analyzed during the Task V I e f f o r t u t i l i z e d a LOX/RP-1 prope l lan t combination as opposed t o the o r i g i n a l Task I 1 Base1 ine LOX/RJ-5 engine. Thi s propel 1 ant combination change resu l ted i n higher chamber heat f luxes than those o f the LOX/RJ-5 engine, and therefore higher coolant v e l o c i t y requirements and r e s u l t i n g pressure drop. The chosen operat ing condi t ions f o r the engine included: PC = 272 atm (4000 ps ia) , MR = 2.9, w i t h a coolant i n l e t temperature se t a t 11 1 OK (200°R). The design incorporates a z i rconium-copper s l o t t e d chamber design t o a nozzle area r a t i o o f 15:1 and a two pass Inconel 718 tube bundle t o the nozzle e x i t area r a t i o o f 40:l. The coolant f low path i s as fo l lows: coolant i n l e t a t nozzle area r a t i o = 1.5:l f lows t o the i n j e c t o r end, i s co l lected, brought ex te rna l l y back t o E = 1.5:1, re-enters chamber channels. f lows t o E = l 5 : l where tube bundle s ta r ts , f lows i n ha l f the tubes t o € e x i t = 40:1, turns and flows i n the other h a l f of the tubes back up L- c = 15:1 where i t i s co l lected. The resu l t an t t o t a l pressure drop fo r the chamber-tube bundle design was evaluated as 126.9 atm (1865 p s i ) .

Slo

tte

d C

ham

ber

AP,

atm

(p

si

Tube

Bun

dle AP,

atm

(p

si)

C

ombi

ned

Slo

tte

d &

Tu

be B

undl

e aP

, at

m (

ps

i)

TASL

E X

CI .

TASK

Y

I - C

OO

LANT

JA

CKE

T SU

mAR

Y (P

RE

SS

UR

E

DR

OP

)

-r -

Bas

e1 in

e M

ode

1 D

ual

~u

ei,

Oxy

gen

Coo

led

Mod

e 2

G

as

Gen

erat

or C

ycle

O

xyge

n C

ool e

d M

ode

1 LO

X/R

P- 1

LO

X/R

P- 1

LO

X/LH

, LO

X/R

P-1

Hyd

roge

n C

ool e

d

(CHA

MBE

R SE

CTI

ON

GEO

MET

RY )

LI

+.I

SLO

ITED

SEC

TIO

N

'T

R~

TUBU

LAR

SEC

TIO

N

7 N O e m - ? z 0 a m m m o* row - . m- r n ~ -0 o o

Rt

=

12.6

49

cm

(4.9

8

in.)

Nom

encl

ztur

e D

efin

ed o

n P

age

261

TABL

E X

CI.

-

COOL

ANT

JACK

ET

SUMM

ARY

(co

nt.

)

(H2/

02

GG C

YCLE

-

HYDR

OGEN

COO

LED

GLM

ETR

Y)

b . Dual-Fuel , LOX Cooled, (LOXIRP-1 , LO,Y/LH~) Engine

Previous analyses i nd i ca ted t h a t the Mode 2 operat ion heat t r a n s f e r cha rac te r i s t i c s d i c t a t e d the cool an t channel conf igurat ions. Therefore, the Mode 2 operat ing condi ti ons were used t o generate the chamber cool an t passage design parameters. A f t e r t he design was generated based on Mode 2 operat ion, the Mode 1 operat ion was analyzed t o determine the amount of al lowable coolant bypass f l ow which would meet the cyc le l i f e c r i t e r i o n and minimize the chanber pressure drop. This bypass f l ow r a t e was determined as approximately 40.8 kg/sec (90 Ib /sec) .

This design a l so incorporates a slot t .ed Zr-Cu chamber t o E = 15:1, and an Inconel two pass tube bundle t o € e x i t = 40:l. It a l so has the same coolant f l ow path as the basel ine design. Mainly, i n l e t ?t 1.5:1, t o i n j e c t o r , 1.5:l t o 1 5 ~ 1 , tube bundle t o 40: l and bac!- to 15:l. The r e s u l t a n t t o t a l pressure drop f o r the Mode 2 operat ion was 84 atm (1235 p s i ) , and f o r the Mode 1 operat ion, 132 atm (1940 p s i ) . The Mode 2 pressure drop i s s i g n i f i c a n t l y lower than t he Mode 1 pressure drop p r i m a r i l y because of the pressure e f f e c t s associated w i t h the 02 heat t r a n s f e r cor- r e l a t i o n . The lower coolant pressures o f the Mode 2 operat ion requ i re lower mass v e l o c i t i e s f o r t he same oxygen heac t r a n s f e r c o e f f i c i e n t as Mode 1 .

c. Hydrogen Cooled, Gas Generator Cycle

The chamber design cons is ts o f a Zr-Cu s l o t t e d chamber t o a nozzle arcs r a t i o o f 25:1, and a one pass A-286 tube bundle t o : ' ex i t = 42.7:l. The coolant f l ow path i s as fo l lows . coolant i n l e t a t nozzle area r a t i o o€ 1.5:1, f lows t o i n j e c t o r end, i s co l l ec ted and brought e x t e r n a l l y back t o 1.5:1, f lows downstream t o 25:l where tube bundle s t a r t s , and then on ou t t o the e x i t area r a t i o o f 42.7 :I , where i t i s .o l lec ted and brought t o the hydrogen-rich gas generator. The l a r g e bu lk temperature r i s e associated w i t h the 9 .O7 kg lsec (20 lbmlsec) hydrogen flow rate, r e s u l t s i n hydrogen e x i t temperatures near 81 1 O K (1000°F). These h igh bu lk temperatures i n the nozzle, preclude the use o f a two pass tube bundle because of the r e s u l t i n g h igh v e l o c i t i e s requ i red t o l i m i t the gas-side w a l l t we ra tu re . A o r~e pass tube bundle design was thus conceived wnich provides successively h igher cool an t v e l o c i t i e s as the coolant f lows through the nozzle t o the l a r g e r area r a t i o s and f i n a l l y t o the e x i t . The r e s u l t a n t t o t a l pressure drop f o r the chamber-tube bundle design was evaluated as 75.2 atm (1105 p s i ) .

6. MAIN INJECTOR DESIGN

The main t h r u s t chapber i n j e c t o r f o r the Mode 1 LOXIRP-1 basel ine and dual- fuel engines i s shown on F igure 83. The i n j e c t o r uses a coax ia l - type eiement which was se lected on the bas is o f Task I 1 r e s u l t s . An acoust i . -avi t y , whose s i ze has been estimated, i s shown as the combustion i n s t a b i 1 i tj upression device. The i n j e c t o r design parameters are summarized ori ~ l e X C I I .

TABL

E X

CII

. -

MODE

1 L

OX/

RP-

1 BA

SELI

NE

AND

DU

AL-F

UEL

EN

GIN

E M

AIN

,

Flo

w R

ate,

kg

lse

c (1

bls

ec)

MR

Type

of

Ele

men

t;

No.

o

f R

ings

No.

o

f E

lem

ents

Ele

men

t

Ele

men

t S

ize,

cm

(i

n.)

Ve

loci

ty,

m/s

ec

(ft/

se

c)

AP,

atm

(p

si)

Ma

nif

old

&

!nle

t

Ve

loci

t.y,

m

lsec

(f

tls

ec

)

AP,

atm

(p

si)

Inle

t d

ia,

cm

(in

.)

Inje

cto

r O.

D.

cm

(in

.)

Fue

l C

irc

uit

249.

3 (5

49.6

)

2.9

Co-

Axi

al

19

1254

,330

(.

130)

d

ia

Oxi

d.

Cir

cu

it

598.

4 (1

,319

.2)

.681

(.

268)

O.

D.

.465

(.

183)

I.

D.

173

(566

.5)

23.7

(3

48)

I f acous t i c c a v i t i e s a re no t requ i red t o s t a b i l i z e the gaseous p rope l l an t combustion process, t he design shown on F igure 84 i nd i ca tes t he s i z e decrease and p o t e n t i a l weight savings t h a t would r e s u l t . The coax ia l element i s d e t a i l e d on t h i s drawing, and i s based upon the Task I 1 t h r u s t per element recommendation. Th:, i n j e c t o r design i s s i r i l a r t o t h a t o f the main i n j e c t o r o f t he SSME, a1 though i t i s sho r t e r because the o x i d i z e r does no t have t o be vaporized.

The a l t e r n a t e Mode 1, gas generator cyc l e engine main l i q u i d / l i q u i d i n j e c t o r design i s shown on F igure 85. Th is i n j e c t o r uses quadle t - l i k e doublet s e l f imping ing elements. The element se l ec t i on i s based on recen t r e s u l t s o f ALRC 1 i q u i d / l i q u i d i r . j e c t i o n development f o r programs such as OMS, MX, and Improved Transtage (ITIP). A summary o f t he a l t e r n a t e Mcde 1 i n j e c t o r design parameters i s presented on Table X C I I I .

H. PREBURNER DESIGN

The Mode 1 base l ine engine uses both f u e l and o x i d i z e r - r i ch preburners . The LOX/RP-1 o x i d i z e r - r i ;h preburner desi g~ i s shown on F i gure 86 and t he f u e l - r i c h RP-1/LOX preburner design i s shown on F igure 87. The i n j e c t o r s use a p l a t e l e t design w i t h l i k e - o n - l i k e imping ing double t elements w i t h t he adjacent u n l i k e fans canted i n t o each o ther . Acoust ic c a v i t i e s are shown as the cornbuc t i o n i n s t a b i l i t y suppression devices a1 though in-depth s i z i n g o f the c a v i t y i s beyond t he scope of t he cu r ren t study. The preburner design parameters are presented on Tables X C I V and XCV,

For the dua l - fue l engine, the f u e l and o x i d i z e r - r i c h LOX/R3-1 preburners a re the same as those showrl f o r the base l ine Mode 1 engine. The hydrogen/oxygen p r e b ~ r n e r designs are s imi l a r 1.2 these and are shown on Figures 88 and 89. The preburner design parameters are presented on Tables X C V I and X C V I I.

The f w l - r i c h gas generator design f o r the a l t e r n a t e Mode 1 engine i s shown on F igure 90. This i n j e c t o r design i s based upon the design developed du r i ns the M-1 engine program (Ref. 76) . The i n j e c t o r uses a codx ia l element which i n j e c t s gaseous hydrogen and l i q u i d oxygen and achieved a 98% combustion e f f i c i e n c y dur ing M-1 t e s t i n g . M-1 t e s t i n g was tonducted over a mix ture r a t i o range f rom 0.6 t o 1.0 and a chamber pressure range of 51 t o 77.9 atm (750 t o 114.5 ps ia ) .

P r i n c i p a l element dimensions f o r the design shown on Figure 90 are:

Ox id ize r Flow Diameter = 0.31 cm (.12ZH) Fuel S l o t Width = 0.208 cm ( .082") Element Wall Thickness = 0.184 cni ( .0725"j

A summary o f the preburner design parameters i s shown on Table X C V I I I .

END

OF

VALV

E3

PLAT

ELET

SU

PPOR

T -F

ACE

PLAT

E

ELEM

ENT

DET

AIL

131 2

ON

19 R

OWS

COOL

ANT

OUT

LET

Fig

ure

84.

M

ode

1 LO

X/RP

-1

Bas

elin

e B

acku

p M

ain

Inje

cto

r

TABL

E X

CII

I. -

ALTE

RN

ATE

MOD

E 1

GAS

GEN

ERAT

OR

CYCL

E EN

GIN

E M

AIN

IN

JEC

TOR

DES

IGN

DAT

A

Flo

w R

ate,

kg

/sec

(I

bls

ec)

MR

Type

of

Ele

men

t

No.

o

f R

ows

No.

o

f E

lem

ents

No.

of

Cha

nnel

s

Ori

fic

e s

ize

, cm

(i

n.)

Inje

cti

on

Ve

loci

ty,

mls

ec

(ftl

se

c)

Ori

fic

e A

P,

atm

(p

si)

Inle

t V

elo

city

, m

lsec

(f

t/s

ec

)

Ma

nif

old

Ve

loci

ty,

m/s

ec

(ft/

se

c)

Ma

nif

old

AP,

at

m (

ps

i)

Inle

t D

ia,

cm

(in

.)

Inje

cto

r O.

D.,

cm

(in

.)

Fu

el

Cir

cu

it

Oxi

d.

Cir

cu

it

21 6

(4

76.0

) 62

1 (1

368.

4)

2.88

Qua

d1 et-

Li k

e D

ou

ble

t S

el f

Imp

ing

ing

14

1464

14

64

14

14

.lo

9

( -0

43

) .I

91

( .

075)

71.3

(2

34)

72.8

(2

39)

34.0

(5

00)

34.0

(5

00)

15.8

(5

1.8)

18

.6

(60.

9)

5.5

(18

) 5.

5 (1

8)

6.8

(100

) 5.

2 (7

7.0)

12.7

(5

.0)

20.3

(8

.0)

63.5

(2

5.0)

'I- O U

E& A0 -

TABL

E XC

IV.

- OX

IDIZ

ER

-RIC

H P

REBU

RNER

DES

IGN

DAT

A SU

MM

ARY,

LOX

IRP

-1

BASE

LIN

E

Flo

w R

ate,

kg

/sec

[i

b/s

ec)

MR Type

of

Ele

men

ts

No.

o

f R

ings

No.

of

Ele

men

ts

Ele

men

t

Ele

men

t S

ize,

cm

(i

n.)

Ve

loci

ty,

m/s

ec

(ftl

se

c)

AP,

atm

(p

si )

Ma

nif

old

Ve

loci

ty,

mls

ec

(ftl

se

c)

AP,

atm

(p

si)

Inle

t d

ia,

cm

(in

.)

Inje

cto

r 0

.D.,

cm

(in

. )

Fue

l C

irc

uit

O

xi d.

C

irc

uit

RP

- 1

LOX

(.01

5 x

.038

) S

lot

(214

)

(900

(.13

0)

Squ

are

(1 55

.8)

(500

)

TABL

E XC

V.

- FU

EL-

RIC

H

PREB

URNE

R D

ESIG

N D

ATA

SUEM

ARY,

LO

X/R

P-1

B

AS

ELI

NE

Fue

l C

irc

uit

O

xi d

. C

irc

uit

R

P-1

LO

X

Flo

w R

ate,

kg

/sec

(1

b/s

ec)

20

4.3

(450

.5)

45.0

(9

9.1

)

Mi?

0.22

0.

22

Typ

e c

,f E

lem

ents

V

-Dou

bl e

t X

-Dou

blet

No.

3f

Rin

gs

7 6

No.

.>

f Ele

men

ts

972

486

Ele

men

t S

ize,

cm

(i

n.)

.I

73

( .

068)

S

quar

e .0

51

x .3

?5

(.02

0 x

.132

) S

lot

Ye

loci

ty,

m/s

ec

(ft/

se

c)

65.2

(2

14)

47.5

(1

55.8

)

AP,

atm

(p

si)

61

.2

(900

) 34

.0

' 500

)

Ma

nif

old

Ve

loci

ty ,

m/s

ec (f t

/se

c)

10.7

( 3

51

10.7

(3

5)

AP,

atm

(p

si)

1.

36

(20

) 1.

36

(20)

Inle

t D

ia,

cm (

in.)

15

.2

(6 0

) 8.

9 (3

.5)

Inje

cto

r O

.D.,

cm

(in

.)

36.5

(1

4.36

) 36

.5

(14.

36)

TABL

E XC

VI . -

02/H

2 OX

1 DIZ

ER-R

ICH

PRE

EURN

ER D

ESIG

N SU

MMAR

Y, DU

AL-F

UEL

Fla

r R

ate,

kg

/sec

(l

bls

ec)

MR

Ty

pe o

f E

lem

ent

No.

of

Rin

gs

No.

o

f E

lem

ents

Ele

men

t

Ele

men

t S

ize,

cm

{i

n.)

V

elo

city

, N

se

c (

f tls

ec)

AP

, at

m (

psi

)

Man

ifol

d

Ve

loci

ty , m

lsec

( ft

lse

c)

AP,

atm

(p

si)

Inle

t D

ia,

cm

(in.

) In

ject

or

O.D.

, cm

(i

n.)

Fue

l C

irc

uit

Jx

l d . C

l rcu

i t

"2

O2-

- -.

3.54

(7

.8)

389.

3 (3

58.3

) 11

0 1 'I 0

X

-Dou

bl e

t V

-hk:

le

t 7

8

642

1284

,038

x

,102

(.0

15

x ,0

40)

Slo

t .2

95

(.11

6)

Squa

re

322

(1 05

7)

43.3

(1

42)

36.9

(5

43)

20.1

(2

96)

-18.

1 C

III

(7.1

2 in

.)

,

.12

in.)

14.0

cm

Figu

re 9

0.

Alt

ern

ate

Mod

e 1

Co-

Axi

al

Gas

Gen

erat

or

TABL

E XC

VI I I. -

ALTE

RNAT

E MO

DE

1 C

O-A

XIA

L GA

S GE

NERA

TOR

DES

IGN

DAT

A SU

mAR

Y

Fuel

Cir

cu

it

Oxi

d.

Ci r

cu

i t

"2

O2

Flow

R~

G, kg

/sec

(1

bls

ec)

WR

Type

of

Ele

men

t

No.

of

Rin

gs

El e

aen

t

Ele

men

t S

ize,

an (

in.)

Co-

Axi

al

1.09

5 (.

431)

O

.D.

0.31

(.

122)

D

ia

.678

(-

267)

I.

D.

VALVE SELECTION AND S I Z I N G

Based upon the controls i den t i f i ed by the s t a r t and shutdown sequence analyses, a valve s i z ing study was performed. These components are shutoff valves o f varying s ize and type. Considering the parameters involved, the majori ty of the valves selected were the poppet type. These valves are actuated from ful l -c losed t o ful l-cpen posi t ion i n the presence o f tank head pressure during engine s t a r t and are closed during engine shutdown once valve i n l e t pressure has decayed t o a reduced leve l .

TranslaLing sleeve valves were selected f o r operation i n the high pressure environment associated wi th valve closure t o i n i t i a t e the engine shutdown. A sleeve valve confi gurat i on can be v i r t ua l l y pressure force balanced by matching the diameter o f the sleeve seal w i th tha t o f the shutoff seal which resu l ts i n a reduction i n power requirements.

The estimated pressure drop and weight f low requirements f o r each valve were defined and converted t o a f l u i d Kw requirement.* An array o f h i s t o r i ca l data regarding the use o f LOX, LH2, and RP-1 valves on past engine programs was 1 i kewi se arranged as a function o f Kw and provided the basis f o r the valve diameter, weight, and envelope estimates shown i n Tables XCVIX, C, and C I .

The dual-fuel engine cycle exhibit: the need f o r preventing the backflow o f main combustion gases i n t o the nonoperatinq fue l - r ich preburner discharge c i r c u i t . An approach t o preventing t h i s backflow i s t o provide check valves i n the fue l - r ich preburner c i r cu i t s , These large, ~ 1 5 . 2 cm (d"), check valves are required t o operate a t high pressures, 272 atm (4000 psia), i n hot gases t ha t are the product of LOX/RP-1 and LOX/LH2 combustion, The weight and djmensions for a 15.2 an (6") check valve were estimated from h i s to r i ca l data and are l i s t e d on Table C I I .

1. Valve Materials

The basic valve materials f o r the valves were deflned which are compatible wi th the propellants designated f o r these engine cycles. Specif ic considerations are e h r i tt lement , metal 1 urgical sta- b i 1 i ty, and chemical compat ib i l i ty . The basic valve materials are 1 i s ted as follows:

Part - Materi a1 Lox - LH2 -

Valve Body Inconel 718 Inconel 718 A-286 CRES Seat Inconel 718 Inconel 718 Inconel 718 Shutoff Seal Phosphor Bronze Phosphor Bronze Beryl1 ium Nickel Spring Inconel 750 Inconel 750 Inconel 35'3 Sleeve ' Inconel 718 A-286 CRES A-286 CRES Be1 lows Inconel 718 Inconel 718 Inconel 718

% = Weight flow/(pressure drop x specif ic gravi ty) l /2

Val

ve

Type

Mai

n LO

X V

alve

S

huto

ff (

We

1)

Mai

n R

P-1

Val

ve

Shu

toff

Mai

n LH

2 V

alve

S

huto

ff

Ox

Ric

h P

rebu

rner

(0

2/R

D-1

LO

X V

alve

S

hu

toff

RP-

1 V

alve

S

huto

ff

Fue

l R

ich

Pre

burn

er

(RP

-1/0

2)

LOX

Val

ve

Shu

toff

Ox

Ric

h P

rebu

rner

(0

2/H

2)

LOX

Val

ve

Sh

uto

ff

LH2

Val

ve

Sh

uto

ff

Fue

l R

ich

Pre

burn

er

(H2/

02)

LOX

Val

ve

Shu

toff

TABL

E C

I. -

VALV

E S

IZIN

G:

DU

AL-F

UEL

EN

GIN

E (M

ODE

1 :

LOX/

RP-

1 )

(MO

DE

2:

LOX

ILH

~)

App

rox .

va1 v

e D

iam

eter

C

on

fig

ura

tio

n R

equi

rem

ent

cm (

in.)

Pop

pet

Pop

pet

Pop

pet

Pop

pet

Sle

eve

Pop

pet

Pop

pet

Sle

eve

Pop

pet

A~

~ro

x.

Val

ve

App

rox.

V

alve

L

.- -

h' A

ctu

ato

r A

ctua

tor

Env

elop

e W

ei h

t D

i am

eter

en

t

*+

cm

(in

.)

L ti:-

)

TABL

E C

II.

- HOT

GA

S CH

ECK

VALV

E W

EIG

HT

AND

ENVE

LOPE

EN

GIN

E CY

CLE

- D

UAL

-FU

EL,

OXYG

EN

COOL

ED

Val

ve

bd

e 1

- F

uel

Ric

h P

rebu

rner

M

scha

rge

Cir

cu

it

Mod

e 2

- Fue

l R

ich

h-e

burn

er

Dis

char

ge

Ci r

cu

i t

App

rox.

V

alve

- bi

amet

er

Type

C

on

fig

ura

tio

n

an

(in

.)

Che

ck

Pop

pet

15.2

(6

)

Che

ck

Pop

pet

15.2

(6

)

App

rox.

V

alve

A

p~

rox

. Val

ve

Env

elop

e D

iam

eter

Le

ngth

cm

(i

n.

j cm

(i

n.)

2. Valve Actuation Forces

The forces t o be overcome include pressure forces, flow forces, f r i c t i o n forces and i n e r t i a l forces o f moving parts. The i n e r t i a l forces become s i g n i f i c a n t f o r the t yp i ca l actuat ion time response requirements associated w i th engine s t a r t and shutdown. A representat ive se t of actuator loads were obtained by comparison t o the ALRC Space Shut t le Main Engine (AJ-550) study.

The actuat ion system should be capable o f producing the fo l lowing approximate 1 inear forces f o r the fo l lowing valves.

14,230N (3200 Pounds)

Main LOX Valves Main RP-1 Valves LOX Valves - Ox Rich ?reburners LOX Valves - Fuel Rich Preburners

7560N (1700 Pounds)

Main LH2 Valves RP-1 and LH2 Valves - Ox Rich Preburners LOX GGV

3. Valve Actuation Methods

Electromechani ca l valve actuat ion has been selected f o r the valves considered i n t h i s study. This select ion i s based upon a trade study conducted ea r l y i n the ALRC SSME program t o select the method o f engine contro l valve actuation. E l e c t r i c a l , hydraul ic and pneumatic systems were evaluated on the basis o f seventeen design considerations ; the primary factors being weight, contamination susceptibi 1 i ty, power re - quirements, fabr icat ion cost and lead time, main ta inab i l i t y , r e l i a b i l i ty and safety. The resu l ts o f t h i s study i t ld icated the three systems were r e l a t i v e l y equal i n t h e i r a b i l i t y t o sa t is fy the overa l l design requirements.

Due t o the close resu l t s o f the i n i t i a l trade study, a more extensive review was conducted i n which greater consideration was given t o vehicle in tegra t ion and system maintenance. This review c l e a r l y revealed the advantages o f the e l e c t r i c a l system which can be sumnarized as fol lowc :

Single Interface: The e l e c t r i c a l system requi res only an e l e c t r i c a l in ter face, an i tem also required by hydraul ic and pneumatic systems i n addi t ion t o t h e i r actuat ion f l u i d interfaces.

R e l i a b i l i t : Due t o i t s s ing le in ter face, the e l e c t r i - -Tl--+ cal system represents e eas complex system. The la rge nunber o f 1 ines

associ atsd w i th conveying the hydraul ic o r pneumatic f l u j d s throughout the

engine are deleted along wi th t h e i r numerous leak paths. A1 though the e lec t r i ca l sys t ~ m requl res more e lec t r i ca l control components, the power levels are su f f i c i en t l y low t o permit the use o f wel l developed switches and control components. A completely redundant e lec t r i ca l sys tem can be provided wi th fewer components than the other systems and a s ingle po in t f a i l u r e i s isolated without i t s a f fec t ing other components, a condit ion not necessarily t rue i n the case o f a hydraulic o r pneumatic component leakage fa i lure.

Contamination Free: The e lec t r i ca l system i s not exposed t o the contamination problems associated wi th hydraul ic and pneumatic systems, problems which are a primary cause o f hydraul ic and pneumatic system fai lures. This i s more sf gn i f i can t as the number o f components i n an overal l system i s increased as, f o r example, i s the case when using the vehicle hydraul ic system for a power source.

' Ease o f Maintenance: The e lec t r i ca l system i s easier t o maintain. No hydraul ic o r pneumatic f l u i d leakage inspection o r clean- 1 i ness checkouts are required. Basic component checkout i s easier since i tems such as hydraulic system bleed-in are deleted. The e lec t r i ca l sys tem a1 so provides f o r easier overal l engine system mat ntenance through the delet ion of f l u i d l i nes which could in te r fe re wi th engine component i ns ta l 1 a t i on o r removal .

' Safet : A double malfunction consist ing of hydraul ic f l u i d and propel lant d ea age could resu l t i n loss o f a hydraul ic system due t o freezing or a potent ia l f i r e hazard.

' Elect r ica l Power Required: The e lec t r i ca l system requires more e l e c t r E a l power than the hydraul i c o r ~ n ~ u m a t i c sustems . however, the t o ta l energy' from the vehicle power source is not s i g n i f i l cantly dif ferent. The overal l power e f f ic iency i s higher f o r the e l ec t r i - cal system.

The estimated power requirements are shown on Table C I I I .

J. MATERIALS SELECTION

The materials selected f o r the major engine components are l i s t e d on Table C I V . These materials were selected t o achieve l ightweight engir.es wi th consideration o f the design and long l i f e requirements and the environmental and propel 1 ant compati b i 1 i t y aspects.

TABLE C I I I - ESTIMATED POWER REQUIREMENTS FOR ELECTROMECHANICAL VALVE ACTUATION

Val ve

Main LOX Valve

Main RP-1 Valve

LOX Valve - Ox Rich Preburner

LOX Valve - Fuel Rich Preburner

Main LH1 Valve

RP-1 Valve - Ox Rich Preburner

LH2 Valve - Ox Rich Preburner

LOX GGV

Estimated Nomi nal Openi ng

Actuation Time ( s e c l

1.5

0.4

1.4

Estimated Average Opening

Power (Watts )

TABLE C I V . - MATERIALS SELECTION

Base1 i ne A1 ternate Component LOX/RP- 1 Dual -Fuel Mode i

1. Low Speed LOX TPA a. Shaft Inconel 718 Same Same b. Impeller & Turbine 7075 T-37 Same Same

A1 mi num A1 1 oy c. Housing A356T-6 Same Same

A1 Al loy d. Bolts A-286 Same Same e. Housing Liner FEP Teflon Same Same

Fused Coating f. Bearings CRES 440C; Same Same

A1 ternate Haynes Star A1 l oy PM

2. Low Speed RP-1 TPA A1 1 materials the same as low speed LOX TPA except Teflon Coating i s not required.

3. Luw Speed LH2 TPA NotApplicable A l l m a t e r i a l s t h e s a m e a s l o w speed LOX TPA except Teflon Coating i s not required.

4. Wgh Speed LOX TPA

Shaft Impel 1 e r High Pressure Pump & Turbine Housing Inducer Housing Turbine Bol ts (pump) Bol ts Turbine Bearings

A-286 Inconel 718 ARM0 Nitronic-50 Inconel 718 Incanel 718

A-286 Waspal oy CRES 440C o r A1 ternate

Same Inconel 718 Same

Same Inconel 718 Same Same Same

Same Inconel 718 Same

Same UDIMET 630

S a m same Same

5. High Speed RP-1 TPA i

l

a. Inducer Housing 5AL-2.5 SnE1 i Same Same i i

Titanium Al loy j

A1 1 other material the same as High Speed LOX TPA i i

TABLE C I V (cont. )

Base1 ine Component LOXIRP-1

A1 ternate Dual -Fuel - Mode 1

6. High Speed LH2 TPA a. Inducer Housing N/A* 5AL-2.5 Sn E l i

Titanium Alloy Same as Dual -Fuel Same b. High Pressure Pump N/A*

Housing c. Turbine N/A* d. Impeller N/A* e. Turbine Housing N/A*

5AL-2.5 SnEl i Ti tan i um A1 1 oy UDIMET 630 Same

Same Same ARMCO

Nitronic-50 f. Shaft g. Bolts (pump)

h. Bolts (turbine)

Same Same Same CRES 440C

Waspal oy

CRES 440C i. Bearings 7, LOX/RP-1 OX-Rich

Preburner a. In jec tor Body b, Chamber

8. RP-1/LOX Fuel-Rich Preburner

ARMCO N i tronic-50 ARMCO Nitronic-50

Same N/A*

Same N/A*

a. In jec tor Body ARMCO Nitronic-53 Inconel 625

Same N/A* Same N/ A* b. Chamber

9. LOXILH2 Ox-Rich PreSurner a. In jec tor Body ARMCO N i t ronic- NIA*

50

b. Chamber N/A* ARMCO N i t ron ic- NIA* 50

10. LOX/LHz Fuel-Rich Preburner o r Gas Generator a. In jec tor Body &

Chamber N/A* AWO Nitronic- Same 50

TABLE C I V (cont.)

Basel i ne Cornponen t LOX/RP-1 Dual -Fuel

11 Thrust Chamber Injector

a. Body Inconel 625 Same b. k n i f o l d s CRES 347 same c. Injector Face Inconel 625 Saae

12. Combustion Chantber ZR Cu ZR Cu 13. Tubes Inconel 718 Inconel 718 14. Nozzle Extension N/A* Col unbium 15. Hot Gas Manifold ARK0 Nitronic-50 Same

A1 temate Mode 1

3 ~ o t Appl i cab J e .

SECTION I X

CONCLUSIONS AND RECOMMENDATIONS

A. CONCLUS ! O M

The conclusions which have been derived from the resul ts o f t h i s advanced high pressure engine study are presented on Table C I V f o r easy reference. These conclusions cover a1 1 study tasks . '

It should be noted that, because candidates are not recommended as coolants, t h i s does not necessarily mean tha t the propel lant combination i s eliminated. LOX cooling would be feasible f o r any o f the study prope l lant c d i n a t i o n s . For example, RP-1 was eliminated as the coolant but selected as the baseline fue l i n a LOX cooled engine. Operational problems my preclude the use o f MMH and N2Hq i n any case.

The dual-fuel engine nozzle i s very long and heavy because the contour was performance optimized. That i s , a two posi t ion nozzle contour which resulted i n performance almost equal t o the base1 ine Mode 1 and Mode 2 engines was selected. The nozzle length and engine weight could be s ign i f i can t l y reduced i f some performance penalty i s acceptable.

B. RECOMMENDAT IONS

The recomnendations f o r technology and fur ther study tha t were iden t i f i ed as a resu l t o f t h i s high pressure engine study are tabulated on Table CV.

o

f 4

CAND

I DAT

E

o B

asel

ine

LOX/

RJ-

5,

LOX

Coo

led

O

LOX/

RJ,

RJ-

5 C

oole

d

o LO

X/R

P-1 ,

RP

-I C

oole

d

o LO

X/M

, M

Coo

led

o LO

X/W

2H4,

N2H4

C

oole

d

o LO

X/CH

4, CH

4 C

oole

d

TABL

E CV

. --

CONC

LUSI

ONS

CONC

LUSI

ONS

FEAS

IBLE

CAN

DIDA

TE

Max

PC

bas

ed u

pon

pow

er b

alan

ce, -, 28

6 at

m (

4200

psi

a)

wit

h 2

0% m

argi

n.

UN

FEAS

IBLE

Can

dida

te f

or

reus

able

, h

igh

pre

ssur

e en

gine

M

ax P

c~1

36

atm

(20

00 p

sia

) ; li

mit

ed

by

coki

ng i

n c

oo

lan

t ch

anne

ls.

UN

FEAS

IBLE

Can

dida

te f

or

reus

able

, h

igh

pre

ssur

e en

gine

. M

ax P

C <

136

atm

(20

00 p

sia)

; lir

nf t

ed

by

coki

ng

in

co

ola

nt

chan

nels

.

NOT

REC

OM

ND

ED

Max

PC

base

d up

on p

ower

bal

ance

s 30

6 at

m (

4500

psi

a)

wit

h 2

0% m

argi

n.

MMH

inco

mpa

tible

wit

h Z

rCu

for

lon

g t

erm

use

. F

uel-R

ich

Pre

burn

er T

emp.

s

1333

'~ (

24

00

~~

) cre

ate

s tu

rbin

e I

lfe

or

desi

gn p

robl

em.

NOT

RECO

MM

ENDE

D.

Max

PC

> 3

40 a

tm (

5000

psi

a);

ba

sed

upon

pow

er b

alan

ce w

ith

20%

mar

gin.

N2Hq

in

com

patib

le w

i M Z

rCu

for

lon

g t

ern

use

. F

ue

l-ri

ch P

rebu

rner

Tem

p. *

10

00

~~

(1

80

0~

~) cre

ate

s tu

rbin

e l

ife

or

desi

gn p

robl

em.

N2H4

va

pors

cre

ate

exp

l or i ve

haz

ard.

FEAS

IBLE

CAN

DI D

ATE

Max

PC

s 27

2 at

m (

4000

psi

a)

base

d up

on p

ower

bal

ance

wit

h 2

0% m

argi

n

De

li ve

red IS

ap

prox

imat

ely

13 s

ecs

gre

ate

r th

an b

ase1

ine

. E

ngin

e w

eigh

t 24

8 kg

(54

6 I b) h

ea

vie

r th

an b

asel

ine.

CAN

DID

ATE

o P

ara

lle

l B

urn,

H

ydro

gen

Coo

led,

LO

XIR

J-5

o H

ydro

gen

Coo

led,

G

as-

Gen

erat

or C

ycle

(L

OX

/RJ-

5)

o A

uxi 1

i ary

Coo

l an

ts

o B

ase

line

Fue

l :

RJ-

5 o

r R

P-1

TABL

E CV

(c

on

t . )

CONC

LUSI

ONS

CONC

LUSI

ONS

VIA

BLE

CAN

DID

ATE

BUT

NOT

RECO

MM

ENDE

D

Max

. PC

Q 34

0 at

m (

5000

psi

a)

base

d up

on p

ower

ba

lan

ce w

ith

20%

mar

gin.

Op

era

tio

na

lly

1 im

i te

d t

o o

nly

par

a1 le

l b

urn

.

Ope

rat

Gim

bal

VIA

BLE

Max

P

C io

n a

ffe

cts

Mod

e 2

engi

ne d

evel

opm

ent.

po

ten

ti a

lly

un

fea

sib

le

CAN

DID

ATE:

RE

COM

MEN

DED

FOR

FURT

HER

STUD

Y

Pag

e 2

of

4 C

c

i

> 34

0 at

m (

5000

psi

a)

base

d up

on c

oo

lan

t an

d po

wer

ba

lan

ce

anal

yses

.

Co

ola

nt

flo

w r

ate

% 9

.O7

kgls

ec

(20

1 bls

ec)

re

qu

ire

s in

cre

ase

d

LH2

tan

k ca

paci

ty .

Li g

hte

r en

gine

wei

gh t ,

327

kg (

722

1 b) ,

com

pare

d to

bas

e1 in

e f

or

eq

uiv

ale

nt Is.

Pre

l im

ina

ry t

rad

e s

tud

ies

show

p

ote

nti

a1 p

ayo

ff c

ompa

red

to b

ase

line

.

NOT

RECO

MM

ENDE

D

WAT

ER

BEST

C

AND

IDAT

E

No

ga

ins

wit

h N

aK

Li t

hi u

rn c

orr

os

i vi t

y a

nd o

pe

rati

on

al

prob

lem

s

A1 1

syst

ems

he

avi

er

tha

n b

ase

lin

e w

ith

eq

uiv

ale

nt Is

Hig

he

r PC

po

ten

ti a1

th

an

bas

e1 in

e .

Incr

ea

sed

sys

tem

co

mp

lexi

ty

RP-

1 RE

COM

MEN

DED

Unk

now

n co

st p

roje

cti

on

fo

r R

J-5

CAN

DI D

ATE

Boo

st P

ump

Dri

ve k

tho

d

o H

ydra

ulic

Tur

bine

o

Gas

T

urbi

ne

o G

ear

Driv

-n

Pre

l im

i nar

y D

esig

ns

o M

ode

1 LO

X/R

P-1

CONC

LUSI

ONS

CONC

LUSI

ONS

o S

elec

ted

Con

cept

: S

impl

e co

ntr

ol

and

lig

ht

wei

ght

o N

ot R

ecom

nend

ed:

(1)

Req

uire

s h

ot

gas

con

tro

l va

lves

, m

u1 ti -

stag

e tu

rbin

e a

nd h

igh

pre

ssur

e h

ot

gas

hous

ings

(2

) V

ery

heav

y

o N

ot R

ecom

nend

ed:

(1)

Re

stri

cts

engi

ne p

acka

ging

. (2

) LO

X co

oled

gea

rs c

onsi

dere

d u

nfe

asi

ble

o P

rom

isin

g C

once

pt:

(1)

Eng

ine

Wei

ght

= 21

12 k

g (

4657

lb

)

(3)

(4)

(5)

o A

lte

rna

te M

ode

1 (G

as

o P

rom

isin

g C

once

pt:

(1)

Gen

erat

or C

ycle

) E

ngin

e (2

)

Eng

ine Is

= 32

3.6

secs

(se

a-le

vel )

= 35

0.6

secs

(va

cuum

) E

ngin

e Le

ngth

= 2

77 c

m

(109

in

)

Req

ui re

s ad

vanc

ed t

echn

o1 og

y e

ffo

rt

Po

ten

tia

lly

sf m

ple

con

tro

l sy

stem

E

ngin

e W

eigh

t =

1758

kg

(393

5 lb

) E

ngin

e Is

= 32

3.5

secs

(se

a-l

eve

l)

= 35

0.7

secs

(v

acuu

m)

Eng

ine

Leng

th =

315

an

(124

in

) M

ost

tech

nolo

gy h

as b

een

dem

onst

rate

d

Add

i ti o

nal

engi

ne c

ompl

exity

h h h L n W h Y Y Y

TABL

E C

VI .

- TE

CHNO

LOGY

RE

COM

MEN

DATIO

NS

Ite

m

-

Tec

hnol

ogy

Sug

gest

i on

1.

Pro

pe

llan

ts

o E

xten

d in

vest

iga

tio

n t

o in

clu

de

oth

er

fue

ls.

o V

eri

fic

ati

on

of

prop

el 1

an

t p

rop

ert

ies

at

hig

h p

ress

ure

2.

Com

bust

ion

o LO

X/H

ydro

carb

on P

rebu

rner

, F

uel -

Ric

h

o LO

X/H

ydro

carb

on P

rebu

rner

, O

x-R

i ch

o LO

X/H

ydro

carb

on T

CA

3.

Per

form

ance

o

Thr

ust

chan

ber

perf

orm

ance

gas

/gas

in

jec

tio

n

4. TC

A H

eat

Tra

nsfe

r o

LOX

reg

en

era

tive

ly c

oole

d TC

A fo

r LO

X/H

ydro

carb

on p

rop

el 1

an

t

o LO

X re

ge

ne

rati

vely

coo

led

TCA

for

LOX/

H2

Pro

pel 1

ants

5. T

hrus

t C

ha

he

r o

LOX

reg

en

era

tive

ly c

oole

d TC

A fa

ilu

re a

na

lysi

s

6.

Turb

omac

hi n

ery

o P

rope

l 1 an

t co

oled

bea

ring

s

o TP

A S

eal

Eva

lua

tio

n

Ob

ject

ive

s D

efin

e lo

w c

ost

syn

the

tic

fue

l fo

r la

rge

qu

an

tity

pr

oduc

tion.

E

xten

d ex

peri

men

tal

da

ta t

o 4

08 a

tm (

6000

ps

i a).

Eva

luat

e ca

rbon

for

mat

ion,

d

ep

osi

tio

n a

nd

com

bust

ion

pro

pe

rtie

s a

t h

igh

pre

ssur

e.

Inve

stig

ate

po

ten

tia

l MR

dis

trib

uti

on

pro

blem

s an

d de

fine

com

bust

ion

pro

pe

rtie

s a

t h

igh

pre

ssur

e.

Ve

rifi

cati

on

of

pro

pe

lla

nt

pro

pe

rtie

s a

t h

igh

pr

essu

re,

408

atm

(60

00 p

sia

) . Ca

rbon

de

po

siti

on

e

valu

ati

on

. V

eri

fic

ati

on

of

chan

ber

len

gth

re

qu

ire

mn

ts,

perf

orm

ance

le

vels

and

sta

bi 1

i ty

of

gas/

gas

inje

cti

on

at

hig

h p

ress

ure.

LOX

he

at

tra

nsf

er

heat

ed t

ube

test

lng

. Dem

- s

tra

tio

n o

f h

ea

t fl

ux

ca

pa

bil

ity

in Z

rCu

thru

st

chan

ber

and

ma

teri

al

con

pa

tab

ili ty

.

Sam

e as

abo

ve.

Eva

lua

tio

n c

f LO

X le

akag

e on T

CA

ho

t w

all

sid

e

wit

ho

ut

film

co

olin

g.

Est

ab

lish

pro

1

lan

t lu

bri

cate

d r

oll

ing

co

nta

ct

be

ari

ng

1 i fe

- != oad

ch

ara

cte

rist

i cs.

D

eter

min

e TP

A se

al n

ea

r an

d 1

i fe

dat

a.

I tern

-

7.

Sys

tem

Ana

lysi

s o

TABL

E C

VI

(con

t. )

TECH

NOLO

GY

RECO

MM

ENDA

TIONS

Page

2 o

f 2

Tech

no1 o

gy S

ugge

stio

n O

bjec

tives

E

xten

d o

f pa

ram

etri

c ra

nge

to

6.67

MN

o

Obt

ain

use

ful

da

ta f

or

heav

y li

ft v

ehf c

le

(1.5

M

il.

lbs)

th

rus

t en

gine

s m

issi

on s

tud

ies

LOXI

CHI

Eng

ine

Pre

l lm

i na

ry D

esig

n o

Obt

ain

ba

selin

e e

ngin

e d

ata

fo

r U

)X/M

etha

ne

engi

nes

Eva

lua

tio

n o

f co

st e

ffe

cti

ve

o

Incr

ease

in

pay

load

thr

ough

eng

ine

wei

ght

chan

ter

life

for

SSTO

m

issi

on

red

uct

ion

De

fin

itio

n o

f A

dvan

ced

Hig

h P

ress

ure

o D

efin

e de

velo

pnen

t le

ad

tim

s,

hard

war

e an

d E

ngin

e de

vel o

pmen

t an

d p

rod

uct

ion

fu

ndin

g ra

te r

eq

ui r

elae

nts

plan

s an

d co

st

REFERENCES

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Roder , H. M. and Weber, L. A. , WRDI 0 ~ n Technology Survey: Vol urn I, Thermophysi cat Properties , NASA SP - , ~ a t f o n a l Aeronautics and Space Administration, Washington, D. C., 1972.

Weber , L. A., Extrapol a t i on o f Thermthysi ca l Propert i es Data f o r 5000 t o 10, 00 psla) a t Low Tenperatures 8 , 121, National Bureau o+ Standards,

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McCarty , R. D. and Weber , L . A., Thermophysi ca; Properties o f Para- h dro en from the Freezing L i u i d Line t o '5600-R t o r Pressures t o

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USAF Propel 1 ant Handbook-Hydrazine Fuel s-Vol ume 1 , AFRPL-TR-69-149, Be1 1 Aerospace Co . , Buf fa lo , New York , Contract 70461 1 -69-C-005, March 1970.

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