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CLASS Z ENGINE INFORMATION

A STUDY OF

ADVANCED

COMPOSITE PROPULSION SYSTEMS

FOR

LAUNCH VEHICLE APPLICATIONS I._.

VOLUME SEVEN

Report 25,194

C ontract NAS7- 377

The Marquardt Corporation

Van Nuys, California

September 1966

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FOREWORD

This report constitutes a portion of the final report documentation

under National Aeronautics and Space Administration Contract NAS7-377.

This and two companion reports (Refs. 1 & 2) present general engine

data derived from the study which are organized to facilitate their

incorporation into concurrent and subsequent advanced systems studies.

Covered here are the Class 2 Engines (_o in number) studied in the

program's concluding phase. More complete study results relating to these

engine concepts, including such areas as subsystem design trade-off

studies and overall vehicle/mission analyses, are given in the main body

of the project report (Reference 3).

The present Volume is one of seven in the total published study documen-

tation. Its orientation in the report sequence is shown below:

Volume i Summary Report

Volume 2 Main Technical Report, Part i

Volume 3 Main Technical Report, Part 2

Volume 4 Class 0 Fact Sheets, Part i

Volume 5 Class 0 Fact Sheets, Part 2

Volume 6 Class I Engine Information

Volume 7 Class 2 Engine Information

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This report results from the contributions of the following

Marquardt Science and Technology Group personnel under

National Aeronautics and Space Administration Contract

NAS7-377:

W, J. D. Escher

B. J. Flornes

W. R. Hammill

J. J. Kuhlmeler

C. R. Orr

B. A. Otsap

A. F. Truitt

Acknowledgement is extended to C. H. Carlson and R. E. Brewster

of Marqumrdt's Product Operations Group who generated the super-

sonic combustion ramjet (SCRAMJET) performance data incorporated

in this report. Appreciation is also expressed to A. Malek and

A. J. Hayek of the same organization for assistance in defining

approaches for engine structural design and cooling.

The Contributions of Rocketdyne, a Division of North American,

Incorporated in the area of the primary rocket subsystems design

effort is noted. Lockheed California Company provided vehicle

integration support which permitted engine sizing and configura-

tion selection for the vehicle model considered here. The assis-

tance of Rocketdyne and Lockheed was received via Marquardt sub-

contracts under Contract NAS7-377.

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TABLE OF CONTENTS

PREFACE

INTRODUCTION

General

Background

Selected Class 2 Engines

Scope and Content of the Report

General Reference Data

SUPERCHARGED EJECTOR RAMJET (ENGINE NO. ll)

Technical Description

Schematic and Operation Mode Block Diagrams

Engine Drawing

Engine Weight Statement

Operation Mode Schematics

Engine Basic Propellant Circuits

Operating Mode Control System

Vehicle Design

Inlet Contour, Pressure Recovery, Capture Area Schedule

Ejector Mode Specific Impulse and Thrust Maps

Ejector Mode Air Flow Map and Capture Area Schedule

Ejector Mode Tabulated Performance Data

Fan Ramjet Mode Specific Impulse and Thrust Maps

Ramjet Mode Specific Impulse and Thrust Maps

Fan Operation Mode Specific Impulse and Thrust Maps

SENSITIVITY ANALYSIS (E_INE NO. ll)

Summary - Bases

Reference Trajectories

Baseline Specific Impulse and Thrust

Sensitivity Parameter Ranges

Summary - Results

Inlet Pressure Recovery Effect, Ejector Mode

Fan Pressure Ratio Effect, Ejector Mode

Primary Rocket Equivalence Ratio Effect, Ejector Mode

Primary Rocket Combustion Efficiency Effect, Ejector Mode

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2

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TABU OF C0 nS (conn mD)

Primary Rocket Nozzle Efficiency Effect, Ejector Mode

Mixing Efficiency Effect, Ejector Mode

Afterburner Equivalence Ratio Effect, Ejector Mode

Afterburner Combustion Efficiency Effect, Ejector Mode

Exit Nozzle Efficiency Effect, Ejector Mode

Exit Nozzle Area Ratio Effect, Ejector Mode

Inlet Pressure Recovery Effect, Fan Ramjet Mode

Fan Pressure Ratio Effect, Fan RsmJet Mode

Afterburner Equivalence Ratio Effect, Fan Ramjet Mode

Afterburner Combustion Efficiency Effect, Fan Ramjet Mode

Exit Nozzle Efficiency Effect, Fan Ramjet Mode

Exit Nozzle Area Ratio Effect, Fan Ramjet Mode

Inlet Pressure Recovery Effect, Ramjet Mode

Combustor Equivalence Ratio Effect, Ramjet Mode

Combustor Combustion Efficiency Effect, Ramjet Mode

Exit Nozzle Efficiency Effect, Ramjet Mode

Exit Nozzle Area Ratio Effect, Ramjet Mode

Inlet Pressure Recovery Effect, Fan Operation

(¢ = O, 0.20)

Inlet Pressure Recovery Effect, Fan Operation

(_ : 0.50, 1.00)

Fan Pressure Ratio Effect, Fan Operation (_ = O, 0.20)

Fan Pressure Ratio Effect, Fan Operatien' (_ = 0.50, 1.00)

Afterburner Combustion Efficiency Effect, Fan Operation(_ = O, 0.20)

Afterburner Combustion Efficiency Effect, Fan Operation(_ : 0.50, l.O0)

Exit Nozzle Efficiency Effect, Fan Operation (_ = O, 0.20)

Exit Nozzle Efficiency Effect, Fan Operation (_ = 0.50, 1.OO)

Effect of + 10% Subsystem Weight Variation on Engine

Th ru st/We ight


Thrus_/We ight

78

79

80

81

82

83

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86

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TABLE OF CO.ZEroeS (com_rmmw)

SCe ACE 22)

Technical Description

Schematic and Operation Mode Block Diagrams

Engine Drawing

Engine Weight Statement

Operation Mode Schematics

Engine Basic Propellant Circuits

Operating Mode Control System

Vehicle Design

Inlet Contour, Pressure Recovery, Capture Area Schedule

Ejector Mode Specific Impulse and Thrust Maps

Ejector Mode Air Flow Map and Capture Area Schedule

Ejector Mode Tabulated Performance Data

Ram0et Mode Specific Impulse and Thrust Maps

SCRP_MJET Mode Specific Impulse and Thrust Characteristics

and Reference Trajectories

SENSITIVITY ANALYSIS (ENGINE NO. 22)

Summary - Bases

Reference Trajectories

Baseline Specific Impulse and Thrust

Sensitivity Parameter Ranges

Summary - Results

Inlet Pressure Recovery Effect, Ejector Mode

Primary Rocket Equivalence Ratio Effect, Ejector Mode

Primary Rocket Combustion Efficiency Effect, Ejector Mode

Primary Rocket Nozzle Efficiency Effect, Ejector Mode

Mixing Efficiency Effect, Ejector Mode

Heat Exchanger Equivalence Ratio Effect, Ejector Mode

Afterburner Combustion Efficiency Effect, Ejector Mode

Exit Nozzle Efficiency Effect, Ejector Mode

Exit Nozzle Area Ratio Effect, Ejector Mode

Inlet Pressure Recovery Effect, Ramjet Mode

Combustor Equivalence Ratio Effect, Ramjet Mode

Page

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TABLE OF CONTENTS (CONTINUED)

Combustor Combustion Efficiency Effect, Ramjet Mode

Exit Nozzle Efficiency Effect, Ramjet Mode

Exit Nozzle Area Ratio Effect, Ramjet Mode

Inlet Pressure Recovery Effect, ScramJet Mode

Combustor Equivalence Ratio Effect, Scram Jet Mode

Combustor Combustion Efficiency Effect, Scram jet Mode

Exit Nozzle Efficiency Effect, Scram Jet Mode

Exit Nozzle/Inlet Cowl Area Ratio Effect, ScramJet Mode


Thrust/We ight


s'_/weig_.t

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LIST OF MAIN REFERENCE ILLUSTRATIONS

Figure No.

1

2

A

B

C

D

E

F

G

H

I

J

K

L

Title

Flight Velocity -Mach Number

Installed Engine Station Nomenclature

Inlet Pressure Recovery Sensitivity Analysis

Range

Exit Nozzle Area Ratio Sensitivity Analysis

Range, Engine No. ll, Ejector Mode


Range, Engine No. ll, Fan Ramjet Mode


Range, Engine No. ll, Subsonic Combustion

Ramjet

Inlet Pressure Recovery Sensitivity Analysis

Heat Exchanger Equivalence Ratio Sensitivity

Analysis Range, Engine No. 22, Ejector Mode

Heat Exchanger Equivalence Ratio Sensitivity


Afterburner Equivalence Ratio Sensitivity


Afterburner Equivalence Ratio Sensitivity



Range, Engine No. 22_ Ejector Mode


Range, Engine No. 22, Subsonic Combustion

Ramjet

Inlet Pressure Recove_F Sensitivity Analysis

Range, Engine No. 22, Supersonic Combustion

Ramje t

ii

12

69

7O

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72

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UNCLASSIFIED Report Z5,194Volume 7

Page 1

PREFACE

This report comprises a major portion of the technical results of the Class

2 study phase of National Aeronautics and Space Administration Contract

NAST-37V, "A Study of Composite Propulsion Systems for Advanced Launch Vehicle

Applications". This phase of the program was conducted under Task IV (of four)

of the contract work statement.

Composite cycle launch vehicle engines, as defined for this study, are single

integrated propulsion systems which are comprised of both rocket (liquid-

propellant) and airbreathing subsystems, e.g., primary bipropellant combustors,

inlets. To date this type of powerplant has received little systematic study

wherein common ground rules are employed to judge the possible merits of the

large number of candidate engines.

The potential advantages offered by the more attractive composite systems in

advanced (reusable) vehicles include the following points: high payload per-

formance (exceeds the advanced rocket, roughly equals the turbomachine-type

airbreather), high operational flexibility across the reusable-cycle mission

profile, ease of development in terms of the indicated major facility require-

ment for competing pure-airbreathing engines (composite engines can be seg-

mented to fit existing or planned ground test facilities which provide high

simulated Mach number airflow capability).

It is the objective of the study to (1) appraise this potential for advanced,

reusable launch vehicle applications, and (2) provide technical guidance for

initiating possible research and development efforts directed toward the

ultimate creation of these systems. The study included consideration of both

single and multistage vehicles, for earth-orbit payload delivery. The study

concentrated on launch vehicles in the 1,000,O00 pound gross weight class

which operate on hydrogen/oxygen propellants. In general, the study was

directed toward propulsion system first availability in the period 1975-1985

and full mission-cycle propulsion requirements from lift-off to landing was

considered. The principal performance criteria for engine ranking purposes

was payload-in-orbit to gross weight ratio. Other criteria were, however,

brought into play as appropriate.

Marquardt, prime contractor, Rocketdyne and Lockheed were associated in this

analytical ana design study effort. The study was extended over nine (9)

months with a final report (of which the present report is Volume 7 ofseven) submitted to distribution in February 1967.

UNCLASSIFIED

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UNCLASSIFIEDReport ZS, 194

Volume 7

Page 2

INTRODUCTION

GENERAL

This report is the third in a series of three, which specifically present

engine information derived from the NASA-contracted study, "A Study of

Composite Propulsion Systems for Advanced Launch Vehicle Applications"

(NAST-377). The three reports (the other two are Refs. 1 and 2) are associ-

ated with the three chronological phases of the study and comprise varying

numbers of engine concepts and various degrees of technical penetration asfollows :

Report Associated Number of Technical

Order Study Phase Engines Included Penetration

1 Class 0 36

(Ref. l)Overall System Analysis

only, performance on

three (3) reference tra-

Jectories based on

"ideal" inlet, important

parameters 'bracketed"

orly

2

(Ref. 2)Class 1 12 Included subsystem con-

sideratlons, performance

presented in map form,

based on realistic inlets,

conceptual designs made,

important engine variables

exercised parametrically

3

(This

Report)

Class 2 2 Effect of varying subsys-

tem and component efficien-

cies and operational points

assessed, performance maps

broadened and refined,

detailed conceptual designs

rendered based on vehicle-

stipulated sizing parameters_

approaches for structural

and thermal design and engine

control investigated.

To this end the present report presents detailed working information on two

(2) engine concepts taken from both the initial candidate listing of 36 con-

cepts reported in Ref. l, and the twelve (12) engine types further treated

in Ref. 2.

The next section will briefly review the two engine concepts. Also the scope

and content of this report will be summarized prior to the two major engine-oriented sections of the document.

'UNCLASSIFIED

.eport 25,194Volume 7

Page 3

BACE_ROUND

Of the thirty-six (36) engines originally ordered within the Class 0 phase

study (Ref. 1), twelve (12) were selected for further study as Class 1 sys-

tems. These twelve (12) systems can be viewed as variations about a single

"parent" multimode composite engine concept; the afterburaing cycle, air

augmented rocket/ramjet system:

SELECTED CLASS 1 ENGINES

(12 SELECTED FROM 36 CANDIDATES)

AFTERBURNING CYCLEAIR AUGMENTED ROCKET/-RAMJET PROPULSION

SYSTEMS

•--_NON AIR LIQUEFACTION SYSTEMS

(FOUR ENGINES)

•.--,_AIR LIQUEFACTION SYSTEMS

tEIGHT ENGINES)

The basic split shown is that of higher performance air liquefaction systems

versus the somewhat simpler non-air liquefaction systems which effects a

grouping of eight (8) and four (4), respectively. Both sub-families are

represented in previous engine types studied by The Marquardt Corporation in

the guise of lightweight, efficient acceleration and cruise aircraft power-

plants. These are the Ejector Ramjet systems (non-air liquefaction) and the

RamLACE systems (air liquefaction). These "parent" engines are further des-

cribed in Ref. 2. Design and performance data for the twelve (12) Class 1

Engines, in essence, comprise Ref. 2.

Report 25,194Volume 7

Page 4


The two engine concepts selected for the Class 2 studies were:

Supercharged Ejector Ramjet (SERJ - Engine No. ii)

SCRAMLACE (Engine No. 22)

With reference to the Class I Selection summary given above, it can be seen

that a continuation of the duality: non-liquefaction/liquefaction engines,

reflected in the Class 1 selection, was also carried forward into Class 2.

That is, both an air-liquefaction engine (No. 22) and a non-liquefaction

engine (No. ll) were further appraised. Each of these is clearly a better

performer in its category, where at the same time additional complexity,

e.g., recycled hydrogen, brought only modest gains in payload for the mission

model used. The significant advantage of the Supercharged Ejector Ramjet

over the basic Ejector Ramjet was notable. This caused it to be chosen as

a Class 2 system despite the additional hardware implication of the fan sub-

system. The Class 2 Engines are summarily characterized in this table:


Supercharged Ejector Ramjet (SERJ, Engine No. II)

Attractive Payload Potential with Minimum

Technology Uncertainties- Providing a

Nearer Term Availability

ScramLACE (Engine No. 22)

Maximum Payload Potential via the Combination

of Air Liquefaction and Supersonic Combustion

Ramjet Operating Modes, which are not fully

Developed Technologies

. _ort 25,194Voluxne 7

Page 5

As can be inferred from this listing, the concept of having two packages

of varying technical risk is apparent in the engines selected. Engine ll,

does represent a significant improvement in potential over the baseline

rocket in a concept that has little technical risk associated with it. On

the other hand, the Engine 22 represents a considerably higher payload gain.

However, two technological risk areas are apparent: (1) the air liquefaction

process and, (2) the SCRAMJET mode operation. These selected engines for the

Class 2 study are schematically shown and further discussed in the following

pages.

SUPERCHARGED EJECTOR RAMJET (SERJ - Engine No. 11)

The supercharged Ejector Ramjet schematically is shown here as a basic

Ejector Ramjet which has mounted before the mixing section a low to moderate

pressure ratio, thin profile low blockage fan subsystem. In concept, the

fan can be driven by either an airbreathing or a bipropellant gas generator.

The fan acts to supercharge the basic cycle in the initial acceleration mode

thereby improving its specific impulse while reducing the rocket subsystem

sizing. Perhaps most important, it provides a mode of operation for low

speed flyback-to-base via a ducted fan mode with or without plenum burning.

(This would be accomplished by relighting the afterburner at various degrees

of lean burning up to stoichiometric.)

SUPERCHARGED

EJECTOR RAMJET - ENGINE NO. 11

m--'-"----- IN LE T "_ '_- M IXER/DIFFU SER _ EXIT --'e'

F[F_- 0_-----.) -"'TURBOPIJMP ASSEMBLY

Report 25. 194Volu._e 7

Page 6

This engine is capable of operating in four discrete modes : (i) supercharged

ejector mode, (2) fan ramjet mode, (3) subsonic combustion ramjet and (4) fan

only operation. The later operating mode is a low-thrust capability applicable

to the flyback and loiter aspect of the mission profile. The engine consists

of a single stage low pressure ratio fan capable of being retracted from the

main engine flow stream. Accompanying the tip turbine fan is a fan drive sub-

system consisting of a remote airbreathing gas generator, or small turbojet

engine. Following the fan, a primary rocket subsystem, mixer, diffuser, after-

burner, and variable geometry exit nozzle are included as in the basic ejector

ramjet engine.

For a typical mission profile the engine is initially operated in the super-

charged ejector mode wherein the fan operates at design speed. The stoichio-

metric primary rockets are operated at full thrust condition and the after-

burner operates stoichiometrically. At a flight condition in the approximate

vicinity of Mach 1 the primary system can be phased off, the engine continuing

in the faro ramjet mode (technically a high bypass ratio, full plenum burning

turbofan cycle). In the vicinity of Mach 2 plus, fan operation is stopped and

the fan is hinged forward and retracted from the flow stream. The engine con-

tinues in the subsonic combustion ramjet mode to the staging condition. Follow-

ing entry and cruise-back in the subsonic combustion ramjet mode, subsonic loiter

and landing is accomplished with fan only operation with little if any plenum

burning.

SCRAMLACE (Engine No. 22)

The SCRAMLACE is schematically represented as having inlet and exit configura-

tions which are compatible with the SCRAMJET mode operation. The heat exchanger

subsystem shown here in the engine flow passage will from practical considera-

tions be located external to the throughput area of the powerplant as will be

noted in the conceptual design drawing provided herein.

SCRAMLACE - ENGINE NO. 22

11-

eport 25,194Volume 7

Page 7

This engine is capable of three operating modes*: (i) liquid air cycle ejector

mode, (2) subsonic combustion ramjet and (3) supersonic combustion ramjet. The

engine consists of a primary rocket subsystem which operates on liquid hydrogen

and liquid air, with the liquid air being supplied from the air liquefaction

unit, consisting of a precooler and condenser. The refrigerant is the liquid

hydrogen total flow supplied to the engine. Following the primary rocket sec-

tion is the mixer, diffuser, afterburner and variable geometry exit nozzle.

Initial engine operation is in the ejector mode with full thrust operation of

the primary rockets at a stoichiometric condition. Air flow, nominally con-

stant, is controlled by hydrogen flow into the engine and the specific flight

conditions, non primary fuel being burned in the afterburner st a significantly

fuel rich condition. At an appropriate flight Mach number the primary rocket

subsystem is shut down and the air liquefaction unit is closed off from the in-

let diffuser. The engine continues to operate with stoichiometric combustionin the afterburner as a subsonic combustion ramjet. At approximately Msch 6

the engine transists into supersonic combustion ramjet operation by simultaneous

shifting of combustion forward into the region of primary rockets (the rockets

are not reignited) and full opening of the aft end of the engine to permit the

normal shock system to pass from the engine. Upon entry, flyback is nominally

accomplished in the subsonic ramjet mode with loiter and landing being achieved

in the liquid air cycle ejector phase operation.

It might be noted here that the geometric criteria for efficient mixing of fuelin the SCRAMJET mode are approximately the same as those involved in the rocket/

air mixing phase for the ejector mode. This implies that the physical geometry

of the rocket structure might in fact be compatible with the SCRAMJET fuel in-

Jection requirement. The presence of the primary rocket subsystem end its

supports, as well as the afterburner fUel injection struts (if these are not

retracted), will affect the supersonic flow stream and these must therefore be

designed with minimum stream shock losses as an objective.

* An inlet closer rocket vacuum mode is feasible for ScramlACE provided

vehicle supplied oxidizer (liquid oxygen, liquid air) is available.

This mode is schematically indicated in the Scraml_CE section of the

present report.

Report Z5,194Volume 7

Page 8

SCOPE AND CONTENT OF THE REPORT

As stated, this report includes a separate section for each of the two (2)

engine concepts described above. The original numerical coding assigned to

these engines as candidates (Class 0 Phase, Ref. I) is retained for continuity,

the engine sections appearing here in numerical order.

The orientation of the engine data presented herein is toward direct user

processing for broad and diversified study activities. Performance, weight,

physical envelope characteristics, operating mode availability, and other in-

formation of this genre is arranged here in a manner intended to promote

effective assimilation of composite engine data by the reader. For this reason,

the documentation of interpretative results of the engine data, e.g. mission

application studies, is left in the main body of the report (Ref. 3). Similarly,

discussions bearing on the trade-off studies leading to selection of engine

design parameters, such as primary rocket chamber pressure, also remain in the

main report, since - per se - these may not be of immediate utility to a sys-

tems analyst striving to assess the applicability of composite engines to his

particular mission requirement.

Therefore, as appropriate, reference should also be made to the main body of

the Study's final report documentation (Ref. 3). There the bulk of the para-

metric analysis which, for example, explore the effect of the internal design

variables, is provided. Also the Study's vehicle integration and mission per-

formance work is represented in these volumes.

Each of the two engin e sections to follow is divided into two parts: (1)

Engine Description, Physical Characteristics, and Performance, and (2) Engine

Sensitivity Analysis - Bases and Results.

In further detail, the topics included, in the order presented, are:

Engine Description I Physical Characteristics and Performance

I. Descriptive Text, Schematic, Operating Mode Block Diagrams

2. Detailed Conceptual Drawing (Includes Numerical Statement of

DesignFeatures)

3. Weight Statement

4. Operating Mode Schematic Diagrams

5- Propellant Flow Circuit Description

6. Vehicle Installation Description

_port 25,194Volume 7

Page 9

T. Assumed Inlet Physical Characteristics and Pressure Recovery

Schedule

e Ejector mode (or supercharged ejector mode) specific impulse,

thrust and airflow maps reflecting the effect of vehicle flight

speed and altitude. These maps are backed up by computer-gener-ated tabular data.

9.* Fan-ramjet mode specific impulse and thrust maps.

i0. Ramjet (subsonic combustion) specific impulse and thrust maps,

including the effect of inlet air precompression (flow field).

II.*_SCRAMJET (supersonic combustion) specific impulse and thrust

data, including the effect of inlet air precompression (flow

field). This information is presented for three reference

trajectories which follow the performance curves.

12.* Fan (ducted) operation specific impulse and thrust maps, reflect-

ing the effect of varying degrees of plenum burning.

Sensitivity Analysis - Bases and Results

i. Reference Trajectories

2. Baseline Specific Impulse and Thrust (both net Jet) Performance

Values derived on the reference trajectories

S. Range and Limiting Values of Sensitivity Parameters, Performance

and WeightC

_. Perturbed Specific Impulse and Thrust - results for each sensitivity

parameter.

Preceding the individual engine sections, and immediately following this

section, a general reference section appears which includes:

i. Mach Number/Velocity Conversion Chart

2. Engine Station Nomenclature Diagram

B. General nomenclature and legends

4. List of references

* Engine No. ll only

**Engine No. 22 only


Page 1 0

SUMMARY - GENERAL REFERENCE DATA

The purpose of this section as noted in the introduction is to provide tech-

nical information and general background material applicable to the two (2)

specific engine sections to follow. Each of the items to be provided in

this general section will be briefly discussed below.

Msch Number versus Flisht Velocity Conversion Chart - Although the basic

engine performance information to be presented in this report is given gen-

erally on the basis of flight velocity (ft/sec, m/sec), much of the general

information as well as the intermediate data is most effectively and conven-

iently stated in terms of Mach number. A conversion plot is provided to

assist in approximate conversions of these two velocity terms. For more

precise computations the use of appropriate tables, however, is recommended.

Ensine Station Desi&nation and Nomenclature - An installed engine schematicis presented reflecting a typical composite engine of the Class 2 series.

The several aerothermodynamically significant geometric stations employed

in the engine general description, as well as in the performance computations,

are called out in this figure.

Standard Efficiencies - The following listed efficlencles have been used as

baseline values for all engine performance computations:

PrimaryRocket:

Combustion, _c* = 0.975

Nozzle, _n = 0.98

Mixer:

Mixing, _m -- O.80

Afterburner or Combustor:

Combustion, _c = 0.95

Exit :

Nozzle _ n m 0.98

Legends_ Nomenclature 2 and References - Within the engine sections certain

diagrammatic conventions have been adopted and these are reflected in both

schematic and tabular form in this section of the report. Also a nomencla-

ture sheet is provided for all symbolic characters employed either in the

presentation of the engine information, or in the computations supporting

the performance provided. Finally a list of references is given at the endof this section.

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Report Z5, 194Volume 7

Page 11

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........ _--.11 ....

:1:_ .... _/-",71:

:::; : :I:1::::..

....... l-t._ ....

e.., ........ =.,o .......... 1-1.,t-:..

::;! |::_:'1:

..... e.,, ....

:::: .... i_..i ....:::: ' :::1-1

iii !i!iii....T;::

;::: i!i!ll .

-.:. .... .:::;

:;:.. '. ,:::

::: ::'j::_:::

: i.li:i!i!_i:i

-._- •.,i:t....:.._

.-._ !_. ,_,..! ....

i:--_. "i!it%!ti:-:.!

___ liii!l_4i

10

VELOCITY, I000 ft/sec

::::lii'iJ i!_i L:!:::] ............::;:1 ::::1 ::."1 MI.... I ............

::::i:L2 iii:: !i::

..i "_ 1"Iii.........

..e ............

.e ......... , , ..

[:: :':: :::- ::,.

: ::: .'.:::::!I

". .:;: :::::;:I!

...... ::L: ....

:.:. .:. :::: -:

......... i;,.0!

:::s:2_!gigi;_il

::: :::: T;xr_:;_!

............ ,...

............ ;.... i

_:; i::i i-._!ii_i]

t:ii ::! :Yl

_::1!..:!..'_12 14

I0

I I I1.0 2.0 3.0

FLIGHT VELOCITY, 1000 m/sec

I4.0

_8

rquard! .,..o,,.o.,,,o,.,.

FIGURE 2 INSTALLED ENGINE STATION

Report 25 194Volume 7

Page 1 Z

NOMENCLA TU RE

Ac

FLOW FIELD i:"

INLET

SHOCK

A2 A31 A4 A6

A3 A5iICOMBUSTOR

_PRIMARY '_THRUST

CHAMBER ASSEMBLY

/FUEL INJECTOR

ASSEMBLY

A (PRIMARY EXIT AREA)P

FOR CONSTANT AREA MIXER A2+ Ap = A3

FOR CONSTANT AREA COMBUSTOR A3a= A 4

: ii!:i_!ii!iil

The basic flow processing situation in a representative composite engine

is schematized in this chart, which also provides the station nomenclature

used in the study. It will be note that both on the inlet and the exit

portions of the aerothermodynamic processes the favorable utilization

possibilities of the vehicle are assessed. _ae precompression of the in-

duced air by the vehicle body (vehicle bow shock system) is fundamentally

important in providing high engine thrust performance, particularly in the

ramjet modes. At the same time, the possibility of expanding the exhaust

flow on the vehicle aft body is of fundamental importance, critically sofor the SCRAMJET mode.

/J

rcluaraz .... _.,..,,,o..,.-- s,Jr:iHll_,U'm",

LEGEND- ENGINE SCHEMATIC

..=port 25. 194Volur_e 7

Page 13

Inlet I Exit: Subsonic Combustion

= INLET _

.--- EXIT---.a

Inlet I Exit: Supersonic Combustion(Including Sub/Super Conversion)

Precooler I

air

Condenser 36"R H 2l

i yH2 liquid air

Precooler I Condenser

air

25"R H?A

i VH2 liquid air

" la r(luard!

IJNCLASSIFIED Report 25 194Volume 7

Page 14

LEGEND- ENGINE OPERATING MODE BLOCK DIAGRAM

Letter Symbols (Within Blocks)

I Inlet SubsystemHE Heat ExchangerF Fan SubsystemR Rocket SubsystemMC Mixer/CombustorMD Mixer/DiffuserC CombustorE Exit

Mixer/CombustorlExit Subsystem

Letter Symbols (Fluids)

H Hydrogen0 OxygenA AirX Exhaust

Graphical Symbols

[ 1I

!

I It j

Functioning Unit

Non-functioning Unit

I i

I I

)

Fluid Flow Direction

Fluid Flow Through a Non-functioning Unit

Flow Mach Number > 1

Flow Mach Number <1

Flow Mach Number is both below and above 1at changing flight speed conditions

UNCLASSIFIED

UNCLASSIFIED

VAN _U_'|. CAI.;FOeNIA

_eport 25,194Volume 7

Page 15

NOMENCLATURE

Nomenclature used in this report is given below. The tabulated computer

printout inforumtion does not include subscripting as will be noted by the

repetition of certain parameter symbols. Refer to Figure 2 for engine flow

area station designations including the distinction n_de between A c_and Ao

where a vehicle flow field is involved (where there is no flow field these

are identical).

AB

A_/A 3

A5

A6

A6/A 5 -

A6/A c

AHX

AO

AOT

BL

CF, CF

H2

HTO

Isp , IS

Mo, M0

M2

NS

o/;P2

Pc

PRf

PT2 , PT2

PT2/PTo

PT0

Afterburner

Afterburner/Mixer Diffusion Ratio

Engine nozzle throat area, ft2

Engine nozzle exit area, ft 2

Exit Nozzle Expansion Area Ratio

Exit to Capture Area Ratio (SCRAMJET) ..

Inlet capture area for heat exchanger air flow, ft2

Inlet capture area for secondary air flow, ft 2

Inlet capture area for total air flow, ft2

Baseline

Thrust Coefficient based on inlet capture area

Secondary air static enthalpy at mixer entrance, Btu/lb

Ambient Total Enthalpy, Btu/lb

Specific Impulse, lbf/lbm/sec (Net Jet)*

Local Mach Number

Mixer entranceMach number

"Normal Shock"" inlet (Includes Normal Shock losses

plus an assumed 90_ diffuser efficiency.)

0xidizer/fuelmass flow ratio

Secondary air static pressure at mixer entrance, Btu/lb

Primary chamber pressure, psia

Fan pressure ratio

Inlet recovered total pressure, psia

Inlet total pressure recovery

Ambient total pressure, psia

UNCLASSIFIED

"'A_/#'I _J/

//imrquamr .,,,o,.,.

UNCLASSIFIED Report 25_ 194Volume 7

Page 16'

R i

Ref

SIS

SPC

T

V6

Vo, V0

WFT

WKX

Wp, WP

Ws/WP, WSWP-

WS, WS

WT

8 -

_C'N" .

_I_ -

Rocket Mode

Reference

Sea Level, Static Conditions

Specific Fuel or propellant consumption, lbm/hr-lb f

Thrust, ibf (Net Jet)*

Exit velocity, ft/sec

Local velocity, ft/sec

Total fuel or propellant flow rate, lbm/sec

Heat exchanger air flow rate, ibm/sec

Primary flow rate, lbm/sec

Secondary/primary flow ratio

Secondary (WS + WHX), lbm/sec

Total air flow

Two dimensional wedge half angle, deg

Combustion efficiency based on enthalpy rise

Characteristic velocity efficiency based on velocity, or thrust

Inlet kinetic energy process efficiency

_M

?IN

_p, PHP

_prec

_sec, PHS -

Mixing Efficiency based on static pressure rise

Nozzle efficiency based on stream thrust

Combustor equivalence ratio

Condenser equivalence ratio

Heat exchanger equivalence ratio

Primary rocket equivalence ratio

Precooler equivalence ratio

Secondary equivalence ratio

Net Jet thrust and specific impulse includes air induction inlet

momentum penalty, but does not include external drag such as cowl,

induced, friction, or spillage drag

bI ED

VdIH NUY$. CAIIFOIIN,a

R_,Jrt 25,194Volume 7

Page 17

LIST OF REFERENCES

la

_o

m

"Class 0 Engine Fact Sheets (Thirty-six Engines)", Contract NAS7-37V,

Marquardt Report 25,19h, Volumes 4 & 5, Sept.1966. CONFIDERTIAL -

Title Unclassified.

"Class 1 Engine Information (Twelve Engines)", Contract NAS7-377,

Marquardt Report 25,19_, Sept. 1966. CONFIDENTIAL - Title Unclassi-

fied.

"A Study of Composite Propulsion System for Advanced Launch Vehicle

A.vp!ications (Main Technical Report) ", Contract NAST-BTV, Marquardt

Report 25,192, Volumes 2 & 3, Sep_ 1966. CONFIDENTIAL - Title Unclassi-fied.

Hrquaror ....o.,=.,,,o..,.

I I.1/RPt /RATIf_


Page 18

SUPERCHARGED

EJECTOR RAMJET, NO. 11The Supercharged Ejector Ramjet (Engine No. ll,

Class 2 Study Phase) is a 215,000 lbf thrust

(sea level, static) engine with Mach 8 flight

speed capability. The propellants are liquid

hydrogen and liquid oxygen and the engine

normally operates in four progressive modes:

(1) supercharged ejector mode, (2) fan ramjet

mode, (3) subsonic combustion ramjet mode and (k) fan operation mode.

As displayed in this section, the engine has an overall length of 371 in.

(9.4 meters), an overall diameter of 142.5 in. (3.62 meters) and a height maximum

of 165 in. (k.19 meters). The unlnstalled engine weight is ll,9i0 lbm, yielding

a sea level thrust-to-weight ratio of 18.0.

The basic design specifiers for the engine are as follows: Design mass flow ratio

3.0 to l, primary chamber pressure 1500 psia, fan pressure ratio 1o3, maximum

internal pressure 150 psia.

The engine features a single stage retractable tip-turbine fan powered by a twin

airbreathing gas generator installation. The fan bypass ratio is l0 to 1. The

primal, rocket is a regeneratively cooled annular bell configuration featuring

a single toroidal combustion chamber fed by separate hydrogen and oxygen pumps.

The turbopump drive operates on the gas generator cycle using self-pumped propel-

lants. A third hydrogen pump provides fuel to the afterburner, during the super-

charged ejector mode, and thereafter feeds the ramjet combustor during high speed

operations. The afterburner fuel pump is also powered by a hi-propellant gas

generator utilizing self-pumped fuel and pressurized liquid oxygen provided frcm

the vehicle.

The basic engine structural components (mixer, diffuser, afterburner and exit

nozzle) consist of regeneratively cooled assemblies employing a ring-stiffened

Rene' kl wire-wrapped, brazed regenerative Hastelloy X tube bundle construction.

Within the mixer is an _lOngated centerbody which structurally connects a fixed

aft plug with the forward thrust ring. The center body and plug assembly is sup-

ported by a multiplicity of radial low drag fuel injector struts commencing atthe afterburner station. Variable exit geometry is accomplished by means of a

translating ring operating continuously to provide two coaxial flow expansion

compartments between the outer exit bell and the fixed plug. This dual throat

design provides a minimum weight, single moving part design and provides high

nozzle performance.

The engine was sized for a 1 million lbm gross weight, horizontal takeoff, two-

stage launch vehicle. The engine was located in a complement of five (5) along the

bottom side of a high fineness ratio, low drag, lifting body boost stage. Air in-

duction considerations for this installation consisted of a moving ramp, two-

dimensional variable inlet with mixed external and internal compression. The inlet

capture plane was located to make full use of body flow-field affects at speeds in

excess of Mach 3- Exit gas expansion is considered to take place solely within the

exit bell of the engine.

_a//

rquaror ...._,,..,,,o..,.I IX )#I_ _VATh'_


Page 19

Eng. No. 11

Engine OperatingSchematic

l/ I _""_ PilMARY TH UST __

.... ] ,,_,,,_ CHAMBER ASSEMBLY--"

_TURBOPUMP ASSEMBLY

Engine Operating Mode Block Diagrams*

@ X

®

©

@

H

I R I HMD' C EL_._" .j L____ _

H

: : L_._JI. .... J ---

H

,,,,,,, , :---",,,o,,,,[---], , ,0 '----'_ i. .... j I. .... J --- L .... a_

H

xI.-_.- J ......

*Note: Mode numerical coding is given on Page 18, first paragraph,

"R" indicates optional all-rocket mode.

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VAN NU1FI. CAiifOINldl

Report 25,194Vohu'ne 7

Page 21

Eng. No. IIWEIGHT STATEMENT - ENGINE NO. ii

Fan Subsystem

Fan AssemblyGas Generators

Frame and Trunnion Unit

Compartment StructureCover

Actuator

Transition Section

Miscellaneous (5_)

12581120

73O36o210

i15

3O67O

4169 ibm

(34.9%)

Primary Rocket Subsystem

Rocket Chamber AssemblySupport Structure

TurbopumpsGas Generator

Ducting and Valves

Starting System

927

316

12788

126

2028

(17.o_)

Mixer/Diffuser/AfterburnerMixerDiffuser

Fuel Injection UnitCombustor

Forward CenterbodyTurbopump and Miscellaneous

io575_o635

315

3OO145

2992(25.1"_)

Exit Nozzle SubsystemExit Bell

Translating Ring Assembly

Fixed PlugActuator UnitMiscellaneous

523

973734!ooll6

2446(2o.5_)

Controls, LinesControl AssembliesValves and Lines

8o225

305(2.5_)

Total Weight, Dry(Thrust = 215,000 ibf)

i1,9_0 ibm (5416 kg)

Thrust/Weight, Uninstalled 18.0

Report Z5Volume 7

Page ZZ

SUPERCHARGEDEJECTORRAMJET (ENGINE NO. 11|PROGRESSIVE OPERATING MODES

(PUMPINGe COOLING AND CONTROL CIRCUITS NOTSHOWN)

1943

SUPERCHARGED EJECTOR MODE

:-:.:-y..:,:.:.:,:-:.:.:.:-:-.":_..:.'..:.:-:.:.:-:-'.:-

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-::::_:::_::::::::::::::::::::::::::::c".'-:.:::::.<::::::::::::::::::::::::::::::

•-.::_:!:_:_:_:_:_:_@_$_!:.• .:..:::::: ::: :::::::

============================

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• :.'.";:'$i:':__$_:_:_::."::::_$:.:i:_:::::.:::

i

FAN RAMJET MODE

R-22, 137

VAN NUV|. CALIPOINIA

Report 25

Volume 7

Page 23

194

SUPERCHARGEDEJECTORRAMJET (ENGINE NO.PROGRESSIVE OPERATING MODES

(PUMPING, COOLING AND CONTROL CIRCUITS NOT

11)

SHOWN)

H 2

RAMJET MODE

H2

FAN

:.:.:;::

_ili!iiii_iiii!i!!i!!i!iiiiiiiii!i!iiiiili!iiii!!iiiiii!iiiiiii!iii!!!ii!i!Y

OPERA TION M 0 D E ...,....,_._..:.

R -22, L 38

.....,..,,,o..,.

__ ._ort 25 1943

Volume 7

Page 24

BASIC PROPELIANT CIRCUIT

SEPJ, ENGINE NO. ii

Eng. No. II

The propellant circuits for Engine No. ll including the pumping, cooling

and primary control elements are displayed on the facing figure. The

engine is supplied hydrogen and oxygen as shown. The airbreathlnggas

generator, which is aerodynmmically coupled to the tip turbine single

stage fan, receives fuel and is controlled for maximum specific horse-

power output.

The three pump units are directly shafted units requiring no gearing.

One gas generator, operating on tapped propellants, is utilized to drive

the primary rocket fuel and oxidizer pump (bottom of figure) to supply

the combustion chamber operating at 1500 psia. A second bipropellant

gasgenerator drives the afterburner fuel pump which provides hydrogen

at an output pressure of 1OOO psia. For the ramjet mode it is required

that auxiliary stored oxidizer be provided to the turbopump gas generator

since oxygen itself is not being pumped in this mode. The primary

rockets are fed through main propellant valves operating in either open

or closed positions. Starting of the primary rocket is accomplished by

turbine blow down from a separate high pressure gaseous hydrogen supply

(not shown). The turbopump exhaust products, being considerably fuel

rich, are conducted into the afterburner of the engine where the excess

fuel is combusted with the induced air. This minimizes the turbopump

drive penalty during the ejector mode. Likewise, the ramjet pump turbine

exhaust is injected into the ramjet combustor. The turbine is therefore

designed for back pressures commensurate with sonic exhaust flow at the

full combustor pressure of 150 psia to preclude back pressure coupling

effects.

The coolant passages consist of two parallel flow circuits to the outer

jacket and to the inner centerbody/plug/nozzle assembly. The outer

engine wall is cooled in two passes first forward and then aft in order

to cool the exit nozzle with warmed hydrogen. The inner cooling circuit

progressively cools the forward centerbody, translating ring and fixed

plug. All coolant flow is injected into the afterburner via the self-

cooled fuel injection struts. Also, the primary rocket external structure

is cooled by regenerative structure during the high speed ramjet mode.

During fan only operation, with little or no plenum burning the engine

is essentially uncooled. For full stoichiometric afterburning (fan ramjet

mode), regenerative cooling of the aft section of the engine is required.

rquardlI t .YAe/4'_4 Y//]W

VAN NUY$. CAUFOINIA


Page 25

Eng. No. ii

mlm

I--'-

61.1

-r.C.Iu,')

I-,-Z

--.I.--.I1.6,1

Q.0

CJI

,<ee_

!

,,-,,I

dZ

ZI

ZILl

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ItIIItIIItII

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R-ZZ, 142

H

rquamr ......,.I I,,YJRICmAI'IfJW

Re .'t25,194Volume 7

Page Z6

Eng. No. 11

OPERATING MODE CONTROL SYSTEM

(BLOCK DIAGRAM)

EJECTOR MODE (NO. 1)

•_. m FAN RAMJET MODE (NO. 2)

.... RAMJET MODE (NO. 3)

.... FAN OPERATION MODE (NO. 4)

OPERATING

MODE

INPUT SIGNAL

CONTROL

LOOP

SELECTOR

FII

t!I

II

COMBUSTOR/

AFTERBURNER

_- CONTROL SYSTEM

I

I _PRLMARYROCKET II CONTROL SYSTEM

I

EXIT NOZZLE

C ONT ROL

SYSTEM

AIRBREATHING

GAS GENERATOR

CONTROL SYSTEM

Engine control is based on manual and/or aut_natic selection of mperating modesvia four active control loops as will be described. The fan subsystem is

controlled to operate at design rpm with protection of both the fan tip-

turbine and the basic gas generator from turbine over-temperature. Retraction

of the fan is effected at mode shift from the fan ramjet to subsonic combus-

tion ramjet mode. The primary rocket is scheduled (pre-orificed) to operate

at design chamber pressure and design oxldizer-to-fuel ratio (8.0 to 1).

The afterburner/combustor controller is designed for operation at an equiva-

lence ratio of l, providing maximum thrust consistent with good performance.

The variable exit nozzle is controlled initially to provide a maximization

of the product of fan pressure ratio and air mass flow (approximately).

Once supersonic flight is reached the variable exit is translated to locate

the inlet nozzle shock at the throat of the inlet. The exit also operates

in an override loop to limit combustor pressure or inlet diffuser pressure

to the maximum design condition of 150 psia.

Report Z5. 194volun_e 7

Page Z 7

Eng. No. 11

COMBUSTOR/AFTERBURNERFUEL CONTROL SYSTEM

I AIRFLOWSENSOR

INPUT

REQ'D

____ EQUIVALENCE FUEL FUELCOMPUTER _ CONTROLLER

AC TUAL

FUEL

FLOW

RATE

I FUELFLOW

SENS OR

The combustor/afterburner fuel control system is a closed loop control system which senses

the air flow rate through the engine and modulates the fuel flow rate to maintain a

required fuel air ratio (_). A signal proportional to air flow is applied at the

equivalence computer which generates a command for the required fuel flow rate. This

signal is compared against a fuel flow rate feedback signal generating an error signal

which is applied at the fuel controller. The fuel controller modulates the fuel flow tonull the error.

ROCKET FEED CONTROLSYSTEM

COMBUSTIONcHAMBER I THROAT

The rocket feed control system is a fixed point control system. A positive feedback

from the pump which is controlled by a metering orifice is supplied to the gas

generator to drive the turbine and pump assembly until the available power is equal to

the required power which is the design condition (full chamber pressure). An external

power source (gas blowdown) is introduced to start the operation of the system.

'A2/V# _H

//I mrquamz ......,,..,,,o..,.I t_Pt _ATIt_

Report 25. 194Volu/ne 7

Page 28I

Eng. No. ii

EXIT NOZZLE CONTROL SYSTEM

J AEROTHERMO-DYNAMIC

FEEDBACK

REF.SHOCK

POSITION

SENSOR &

COMPUTER

PEAK

HOLDING

CONFROLL ER

AC TUATIONSYSTEM

DUCT

FLOW I

RATE J FLOW

J SENSOR

T FANPRESSURE

RATIO

EXIT

NO Z ZLE

AEROTHER/VIO-DYNAMICS

The exit nozzle control system performs two functions, i.e., positions the inlet

shock at its optimum location and maximizes the combination of fan pressure ratio and

duct flow rate by modulating the throat area of the exit nozzle. A pressure signal

indicative of the position of the inlet normal shock is compared against an actual

feedback signal generating an error which is applied at the actuation system to null

the error. The actuation system receives another signal from the peak holding control-

ler to maximize the combination of duct flow and fan pressure ratio. The actuation

system will be designed with the capabilities of selecting the correct signal in theevent there is a conflict between the two commands.

JJrquam[ ....,,.,...,,,o,.,.


Page Z9

Eng. No. II

AIRBREATHING GAS GENERATOR CONTROL SYSTEM

REF.

ITEMP.SENSOR F

P •

FUEL

SUPPLY

CONTROLLER __GAS GENERAT OR

RPM

The airbreathing gas generator control system which is a closed locp ccntro! zyztem,

maintains the fan speed (RPM) at a selected reference speed by throttling the fuel

rate to the airbreathing gas generator. The reference speed is compared agalnz=

the actual speed (feedback), generating an error which is fed to the fuel sul=l:

controller to modulate fuel flow to the gas generator. This, in turn, controls

the fan speed until the error is nulled. A temperature override loop is included

to limit the temperature of products of combustion from the gas generator and prc=ect

the fan tip-turbine.


Page 30J

Eng. No. l l

VEHICLE DESIGN - SERJ, ENGINE NO. ll

This figure shows the final Class 2 vehicle design utilizing super-

charged Ejector Ramjet engines. This 1.0 million pound gross weight

lifting body vehicle was determined to be substantially superior in

performance to the other vehicle types considered.

The lifting body shown features high slenderness ratio, elimination of

the second stage base drag through submergence, and attainment of

stabilizing surface at low unit weight. For the BERJ installation the

aft vehicle extension contains no propellants but provides for

increased slenderness ratio. This SERJ Mach 8 vehicle has a complement

of five 215,000-1b SLS thrust engines (1.075 T/W), with a capture areg

of 350 ft 2, integrated beneath the fuselage. The second-stage gross

weight is &_5,05& pounds for Mach 8 first stage cut-off conditions.

The lifting-body configuration employs a modified conical fuselage

where the forebody is a blunted cone with a depth-to-wldth ratio of

0.& at any station. Maximum cross section of the fuselage is at 73

percent of the body length, as measured from the virtual nose (apex).

The fuselage nose radius is one foot, and the body planform area is13,612 ft 2.

The horizontal stabilizer has a leading edge sweep of 65 degrees,

and an area of 2612 ft 2. The airfoil section is double wedge, with

a two-inch leading edge radius. The movable horizontal control sur-

faces comprise 2000 ft2. The horizontal control surface rotates against

the vertical stabilizer with forward extending dorsal fins, to alleviate

the thermal problem associated with the sharp edges of the control sur-

face under high-speed deflected conditions.

.

The twin vertical stabilizers have a total exposed area of 1200 ft 2,

with a leading edge radius of two inches. No toe-ln is provided for

the verticals, rather, a concept of utilizing small outward rudder

deflections to load the surfaces during hypersonic operation where

the control surface lift curve slope is zero at zero deflection is

proposed, in order to maintain minimum vehicle drag. All panel surfaces

have a thickness ratio of 5 percent.

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Eng. No. 11

_L__

VAN NUVl;, CA_IWORNIA

eport 25,194Volume 7

Page 32] i

Eng. No. II

INLET PRESSURE RECOVERY

oI--0.

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INLET MACH NUMBER, M0

VAN NUI'$. CA&IFOIINIA

Report 25.194Volu_e 7

Page 33

Eng. No. Ii

INLET CAPTURE A REA

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VAN PlUY|, CA_tlOe_iJ

,_eport Z5 194Volume 7

Page 34

EJECTOR MODE SPECIFIC

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Report 25,194Voluyne 7

Page 35

Eng. No. II

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Page 36

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VAN MUV|. CALIFOmN|A

Report 25. 194Volu._e 7

Page 3 7

Eng. No. 11

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Page 43

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VAN NUt'S. CAiJFOI_M


Page 44

FAN RAMJET SPECIFIC IMPULSE

Eng. No, ii

4400

4000

t,.,}

u

3200

-- 28000

o

2400

W

w 2000Z

1600

1200

I0

0.5

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1.5

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i+iti!;!l i :....................... :::=:::: il;.=l;-4?]i! iEi _!!! !!i_ i :ii_.:;; ii:ir_:;: .., .....i!iiisii_ii!!.:;iiALTITUDE, I000 ft

................ ii i 36' '-i_!;i!!! !i!! ;ii! i-il -70 ;1:!-iTii'-i]

.... _ : It..; ............. , ............ / .I

,,:. _;:: .......................... _ ,., .... 1....

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................. _"_ :"_ .... ;i"1 .... I .... 1"1" I

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;:._; ,_,_ .... _...... . :.._ .._:t..,, .................

...... i............... ,.....................::'; ":': :L::J_::: _T'.:.??';;_.:' :TTT T:::I;:':I:::: :;J. :'.':|:.sJTl

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!i'2. iE÷ ii:_]iil i!!_'ii_ii!- _!ii iiTiEi!iti:._:i i'ii ::::_:::....... : ...._.............. :::::...............................i_ .................. :::;_::-........... _:;_; :.!;:i ....................,-:::_; ,.,,,;:::r::_..:::: ."-::-i:=::.....;::; ....!E'P.'i_._-- ....._ "'_.:r.-::'1_' _*'":v'" ' '_-""-- .... _-"

2.0 2.5 3.0

VELOCITY, 1000 ft/sec

I0.2

I I I

0.4 0.6 0.8

FLIGHT VELOCITY_ 1000 m/sec

VAN NUY|. CALIFORNIA

Report 25.194Volun%e 7

Page 45

Eng. No. ii

FAN RAMJET THRU ST

c-o

e-

o,im

E

I--I.u

I--IJjZ

1.4 --

1.2 --

1.0-

0.8-

0.6-

0.4--

0.2-

0-

320

280

240

w,,,,

..=

o 200o0r-4

160

e,.-m}-.

I--IJJZ

80

40

0

I0

0.5 1.0 1.5

FLIGHT VELOCITY, 1000

2.0

ft,/sec

I I I0.2

FLIGHT

0.4 0.6

VELOCITY, I000 m/sec

I0.8

VAN NUV|. C&LIFOflN|A

"-port 25,194Vol_e ?

Page 46

Eng. No. 11

FAN RAMJET CAPTUR E AREA

E

r_<:

IJJ

p.r_

l--UJ

Zm

6

5

4

3

2

0

,,¢I.IJev,,,¢

IJJ

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6O

5O

4O

NOTE: CURVE = UPPER LIMIT

i_:..::.__ _._2-_.:-l!i!t!!!!t!!ii:;:_;::::_-:: ;::_.======================I_:s_:_s."!!!,i_!T!l!!!it:!il;!!!I:_!]:-!l!_!!,!i.!t!!

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"'-.................="..........'..........." ...........'....._iilIiiii,!i_llil•,'.r_* -_ .... r.r_. ,._..+._. ,s..._ _ _-_-,--,-'_,_::_-. :.:. -_.*t.-.:I- .... _,. ,_:I:,]. +.,I-'-t,

;:;.._ _='_..................................................I.... A_ ............ :r'.:Z:.._,'_'jl. ,. . . _! _ .... _', .t ,.,! .'_-.I .I: ,II_1t_ -: ::::|:':

FLIGHT MACH NUMBER

_a__//

rquar(/z ...._.,,..,,,o..,.It:f_/v_4TNJ'o

UNCLASSIFIED

This Page Intentionally Blank

. UNCLASSIFIED

VAN NU¥$, CA&IFO|NIA

Report 25. 194Volu_n e 7

Page 47

Eng. No. II

RAMJET SPECIFIC IMPULSE

SUBSONIC COMBUSTION

NO PRESSURE FIELD

4400

4000

¢3

:' 3600

m

m

. 3200L_J¢n.J

0.

-_ 2800

m

=ram

1.1.1" 2400

_ 2000

16000

I0

ALTITUDE,

1 2 3 4 5

FLIGHT VELOCITY, 1000 ft,/sec

I I I0.5 1.0 1.5


6

I2.0

7 8

J2.5

VAN NU_'S. CAi.fPO|NPA

RA M JET THRU ST

SUBSONIC COMBUSTION

NO PRESSURE FIELD


Page 48

Eng. No. ii

1.4 - 320

1.2

Ult-O

=1.0C0

m

E

_0.8

-r-F.-

'"0.6

I--U.IZ

0.4

0.2

0

280

01-4

200I--u'l

,y-I-

_" 160

ELI

LUz 120

40

00 1

I0

I0.5

100!

2 3 4 5FLIGHT VELOCITY, 1000 P,./'sec

I 11.0 1.5


6 7

:120

8

Z2.0

I2.5

3800 ........

3600 .......

i:': _'-+'

3,400 ::::::::.;i-: _i-i..... ?+1

fJ :::: ':;;

:::: ....,,, 3200 .... :!i'::T'; ::::

• 3000 i!!:.:'::

(J) :::; ::-:

2800 ........

• o,

eL 2600 .... i!i;(J") :::: ::::

:::: :!!:

Ill ::.:: :"::

F,- 2400 i_:.ii--!IiI ........

!!!: ;'-'

2200 ::":C.=._.... ,oo,

.... ,o.,

2000 '-' '_0

VAN NUY|, C&LIPOQN|Ai

RAMJET SPECIFIC IMPULSESUBSONIC: COMBUSTION

EFFECT OF PRESSURE FIELD

port Z5,194Volume 7

Page 49

Eng. No. 11

................ ; ........ !

ii!i: Ni :":i:.'_i: F:.il .... i

'.:'i :-:: 2.--..: :;:: ;:-',c" :::;}

............ :.- .......... J

.............?i;..........:::i :::1 :T:- ' +'t

![_Z:._i_.!'E::A:.i i'.':.!! !ii_::;r;,;" ":T+'. "i"+: ::::l

:::: :T'-!:::: ,'::: ::TT:::7!

++-+ !i!+t°., ..................

' ....... ' ............ I

.++ +, +._ ._ .............. +

..................... it

:;. "2::;._2::;..:::!..]

.... _ ...... ,";;";; '"1

..... o.+iii _ .+ .... ;; .... :: +

"if ....*.... o+* o,.i,

1

_::: ::---_ .... _::: ::::' .....1

"""* ...... ;,'1" 'T ..... 1

............ • ..... ii!]

............. It,'. , i

i

"_:.:_Z" -7_!212:+21!tt::,

++..-+__++;++.+_+*+++++++++

7 8

I I ! I I I0 0.5 1.O 1.5 2.0 2.5


VAN MUY$, ¢ALfFOIINIA


Page 50

Eng. No, 11

RAMJET

SUBSONIC

THRUST

COMBUSTION


3.0

2.5

In

_2;0

iI

_z.5

==

Gz.ch=LgZ

0,5 -

6oo, , ,: ,,,,• , t _ * ' !i:i:i I iiill !:I_i_ I I ¸ _ii:Ii: i I _ i i : ...... ::: _I_: :_!_

:lZ° -'--_:,_ !I I!!!I:: ! 1:;.

400:4:I I i.:l:l_::l_._l_l:.-!:i#l:,:, i;_:,:_:_;::-:_:_;:::i_;:_;'lh__i_I:_: l I I_:..i:::I_!:.:.I!::i!I:!i::lii::Ii:_#:/!i!::Tili:ili:;::_li:Fi::ii::-Ii:_::_:,:I;:!:_i:i::!:-:.i:.;;I::

_ALTITUDE, 1000 ft-:Iii |_!! : :! :: Ii:::i_:::::;::i_! I ii:!::i:I::: :i:;i::ir ' _' :_. f::- I ' ' i300 ,: i: :I. l:-:I ::I::":F:: !.:i: 50 " :: :: :i--: l::..::::_:::r:::!..l: !::I: .I:::: :::I::l::::l::

._ :: . ,:_ :l:_:l::_:li_:!li_:i:::.: I ::::::::::::::::: ::::::::::::::: : :::::::I : ::_:,.;_:" "' 2oo : =:: :I: I::i i:.::ili!!::lii!I;::i:I:i!!::_l i _:!?_ i_i =:i !:;:it.:_;;.i!:.ii_!i!..ii_i_!r:!'ii_:_:_iZ • • :::I: , ::J:.-- •

i_ii!:i::I_::_:":::..... -- i:'_.--:_:::.::::_._ _.:..-I:i:l: : ::::::;.I,_ _'T.. ::I:: _::::_:::: :::_:!:: .... ! ..

100 __:"" ..............................

.:i!hii!llO- --,--_,,.... I._ ........... ,....... ".:-

O- 0 6 7 8

_:_i::I iil:iiliiilI:i:i}i_i!lii!:l._iiI!:Ii:_:::i_I:i:•

0 1 2 3 4 5

FLIGHT VELOCITY, lOOO ['t/sec

I I I I I I

0 0.5 1.0 1.5 2.0 2..


VAN NUY|, C&L|PO(INI&

R=port 25,194Voltage 7

Page 51

Eng. No. 11

FAN OPERATION SPECIFIC IMPULSE

m

,,.;U3.J

Q.=E!

m

IJ.m

t.)uJG.u3

I--I.iJ

I,.-uJz

24,000

20,000

16,000

12,000

8,000

4,000

12,000

8,000

4,0000

FLIGHT MACH NUMBER

VAN NUTS. CAiiFOINIA


Page 5?

Eng. No. ii

FAN OPERATION THRUST

150 -

30 1 { !: I i:l i' 1 ! 1 i f ; [ . I ; :i t: _PLENUM BURNING (_P = O)t , ':

! '. :: :i i i : j,::: :::; :::: :,'. :::: I ! ": I I t

! t::,: :.::.J:.,: ,"_L2I::ALTITUDE, I000 Ft .

'_ _!,]hi_ii!:i:I:_4,oiri:,l,_i:i0,::i_;l_ii-iii_li:i:!:l:ii: l , I !

200 _ ,:::st:/:151!t;ii:.ii:!!::::;ilii_::i:: I]! _i::

= _iit331t:i:H_3.:t:.--3T.33iFr:_!t]3i:l3T_:!':::rr_!;-:3::i: 71_i :I: !3.5o 13t: 1_:::::3337!_ii:fi:i!t!::::it!!!:.t3i3-t:.:t:i; :433!:i i _ i I i

i:::Nj:l:;iil:iiiti_i;l;7:::tT?iit+:i:t::Li::i;i !I:;i-:::! I: : i :..... _ti3:3lTii::':ii{!iii':i::3!i::-ii :77i37":{:7_17......i_:- ,

i: i_i]_iii,o i l;iiil;Lli/l i;l io - o -_:_!-.--_ii_-_?_ri_i_.-_.-:::_i__

0 0.2 0.4 0.6 0.8 i.0

FLIGHT MACH NUMBER

VAil NUY|, CALIFORNIA


Page 53

Eng. No. tl

FAN OPERATION SPECIFIC IMPULSE

0

5200

4800

4400

,2Lt),,.I:Da.

m

IJ.m

t_I,,1.1a.0'3

I--LU

I--LI.IZ

4000

3200

2800 :,:

2400

20000 0.2

.... , ::: ::; .... -_.........

.... ;_i ]_:!...........BURNING (4, =

:::::::::::::::::::::::::::::::::::

,.,a 4

0 20) ::" ::::::"..... ::_.:_

, I000 ft

_36.2

_>20

0.4

,_ _i!!_:il

.... = ill_ _''' "[, :I:;::

:: ;..": ;,,_..::: :,,_ . ,-tr .....

• ..H ..... **-.

_'_'! '_ "_' ';"_'_i

0.50)_ i_:_!:i_

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.... _::1 .... :1:.1:

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0.6 0

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!ii:....

4:= :.:."

.8

:::i!i

i_i ili._, ..,

:T: :-; ;

= i

=2 .,:;

:T: ::::

1.0

10

2O

FLIGHT MACH NUMBER


Page 54

Eng. No. 11

FAN OPERATION THR UST

300 -

200 -

100 -

i

o 00 --0

400-l--I.U,--)

I--I.U

z 300-

200 -

I00 -

O-

60

00

= 0.20)

0.4 0.6 0.8

FLIGHT MACH NUMBER

40

VAN NUlrlI. CALIFORNIA

oft 25 194Volume ?

Page 55

Eng. No. Ii

FAN OPERATION S PECIFIC IMPULSE

¢,)

i

I

_I

¢%.

i

¢,3I

u.i

¢.)UJa.el')

I-.U,.I

I--L,IJZ

3000

2600

1800

1400

2600

2200

1800

:.: l: : :::::::::::::::::::::: 1:::l::::1:. I: 71::. i..::l:;: ; :: .... :: :.I :;

! LENUM BU NING (_= 0.75)!:_ :t:.:l._::_:_z-:

:_::_!!_!!_ii_:_!_:i::::_::_::i_ii!i!_i_!'_i?:_::ii_::i:_::i:!_:.!_:_i_!i::_i:::,,ii__ _/i!_.!!..................................................... ' ...._......... _ihiii:-:_....... :.l_i::.l_iiili:::!li::iiiii:.ili:.iit:.ii!tiiiF_i:iiii!ii::!it!iii!_ii_i: i!::_tiiii::ii_:_:_:.i_tlilil::!::iF:ii!li_!!ti_iilii!iliiiitiiiliii_iti::!_tii:!'__fiti_:_::I _:::liiii:':I: :::::::::::::::::::::::::::::::::::::::::::::::::::::::::::_,::::_:.::,_:::_::::r:::_:_::_:::'::::'.:::":_": :_ _ _ i!ii.i!i._!i!if_i::!

2200i i Ji_:1iitiiiil_:-::::tiii::ti!iit::ii::r:_ili::ii:l:_ilitii?_ / ..,il_:i_i__i::iF:iii............

ALTITUDE_ 1,000 _ff t: ::!;::::_ =====================_!ii:: :.,,. :..:.:.::::t ::,:_:: :::_:_.,::::

i''ii iii:llli:iiliiiitiiitiiii__i :l;iiii::;i?If.:-:.-,-:_:_

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::i::l:i:::l_ilil::_iii!:._il::i::il_ililii::::Iiiii,::i::i!i_Liiii i:iiliiii_i_!ii::i iiii._::ii-_ii_L_:_

PLEN ,,L_,,BURNING .(_ = !.00) _::ij':_!::iiii:.-_

: : I! .:1 ' : ;:!i :": ::::::::::::::::::::::::: ::;:t',::; :::; :::: "::: ::::......._............................_........, ........_!!i]:,_:"_::..,_:.,_I:_.._I:_:_I_i_::_::::::::::::I:-:::,.....I:_:.I:::;_:::I_::__i:._.!h.i: :t : X.i_:::t::iiilii::iti::i::l::i!!L_i::_F:iiit!ilit:::::.itliiitil _it_:_.i:_i_ _i!,::i::___'ii_

_.,_r..,::h:: :',::::'.:::_':::

ii::il:/:i::l_i:::_::::M:::i'ii_!i::i::_i::i_iiii!::::;:ti::::t __i"_ _ _::ii:_i!ii!i.i_i_ ::F.ii_i::_:i::i::I::,,_:__t_._t_i!_::i!!_:_i,_i: _i _:::_::_" _i_ i::it:_-

': ..........t":"'*": .....i_i_!i_!iiii'::!:.it!i::!t!i_t::]!itii!::li_i:_!:::_: !_[i!_._-_!.....

i:_. ;:-."1::::_::::- ,_,.-:_._ :_::_::::::'.:!!i"_b'_i ::!i':_.........' ..................:.

.... _ ............... i::'-"_t:::_J:_: :_:::,I=:i:::_ii_ii!Jiiii _!ii ::.: _' ................ ::;: .::_::_', :.*i _:::l.'_:_ '+_ :::; ::::'.:--',::H ";_

!_'-:ih_.:::' _::_:'_. ............

0.2 0.4 0.6 0.8 1.0IO001SIt_ ""_ ::.,i:._..":i_i!_:_.,,HiTi'_'i:=,,_:E_ !!:: ::;

FLIGHT MACH NUMBER

V4N NUV|, CAIIPORNI&


Page 56

Eng. No. ii

FAN OPERATION THRUST

e-

.9,°

000

i--

re.-mI--

I--

-.j

I--

Z

600

4OO

2OO

0

800

600

400

2OO

0

120

8O

4O

i

o 0OOi--(

p-(n

= 200-r-I-

F-LU

160p.WZ

120

8O

40

00

_i_ii__......_ _

1.00) ,

FLIGHT MACH NUMBER

UNC_SSIFIE_D_,-

va_ NUV|. ¢&ilFOIl_ea


UNCLASSIFIE_

VAN NUY|° CAAtQOINIA

Report Z5 194Volume 7

Page 57

Eng. No. II

SUMMARY: SENSITIVITY ANALYSIS - BASES

The performance data shown in the present report, as well as in the

Class 0 and 1 engine documentation (References 1 and 2), was computed

on the basis of a singular set of internal component efflclencies, as

well as stated operating points (e.g., design mass flow ratio, Ws/Wp).Component sensitivity studies were conducted as a major effort withzn

the Class 2 study phase. The bases for the analysis are given here,

followed immediately by the results.

The approach used was to define baseline performance, specific impulse

and thrust (both net Jet), for a reference trajectory. This was

accomplished for each of the engine's operating modes over the normal

range of flight velocities for that mode. It is appropriate here, to

comment to the point that sensitivity studies of trajectory effects,

per se, are already intrinsic in the previously displayed performance

maps.

Proceeding from this basis of specific impulse and thrust discrete

trends, each of the important engine variables was perturbed from the

baseline value, e.g., afterburner combustion efficiency: Baseline

value - 0.95, sensitivity excursions - 0.85 and 1.OO. All of the remain-

ing variables were essentially held at the baseline, or nominal value.

Any exception to this resulted from the engine performance computer

program's automatic compensation characteristics which, in some instances

"retunes" some of the engine internal variables. The extent and impli-

cations of this situation are covered in the main technical report

(Reference 3).

This section presents the following base_.___sfor the sensitivity analysis

results to be given subsequently:

i. Reference trajectories

2. Baseline specific impulse (on reference trajectory)

3. Baseline thrust (on reference trajectory)

_. Ranges of sensitivity variables, with reference to the baseline

values (both curve and tabular presentation)

VAN NUI'|. CAlfllOaNIA


Page 58

Eng. No. Ii

REFER ENCE TRAJECTORYSENSITIVITY ANALYSIS

120

_ = I00

80 /

I--

/40

/20

0 1 2

I I0 0.5

//

3 4 5 6

FLIGHT VELOCITY, 1000 ft/sec! I I

1.0 1.5 2.0


7

/

8

I2.5

VAN NUF|, CAL|FOIIN|A


Page 59

Eng. No. Ii

R EFERENCE TRAJECTORYSENSITIVITY ANALYSIS

FAN OPERATION ONLY

5O

f,/}

E

0

LI.I,,-,,

I--i

I--.-I

12

8

4

0

'..=_

=o

m

,3

I--

,-I

4O

3O

2O

10

00

I

0

J

0.2 0.4FLIGHT

I

0.I

FLIGHT

//

/

0.6 0.8VELOCITY, 1000 ftYsec

I0.2

VELOCITY, I000 m/sec

1.0

1

0.3

YAM NUYS. CALIFORNIA


Page 60

Eng. No.. II

BASELINE SPECIFIC IMPULSEEJECTOR MODE

1000

900u_

E.o

,.o

-- 800

uJu3_J

¢1.:_ 700

tOIJ.m

tO

"' 600o.(/)

I--la.I

_- 500uJz

Jf

//

///

/

///

4000

L0

0.4 0.8 1.2 "I 1.6 2.0

FLIGHT VELOCITY, 1000 ft/sec

I I I0.2 0.4 0.6


2.4

I

0.8

2.8

_T_rTF.rr;_::]FrT--_nrq

VAN _UY$. CALIFORNIA

Report 25, 194Volume 7

Page 61

Eng. No. 11

BASELINE THRUSTEJECTOR MODE

2.0

U1e-0

e-

: 1.50

E

m,,-i-I--

I--i,i": 0.5I-=i,iz

0

500

4OO..0

300

200

I-,LI.I

--I-'-

I00

- 00

I0

0.4

/

f/

f

0.8 1.2 1.6

FLIGHT VELOCITY t 1000

I I0.2 0.4

FLIGHT VELOCITYt I000

2.0 2.4

ft/sec

0.6

m/sec

I0.8

2,8

___/]rquarez ,,. _..,,,.,.,.

It,ltmR_A_i_


Page 6Z

Eng. No. 11

BASELINE SPECIFIC IMP ULSEFAN RAMJET MODE

4OOO

t,,,}

3600

_. 3200

/i

I.u

.J

m

m

u.

uJ

u')

k-

p-wz

2800

2400 /

2000 /

1600

//

/

/

12000

I

0

0.4 0.8 1.2 1.6 2.0

FLIGHT VELOCITY, I000 ft/'sec

I I I0.2 0.4 0.6


2.4

I

0.8

2.8

VdN NUYS. CAtlFOIINIA

Report 75 194Volu_ne 7

Page 63

BASELINE THRUSTFAN RAMJET MODE

Eng. No. ii

1,4 --

1.2

O,mJ

1.0c-

o

E 0.8

U3

e,,-rI'-

0.6)--UJ

I-.-uJZ

0.4

0.2

0

28O

240

J

w.,

°o 2000,-4

B

_c0'3

160ew

-_ 120

M-WZ

80

- 40 ¸

|0

/

0.4

//

/

0.8

FLIGHT

I0.2

FLIGHT

//-

/

//

/

1.2 1.6 2.0


! |

0.4 0.6

VELOCITY, 1000 m/sec

/

2.4 2.8

I0.8

VAN NUY$, CALIFORNIA

i

}_port 25, 194Volume ?

Page 64

Eng. No. II

BASELINE SPECIFIC IMPULSESUBSONIC COMBUSTION RAMJET

_000

3600

3200

2800

2400

200O

16000

LI

0

i/

f \/ \/ \

/

I0.5

\\\

2 3 4 5


I I1.0 1.5

FLIGHT VELOCITY, 1000 m/see

\

6 7

I2.0

\

8

I2.5

VAN NuYS, CALI#O|NIA

Report 25 194

Volume 7

Page 65

Eng. No. 11

BASELINE THRUSTSUBSONIC COMBUSTION RAMJET

1.2

(nm-

:I.0c-oo_

E

I--'0.8or)

m,,-r-l--

l--uJO°6

Z

0.4

0.2

0

280!

"- 240,.Qm

-- 00(Dp4

200k-or)

..p

160W-W

i,lz 120

8O

4O

00

[0

// \

\\

\

1 2 3 4 5

FLIGHT VELOCITY, I000 ff./sec

I I I0.5 1.0 1.5


\\

\

6

\

7

I2.0

8

I2.5

25,000

SPECIFIC IMPULSEOPERATION

._eport 25,194Volume 7

Page 66

Eng. No. ii

20,000

15,000

U

u_

E

I0,000

--_ 9,000

... 8,000u_

7,000

__ 6,000

,'i" 5,000

I,u

o. 4,000

_ 3,000IIW"Z

2,000

1,500

1,0000

I0

0.2

I0.05

/

PLENUM BURNING, _ -

./

//

0.2

0.4 0.6 0.8 1.0

FLIGHT VELOCITY, 1000 ft./sec

I I I I I0.i0 0.15 0.20 0.25 0.30


VAN NUY$. I_ALIFOIINfA


Page 67

Eng. No° ii

BASELIN E THRU ST

FAN OPERATION

200

io

o 150c

ooo

ee

Or)-_ 100

I--

I--"' 50

0 00 0.2 0.4 0.6 0.8

FLIGHT VELOCITY, 1000 Pjsec

I I t I I I0 0.05 0.i0 0.15 0.20 0.25

1.0

I0.30


H

rcluamr ....o,,.,.,,,o,.,.

f-k

._eport ZS. 194Volur_e 7

Page 68

SENSITIVITY ANALYSIS RANGES

Ejector Mode:

Inlet - Pressure recovery, Pt2/Pto

Fan - Pressure ratio, PRf

Primary

Equivalence ratio,

Combustion efficiency, _ c*

Nozzle efficiency, W N

Mixer - Mixing efficiency, _ M

Afterburner

Equivalence ratio, _AB

Combustion efficiency, _c

Exit

Nozzle efficiency, N

Exit area ratio, A6/A 5

Fan Ramjet Mode:

Inlet - Pressure recovery, Pt#Pto

Fan - Pressure ratio, PRf

Afterburner


Combustion efficiency, _ c

Exit

Nozzle efficiency, S N


Subsonic Combustion Ramjet:


Combustor

Equivalence ratio,


Exit



Fan Operation:

Inlet - Pressure recovery, Pt2/Pto

Fan - Pressure ratio, PRfAfterburner - Combustion efficiency, _ c

Exit - Nozzle Efficiency,

Eng. No. II

Base-Iline I Range

Figure A

1.30 1.50 i.i0

1.00 l.lO 0.90

0.975 1.O0 0.92

0.98 1.O0 0.95

0.80 1.00 0.50

1.O0 1.50 0.50

0.95 1.00 0.85

0.98 i.oo 0.95

Figure B

Figure A

1.30 1.50 i.i0

1.00 1.50 0.50

0.95 1.00 0.85

0.98 1.oo 0.95

Figure C

Figure A

1.O0 1.50 0.50

0.95 1.O0 0.85

0.98 i.oo 0.95

Figure D

1.oo 0.95 0.9o

1.30 1.50 1.100.95 i.oo 0.85

0.98 1.oo 0.95

VAN NUYS. ¢A&IFO4PIIA


Page 69

Eng. No. II

Figure A INLET PRESSURE RECOVERYSENSITIVITY ANALYSIS RANGE

I--

¢M

m

>-m,,u.i

Q¢Jhi

I,Im,"

O0O0I,Im,"

1.0

0.8

0.6

0.4

0.2

00

I0

RED I ICED \ _SUBSONIC- \ X

REGIME

RECOVERY -

BASELINE

\',,,

\

\NORMALSHOCK

ID

MIL-E-5OO8B%\ -\

7/KE = 0.9_" _"

1

]0.5

2 3 4 5

FLIGHT VELOCITY t lOOO ft/sec

I I1.0 1.5

FLIGHT VELOCITYt I000 m/sec

6 7 8

I2.0

rqu ,_ .c°/1#/m.At_r_

•port 25,194Volume 7

Page 70

Eng. No. Ii

Figure B EXIT NOZZLE AREA RATIOSENSITIVITY ANALYSIS RANGE

EJECTOR MODE

2.8

_D

2.00m

I.--.¢

<:1.6ILl

<

I-"

Xl.2hi

Ul.JNN

°008z

_ v

i

If

,a

.1.25 BL

i/ _ _ • BASELINE CBL)

• f

/ .._ ..L... -_" 0.75 BL/ I

/ /

ql, '"'

0 o

|

0

0.4 0.8

I

0.2

1.2 1.6 2.0FLIGHT VELOCITY, 1000 ft,/sec

I I

0.4 0.6


2.4 2.8

I

0.8

3.2

I

1.0

VAN NU_|. ¢AL#FOIN/A


Page 71

Eng. No. II

FIGURE C EXIT NOZZLE AREA RATIOSENSITIVITY ANALYSIS RANGE

FAN RAMJET MODE

2.8!

2,4

U3<:

,,o<:

,.2.00m

I--<:ev,

<:16IaJ "ev"<:

I'-

X,.,1.2

-JNNOz0.8

/

J

J

//

/¢

/ f

/ /

/i / // ./

J/

Jv

I

•1.25 BL

" BASELINE (BL)

• 0.75 BL

0.4

I

0

0.4 0.8FLIGHT

I0.2

FLIGHT

1.2 1.6 2.0

VELOCITY, 1000 ftv'seci I

0.4 0.6


2.4

I

0.8

2.8 3.2

Io0

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

,eport 25,194Volume 7

Page 72

Eng. No. Ii

Figure D EXIT NOZZLE AREA RATIOSENSITIVITY ANALYSIS RANGE

SUBSONIC COMBUSTION RAMJET

20U3

,<

,,D,<

,,16Om

I--,<==

< 12ILl

I"

X,,, 8

IJJ--INNOZ

4

00

I0

1

fIp

2 3 4 5


I I

6

12.0

7

0.5 1.0 1.5


8

I2.5

Report 25.194Volu__e 7

Page 73

SENSITIVITY ANALYSIS - RESULTS

For the reference conditions stated in the previous section, resulting

specific impulse and thrust perturbations are presented here. Performance

is normalized to the baseline trends given over the appropriate flight

velocity ranges.

The specific impulse and thrust data are both displayed on individual sheetsfor each sensitivity variable. On the same sheet, a miniature plot of the

absolute specific impulse and thrust baseline characteristic is shown for

nominal reference purposes. For precise readings, the full-sized curve

appearing previously (its page number is indicated) should be referred to.

The section concludes with a plot reflecting subsystem weight variations

on uninsta/_led engine thrust/weight ratio.

van wilY|, C4¢O#OINIA


EJECTOR MODE

Report 25,194Voluro e 7

Page 74

EFFECT

kq. If, U _I¢ii

Page 60,....,.,,,..,,.Eng. No. II

|JICl$I ImM

Page 61Og$|Lm! t_m$1tJt_O| moot

1.60

'-'1 i i i i ! i i _ I _ I" I

i__[: _, ! + I : _ I/ i lj-_Vi _ i I : ; i 11,,"] i i}i_l_ Li i I iYL [ i i Ii I J I [ i [IY' ! ] i i 1

i +, tfl i[ i i I tl1.._ i k VE i { ; i i [ i I| ;

'. , i.!. i!, I,+I :. _,!._,!.l,_It w_ll_,, NIl lml

++ ,_ ,_, ,.'+ .,',

I,i..i_ ._lll_ iml

Baseline

PT2/PT0:

Figure A(Page 69)

g" I

I ='+F a __

: r'llHi>-T, iiiii i ii '.'_-l._--+'T t I I I I + I ( I I" i_ I I t I I I I l.J,I l I I I l

..,_-,,,I 1 I I 1 I I ! i ) _ 1 1 I !i " I

i,_iI lll,ii_. _ml ,,.i

=m

1.40

1.2oa,.

_= 1.oo

_1_ 0.80

]°el) {,0

.

1.60

1.40

I,-1.20

==1.oo0 80

O3

0.60

0.400.5

L I

0 0.2

= 0.92

NS __

-- ,gKE = 0.92

1.0 1.5

FLIGHT VELOCITY_

I

0.4

1000

2.0

flVsec

0,6

2.5 3.0

0.8

FLIGHT VELOCITY m/_c

RATIO

MODE

EFFECT

Page ? 5

1.15

I.I0

1.05

1.00

0

0.90

.

1.15

_: I.I04;O_

M-u_ 1.05

= 1.00F-

F--u_ 0.95

O(

F- 0.90

Page 60

'"'"" """' '-'" EngIJ|=TOU _N No. 11

,m i II[ i: : i[Ik_l1_J _ } j _ ; _ I J i/I i I| I i : i i _ i [ ] [i/1 i [ Im...J I _ i i £ i [ /,'I I ] f It It : ii i il/i !I] II|,,,[ ; ; [ [ I L_'! i L I = I i, [i ; ! ! tfi t I I I L :1!.,1 ! i [ I Vi [ ! I [ I I i t:' ! i !_ I I i i i i I _ I_ [ I :_R'I i I I i I i ! ! I

Baseline

PRf = 1.30

!

J! so

I.I0I

1.50

0"850 0.5 1.0 1.5

FLIGHT VELOCITY, I000

I

0I I

0.2 0.4

Page 61

gASltlml rw_us!(JlcTom mo|t

li.ZI i.JP'_

' i""I "F?"

eL

Ii [I IIIll i ! ill

I ! L 1 ; i t L_I_ i J ! i

I I I 1_I_ ! I [ i ! t i i-_li i i I I ! i ! I l

ll!l!liil_i_!!LII!!I llli _ [iiI.!.I.LI!',!.,?._!._.!

mu_ _. ,m e#m

2.0 2.5

ft/sec

I

0.6I

0.8

3.0


lq

1.02

EQUIVALENCE RATIO

Report 25 194Volu_ne 7

Page 76

EFFECT

EJEC TOR MODE

iml. _ 11

Page 60_S|LII SP[¢IFI¢ l.gVL||

IJlCI_ m@M Eng. No IIPage 61

IASILIU! tUmSl[J[Ct@llmOllI

--l..J

¢._ C.)

r_ia.

1.01

1.00

0.99

llmO

i_

i-II"

i_il 'liILil

] I Y; i I _ i_i i I I i I I

i!iL!lllli ._,,'" . : i L ! ] i i I I

]iiil_ll!!l

# ._ ... L,' _,

Ih

V

Baseline

4,= 1.O0

I.I0

me , ,

"F ._

|_.r!..(I({((_ I , _ I I I|$ I I I ! IAII I I I I I I 1 i

I "t'|.-_-_--_ I I ! I ! I i I I ]. (_ |JlllJllJ JJjJlJ

'"!i"l ' I I I I I lLill t

.C ,t '.!. _.!.',_,I,!._!. I!. I,.

m_Dw ,e_mmw wu .,i,

0.98

O.

1

. ._ ..

__.._ .._ 0.901

,2

k-

l--

p-

k-

1.02

1.00

0.99

0.98

0.97

|0

_y

I.I0

.5

f

1.0 1.5

FLIGHT VELOCITY,

I i

0,2 0,4

1000

0.90

2.0

ft,/sec

I

0.6

2.5

I

0.8

3.0

FLIGHT 1000 m/see

uard! .. +, .,oo°_.' II,Y_/R4/'X_

PRIMARY ROCKET COMBUSTION

EJECTOR MODE

EFFIClENCY


Page 77

EFFECT

I.,i.I¢.¢1

_+n-,

1.06

_I.04LUU_

l.o2__ 1.00

,r 0.98

_ o.96_

0.941.06

_: 1.94el,

l--u'l 1.02 -=3m,.= 1.00l--

h-0.98

,w

:= 0.96

0.94

&0

Page 60

_SlAJN SPI¢IFtC _MA$|IJ|GTg|14111 Eng. No. 11

Baseline

7/. =0.9"/5C

_JI

JI

1.00

I

1.00

0.5

!F;'I!+If-:

&

Page 61

IAStLII! i_nsvll|ci@lm@D!

_ mmmmmm_ mmmmmmmm m

1.0 1.5 2.0

FLIGHT VELOCITY, 1000 ff/sec

I ! I0.2 0.4 0.6

FLIGHT m/sec

2.5

I0.8

3.0

PRIMARY ROCKET NOZZLE EFFICIENCY

EJECTOR MODE

keport Z5 194Volume 7

Page 78i i

EFFECT

Page 60

""'"''"""',.°--- Eng. No. 11

'--i t : i i i t t I I [ I I,_1_!IL!!!111_ IIj--I I i ; I ! ! i i I _ ] ! Il_l i ! t j i i ! iYl I t I I

;-I i i i [ lit i t I ] I J I!_l ; i i I VL [ ] I f ! i I I=--r; i i _ ; i i ! I _ I i t_lI: _ lliilili]li _F"T"_I i i i I J

.[ i.]._ .[,,_ ,[. ; _.r,!. I.!.

Baseline

?7N = 0.98

Page 61I_H&III t#mu, I

l!I. ' " Iz"z i 1 _'_

J• .,'! I "

D.

¢,JE

Ji.o2

__ 1.oo

E 0.98

0.96(/1

0.94

1.06

-- 1,04rs,

_n 1.02

= 1.00h-

0.98

:= 0.96I---

0.94

li_lia_m m _miil_m mmmmm_a

j_

_ m m _lmmmm _

_.,I

_.,,,,... _

|0

5 1.0

FLIGHT

I

0.2

FLIGHT

... 1.00

0.95

1.00

0.95

1.5

VELOCITY,

|0.4

i m

mu

1000

1000

i

2.0

ft/sec

i0.6

m/_ec

---- ...- ..__j--__J

2.5 3.0

I

0.8

VIN NUYS. CIIJFOIIN/A

MIXING EFFICIENCY

EJECTOR MODE

Ixl el, U

Page 60

"'"'""'<,,.,,,........ Eng. No. II

i:

i

I I Z [ J ] _ i i i,g ] ] Ili ! i i I _ i]/ !l ! lI]i i ] ! LIF] Jillri i i f i ;_,'1 l [ I I Ili;IIl/ll; Ill]_l _1 i i ; I I I I[,]" 7 [[[1 ]tll

!i.!.i,!,i,Sl,!.',l,!.I,}.

n.,e..m_ *m

Base Iine

7/a = 0.80

EFFECT

Report 25 194

Volume 7

Page 79

Iq. _ 11

Page 61

IASILIWl fMllSTIJICTO| mO0¢

=.i r

I i,"I*.+!.

"P :+.o!i.i IL w

o i

I l 1I I', I'G_,_'! _I

1'!11 _1111I_ _1o., I.i t._ :.+ ;l z.4 z|

• '- ,.+, ' ,'.= ,.iiqmlln% lm _

T(l) ¢/)-J-- 1.01

0._.

1.00 "

0.99

0.98

0.97

1.03

-- 1.02O:

k-oo 1.01

= 1 O0I--

I--0.99

-T- 0.9p-

I I

0.97 0.5

.I0

1.00

0.50

1.00il

0.50

1.0 1.5 2.0

FLIGHT VELOCI_, 1000 _c

I I, I0.2 0.4 0.6

FLJGHTVELOCi_ )00 _c

i

2.5

I

0.8

3.0

AFTERBURNER EQU IV ALENCE

EJECTOR MODE

RATIO

Report Z5,194Volu_e 7

Page 80i

EFFECT

f_ktl

Page 60 Page 61

""....'""'""' Enq No 11I)|¢141 lIM I_SCLIM IN_LST• • ¢JI¢TN IIO0!

1.15

_1.10

1.05

1.oo

i: Baseline . _llt _(_}

i i i I IFi I _ I i I1 lYl I I I lL ].Icq i I J ¢ 'I _¢_ , [.I II I I I I I LcI_i I I

_. j_ I I 1 1 n..¢'l I I I _ I I I II![ I [ Ii/,i! i ; I I L [1¥' _ i I I I I I i_ffl--_ I I 1 I I I I I i

i;-iltl i Illltilll; i ' l/_ i i i I l t IJ i I I I I 1 I I I I: ._, _ ii_ ..'_"ri-Illli )llli_lll• i I I I I

" ' ":L. '_._','_.,".L ...... ,_-",-,_'-- ',f- .....=:. I _ ,-'o ,L. .'-

0.70 .5 1.0 1.5 2.0 2.5 _.0

FLIGHT _LOCIW, 1000 _c

I I i I i0 0.2 0.4 0.6 0.8

¢1 IP.MT Vl:'l I'IPII"V_ l l_fI m/_u=e

AFTERBURNER COMBUSTION

EJECTOR

EFFICIENCY

MODE


Page 81

EFFECT

u'}u')--I...I

u'3_

1.03

1.02

1.01

1.00

Page 60I_$1LIM SP¢Ct;JC _ULU

EJlCT@ll alOM

, i l!_,ll I!/_[!• IL i T_ I]

i ! iili_/_ If!i r W'!L i]_

! *i i/! _II: /;iI_ J{{

S i ' i i/ ' [ ; { ',I i

iTLii ,i ] LI_i.* a.,

Eng. No.

Baseline

= 0.95C

11

1.00J

I--c¢}

l--

c_

I--

O.

O.

0.

1.03

-- 1.02

1.01

1.00

0.99

0.98

O.

",.,.w

0.85

Page 61_A$|L_mE 1_lusl

1.00

i

0

0.5

__ 0.85I

1.0 1.5

FLIGHT VELOCITY,

| 1 ,,

0.2 0.4

FLIGHT VELOCI

2.0

1000 ft./sec

0.6

)0 m/sec

_mm

2.5

!

0.8

:5.0

VAN Nsrr$. ¢AJlPOQMMi

EXIT NOZZLE EFFICIENCY

EJECTOR MODE

EFFECT

i---..

Report Z5 194)

Volume 7

Pa. ge 8Z

1.06

N

-_ 1.02Q.a.

_m z.oo '

0.941.06

_: 1.04

1_02

z.ooF-

0.98

= 0.96F-

0.94

"i, iU

Page 60

'"""'"''',.,,.,-., Eng. No. ii

l,,r:

!.i.!I"11

[

0

.5 1.0

FLIGHT

I III

0.2

FLIGHT

Page 61|_S|LOml _N|UST

I,I[¢tlle liOl) !

Base line

_/N = 0.98

. "II I I II

I [,,,,I I I I I I/ = I!I111

,_--_.1 ] I ! 1 I|' Illl_l

fll![[tl] I ] L,,,,_ I IIL_]]i]

,,"f"q I I I I _

iii tii,_,L_,_,._,I.,'_,!.

1.00

0.95I

1.00

0.95

I1.5

VELOCITY,

i

0.4

2.0

1000 ft/se¢

I0.6

O0 m/=ec

2.5

!

0.8

3.0

EXIT NOZZLE AREA RATIO EFFECT

Report 2,5 194Voluz'h e 7

Page 83

EJECTOR MODE

Ilq, m U

Page 60

|ISIL¢_ SP(¢Irr{ r_H,S|IJlCT@l m.N

Eng. No. II Page 61

|AStLIUl rwmusrEJlgIOi moll

1.06

I_ 1.04

:r:_ 1.02

Q,. l_.

1.00

c,,)(J,TIE 0.98

I_LIJ_._. 0.96

0.94

1.06

_: 1.04n,,

k-u_ 1.02

" 1.00p-

F- 0.98u'}

" 0.96p-

0 °94

"l ] , : : ! i i i [ ] l [ I I,dt ; i L + iI iv,'rll}-II _i i /li_II ] ] ] t i t !/1 I I I I

W--I C i i I _ i IF'I I I I i 1|..! i + ! t . ',,,t"l I I I I I ], |_, _ 1 I I /I _ [ t I I i I!,.Fh i / l _,"l i i I I I I I I

+-_Fi i!_ll_[ [ ! 1"t' ,.+ ..,' ,+ ,. ,]* ,., t,

m+.l_ ,,ll,l_, lul _,ii

euml, _qumw. imlO._

Baseline

A6/A5: I"[",.,kl.

(Page 701 i'" "i.| _Lw,

.l.

i

J

+ IlItrl i]

t i t t.Jr'l t t I + I _ ]-b.b.-,_ I I I I I i i I l]tllll_tttt ttllIlii I ttl k t ]

rli(E iII[ll]]IIt]! !l

0.5J

1.25 BL _,.,

r-

/

f

._ / _'_ _ Oj 75 BL

1.25 BL! _,

f

__ _?__ 0.75 BL

_-- ---._.

1.0 1.5 2.0


2.5

I I I | I

0 0.2 0.4 0.6 0.8

3.0

FLIGHT VELOCITY 1000 m/see

INLET PRESSURE RECOVERY EFFECT


Page 84i

Page 62|IsCLtm[ SPI¢OFOC JI_'ULSt

_,le nAmJt*l uol!

FAN RAMJET MODE_IL II

Eng. No. II

!

1

Baseline

PT2/PTo:

Figure A(Page 69)

1.60

_o_

_uem_,m_

mu__ 1.40

1.20_ 1.00

u. 0.80 m /__0

_'i [" /a. _. 0.60 f

0.40

1.60

Page 6.3I*S[tOmt I_IUS_PA_ 1,lJll lOlL

i n _ l i

"; '°,, 'o2, _,J ....

A ,I, ,'.. ,L, v.,

1.40

1.2o

1.oo

F- 0.80

= 0.60

0,40

/

.5

/

1.0 1.5 2.0

FLIGHT VELOCITY, 1000 ft./sec

2.5

1 I | I I

0 0.2 0.4 0.6 0.8

0.92

3.0

FLIGHT _/ELOClI_I_O m/see

VAN NUVS. CAL|_O|NJA

FAN PRESSURE RATIO EFFECT

Report: 25,194V 01 _Z.I%'3e {

Page 85

I

FAN RAMJET MODE

Page 62 _" "

..........,..,.-,,,'*..........=,," Eng. No. 11

, ,:,_-',..=--"-:-

Baseline

PRf = 1.30

L.i,

I

i i.o

_t

L

Page 63 .....IA$I_I_I twIi sir_AII llal, l_ll MOOI

i,i

,,q

1.60

1.40

"' _ 1.5o1.2o

'_ 1.00 ---- .....m

,;" 0.80

o6o /0.40

1.60

-- 1.40

_ 1 20_ •

_" 1.00F--

F.- 0.80--(n

= 0.60

0.400.5

/f_

/ 1.10

" I1.0 1.5

FLIGHT VELOCITY,

u_mi m

p j--------

2.0 2.5

1000 ft/sec

3.0

I0

I

0.2i

0.4I

0.6I

0.8


EQUIVALENCE RATIO

FAN RAMJET MODE

1.60

1.40

1.20

:_ 1.00

Page 62 _"= "

IASlLlUt SPI¢I_I¢ tmp_Ls!FIk I&mJll ao#!

Eng. No. II *";J

I i,

Baseline :°'_

(_AB = 1.00 _°4I

0.50

1.50

Report ?.5 194Volu/In e 7

Page 86

EFFECT

Page 63 .....I_$[LINI l_lUs?PA_ IA_J[T moo(

_Cl 40"

J,vI'--u_ 1.20

n,,-r-_. 1.00

0.80

,Y= 060I-"

0.40 0

i0

1.50

0.50_ _ / III|

0.5 1.0 1.5 2.0

FLIGHT VELOCITY, 1000 ft/sec! l I

0.2 0.4 0.6


2.5

I

0.8

3.0

V4N NUI'S. CALIFOINfA

AFTERBURNER COMBUSTION EFFICIENCY


Page 87

EFFECT

I..1.,I

E..

f,.;

(,/3

1.15

I.I0

.j 1.0

O.

O

0.95

0.90

0.85

1.15

Page 62

|,slt_.! $PI¢_FI¢ ,IPULS!_Am mJm_(l mOH

FAN RAMJET MODE

Eng. No. II

I"- Ba se!ine _"

;'- _/ = 0.95,. C

v

Page 63 .....|As|_im¢ IwtuS!_i_q mAa*J[l" _oo(

M.o, u. _m

1.00___ BIB

0.85

,.24Jo_

p.oo

c¢

p-

p-u$

I.I0

1.05

1.00

0.95

0.90

0.85

I

0

1.00

0.85

0.5 1.0 1.5 2.0

FLIGHT VELOCITY, 1000 f't,/sec

I I I

0.2 0.4 0.6

FLIGH_ m/sec

2.5 3.0

I

0.8

VAN NVV$o CAUfOINIAi I


FAN RAMJET MODE

Page 62 _"" "I_$[_INI SPICIFIC ,ImPUL$!

i,s, |AmJf_ tool)!

_an ,m,a_ry° _m

_ .e.lo_ Lea.

Eng. No. 11

Baseline

_N = 0.98

iI;.!

i

.L

1.06

J

0.9

0

O.!

1.06

.: 1.04

_-_ 1.02

1.00

_" 0.98

= 0.96

0.94

I0

S

I

0.2

1.0 1.5

FLIGHT VELOCITY,

I0.4

1000


Page 88l u l i

EFFECT

Page 63 _,_,,fill&Ill l_lUS1IAh IAmJ(1 mool

1.00

i

LI

0.95

F

1.00

0.95

112.0

i0.6

FLIGHT VELOCITY 1000 m/sec

2.5

!0.8

3.0

VAN MUV$. CAI.IFOilNM



Page 89

W,

1.15

I.I0

no.

',JUJ

u_oO

1.00

0.95

0.90

0.85

1.15

_: I.I0

k-_n 1.05

"_ 1.00k--

_" 0.95O0

_" 0.90k-

0.85

Page 62I*$1Lum! SPiCb_l¢ aM*u_s(

_a_ mAIJ[t MOO[

FAN RAMJET MODE

b_ N,L II

Eng. No. III

Figure C(Page 71) I

1.25 BL,/

/

1.25 BL,_

//

0.5

I

Page 63 .....

F*m eAMJ!t mOO[

"__0.75

_ "'_'_ _ 0.75

BL

1.0 1.5 2.0 2.5


BL

3.0

| I I I |

0 0.2 0.4 0.6 0.8

FLIGH_O m/sec

Report 25,194_tA_ Volume 7

/A""" __,.,.,,,o.., . .


1.30

1.00

0.80

O.7O

1.30

,.:1.20

u_ I.I0

a

I"

i11.

_m

mw

1.oo-----

090== o.8oI-

f,q. _ II

RAMJET MODE

Page6b Eng. No. 11|A|I&INI SPICOWI(: ImPULI!

Sm|SOeI¢ ¢omluSllom |*mJcT

m m mmmmmmmm m

Baseline J,*

PT2/PTo:II

Figure A ...(Page 69) ._

_._ _.____.___--.__ _

_ _i _'"- I?KE = 0.92

i,..

Page 65 _" "|AStLll( twlU$t

IMllOml¢ ¢OlOIIIllOI |AMJI1

MIL-E-5OO8B

_...--

MIL-E-500

= 0.92

v

3 4 5

FLIGHT VELOCITY, 1000 ft/secI t

1.0 1.5


O. 70_ I 2

I0.5

I

0

P

_J

J

8Bi

==,--

_JV

_a

6 7 8

I2.0

I2,5

vJN NUY$. ¢ALJlQmN|4


Page 91n, ,

AFTERBURNER EQUIVALENCE RATIO EFFECT

RAMJET MODE

-,- ,, Page 65 .....

Pag_6_. Eng No. 11 '"...........SU|SO_lC CBM|U$110N IIJ*iJ(Tla|tLim( SPfGI_C I_pULI! •su|somtc COmOUSl;ob tJmJ[T

) i

ll.m t _

| ''

_-II/

1.60

atom

i,ii,J

(J} (f}

1.40

1.20 ----

1.00

0.80 "

0.60

0.401.60 ....

£., _% ;!, z:, ,],

\

Baseline

4)= 1.00

• _9_o_._

0.50

J

1.40

,w

F- 1.20

= 1.00F--

F- 0.80

F- 0.60

0.40

L

0

i i_ _ m

V

.0.50__

1

I

0.5

2 3 4 5

FLIGHT VELOCITY, 1000I I

['t/see

1.0 1.5


6 7 8

2.0I

2.5

-J

=__=I

u.

Q.U'}

.......,_,.AFTERBURNER COMBUSTION EFFICIENCY

RAMJET MO DE

Report 25_ 194Volu_e 7

page 92

EFFECT

tick I]

Page6_ Eng. No. IIIAIILI_I SPICIfI¢ IIIPU&$1$wIl_Ii{ G_I_STIO_ IMIJlt

Page 65

lJs[Llm[ 1#|uslIuoswli¢ ¢OIlUSTImI |Jllj[!

I*i IIL ii

F'it,it_lJ, I/i-- I i

• I

i i1_ I1!]1" fill]

:::: %Jl Ill

' i [_111IINI I !

i_IL I I N.! 1

t[ ' " ' I ! I N.Iiii Ltill[... III11

l I I _ I I ] l Ir • i , 1 •

Baseline

_c = 0,95

L.I,

11.*

Ii"I

i.i

*J

1.15

"_ I.I0

i.o.:

__ 1.00

-- 0.95

0.90

0.8.=

1.15

,.:I.I0q;

p._n 1.05

n.

= 1 00k- •

_" 0 95

n,= 090

m_

e_

y

m

m

1.00

0.85

1.00

0.85

J

h

I

0.850

I

0

i

I

0.5

2 5 4 5

FLIGHT VELOCITY, 1000 ft/secI i

1.0 1.5


6

I

2.0

7 8

I

2,5

VAN NUV j. ¢AiI/OIIINiA

Report 25 194Volun_e 7

Page 93

EXIT NOZ ZLE EFFI C IENCY

RAMJET MODE

EFFECT

Page 6hIAiELIm! |_(CIFIC ImPULS[$UlSOmac com|ust*oN |AmJ[T

[a_. ql il

Eng. No. II

I •

,,161:io0000 1 05.,Jlo

=_ 1.00

M_° 0.95 ----

,o,o,0.90a.o.0003

0.85

1.15

i l_U/ I! I I

i ] i /I [ i ] "_ t I_ I I/1 I I i i / N ], i/I ] E i i i i L_,I

| nil]lil/i l iI

illll[ILI ili

lllltll I I!1ttiLi!l!l!Ll

• '., L_ _.,

II

,I I\J J

II1lil

Baseline

_N = 0.98

i:!,-m,.

_m

Page 65 .....IAS[LJn[ t.mUSV

m,_,m

0.95

I

1.00 __-

-: I.I0,v

_n 1.05 __

= 1.00p-

_" 0.9500

= 0.90t--

0.850

|i

1

I

0

n

I

0.5

2 _ 4

FLIGHT VELOCITY,I

1.0

FLIGHT VELOCITY,

m

0

5

I000L

1.5

I000

1.00

.95 _---

6

ft/sec

I

2.0

m/see

7 8

I

2.5

van NUlrl, C4LWPO|NJA


Report Z5_ 194Volur_ e 7

Page 94

RAMJET MODE

Page 62|ASttll| SPl_lfI¢ ImPULS[

SU|S_O( COmOUSIIO* mAm_tTEng, No. ii

Page 65O*$tLIO0! TStUST

Su|SOilC ¢Om|USTIOI* IAmJtl

1.06

i0

i! ' ' -I

: ILtl

",:\ It1!, I_II

n NJ ! t

. i i "N.I

I I I I I

i l]L

_um. _IMeW,, *ml _

/

I0.5

Baseline

A6/A 5:

Figure D(Page 72)

1.25

i

I

i

i

J_0.75 BL

1.25 BL_". _.,.. _._ _ ...._ _

/0.75 BL

2 3 4 5 6 7 8

FLIGHT VELOCITY, 1000 ft/secI i I

2.01.0 1.5


I

2.5



Page 95

EFFECT

FANPage 66

_ :_ol .... _-_

1.60

OPERATION

Eng. No. II

Baseline

/ -1PT 2 PTO-

.00

Page 67 .....IA_LaNF r.qusr_AN 0PE_rl0_

ILl

.J

O.U3

..J

I

u.

o.u')

l.-u3

e,,

I--

l--(,f)

e,,

I,-

1.40

1.20

1.00

0.80

0.

0.40

1.60

1.40

1.20

1.00

0.80

0.60_m

0.400

I

0

0.2 0.4 0.6


0.i 0.2

0.8

ft/sec

1.0

I

0.3

PT 2

PT 0

0.95

0.90

0.95

0.90

0.95

0.900.95

0.90

0.200.200

0

0.20

0.200

0




Page 96

EFFECT

Page 66IA$[tOh{ SPIC_FI¢ imPeLS! b,& NIL n

flm OPtaAVO0_

I

:5,_ i

' 4

,.'- I _ _-_'1

I_ *.s G.4 u.* o,l _

wu4_* vl_,. _m

FAN OPERATION

Eng. No. 11

Baseline

- 1.00PT2/PT O-

Page 67 '_" "UA$ILIml iwmu$1_Ah OP[lJIIOh

iI

RI.I.

nr,o

PT

1.oo

L

Ug

,vF-

n,,

M-

U_

_AB

0.40

1.60

0.95 I_00

o.gs o.so0.90 1.00

_'- _'_ 0.90 0.50/

1.40

1.20

1.00

0.80

0.60

0.400

I0

_ jr

0.95 0.50__ "-" &l.O0

"-- 0.90 0.50

& 1.00

i

0.2 0.4 0.6 0.8

FLIGHT VELOCITY, 1000 ft/se=I I

0.1 0.2

1.0

I

0._


VAN NUIL CAL|FOmNIA


Page 97


i.s,l*

_.o**

c.#') (J')_J_l=L

m u

LLJI.U

_u

i--or)

Q_

i--

I--U')

r,,

i--

FAN

Page 66

_*. OPlmAr_O.

[F

=

...i_ } :. L, ,!, o:.......

1.60

OPERATION

Eng. No. II

Baseline

PRf = 1.30

1.40 ----

1.20

Page 67 .....IAS[_N¢ I._USr

PRf 4)AB

__ 1.50 0

-- 1.50

1.00

0.80

0.60

0.40

1.60

1.40__

1.20 ....

0.20

._ I.I0 0.20

I.I0 0

1.50 0

1.00

0.80

0.60

O.

_fI

_Y

I0

0.2 0.4 0.6 0.8


0.I 0.2

_w/sec

1.50 0.20

I.I0 0.20

I.I0 0

1.0

I

0.3

FLIGHT VELOCITY, !000 m/sec

van I, IUSPS. CJ&IPOINIA



Page 98

Page 66IAS[LI_t SPICtHC ImpULSt

F_ OP[IAVtak

ii

_ Iom_' _bw Ivwac. * _

|,.-

i.o_,d

o.z o.+ +_* =.l

' °.,% oi= +.h +._ 01,,+ o._+

,.,lJ'+r".jl

I.i.I l,.iJ

1.60

1.40_-

1.20

1.00

0.80

0.60

0.40

1.60...._

1.40

ew_- 1.20u'l

.- 1.00F--

F- 0.80

" 0.60Zp-

0.400

I0

FAN OPERATION

Eng. No. II

Baseline

PRf = 1.30 l"f!

I"'!

i .L'

.L

Page 67 .....II$ltltl TNIUST

FII OlllAllOk

r,-i..! i i i ! i i _/ ! _ i _ ! + I

I _ ! i'b.l ; I'*._,I

I r'*.i I I i 2>._.- i[ : [ ,7['._11 ! ! }

i I I 11

0.2

FLIGHT

PRf _AB

1.50 0.50

1.50 1.00

_ I.I0 0.50

/ I.i0 1.00

1.50 0.50

1.50 1.00

I.I0 0.50/_'- I.i0 1.00

0.4 0.6 0.8 1.0

VELOCITY, 1000 ft,/secI I I

0.i 0.2 0._

FLIGHT VELOCITY m/sec

- Report Z5 194

Volur_ e 7._/J///

iarqua,,../u._ ....u,,.¢.,,,.,,.,. page99

AFTERBURNER COMBUSTION EFFICIENCY EFFECT

! L=,a_mI +,._m

i +,¢¢G

Page 66|A_tL_"[ _PEClf*¢ tMPULS[

_*N O_I|MTIOm

FAN OPERATION

---_____ ", .-.----T _"

• L. t

, Page 67 .....ii IA$1LIII| t _llu s/

Eng. No. ll ............

Baseline

= 0.95-c

a.es o.to o.xs +.,.i +.+s a.m

I(hen_J.._

uJu_

1.06

1.04

1.02

1.00

0.98

0.96

0.94

1.06

-----.

7_C

1.00

0.85

_AB

0.20

0.20

neI-

(#}

n_

"-rI-

1.04

1.02

1.00

0.98

0.96

0.940

I

0

__..

0.2 0.4 0.6 0.8

FLIGHT VELOCITY, 1000 ft/secI I

0.I 0.2

_---- 1.00

0.85

1.0

0.3

0.20

0.20

FLIGHT m/sec

AFTERBURNER COMBUSTION EFFICIENCY


Page 100

EFFECT '

FANPage 66

fA_ Op[|A11@m

1,1_ Dx = ii

,.h .I° o.h o._,

OPERATION

Eng. No. 11

Baseline

_/ =0.95C

Page 67 .....IA$[LI_I INIUST

_J_ OPlmArt_

"r_,_.,...[[ i !

i,._-',.I J i I'_i

I i°L _, .....,:.,-

._ ,:. ,._, ,!.

I I ! _ i_ i i i ,

I"N. T_.'!! i"h_._.,!

I ,r.._o.,:

i r-,,,,_l ol

,.. ,:. *._

.J_l

IdJUd

1.06

1.04

1.02

1.00

0.98

0.96

0.94

1.06

/I

J><

_'/c _)AB

1.00 0.50

1.00 1.00

0.85 0.50

0.85 1.00

.2

l--on

I--

I-

1.04

1.02

1.00

0.98

0.96

0.940

I

0

JY

0.2 0.4 0.6 0.8


0.I 0.2

1.00 0.SO

1.00 1.00

0.85 1.00

0.85 0.LO

1.0

I

0.3

FLIGHT VELOCl .000 m/sec

EFFICIENCY EFFECT


Page 101

_109_J.J

ff

page ooIA_(_IM[ _plClF+C 'MPuLSl

rA. O_[IAIlO_

_iii'

1.15

1 .I0

1.05

-=_'_Iz.oo

c_ 0.95

_ 0.90,Jl,,,

0.85

1.15

_: Z .I0

F- 1.0509

"_ 1.00-r-

_- 0.9509

=: 0.90

O.

I

0

FAN OPERATION

Eng. No. Ii

Baseline

"qN -_'= 0.98 +

; .

/I

jl

Page 67 .....

la_t_+NI _weus_F_ oPllallO_

= +.o+ +.1+ +++s +.++ +.z+ +.m

// 1.oo

'F/N

-- 1.00

//

0.95

j_

/I

0.95 0

1.00 0

0

0.20

0.2

FLIGHT

0.20

.--- --- 1.00 0.20

_" '--- ---- 0.95 0.20

0.95 0

1.0

I

0.3

0.4 0.6 0.8

VELOCITY, I000 ff/secI I

0.i 0.2

FLIGHT _ )00 m/sec

EX IT NOZZLE EFFIC IENCY EFFECT

Report 25,194Voluro e 7

Page IOZ

FANPage 66

IA_i_lh! SPfCIIl¢ Im_LSl [_ _A LL

fan OPEmM fDi

a_,oao ; i

i ,.i

rLOT _V,, _ON _

OPERATION

Eng. No. 11

Baseline

_N = 0.98

Page 67 _"*"|A$[LII! TN_US_

I.I i'_! ! I I !_.,i

' ._ ! i i ! i i ,_1l _ _ 1,4 4.1 O,I 1.1

1.06

l_ 1"°4_oo 1.02 _'"_J.J

o.o. 1.00

o_ 0.98

uJ_ 0.96--

0.94

1.06

_N _AB

1.00 0.50

1.00 1.00

.... 0.95 1.000.95 0.50

,.,: 1.04¢l,t

_- 1.02 _

" 1.00:%:I-

_.. 0.98U'}

0= 0.96_-rI-

0.940

I0

1.00 0.50__-_'_ _ __----_" --- 1.00 I.O0

_ _._..._..___. =_. 0.95 !.00"----_--'" 0.95 0.50

0.2 0.4 0.6 0.8 1.0

FLIGHTVELOClW, IO00 _secI I I

0.I 0.2 ,0.3


rquaruz ..._,,._.,,,o..,.


Page 103

Enq, No. 11

EFFECT OF SUBSYSTEM WEIGHT VARIATIONON ENGINE THRUST/WEIGHT

18.8

+ 10 PER CENT WEIGHT VARIATION

18.6

-- 18.4

_- _ OES,_NPO,NT

_" 18.0

LU-J-J

<

(11

zi

z

_z

zi11

SUBSYSTEM __-@,'CONTROLS, LINES _'_-

PRIMARY ROCKET SUBSYS EM I_-

_-(_-IEXIT NOZZLE I SUBSYSTEM " _ (_)-

I t J I k_®--,(_- MIXER/DIFFUSER/AFTERBURNER

i \®__-@- FAN SUBSYSTEM

17.8

17.6

17.4

17.20.90 1.00 I.I0

NORMALIZED SUBSYSTEM WEIGHT

_8.dport 25.194

volun_e 7

Page 104

EFFECT

23+m

I

Eng.

OF SUBSYSTEM WEIGHT VAR IATION

ON ENGINE THRU ST/WEIGHT

50 PER CENT WEIGHT VARIATIONI I I I I

No. 11

22

• \o

/o s,o.

1/! \_i._ I ! , 0

SUBSYSTEM •

17 "_)-CONTROLS, LINES _l _"_.'_' '

-- .r_-PRIMARY ROCKET SUBSYS"EM .

:- _1 i i i i i i i,,\-,,,_16 ---_) EXIT NOZZLE SUBSYSTEM _ _ (_).

_.1 I I I I I I I "_._,_- M IXER/DIFFU SER/ AFT ERBURNER --

15---@ FAN SUBSYSTEM-

140.50 1.00 1o50


/V I __ J1/,4Jarquaral ....,,., .,,.o..,.

- iiXiRplkftAtll_


Page 105

SCRAMLACE, NO. 22

The ScramLACE powerplant (Engine No. 22, Class 2

Study Phase) is a 173,000 ibf thrust (sea level

static) engine with Mach 12 flight speed capabil-

ity. The fuel is liquid hydrogen, with an auxil-

iary supply of liquid oxygen required for gas

generator drive purposes. The engine normally

operates in three progressive modes: (i) liquid air cycle ejector mode, (2) sub-

sonic combustion ramjet mode and (3) supersonic combustion ramjet mode. Primarily

because of SCRAMJET considerations, the engine has been packaged in a two dimen-

sional configuration. The uninstalled engine weighs 10,457 lbm, providing a sea

level static thrust/weight ratio of 16.5.

The basic design specifiers are: Design mass flow ratio 1.5 to l, primary chamber

pressure 1000 psia, maximum internal pressure 100 psia, and an air liquefaction

heat exchanger equivalence ratio of 8 to 1. The overall length of the uninstalled

engine is 300 in. (7.65 meters), the width is 142 in. (3.6 meters). The overall

height is 102.5 in. (2.6 meters).

The engine comprises a light-weight air liquefaction heat exchanger assembly con-

sisting of a precooler and condenser unit ducted together in a low pressure shell

constructed of reinforced plastic. All pumps are driven by hi-propellant gas

generators. The heat exchanger assembly is capable of being closed during the

high speed modes.

The primary rocket assembly consists of eleven (ll) regeneratively cooled vertical

two-dimensional linear bell rocket strips. These units act, also_ as mechanical

supports for the supersonic combustion ramjet fuel injectors. Aft of the rectan_a-

lar mixing section and diffuser is another series of cooled vertical struts which

inject the afterburner and subsonic burning ramjet mode fuel.

Exit throat area control is effected by four vertically hinged cooled exit panels

which close from the engine walls and a center structure. This throat variability

is consistent with ramjet (subsonic combustion) and ejector mode performance cited

herein. For supersonic combustion the panels are faired in line providing minimum

drag losses.

All internal engine surfaces are regeneratively cooled during all modes of engine

operation. The basic panel structure of the engine consists of light weight com-

posite structure consisting of a thin-gage, multiple wall internal surface cooled

by double passed hydrogen, supported through a bonded compliant layer of elastro-

meric compound by an externally insulated berryliumhoneycomb structure.

The engine was sized for a 1 million lbm gross weight horizontal takeoff two-stage

launch vehicle. The engine was utilized in a complement of six (6) units mounted

along the bottom side of a high fineness ratio, low drag lifting body design. The

inlet comprised a two-dimensional moving ramp, variable geometry, mixed external

and internal compression unit. Exit gases are considered to be further expanded

during the high speed flight modes against the underside of the vehicle in order

to maximize supersonic combustion ramjet performance.

_a/I

rquanTr ...._.,.:.,,,o,.,.14,YMbM_ATA'3W


Page 106

En_ No, 22

Engine Operating SchematicHEAT MIXER/ COMBUSTOR

..... I EXCHANGER I LDIFFUSERI_ _ .....

= |NL_I 7= --I I-- -_- --IZ _..,A I i ,-.,""4n

F_'_)_.-._""---TU.O,,U.,,ASSEM.LY

Engine Operating Mode Block Diagrams *

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®

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--'1 XLom ,J

*Note: Mode numerical coding is given on Page 105,

"R" denotes optional all-rocket mode.

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:42.0

io_Va_i Q_r_D m


Page 108

Air

WEIGHT STATW_MENT - ENGINE NO. 22

Liquefaction SubsystemPrecooler Core

Condenser Core

Forward Shell

Center Shell

Aft Shell

Sump

Boost Pump, Ducting

Catalyst (para/ortho)

Closure and Transition

526

_65

25O284

130

100

307969

53o

3561 ibm

Eng. No. 22

Primary Rocket Subsystem

Rocket Chamber Assembly

Support Structure

TurbopumpsGas Generator Unit

Ducting and Valves

Starting System

Mixer/Diffuser/Afterburner Subsystem

Mixer

Diffuser

Fuel Injection Unit

Combustion Chamber

Miscellaneous

Exit Nozzle Subsystem

Moving Plate Exit Nozzle

Actuation AssemblyExit Nozzle

Miscellanous (5_)

Controls, Lines

Control Assemblies

Valves and Lines

Total Weight, Dry

(Thrust = 173,000 lbf)

Thrust/Weight, Uninstalled

588I089

284

149

76148

605

585460

_95

107

11852_

5_

8o185

233_

(22.3_)

2252

(21.5_)

2025

(19.6 )

265

m

lO, 457 ibm

16.5

(2743 kg)

rquamr __,. o.,,._.

Report 25 194Volume 7 '

Page 109

SCRAMLACE (ENGINENO. 22)

PROGRESSIVE OPERATING MODES

(PUMPING, COOLING AND CONTROL CIRCUITS NOT SHOWN)

H2 _ -.-L-AIR

ii lllliilll 1,- ,_,____-_w_, _.._]_ . all-l i "l" __,_-"_]

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5

SUBSONIC COMBUSTION RAMJET MODE

R ">-' 1 3 c)


Page Ii0

SCRAMLACE (ENGINE NO. 22)

PROGRESSIVE OPERATING MODES

(PUMPING, COOLING AND CONTROL CIRCUITS NOT SHOWN)

H2

SUPERSONIC COMBUSTION RAMJET MODE

::.- OR L-.,-t,IR

i

J_ ':::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::: iiiiii!i!iiiiiiiii!iiiii!!iiiiiiiiiiiiiiii!iii!!

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ROCKET VACUUM MODE

R-22 1 40!

J2

rcuarar ..,,,=.,.

t _ .

.,_port Z5,194Volu_e 7

Page 111

Eng. No. 22

BASIC PROFEI/ANT CIRCUIT

SCRAMIACE, ENGINE NO. 22

The propellant flow circuit including pumps and primary controls

is reflected in the adjoining figure. The circuits shown are

essentially for hydrogen fuel_ the exception being the llne from

the liquid air condenser through the liquid air pump and to the

primary rocket unit and the liquid oxygen auxiliary supply. The

several valve Junctures shown in the propellant circuitry are to

accomodate bypass of fuel a_ appropriate for the individual

operating modes.

Three turbopump assemblies are shown; one for liquid air, the

remaining two for liquid hydrogen. Each pump is directly con-

nected to its drive turbln_ These are driven by bipropellant gas

generators as shown. The lower set of pumps (figure) operates to

provide propellants to the primary rocket and are driven by a

common gas generator. The exhaust products of the gas generator are

injected into the afterburner for further combustion of the fuel

rich gases. A single gas generator is associated with the ramjet

fuel pump which operates during the high speed modes following

primary rocket shutdown. Fuel for the gas generators is self-

pumped, with an auxiliary supply providing the oxidizer as shown.

All hydrogen, once pumped to pressure, passes through the condenser

and the pre-cooler of the air liquefaction unit (in that order)

prior to any other routing. A high pressure circuit of the order

of 2000 psia conducts fuel through the heat exchanger and to the

primary rocket assembly. Liquid air taken from the condenser is

first pumped by an in-sump boost pump from which it goes to the

suction connection of the high pressure liquid air pu_p. The pumped

out liquid air is routed directly to the primary rocket chamber. A

significantly higher flow of hydrogen is applied to the lower pressure

circuit (1000 psla) which, after passing through the heat exchanger,

is used for regenerative cooling of the entire engine. For the lower

speed modes, the hydrogen coolant is collected and injected via the

vertical fuel injection strips in the afterburner region. For super-

sonic combustion ramjet mode operation, the fuel is routed forward to

the region of the primary rocket assembly where fuel injection takes

place in the supersonic air stream. This shift is effected in order to

accomplish combustion at the maximum contraction point in the engine.

/Jrquara[ ...._.,..,,,o..,.

• II,Y_q_I_ATIIgN


Page 112

Eng. No. 22

i

I--

w-1-

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(

R -22, 141

UN_FIEDi,_-

VMN NUTS, C&tOOOeNIJ


UNCLASSIFIED.

rquardl, lf(.T_9,cf_l_W VAN NUV|, ¢_¢J;OiINIA

Report 25 194

Volume 7

Page 113i i,,

OPERATING MODE CONTROL SYSTEMEng. No. 22

(BLOCK DIAGRAM)

J LIQUID AIRLEVEL

--[ CONTROLSYSTEM

OPERATING_

MODE CONTROL

INPUT LOOP

SIGNAL SELECTOR

EJECTOR MODE (NO. I)

--------- RAMJET MODE (NO. Z)

"---------- SCRAMJET MODE (NO. 3)

r--I

I '

III

II

NORMAL

SHOCK

POSITION

C ON TROL

SYSTEM

ROC KE T

FEED

CONTROL

S YST EM

,,, i

EXIT I

NOZZLE

CONTROL

SYSTEM

FUEL

C ON TROL

SYSTEM

The engine is controlled by manual and/or automatic inputs for mode

selection with several active control loops being used. Primary rocket

control is based on a scheduled or pre-orificed system and has no active

control as such. It is set to operate the system at a chamber pressure

of 1000 psia and an air/fuel ratio of 34 to 1 (stoichiometric). An active

loop controls the liquid air sump level. This is accomplished by propor-

tional control of total hydrogen heat exchanger flow since this controls

the air liquefaction rate. During the ejector mode the afterburner handlesall of the excess hydrogen from the heat exchanger (fuel rich condition).

During subsonic and supersonic combustionj ramjet mode, an active control

establishes a nominal equivalence ratio of stoichiometric burning. This

is accomplished by air and fuel mass flow sensing and consequent fuel

control. The variable exit accomplishes normal shock location in the inlet

throat during supersonic flight speeds. The exit also operates in an over-

ride loop to limit combustor pressure to the maximum design pressure of

100 psia.


Fag e 114

Eng. No. 22

LIQUID AIR LEVEL CONTROLSYSTEM

REF.

_ ,VALVE__ RPMLEVEL

r__)---_ COMB. _-_ TU_mE = PUMPRATE HEAT

EXCH.

SYST.

L. AIR

LEVEL

The liquid air level control system maintains the required liquid air level in the heat

exchanger sump by controlling the flow of total liquid hydrogen through the heat

exchanger, hence to the engine. The error signal which is generated by the difference

between the required and actual level is employed to modulate the output of the gas

generator controlling the speed of the turbine and pump. This results in flow modu-

lation of liquid hydrogen through the heat exchanger to null the error, i.e., to adjust

the air liquefaction rate to that being fed to the primary rocket.

NORMAL SHOCK POSITION CONTROLSYSTEM(EXIT NOZZLE MODULATION)

REF. co poTI oH ACTUATIONINETWORK ASSEMB LY NO Z ZLE

I AEROTHERMODYNAMICS iFEEDBACK

The normal shock position control system modulates the exit nozzle to position the

normal shock in a predetermined optimum location, normally the inlet throat. Pressure

signals indicative of the actual position of the shock are compared against a reference

signal and the error is fed to a computing network which generates the required signal

to the actuation system to null the error. The loop is closed by the aerothermodynamic

feedback on the engine internal air flow.

rquard/ _ ..,,..,,,o,.,.' It,YHIPIIIt,4TIt_ '

ROCKET FEED


Page 115

CONTROL SYSTEM Eng. No. 22

EXIT NOZZLE CONTROL SYSTEM

AREA

_PEAK HHOLDING

CONTROLLER

FLOW RATE

ACTUATIONSYSTEMEXIT INO Z ZLE

I AEROTHERMODYNAMICS

The exit nozzle control system is designed to maximize the combination of exit nozzle

area and air flow rate by modulating the exit nozzle throat area. A signal proportional

to air flow rate ahd exit nozzle area is introduced at the peak holding controller which

signals the actuation system to modulate the exit nozzle actuator to maximizej or peak,

the product of throat area and air flow rate.

FUEL CONTROL SYSTEM

•- 4) INPUT

CONTROLLER I

ACTUAL

FUEL

FLOW

RATE

I FUEL

FLOW

SENSOR

The fuel control system is a closed loop control system which senses the air flow rate

through the engine and modulates the f_,el flow rate to maintain a required fuel air

ratio (_). A signal proportional to air flow is applied at the equivalence computer

which generates a command for the required fuel flow rate. This signal is compared

against a fuel flow rate feedback signal generating the error signal which is applied

at the fuel controller. The fuel controller will then modulate the fuel flow until the

error is nulled.

RePOrt Z5,194Volume 7

Page 116

Eng. No..22

VEHICLE DESIGN - SCRAMIACE, ENGINE NO. 22

This figure shows the final Class 2 vehicle design utilizing ScramLACE

engines. This lifting body vehicle was determined to be substantially

superior in performance to the other vehicle types considered.

The lifting body shown features high slenderness ratio, elimination

of the second stage base drag through submergence, and attainment of

stabilizing surface at low unit weight. The vehicle incorporates an

aft hydrogen tank and a propulsion package consisting of six enginemodules of 173,000-1b thrust each (1.038 T/W), with a 408-ft 2 total

capture area. A vehicle affixed nozzle contour is effected to

accommodate the supersonic combustion mode. The system second-stage

gross weight is 397,573 pounds for Mach l0 cut-off conditions (payload maximum).

The lifting-body configuration employs a modified conical fuselage

where the forebody is a blunted cone with a depth-to-width ratio of

0.h at any station. Maximum cross section of the fuselage is at 73

percent of the body length, as measured from the virtual nose (apex).

The fuselage nose radius is one foot, and the body planform area is

13,612 ft2.

The horizontal stabilizer has a leading edge sweep of 65 degrees, and

an area of 2612 ft 2. The airfoil section is double wedge, with a

two-inch leadingedge radius. The movable horizontal control surfaces

comprise 2000 ft2. The horizontal control surface rotates against the

vertical stabilizer with forward extending dorsal fins, to alleviate

the thermal problem associated with the sharp edges of the control

surface under high-speed deflection conditions.

The twin vertical stabilize_shave a total exposed area of 1200 ft 2,

with a leading edge radius of two inches. No toe-in is provided for

the verticals, rather, a concept of utilizing small outward rudder

deflections to load the surfaces during hypersonic operation where the

control surface lift curve slope is zero at zero deflection is pro-

posed, in order to maintain minimum vehicle drag. All panel surfaces

have a thickness ratio of 5 percent.

i

SECTZO_,A-- A

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V,.] "J_-r_c T

BCDY PLANFORM AREA 136t2 FT _

MORIZ STAB. AREA 20CC FT '_

STU§ 'A/tNG AREA 612 FT _

TOTAL FLANFORM AREA 16;'24 FT _

VERTTCAL FIN AREA 1200 FT _

A¢ 4C}8._ FT ¢

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Page 118

No. 22

YAW NI/YS, ¢ALIFOINI&

Re_rt 25_194Volu_ne 7

Page 119

Eng. No. 22

INLET PRESS URE RECOVERY

i-e,l

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;:SU ERSONIC i!:: 2_i:'COMBUSTION _i!i'"'

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VAN NUV| CALIFORNIA

Report 2-5,194Volume 7

Page I;)- 0

Eng. No. 2Z

INLET CAPTURE AREA

6

5

E

"¢ 4I,,0

3t_

I-"W

z 2i

1

0

l

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W

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, _,.,. f: - .-i, --- I-.SUBSONIC .........................d]; _: ..... ::::': : '!:: ...... _,,,r..... COMBUSTION ::: i:i : i':i i:'i i i;

i;:iii':::i:'i-:.ii"u::::,: e._:::.:t :: i:i il; 'i_i__ii iii !i_i i_i__i_i_:;i _t:i_ _'_ :

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..... :::::::::::::::::::::::::: illi: ::.:::o 2 4 6 8 I0 12


vJN NuVlI, CA&WOII_IA

Report 25,194Volume ?

p_g?Izl

Eng. No. 22

EJECTOR MODE SPECIFIC IMPULSE

4400

4000

o

3600

m 3200

J

-- 2800

W

m 2400

_ 2000Z

1600

12000.5 I.i

FLIGHT

!_:=::: 140!• ;;6

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3.0

FLIGHT VELOCITYs 1000 m/_c

I I I I I0 0.2 0.4 0.6 0.8

YAW _UV$, CJLJPOIIf4I "L


Page 122

Eng. No. 22

EJECTOR MODE THRU ST

3

(g

2

iim

E

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P#-t-I-

l--U.l

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700

600

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l-l,iJ

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100

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VAN NI/'rL C&Lfl_NINIA


Page. 123

Eng. No. 22

EJECTOR MODE A IRFLOW

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EJECTOR MODE CAPTURE AREA

Report 25,194Volm_ne 7

Page 124

Eng. No. 22

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Page 131m

ENGINE 22

_ARY DATA

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0.75 lllS. 520. 1635. 17.41 8.12 25-53

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ALTITUDE - i0000. FEET

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1.4o 1468. 520. 1988. 17.14 6.07 23.21

1.60 1816. 520. 2336. 18.52 5.30 23.82

ALTITUDE - 20000. FEET

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1.80 16&6. 520. 2166. 21.55 6.81 28.36


1.17 553. 327. '880. 16.65 9.85 26.50

1.30 6&0. 392. 1032. 17.27 10.58 27.85

1.40 716. _d_0. 1156. 17.93 11.02 28.95

1.50 802. 520. 1322. 18.75 12.16 30.911.90 1297. 520. 1817. 23.88 9.57 33.452.40 2345. 520. 2865. 34.04 7.55 41.59


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1.70 789. 478. 1267. 20.9& 12.69 33.63

2.O0 i138. 520. 1658. 25.64 11.72 37.36

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YAN NUY|, f_AL|fOflN|A


Page 132

ENGINE 22 CONT'D.

SUP_ARY DATA

MO WS WHX WT AO AHX AOT


1.52 525. 317. 842. 18.92 11.42 30.34

1.65 614. 373. 987. 20.32 12.34 32.66

1.85 782. 478. 1260. 23.08 14.11 37.19

2.00 940. 520. 1460. 25.64 14o18 39.82

2.50 1738. 520. 2258. 37.76 11.30 49.06

3.00 3126. 520. 3646. 56.34 9.37 65.71

ALTITUDE - 50000, FEET

1.86 490. 292. 782. 23.22 13.84 37.06

2.00 583. 347. 930. 25.64 15.26 &0.90

2.10 659. 385. 1044. 27.61 16.13 43.74

2.20 746. 436. 1182. 29.80 17.42 47.22

2.50 1077. 520. 1597. 37.76 18.23 55.99

3.00 1937. 520. 2457. 56.34 15.12 71.46


2.19 457. 255. 712. 29.57 16.50 46.07

2.25 492. 281. 773. 30.98 17.69 48.67

2.35 556. 304. 860. 33.51 18.32 51.83

2.40 591. 329. 920. 34.87 19.41 54.28

2.70 848. 429. 1277. 44.33 22.43 66.76

3.00 1201. 520. 1721. 56.34 24.39 80.73


2.53 430. 203. 633. 38.79 18.31 57.10

2.6O 466. 225. 691. 40.91 19.75 60.66

2.70 525. 256. 781. 44.33 21.62 65.95

2.80 590. 294. 884. 48.04 23.94 71.98

2.90 663. 326. 989. 52.03 28.58 77.61

3.00 743. 347. 1090. 56.34 26.31 82.65



Page 133

Eng. No. 22

SUBSONIC COMBUSTION

NO PRESSURE FIELD

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VAN NUV|. CALIFOINtA


Page 134

RAMJET THRUSTEng. No. 22

SUBSONIC COMBUSTION

NO PRESSURE FIELD

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Report Z5_ 194Volu_ne 7

Page 135

Eng. No. 22

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Report 25. 194V olubne 7

Page 136

Eng. No. 22

RAMJET THRUST

SUBSONIC COMBUSTION


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VAM MUY|, CAL#@O|MIA

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Report Z5_ !94Volu_e 7

Page 137

Eng. No. 22


SUPERSONIC COMBUSTION

NO PRESSURE FIELD

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iT..=_f_i_i!.......- .................................... ..........!_i..._.L_i`ii_-_._T_._:._._."_-:!_..i_!_im.._-_.i_i_i.!.i_i_-_!_ii_-_f_7_+_.........,.............. _...... ,...._.................. ,...=::,:--.i:LS=-:5:'-i=_H=_ ::::::::::::::::::::::: ::::: ....... :.,-:.............................................. : .... _... ::

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:. :: =:=-'--,,=_,,-:...............:_-!Ti!ifi77ili_i_t!_::!]_:.=:l::-_:_.i_:.=.-l::_!_l_!::liT_il!ii::::?_i17:___i!_-.-::=#->f_-!:.i;.'__!_L:==gli!i_'_.-7i_'.:iii!_:i::i::i!ii!l::_!lli!_!;-i::_-!'::!::_':::iiii!ii'i!!i'i-_i'_::_ii::-':..ii:'i!i'iii:.li_ _::i!i_ii_=::,=:::=::=: _._"t!!_! "!:.=..u=._ i_iliiif:!i_;z_i-'!i!i .....::iii ii!ili::il!i::_l!iiii_.:-!'i!i::!i:.-!!-ii!-

_=_.:._._'_:_'2__=-"k_..l,.-:=::==:=;.==._::=.:..:.;.=='"__=::_==:'ii::i1::?:_li::-ii_::_:i_!_i:li::::ill!::il_i!_ii[::_=-'.'_-:._i:_li::::-i!iiii:.i:::-i_iii

6 7 8 9 10 11 12

FLIGHT VELOCITY_ 1000 ft/secI I I I I

1.5 2.0 2.5 3.0 3.5


13

I

4.0

VAN NU¥|. CALIPO|NIA

RAMJET

SUPERSONIC

THRUST

COMBUSTION

Report 25 194)

Volume 7

Page 138

Eng. No. 22

P

NO PRESSURE FIELD

1.2 -

1.0

- 0.8

_=o.6p.

l--hl--j

0.4,,iz

0.2

0

i,i-..)

-- l.IJZ

280

|

240 * :_i % ; : :

' *" ' l::: ! , :liiiii:i:it::ii!ii::ili:::I:iiil!::i_i::iiii:;:i:::i_li!ii!:.::iii

: i I " I J t:; :;;t. I :!_::::t:i ]: li_f::_;iti:::.;Y:,iii;;::it:!ii_l!_i::ii::i;::_;]:::i!fili;i:_;ii! / I ]: "_:t i:i:i:i t l t i ;1_;1:i::;::t?::it!;!t::i!ti:_!l:iil;i:t::!:i;:!':::1:L_

; _6° _ l _ __ :ii:iii::iit{::_:]\t:_:_t; l::_1::;_t_:::! i:l:i:::t:li!:t::!::::tt::i{liiitfiii::t::::::::t:iti!{::l_:iil:_i.... I::!:_i::!i:_

II _ II: t:i: i! i[:i]tti:il:]::I:i..,_ i:i::!: f [::i: :'i::Jiii:.L:iii!::ii:.F:i!l:::i I::[i:L:I/i: : l _ : _ : ' i : { _ : _ : :

•liiiil :::ti:::1_:::I_,f:_:1::t:::::f_::]_ c. _1:, :ii!ii:;t;_l::i_i_!::iP_#liiiiI:ifit{i::f;!::_{i]iliii!: :tiill'I i _1 i?::*_:'_1_::':_:_:_:t: i i i:_:{1; I:ii::i:ti{lii :t:i::P:i::tP:ii::!t::ti::::t::i{il::{::;l:i#!i!ii::::: ;iii

8O

40

[iii! :::+,'i_'::_:_:::_::_::::::::::::::::::::::::

5 6 7 8 9FLIGHT VELOCITY,

!ii::F:lili_li:;it:,:::,::iii_::tiiiit:::::::f!iiii:i]::t!:::!:_i_;;ili:_:::: ;::: =:t::::|:::: :::: :::::::::::::::::::::::::::::::::::::::::::::

0

i::!i;:i]i!ii_-_!ili.i:!!L::i_it::!i!I:::i!!:ii!F:::iil;::-t::!iii!iliti!!ii

_iiiii!!i:

i0 IIi000 ft/sec

::: ":::r::::t:::!i: :: ::::=;:: ::::::F::::::::F::!I::[F:;: ::::i:i::

_iii :iiiili:!!:i,iiii{::i_l::ii:i

12 13

I , I I I I I I

1.5 2.0 2.5 3.0 3.5 4.C


VAN NUVS. CAI.tflOIINJ&


Page 139

Eng. No. 22




TRAJECTORY b

3000 L-L:I;I[.I .....

'" , ::':1 .... i!_._ii::z) _: . :; :: !i: il_i} tl;il :I ; ! i: : ! !i:::;!i ii1 : il!l::ii:_t!i!!ii!!!tii::::i::ilhiiiliii_tli::ilii?igi!@-ti!i_

i = :i:i :_i:,::N_:i i:l: i !: : _ _li:ii:, l!:_?i_lili!tiiiilii::i;iiil,i::i::!;iiliiii_iiil::ii!lliliti:.-.:iti_.:.!!!ii2500 I:_.,::::1 I :. ,: ._:t ::I:! ,.::!::,:1::::i=================================================================================

I . 1 ..... ' .......... ; .. .; , ............. '. ..- ;.

,-_ !::i ::! I I:i:t:_ :i:_l::l _t_:_l:ii:it;:l_;it_;si!ii::l!il;l;iiii:;itli:-it_iiit::;i_tii:J;i;ii::ii_livl

-" 2000 =: -_ :: _::t_!!,;_.,_::,_:::J:I. :li_.,_ ............................... ' .............! . !_ !,:: -..,. , ,-i_: :l_:_il_;i!t_:_:li!:.ili:iii!_iiT:i:i!I_iiifi!i! "::.-:.il::::_::::_s_i:_-'ii!,,," _ i: :7i: ::i_i;t!::_il{:._ti!!::l:iiJ: _i::iii::_iF!:t : _ _:'if:"t::!::::t::i:._t::!#!'::i::i::i::il!?iitiiiit::::ifi::iiti!f!i{iliii!

' I !: :_:!i*,iil:i2:l!:_i! :::: : :: Ii:£"_'.._::i !!iii!!:.i! il;ij_ = 0 o AND 4°"

=_. t; ' I:::: "ii_!it_:::::lii?!l?I]I_::1!; :i.i::ili!il :!i: i:.i:,li!!i,..,y ::li:ii._.:_:i:ii _ i-t_i_i:..!TF:git!:.!]:.iii":ii:. • -,,,a.. -: ......... ;;I ...... ......

- 15oo i: _' I :I:::::t!:!l!:!l;i!:l_;!l:__:1::f:::_;t::_:;::;i!ii;i!i;::_:.ii:ii;!i_i;:l!i!:f;__ _:_::!::;;:4!:!' : I :;_I;;:t:::::;I;i_t i/t: _I::_t :l::i!lli_t .ij:::_!ii::i;i::;::i!i:_[ii-ii::li:::!ii: [!?::;-_:-__i_!ii;i :_;

"' _::_:,:i:_,_:.,i!:,il:iiili::!lii::ilii::il:_;!i!_ii:.i_iii_ii::ili::is!::_l!:_i_::::i'4s:.!iliiiilLi._ili::ii.i!:i:iii"i::itii::it::iiil_si::t;::2il_:_iii_?1_i::t;i::::t:,:/f!ilil::.::I_ii;.:i:.i;li_!_l_:_:t::::::i;i.::lli;I,;i;lii:/:t!iiit_::iiliiiifi_iiti!:ii::ilii;;l::!!itii!!ti::!il;i::::::iii!.--7..-'::;i

_. . ...............,'..........................................................., , _ _ , , , _ _ ilil, _,,=_l!i:;I:i:.t::::iSlS!iit::!i::l::i::q::i::t::_I i_::li::iit!i!ili!!:41::_:.l_i;iir:t::!_.:_ii_::!i!!i?i_!i!iiii::!_:i!i;!i!!iigii_::i_ i_M;_!

ad i ! i ! !-T_I!!iF_E[_!H:_'_IE÷__iiii!i i_i!-ii!':u:;ii;ii .....z ...."-" .... " ..... '""" ....... '"' ....... !ii!i!!-! !i!I I _ I | | I II: _:1_]i_i::t! ifi::::::liiiitii::iti_i::t!_if! :_t:i:il_i!:.t::i::.:i::i?it:::_i'::_'iii_,:::!_lSiitii.:-.T-.-::ii.::g_]ii::i 2:;:_i!!iiii:/i!i i:_-;!lig_g_==i

500 ::. iiiiiiiii.liiiii!ililiiiii_i:i';i!i:::.ii_!iilF!ili!iiii!!!ii':_:i!iiii!! t!i]! ::::_::;;::!lik_li_{_i_i_;.i]!iii[iii!!!i'iiii'ii]!!':i]l _IF.I:_÷i:ii;

0

_::I_1 :_t_i_!t:i_f_;_-L:;::_:::: ;.::':ili::ti:::il:;;:i':ii_::!:::'ii!::_;ii ::_i!l;:::':::;';-:::=_,_ii!il ,;.:::::::::::::::::::::::::;I;;::;.;_: .:: .,_;;;:;::;...:::::, , , . ._;'; ':::';::;_ .... _...............

....._....:...:_I:_.................. ;:_il L_._ .................!....!...! ....t....,..._..... _,.....t,.-.':'::::: :.-:_ :_:..-.,_.=--._ !iiil ::i=,i!!_!i'_:_._i__ :--_"_;5 6 7 8 9 I0 11 12 13

FLIGHT VELOCITY, I000 ft,/secI I I I I I

1.5 2.0 2.5 3.0 3.5 4.0FLIGHT VELOCITY, I000 m/sec

VAN NUS'$. ¢AiWPOIINM

Report Z5,194Volu.m e 7

Page 140

Eng. No. 22

RAMJET THRUST



TRAJECTORY b

1.0

O'l

_0.8

{

_.0°4

I,,-w

I--wzO.2

0

24o,......_ ,, 11 iI_I_t _!J '......_ ¸¸' __:: ...............:........ _!J _i_I:i I_II _I_].. ,.,.._..,_,. ' ....... ::;: :::'"::: ;::: I:: i

_i_i-:_..!L'.'._;_:l_l_:!_:_l_i_i_'_-.i_!_!_ ._:I _ l_:J::_ _ I_:: :, ........:,::-_................. i................... _.................. I _ _ , I _ ii_illl200 ::::!:::;....................... k'_'-:J'"'!' "'l ........... J I. , {ii:" ; , :: ;:;

...................... ;'rq'-1--'" ; .................. , ' : j: :: ..:-: ',!_,_:__'_:_'__:_'_iI:___i.........! _ Ii"ii! .] iiilii] - :; .... ' ...... i...... _'h.... [[i ] ...... ' : .....

':if_i_.-:.i:!_L::iLij::ii_;qiT:_=i#_:!i_#!iii!ii#i_!i::ii:: i::ii:.iit!:.!_ti!i ; :_ ._:::= -

x6o _i_.:i.!_i:iI_ii_!i;:iii_T_:!iiF:ii3!i!_iiiI!:i_i._ij!_!:.i_:i`!_:i:_i:_:_ii :, IL:ii!:.l:_ :i!ii_:lii:: ilii:i_ i_j_i..--_iI.i_j:_!_!_i_i_::_i!_:_i_ii_i!iji!_.:!_:!ii;_!i:::.!_i!;_iiii_i_j__ J !_:ililJi! _ _!!_i LL:!II:.IJiI:L_ii

120 ..:.....'-.':.-- ................... =,.=_,:_, .... :.: ._:,::::_::,:::::::..:_ ._:i:::.::_! . _:t ::_.:,::::-:: :_.

IT::: _ ..... i "i I' '_[..', ... ._...... L ... ............ -.. ....... " ._ .... : ..... _'_ ............... , ......."' _i:i:._:L_-_!_!_,_:_::!:_,_._,_i- z_o::_!_:._I_: _I_ I:_!._4, 'I _! , _::_;_ i !

: :i .....: !ii!:i;

5 6 7 8 9 10 1Z 12 15


I l I I I ',

1.5 2.0 2.5 3.0 3.5 4.6


_a

rquardlVAN NUf|, ¢AIIPOItNIA

i

Report ZS, IVolume

Page I,

Eng. No. 2;

5O

4O

Ul

20m

I0

0

REFERENCE TRAJECTOR IES

SUPERSONIC COMBUSTION RAMJET

160•:.!...._-_;-_....-,1.... :!:.!-:-:!--..-F-;F...,!:,:,_.-..!:_:!.,.!.;:_:4DYNAMICPRESSURE, psf_ 250_

i................................. 5-............................................................ I. ,,L::............I...m.._.L-i ....,.........I..]-..] ........,,,:-..:,...f....I...I...t....L_ ......t I _.!....I..._....I ..;....I ...... ___ ....

;_ :i:1:I=I; .%==:..'" :; ;::: " :':.'I,:. _..:; : :: , . :: ::: : ', "" _' " _,i_: . ,, :::' : :: ::: :::: "" .;..,.,.: .,:",.._.,,,,F=-_..,L.t_,L__.,,_,_..,,: ,.,, ,:,,, ._......._ ...._,,_ .....!_ ....t....._ _i ........

120 '::'::i.=:=.-::- -:::_.:-:::..:r:: .................... _ " _-_

i__,_-'-_:._ii._._..:_ili_ __-_..::,_T:_ ..1............. _,_._, !_::'_;_)::T::.::i):/i 1500..........'......._'" _" _li!_::!!,i!!!,_;ii,ili:t_;_ _ _ _!l!i::_.........;;_;;;_ ::__ 2ooo.:4.

:_;..........a"_,'...... _ : " _' _;i: _ii;::.::ii ....... i "_ ,_ 100 iiii(!i_',_ :.'_itii_Yi)_il:;ii:_:i'_:'r':_ 'ml':: _!:_':".i _.'i!))! _ _ !) iii::-"i_tii:_:::_ '::..... ....... ;_,000_

............ I..... ' ........ _ ........ , ......... I"" _ ..... ::":'_: .... "_ :i: ::: ::::::i:::t::):::::: - ,:::_ .- I_:- :_T : "" '.:" ";::..:_'::.t_.'_::', " .... ':"1 :: : ::/'=1 T::: T._.' _L_ i::i L:_(;::. :: ".':':: i_, _ : . :T::_t:T:I::::L::..::::h:'_::

............................... ......................................i....I.................................................. I ....... _ ......... =-...... I.............................. ] ..................... l ......._...... _ ......L.. _,,,... _ ...... _ 4000 ............ L-.t........t....t....................!......._...........:::; :::- :.._.= =.:=::=:- :--r:::." . :"=: L_.. *:=: _:;';:-L:' ;:.I._i.":;::$ ' ='= .'.;; :::: :::: ;:: '::' :::; :=: ;::: :':, :.::':L;:_':-: :':" :?:" ::;:........ _.......... _ _,_.--: .... _................. _.= ............... I _.................... _...... J ............

_:_-_,_= :::_=::iL_:_:::;,;;::_.:_,=.-_::ztq:u::i,::::.:::='-|==::J-:_::-_':;:;:.:::!:":h::::-:--==.-_::-:-:-='--:'I::"

40 !::""::::':::":=:=_'=1 :'.;:::-=_-"-_'=_t 'T_''_]il]::iii'i:._T_::'_=_:'-'''_iii i_ii.... : _=_::=1i:.::!t_:;iitii_iK'.-!t:]_::!/_,:.i!:i!iliI:!::f._iit ii i]::ili::_K4._l::iiiI_::it_ill_iii

::":*:-::-_==:='-.iL_.ii_i=)_i=_:!i-t;i/-:=:,_;:,.:_m_=-_:-:r_;F:i)itihli-i)"_:::';i]!_i7!I!:.:iiii-)_,_:_t_.:;f::t_J_(ifil........................................... _,.,_._......... : :-ii:i;]!::!_::::::_!i::!iii:!_)::i!i!!]]i--:_i]i_i_::)::;i!

20 ':i::i:: :::::::::::::::: ::: ::_::=_:.-. T:":_:_::: h: _!l!iii;!_!i!i li"_iii_i_l!::!!l_!'i1::_i_!iii!i_ i!iii !!i!l_!i l!iiil:!i]: :iili_i_I!::i_!::_!i,iiili-i _i!I:::ili_! _.:_ !._i _!::.._,i_!i ;.:; _iiii,;!_._.._: -:.,; ..... :::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::-F.ii.:_;. :ii!!i-_ii

_.._ r+-_- r_'t'i * "T, -

- 0 _=:=:'_';';";"_:'--'l_="_'_i':'"_':l""=:';:_":':':_; ':.:;'::::;:- =:::_'_"_'_z-='_::=:"._ii_]t]i:.l ::i!il)_':]:#I;:::_:::_,.-,-.::,=:_;;:::::'=:'_:::.::::_,:::..:::,::::="=:':.:":_(:::_ii_}T_f_lii_i. ._.,..,._iii5 6 7 8 9 10 11 12 15


I I I I I a J

1.5 2.0 2.5 3.0 5.5 4.C


rquarar .....,, ..,,,o..,.


Page 142

Eng. No. 22

SUMMARY : SENSITIVITY ANALYSIS - BASES

The performance data shown in the present report, as well as in the

Class 0 and 1 engine documentation (References 1 and 2), was computed

on the basis of a singular set of internal component efficiencies, as

well as stated operating points (e.g., design mass flow ratio, Ws/W_).

Component sensitivity studies were conducted as a major effort withln

the Class 2 study phase. The bases for the analysis are given here,

followed immediately by the results.

The approach used was to define baseline performance_ specific impulse

and thrust (both net Jet), for a reference tra_ector 2. This was

accomplished for each of the engine's operating modes over the normal

range of flight velocities for that mode. It is appropriate here, to

comment to the point that sensitivity studies of trajectory effects,

per se, are already intrinsic in the previously displayed performance

maps.

Proceeding from this basis of specific impulse and thrust discrete

trends, each of the important engine variables was perturbed from the

baseline value, e.g., afterburner combustion efficiency: Baseline

value - 0.95, sensitivity excursions - 0.90 and 1.O0. All of the remain-

ing variables were essentially held at the baseline, or nominal value.

Any exception to this resulted from the engine performance computer

program's automatic compensation characteristics which, in some instances

"retunes" some of the engine internal variables. The extent and impli-

cations of this situation are covered in the main technical report

(Reference 3)-

This section presents the following bases for the sensitivity analysis

results to be given subsequently:

i. Reference trajectories

2. Baseline specific impulse (on reference trajectory)

3. Baseline thrust (on reference trajectory)

4. Ranges of sensitivity variables, with reference to the baseline

values (both curve and tabular presentation)

VAN NUY$. CA|IPOIINIA


Page 143

Eng. No. 22

REFERENCE TRAJECTORYSENS IT IV ITY ANALY S I S

SUBSONIC COMBUSTION MODES

E

oo

,,-I

::)I-,!

I.-.J

30- I00

8O

00,._60

t,,

e_

40

2O

0o

I0

//

//

/

/

2 3 4 5

FLIGHT VELOCITY, 1000 _sec

L I !0.5

FLIGHT

6

1.0 1.5


I2.0

VAN NUY|. CAAtPQIIN#_


Page 144

Eng. No. 22

REFERENCE TRAJECTOR YSENSITIVITY ANALYSIS


140

4O

t_

_J

s 30

20

..I

10

B

120

80 "/

-_ 60

4O

20

B 05 6

I

1.5

TMC A

/

7

I2.0

/

/

8 9


I

2.5 3.0


//

/

I0 II 12 13

ff./sec

m/sec

I

3.5 4.0

rquamz....,, .,,,o,.,.I III, IRPI _ATll'_

Report 25,194Volun_e 7

Page 145

Eng. No. 22

BASELINE SPECIFIC IMPULSEEJECTOR MO DE

3600

Em

QI,

It).,.I:Dr_

m

I.i.t_l.l.I

U3

I--

--}

I--I.l.Iz

3200

2800

2400

2000

1600

12000

_i

0.4 0.8 1.2 1.6

FLIGHT VELOCITY,

/

2.0 2.4

I000 ft/sec

//

2.8 3.2

i0

I0.2

I I0.4 0.6


10.8

I1.0

VAN NUYS. CALtPOJNIA


Page 146

Eng. No 22

BASELINE THRUSTEJECTOR MODE

1.6 - 360

,,,1.4c0

e-

e-

._o.1.211 .,1

E

_" 1,0

I-,

z

0.6iIiI

0.4 j

i320

o° 280of-I

1-

240

I--

h-

I..-IJJ

--Z

4k, '

200

160

120

8O0 0.4

I

Jf

J

/

//

0.8 1.2 1.6 2.0


/

2.4

//

2.8 3.2

I0

I I I0.2 0.4 0.6


I0.8

VAN NUY$. ¢ALIIIOINIA


Page 147

Eng. No. 22

BASELINE SPECIFIC IMPULSESUBSONIC COMBUSTION RAMJET

4400

B

m

=h

LUu3_1

n

u.¢Di.g¢3.¢.D

I-

F-LUZ

4000

3600

3200

2800

2400

2000

1600

//

/ \\\

\\

12000

I0

1 2FLIGHT

3 4 5VELOCITY, 1000 ft/sec

6

I I I I0.5 1.0 1.5 2.0


VAN NUYS. CALIFO|NIA

Report Z5 1942

Volurn e 7

Page 148

Eng. No. 22

BAS ELINE THRUSTSUBSONIC COMBUSTION RAMJET

240

i000 -

8OO

o4.J

Oo 600O,-4

u3

,,- 400

I-

I-,

F- 200Z

0

.oo /\16o J

F-

_-:=_,,_120 /-_ 80

/ul

4O

00 1 2 3 4


I I I0 0.5 1.0


\\

\\

5 6P,.,/sec

L1.5

m/seo

2.0

H

rquarar ... _. ,.,,,o..,.


Page 149

Eng. No. 22

BASELINESUPERSONIC

SPECIFIC IMPULSECOMBUSTION RAMJET

2800,

O

E2400

w

2000

1600

w

w 1200

Wz

8OO5

\

6

\

7 8 9 10 11

FLIGHT VELOCITY, 1000 ft/_c

12 13

I1.5

!

2.0I I

2.5 3.0


|_.5

!4.0

VAN NUY|. CAAIPOIINIA

Report 25.194Volu/ne 7

Page 150

Eng. No. 22

BASELINE THRUSTSUPERSONIC COMBUSTION RAMJET

400 -U_

e"0

¢,-

o 300 --00,-I

"_ 200-"-I-l--

1--

100 -l'-LgZ

O-

I00

80

=--I

.60I,,-

Z1-

40k-

l--LIJ

z 20

\

\

0

t1.5

6

I2.0

FLIGHT VELOCITYe

I

2.5

FLIGHT VELOCITY,

#

IIl_=llllllll ila=,... -

1

I000 ft/se¢

i3.0

I000 m/_c

I

3.5

12 13

4°0

n

rCuamr ...,,..,,,o,.,,I r:tk_Pf_ArlfW

SENSITIVITY ANALYSIS RANGES

Ejector Mode:


Heat Exchanger

Equivalence ratio, _HX

Equivalence ratio, _HX

Primary

Equivalence ratio,

Combustion efficiency, _ c*

Nozzle efficiency, _ N

Mixer - Mixing efficiency, _ M

Afterburner


Equivalence ratio, _ AB

Combustion efficiency, W c

Exit



Subsonic Combustion BamJet:


Combustor

Equivalence ratio,

Combustion efficiency, W c

Exit



SupersonicCombustion Ramjet:


Combustor

Equivalence ratio,


Exit

Nozzle efficiency, 9 N

Exit/Capture area ratio, A6/A c

Base- Iline Range

Figure E

Report 25.194Volu_ne 7

Page 151

Eng. No. 2z

Figure F

Figure G

1.00 i.i0 0.90

0.975 1.00 0.92

0.98 l.OO 0.95

0.8o l.OO 0.50

Figure H

Figure I

0.95 1.oo 0.85

0.98 1.oo 0.95

Figure J

Figure E

1.00 1.50 0.50

0.95 1.O0 0.85

0.98 1.OO o.95

Figure K

Figure L

l.OO o.75 0.5o

o.95 0.9o 0.85

0.98 1.00 0.96

1.50 4.0o 1.oo


Page 15Z

Eng. No. 22

Figure E INLET PRESSURE RECOVERYSENSITIVITY ANALYSIS

0

Q.

M-f:.

>.m_u.i>0(.1i,i

him_

U3U3l,Im,,rt

1.0

0.8

0.6

0.4

0.2

0

,NEI_" SE

I--REDUCED-

SUBSONIC--REGIME

RECOVERY

¢! _.

\\\

' \ _,_, MIL-E-5OO8B

NORMAL\ SHOCK _

\\ %.. "_k_._,: o._

\

0 1

I0

2 3 4 5


t I ]0.5 1.0 1.5


6

12.0

'VAN NUY|, CALIPOONIA

_port 25,194Volume 7

Page 1 53

Eng. No. 22

Figure F HEAT EXCHANGER EQUIVALENCE RATIO

SENSITIVITY ANALYSIS RANGE

EJECTOR MODE

X-.i-

m

I--,,¢n,,

LIJ(.,)=,,1.1.1,.,I,,¢_>

O'1.1.1

,vu4

z

0Xu4

U4

10

6

5 0

I0

_.. EL + "_'I.,)CHX

.4

___ BASELINE (BIJ

BL- .vZ.O CHX

1.2 1.6 2.0FLIGHT VELOCITY, lO00 ft,/sec

! I I I0.2 0.4 0.6 0.8

FLIGHT VELOClTYt 1000 m/sec

2.4 2.8 3.2

I1.0


Page 154

Eng. No 22

Figure G HEAT EXCHANGER EQUIVALENCE RATIOSENSITIVITY ANALYSIS RANGE

EJECTOR MODE

EFFECT OF INLET PRESSURE RECOVERY ON _(;OHXFOR CONSTANT LIQUEFIED AIRFLOW

X..p

"e-

c_m

I--<I:n,,.

IJJ

zI.iJ.._1<:_>g

O'ILl

e,,,LI.If,DZ<:.-r

XI.IJ

I.-.-<:I.U"-r"

10

9

7nmm_

//

/

_wmm _m_

//

0.92

•'_'=" F

BASELINE

6

50

I0

0.4 0.8 1.2 1.6 2.0

FLIGHT VELOCITY, 1000 ft/secl l l

0.2 04 0 6


2.4 2.8

I0.8

3.2

Ii0

VAN NUt'S. CALPFO|NIA

Report 25 lg4Volume 7

Page 155

Eng. No. 22

Figure H AFTERBURNER EQUIVALENCE RATIOSENSITIVITY ANALYSIS RANGE

EJECTOR MODE

6

BASELINE INLET PRESSURE RECOVERY

_5"-

m

Or_J

uJz

_n

Ld

L_.<C

_ _._ BL + " 1.0 CHX

3

.RASELINE @HX

2

1

. . _

00

I0

0.4 0.8 1.2 1.6 2.0 2.4 2.8FLIGHT VELOCIW, ZOO0 _sec

I I ! I0.2 0.4 0.6 0.8

FLIGHT VELOCI_, 1000 m/sec

'¢AN NUI_|o CALIPO|NIA

Report 25,194Voluxne 7

Page 156

Eng. No. 22

Figure I AFTERBURNER EQUIVALENCE RATIOSENSITIVITY ANALYSIS RANGE

EJECTOR MODE

rm<:6

0

m,,

_Q,,=

._ 4 \\

LLI

ILlZr_

rmm-LLII-Ll.<:

3

//

/

%>,, +.. I•_ 4 . _ " HX

' __ _

BASELINE _HX-_

1

00 0.4

I

0

0.8 1.2 1.6 2.0FLIGHT VELOCITY, 1000 ft/sec

i I I0.2 0.4 0.6

FLIGHT VELOCITYo 1000 m/sec

2.4

I

0.8

2.8 3.2

11.0

VAN NUY|° CALIFOa*NIA

_leport 25,194Volume 7

Page 157

Eng. No. 22

Figure J EX IT NOZZLE AREA RAT I0

SENSITIVITY ANALYSIS RANGE

EJECTOR MODE

6

,< 5,,D

'_ Ir

Om

F- 4e,,,,

I,,,I.i

"_ .3I--g

X

IJJ 2NNOz

i Ii il i i

1

/

/

/

mnmnW _ _ _ -- /

_I

O0 0°4 0,8

I I0 0,2

1.2 1.6FLIGHT VELOCITY,

I0.4

FLIGHT VELOCITYa

2.0 2.41000 ft/sec

I0.6

1000 m/sec

2.8 3.2

I !

0.8 1.0

VAN NUVS. CALIFORNIA


Page 158

Eng. No. 22

Figure K EXIT NOZZLE AREA

SENSITIVITY ANALYSIS

RATIO

RANGE

SUBSONIC COMBUSTION RAMJET

24

U3,¢1:

<

0m

.¢Cn.

<LUn-<

k-

XUJ

UJ-JNNOZ

2O

16

12

8

4

00

I0

_mlm

/\%

J,/4

/,

1 2 .3 4 5FLIGHT VELOCITY, I000 ft/sec

I I i0.5 1.0 1.5


6

I

2.0

VAN NUV'S, CALIFOIINiAu

Eeport 25,194Volume 7

Page 159

Eng. No. 22

Figure LINLET PRESSURE RECOVERY

SENSITIVITY ANALY SIS RANGE


1.0

0I,--

t_

c_JI-.

O.

q,b

>.e,.IJJ

0f.J1.1.1e,-

UJe,-

u')u9I.tJn-IZ.

0.8

0.6

0.4

0.2

05

L

1.5

6

_._ _ _ _ '8 _h.. _

_KE -965 -'-' "

, I7 8 9 10 11 12

FLIGHT VELOCITY, 1000 fV_c

I I I I I

2.0 2.5 3.0 3.5


13

H

rquamz


Page 160

Eng. No. 22

SENSITIVITY ANALYSIS - RESULTS

For the reference conditions stated in the previous section, resulting

specific impulse and thrust perturbations are presented here. Performance

is normalized to the baseline trends given over the appropriate flight

velocity ranges.

The specific impulse and thrust data are both displayed on individual sheets

for each sensitivity variable. On the same sheet, a miniature plot of the

absolute specific impulse and thrust baseline characteristic is shown for

nominal reference purposes. For precise readings, the full-sized curve

appearing previously (its page number is indicated) should be referred to.

The section concludes with a plot reflecting subsystem weight variations

on uninstalled engine thrust/weight ratio.

m

_J

E.

Is1

Report 25 194rMt.4._ Volur_e 7

Page 161


EJECTOR MODE

Page iI_5 _" _|ASILII[ SP[Ctf*¢ I_UL$[

CslCt_ _@D!

Eng. No.

-'_-I L I ! I i t 1 Z ! t i ] ; I l_z_J_I_:J] Baseline

-1i J

j l'iilJi!llh_l PT2/PTo:_1 1 I i I 1 I I I ! I I_t'l I I I-I 1 t I _ _ I 1 ! t I/l- ! I I I I

I _tt t I tl t t_ri it i _l I Figure EI11 I i I 1.14 _ E { I I 1 ti,.I I ! I 1 _ill {.t I ! I il (Page 152)

... ,1-'I'_! I I i I I I I I I T ! I 11I i *.* i.| iJ l._ I.* I.* I,e =,1

22

|,-

j_

I*

I

a i

:= !

I

r- i

I

A

Page lh6 " " "|_$$LII[ IH|U$1

#. 1.4o-, 1

'riKE = 0.92

u. 0.80P._. 0.60

0,401.60

F-u_

.-I-F-

I--{,o

e,,

1.40

1,20

1.00

0,80

0,60

0,40

!

0

am

'l

mm

,5

_---- _/KE = 0.92

, I

0.2

I

1.0 1.5 2.0


1 !0.4 0.6

m n

2.5

I

0.8

FLIGHT VELOCI' O0 m/sec

i i

EQU I VALENCE RAT I0

EJECTOR MODE

Report 25.194Volur6 e 7

Page 16Z

EFFECT

1.15

_ 1.10uJw i

_!_ 1.o5

__ 1.oo!

._..u._-o.951--I

o.9oi

o.8511.15

0.85

L0

Page IL_5 " " _IaSlL_umt SPt¢IFIC _nPU_S!

i Jl c'I@ll mOOl

]d +tm

| ,,,,,

i-

+,NI

J"F+ i

_+ =.+ 4.1 l.l

_m

i •

I.I0

m

m

1 .I0

0.5 1.0

FLIGHT

I

0.2

FLIGHT

Eng. No.

Baseline

9_ = 1.00

22

Page i£6 '_" °IA$_LIN[ I.|USr

(JlCtOl mOO1

i I II

0.90

_______.n

0.90

1.5 2.0

VELOCITY, I000 ft,/sec

1 I

0.4 0.6

VELOCITY 1000 m/see

mpmm

2.5 3.0

0.8

te

nluan#J JI.Y_'Ir_4F_

PRIMARY

VAN NU¥|, ¢&LIFOINI&

ROCKET

i

COMBUSTION EFFICIENCY

EJECTOR MODE

Report 25, 1 94Volume ?

Page 163i i

EFFECT

Page 145 _""IISILIII SIICIIIC IItVLS|

t Jl_tOl moll

i !!,+""I II i| ._J i i ;

I}. t l

r i j,..,.-

Eng. No. 22

I ] I_'I I

IY[ ili_ .X!!, !+-i ;iii',Is i" i ....

i

i i I I I, 1.1 i+I 31.1

vl&ll_, ;iiii l_m

1.06

o.+ 1uJIm 0.96Q.Q. -

0.94

Baseline

*= 0.975C

!,|,., .I =

. I

Page 146 _ " "Idllll.lll IIIIU It

IJl¢_Ol Hit

1.06

1.041.02

+ __0.98

= o.9a 1 -I n I L _,/

0.94 0.5 1.0 1.5 2.0 2.5 3.0

FLIGHT VELOCITY, 1000 if/see

I , I I I , I0 0.2 0.4 0.6 0.8

1000 m/sec

PRIMARYROCKETNOZZLEEJECTOR

EFFICIENCY

MODE

Page 1_5|ASlL_N! SPt¢4_tC impu_$!

lJ[¢_Qt MOO1

Eng. No. 22

Page 146IASILIN[ r_RUST

lllCtOl moa!


Page 164

EFFECT

1.03

1.02WI,,

1.OlIz._.

____ 1.oo

E:,'T 0.99

a,o. 0.98¢/I_U')

0.97

1.03

_: 1.02

I--_n 1.01

,w

= 1.00p.

p-_n 0.99

= 0.98p-

0.97

I0

Baseline

_N = 0.98

J

i

0,5

j_

1.0 1.5

FLIGHT VELOCITY,

I I

0.2 0.4

FLIGHT VELI

1.00

/_'- 0.95

/

1.00

.._a20.95

2.0 2.5 3.0

I000 ft,/sac

I0.6

000 m/sac

I

0.8

ugln_J_Jd=riO.

_m_m

f.)f.._

,r,rml i

(._uf.}Iiii

{hi cn

/

Report 25 194

_. Volun_e 7

/.4 mKlUOrdl .........,,,=.,. P_ge_6s

MIXING EFFICIENCY EFFECT

EJECTOR MODE

Page 145 _ ""I&sli, ll( $P|¢111c lifgksf

11|¢I01 moil

i ! i i ...r, _1

;-, t. _!!!!!ill_ j,=; ,:,. j,ij.iii2.iL.ii• Ii.I *.z

l,t,llm wl,lCl_. Iiii i_

A ,_, .t. ,L. .'., ,I,n_ll_ _, ,iiol,ii

Eng. No. 22

Baseline

19M = 0.80

Page i_,6|ASILOml lU|USTIJI¢I01 m@|[

!,I"--'.:=.r

;++l.l ,III

J.i.I • ..,.+

;_j

& ,4, ,., 4, ,.,_m,s w. m m

,,2

n,I--

,Y-r-F-

I--(/)

n,..i-

1.06

1.04

1.02

1.00

0.98

0.96

0.94

1.06

1.04

1.02

1.00

0.98

0.96

0.94

m

i

//

i i

--m

_,.,,i._ _

1.00

"-T-o.50

n"--

|

0

/f

i|

/i i

0.5 1.0 1.5 2.0


I I I ,

0.2 0.4 0.6

FLIGHT VELOCI I00 m/see

j,i

zoo__0.50 _.

w._i/"

2.5

I

0.8

m

3.0

HEAT EXCHANGER

EJECTOR

EQUIVALENCE

MODE

RATIO

Page 1_5IASlUlII SPI¢IFI¢ ImpULSl

{JtCt@l mO|l

Wiq

,=1!!Ii iii¸

Illl : : :

I ,

i "3t_l _,i_ p o.4 ' *i

Eng. No. 22

Baseline

4)HXFigure F

(Page 153)

Page 146IASlLSUl _wlust

(_fctom mop(

1.30

UJ

= ..j I.I0

__m 1.oo

°J°IT,;" 0.90

"'"' 0.80D.(_.

0.701.03

7

9

-- 1.02

p.u_ 1.01

= 1.00l--

P" 0.99

= 0.98

0.97

1

0

0.5

9

7-_

1.0 1.5

FLIGHT VELOCITY,

I0.2

I

0.4

I000

I000

/

2.0

ft/sec

0.6

m/see

/


Page 166

EFFECT

2.5 3.0

I

0.8

_.I_Ir,_cj

,";"T_,.

AFTERBURNER COMBUSTIONii

EFFICIENCY

EJECTOR MOD E

Page 145 _"" °IIAS|LIm¢ SP(l_lfll_ IglPUL$|

I_IClO| moll

Eng. No. 22

Baseline

7/ =0.95C

z.;

i

e_ i l-

i

o., L

Page 146|AStLm! IMn$!

IJIC?OI ION

1.06

1.04

N 1.o21.00

0.98

0.96

0.94

1.06

.: 1.04

a-

P" 1.026")

= 1.00p-

_" 0.98

,w

::: 0.96p-

0,94

0

J_

I

0.5

f

I I

1.0 1.5 2.0


0.2 0.4 0.6

FLIGHT VELO( 1000 m/=ec

I

f


Page 167

EFFECT

1.00

1.00

1

0.85

2.5 S.O

I0.8

VAN NMY$. CAtlPOONM


EJECTOR MODE

EFFECT

Report 25,194Volurne 7

Page 168

1.06

_|uJ

_.J 1.02

____ 1.oo

°t°,r_. 0.9

LUUJ_._. 0.9¢nu_

0.9

1.06

,c 1.04

F--u_ 1 02_ •

" 1 00

_" 0.98oo

"" 0.96I,-

Page 125 _" =IiilLim( $PI=i_I¢ tmpuis!

[Jiclon moo(

Eng. No. 22 ,.,

Page 126$ASlLbm! t.lU STlilC10e iOOÁ

1.00__

m

m_

1.00__

t.5 1.0 1.5 2.0 2.S


i I I I

0 0.2 0.4 0.6

FL_G_T, ,000 m/sec

0.95-"T-"

B.O

I

0.8

/

AREA RATIO

MODEEJECTOR

EFFECT

Report 25_ 194Volur_e 7

Page 169

.J

E

1.06

1,04

..i 1,02

0.

__ l.oo

E 0.98

0.96on

0.94

1.06

1.04

1.o2

1.00

k-0.98

o¢= 0.96

0,94

I0

Page 1_5 _ ""IASILOlll SPECIFIC Im@_L$(

IJICIOI mill[

i -IIIl_1 I I i i [ i I l 1 _ vl I

i I I I !] I I _ _ i /_ i I: II_ till

l_t+ JillI _ i I + I i._ I I i t I i I

i I I i L,,_I t ] i I I i I i I

-r:_,_. i L ] ,!, L L,Ll l ,!.i,!

_, ,_, ,_. _,

Eng. No. 22

Baseline

A6/A5:

Figure J(Page 157)

z;,

0.75 B_,_ _-

1.25 BL /'v/

.5

Ol. 'sBL/f1.25 BL /

/

1.0 1.5


I0.2 0.4

FLIGH" 000

2.0

ft/sec

i0.6

m/sec

Page i_6 _""las|Loml 1NIU|T

IJ1¢I_1 moll

2.5

I

0.8

3,0

INLET PRESSURE

RAMJET

RECOVERY

MODE


Page 170

EFFECT

Page 147|_SlLIm! splcJp,¢ i_pu_s!

sulsom_¢ ¢oml_tlOl i_mJHPage 148IASILim[ TWIUS!

_U$$_I¢ C@mO_|tlGI nA@J(r

1.30

F'.o

Ioo__I__ 0.90-l--

°J°uJ,,, 0,80Q,n

0.70

1.30

su_r _tucar,. L_ vm

Eng. No. 22

Baseline I

PT2/ : _-PT 0 i

Figure E _"_(Page 152) !

[

/J/

M IL-E-5OOSB --_ '_ /

KE

IXp-

IX

k-

k-

1.20

I.I0

1.00

0.90

0.80

0.700

L0

J

MIL-E-5OO8B -'_" / /

_KE

I1

= 0.92

0.5

2 3 4 5 6


1.0 1.5


7 8

2.0 2®5

COMBUSTOR EQUIVALENCE

RAMJET MODE

RATIO

Report 25 194J

Volu.n'_ e 7

Page 171

EFFECT

1.60

_="J'Ju_l_1.20:Era 1.00

0.80ElL

°F_ _ 0,60'

0.40

1,60

1.40

z.2o

l-¢n 0.80

= 0.60k-

0.40

PB ge 147 " " =

$_|somlc c|=|ust¢om |AmJlt

i i : ; !

i : "N

. i!! !__n,N ! }1,!._ ii

. . I I

'" I [ "

1_-: i i i ' , •

A ", _I, ,2, Z,

0 1 2

I

0I

0.5

Eng. No. 22

Baseline

= 1.00

Page i_ _ " =sal |LOmt tNmU|T

su|s_lc ¢oMlu$_lo_ mLmhll

Ii i i/kTi ] i

I I/i I I_L [ I !

,I vl ! I i'i,I _

] i/ [ I i i : i n i'!

/

i i III] Ii i ] I

_ 0.50

..j--

1.50

1.50

0.50

r.,,... ,.,.._ _. _

3 4 5

FLIGHT VELOCITY, 1000 ft,/secI I

6 7

I

2.01.0 1.5


8

I

2.5

VAN NMY|. CALIFQINtA

COMBUSTOR COMBUSTION EFFICIENCY

RAMJET MODE

Page 147 _ " _lasl%pq! spq¢_i¢ AmP.LS(

su|som4c COMlUST_O_ aAMJlT

Eng. No. 22

Baseline i!

=o.95 j; I

,L

1.15

r 1.1oujluj_U_l_n 1.0 5

..i.Jl

rim..

____ _.oo

(-'__ 0.95u.tu-

uJ[_ o 900.85

i.I

1.00i

__ --" "-------_ _ 0.85

-.:I.I0

F--u', 1.05

" 1.00I,,-

I--u_ 0.95

O_"" 0.90F-

1.00i

_-_ .__ .... 0.85

0.850

I

0

i 2

I

0.5

3 4 5


ft,/sec

m/sec


Page 17Zii

EFFECT

Page 114.8 ....IA s [L*m! i.llu_1

6 7 8

I

2.0I

2.5

Report 25 194

Volu_e 7

rquardt ....., ., ,.., page73I t,TJll_#Mtl_

EXIT NOZZLE EFFICIENCY EFFECT

RAMJET MODE

Page IL_V "-" "i_SliIIl $_IGlflC mPULSl

iUlSOql¢ COmIUiTIOU e_aill

Eng. No. 22|

1,. Baseline ! "i l"

I/N = 0.98 j.i

lloT mJc_v, ll_ lyre

r_ _ Im• wm

Page 148 ....

|A$|LIq| THinS1lll$OlllC lCIM|U$ltOi llIl[l

,.I I t I i # _ki 1 iI i i i [/! I_] i ' !

_Ill V Ii_, iI ! ] I.,4 ] : !"% i

] 1 _ _ ! ! i i : --i

_1 i I,"I, i i i : ;I!lI![ i Ii,'

.{i!,}_!,_Ji :

.: I.I0

1.05--

1.ooI,-.

0.95

=: 0.90k-

0.850

I

0

_I.00-

|i

___0.95_

_m

_mm_

1.00

.0.95

1 2

I

0.5

3 4 5

FLIGHT VELOCITY, 1000 ft,/se¢I l

1.0 1.5

FLIGHT I000 m/see

6 7 8

!

2.0• I

2.5

AREA RATIO

MODE

EFFECT

Report Z5,194Volume ?

Page 174

Page lit _ _ =|ASEtlml SPt¢IFOC tmPuL|l

SUgI$OIIC Comoustno0o llAmJ{f

• , , m

Eng. No. 22

Baseline

A6/A 5 •

Figure K(Page 158)

1.03

jj,o.(Z)o_..nJ_ 1.01

_o/

_ 0.99 BL //"_/-o981.25 BL

0.97 ' J

1.03

1.02

0¢k-

1.01

=: 1.00k-

k--¢n 0.99

"_ 0.98f-.

0.930

I

0

f

1 I

1 2

I0.5

5 4 5 6


1.0 1.5

FLIGHT 1000 m/sec

Page li8 '__ =

|ASILIN! l.lus_Sul$Omt_ CO_IuSlION IAM*£T

7 8

!

2.0I

2.5

/,, ......,..,,,..,.INLET PRESSURE

SCRAM JET

RECOVERY

MODE

EFFECT

pk

Keport 35 194Volun_e 7

Page 175n

Page ib-9

IalILI_E _Ptc,;l_ u_Pu_S!SuplISOml_ ¢OmIUSllON iAmji1 Eng. No. 22

Page 150 '':

IASILItl IWIUII_,uPtlsomlc ¢olleu$11OlO IAmjI1

,LW_' VtLIC.I,. *IO0 *..

:_+ l'o ZI +'.* ];_ ,,

Baseline

PT2/ :PT 0

Figure L(Page 159 )

ui_ _lUlO_ 10w ,n,

zl, _.i ,_e ,'i ,[.6

1.15

1.05 "-- I'/ = 0.985m KE

1.00 '--" -" " .....

l1.5

•r/KE- 0.985

6 7

I

2.0

965

___ _ I

.,._._,KE = 0.965

8 9 I0 ii

FLIGHT VELOCITY, I000 K/secI I

2.5 3.0

FLIGHT VEL _ITY, 1000 m/see

12

I3.5

13

14.0

%L

COMBUSTOR EQUIVALENCE


Page 176

RATIO EFFECT

SCRAM JET MODE

Page 149

IAS[Llml SPI¢_F_¢ hI*PU_S!SUP[ISOmI_ ¢OmlUST_Ot t*mJlt

[,4 'IL _

Eng. No. 22

Baseline

= io00

Page 150 ....

llSlLI_I twlU$+SU*IRSOmlC Comlustlom lJ_Jlt

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12 13

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z_eport 25 i94

rs_ Volu_e 7

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COMBUSTOR COMBUSTION EFFICIENCY EFFECT "_

SCRAM JET MODE

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m m ----------/

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FLIGHT VELOCITY, 1000 ft,/secI I

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Report 75 194Volu_n5 e 7

Page 178

EFFICIENCY EFFECT

SCRAM JET MODE

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Page 149 .... Page 150 ....

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,,u_,, vlu_'r,,. L_ .,.

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"_ 1.20

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• Report Z5,194

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Page 14-9 _" = Page 150 _""

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Eng. No, 22

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Page 180

Eng. No. 22

EFFECT OF SUBSYSTEM WEIGHT VARIATION

ON ENGINE THRU ST/WEIGHT

+ 10 PER CENT WEIGHT VARIATIONn

17.2

17.00

_ 16.8 f DE IGN POINT/

7 _(D__, 56.4 SUBSYSTEM "N_ _,_--t ('_ LINES, CONTROLS , .

__L(_ EXIT NOZZLE SUBSYSTEM. •=_ 16.2_p I I I I I I

_z ---(,_) M IXER/DIFFUSER/AFTERBURN ER

,., 16.0 ---(_ PRIMARY ROCKET SUBSYSTEM_1 I I I \®_)HEAT EXCHANGER UNIT

15.80.90 1.00 1,10


EFFECT OF SUBSYSTEM WEIGHT VARON ENGINE THRUST/WEIGHT

. eport Z5,194Volume 7

Eng.

IATION

Page 181tl

No. 22

+ 50 PER CENT WEIGHT VARIATION

21

20

0 19 ' •

=_ 18y DESIGN POINT

o _,,,_E,,,, ,,__

14 -

130.50 1.00 1.50


UNCLASSIFIED

RePort 25.19_

Volume 7

DISTRI_JTION

Copy No.

i.

.

.

4to7.

8 to 32.

33.

34.

35.

Transmitted to

NASA Office of Resident Representative

Jet Propulsion Laboratory

4800 Oak Grove Drive

Pasadena, California 91103

Attn.: Office of Technical Information


Jet Propuls ion Laboratory



Attn. : Contracts Management Division


Jet Propulsion Laboratory4800 Oak Grove Drive


Attn. : General Counsel for Patent Matters

NASA Headquarters

Washington, D. C. 205_6

Attn. : Chief, Liquid Propulsion Technology, RPLOffice of Advanced Research and Technology

NASA Scientific and Technical Information Facility

P. O. Box 33

College Park, Maryland 20740

NASA Headquarters

Washington, D. C. 205h6

Attn. : Mr. Vincent L. Johnson

Director, Launch Vehicles and Propulsion, SV

Office of Space Science and Applications

NASA Headquarters

Washington, D. C. 20546

Attn.: Mr. Edward L. Gray

Director, Advanced Manned Missions, Mr

Office of Manned Space Flight

NASA Ames Research Center

Moffett Field, California 24055

Attn.: Mr. Clarence A. Syvertson

Mission Analysis Division

UNCLASSIFIED

rquamz ,.,,,o..,.01.wYlv_4rXa% "

UNCLASSIFIED

Report, 25,194

Volume 7

DISTRIBUTION (Continued)

Copy No.

36.

Transmitted to

Marshall Space Flight Center

Huntsville, Alabama 35812

Attn. : Technical Monitor

Mr. Donald Thompson

37. Langley Research Center

Langley Station

Hampton, Virginia 23365

Attn. : Mr. Kenneth Pierpont

38. NASA Ames Research Center


Attn. : Mr. Thomas Gregory

Mission Analysis Division

39, 40. NASA Ames Research Center


Attn.: Harvey J. AllenDirector

hl, 42. NASA Goddard Space Flight Center

Greenbelt, Maryland 20771

Attn.: Merland L. Moseson

Code 620

43, 44. NASA Jet Propulsion Laboratory

California Institute of Technology4800 Oak Grove Drive

Pasadena, California 9]-105

Attn.: Henry Burlage, Jr.

Propulsion Div., 38

45, 46. NASA Langley Research Center

Langley Station

Hampton, Virginia 23365

Attn.: Dr. Floyd L. Thompson

Director

_7, 48. NASA Lewis Research Center

21000 Brookpark Road

Cleveland, Ohio 44135

Attn.: Dr. Abe Silverstein

Director

UNCLASSIFIED

rquarer ....o.,.,,,o,.,.

UNCLASSIFIED

Report 25.19h

Volume 7


Copy No. Transmitted to

49, 90. NASA Marshall Space Flight Center

Huntsville, Alabama 398L2

Attn.: Hans G. Paul

Code R-P&VED

51, 92. NASA Manned Spacecraft Center

Houston, Texas 77001

Attn.: Dr. Robert R. Gilruth

Director

53, 54. NASA Office of Resident Representative

Jet Propulsion Laboratory



Attn.: Robert W. Kamm

Director

55, 56. NASA John F. Kennedy Space Center

Cocoa Beach, Florida 32931

Attn. : Dr. Kurt H. Debus

57. Aeronaut ical Systems Divis ion

Air Force Systems Command

Wright-Patterson Air Force Base

Dayton, Ohio 49433

Attn. : Do L. Schmidt

Code ASRCNC-2

_8. Air Force Missile Development Center

Holloman Air Force Base, New Mexico

Attn.: MaJ. R. E. BrackenCode MDGRT

59. Air Force Missile Test Center

Patrick Air Force Base, Florida

Attn. : L. J. Ullian

60. Air Force Systems DivisionAir Force Unit Post Office

Los Angeles 49, California

Attn. : Col. Clark

Technical Data Center

UNCLASSIFIED

uard/van NUV|. CA¢IWOeNeA

UNCLASSIFIED

Report 2_, 194

Volume 7

DISTRI_JTION (Continued)

Copy No.

61.

Transmitted to

Arnold Engineering Development CenterArnold Air Force Station

Tullahoma, TennesseeAttn. : Dr. H. K. Doetsch

62. Bureau of Naval Weapons

Department of the Navy

Washington 2_, D. C.

Attn. : J. Kay_r_s-41

63. Defense Documentation Center Headquarters

Cameron Station, Building 59010 Duke Street

Alexandria, Virginia 22314

Attn.: TISIA

64. Headquarters, U. S. Air Force

Washington 25, D. C.

Attn.: Col. C. K. Stambaugh

AFRST

65. Picat inny Arsenal

Dover, New Jersey 07801

Attn. : I. Forsten, Chief

Liquid Propulsion Laboratory

SMLrPA-DL

66. Air Force Rocket Propulsion Laboratory

Research and Technology Division

Air Force Systems Command

Edwards, California 93523

Attn.: H. Main, RPRR

67. Research and Technology Division

Air Force Aero Propulsion Laboratory

Wright-Patterson Air Force Base, Ohio 45433

Attn. : H. C. Long, APRP

68. U. S. Atomic Energy CommissionTechnical Information Services

Box 62

Oak Ridge, Tennessee

Attn.: A. P. Huber

Oak Ridge Gaseous Diffusion Plant

(ORGDP) P. O. Box P

UNCLASSIFIED

/i Ua/'tH

UNCLASSIFIED

van NUV$ ¢ALIFOIIN#a



9. U. S. Army Missile CommandRedstone Arsenal

Alabama 35809

Attn. : Dr. Walter Wharton

70. U. S. Naval Ordnance Test Station

China Lake, California 93997

Attn.: Chief, Missile Propulsion DivisionCode 4962

71. Air Force Systems Command

Space Systems Division

Ing!ewood, California 90045

Attn. : Lt. Col. Kenneth Mc Aaron

72. Chemical Propulsion Information Agency

Applied Phys ics Laboratory

8621 Georgia Avenue

Silver Spring, Maryland 20910

Attn. : Nell Safeer

3. AeroJet-General Corporation

P. O. Box 296

Azusa, California 91703Attn. : L. F. Kohrs

74. AeroJet-General CorporationP. O. Box 19_7

Sacramento, California 95809

Attn. : R. St iff

Technical Library, Bldg. 2019, Dept. 2410

75. Aeronutronic

Philco CorporationFord Road

Newport Beach, California 92663

Attn.: D. A. Carrison

76. Aerospace Corporation

2400 East E1 Segundo Boulevard

P. 0. Box 99089

Los Angeles, California 90049

Attn.: John G. Wilder

Propulsion Dept. _-2293

UNCLASSIFIED

ualdl_t,l_A.ifl_rt4rHt_ * van NUW|. CaLtwOIImfa

UNCLASSIFIED


Copy No.

77.

Transmitted to

Arthur D. Little, Inc.20 Acorn Park

Cambridge, Massachusetts 02140

Attn.: E. Karl Bastress

78. Astropower Laboratory

Douglas Aircraft Company2121 Paularino

Newport Beach, California 92663

Attn.: Dr. George Moc

Director, Research

79. Astrosystems International,

1279 Bloomfield Avenue

Fairfield, New Jersey 07007Attn.: A. Mendenhall

Inc •

0. Atlantic Research Corporation

Edsall Road and Shirley Highway

Alexandria, Virginia 22314

Attn.: A. Scurlock

81. Beech Aircraft Corporation

Boulder Division

Box 631

Boulder, Colorado

Attn.: J. H. Rodgers

82. Bell Aerosystems Company

P. O. Box 1

Buffalo, New York 14240

Attn.: W. M. Smith

Bendix Systems Division

Bendix Corporation

3300 Plymouth Road

Ann Arbor, Michigan

Attn.: John M. Brueger

84. Boeing CompanyP. O. Box 370T

Seattle, Washington 9812hAttn.: J. D. Alexander

UNCLASSIFIED

rquamt .,..o,,.:.,.,o..,.le:t_lf_4rxl_

UNCLASSIFIED

Re_ort 25, 19h

Volume 7

DISTRI_JTION (Continued)


89. Missile Division

Chrysler CorporationP. O. Box 2628

Detroit, Michigan 48231

Attn.: John Gates

86. Wright Aeronautical Division

Curtiss-Wright Corporation

Wood-Ridge, New Jersey 07079

Attn.: G. Kelley

87. Missile and Space Systems Division

Douglas Aircraft Company, Inc.

3000 Ocean Park Boulevard

Santa Monica, California 90_06

Attn.: R.W. Hallet

Chief Engineer, Advanced Space Tech.

88°

89.

Aircraft Missiles Division

Fairchild Hiller Corporation

Hagerstown, Maryland lO

Attn.: J. S. Kerr

General Dynamics/AstronauticsP. 0. Box 1128

San Diego, California 92112

Attn.: Frank Dore

Library& Information Services (128-00)

90. Re-Entry Systems Department

General Electric Company

3198 Chestnut Street

Philadelphia, Pennsylvania 19101

Attn.: F. E. Schultz

91. Advanced Engine & Technology Dept.

General Electric Company

Cincinnati, Ohio k5215

Attn.: D. Suichu

92. Grumman Aircraft Engineering Corp.

Bethpage, Long Island

New York

Attn.: Joseph Gavin

UNCLASSIFIED

UNCLASSIFIED

_eport 2.5, !9h

Volume

DZSTRZ O ZON(Continued)


95. Ling-Temc o-Vought Corporation

Astronautics

P. 0. Box 5907

Dallas, Texas 75222

Attn.: Warren C. Trent

94. Lockheed-Californla Company

2555 North Hollywood Way

Burbank, California 91503

Attn.: G. D. Brewer

95.

96.

Lockheed Missiles and Space CompanyP. O. Box 504

Sunnyvale, California 94088

Attn.: Y. C. Lee

Technical Information Center

Lockheed Propulsion CompanyP. O. Box iii

Redlands, California 92574Attn.: H. L. Thackwell

9T.

98.

Baltimore Division

Martin-Marletta Corporation

Baltimore, Maryland 21203Attn.: John Calathes (3214)

Denver Division

Martln-Marietta Corporation

P. O. Box 179

Denver, Colorado 80201

Attn.: J. D. Goodlette (A-241)

99. McDonnell Aircraft CorporationP. O. Box 516

Municipal Airport

St. Louis, Missouri 65166

Attn.: R.A. Herzmark

i00. Space & Information Systems Division

North American Aviation, Inc.

1221_ Lakewood Boulevard

Downey, California 90241

Attn.: H. Storms

UNCLASSIFIED

le;rwl_Har#a_

UNCLASSIFIED

Report 25,194

Volume 7

DISTRI]_JTION (Continued)

Copy No.

i01.

Transmitted to

Rocketdyne

North American Aviation, Inc.

6633 Canoga Avenue

Canoga Park, California 91304

Attn. : E. B. Monteath

Library, 586-306

102. Northrop Space Laboratories

3401 West Broadway

Hawthorne, California

Attn. : Dr. William Howard

103. Astro-Electronics Division

Radio Corporation of America

Princeton, New Jersey 08940

Attn.: S. Fairweather

104. Reaction Motors Division

Thiokol Chemical Corporation

Denville, New Jersey 07832

Attn.: Arthur Sherman

105. Republic Aviation Corporation

Farmingdale Long Island, New YorkAttn. : Dr. William O'Donnell

io6. Space-General Corporation

9200 East Flair Avenue

E1 Monte, California 91734

Attn.: C. E. Roth

107. Stanford Research Institute

333 Ravenswood Avenue

Menlo Park, California 94025

Attn.: Lionel Dickinson

108. Space Technology Laboratories

TRW Incorporated

One Space ParkRedondo Beach, California 90278

Attn. : G.W. Elverum

UNCLASSIFIED

vm,_ UUYS CAl._POIrNtdl

UNCLASSIFIED

DISTRI_JTION _Continued)

Copy No.

109. TAPC0 Division

TEW, Incorporated

23999 Euclid Avenue

Cleveland, Ohio 4hll7

Attn.: P. T. Angell

Tramsmitted to

llO. Thiokol Chemical Corporation

Huntsville Division

Huntsville, AlabamaAttn.: John Goodloe

lll• Besearch Laboratories

United Aircraft Corporation

400 Main Street

East Hartford, Connecticut 06108

Attn.: Erle Martin

United Aircraft Corporation

Pratt &Whitney Aircraft Division

East Hartford, Connecticut 06108

Attn.: W. Sens

]_13. UnitedAircraft Corporation

Pratt &Whitney Aircraft Division

Florida Research and Development Center

West Palm Beach, Florida 33_01

Attn.: R. Coar

UNCLASSIFIED

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