SUMMARY OF MACH 4 INTEGRAL RAMJET STUDY DURING THE PERIOD 1 JANUARY TO 15 JULY, 1959

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• Document Creation Date: December 28, 2022
• Document Release Date: February 9, 2017
• Case Number: F-2015-02619
• Publication Date: February 8, 1960

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DATE 8 February, 1960
REPORT 5808
This Document contains Information affecting the Notional Defense of the United States within the meaning of
the Espionage Act .50 U.S.C., 31 and 32, as amended. Its transmission or the revelation of its contents in any
manner to an unauthorized person is prohibited by low.
CONF NU AL
(Title -- Confidential)
SUMMARY OF MACH 4 INTEGRAL RAMJET STUDY
During the Period
1 January to 15 July, 1959 (b)(3)
Contract
Project 216
(b)(3)
CONHDYNT IAL
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THE
III ENGINE DEVELOPMENT PROGRAM 6
A. Configuration Development 6
10
IV CONTROLS 11
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REPORT 5808
CONTENTS
Section Page
I. INTRODUCTION 1
II GENERAL CONCEPTS 1
A. Applications and Performance 1
3
B. Components and Materials
B. Full Scale Design
A. System Functions
B. System Concept
C. System Design
D. Environmental Considerations
E. Installation and Ground Check Features
F. Air Turbine Motor Accessory Drive
G. Development Status
13
14
15
19
20
21
23
V CONCLUSIONS 23
TABLE II --
TABLE III -
TABLE IV --
APPENDIX A
TABLE I -- Typical Trajectory Variables 2
Results of Small Scale Burner Configuration
Development Test 8
9
12
-- Engine Model Specification Including Air Induction
Control and Actuation System 62
- Results of 30-inch Scale Model Engine Burner Develop-
ment Tests
Weight Breakdown for Mach 4 Integral Cruise Type
Ramjet
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ILLUSTRATIONS
Figure
Effect of Mach Number on Range Parameter
Typical Operating Envelope
Page
1.
2.
24
.25
3.
Acceleration and Cruise Performance
26
4.
Mach 4, 30-inch Diameter Structural Test Engine, Side View
27
5.
Schematic of Mach 4 Cruise Type Ramjet
28
6.
Diffuser Total Pressure Recovery
29
7.
Mach 4, 30-inch Diameter Structural Test Engine, Looking Aft. . . .
30
8.
Material Properties at Elevated Temperatures
31
9.
Nozzle Efficiency vs. Secondary Flow
32
10.
Segmental Burner Tests, Combustion Efficiency Evaluation
33
11.
Segmental Burner and Components
34
12.
Combustor Performance, 30-inch Diameter Engine
35
13.
Major Subassemblies
36
14.
Maximum Operating Metal Temperatures
37
15.
Material Selections
38
16.
Prototype of Flight Engine
39
17.
Schematic of Propulsion System Inputs and Variables for Control . .
4o
18.
Thrust and Specific Fuel Consumption Sensitivity to Engine
Performance Characteristics
41
19.
Block Diagram of Inlet Control System
42
20.
Block Diagram of Engine Control System
43
21.
Signal Parameter Suitability, Inlet Control System. � � .
44
22.
Schematic of Inlet Control System
45
23.
Schematic of Bypass Control System
46
24.
Engine Control System Schematic
47
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ILLUSTRATIONS (Continued)
REPORT
98o8
Figure
Page
25.
Manual Inputs, Propulsion Control System
48
26.
Controlled Engine Output Power Characteristics
49
27.
Fuel-Air Ratio Accuracy, Engine Control System
50
28.
Air Mass Flow Computer Performance
51
29.
Acceleration and Cruise Control Characteristics
52
30.
Turbopump Fuel Flow, Pressure Rise, and Speed Characteristics . . .
53
31.
Bypass Door Operating Characteristics
54
32.
Exit Nozzle Areas During Ignition and Maximum Power Operation . . .
55
33.
Layout of Control Package
56
34.
Control Performance at Elevated Fuel and Air Temperatures
57
35.
Setup for Elevated Temperature Test of Control
58
36.
Diaphragm Motor Life -- Temperature Relationship
59
37.
Schematic of Air Turbine Motor Unit
60
38.
Layout of Air Turbine Motor Unit
61
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I. INTRODUCTION
In 1958, the state of the art of development of materials and ramjet
components had reached the stage wherein a high speed (Mach 4) and high altitude
(100,000 feet plus) ramjet engine appeared feasible for development and applica-
tion to long range cruising vehicles. Aerodynamic test data, coupled with engine
component data, revealed that long range capabilities increase rapidly with super-
sonic Mach number as shown in Figure 1 and there is appreciable advantage to
pushing cruise speeds as high as material technology will allow.
Studies and component tests made in 1957 and 1958 of an integral
cruise type ramjet, as applied to the Super Hustler vehicle, showed that a light-
weight ramjet engine could be rapidly developed using existing state of the art
knowledge. Consequently, The Marquardt Corporation and the Air Force entered into
a program to do further development of engines of this general type which could
have application to future ground or air launched cruise vehicles. This study was
initiated in January, 1959 and completed with the fabrication of a prototype
engine on 15 July, 1959.
GENERAL CONCEPTS
A. Applications and Performance
A representative altitude�Mach number operating envelope for such
an advanced cruise engine is shown in Figure 2. For air launched missiles or
manned aircraft, the initial engine operation could occur at subsonic Mach numbers
to supplement booster thrust. At some Mach number between 1.5 and 2.0, depending
on the relative size of the vehicle and engine, the ramjet could take over and
accelerate the vehicle to cruise conditions. In the case of a supersonic air
launch, no supplemental booster system would be required. Another possible appli-
cation of the engine would be with ground launched vehicles, wherein the engine
again would ignite subsonically to augment boost thrust and self-accelerate from
the region of Mach 1.5 to 2.0 to cruise Mach number.
The performance capabilities of an integral ramjet engine of this
type are shown in Figure 3 which shows acceleration thrust and throttled cruise
specific fuel consumption. A minimum cruise specific fuel consumption of 1.86
rbs fuel/lb thrust per hour is obtainable at Mach 4. Tests made in August, 1958
of a full scale engine at the Mach 4 condition demonstrated that such a minimum
specific fuel consumption was attainable. The engine was a flight weight type and
it incorporated the salient features of the Mach 4 integral cruise type engine.
A photograph of this engine is shown in Figure 4.
Table I lists performance variables along a typical trajectory.
Appendix A is a preliminary engine model specification with complete engine per-
formance curves presented on a gas generator basis. Component performance levels
referred to hereinafter as "estimated values" are those used in arriving at the
over-all engine performance presented in the specification.
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MAC A673
TABLE I
TYPICAL TRAJECTORY VARIABLES
Time
(min)
Mo
Alt.
(ft)
Ao
(sq ft)
V8
(PPs)
Tt2
(�F)
Pt2
(Psia)
o
(Psis)
o
(�F)
*
M2
t2
F0
F/A
Pt4
(psia)
A5
Wf
(pps)
F6
(lbs)
F0
(lbs)
CF
NJ
***
NJ
(lbs)
SFC
(1b/hr/1b)
a
Ft2
TE
Pt2
(sq in.)
F2
(lbs)
p2
(psis)
p4
(psis)
sec
ACCELERATION AND
CLIMB
Cold
--
2.0
36,500
2.610
112.9
242
15.78
3.2294
-70.0
0.248
0.625
0
flow
13.88
0.400
o
3,948
6,798
-.233
-2850
c
7.160
527.0
1,124
15.12
13.35
2.0
36,500
2.610
112.9
242
22.82
3.2294
-70.0
0.168
0.904
0.0603
0.90
19.66
0.710
6.676
15,349
6,798
0.6990
8,549
2.811
4.949
864.0
17,298
22.38
17.14
o
2.2
40,000
3.191
128.4
3o7
26.00
2.7305
-70.0
0.176
0.891
0.0590
:-.,.90
22.31
0.710
7.433
18,683
8,509
0.8130
10,173
2.630
4.940
860.7
20,561
25.44
19.45
0.54
2.4
44,500
3.915
138.6
377
28.13
2.2015
-70.0
0.183
0.874
0.0577
0.90
24.04
0.710
7.837
21,137
10,001
0.927611,136
2.531
4.927
857.3
22,942
27.48
20.96
0.84
2.6
48,00o
4.510
146.3
454
31.83
1.8620
-70.0
0.179
0.856
0.0562
0.90
27.76
0.648
8.058
23,265
11,445
0.9917
11,819
2.454
4.596
810.0
26,475
31.13
24.90
1.11
2.8
50,000
4.765
151.2
536
38.52
1.6915
-70.0
0.159
0.837
0.0545
0.90
34.92
0.533
8.076
25,075
12,738
0.9825
12,337
2.357
3.926
710.4
32,374
37.84
32.60
1.26
3.0
52,000
5.019
155.1
623
46.30
1.5372
-70.0
0.142
0.816
0.0526
0.90
43.13
0.441
7.994
26,520
13,998
0.955912,521
2.298
3.350
617.7
39,206
45.65
41.21
1.50j3.2
55,000
5.296
151.2
715
52.77
1.3319
-70.0
0.126
0.795
0.0507
0.90
50.08
0.369
7.515
27,467
15,0600.9289
12,406
2.255
2.866
537.4
44,920
52.19
48.55
1.80
3.4
58,1oo
5.625
147.2
811
59.45
1.1485
-70.0
0.114
0.773
0.0488
0.90
57.12
0.315
7.038
26,415
15,055
0.9037
11,360
2.230
2.476
470.5
50,800
58.91
55.86
2.19
3.6
62,200
5.999
136.6
912
63.33
.94395
-70.0
0.104
0.751
0.0468
0.90
61.40
0.271
6.265
24,928
14,795
0.8746
10,130
2.226
2.157
413.8
54,274
62.86
60.40
2.61
3.8
67,500
6.433
120.1
1018
62.98
.73318
-70.0
0.095
0.728
0.0447
0.90
61.42
0.238
5.261
23,338
14,401
0.8498
8,936
2.223
1.907
369.0
54,097
62.59
6o.66
3.00
4.0
71,000
6.657
110.7
1128
68.01
.62017
-70.0
0.084
0.705
0.0425
0.90
66.75
0.201
4.609
20,775
13,317
0.7939
7,459-
2.224
1.627
317.8
58,481
67.68
66.16
3.00i
4.o
71,000
7.106
118.1
1128
68.69
.62017
-70.0
0.090
0.712
0.0425
0.90
67.27
0.213
4.919
22,138
14,215
0.8433
7,923
2.235
1.720
334.5
59,110
68.28
66.60
3.06
4.0
75,000
7.106
97.60
1128
56.76
.51245
-70.0
0.090
0.712
0.0418
0.911
55.59
0.212
3.998
18,242
11,746
0.8367
6,496
2.216
1.720
333.6
48,844
56.42
55.03
3.15
4.0
80,000
7.136
77.21
1128
44.77
.40370
-70.0
0.091
0.713
0.0410
0.925
43.83
0.213
3.102
14,434
9,292
0.8411
5,142
2.172
1.724
334.6
38,530
44.50
43.39
3.60
4.0
85,000
7.398
62.77
1144
35.63
.31831
-66.7
0.093
0.718
0.0398
0.940
34.85
0.217
2.448
11,692
7,595
0.8497
4,097
2.151
1.762
340.3
30,672
35.42
34.49
4.32
4.0
90,000
7.719
51.33
1175
28.51
.25204
-57.5
0.095
0.723
0.0385
0:956
27.86
0.222
1.937
9,508
6,276
0.8464
3,232--2.157
1.800
345.4
24,556
28.34
27.56
NOMINAL CHuisE
7
4.0190,20o
7.345
48.38
1176
28.14
.24978
-57.2
0.091
0.717
0.020
0.961
27.58
0.203
1.5283
8,596
5,918
0.707
2,678
2.054
1.719
317.5
24,220
27.98
27.36
Tt4 3190�F
leakage . 0.02
except for nominal
cruise where leakage
' 0.0129
M2 based on actual A2 where A2/A3 0.86182
** Fi...Pi Ai (1+ riMi) - PGA'
*** FNJ = CF NJ qo A6
TE P6A6 (1 42r6M62)
rIVI IN= 3NO 0
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Amu 5808
It was originally intended to use a full length variable plug exit
nozzle in the engine and the data in Appendix A represent this concept. However,
it was decided later in the program to incorporate a blunt plug variable exit
nozzle with its attendent advantages of decreased length and weight. For purposes
of saving time Addendum II to the preliminary engine model specification (Appendix
A to this report) was prepared to reflect the weight, length, and performance es-
timate changes. The engine, as described in the remainder of this report, in-
corporates the blunt plug exit nozzle.
B. Components and Materials
A schematic of an engine designed for the envelope of operation of
Figure 2 is shown in Figure 5. The major components of the propulsion system are
1. Inlet diffuser
2. Fuel injectors
3. Combustor
4. Exhaust nozzle
5. Fuel pumping and control system and nozzle actuator and control
system
1. Inlet Diffuser
Although this component is a very important part of the propul-
sion system, the diffuser would be part of the airframe itself for an integral
engine and is of interest only insofar as its performance affects the engine de-
sign. Specifically, the maximum attainable inlet total pressure recovery and mass
flow variation with Mach number determines the variation of engine exit nozzle
throat size. Secondary considerations are the effect of diffuser outlet velocity
profiles on engine performance and control interrelationships between the engine
fuel and nozzle geometry controls and the inlet geometry control.
Figure 6 presents a compilation from a literature survey of inlet
pressure recoveries for variable geometry inlet configurations tested in the range
of Mach numbers of interest. To minimize external drag, an inlet with internal
compression is required at Mach numbers as high as 4.0. Complete internal com-
pression type inlets require considerable bleed and bypass flow to give good per-
formance. Consequently, a mixed internal-external compression inlet was con-
sidered optimum for this application. It has the following advantages:
1. The diffuser boundary layer bleed for high pressure recovery
is small.
2. The variable geometry sections used to obtain high recovery
are relatively small as is their motion.
3. The external compression portion yields a variation in mass
flow with Mach number that tends to match the engine require-
ments. A moderate amount of additional bypass at low Mach
numbers would also be required for complete matching, however.
4. External drag is very low.
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Based upon the data shown in Figure 6, a pressure recovery level
was assumed as shown. This level of inlet performance is considered to be con-
sistent with the attainable performance of other major engine components and
materials.
2. Fuel Injection System
In order to obtain a maximum number of fuel injection points to
facilitate good mixing of fuel and air in a short length, a spray bar system was
selected. The engine flow passage is of annular shape, this being dictated by use
of the cantilevered plug type variable nozzle. Consequently, the burner itself is
annular in shape and there are three circumferential fuel spray bars: one to
supply fuel directly into the burner pilot zone, the other two to supply fuel to
the outer and inner burner annular passages. These fuel manifolds are referred to
as the pilot manifold (center bar) and the main fuel manifolds (outer and inner
bars), respectively. Figure 7 is a view of the engine looking downstream showing
the spray bars.
3, Combustor
The requirements for high combustion efficiency over a very broad
range of burner inlet temperatures, air mass flow, and at fuel-air ratios both
lean and rich dictated selection of a can type burner. Development tests of such
burners at Marquardt for the RJ59 Mach 3 and Mach 4 engine series under Contract
AF 33(600)-22985 provided a wealth of experience and data which not only defined
this burner type as the most feasible for this application, but enabled immediate
design of a configuration of high performance.
The burner, although annular in shape, is divided circumferential-
ly into three separate sements. These are separated by the longerons which sup-
port the center body section and they are placed in the burner section, as shown
in the photograph, Figure 7, to minimize engine length 'and weight.
14. Exhaust Nozzle
To obtain efficient cruise performance at Mach 4, a nozzle of
high thrust efficiency is mandatory. An increase in nozzle thrust efficiency of
1 percent results in a reduction in specific fuel consumption of about 5 percent
and a resultant range increase of about 5 percent. To obtain the large variation
in nozzle throat-to-combustor area ratio required for maximum low speed thrust
(71% Athroat/
Acombustor) and efficient Mach 4 cruise operation (18% A throat/
a plug type exit nozzle was selected as the most desirable. The plug
Acombustor)
itself is segmented and very short, as shown previously in Figure 5. The varia-
tion in area ratio can be obtained in a short length with high nozzle thrust ef-
ficiencies at all area ratios.
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5. Fuel and Geometry Control Systems
REPORT
5808
The center body type engine resulting from the above arrange-
ment of nozzle and combustor provides a convenient location for the fuel and
geometry control systems. The actuation unit for the variable exit nozzle is
located in the aft portion of the center body and the fuel pumping and control
system is located in the forward section. The various elements of these systems
and their functions for manned aircraft or missile application are discussed in
Section IV of this report.
Fundamentally, the control system keeps the nozzle in the open
position and the fuel-air ratio near ztoichionetric for high thrust during
initial acceleration up to Mach 2.5. From Mach 2.5 to 4, the control system re-
duces fuel-air ratio and exit nozzle size to maintain high thrust but not over-
temperature the engine. At cruise conditions, the fuel-air ratio is reduced
further, as is the exit nozzle throat, to maintain optimum cruise specific fuel
consumption.
6. Materials
Materials technology had advanced to the stage where not only
were adequate materials available to fabricate an engine for extended cruise
operation at Mach 4, but a relatively lightweight structure could be developed
using these materials. Temperatures which were calculated for different parts of
the engine revealed that the nozzle throat area would be the hottest part of the
engine required to withstand load and maintain shape. The maximum temperature
here would not exceed 1800�F.
The particular materials selected for certain parts of the engine
are based upon the maximum operating temperature design life, and, of course,
loads. These items are discussed further in Section III.
The materials of particular interest for the engine application
are Rene.' 41 and Udimet 500, which were planned for use in numerous parts of the
engine. .These materials, being newer alloys, were not completely documented as
to short time tensile and creep data. Consequently, a program was initiated to
collect such data using the Marquardt High Temperature Testing machine. The
materials investigated were
1. 422M stainless steel
2. 6A1-4V titanium
3. MST821 titanium
4. 16V-2.5 Al-titanium
5. A286 iron base alloy
6. .AF71 iron base alloy
7. N-155 mixed base alloy
8. R-235 nickel base alloy
9. 1,605 cobalt base alloy
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10. M252 nickel base alloy
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11. Udimet 500
12. Rene' 41
13. Waspaloy nickel base alloy
14. Commercially pure molybdenum
15. 0.5% Ti-molybdenum alloy
16. Tantalum
17. 0.5% Zr-columbium alloy
18. Tungsten
Figure 8 is a summary of the tensile strength-to-weight ratios
at elevated temperatures for several alloys.
In addition, the fabricability characteristics were studied in-
cluding as radial draw forming, flow turning, impact forming, hydroforming, roll
forming, and spinning, as well as fusion, flash, and spot welding.
ENGINE DEVELOPMENT PROGRAM
A. Configuration Development
1. Exhaust Nozzle
Small scale nozzle model tests were initiated early in the
development program to define the most efficient variable exhaust nozzle configu-
ration. As mentioned previously, a nozzle of high thrust efficiency was mandatory
since a small increase in nozzle efficiency is magnified by a factor of about 5
in increased range. Highly efficient nozzles tend to be long, however, and the
variable geometry requirement would make a long nozzle very heavy.
A plug type nozzle was selected for this application since high
efficiency is obtainable in a relatively short length with a plug type nozzle as
compared to a conventional convergent-divergent nozzle. Tests of short length
plug nozzles revealed that a high component efficiency could be obtained with a
plug nozzle with virtually no physical divergent section downstream from the
throat.
A sketch of such a nozzle is shown in Figure 9 together with the
over-all nozzle efficiency with secondary flow through the base of the plug. This
secondary flow forms an "aerodynamic" taper to the plug which results in high per-
formance with a very short length nozzle. The secondary flow could be diffuser
bleed air which has to be discharged overboard, or it could be air taken on board
by enlarging the inlet and ducting the air directly through the engine center body
from the engine face. The ram drag penalty has been accounted for and the re-
sulting nozzle efficiency shown in Figure 9 is that component efficiency which is
applied to the engine gas flow directly.
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2. Combustor Designs and Performance
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Early tests of a 30-inch diameter plug nozzle type engine in
August, 1958 under Contract AF 33(600)-33517 indicated that.high combustion effi-
ciency and burner total pressure recovery at the Mach 4 conditions should be
relatively simple to achieve. The relatively small exit throat at cruise results
in low combustor velocities and law pressure losses. The high inlet temperature
(1200�F) is ideal for high combustion efficiency. These early tests revealed that
efficiencies above 95 percent were obtainable. Configuration development tests
were then concentrated in the low Mach number area (Mach 2.0 to 2.5) where the
large exit throat, high combustor velocities, and the low inlet temperature
(250�F) made attainment of the target objective of 90 percent combustion effi-
ciency more difficult.
In developing the combustor configuration for the full scale
prototype engine, use was made of small scale burner component configuration
development tests. Data were obtained utilizing a segment of the full scale
burner in the Marquardt Aerothermo Laboratory as well as complete large scale
engine testing with a 30-inch diameter engine ihAhe,Marquardt Jet Laboratory..
Table II lists the test periods, number of runs, variables investigated, etc., for
the small scale component development tests. Figure 10 shows typical combustion
efficiency test results obtained from the small scale segmental burner tests and
Figure 11 illustrates the segmental burner and typical components that were used.
Promising configurations from these tests were integrated into
the 30-inch diameter engine design and evaluated. Table III lists the 30-inch
engine test periods, runs completed, total burning time, variables, etc. As can
be seen, nearly all of the burner tests were performed at the low inlet tempera-
ture condition of 250�F.
At the conclusion of the limited engine configuration development
test period, a burner configuration was evolved which gave essentially 90 percent
combustion efficiency at the law Mach number, low inlet temperatures condition as
required. The performance parameter burner drag coefficient (Cab) was also de-
termined from test results to be of the corresponding proper magnitude of 4.0 at
the operating inlet Mach number to the combustor. Figure 12 lists pertinent
combustion efficiency results and gives the burner drag performance of the final
burner configuration.
Pentane, 80-octane gasoline, JP-41 and RJ-1 fuels were evaluated
in developing the burner for the low temperature (250�F) operation. The high
temperature RJ-1 fuel is planned for use in the extended cruise mission.
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MAC A473
rIVI,LNaGI
1
,
,
RESULTS OF SMALL SCALE
Marquardt
TABLE II
,
DEVELOPMENT TEST
Laboratory
BURNER CONFIGURATION
Aerothermo
Phase
No.
Test Dates
Number of
Runs
Completed
Inlet
Temperature
Range
(�F)
Variables Investigated
h
qc
Fuel-Air
Ratio Limit
c
db.
Pilot Fuel
Injection
Main _Fuel
Injeetion
Burner
Geometry
Fuel Effects -
JP, 80 Octane,
RJ-1
I
2-12-59 to
3-5-59 and
3-25-59 to
4-3-59
54
250 to 450
x
x
x
x
x
--
--
II
4-7-59 to
4-15-59
22
250 to 400
x
x
x
x
x
x
x
III
4-29-59 to
5-29-59
18
250 to 400
with A5/A3..65
500 to 1175
with A5/A3=.14
x
x
x
x
x
--
x
Iv
6-16 and
6-17-59
19
700 to 2500
x
x
--
x
--
x
x
Total number of runs . 113
LT)
0
'CS
'CS
S
)
CD
-h
0
-5
CD
CD
SI)
CD
. .
cD
cD
cD
cD
CO
MAC A673
'IVIINaCEI3NOD
RESULTS OF 30-INCH SCALE MODEL
TABLE III
,
DEVELOPMENT TESTS
ENGINE BURNER
Marquardt
Test No.
Test
Dates
Number
of
Runs
Corn-
pleted
Burn-
ing
Time
(min)
Inlet
Temperature
Range
(�F)
Number
Variables Investizated
of
Usable
Data
Points
71c
Fuel-Air
Ratio
Limit
Ignition
Cd b
Pilot
Fuel
Injection
Main
Fuel
Injection
Burner
Geometry
Fuel Effects
80 Octane,
JP, RJ Fuels
2288
Cell 3
2-13 to
2-27-59
16
18.5
230
99
x
x
x
x
x
x
x
--
2406
Cell 3
3-24 to
3-26-59
3
3.6
250
36
x
--
-
--
x
--
x
--
2425
Cell 3
5-8 to
5-13-59
6
8.5
250 to 427
75
x
x
x
x
x
x
x
x
2290
Cell 8
6-2 to
6-5-59
11
46.7
80 to 300
44
x
x
x
x
x
x
x
Totals
36
77.3
254
\J1
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B. Full Scale Design
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MORT 5808
Design of a flight prototype engine was completed during the contract
work period of January to July, 1959. Fabrication of a prototype engine con-
sisting of flight engine components wherever possible was also completed.
Studies made during the RJ59 engine programs revealed that ramjet
engines delivered more thrust and better specific fuel consumption per pound of
engine weight as engine diameter increases. The RJ59 series was developed in
36-inch engine size, since this was considered to be the largest practical engine
diameter consistent with test facility limitations. Facilities considered were
primarily the Arnold Engineering Development Center, Ordnance Aerophysics Labora-
tory, and the Marquardt Jet Laboratory. The combustor flow area of the RJ59
series was approximately 1000 sq in. and the integral cruise type engine is de-
signed with the same flow area, which is a measure of required air flow rates, and,
hence, facility reqUirements.
1. Flight Design
A sketch of the resulting design of the flight type engine is
shown in Figure 13. The engine consists of several subassemblies exclusive of
the fuel and geometry control packages, which are discussed in Section IV. The
forward outer shell is the main structural subassembly and it would transmit axial
loads to the airframe at the forward ring which is designed to attach to the air-
frame with a V-type clamp. The main structural ring would transmit normal
maneuver loads to the airframe at three points through rollers. This whole struc-
tural assembly is exposed solely to inlet air temperatures and receives no heat
from the combustion section.
The longeron--center body assembly transmits all nozzle plug
forces and inertia loads from the center body with enclosed fuel and geometry con-
trol package to the outer structural assembly. The longerons, three in number,
separate the annular burner into three segments and the longerons receive little
or no heat from the combustion region.
The variable plug assembly is of leaf or "iris" type design. As
shown in Figure 13, the aft portion, which is leafed, rotates about hinges and it
changes the effective throat area of the exhaust nozzle between 18 and 71 percent
of the combustor flow area.
The outer combustor and nozzle assembly is simply skinned
material primarily carrying bursting loads. The cooling liners shown duct fuel
free inlet air aft to the nozzle entrance on the outside shell as well as the
center body. These liners are louvered in such a manner that some of the air in-
side escapes and film cools the liner itself. The remainder exits at the liner
end and film cools the center plug and outer nozzle assemblies.
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REPORT 5808
The maximum steady state operating temperatures of the major
parts forming these subassemblies are shown in Figure 14. These temperatures were
determined for a representative trajectory wherein the engine accelerates from
Mach 2 to 4 at maximum power, climbs to cruise altitude, and operates for 1 to 3
hours at cruise power settings. Figure 8 summarized the performance of the
various materials at elevated temperatures. The material selections resulting on
the basis of these temperatures, loads, etc., are shown in Figure 15. Much use is
made of Rene' 41, which appears to be optimum for many of the parts considering
manufacturability as well as material performance. Adequate creep or "life" data
for the more attractive material are not yet available and ultimate analysis may
reveal one of the materials other than Rene' 41 more suitable.
Utilizing the estimated operating temperatures, material proper-
ties, load factors, etc., to select optimum materials and shapes, a resulting
engine weight of 880 rbs is estimated. This weight breakdown is shown in Table IV.
2. Prototype Engine
For early structural and aerothermodynamic development testing,
a prototype engine was fabricated which was of flight engine design wherever
possible. A photograph of this engine is shown in Figure 16. The engine was com-
plete except in two respects, namely it had no control package since long lead
times are required for designing and making numerous castings, and it had no
variable exit plug for the same reasons. Two plugs were fabricated simulating the
variable plug in the maximum power position and in the cruise power position. In
addition, N-155 alloy was substituted for other materials in some areas, again due
to long lead time requirements for the correct materials.
The engine was completely instrumented and ready for test at the
end of the contract work period.
IV. CONTROLS
The fuel and control system for the Mach 4 integral cruise engine was
designed to provide optimized control functions for the complete propulsion system
which included the variable geometry air induction system and the ramjet engine.
This section summarizes the concepts and design principles of the
over-all power control system.
The control system design presented in the subsequent discussions was
conceived to be fundamentally suitable to both piloted and nonpiloted vehicle ap-
plications. All control functions are completely automated and only the mode of
propulsion system operation as required for specific missions or trajectories is
selected by either manual or further automated means. The piloted application is
used as the primary reference in these discussions.
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MORT 5808
TABLE IV
ESTIMATED WEIGHT BREAKDOWN FOR THE
MACH 4 INTEGRAL CRUISE TYPE RAMJET
(Incorporating Blunt Plug Exit Nozzle)
Component
Sheet
Metal
Ring &
Mach. Parts
Castings
Purchased
Parts
--
Totals
Fwd. inner body
16.10
20.10
36.20
Aft inner body
28.50
22.40
50.90
Fwd. inner body liner
7.75
7.75
Aft inner body liner
24.65
24.65
Longeron assembly
42.50
6.00
48.5o
Fuel delivery
11.70
11.80
23.50
Burner
93.4
5.9
99.30
Fwd. liner outer
22.00
22.00
Aft liner outer
45.5
45.50
Diffuser assembly
37.3
59.40
96.70
Tailpipe
89.5
4.o
93.50
Exit plug
38.00
27.00
44.00
10.0
119.00
Miscellaneous fasteners
9.0
9.00
Miscellaneous
15.0
15.00
Package including
actuators & ignition
system
188.50
Total 880.00 rbs
CO
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REPOU 5808
The design of an engine bleed air turbine powered unit for hydraulic
and electrical accessory power is also presented.
A. System Functions
Figure 17 indicates the type of propulsion system (with major input
and output variables affecting control design) for which the subject control
system is designed.
The general requirements performed by the power control system can
be summarized as follows:
1. Induction System Controls
a. Position inlet geometry and engine bypass duct for minimum
drag during engine nonoperating phases.
b. Position inlet geometry to provide optimum pressure recovery
and capture area at low supersonic speeds.
c. Position inlet geometry in such a manner that external com-
pression shock waves are held in stable locations and so
that maximum pressure recovery is available to the engine.
d. Provide starting capabilities of the propulsion system any-
where within the flight envelope and restarting capabilities
in the event of diffuser shock expulsion.
e. Regulate engine bypass duct flow to provide proper matching
of inlet and engine air mass flow characteristics.
2. Engine Controls
a. Regulate ignition and reignition fuel flows.
b. Control desired modes of fuel distribution to the combustor.
c. Limit exhaust gas temperatures during accelerating thrust
conditions at all Mach numbers.
d. Limit exhaust gas temperatures during emergency thrust
operation.
e. Optimize acceleration and cruise specific fuel consumptions
through control of pressure recovery, fuel flows, and exit
nozzle size.
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REPORT 5808
Further specific required functions and design considerations for the
power control system are reviewed separately.
Precision operation was specifically designed for the flight ranges
of Mach 2 to 4 at 30,000 to 100,000 feet altitude. The degree of precision out-
side of this nominal envelope was not specifically examined.
B. System Concept
The inlet controls and engine controls are, by function, conveniently
separated into two subsystems. The basic function of the inlet and controls is to
provide the optimized inlet capture area and ram pressure recovery potential. The
engine controls (in this case regulation of engine bypass duct air, heat addition,
and exit nozzle area) then are charged with maintaining maximum potential pressure
recovery while delivering acceleration and cruise thrusts at minimum specific fuel
consumption.
Even though the inlet and engine controls are not integrated through
common loops, they of course must act synergistically during operations such as
ignition, possible diffuser shock expulsion, etc. Therefore, the inlet control
system was designed to function independently and to compliment engine controls
in cases where normal propulsion system operation need be established or re-
established.
The engine control system involves the regulation of three variables:
bypass air, fuel flow, and exit nozzle area. The basic criteria require a system
arrangement which assures maximum thrust potential for acceleration at low or ram-
jet takeover Mach numbers and accurate optimization of specific fuel consumption
during the high Mach number cruise operation. Consideration of the sensitivity of
engine characteristics to possible controlled variables (See Figure 18) and
physical limitations of both engine and control determined the arrangement of con-
trol functions and loops as shown in Figure 19 in order to best satisfy per-
formance accuracy requirements.
The system concept reflected by the control system design (Figures
19 and 20) provides for closed loop control of functions such as exit nozzle area,
bypass air flow, and inlet geometry, whereas engine fuel flow is controlled by an
open loop system. This type of arrangement is indicated by relative sensitivity
of engine performance to the controllable variables and also by the difficulty in
determining effective areas of variable geometry engine components such as the
exit nozzle under conditions of thermal expansion, thermal creep, exhaust gas
leakages, change coefficients, etc. Closed loop controllers automatically com-
pensate for these variations.
Conversely, the open loop function of the engine control system
(fuel-air ratio regulation) can be independent of engine and induction system per-
formance deviations. Therefore, the fuel flow control loop can conveniently ac-
cept external commands for selecting thrust levels and start up and shut down se-
quencing.
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1. Inlet Geometry Control System
REPORT 5808
By selection of proper pneumatic pressure parameters available
from external and internal compression fields of the induction system, a closed
loop control system which maintains a fixed pressure ratio was made possible in-
stead of a more complicated open loop scheduling system which would schedule inlet
position with Mach number. The suitability of the parameter (a fixed ratio of
external diffuser pressure to throat pressure) is shown in Figure 21 wherein the
command geometry is positioned in such a manner that ideal performance is closely
matched over the design Mach number range.
The inlet geometry control system is shown in Figure 22. The
system consists of hydraulic (3000 psi) and pneumatic (ram pressure) components.
The oil--hydraulic power source was used because of the significant heat transfer
problems under the 1200�F environment involved in supplying power from the re-
motely located engine. Hydraulic power from the air turbine motor unit would be
used.
The control and actuator system consists of a proportional plus
integral pneumatic control unit which senses the signal pressure ratio. The out-
put pneumatic signal is received by a pneumatic signal booster unit which provides
a position output and drives the actuator hydraulic servo through mechanical
linkages. The resultant actuator motion and position is fed back mechanically to
the signal booster-servo valve linkage, thus making it a proportional element.
This arrangement avoids a double integrating system (controller plus servo valve
actuator) while still maintaining the zero steady state error characteristic of
the proportional plus integral system.
Full extended or retracted actuator positions (minimum or maxi-
mum induction inlet areas) can be commanded by means of separate bias to the
signal booster unit of the system.
2. Engine Air Bypass Control System
The engine air bypass system functions only to match diffuser
and engine air flaw characteristics (See Figure 31). Except for low Mach number
operation wherein bypass may be required even though the engine were controlled
to consumed maximum possible air flow, the need for bypassing air is dependent
upon the type and accuracy of control of the engine operating variables. There-
fore, control of bypass air is integrated with the engine control system. This
control means is discussed within the engine control section of this report.
However, due to the remote location of the bypass system from
the engine proper, the bypass controls are not physically integrated with the
engine control system and they also receive actuator power from a vehicle hy-
draulic power source (air turbine motor unit) rather than from engine accessory
units.
The bypass control system, as shown in Figure 23 is composed of
components of the same design as those for the inlet geometry control system ex-
cept for variation in sizing, gains, and calibration characteristics.
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REPORT 5808
3. Engine Fuel Control System
The fuel flow control system (Figure 24) is designed to operate
with several distinct modes of control. A simplified representation indicating
the various loops and modes of control is shown in Figures 25 and 26. Figure 25
represents a manual input arrangement for selecting the desired mode of operation
whereas Figure 26 shows the resulting control performance with input variations.
The fuel delivery characteristics of Figure 26 are fixed irrespective of flight'
Mach number, altitude, or day temperature.
Briefly, in reference to Figures 25 and 26, the control system
modulates fuel-air ratio directly during ignition procedures and minimum thrust
demand conditions in order to match engine combustor characteristics at lean
burning operation. Intermediate and maximum power operation (and emergency power)
are governed by the control system so that combustion chamber temperature is
maintained at the pr6kribed calibrations irrespective of ram air temperature (al-
titude, Mach number, and ambient air temperature). The third control loop is re-
quired during long cruise durations wherein one of the two fuel injector rings is
made inoperative (for added combustor structural life) and a high gain (thrust
versus speed) characteristic is provided by the control for convenient speed
regulation. This high gain thrust control again controls combustion temperatures
on an open loop basis, but, due to the narrow band, high gain characteristics,
maximum temperature limits are maintained by a fuel-air ratio override loop. Ac-
curacies determined for the engine control system are described in Figure 27.
The fuel controllers are of pneumatic, hydraulic, and mechanical
design which are packaged and housed within the ramjet engine center body (See
Figure 33). The controls require no external power source for operation because
they utilize ram air and fuel as working fluids for computator and actuator power.
(The exceptions which use external power sources are intermittent electrical
power requirements for combustor spark ignition, and in the case of manual selec-
tion of operational modes, inputs through a mechanical shaft are required.)
For convenience, the fuel flow and control system can be dis-
cussed in terms of four subsystems. These indlude: the pneumatic computing
system, the fuel metering and injector system, the power mode selector system,
and the fuel pumping system.
The heart of the pneumatic computing system is the engine air
mass flow computer. It operates by sampling engine air flow at the engine inlet
(downstream from the diffuser bleeds and engine air bypass ducting). A fixed
percentage of engine air is captured by the sampling probes. A pressure signal
which is proportional to sampled air (and therefore engine air flow) is obtained
through manipulation of the sampled air, by fuel-to-air heat exchangers, prior
to exhausting through calibrated choked nozzles. Experimental performance of the
air mass flow computer component is presented in Figure 28.
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won 5808
The intelligence provided by the engine air mass flow computer
provides a fundamental reference from which all fuel flow functions may be re-
lated to engine operation. Accurate fuel-air ratio regulation to the primary
injectors for ignition, combustor zone stabilization, minimum fuel-air ratio
limits, and to both primary and secondary injectors for maximum fuel air limits
are readily achieved.
Fuel-air ratios are varied by automatic and manual means to
attain acceleration maximum thrust schedules and to select desired thrust levels
for cruise, deceleration, and emergency power conditions by processes which bias
the basic air mass flow signal pressure. The signal is, in general, modulated
by a series of pneumatic pressure divider units, each of which delivers a separate
output pressure which is a function of the air mass flow computer signal and the
manual or automatic demand input to the variable pressure dividers.
The automatic inputs, with reference to the system schematic of
Figure 24, are governed by stagnation air temperature sensers. One temperature
senser operates a pressure divider unit so that the output signal varies engine
acceleration fuel-air ratios so as to maintain a constant combustion chamber
temperature for all Mach numbers and elevated thrust demands as illustrated in
Figure 26. A second temperature senser, used only during cruise speed operation,
biases the control signal in order to vary engine thrust inversely with vehicle
speed and therefore provide vehicle speed stabilization (See Figure 29).
The mode selector system consists of a complex of switching
valves which port the desired pneumatic signals to the fuel metering valves and
also includes variable pneumatic pressure divider units which receive commands
for adjusting fuel-air ratio and intermediate engine thrust levels. All pneu-
matic switches and pressure dividers are synchronized to operate from a single
rotary input shaft. An interlocked push-pull mechanical input is provided for
selecting the cruise operating mode of control operation.
Engine bleed air powers the turbine driven centrifugal fuel pump
which raises fuel pressures from tank pressure at the engine inlet to that re-
quired to operate the fuel controls, actuators, and fuel injection system. The
turbine air power is controlled by throttling the bleed air upstream from the
stators so that the pump head rise does not exceed the requirements of the
system.
Control of the pump serves three additional Objectives. First,
it reduces system pressures to a minimum, which allows use of lightweight mag-
nesium fuel component castings under the extreme temperature environments (up to
500�F fuel temperatures). Second, the throttling of turbine supply air in this
manner reduces engine bleed air at cruise speeds and gives an incremental improve-
ment in specific fuel consumption. It also makes practical the adoption of a
small high pressure pump run directly by the turbine pump shaft, since the pump
control limits maximum shaft speeds to slightly over 20,000 rpm. This small, high
pressure pump supplies hydraulic power to the engine exit nozzle actuator system.
The compatability of turbopump characteristics with systems requirements is
demonstrated in Figure 30.
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REPORT 5808
A portion (approximately 25 percent) of the engine pump fuel
output is recirculated to an ejector at the pump inlet which significantly in-
creases the suction specific speed to allow minimum fuel tank pressurization.
The two fuel metering valves are of nearly identical design and
differ only in internal port sizing as required capacities differ slightly. As
previously indicated, they are arranged in parallel and they independently meter
fuel to the primary and secondary injector systems. Simple orifice type fuel
nozzles are installed in both primary and secondary injectors. However, the
primary injector, because of its larger flow range requirements, incorporates
pressure sensitive switching valves between two sets of nozzles so that maximum
system pressures are minimized. The more mechanically complicated, spring loaded,
variable area type fuel nozzles were not deemed practical since, under certain
engine operating conditions, the injectors encounter 1200�F environments without
benefit of fuel flow cooling.
The fuel metering valves are the flow regulating type (volu-
metric) which deliver a scheduled fuel flow characteristic in accordance with
the input pneumatic differential signal. A constant fuel pressure drop is main-
tained across the variable area metering orifice by a servo controlled throttling
valve. The metering orifice area is governed by a positioning servo loop which
is in turn positioned by a spring loaded diaphragm which receive the pneumatic
demand signal. The control is fuel temperature compensated. Therefore, the unit
regulates the fuel weight rate flow for any specific fuel.
4. Variable Exit Nozzle Control System
The exit nozzle throat area is controlled to maintain approxi-
mately 97 percent of diffuser critical pressure recovery within limitations of
full nozzle area excursion. As previously indicated, the engine air bypass system
and the exit nozzle system (Figures 23 and 24) are complimentary toward main-
taining critical diffuser pressure recovery under certain conditions. During low
Mach number operation, at intermediate and high power levels, engine bypass con-
trol is required even with full open nozzle as shown in Figure 31. Second, the
high response critical control circuit is placed in the bypass system in order to
minimize actuator size and power to the more massive and higher loaded exit noz-
zle. Both the exit nozzle and bypass control systems operate from the same dif-
fuser probe pressures which describe critical pressure recoveries. However, the
two systems are calibrated with an incremental difference so that the bypass
system is not activated except during conditions wherein the exit nozzle is in-
capable of maintaining critical recoveries. By these means, engine thrust and
specific fuel consumption are optimized except during brief transient periods
(Note Figure 31).
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The exit nozzle control is also relied upon to set engine aero-
dynamic flow conditions within specified limits required for ignition at all alti-
tudes and Mach numbers. These burner conditions are satisfied by adjusting the
exit nozzle in such a manner that approximately 65 percent of critical pressure
recovery is maintained under nonburning operation. This setting also assures
supercritical diffuser operation during the transition from cold flow to burning
operation. The exit nozzle area control for ignition is regulated through the
normal control system which is biased to maintain the lower pressure recovery
setting. Nominal exit nozzle positioning for maximum power and ignition scheduline
is shown in Figure 32.
The exit nozzle controller is an integrating type control with
velocity feedback. The controller unit consists of diaphragm motors which receive
the pressure recovery pneumatic signals and the velocity feedback signal. The
integrating diaphragm reuses the diffuser demand signals, one of which passes
through a restrictor thus providing the integral characteristic. The controller
is stabilized by the second diaphragm which receives an exit nozzle position
signal from a pneumatic variable pressure divider and ports the signal across the
diaphragm through a restrictor to provide an opposing force during transients.
The diaphragm motor system actuates the two-stage hydraulic servo valve to govern
exit nozzle actuator motion.
Fuel is used as the hydraulic working fluid for the controller
and piston type nozzle actuator. A high pressure hydraulic source (1500 psi) is
provided by the small (3 gpm) positive displacement pump which is directly driven
by the air-turbine-driven main fuel pump shaft.
D. Environmental Considerations
All fuel system and exit nozzle control system components are inte-
grated into one package assembly which is installed within the engine center body.
The system layout (Figure 33) illustrates the installation of the system. The
entire assembly is fuel cooled by the metered fuel flows and by the turbopump
bypass flow to the pump supply ejector. In addition, molded thermal insulation
is applied to external surfaces.
These cooling techniques make possible the use of magnesium castings
for most control housings under conditions of 1200�F ambient environment when
supplied with fuel at 500�F. Components such as the turbo pump inlet ducting and
the exit nozzle actuator, which are subject to convection and direct radiation
from combustion chamber and exit nozzle walls, are fabricated from high tempera-
ture steels. Special Laminated high temperature spacers are used at the package
tie-down points to minimize heat conduction.
Addition of heat to internal control parts through flowing ram air
signal lines is avoided since all air is cooled by the fuel-to-air heat exchanger
which also provides the computing function in the air mass flow computer circuit.
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REPORT 58o8
Conversely, heat conduction effects from the package body to the
bimetallic temperature sensers were minimized by locating the sensers at the ex-
treme forward end of the package, ahead of the steel fuel injector manifolding in
the package assembly, remote from the magnesium fuel cooled control sections.
The suitability of this approach toward achieving environmental re-
sistance was exhibited experimentally by subjecting pneumatic computers (with
diaphragm motors) and fuel flow regulators, in a packaged assembly, to the maxi-
mum environmental temperatures and heat transfer rates. The fuel metering valves
and cooling passages of the magnesium castings were supplied with fuel at near
maximum temperatures. Control performance and structural integrity were satis-
factory after steady state temperature distribution was achieved and maintained as
shown in Figure 34. A photograph of the environmental test stand and the engine
model (control test cell) is shown in Figure 35.
Nonmetallic elements such as diaphragm motors and seals were further
evaluated experimentally in order to select the most reliable materials and
fabricating techniques for the required temperature operating range. Figure 36
describes the environmental life of the selected diaphragm material, which was
DuPont Fairprene elastomer on glass fabric. The manufacturing process was noted
to be the most significant factor in achieving satisfactory performance and life
at high temperatures for a given combination of materials.
E. Installation and Ground Check Features
The complete engine control, pump, and nozzle actuator assembly is
installed and removed through the center plug of the variable area exit nozzle.
The package tie-down point is located at the forward end of the exit nozzle where
steel supports, cast integrally with the exit nozzle actuator (See Figure 33) are
bolted to an engine structural ring. Forward package shear support is provided
between the forward steel fuel manifolding casting of the package assembly and
the forward engine inner body structural ring.
The leading edge of the inner body aerodynamic shape is incorporated
into the package design in order that fuel lines to the injector nozzles could be
attached to the package through slip joint seals. Thus, package installation and
removal is facilitated and connections remain sound under environmental tempera-
tures where axial differential expansion between the package and engine inner body
occurs.
All electrical lines, fuel supply lines, manual control input shafts,
external pneumatic signal lines and ground check lines are carried from the engine
attach pad through a single engine longeron to the inner body and control package.
Consequently, the control system installation is performed by attaching lines at
two points (fuel injector lines at the engine face and all other connections at
the engine attach pad) and bolting the assembly in place through the exit nozzle
at the aft engine structural ring.
CONFIDENTIAL
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REPORT
5808
Provisions are made to ground check the operation of the fuel con-
trol and pumping system and the exit nozzle actuator system while the engine is
installed on a vehicle. All necessary connections and lines (additional to
normal flight connections) including auxiliary fuel control discharge ports,
pneumatic signal inputs, a turbo pump ground check air supply line, and a pneu-
matic control circuit vacuum line are provided at the engine attach pad. A
quantitative check of the package performance may be conveniently conducted by
use of these provisions for ground supplied hydraulic and pneumatic services.
F. Air Turbine Motor Accessory Drive
Accessory hydraulic and electrical power is provided through use of
an air turbine motor drive unit as the prime mover. Propulsion system diffuser
bleed air (See Figure 17) is ducted to the turbine motor unit. The unit is de-
signed to deliver full power requirements for the propulsion system and a vehicle.
Design horsepower outputs are a maximum of 74 horsepower for continuous operation
and 29 horsepower during average conditions.
The unit (and air ducting) consists of an upstream air inlet
throttling valve (and associated speed and overspeed controls), a single stage
turbine, hydraulic lubricating and scavenging pumps, hydraulic recirculating
pump for the alternator cooling system, a gear box and the two output power pads
for the alternator and hydraulic pump. The air turbine motor system is shown
schematically in Figure 37.
1. Inlet Power Control Valves
The inlet air valving is basically the turbine inlet duct. The
duct contains two valves capable of throttling bleed air flow to the turbine.
The forward valve is an on-off valve used for normal start and shut down functions
as commanded by a signal to the electrical actuators. In addition, it receives ar
electrical signal from the overspeed governor to command emergency shutdown.
The aft valve is positioned by a hydraulic actuator to regulate
air bleed power to maintain constant turbine speed under transient and steady
state conditions of accessory power demand and available bleed air pressure
ratios and mass flows from the engine.
2. Turbine Assembly
The turbine is an axial flow, single stage, reaction type unit.
The turbine operates with 100 percent admission with the speed controlled (32,000
rpm) by the inlet duct valve which varies the available air horsepower. Materials
for the disk and blade are forged Rene' 41 and investment cast 713C alloy, re-
spectively. Choice of these materials made design possible for safe operation at
130 percent overspeed at maximum temperature (1270�F) and yet control possible
failure within a narrow band of overspeed values at all operating temperatures.
Since failure would occur at the blade roots, all parts can be contained within
the exhaust duct section in the event of failure. The turbine shaft is mounted
In a ball and roller combination.
CONFIDENTIAL
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CORPORATION
VAN NUYS, CALIFORNIA
REPORT 5808
3. Speed Controls
Design speed is maintained (to an accuracy of +1 percent) with a
fly ball type hydraulic governor driven by the turbine. The speed governor posi-
tions the hydraulic actuator on the inlet air valve to control the turbine input
ram air power.
A mechanical fly ball governor is mounted on the turbine shaft
which actuates the shutoff air valve through a mechanical linkage in the event
of overspeed. The spring loaded governor initiates valve closing at 110 percent
overspeed and the valve is full closed at 120 percent overspeed.
4. Gear Box
Power from the unit is provided at the output pads for the
alternator and hydraulic pump mounting, each turning at 8000 rpm. Speed reduc-
tion to the output pads is accomplished from a single reduction spur gearing
with the turbine shaft pinion and the output shaft drive gear. A secondary ac-
cessory pinion on the turbine shaft mates with a gear which mounts on the shaft
which drives the speed control governor at 6000 rpm. The governor drive shaft
has a pinion that mates with two additional gears for driving the two lubricator
oil cooling pump arrays at 6000 rpm.
5. Lubricating and Cooling Systems
The lubricating and cooling systems (Shown in Figure 37) consist
of the turbine and gear box lubricating system, a generator cooling system, and
an oil cooling system. All system components, such as supply and scavenge pumps,
filters, oil sumps and relief valves, are integral with the air turbine motor de-
sign as indicated in Figure 38. The aft half of the gear case includes pads to
mount pumps, governors filter, etc. Oil lines are based in it to eliminate ex-
ternal plumbing, and the hot and cold sumps are cast on the front of the turbine
housing.
The lubrication and cooling system for the turbine shaft and
bearings is designed to provide for operation under the severe thermal environ-
mental conditions of 300�F ambient temperature, -65�F to 300�F oil supply tempera-
ture, and 1270�F turbine air supply temperature. Lubrication and cooling is
accomplished by pumping the oil (Specification MIL-L-7808C) through a nozzle
jetting in to the end of the turbine shaft. Centrifugal force aids in discharging
the oil to three different locations. The first two are small lubricating jets
discharging horizontally to the bearings. The third is for cooling purposes. The
oil flow (approximately 2 gpm) absorbs heat travelling through the shaft toward
the front bearing. It is then discharged from the shaft forward of the front
bearing through holes drilled in a high thermal conductivity copper disk holding
carbon nozzle seals. The oil flow returns through the turbine housing around the
outside of the bearing, thus cooling the bearing internally and externally.
CONFIDENTIAL
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CONFIDENTIAL
G. Development Status
arquard
CORMRATION
VAN NUTS, CALIFORNIA
WORT
9E1�8
Completion status of the inlet control system, engine fuel and con-
trol system, and the air turbine motor unit at the termination of this study is
tabulated below.
Phase
Inlet
Control
System
Engine
Fuel &
Control
System
Air
Turbine
Motor
Unit
a.
System concept and design approach
95%
98%
90%
b.
System and component detail design
analysis
90%
95%
80%
c.
System and component detail design
and release
50%
50%
4o%
d.
Heat transfer analysis
50%
50%
60%
e.
Materials and stress analysis
80%
85%
85%
f.
Component and element testing
5%
8%
0%
g�
Systems testing
0%
3%
9%
h.
Manufacturing investigation and
tool engineering
98%
98%
90%
V. CONCLUSIONS
As a result of the six-months study of the Mach 4 integral cruise
engine, it has been concluded that
1. A combustion system can be developed which can be spark ignited
and which will give combustion efficiencies up to 90 percent during near stoichio-
metric operation during climb and acceleration and 95 percent during Mach 4 cruis-
ing at lean fuel-air ratios at altitudes on the order of 90,000 feet.
2. A lightweight ramjet engine structure, made largely of Rene' 41
alloy, can be fabricated and should withstand the environments imposed during long
periods of cruising operation at Mach 4 (incorporating an overlapping leaf,
variable plug exit nozzle).
3. An engine fuel pumping and power control system can be built
largely from modified XRJ43-MA-9 (Bomarc B) components which will provide neces-
sary fuel pressurization and power control during long periods of cruising opera-
tion.
4. Stated more generally, it is concluded that the Mach 4 integral
cruise ramjet engine state of the art has been sufficiently well established to
be used as a foundation for immediate development of flight equipment.
CONF IDENT IAL
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711a rquadi
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REPORT 5808
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Approved for Release: 2017/02/01 C05811775
FIGURE 15
Approved for Release: 2017/02/01 C05811775
THE
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FIGURE 16 - Prototype of Flight Engine
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Approved for Release: 2017/02/01 C05811775
Approved for Release: 2017/02/01 C05811775
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FIGURE 17
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THRUST AND SPECIFIC FUEL CONSUMPTION
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Approved for Release: 2017/02/01 C05811775
FIGURE 18
Approved for Release: 2017/02/01 C05811775
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Approved for Release: 2017/02/01 C05811775
FIGURE 19
Approved for Release: 2017/02/01 C05811775
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