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39* AIAA/ASME/SAE/ASEE Joint Propulsion Conference 
Huntsville, Alabama 20 - 23 July 2003 



NON-TOXIC DUAL THRUST 

REACTION CONTROL ENGINE DEVELOPMENT 

FOR ON-ORBIT APS APPLICATIONS 

Philip J. Robinson* and Eric M. Veith^ 

Aerojet Propulsion Division 

Sacramento, California 



ABSTRACT 

A non-toxic dual thrust proof-of-concept 
demonstration engine was successfully tested at the 
Aerojet Sacramento facility under a technology contract 
sponsored by the National Aeronautics and Space 
Administration's (NASA) Marshall Space Flight Center 
(MSFC). The goals of the NASA MSFC contract 
(NAS8-01I09) were to develop and expand the 
technical maturity of a non-toxic, on-orbit auxiliary 
propulsion system (APS) thruster under the Next 
Generation Launch Technology (NGLT) program. The 
demonstration engine utilized the existing Kistler K-1 
870 Ibf LOX/Ethanol orbital maneuvering engine 
(OME) coupled with some special test equipment 
(STE) that enabled engine operation at 870 Ibf in the 
primary mode and 25 Ibf in the vernier mode. Ambient 
testing in primary mode varied mixture ratio (MR) from 
1.28 to 1.71 and chamber pressure (Pc) from 1 10 to 181 
psia, and evaluated electrical pulse widths (EPW) of 
0.080, 0.100 and 0.250 seconds. Altitude testing in 
vernier mode explored igniter and thruster pulsing 
characteristics, long duration steady state operation 
(>420 sec) and the impact of varying the percent fuel 
film cooling on vernier performance and chamber 
thermal response at low Pc (~4 psia). Data produced 
from the testing provided calibration of the 
performance and thermal models used in the design of 
the next version of the dual thrust Reaction Control 
Engine (RCE). 

INTRODUCTION 

The National Aeronautics and Space 
Administration's (NASA) George C. Marshall Space 
Flight Center (MSFC) has contracted with Aerojet to 
develop and expand the technologies necessary to 



support non-toxic, on-orbit auxiliary propulsion system 
(APS) needs for the Next Generation Launch Vehicle 
Technology (NGLT) program. Contract NAS8-01109 
has been issued to Aerojet to develop a dual thrust 
Reaction Control Engine (RCE) that utilizes oxygen 
and ethanol as the propellants. For the purposes of this 
contract, thrust levels of 870 Ibf and 25 Ibf have been 
selected for the primary and vernier mode, respectively. 

The Aerojet contract is divided into three 
separately funded phases: Basic, Option 1 and Option 
2. The completed Basic phase tested the existing 
Kistler K-1 Orbital Maneuvering Engine (OME) with 
contract designed STE hardware in vernier and primary 
modes, the results of which are discussed herein. 
Option 1, which is currently in work, uses the resulting 
test data from the Basic program to design, fabricate 
and test an integrated dual thrust RCE demonstration 
engine. Option 2, which is a future effort, utilizes the 
Option 1 test data to fine tune the dual thrust RCE 
design, leading to a Critical Design Review (CDR) and 
to the final fabrication of three deliverable prototype 
thruster assemblies for the NASA White Sands Test 
Facility (WSTF). These three thrusters are to be 
installed in an on-orbit APS simulator and to undergo 
system level hot-fire testing in GFY 2005. 

The feasibility of the dual thrust RCE was 
demonstrated by testing performed at NASA WSTF in 
New Mexico.' A 600-lbf GOX/Ethanol Reaction 
Control System (RCS) workhorse thruster developed by 
Aerojet in the I980's for NASA Johnson Spaceflight 
Center (JSC) was utilized for the WSTF testing. The 
workhorse RCS thruster incorporated a center-mounted 
spark-initiated torch igniter. Some modifications were 
made to the RCS thruster hardware to permit igniter 
operation in conjunction with reduced auxiliary GOX 



Senior Member A.LA.A., Engineering Specialist 
* Senior Member A.I.A.A., Program Manager 



1 
American Institute of Aeronautics and Astronautics 



flow through the injector face, and with reduced ethanol 
boundary layer cooling (BLC). These hardware 
modifications permitted simulation of the function of 
the current Space Shuttle 25 Ibf vernier reaction control 
subsystem (VRCS) engine for short test durations (-30 
sec). 

Hot-fire testing of the existing Kistler K-1 
OME hardware has been performed during the Basic 
phase of the contract to continue in the path of the 
NASA JSC sponsored WSTF testing. The Kistler K-1 
OME is an 870-lbf engine that utilizes LOX and 
ethanol as its propellants, and incorporates a 
GOX/Ethanol center-mounted spark-initiated torch 
igniter for igniting these propellants. Although the 
Kistler OME has undergone some previous testing in 
steady state mode, as well as limited testing in pulse 
mode,^ additional testing of the Kistler OME hardware 
during the Basic phase of the contract has extended the 
test database. This additional testing has achieved 
significantly longer vernier steady state durations and 
some vernier pulse mode operation. The primary mode 
testing evaluated a variety of mixture ratio and chamber 
pressure conditions, as well as pulse mode conditions. 
The completed testing has provided the data necessary 
to support the subsequent design effort in Option 1 for 
the dual thrust LOX/Ethanol RCE. 

TEST OBJECTIVES 

There were multiple test objectives associated 
with the planned testing of the Kistler OME hardware 
at 25-lbf thrust (vernier) and 870-lbf thrust (primary) 
conditions. One main objective of the vernier testing 
was to evaluate the Kistler torch igniter with and 
without augmented propellant flow at the vernier 
operating conditions. The igniter produced 5 to 6 Ibf of 
thrust by itself, so it was necessary to provide auxiliary 
fuel and oxidizer to achieve the 25-lbf thrust level for 
the vernier mode. Initial tests evaluated igniter only 
operation, while subsequent tests determined the effects 
of coupling augmented propellant flow with the igniter 
flow to achieve the desired 25-lbf thrust. Two 
additional objectives of the vernier testing were to 
determine the effect of varying the percentage fuel film 
cooling (FFC) and the effect of two different chamber 
L' lengths (5.75 and 9.00 inches) on engine 
performance and thermal response. Finally, vernier 
pulse mode response and repeatability characteristics 
were to be evaluated for the pulse train definitions of 
Table I. 

There were several objectives associated with 
the primary testing of the Kistler OME hardware. One 



of the objectives was to determine engine performance 
and thermal response for varying mixture ratio and 
chamber pressure conditions utilizing an injector 
developed by Kistler. In addition, engine performance 
and thermal response were to be determined for two 
different chamber L' lengths (4.50 and 8.50 inches). 
Finally, primary pulse mode response and repeatability 
characteristics were to be evaluated for the pulse train 
definitions of Table I. 

The hardware thermal response to be evaluated 
in vernier and primary modes consisted of the thrust 
chamber heat load and the injector head-end heating for 
the conditions tested. The chamber heat load 
characteristics were determined from thermocouple 
temperature data measured at the chamber backside 
wall and the injector head-end. The temperature data 
established heat transfer coefficients and recovery 
temperatures to correlate the Aerojet SCALE thermal 
prediction code. This correlation of the SCALE code 
will enable predictions to be made for different thrust 
levels and chamber configurations, as defined by 
system architectures evaluated for the NGLT program. 
These thermal predictions will provide temperature 
profiles for use in design related activities such as 
fatigue life analyses, as well as in assessing engine 
head-end heat rejection to the propellant valves under 
steady state, pulsing and heat soakback (after 
shutdown) conditions. 

Table I. Pulse Train Definitions 



Pulse 


Characteristics 


Vernier 


Primary 


A 


On Time, EPW - sec 


0.250 


0.250 


Coast Time, ECW - sec 


LOO 


LOO 


Duty Cycle - % 


20.0 


20.0 


Pulse Quantity 


10 


10 


B 


On Time, EPW - sec 


0.250 


0.100 


Coast Time, ECW - sec 


0.500 


0.250 


Duty Cycle - % 


33.3 


28.6 


Pulse Quantity 


30 


30 


C 


On Time, EPW - sec 


0.250 


0.080 


Coast Time, ECW - sec 


0.250 


0.100 


Duty Cycle - % 


50.0 


44.4 


Pulse Quantity 


50 


50 



TEST HARDWARE DESCRIPTION 

The test hardware consisted of the Kistler K-1 
OME hardware, and some additional hardware 
consisting of an injector simulator, a special test 
equipment (STB) FFC ring, a FFC ring adapter, and a 
4-inch L' spool section. The Kistler OME hardware of 



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Figure 1 was comprised of an igniter assembly, igniter 



IH^Silicide OMted C-103 
ColumbfutuThriKt Chamber 



Spark Plus . 




Igniter 

Oxidizer Valve sectkm o~a 

valves, a spark exciter, an injector assembly, injector 
valves and a thrust chamber. 

Figure 1, Kistler K-1 OME Hardware 

The Kistler igniter was a GOX/Ethanol spark- 
initiated torch type, the design of which was well 
characterized by extensive testing on a variety of 
programs^'"*'' for over 30 years, and specifically for 
GOX/Ethanol on NASA Contract NAS9- 16639 in the 
1980's.* The igniter valves were direct-acting solenoid- 
type. 

The Kistler injector assembly consisted of 
individually photo-chemically machined nickel platelets 
that were diffusion bonded to a machined Inconel 
injector body. The injector body provided for direct 
mounting of the flanged thrust chamber valves and the 
igniter assembly, as well as provided for 
instrumentation ports for high frequency (Kistler) and 
standard (Tabor) pressure transducers. The injector was 
developed during the Kistler K-1 program, as were the 
thrust chamber valves. 

The Kistler thrust chamber was machined from 
a forged billet of C-103 columbium, and was di-silicide 
coated to provide oxidation protection. When attached 
to the Kistler injector, the resulting chamber L' was 4.5 
inches, which was the shortest L' tested. The 15-degree 
half-angle conical nozzle was truncated at an expansion 



area ratio of 3:1 to accommodate ambient testing and to 
facilitate extrapolation of performance (Isp) to higher 
area ratios. 

To test the Kistler OME hardware in vernier 
mode required some adjustments, as well as some 
additional test hardware. The Kistler injector was 
originally designed for the 870-lbf thrust level only, 
with a nominal LOX flow rate of 1.8 Ibm/sec, and was 
suitable for primary testing; however, the nominal 
augmented LOX flow rate to achieve 25 Ibf thrust in the 
vernier mode was 0.046 Ibm/sec. Therefore, the 
desired vernier LOX flow rate was only 2.6 percent of 
the original Kistler design LOX flow rate. This 
extreme difference between vernier and primary LOX 
flow rates would require virtually all of the oxidizer 
orifices to be blocked off to achieve the same margin of 
chug stability during operation in vernier mode. 
Blocking this many orifices defeated one of the 
intended purposes of having augmented LOX flow 
from the face of the injector, namely to provide injector 
face cooling during vernier operation. Therefore, 
Aerojet tested the Kistler injector using GOX to 
augment the igniter oxidizer flow in the vernier mode, 
thus permitting full flow through all of the oxidizer 
orifices. 

The Kistler injector was similarly oversized on 
the fuel circuit for vernier mode testing, having orifices 
far too large to accommodate the reduced fuel flow rate 
required to augment the igniter fuel flow. In lieu of 
using the existing fuel orifices on the Kistler injector, 
the STE FFC ring was designed to provide the reduced 
auxiliary fuel flow rate required for augmenting the 
igniter fuel flow during vernier mode operation. This 
FFC ring contained two separate fuel circuits, with each 
circuit having 18 fuel elements on the face. Half of the 
elements were directed along the chamber wall to 
perform fuel film cooling and were denoted FFCw for 
wall. The other 18 fiiel elements were directed inward 
toward the GOX flowing through the injector face, to 
enhance mixing and thus performance. These inward 
directed elements were denoted FFCc for core. The 
two FFC ring fuel circuits, FFCw and FFCc, were 
independently controlled to determine the sensitivity of 
the venire mode performance and thermal 
characteristics with respect to the flow splits on these 
two circuits. 

The STE FFC ring was designed to fit between 
the Kisder injector flange and the C-103 chamber 
flange, occupying the volume of the original Kistler 
acoustic cavity, as shown in Figure 2. Due to some 
dimensional constraints in the FFC ring design, an 



American Institute of Aeronautics and Astronautics 



2FFC Inlets, 180° Apart 
(FFCwALL and FFCcore) 



KIstler 
Injector 




Kistler 

Columbium 

Chamber 

Acoustic Cavity 



FFC 



LOX and Ethanol 
Elements 



KIstler 

Columbium 

Chamber 



KIstler 
Injector 



Igniter 



Original Kistler (870 Ibf) 
Configuration 




FFC Ring 
(CRES 304L) 

7ZZ7 



Igniter 



RCE Vernier (25 Ibf) Test Configuration 
with Kistler Hardware 



Figure 2. Special Test Equipment FFC Ring for Vernier Testing 



adapter ring was designed to provide a transition 
between the FFC ring and the Kistler C-103 columbium 
chamber. Some reduced size radial acoustic cavities 
were machined in the backside of the FFC ring to 
provide some damping capability, if it was required. 

The remaining additional test hardware 
consisted of the injector simulator and the 4-inch L' 
spool section. The injector simulator had no propellant 
circuits, and was used for the initial igniter-only testing 
in vernier mode. The 4-inch L' spool section enabled 
two different chamber L' lengths to be tested in each 
operating mode: 5.75 and 9.00 inches in vernier mode 
and 4.50 and 8.50 inches in primary mode. 

TEST SETUP 

The testing required two different test bays. 
The primary testing could be performed at ambient 
conditions with the Kistler OME hardware; however, 
the vernier testing required an altitude facility because 
of the very low vernier chamber pressure (Pc -4 psia). 
Therefore, the vernier testing was performed in the 
altitude facility of Test Cell A-2 at Aerojet, which was 
able to simulate approximately 175,000 feet at the 
beginning of a test, and to maintain a sustained 



equivalent altitude of 90,000 feet. The primary testing 
was performed at ambient conditions in the Aerojet 
Test Cell A-5. 

Vernier Test Hardware Configurations 

Three different test hardware configurations 
were used during the vernier testing. The first 
configuration consisted of the Kistler igniter assembly, 
the injector simulator, and the Kistler chamber, which 
was used to perform the initial igniter-only testing. The 
second configuration incorporated the Kistler igniter 
assembly, the Kisder injector assembly, the STE FFC 
ring and adapter, and the Kistier chamber. This 
configuration, shown in Figure 3 resulted in an 
effective chamber L' length of 5.75 inches, the shortest 
L' tested during the augmented flow tests in vernier 
mode. Finally, the long L' hardware configuration of 
Figure 4 substituted the 4-inch L' spool section for the 
FFC ring adapter, resulting in a chamber L' length of 
9.00 inches. A close-up view of the second test 
hardware configuration is shown in Figure 5, mounted 
within the test cabin of Test Cell A-2 at Aerojet. An 
overall view of the test cabin in the altitude test cell is 
shown Figure 6. 



American Institute of Aeronautics and Astronautics 




Figure 3. Vernier Test Conflguration - S.75-incli L' 




Figure 4. Vernier Test Configuration - 9.00-incIi L' 



Primary Test Hardware Configuration 

Two test hardware configurations were used 
during the primary testing. The first configuration was 
the basic Kistler K-1 OME hardware, which consisted 
of the igniter assembly, the spht triplet injector, the 
main thrust chamber valves, and the columbium 
chamber. This configuration, shown in Figure 7, had an 
effective chamber L' length of 4.50 inches. The second 
configuration added the 4-inch L' spool section 
between the injector and chamber, as shown in Figure 
8, resulting in a chamber L' length of 8.50 inches. 
Views of the first and second configurations installed 
on the test stand are shown in Figures 9 and 10, 
respectively. Primary testing was conducted at ambient 
conditions in Test Cell A-5 at Aerojet. 




Figure 5. Vernier Hardware Mounted in Test Cell 



-■m'i I'm 

:, 7, ', . •";-■■■ -U 

■■■'■ :,.:..- ---I 

■." i, '■/ ... ;■' 





Figure 6. Vernier Altitude Test Cell with Cabin 
Doors Removed 



American Institute of Aeronautics and Astronautics 




Figure 9. Primary Test Hardware with 4.50-inch L' 



Figure 7. Primary Test Configuration - 4.50-inch L' 




Figure 8. Primary Test Configuration - 8.50-inch L' 
TEST RESULTS 

Vernier Testing 

A total of 51 vernier tests were conducted. 
These hot fire tests included steady state and pulsing 
tests in igniter only mode, as well as in the augmented 
flow mode, for the L' length of 5.75 inches. Several 
steady state tests were completed for the L' length of 
9.00 inches for the purposes of comparing performance 
and thermal characteristics. A summary of the vernier 
tests is provided in Table II for the types of tests, test 
quantity, and test duration. A view of one of the 
vernier hot fire tests is captured in Figure 11. 




Figure 10. Primary Test Hardware with 8.50-inch L' 
Table II. Summary of Vernier Testing 



Test Description 


No. of 
Tests 


Duration, 

sec 


Igniter Only Tests 


wmm 


MMM 


Steady State 


9 


126.25 


Pulsing 


3 


32 


Subtotal: 


12 


158.25 


Augmented, 5.75-inch L' 




Steady State 


10 


340 


Long Steady State 


2 


689 


Pulsing 


7 


419.7 


Subtotal: 


19 


1,448.7 


Augmented, 9.00-inch L' 


WKKBKSUSKKMSBB' 


Steady State 


4 


195.5 


Subtotal: 


4 


195.5 


■■■■■■■■■■■■■HHVlHHWHHHI 


Total for Vernier Testing: 35 1,802.45 



American Institute of Aeronautics and Astronautics 




Figure 11. Vernier Hot Fire Test 

The nominal vernier operating conditions at 
the nominal vernier thrust level were a mixture ratio 
(MR) of 1.5 and a chamber pressure (Pc) of about 4 
psia. During the augmented flow tests of the vernier 
testing, the percent fuel film cooling (%FFC) was 
varied from 37.7 percent to 19.6 percent, with values 
between approximately 26 and 32 percent providing 
acceptable chamber thermal conditions at the smaller 
chamber L' of 5.75 inches. The total augmented fuel 
flow rate was held constant, so if the %FFC were 
increased, there would be a corresponding decrease in 
fuel core flow, and vice-a-versa. A long steady state 
test of 422 seconds allowed the thrust chamber to 
achieve thermal steady state, as evidenced by the 
relatively flat temperature plots of Figure 12. This long 
duration test was a major achievement of the vernier 
testing. 



Vfntf T—t 130- Long Durllow WM iil n i (32.1% FFCw) 
ChamlMr and bijador TwnpMvlura PiBili a 



2000 

.»» 

1000 
f- 1400 






H 


M 


i- 

1200 

1000 

, 800 






^■H 


m 




Sh^IHm 


MO 






^^i^m^H 


Bs^ 


400 






H^^^I^H 


MhU 


200 






^BK^KI^m 











nwnvMWfMM^ 





so 100 150 aOO 2S0 300 

T1nM.|.a«e 



400 450 



Figure 12. Chamber Temperatures for 422 sec 
Steady State Vernier Test 

Augmented pulsing tests consisted of 219 
pulses at an electrical pulse width (EPW) of 0.250 
seconds and 20 percent duty cycle, and of 85 pulses at 



an EPW of 0.250 and 33.3 percent duty cycle. Pulsing 
tests were discontinued after the main injector GOX 
valve failed to close properly on two pulses during the 
last pulse train attempted (Pulse Train B). Pulse trains 
were relatively short, i.e., 40 to 85 pulses, due to the 
difficulty in running the pulse tests with the STE FFC 
ring and its disproportionately large dribble volume. 

Primary Testing 

The nominal primary operating conditions at 
the nominal primary thrust level were a mixture ratio 
(MR) of 1.5 and a chamber pressure (Pc) of 150 psia. 
In the primary mode, a total of 31 hot-fire tests were 
attempted, with 21 yielding meaningful data. These 
hot-fire tests consisted of a series of steady state tests 
covering chamber pressures (Pc) from 110 to 181 psia 
and mixture ratios (MR) from 1.28 to 1.71 using the 
short chamber L' of 4.50 inches. The FFC was fixed at 
11.5 percent. A significant number of pulsing tests 
were also run with the short L'. A single steady state 
test was run for the long chamber L' of 8.50 inches for 
the purposes of comparing performance and thermal 
characteristics. A summary of the primary tests is 
provided in Table III, listing the types of tests, test 
quantity, and test duration. A view of one of the 
primary hot fire tests is pictured in Figure 13. 

A Pc / MR survey was performed with the 
short L' to determine the impact on chamber wall 
temperatures and engine performance. Figure 14 shows 
a sunmiary plot of the Pc and MR for each of the tests 
performed. The maximum chamber wall temperatures 
were between 1,900 and 1,950 °F, not varying 
significantiy with test conditions. By contrast, the 
maximum chamber wall temperature was about 2,280 
°F for the one steady state test performed at nominal 
conditions with the long chamber L' of 8.50 inches. 

Some long pulse trains were run in the primary 
mode with the short L'. The primary pulsing tests 
included the following: 325 pulses at an EPW of 0.250 
seconds and 20 percent duty cycle, 200 pulses at an 
EPW of O.ICX) seconds and a 28.6 percent duty cycle, 
and 806 pulses at an EPW of 0.080 seconds and a 44.4 
percent duty cycle. The pulse trains were of sufficient 
quantity/duration to achieve thermal steady state in the 
combustion chamber for all three EPW's tested. No 
appreciable difference (-50 °F) in the maximum wall 
temperature was detected between steady state tests and 
long duration pulsing tests where thermal equilibrium 
was achieved. The shortest EPW of 0.080 seconds had 
the highest chamber temperature for any of the short L' 
tests of approximately 2000 °F. 



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Table III. Summary of Primary Testing 



Test Description 


No. of 
Tests 


Duration, 

sec 


Tests with 4.50-inch L' 


MHHMHHHH 


Checkout 


3 


7 


Steady State (Pc & MR) 


9 


270 


Pulsing 


8 


666.2 


Subtotal: 


20 


943.2 


Tests with 8.50-inch L' 


ifeHHiiiHHiiiiiii 


Steady State 


1 


18.9 


Subtotal: 


1 


18.9 


flBilili^^ 


Total for Primary Testing: 21 962.1 




Figure 13. Primary Hot Fire Test 

Vacuum corrected specific impulse (Isp, vac) 
for the Kistler 3:1 conical nozzle was measured 
between 239.6 and 245.7 Ibf-sec/lbm for the short 
chamber L'. For the long chamber L', the Isp, vac was 
determined to be 249.6 Ibf-sec/lbm, slightly higher due 
to the increased chamber length. A summary plot of 
these performance values is given in Figure 15. 
Extrapolated primary performance for a chamber L' of 
4.50 inches, an 11. 5 percent FFC, and a 75 percent Bell 
nozzle with a 22:1 expansion area ratio was estimated 
to be 282 Ibf-sec/lbm. 



along the chamber wall. Specifically, the heat transfer 
coefficients and recovery temperatures were inferred as 
a function of chamber axial position from the vernier 
and the primary test temperature data. These heat 
transfer coefficients and recovery temperatures are 
shown in Figures 16 and 17 for the vernier tests, and in 
Figures 18 and 19 for the primary tests. There was 
significant data scatter for the vernier tests, but the 
curves for the heat transfer coefficient and recovery 
temperature were showing the correct trend. In 
contrast, the data for the primary tests was better 
behaved, and the curves for the heat transfer coefficient 
and recovery temperature have matched the data very 
well. 




1.40 1.50 1.60 

Ovwill MMim Ratio, O/F 



Figure 14. Summary Plot of Primary Mode Pc and 
MR Test Conditions 




MMuralMlo,MR.CVF 



Figure 15. Vacuum Corrected Is? for Primary 
Mode Tests 



POST TEST ANALYSIS 

Upon completion of the vernier and primary test series 
using the Kistler K-1 OME hardware, a post test 
analysis of the chamber thermocouple temperature data 
was performed to establish the gas-side conditions 



8 
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•t- 0.00025 



3000 




-4-3-2-1012 
Axial Distance from Throat, Inches 



Figure 16. Inferred Heat Transfer Coefficient for 

3000 




Figure 17. Inferred Recovery Temperature for 
Vernier Tests 




-4-3-2-10 1 2 
Axial Distance from Throat, Inches 



Figure 18. Inferred Heat Transfer Coefficient for 
Primary Tests 



-5 -4 -3-2-10 1 2 

Axial Distance from Tliroat, Inches 

Figure 19. Inferred Recovery Temperature for 
Primary Tests 



The inferred gas-side environments were used 
to correlate the Aerojet SCALE thermal prediction 
code. This correlation of the SCALE code compared 
very favorably with the primary mode temperature data 
for the various MR and Pc combinations tested. A 
comparison for the MR sensitivity in primary mode is 
given in Figure 20, where the solid color lines represent 
the SCALE predictions and the color markers are the 
actual test data. Similarly, a comparison for the Pc 
sensitivity in primary mode is shown in Figure 21, 
again where the solid color lines represent the SCALE 
predictions and the color markers are the actual test 
data. Both sets of SCALE predictions tracked very 
closely to the measured temperature data. The benefit 
of such a correlation will be the ability to perform 
thermal predictions for other thrust levels and chamber 
configurations, depending on system architecture 
requirements. This correlated SCALE code was used in 
the thermal design of the iteration of the dual thrust 
RCE, which is scheduled for primary and vernier 
testing in late summer 2003- 



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(Pc = 153-159psia) 




-4 -3-2-10 1 2 

Axial Distance from Throat, Inches 

Figure 20. SCALE Code Predicts MR Sensitivity in Primary Mode 
(MR = 1.47 -1.49) 




-3 -2 -10 1 

Axial Distance from Throat, Inches 

Figure 21. SCALE Code Predicts Pc Sensitivity in Primary Mode 



10 
American Institute of Aeronautics and Astronautics 



CONCLUSIONS 



REFERENCES 



Vernier Testing 

The following conclusions were drawn from 
the results of the vernier testing: 

• The vernier testing was successful in achieving 
long duration steady state firings with the modified 
Kistler OME hardware; 

• The vernier FFC can be reduced from values 
around 32 percent to perhaps 22 - 24 percent by 
improving the FFC injection velocities at the 
vernier design point, and by not having the STE 
hardware between the injector and the thrust 
chamber; 

• No chamber head-end heating was observed during 
limited pulse testing; 

• The low vernier chamber pressure influenced the 
definition of the pulse profiles during pulse testing. 

• Calibrated chamber thermal model correlates fairly 
well with the observed temperature data, enabling 
reliable predictions to be made for a design update 
to a fiilly dual thrust RCE. 

Primary Testing 

The following conclusions were drawn from 
the results of the primary testing: 

• Variations in Pc and MR had little effect on the 
chamber wall temperature profiles, as well as on 
the maximum wall temperature achieved; 

• Maximum chamber wall temperatures varied by 
+50 °F to -200 °F from the nominal steady state 
baseline test for the short L' configuration, over the 
entire range of Pc and MR tested, as well as for all 
three pulse trains (A, B, and C); 

• Estimated performance extrapolated for a 22:1 
nozzle was approximately 282 Ibf-sec/lbm; 

• There was no evidence of any chamber head-end 
heating during any of the pulse testing, as the 
ethanol appears to be a very benign fuel film 
coolant, i.e., no decomposition issues; 

• The chamber L' of 4.50 inches provided good 
performance and acceptable wall temperatures with 
an FFC of 1 1.5 percent; 

• Calibrated chamber thermal model correlates very 
well with the observed temperature data, enabling 
reliable predictions to be made for a design update 
to a fully dual thrust RCE. 



6. 



NASA Report WSTF-TR-0954-001-01, 
"Nontoxic Orbiter Upgrade, Dual-Thruster 
Reaction Control Subsystem (RCS) Engine 
Test," Lyndon B. Johnson Space Center, 
White Sands Test Facility, dated 11 April 
2000. 

May, Lee, "Orbiter Nontoxic RCS Thrust 
Chamber Assembly Final Report," Final Test 
Report for Contract 9-NRA-96-BE-1, Aerojet, 
Sacramento, California, dated 8 September 
1998. 

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American Institute of Aeronautics and Astronautics