The performance and heat transfer characteristics of a doublet element type injector for the space shuttle orbiter maneuvering engine thrust chamber were investigated. Ths stability characteristics were evaluated over a range of chamber pressures and mixture ratios. The specific objectives of the test were: (1) to determine whether stability has been influenced by injection of boundary layer coolant across the cavity entrance, (2) if the injector is stable, to determine the minimum cavity area required to maintain stability, and (3) if the injector is unstable, to determine the effects of entrance geometry and increased area on stability

Pauckert, R. P.

English

NASA Technical Reports Server

ASR74-204
Page 1
INTRODUCTION
The like-doublet element type injector is one of the two main candidates
for the Space Shuttle Orbit Maneuvering Engine Thrust Chamber. Rocketdyne's
L/D #1 injector has been extensively tested at Rocketdyne and NASA/White
Sands test facilities in solid wall and regeneratively cooled chambers to
determine its performance and heat transfer characteristics. However, the
only bomb tests conducted to determine its stability characteristics were
performed without film coolant being injected across the cavities which were
tuned to the first tangential mode. The L/D #2 injector which has undergone
extensive stability tests, did not have boundary layer film coolant injected
across the cavity entrance as does L/D #1.
The injector is stabilized by acoustic cavities which have gradual sloping
entrances to facilitate regenerative cooling in this area with minimum pres-
sure drop and low fabrication cost. Tests on other NASA sponsored programs
have indicated this inlet configuration to be unstable with certain injectors
while a sharper-edge entrance was stable.
The purpose of the present program is to define the stability characteristics
of the L/D #1 injector over the range of OME chamber pressures and mixture
ratios. The specific objectives are as follows:
1) To determine whether stability has been influenced by injection
of BLC across the cavity entrance.
2) If the injector is stable to determine the minimum cavity area
required to maintain stability.
3) If the injector is unstable to determine the effects of entrance
geometry and increased area on stability.
The injector and cavity configurations will be bomb-tested in solid wall
thrust chamber hardware typical of a flight contour with fuel heated to
simulate regenerative chamber outlet temperatures.
(NASA-CR-134359) SPACE SHUTTLE
MANEUVERING ENGINE REUSABLE THRUST CHAMBER 74-3127
PROGRAM. TASK 11: LOW EPSILON
STABILITY TEST PLAN (Rocketdyne) )8 p Unclas
HC $4.00 Unclas
CSCL 21H G3/28 46166
https://ntrs.nasa.gov/search.jsp?R=19740023163 2020-03-23T04:25:21+00:00Z
ASR74-204
Page 2
TEST HARDWARE
The test hardware consists of the L/D #1 injector, solid-wall thrust chamber
and cylindrical extension, the fuel manifold, and replaceable acoustic cavity
rings.
The L/D #1 injector, shown in Fig. 1, has 186 elements arranged in 9 rows.
Oxidizer orifice diameters range from 0.032 to 0.038 inches while fuel orifice
diameters range from 0.028 to 0.033 inches. There are also 68 fuel orifices
(0.020 inch diameter) to supply boundary layer coolant amounting to 2.7 percent
of the total propellant flow. Under nominal conditions the injector pressure
drops are 56 psi on the oxidizer side and 62 psi on the fuel side. The diameter
of the injector face is 8.2 inches. The injector has been fired 284 times for
a total duration of 1695 seconds. It has been dye-penetrant inspected to verify
that all welds are sound.
The solid wall chamber and combustor extension are shown in Fig. 2. The dis-
tance from the injector face to the throat is 16 inches. The heat sink capa-
bility of the chamber allows firing durations of up to approximately 5 seconds.
The chamber and extension are instrumented for high response chamber pressure
measurements. Steady-state chamber pressure measurements are made in the
acoustic cavities. Circular grooves are trepanned into the wall at several
locations to provide backwall transient temperature measurements which are
indicative of essentially one-dimensional heat flux values. Two ports for
bombs are located in the center of the four-inch long extension.
The fuel manifold (Fig. 3) distributes fuel to the injector and retains the
acoustic cavity inserts. The acoustic cavities are formed by the injector and
the replaceable two-piece cavity rings shown in Fig. 4. The aft ring defines
the inlet geometry of the cavity and can be replaced with a new ring to provide
a different inlet geometry without machining the forward ring. The forward
ring defines the cavity width and depth. Only the forward ring need be modi-
fied to change the cavity depth. The rings are pinned together and to the fuel
manifold to assure consistent orientation.
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ASR74-204
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The initial configuration of the rings provides the same configuration as
that of the regeneratively cooled integrated thrust chamber. This provides
a gradual entrance to the cavity which results in a gradual turn of the
regenerative coolant passage in this area which maintains a low pressure
drop. The same milling cutter is used in this area as for the throat and
forward end of the chamber thereby eliminating additional set up costs
during manufacturing. All cavities are of the same width (0.5 inches).
Eight of the cavities are tuned (1.75 inch depth*) to stabilize the first
radial mode; the remaining four cavities are 0.92 inches deep* to stabilize
the third radial and first tangential modes.
TEST FACILITY
The tests will be conducted at the Victor Test Stand of the Rocketdyne
Research Test Facility at Santa Susana. A schematic of the feed system is
shown in Fig. 5. NTO and MMH will be supplied from pressurized tanks having
maximum pressure capabilities of 2500 and 1500 psia respectively. The
oxidizer flows to the engine at ambient temperature while the MMH is batch
heated in the quantities required for a single firing through the use of a
4.5 gallon heat exchanger (limited to 430 psia) located immediately upstream
of the main fuel valve. In this heat exchanger, hot water flows inside four
concentric coils of one-quarter inch O.D. stainless tubing and provides a
temperature limited heat source for the fuel. The heating water is circulated
in a closed system from a steel reservoir tank through a 2.5 gpm Burke pump,
past an 18 kilowatt Chromalox electrical heater, and then through either the
heat exchanger or a bypass loop back to the reservoir. An alternate supply
of cold water can be introduced into the system to quickly cool the heat
exchanger between tests and thus permit test personnel to work in the immediate
vicinity of the heater test stand. Heat up and cool down time for the system
are both approximately 10 minutes and conveniently fit into the pretest and
posttest procedures so that negligible delays in the firing schedule are
introduced.
The NTO and MMH pas through 4 0 M filters before entering the engine valves.
GN2 purges are supplied downstream of the engine valves.
*Effective acoustic depth
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400 PSIG
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FILTER P GN g.2 oo
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Figure S Propellant Feed Systems and Instrumentation Schematic
ASR74-204
Page 9
INSTRUMENTATION
High response pressure pickups will be used to monitor chamber pressure.
Three type 317 Photocon transducers will be mounted in the cylindrical
spool approximately 2 inches from the injector face at 12, 108, and 228
degrees location. The steady-state value of chamber pressure will be
measured using two Taber type transducers with sensing ports located in
the acoustic cavities. These same type transducers will be used to
measure the fuel and oxidizer injection pressures and the feed system
pressures. The temperature of the gas in the acoustic cavities will
be measured using tungsten/rhenium thermocouples. Chromel/alumel thermo-
couples will be used to measure the thrust chamber wall temperature.
Propellant feed system temperatures will be measured with iron/constantan
thermocouples. Two turbine flow meters will be used to measure each pro-
pellant flow rate. Thrust will also be measured for computation of c*.
The instrumentation is listed in Table 1. The estimated precision of each
of the critical measurements (thrust, chamber pressure, and flow rate) is
0.25 percent.
High response data will be recorded on tape and oscillograph. The oscillo-
graph will also be used to record the slower responding chamber pressure
measurements, the flow rates, and the injection pressures. Most data except
the high speed data will also be recorded on a digital tape. Direct inking
charts will be used to provide quick-look data.
TEST PROGRAM
The test program is structured to investigate three different acoustic cavity
configurations with multiple tests scheduled for each configuration. The
logic for selecting the configuration is shown in Fig. 6. Initial testing
will be done with the acoustic cavity configuration which is incorporated
into the integrated thrust chamber, i.e., 14.8 percent open area with gradual
entrance. In order to minimize inactive time on the facility the two subse-
quent configurations to be tested will be fabricated based on the results of
the first test series. If the ITC cavity configuration is stable two con-
TABLE 1
INSTRUMENTATION LIST FOR L/D #1 STABILITY PROGRAM
Recordin System .
Beckman Digital Direct Read- Oscillo-
Parameter/Measurement Symbol Transducer Employed Data System ing Recorder graph Tape
MMH (Fuel) System
MMH Tank Pressure PFT Taber* X
Fuel Flowrate #1 WF-1 Turbine Flowmeter X X X
Fuel Flowrate #2 WF-2 Turbine Flowmeter X X X
Fuel Line Pressure PFL Taber X X
Fuel Line Temperature TFL I/C TC** X X
Fuel Heater Temperature #1 TFH-1 I/C TC X
Fuel Heater Temperature #2 TFH-2 I/C TC X
Fuel Heater Temperature #3 TFH-3 I/C TC X
Fuel Injection Temperature TFI I/C TC X X
Fuel Injection Pressure PFI Taber X j X X
N20O (Oxidizer) System
N204 Tank Pressure POT Taber X
Oxidizer Flowrate #1 WOX-1 Turbine Flowmeter X X X
Oxidizer Flowrate #2 WOX-21 Turbine Flowmeter X X X
Oxidizer Line Pressure POL Taber X X
Oxidizer Line Temperature TOL I/C TC X X
Oxidizer Injection Temperature TOI I/C TC X X
Oxidizer Injection Pressure POI Taber X X XL~ ~ ~ ~ ~... .......... . .......
TABLE 1 (Cont'd)
INSTRUMENTATION LIST FOR L/D #1 STABILITY PROGRAM
Recording System
Beckman Digital Direct Read- Oscillo-
Parameter/Measurement Symbol Transducer Employed Data System ing Recorder graph Tape
Thrust Chamber
Cavity Temperature #1 TC-1 W/R TC*** X X
Cavity Temperature #2 TC-2 W/T TC X X
Cavity Temperature #3 TC-3 W/R TC X X
Cavity Temperature #4 TC-4 W/R TC X
Cavity Temperature #5 TC-5 W/R TC X
Cavity Temperature #6 TC-6 W/R TC X
Chailber Pressure #1 PC-l Taber X X
Chamber Pressure #2 PC-2 Taber X X X X
Thrust F Load Cell X X
Chamber Photocon #1 )CPH-1 Photocon X X
Chamber Photocon #2 CPH-2 X X
Chamber Photocon #3 'CPH-3 X X
Miscellaneous
Water Temperature @ Water Tank TW-WT I/C TC X
Water Temp @ Water Heater outlel TW-WHO I/C TC X
Water Temp @ Fuel Heater Inlet TW-FHI I/C TC X
Water Temp @ Fuel Heater Outlet TW-FHO I/C TC X
Reference Junction Temperature RJT I/C TC X
Fuel Main Valves Power & Travel--- ------ X CD -4
Oxid. Main Valve Power & Travel --- ------ X X X
Film Coolant Flowrate WBLC Turbine Flowmeter X X X 4
*Taber Strain Gage Pickup
**Iron/Constantan Thermocouple
***Tungsten/Rhenium Thermocouple
IF GRADUAL ENTRANCE GRADUAL ENTRANCEIF
STABLE r 12% OPEN AREA* r~ 9% OPEN AREA*
INTEGRATED T/C
CAVITY CONFIGURATION
GRADUAL ENTRANCE
14.8% OPEN AREA*
SHARP ENTRANCE GRADUAL ENTRANCE
IF
UNSTABLE 14.8% OPEN AREA* 21% OPEN AREA*
0 (n
(D -4J
*AREAS OF PRIMARY CAVITIES - SECONDARY o
CAVITIES HAVE 1/2 THE AREA OF THE PRIMARY
Fig. 6. Configuration Logic Diagram
ASR74-204
Page 13
figurations having smaller areas will be fabricated and tested to determine
the stability bounds of this configuration.
Should the original configuration prove unstable two variations will be
made to enhance the stability. One variation will be to increase the cavity
area while maintaining the gradual entrance type geometry. The other con-
figuration will maintain the original area in conjunction with a sharp-edged
entrance configuration.
Three test periods will be scheduled. Each cavity configuration will be
tested over the range of conditions shown in Table 2 as time allows. Since
the duration of the test period may not permit conducting all five tests, the
tests will be conducted in the sequence indicated by the numbers. Two 6.5
grain bombs will be detonated in the chamber during each test.
ASR74-204
Page 14
TABLE 2
TEST MATRIX
MIXTURE RATIO
1.45 1.65 1.85
2 3
__I____ ___
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Space shuttle maneuvering engine reusable thrust chamber program. Task 11: Low Epsilon stability test plan

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