Engine 0209, the certification engine for the new Phase 2+ Hot Gas Manifold (HGM), showed severe deterioration of the Main Combustion Chamber (MCC) liner during hot fire tests. One theory on the cause of the damage held that uneven local distribution of the fuel rich hot gas flow through the main injector assembly was producing regions of high oxidizer/fuel (O/F) ratio near the wall of the MCC liner. Airflow testing was proposed to measure the local hot gas flow rates through individual injector elements. The airflow tests were conducted using full scale, geometrically correct models of both the current Phase 2 and the new Phase 2+ HGMs. Different main injector flow shield configurations were tested for each HGM to ascertain their effect on the pressure levels and distribution of hot gas flow. Instrumentation located on the primary faceplate of the main injector measured hot gas flow through selected injector elements. These data were combined with information from the current space shuttle main engine (SSME) power balances to produce maps of pressure, hot gas flow rate, and O/F ratio near the main injector primary plate. The O/F distributions were compared for the different injector and HGM configurations

Chik, J.

Dill, C.

Mahorter, L.

McDaniels, D.

English

NASA Technical Reports Server

53 2 (	 /VA5#-
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AIRFLOW MODEL TESTING TO DETERMINE THE DISTRIBUTION OF HOT GAS FLOW MJD 0/F RATIO 
ACROSS THE SPACE SHUTTLE MAIN ENGINE MAIN INJECTOR ASSEMBLY
	 /q '775 
L. Mahorter, J. Chik, D. McDaniels, and C. Dill 
National Aeronautics and Space Administration 	 jj'. / b Marshall Space Flight Center, Alabama 
ABSTRACT
	 ,&rt 71 
Engine 0209, the certification engine for the new Phase 11+ Hot Gas Manifold (HGM), showed severe 
deterioration of the Main Combustion Chamber (MCC) liner during hot fire tests. One theory on the cause 
of the damage held that uneven local distribution of the fuel rich hot gas flow through the main injector 
assembly was producing regions of high oxidizer/fuel (0/F) ratio near the wall of the MCC liner. 
Airflow testing was proposed to measure the local hot gas flow rates through individual injector 
elements. The airflow tests were conducted at the Marshall Space Flight Centers Air Flow Dual Leg 
Facility using full-scale, geometrically correct models of both the current Phase II and the new Phase 
11+ HGM5. Different main injector flow shield configurations were tested for each HGM to ascertain their 
effect on the pressure levels and distribution of hot gas flow. Instrumentation located on the primary 
faceplate of the main injector measured hot gas flow through selected injector elements. These data were 
combined with information from the current space shuttle main engine (SSME) power balances to produce 
maps of pressure, hot gas flow rate, and 0/F ratio near the main injector primary plate. 0/F 
distributions were compared for the different injector and HGM configurations. 
INTRODUCTION 
This series of tests was initiated when unusually severe deterioration of the MCC liner developed 
during hot fire testing of the certification engine for the new Phase 11+ Hot Gas Manifold design. 
Damage to the MCC included severe cracking on both the fuel and oxidizer sides, wall blanching, and 
baffle erosion. The goal of the airflow tests was to evaluate uneven local 0/F ratios as a potential 
cause of damage to the MCC liner by deriving and comparing the local hot gas flow rates through the main 
injector elements of both the Phase II and Phase 11+ HGM5. Indications of potentially damaging 0/F 
ratios would be large differences in 0/F distributions or 0/F ratios above stoichiometric (8) which could 
provide free oxygen to react with the chamber wall. 
FACILITY DESCRIPTION 
The airflow tests were conducted at MSFC's Airflow Facility (Dual Leg) in building 4777 (north 
addition) . Flow into the fuel and oxidizer legs is individually controlled by valves which regulate the 
downstream model inlet pressure to a specified value, usually 1.72 x 1o 6
 N/m2 (250 psi). Back pressure 
in the model is maintained by an exit valve located downstream of the model MCC. For this series of 
tests, the exit valve operated with an open area of .0124 sq. m. (19.2 sq. in.) . These conditions allow 
approximately 30 kg/sec (66.3 lbm/sec) of mass flow through the model. The inlet pressures are adjusted 
to produce a flow split of 71% of total mass flow on the fuel side and 29% on the oxidizer side. At 
these conditions, the facility is capable of providing steady state flow for at least 15 seconds. 
Data acquisition equipment in the facility includes a Pressure Systems, Inc. (PSI) system and a 
Hewlett-Packard (HP) computer, as well as various other measurement devices. The PSI system has 
individually calibrated transducers for every model pressure measurement. Facility and data measurements 
are manipulated and stored by the HP computer, then sent from the HP to the VAX 6310. 
MODEL DESCRIPTIONS 
Both the Phase II and the Phase 11+ test HGM5 are modular, full-scale, aluminum models. Major 
differences between the Phase II and Phase 11+ designs are: the change from three to two fuel transfer 
ducts (FTD) with a larger total flow area, an increase in the fuel turn-around duct (TAD) cross-sectional 
area, a larger fuel bowl volume, a larger injector bowl volume, and more efficient, aerodynamically 
rounded duct inlets and outlets. The interior contours of the models match those of the engine designs 
from before the turbine inlets to the MCC. Figure 1 shows the Phase II and Phase 11+ injector bowl 
geometry. 
Both models have turbine simulators on the fuel and oxidizer sides. The turbine simulators consist 
of a porous plate to produce a pressure drop and swirl vanes to simulate turbine exit swirl conditions at 
Full Power Level (FPL or 109%). 
The same main injector is used by both models. The injector is actual flight hardware modified for 
use with the airflow models. Interchangeable flow shields can be configured to produce a variety of flow 
Approved for public release; distribution is unlimited.
117 
PRECEDiNG PAGE BLANK NOT FILMED
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shield configurations. Figure 2 shows a schematic of the flow through the main injector elements. Due 
to the design of the model, neither the LOX flow through the LOX posts nor the hydrogen coolant flow in 
the interpropellant plate region are simulated. 
INSTRUNENTATION 
Both models use similar instrumentation. For these tests the instrumentation was concentrated at 
the turbine exit and in the main injector region. Turbine exit pressure taps are located just downstream 
of the turbine simulator swirl vanes on the hub wall, eight on the fuel side and twelve on the oxidizer 
side. These measurements were used to calculate the turbine exit conditions. 
In the injector region, some special instrumentation was developed specifically for these tests. 
Plugs, Fig. 3, were used in some locations to replace the primary faceplate nut. The plugs measured the 
MCC pressure at the injector primary plate and the stagnated hot gas pressure of a non-flowing element. 
The basic instrumentation in the injector region is the same although the Phase II and Phase 11+ 
models use different injector bowls and flow shield configurations. The injector outer racetrack has 
pressure taps and instrumentation plugs located on three circumferential planes. The injector itself has 
eighty-five pressure taps in various locations. Twenty-four of these pressure taps are located on the 
upper face-plate of the injector, while the others are located on the front and back of the flow shields. 
The pressure taps on the flow shields are mainly located at the center height of the shields but some are 
located closer to the secondary injector plate and some closer to the LOX dome. Most of the instrumented 
shields are grouped near the fuel transfer ducts or near the oxidizer transfer ducts. 
TEST PLAN 
All testing was done with inlet line pressures of approximately 1.72 x 106 N/rn2 (250 psi) and 109% 
power level swirl vanes in the turbine simulator. In all, five engine configurations were tested: 1. 
Phase II HGM with Phase II flow shields; 2. Phase II HGM with no flow shields; 3. Phase 11+ HGM with 
Phase 11+ flow shields; 4. Phase 11+ HGM with Phase II flow shields; and 5. Phase 11+ HGM with no 
flow shields. For each engine configuration, injector element plugs were located in at least 100 
locations. In order to minimize disturbance of the flow, no more than 20 injector element plugs were 
installed in a well distributed pattern at any given time. 
DATA REDUCTION 
For every run, the facility data acquisition system collected and stored seven frames of data per 
label. Each frame consisted of the mean of ten measurements made by the PSI system. After verification 
of test conditions, all seven frames were transferred to the VAX, where they were put into the database. 
The mass flows through the fuel and oxidizer legs, M and Mx respectively, were calculated at the 
facility from the following equations: 
	
=1.098 Cf	 flPTF)()
	
Cf = .99, Af = .0559 m2 (86.59 in2) 
	
p =1.o98 C,	 = .99, Ax = .0285 m2 (44.18 in2) 
The pressure data was non-dimensionalized and scaled to engine conditions using a coefficient of 
pressure, C, linked to the fuel turbine exit conditions. The mean of the steady-state data frames was 
calculated for each measurement of each run. Fuel turbine exit static pressure, fte' was calculated 
from the mean of the turbine exit hub wall statics. The mean turbine exit pressure was then used to 
calculate 1-dimensional turbine exit dynamic pressure, qfte. 
C' ___ 
-w 
For plug data, a CAp was calculated from the difference between the MCC pressure and the hot gas 
pressure, both measured by the primary plate plug, divided by qfte 
118
The pressure drop across the injector element is scaled to engine conditions by: 
scaled = Cp * Qte engine 
An inlet velocity for the retainer holes can be calculated and used to calculate the mass flow of 
hot gas through the injector element. 
= 
I_tPed 
/1photgasK	 thgasuphotgasA 
The hot gas mass flow is then divided into hydrogen and oxygen mass flows. 
thhot gas 
rnfuel In hg
F hotgas 
The amount of hydrogen and oxygen in the hot gas is dependent on the origin of the hot gas, since 
the fuel and oxidizer preburners have different mixture ratios, (0/F)hot gas- The flow split between the 
two sides is shown in Fig. 5. 
A final 0/F ratio for that element is then calculated to include the oxygen flow from the LOX dome, 
rnLOX, and the hydrogen coolant flow through the secondary plate into the hot gas, thcootsnt. The 
theoretical distribution of LOX flow by injector row is: .967% of total LOX flow for each element of 
rows 1 and 2; .974% on row 3; 1.006% on rows 4, 7 and 8; and .998 on rows 8 through 13. 
()	
= rnLOX+rnoxlnhg 
F element rnfuellnhg+rncoolant 
The scaling factors used, table I, were taken from the 109% power level conditions in the Phase 11+ 
power balance 86A, ref. 1 and the Phase II power balance, ref. 2. 
Table I. Power Balance Data Used
Phase II Phase 11+ 
Qte engine 5.64x105 N/rn2 	 (81.82 psia) 5.6lxlO5 N/rn2 	 (81.4 psia) 
Photgas
12.10 kg/rn3 (.7554 lbm/ft 3 ) 12.17 kg/rn3 	 (.7597 lbm/ft3) 
______________________________ 
thcooiant
______________________________ 
1.93 kg/sec (4.26 lbm/sec) 1.89 kg/sec (4.33 lbm/sec) 
_________________________ 
rnLOX .6545 kg/sec (1.443 lbm/sec) .6559 kg/sec (1.446 lbrn/sec) 
(O/F)hot gas .8006 .7872 
mean (O/F)element 6.45 6.45 
w/ out secondary plate coolant: 
(O/F)hot gas .8269 .8135 
(O/F)hg (fuel side) .8789 .8599 
(O/F)hg (ox. side) .7108 .7101 
mean (O/F)element 6.67 6.67
119 
RESULTS 
Most injector plug data was collected on rows 13, 12, 11, 9, 7, 5, 3, and 1. Figure 5 shows the 0/F 
ratios for these rows for the Phase II HGM with Phase II flow shields configuration. The drop in 0/F 
ratio occurring between 1300 and 2300 is due to the low 0/F ratio of the hot gas coming from the oxidizer 
preburner. The 0/F ratios on rows 1 through 12 are grouped with a spread of approximately .5 0/F ratio, 
while row 13 has a lower 0/F ratio than these inner rows. This 0/F ratio distribution is typical of all 
of the configurations, but the model configurations with no flow shields on the injector tend to have a 
slightly more uniform 0/F ratio distribution. For all configurations the 0/F ratios are less than 7.2. 
Baseline Configurations Figure 6 shows the comparison between the Phase II baseline configuration 
(Phase II HGM with Phase II flow shields) and the Phase 11+ baseline configuration (Phase 11+ HGM with 
Phase 11+ flow shields) on row 13. On this row, the Phase 11+ configuration generally has slightly 
higher 0/F ratios than the Phase II configuration. Exceptions are mainly in the regions near 00 to 30° 
and 330° to 360°, where differences in fuel side transfer ducts are greatest. On the oxidizer side, the 
Phase 11+ shows 0/F ratios that are generally .3 0/F higher. Figure 7 shows the same comparison for row 
12. On this row the two configurations are very similar, although there is still a difference of 
approximately .2 0/F on the oxidizer sides. This oxidizer side difference in 0/F is almost nonexistent 
by row 11. The difference between the two configurations are generally small on all of the inner rows. 
Phase 11+ HGM Configurations Figures 8 and 9 show comparisons of rows 12 and 13 for the three 
different flow shields configurations tested on the Phase 11+ HGM. On both rows, the no flow shields 
condition has a higher 0/F ratio, although the difference is smaller on row 12. On rows 11 and 9 the 0/F 
ratios are comparable, while the no flow shields condition produces slightly lower 0/F ratios on rows 7 
through 1. Comparisons of the Phase 11+ HGM shielded configurations show very little difference in the 
0/F ratios produced by the Phase 11+ and the Phase II shield configurations. 
Phase II HGM Confiourat ions Figures 10 compares the two Phase II HGM configurations with the Phase 
II flow shields and with no flow shields for row 13. The no flow shields condition for the Phase 11+ and 
the Phase II HGMs show very similar trends with higher 0/F ratios on the outer rows, comparable 0/F 
ratios in the middle rows, and lower 0/F ratios in the inner rows. 
Figures 11 and 12 show the radial 0/F ratio distributions at 356° and 140°. The difference in 
overall 0/F level between the two plots is due to the differences in fuel and oxidizer side 0/F ratios. 
The elements on the 356° radial receive their flow from the fuel side preburner, while those on the 140° 
radial receive theirs from the oxidizer side. Row 13 has the lowest 0/F ratio, with the 0/F ratio 
generally rising from row 12 to row 7. The drop in 0/F ratio shown in rows 1 through 3 is due to the 
lower mass flow of LOX through these elements.
CONCLUSIONS 
For all of the configurations row 13 had the lowest 0/F ratio, usually below the mean injector value 
of 6.45. If it were desired to further lower the row 13 0/F ratio there is no obvious flow shield 
configuration which would produce these results, since the shieldless injector configurations produce 
higher 0/F ratios on row 13. 
Since all configurations tested showed a maximum 0/F ratio of less than 7.2 and the differences 
between the Phase II and the Phase 11+ baseline configurations were small, it is unlikely that the 
differences in 0/F ratio caused by hot gas flow are the primary cause of MCC wall deterioration in Engine 
0209. It is possible, however, that the hot gas distribution may be contributing to deterioration caused 
by another source. 
REF: 1. IL 88-05-019, Rocketdyne, "SSME Phase 11+ Performance Prediction - Power Balance Model Version 
SSME86A', Process Date May 4, 1988. 
2. OL 87RC09038, Rocketdyne, SSME Phase 	 Predicted Engine Performance", Process Date April 27, 
1987. Phase II Power Balance. 
120
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ransfer 
- 4-4- 180° 
Injector Bow' 
Phase H 
Phase 11+ 
900 
Fuel Transfer Ducts
Figure 1. Schematic of Main Injector Bowls 
(view from MCC towards LOX dome) 
Oxygen	 Oxygen 
4,	 4, 
I III	 Ii Hot Gas = Hydrogen + Steam 
3527.0 psia
Secondary Plate 
liii	 Hydrogen	 liii
Primary Plate I II,	 II 
Iii,	 II 
3276.5 psia
Standard 
Baffle	 Injector 
Element	 Element 
Figure 2. Schematic of Main Injector Element Flows
121
hot gas 
Hot Gas-
	
/Pressure 
öX"""""I	 instrumented 
tube r-	 plug 
-. MCC 
pressure 
primary faceplate 
Figure 3. Instrumentation Plug for Primary Face Plate Nut 
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Mass Flow	 )Q	 Q000Q O 0000000000 
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00000	 Q,JLILI \Q0000Q0'5a0ik'00 
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—..._.-
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•.•.• •• 
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o all fuel side 
0 75% fuel side 
25% LOX side 
• 75% LOX side 
25% fuel side 
• all LOX side
w
LOX Side
Mass Flow
28.7% 
Figure 4. Fuel to Oxidizer Side Flow Split 
122
Phase II HGM with II Flow Shields 
-	 -	 .1	 I	 I	 I	 I	 I	 I	 I	 I	 I7.2-
7.0-
6.8: 
6.6 
6.4: 
6.2-
0/F
6.0: 
5.8 
5.6: 
5.2-
C A.
I I	 I	 I	 I •I I	 I	 I	 I 
--
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4 
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..-	 I I.	 •...W--.-.. 
o	 R13 
•	 R12 
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o	 R09 
a	 R07 
a	 ROS 
V	 R03 
•	 ROl 
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F 30	 F
I	 I	 I	 I 
90	 120	 X	 180
I	 I	 I 
X	 240	 270
I 
F	 330 F 
C 1.. R L a c 
T T T T T 
D D 0 0 D D 
angle	 (dig.) 
Figure 5. Phase II HGM with Phase II Flow Shields 
0/F Ratio Distribution on Row 13 
-	 -	 I	 I	 I	 I	 I	 I	 I 0.0 
6 .5-
6.4-
6.3-
6.2-
6.1-
6 .0-
5.9. 
0/F 5.8-
5.7-
5.6-
5.5-
5.4-
5.3-
5.2-
5.1-
I I I I	 I I I	 I I I	 I	 -
0	 II /11 
11+111+ 
I	 I	 I	 I	 I	 I	 I	 I	 I	 - 0	 30 60 90 120
	 x 180 X
	 240 270 300 330 360 
R	 L 
T	 T 
D	 0 
angle	 (dig.) 
Figure 6. Comparison of Baseline Configurations

Row 13
123 
0/F Ratio Distribution on Row 12 
7.1-	 I	 I	 I	 I	 I	 I	 I	 I	 I 
7.0-

6.9-

6.8-

6 . 7-

6.6-

6 .5-

6 .4-

0/F 6.3-
6 .2-
6 . 1-
6 .0-
5.9. 
5.8-
5.7. 
C C.
o	 II /11 
•	 11+/11+ 
-I	 —I11	 I	 I	 I	 I	 I	 I	 - 
0	 30 60 90 120	 X 180 X
	
240 270 300 330 360 
R	 L 
T	 T 
D	 D 
I	 I	 I	 I	 I I I	 I	 I	 I 
p _ 
I	 I	 I
•
I I	 I	 I 
angle	 (dog.) 
Figure 7. Comparison of Baseline Configurations
Row 12 
0/F Ratio Distribution on Row 13 
o .o 
6 .5 
6 .4 
6 .3 
6.2 
6 .1 
6 .0 
59. 0/F
5 7. 
5.6 
5 .5 
54. 
5.3. 
5. 2 
C 1.
-
I	 I	 I	 I	 I I I	 I	 I	 I - 
•	 11+/11+ 
£	 11+/It 
•	 11+/no 
• I	 I	 1	 I	 -	 I I	 u	 - 
0	 F 60	 90	 120	 x	 180 X 240	 270	 300	 F	 360 
R R 
T T T 
D 0 0 D
angle	 (d.g.) 
Figure 8. Comparison of Phase 11+ HGM Configurations

Row 13 
124
0/F Ratio Distribution on Row 12 
]
7.1 
7.0 
6.9 
6.8 
6.7 
6.6 
6.5 
6.4 
0/F 6.3 
6.2 
6.1 
6.0 
5.9 
5.8 
5.6
--	 liii
•	 II+/II+ 
a	 11+/Il 
•	 11+/no 
F	 60 90 120	 X 180 X	 240 210 iOU	 r .ibU 
R	 I.	 R 
'V	 'V	 'V 
D	 D	 D	 D 
angle	 (dog.) 
Figure 9. Comparison of Phase 11+ HGM Configurations 
Row 12 
0/F Ratio Distribution on Row 13 
0.0 
6 .5-
6.4-
6 .3-
6.2-
6.1-
6.0-
5.9-
0/F 5.8 
5.7. 
5. 6 
5.5. 
5.4. 
53 
5.2-
5.1-
I I	 I	 I	 I I I	 I	 I I 
:' f""H 
o	 it /11 
A	 It/no 
F 30	 F 
C 
'V T 
D D
90 120	 X 180 X 
R	 L 
'V	 'V 
D	 D 
240 270 F 330	 F 
R C 
'V 'V 
D D
angle	 (dog.) 
Figure 10. Comparison of Phase II 11GM Configurations 
Row 13
125 
0 
A 
a 
. 
7.2 
7.0 
6.8 
6.6 
6.4 
6.2 
0/F
6.0 
5.8 
5.6 
5.4 
5.2 
5.0
'I/I' 
I I/no 
11+111+ 
11+/Il 
11+/no 
0/F Ratio Distribution at 356 Dog.

Between Left and Right Fuel Transfer Ducts 
C 
ft	 ft ft R ft B R ft ft ft R R	 ft 
0	 0 0 0 0 A 0 0 0 1 1 1	 1 
1	 2 3 4 5 F 7 8 9 0 1 2	 3 
F 
L 
E 
row number 
Figure 11. Radial 0/F Ratio Distribution at 356 Deg. 
0/F Ratio Distribution at 140 D.g.

Next to Battles at Exit of Oxidizer Right Transfer Duct 
7.2 
7.0 
6.8 
6.6 
6.4 
6.2 
0/F 6.0 
5.8 
5.6 
5.4 
5.2 
5.0
R	 R R R R B R R R R ft R	 ft 
0	 o 0 0 0 A 0 0 0 1 1 1	 1 
1	 2 3 4 5 F 7 8 9 0 1 2	 3 
F 
L 
£
0
	
'I/I' 
A
	
It/no 
.
	
11+111+ 
£	 11+/Il 
11+/no 
row number 
Figure 12. Radial 0/F Ratio Distribution at 140 Deg. 
126


Airflow Model Testing to Determine the Distribution of Hot Gas Flow and O/F Ratio Across the Space Shuttle Main Engine Main Injector Assembly

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