NASA Technical Reports Server (NTRS) 19910014990: Carbon monoxide and oxygen combustion experiments: A demonstration of Mars in situ propellants

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NASA Technical Memorandum 104473
AIAA-91-2443
Carbon Monoxide and Oxygen Combustion
Experiments: A Demonstration of Mars
In Situ Propellants
Diane L. Linne
Lewis Research Center
Cleveland, Ohio
Prepared for the
27th Joint Propulsion Conference
cosponsored by AIAA, SAE, ASME, and ASEE
Sacramento, California, June 24-27, 1991
NASA
CARBON MONOXIDE AND OXYGEN COMBUSTION EXPERIMENTS:
A DEMONSTRATION OF MARS IN SITU PROPELLANTS
Diane L. Linne
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio 44135
Abstract The atmosphere of Mars is
95 percent carbon dioxide. Several
The feasibility of using carbon detailed studies have defined a pro-
monoxide and oxygen as rocket pro- duction system that would separate
pellants was examined both experi- the carbon dioxide into carbon monox-
mentally and theoretically. The ide and oxygen, and then store the
steady-state combustion of carbon propellants at cryogenic temperatures
monoxide and oxygen was demonstrated (Refs. 1 to 4). Mission analyses
for the first time in a sub-scale have shown a potential for signifi-
rocket engine. Measurements of cant reductions in Earth launch mass
experimental characteristic velocity, for both Mars precursor missions
vacuum specific impulse, and thrust and for the manned Mars missions
coefficient efficiency were obtained (Refs. 5,6). These analyses have
over a mixture ratio range of 0.30 to been based on theoretically calcu-
2.0 and at chamber pressures of 1070 lated values for combustion perform-
and 530 kPa (155 and 77 psia). The ance of a carbon monoxide and oxygen
theoretical performance of the pro- engine system. Before the use of in
pellant combination was studied para- situ propellants can be planned into
metrically over the same mixture the SEI mission architectures, exper-
ratio range. In addition to one- imental testing and more rigorous
dimensional ideal performance predic- theoretical performance predictions
tions, various performance reduction will be needed.
mechanisms were also modeled, includ-
ing finite-rate kinetic reactions, Some parametric calculations have
two-dimensional divergence effects, been conducted for the performance of
and viscous boundary layer effects. a carbon monoxide and oxygen rocket
engine as a function of mixture
Introduction ratio, chamber pressure, and nozzle
area ratio (Refs. 7,8). These para-
As currently envisioned, the metric studies have all used one-
Space Exploration Initiative (SEI) dimensional chemical equilibrium
presents an ambitious plan of assumptions such as those in the CEC
expanding human presence into the computer code (Ref. 9). some mission
solar system. The ultimate goal of analysts have reduced these predicted
manned missions to Mars will impose specific impulses by arbitrary
heavy burdens on both financial amounts to account for the various
resources and launch capabilities. performance losses inherent in a
Many new technologies have been pro- rocket engine.
posed for the Mars portion of the SEI
scenario which would offer signifi- The ignition characteristics of
cant reduction in both fiscal and carbon monoxide and oxygen in a spark
launch vehicle requirements. One torch igniter have been evaluated
such proposed new technology is the experimentally (Ref. 10). The igni-
use of indigenous space materials to tion boundaries as a function of
produce propellants for the return mixture ratio and percent hydrogen in
trip from Mars. the carbon monoxide were established.
No experimental programs studying rate kinetic chemistry is assumed in
CO/Oz combustion in a rocket engine the nozzle.
have been performed prior to this
study. The TDK module predicts the
inviscid, two-dimensional expansion
This paper discusses both theo- of the gaseous combustion products.
retical predictions and experimental A finite-difference mesh comprised
results for carbon monoxide and of left runnino characteristics and
oxygen combustion in a rocket engine. streamlines is usedto model the
Theoretically predicted performance divergence losses in the nozzle. The
losses caused by finite rate kinet- initial line for the calculations is
ics, two-dimensional geometry, and generated in a transonic flow module
boundary layer effects are examined. using the output from ODK. As in the
Experimental results are presented as ODK module, finite-rate kinetic chem-
a demonstration of steady-state istry is assumed in the nozzle.
combustion, and are compared to the
theoretical predictions. The MABL module is a boundary
layer module that models the growth
Description of Computer Code of the viscous boundary layer in the
chamber and nozzle. Because the test
The Liquid Propellant Program hardware was a small, low pressure
(LPP) computer code (Ref. 11) was engine, frozen chemistry was used to
used for most of the theoretical pre- generate the necessary gas properties
dictions. This code uses a chamber tables. To simulate the expected
and nozzle geometry together with wall conditions, an estimated wall
thermodynamics and kinetics to pre- temperature profile was input based
dict the various performance losses on knowledge of heat flux profiles in
that an actual engine will experience rocket engines and known operational
in normal operation. The code con- limits of the engine material. For
sists of several modules, each of the theoretical analyses presented in
which models a different type of this paper, MABL calculates the dis-
performance loss. These modules are placement thickness for the actual
One-Dimensional Equilibrium (ODE), chamber and nozzle geometry and uses
One-Dimensional Kinetics (ODK), Two this to obtain a displaced, or invis-
Dimensional Kinetics (TDK), and Mass cid, wall contour. The TDK module is
Addition Boundary Layer (MABL). All then rerun with the new contour. A
modules assume complete combustion in new mass flow rate is obtained by
the chamber, that is, no loss in integrating the new initial line.
energy release due to slow vaporiza- This mass flow is then used in the
tion or nonuniform mixing. calculation of characteristic vel-
ocity, C*, along with the actual or
The ODE module predicts the ideal geometric throat area, to obtain a
engine performance for an input pres- theoretical value of C*. These
sure and enthalpy. The calculations values are referred to as TDK/MABL
are performed using the minimum-free- predictions in the rest of this study
energy methodology. Performance is to indicate that the TDK module was
calculated based on equilibrium chem- rerun with the displaced wall pre-
istry (i.e., infinite reaction dicted by the boundary layer module.
rates). The ODK module predicts the
inviscid, one-dimensional expansion Theoretical Analvsis
of the gaseous combustion products
through the converging-diverging Several mission analyses have
nozzle. Equilibrium chemistry is been performed recently that have
assumed in the chamber and finite- investigated the potential benefits
2
of using carbon monoxide and oxygen used to predict performance losses
produced at Mars for portions of a associated with finite-rate kinetics,
round trip mission (Refs. 5,6). Most two-dimensional flow, and boundary
of these missions have assumed a layer growth.
specific impulse of 260 to 280 sec
for the expected performance of this Predicted Performance Losses
propellant combination. In this
range of specific impulse, a 10 sec To predict performance losses,
change in delivered performance can the LPP computer code requires that
significantly affect the results of an engine geometry be specified. The
the mission analysis. A parametric required parameters include chamber
evaluation was performed to determine radius and length, throat radius,
the expected performance of carbon upstream and downstream radius of
monoxide and oxygen. curvature at the throat, nozzle con-
tour, and nozzle inlet and outlet
One-Dimensional Equilibrium angles. Because no CO/O2 rocket
engine has been designed to date, the
A one-dimensional equilibrium geometry from an RL10 rocket engine
computer code (Ref. 9) was used to was used for this part of the analy-
calculate vacuum specific impulse as sis. This version of the RL10 has a
a function of mixture ratio, chamber throat radius of 6.53 cm (2.57 in.),
pressure, and area ratio. Figure 1 an expansion ratio of 205, and is
shows the results of this parametric regeneratively cooled with hydrogen
study for a mixture ratio range of to an area ratio of 60. For the
0.25 to 2.0, chamber pressures of 1.4 CO/02 analysis, a chamber pressure of
and 20.7 MPa (200 and 3000 psia), and 1.4 MPa (200 psia) was assumed, and
area ratios of 10, 60, 100, 200, and liquid carbon monoxide was used as a
500. The curves exhibit typical coolant (Ref. 8). Liquid oxygen is
liquid rocket engine behavior, with also a viable option as a coolant for
peak specific impulse occurring a carbon monoxide/oxygen engine.
between a mixture ratio of 0.40 and
0.60 (stoichiometric mixture ratio is Figure 2 shows the predicted
0.571). As expected, chamber pres- vacuum specific impulse for the
sure has a small effect on specific carbon monoxide/oxygen propellant
impulse, with only a 5 or 6 sec combination as a function of mixture
increase in specific impulse gained ratio. The four lines represent the
with an increase in chamber pressure performance predicted by the various
from 1.4 to 20.7 MPa. modules of the computer code; they
represent the different types of per-
The figure shows that theoretical formance losses that are obtained in
specific impulses as high as 313 sec an actual engine. The top line (ODE)
are predicted for a low pressure in the figure represents the ideal,
engine with a nozzle expansion ratio one-dimensional equilibrium perform-
of 500. These higher predicted spe- ance; these values are the same as
cific impulses could have a signifi- those in Fig. 1 for an area ratio of
cant effect on the results of the 200 and a chamber pressure of 1.4 MPa
mission analyses that assumed only a (200 psia). The second line (ODK) in
260 to 280 sec specific impulse. The the figure represents the one-
313 sec, however, is an ideal theo- dimensional performance with finite-
retical prediction, and an actual rate kinetics assumed instead of
engine would not be expected to chemical equilibrium. The third
deliver this performance. To predict line (TDK) in the figure adds two-
the performance losses that may occur dimensional flow losses. Finally,
with the operation of an actual the fourth line (TDK/MABL) in the
engine, a second computer code was figure represents the predicted
3
performance with the effects of pressures. As chamber pressure
boundary layer growth also included increases, gas density also
increases, and species production
For the conditions modeled here, increases. The low ODK efficiency
significant performance losses are shown in Fig. 3, therefore, is caused
predicted, and the ideal specific by the low rate of recombination at
impulse of nearly 305 sec is reduced the low chamber pressure.
to 260 sec. Figure 3 shows the spe-
cific impulse efficiency predicted by The ODE and ODK calculations were
the various computer modules. These rerun over the entire mixture ratio
values were obtained by dividing the range at a chamber pressure of
predicted specific impulse by the 20.7 MPa (3000 psia), and much higher
ideal values from the ODE module. ODK efficiencies were obtained. Fig-
The figure shows a significant ure 5 shows the ODK efficiencies
decline in efficiency as the stoi- obtained at the two different chamber
chiometric mixture ratio of 0.571 is pressures. The increase in chamber
approached. pressure reduces the predicted
kinetic losses at the stoichiometric
It can be seen from Figs. 2 and 3 mixture ratio from greater than
that the largest losses occur when 8 percent to a little more than
finite-rate kinetics are included in 3 percent. This figure shows that
the calculations. These losses are although chamber pressure had a neg-
caused by the high rate of dissoci- ligible effect on predicted ideal
ation of the carbon dioxide and the specific impulse (Fig. 1), it can
slow rate of recombination. Figure 4 have a significant effect on actual
shows the mole fraction of carbon specific impulse.
dioxide for different locations in
the engine, as indicated by area The results of the theoretical
ratio. Both ODE and ODK predictions predictions of the performance losses
are shown for three chamber pressures made by the LPP computer code indi-
at a mixture ratio of 0.55. This cate that slow recombination rates
graph illustrates several chemical could cause significant reductions in
reaction patterns. First, most delivered specific impulse at low
recombination occurs very close to chamber pressures. To compare actual
the throat area. Recombination engine performance with the theoret-
begins upstream of the throat when ical predictions, an experimental
the temperature begins to drop and program was conducted to measure the
shifts the equilibrium constant. performance of a low-pressure carbon
Recombination ends shortly downstream monoxide/oxygen rocket engine.
of the throat when the temperature
becomes too low for further reaction. Test Apparatus and Procedure
Because the ODK module considers
finite reaction rates, it does not Test Facilit
predict as much recombination as the
ODE module with its assumption of The experimental tests for this
infinite reaction rates. study were performed in Cell 21 of
the Rocket Lab at the NASA Lewis
Second, the figure clearly shows Research Center. This facility con-
that at higher pressures, there is tains a low thrust rocket engine test
less dissociation in the chamber. stand with supporting fluid systems
This in itself will give a higher that allow precise flow of several
predicted performance, as was seen in fuel and oxidizer combinations. Four
Fig. 1. Finally, the figure shows separate propellant lines were used
that more recombination is predicted for this research program: one oxy-
in the ODK modules at higher chamber gen supply line (primary) to the
engine, one oxygen supply line (sec- 40, where a standard spark plug ini-
ondary) to the spark torch igniter, tiated combustion. The hot gases
one carbon monoxide fuel supply line then travelled down a tube through
to the engine, and one hydrogen fuel the injector manifolding and into the
supply line to the igniter. combustion chamber. At the exit of
the igniter tube, additional gaseous
The flow rate of each of the hydrogen, which had been used to cool
gases in the system described above the outside of the igniter tube, was
was controlled with a sonic orifice. added to the hot gases to increase
Inserted as a component of the pro- the flame temperature. The total
pellant line, each orifice insured a igniter mixture ratio at the exit of
constant flow rate of gas, indepen- the igniter tube was approximately
dent of downstream pressure pertur- 7.5.
bations. By measuring the line
pressure and temperature at a point An eight element triplet injector
just upstream of each sonic orifice, design was used to inject the primary
and using orifice calibration curves, propellants into the combustion cham-
gas flow rates were calculated. Dif- ber. Each triplet element was a
ferent diameter orifices could be fuel-oxygen-fuel (F-O-F) design. The
easily interchanged in the system so eight elements were arranged in a
that the gas flow rate range could be mutually perpendicular manner sur-
varied throughout the test program. rounding the igniter outlet orifice
The primary oxygen flow rate ranged to promote inter-element mixing. The
from 10.9 to 68.0 g/sec (0.024 to two outer orifices had an impingement
0.150 lbm/sec). The carbon monoxide angle of 500 (inclusive). Because
flow rate ranged from 16.8 to the pressure and density of the gases
75.6 g/sec (0.037 to 0.160 lbm/sec). will vary rapidly as mass flow rates
The total flow rate was held rela- change, two injectors were used to
tively constant at 47.5 and cover the desired mixture ratio range
95.3 g/sec (0.105 and 0.210 lbm/sec). of 0.30 to 2.0. Injector 1 was used
ODE predicted chamber pressures for for mixture ratios of 0.30 to 0.80,
these two total flow rates are 520 and injector 2 was used for mixture
and 1240 kPa, respectively (90 and ratios of 0.90 to 2.0.
180 psia). Actual chamber pressures
achieved were approximately 530 and A copper heat sink chamber and
1070 kPa (77 and 155 psia). nozzle were used. The chamber had
an interior diameter of 5.22 cm
Test Hardware (2.055 in.), and was 20.3 cm (8 in.)
long. A chamber pressure tap was
The test hardware for this exper- located at the entrance of the cham-
iment consisted of standard liquid ber next to the injector. The nozzle
rocket engine hardware including an had a throat diameter of 1.15 cm
igniter, injector, chamber spool (0.454 in.), and an exit area ratio
piece, and converging-diverging noz- of 2.363. The diverging nozzle con-
zle. Figure 6 shows a schematic of tour was a cone, with an exit half-
the assembled engine, the injector angle of 150 . Figure 7 shows the
face, and an injector element. engine mounted on the thrust rig
during a test.
A hydrogen-oxygen spark torch
igniter was used to initiate com- Test Procedure
bustion. Gaseous oxygen and gaseous
hydrogen were injected into the To insure a uniform run profile
igniter chamber at an oxygen-to-fuel throughout the duration of the test
mixture ratio (O/F) of approximately program, each firing of the igniter
5
was sequenced by a programmable line imental tests. The first was
controller. Each test run started characteristic velocity, C*, which
with the initiation of the secondary was calculated based on the measured
oxygen and the hydrogen flows to the chamber pressure and propellant flow
igniter and the primary oxygen flow rates. The second measure of per-
to the engine. One tenth of a second formance was the vacuum specific
later, the spark was started, fol- impulse, which was calculated based
lowed one tenth of a second after on the measured propellant flow rates
that by the carbon monoxide, at which and measured thrust corrected to vac-
point the main combustion was initi- uum conditions by adding the nozzle
ated. After combustion started, the exit pressure force. Both of these
secondary oxygen and the hydrogen to performance measurements were com-
the igniter were stopped, and the pared to theoretical values predicted
test continued for 1.2 sec with no by the LPP computer code.
hydrogen flowing. This sequencing
allowed for hydrogen to be present Some of the experimental results
during ignition of the engine to aid are tabulated in Appendix A. For
in the ignition of the dry carbon each of the two chamber pressures,
monoxide and oxygen mixture three tests were performed at each
(Ref. 10). The steady-state portion mixture ratio. Only one test at each
of the test run from which the data mixture ratio is shown in the tables
was taken, however, was after the as a representative value.
hydrogen flow had been terminated,
demonstrating steady-state combustion Figure 8 shows the experimental
of dry carbon monoxide and oxygen. and theoretical values of character-
istic velocity for two chamber pres-
Experimental data was gathered sures over a range of mixture ratios.
during the test runs by a high-speed As seen before, the chamber pressure
data acquisition system. In addition has little effect on the theoretical
to the instrumentation on the hard- predicted C*. For the experimental
ware, pressure transducers and results, different symbols are used
thermocouples were applied to the to denote results from the two dif-
facility feed systems to properly ferent injectors. As can be seen in
measure the propellant flow rates and the figure, a discontinuity exists
temperatures. A total of 100 instru- where the injectors were changed.
mentation channels were scanned at Table 1 lists some of the operating
the rate of 100 times per second per characteristics of the injectors.
channel. For each channel, every ten Pressure drop as a percent of chamber
readings were combined to provide pressure, injection velocities,
approximately ten averaged data velocity ratio, and momentum ratios
points per second. All values quoted are listed at each chamber pressure.
in this analysis were obtained by Velocity and momentum ratios are cal-
averaging together three of these culated as fuel to oxidizer ratios.
averaged data points. Therefore, Each injector was designed for the
each value quoted is an average of 30 midpoint of the mixture ratio range
readings of the instrument by the at which it would be used. Because
data system. The data reduction was the densities and pressures of the
performed by a FORTRAN 77 computer gases vary rapidly as mass flow rates
program hosted on a VAR cluster. change, the upper and lower end of
each injector's operating range may
Experimental Results produce nonoptimum injector perform-
ance. This was the cause of the dis-
Two measures of engine perform- continuity between mixture ratios of
ance were taken during the exper- 0.80 and 0.90.
6
It can be seen in Fig. 8 that the cant here. As in the figures of C*,
experimental C* curve has the same the experimental I curve has a
vac
general shape as the theoretical similar shape to the theoretical
curve, but appears to peak at a curve. Figure 11 shows the experi-
higher mixture ratio and is signifi- mental and theoretical vacuum
cantly lower. To quantify the dif- specific impulse efficiencies as a
ference between experimental and function of mixture ratio. The theo-
predicted C*, the experimental values retically predicted efficiencies are
were divided by the one-dimensional about 93 to 95 percent, while the
equilibrium values, and C* efficiency experimental efficiencies are 85 to
was plotted in Fig.9. The theoreti- 89 percent. This difference is most
cal efficiencies plotted are the likely caused by the incomplete
TDK/MABL predicted values divided by energy release that was observed in
the ODE values. The LPP code pre- the C* efficiency graph. In Fig. 9,
dicts C* efficiencies between 95 and the difference between the theo-
97 percent (note the increase in pre- retical and experimental values of C*
dicted efficiency with increased efficiency average 6 percent. If the
pressure). These theoretical effi- theoretical vacuum specific impulse
ciencies are much higher than those efficiencies in Fig. 11 are reduced
shown in the previous section because by 6 percent to account for losses
the expansion area ratio of the test caused by incomplete energy release,
hardware was only 2.4, and the pre- then the theoretically predicted
dicted kinetic recombination rates efficiency would be approximately
are still close to the equilibrium 88 percent. This is right in line
values to this point. The experi- with the experimental values shown.
mental values fall between 89 and Therefore, the difference in Fig. 11
93 percent. The theoretical predic- between the theoretical and exper-
tions of C* efficiency account for imental curves is again probably
expected performance losses caused by caused by the incomplete energy
finite-rate kinetics, two-dimensional release that the LPP computer code
flow, and boundary layer growth. The does not take into account.
LPP code, however, assumes complete
combustion in the chamber. The dif- Theoretical and experimental
ference between the theoretical and thrust coefficient efficiencies are
experimental C* efficiencies in graphed in Fig. 12. Thrust coeffi-
Fig. 9, therefore, are most likely cient is dependent on the hardware
caused by incomplete energy release geometry, and theoretical and experi-
in the chamber. Because both propel- mental thrust coefficient efficiency
lants were gaseous, the most probable should coincide. The theoretical and
cause of incomplete energy release is experimental values were obtained by
poor mixing between the gases. A dividing the TDK/MABL predicted val-
more optimum injector design would ues and the experimental values of
most likely increase the experimental thrust coefficient by the ODE pre-
efficiency toward the level predicted dicted value. In Fig. 12, the
by the LPP computer code. experimental thrust coefficient effi-
ciencies obtained with injector 1 at
Experimental vacuum specific low pressure (530 kPa) coincide with
impulse was also measured, and is the theoretical predictions. The
shown in Fig. 10 along with the ODE experimental thrust coefficient effi-
theoretical values. It should be ciencies obtained with injector 2 at
noted that because the expansion area low pressure and with both injectors
ratio of the experimental hardware at high pressure (1070 kPa), however,
was only 2.4 the actual magnitude of are approximately 2 percent lower
the specific impulse is not signifi- than theoretical predictions.
7
—Ad
Reexamining Figs. 8 and 10, the the carbon dioxide at the lower pres-
experimental C* for the higher pres- sures was the main cause of the
sure (1070 kPa) is slightly higher kinetic inefficiencies. The results
than that for the lower pressure of the theoretical analysis indicate
(530 kPa), agreeing with the theoret- that a specific impulse in the range
ical predictions. In Fig. 10, how- of 260 to 280 sec is realistic for
ever, the experimental specific the assumption of a low pressure
impulse at the lower pressure is engine. Specific impulses of 290 to
higher than that for the higher pres- 300 sec should be used, however, for
sure for tests run with injector 1. the assumption of a higher pressure,
Because the low pressure, injector 1 pump-fed engine.
set of data is the only set that
coincides with theoretical thrust Gaseous carbon monoxide and oxy-
coefficient efficiencies (Fig. 12), gen were combusted in a sub-scale
it is possible that a bias error was rocket engine, demonstrating steady-
introduced for the remaining specific state combustion of this potential
impulse measurements. An examination Mars in situ propellant combination.
of the raw experimental data and the C* and vacuum specific impulse effi-
thrust stand calibration curves did ciencies of 89 to 93 percent and 85
not disclose any obvious source of to 89 percent, respectively, were
this bias error. obtained from the experimental pro-
gram. These experimental efficien-
('nnn l n c i nn cies are approximately 6 percent
lower than the efficiencies predicted
An engine performance computer by the theoretical computer code for
code was used to parametrically study this specific test hardware. This
the theoretical performance in a car- discrepancy between the theoretical
bon monoxide/oxygen rocket engine. and experimental values is most
Losses caused by finite-rate kinetic likely caused by incomplete energy
reactions, two-dimensional flow release in the chamber due to nonuni-
effects, and boundary layer growth form mixing of the gases. The com-
were calculated. At a chamber pres- puter program results assume complete
sure of 1.4 MPa (200 psia) and an combustion in the chamber.
expansion area ratio of 205, the code
predicted vacuum specific impulse The results of the theoretical
reduction from ideal (ODE) of as much parametric studies and the experimen-
as 14 percent at a stoichiometric tal tests indicate that with careful
mixture ratio. More than 8 percent engine design, a carbon monoxide/
of these losses were caused by oxygen rocket engine can be developed
finite-rate kinetic reactions in the to perform with reasonable effi-
expanding nozzle. Further paramet- ciency. Such an engine will allow
rics indicated that the kinetic los- the use of in situ propellants for
ses were reduced to 3 percent if the return trip from Mars. This
chamber pressure was increased to could significantly reduce the launch
20.7 MPa (3000 psia). This indicates vehicle requirements of future manned
that the high rate of dissociation of Mars missions.
8
Appendix A
TABLE Al. - EXPERIMENTAL DATA FOR CO/O2 COMBUSTION TESTS
(SEE BELOW FOR KEY TO COLUMN HEADINGS)
[P = 530 kPa (77 psia); Injector 1.]
r
Rdg O/F C* I C* ,
,/C" Ivacth' Ivacx' ,/Ivac' Cfth Cfx , Cf'
m/^s m/s percent sec sec percent percent
(a) (b) (c) (d) (e) (f) (g) (h) (i) (j) (k)
123 0.304 1347 1198 89.0 207.4 180.3 86.9 1.510 1.476 97.7
124 .346 1355 1213 89.6 209.3 182.7 87.3 1.515 1.477 97.5
125 .405 1358 1223 90.1 210.1 184.4 87.8 1.517 1.478 97.4
126 .452 1359 1226 90.2 210.2 184.9 88.0 1.479 97.5
127 .502 1358 1227 90.4 210.0 185.5 88.3 1.482 97.7
128 .554 1356 I 90.4 209.8 185.1 88.2 1.480 97.6
129 .614 1353 1 90.6 209.3 185.4 88.6 1.482 97.7
130 .648 1351 90.8 209.1 185.4 88.7 1.482 97.7
131 .699 1349 1225 90.8 208.6 184.6 88.5 1.477 97.4
132 .807 1342 1222 91.0 207.6 184.2 88.7 1.479 97.5
[Pr = 530 kPa (77 psia); Injector 2.]
Rdg O/F C* h, C* ,
^C*' Ivacth' Ivacx' ./Ivac' Cfth Cfx ^Cf'
m^s m/s percent sec sec percent percent
(a) (b) (c) (d) (e) (f) (g) (h) (i) (j) (k)
205 0.855 1338 1203 89.9 206.9 177.8 85.9 1.517 1.450 95.6
204 0.957 1331 1199 90.1 205.8 177.3 86.2 1.516 1.450 95.6
203 1.190 1315 1185 90.2 203.1 174.8 86.1 1.515 1.446 95.4
202 1.400 1298 1170 90.1 200.5 172.5 86.0 1.514 1.446 95.5
201 1.650 1279 1150 89.9 197.2 169.3 85.9 1.512 1.444 95.5
200 1 1.940 1255 1121 89.4 192.9 165.3 85.7 1.508 1.446 95.9
[P = 1070 kPa (155 psia); Injector 1.1
Rdg O/F C* I1I C* ,
^C*' Ivacth' Ivacx' , ivac' Cfth Cfx ,/Cf'
m^s m/s percent sec sec percent percent
(a) (b) (c) (d) (e) (f) (g) (h) (i) (j) (k)
154 0.321 1358 1216 89.5 209.3 179.1 85.6 1.511 1.445 95.6
155 .373 1366 1230 90.0 211.1 181.0 85.7 1.515 1.444 95.3
156 .431 1369 1236 90.3 211.6 182.0 86.0 1.516 1.444 95.3
157 .481 1369 1238 90.4 211.6 182.5 86.2 1.516 1.446 95.4
158 .537 1367 1240 90.7 211.4 182.9 86.5 1.517 1.447 95.4
164 .593 1365 1246 91.3 211.1 183.5 86.9 1.517 1.444 95.2
163 .647 1362 1244 91.4 210.7 183.3 87.0 1.517 1.444 95.2
162 .681 1360 1243 91.4 210.3 183.2 87.1 1.516 1.445 95.3
161 .727 1358 1241 91.4 209.9 182.8 87.1 1.516 1.445 95.3
160 .836 1351 1232 91.2 208.8 181.8 87.1 1.516 1.447 95.4
9