Взрыв под лунным модулем

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APOLLO EXPERIENCE REPORT
ASCENT PROPULSION SYSTEM
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REQUIREMENTS
To meet guidance requirements, the APS was required to produce 3500 pounds of
thrust, to fire for a total duration of 550 seconds, to develop 90 percent of the rated
thrust within 0.450 second after the start signal, and to decay to 10 percent of the rated
thrust within 0. 500 second after the cutoff signal. The maximum allowable combustion chamber
pressure during start transients was 177 percent of the nominal combustion chamber
pressure. This was a vehicle structural limitation. To minimize the heat
transfer to the chamber, the magnitude of any periodic or uniformly cyclic chamberpressure
fluctuation or oscillation that occurred at a frequency of 400 hertz or less
could not exceed ± 3 psi, and those variations that occurred at a frequency greater than
400 hertz could not exceed ± 6 psi. In addition, the engine was required to be stable
dynamically in the presence of all self -induced or artifically induced disturbances,
thereby causing fluctuations of 175 percent of nominal chamber pressure in the combustion
process. Recovery time to a stable steady-state operation could not exceed
0.020 second....

ENGINE ASSEMBLY
Analyses of the heat-shield structural margins during fire-inthe-
hole (FITH) testing of the engine indicated a negative margin because of the
combustion-chamber-pressure peaks during the start transient of the engine. The contractor
investigated several means of reducing the chamber-pressure overshoot without
changing the valve design, but all solutions tried were unsuccessful. Therefore, the
heat-shield structure on the vehicle was modified to accommodate the higher chamber
overshoot during FITH testing. Similar problems were experienced by the backup
contractor....

PROTOTYPE TESTING
Most of the prototype testing of the APS was conducted on the
WSTF test rig PA-1. This test rig was an ascent-stage structure that incorporated
the APS and RCS equipment in an approximate flight configuration (fig. 13). Simulation
was used except where it would affect the APS performance. The test rig accommodated
both the pressure-fed, 3500-pound-thrust ascent engine for firing in a
downward position and the 16 RCS engines, using a separate propellant-supply system
interconnected with the APS propellant tanks. The PA-1 ascent stage was constructed
of aluminum alloy and consisted of three major sections: the forward cabin, the midsection,
and the aft equipment bay. The structure included provisions for installing heat shields
for the FITH testing. A summary of the tests conducted on test rig PA-1
is given in table 11. These tests included all normal mission requirements and a
number of off-limit tests of possible problem areas such as FITH tests, pressure
overshoots, and component-abort tests. Although all of these hot-firing tests were
not conducted with a qualified engine injector, the tests proved successfully the integrity
of the propellant and pressurization sections.

 

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With flight-qualified engines, 57 tests were conducted over a total firing time
of 3392 seconds. Five engines and six thrust chambers (three heavyweight and three
lightweight) were used. The following are the results of these tests.
1. The mixture ratio of the ascent engine could be predicted within 0.6 percent
for engine operation at chamber pressures of 112 to 130 psia and temperatures of
40' to 100' F with helium-saturated and unsaturated propellants.
2. Combined with RCS operation, the overall APS and RCS mixture ratio could
be predicted within 0. 75 percent.
3. The validity of the propellant-feed- system cold-flow calibrations was verified,
and the flight-engine characteristics were confirmed.
4. The FITH starts, at simulated lunar-launch conditions, were performed
with no structural damake or adverse effect on engine performance.
5. Abort starts were performed safely at ullage pressures of 62 to 215 psia.
6. Engine operation was not affected adversely by operation on redundant
regulators.
7. The transition to the adjusted system-pressure levels was smooth and
gradual.
8. The engine operated safely in the tank-ullage-decay (blowdown) mode from
nominal chamber pressure to 8 psia, and it could be safely shut down by means of
propellant depletion.

 

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APOLLO 15 MISSION REPORT
SUPPLEMENT3
ASCENT PROPULSION SYSTEM FINAL FLIGHT EVALUATION
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7. ENGINE TRANSIENT ANALYSIS
An analysis of the start and shutdown transients was performed with
the primary intention of determining transient total impulse. Figures 11
and 12 are traces of engine chamber pressure, measurement GP2010, during
start and shutdownof the lunar liftoff burn, repectively. No data were
available from the TPI burn.
The time from ignition signal to 90 percent steady-state thrust was
0.345 seconds, well within the specification limit for unprimed starts
of 0.450 seconds. Total start transient impulse was 27 Lbf-sec. The
chamber pressure overshoot exceeded the upper limit of the measurement
range (150 psia); however, there were no indications of rough combustion
or other abnormal performance.

 

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