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of gas inflow. The first pressure increase was detected after 100 ms. The likely cause of this rise was slow inside the described, the first hot-fire test experienced a high O/F ratio of 9.3 due to the higher oxygen tank pressure. This caused the oxygen mass flow rate to be greater than stoichiometric. Therefore, during the test, the O/F ratio dropped slightly. Subsequent tests showed lower O/F values, with an approximate equilibrium reached at an O/F of 7.5. The variation in mixture ratio was caused by the particular geometry used in the bench test, where the hydrogen storage volume was more than twice the oxygen storage volume. A configuration designed for optimum performance is not expected to show this large variation, but is expected to operate at a nearly constant O/F of --8. Fig. 9c shows the C* efficiency. It shows that the maximum C* efficiency was obtained after in delay dynamics injector. increased (100-150 ignition ignition undesirable step was 0.45 N thruster. 2 Further measured pressure in the pressure sensing port During the next 100 to approximately 69 ms) hesitation was occurred. The pressure was not smooth. Such for performance reasons. A similar with a that found in testing at Marquardt development program undesirable ignitor. The step was copper chamber tests, and the lr/Re thruster tests, suggesting that it was injector/igniter design which for these laboratory experiments. succeeded phenomenon in by a redesign during this of the both the Ignition occurred at approximately test initiation, after which the The maximum C* level was were efficiency was sharply reached. decreased decrease and less mass flowed into the chamber. The abort acceptable blowdown limit was combustion tests. This selected to provide an until hot test Subsequently, the as the propellant chamber pressure supply pressures design of injector, and chamber. water-cooled adapter section, for optimum pressure. This is present equilibrium of performance Fig. 9a show typical combustion chamber pressures during the Ir/Re thruster tests. All the hot-fire tests show the same step in combustion chamber pressure increase that was shown in Fig. 8. Again this was attributed to the fact that the ignition was not optimized. Such a step should not present an issue in a flight type system. As a result of slightly different chamber dimensions in the Ir/Re thruster, as compared to the copper chamber, the cold gas pressure buildup reached a higher pre-ignition equilibrium level, -78 vs. -68 kPa; and at a later stage, -400 ms vs. -200 ms. Ignition always occurred, with delays varying from 50 to 150 ms. The ignition delay is shown in Fig. 9a. Pressure rise after ignition was slow. A maximum pressure between 173 kPa and 190 kPa, depending on mixture ratio, was reached -1 s after test initiation. After that, the chamber pressure gradually dropped as the result of decreasing propellant supply pressures and thus mass flow rates. The hot-fire, low-pressure abort limit for this specific series of tests was set to 136 kPa, which ended the tests. The thruster performance was reached was designed at 170 kPa chamber at approximately alternate test series done with abort limits of 68 kPa and 34 kPa. Fig. 9b shows the ratio) during the propellant series of mixture tests. As ratio (O/F previously higher chemical NASA TM-113157 9 ms, the kPa. A slight detected before rise through a "step" was under eliminating caused was efficiency was corroborated by pressure by the not optimized 250 ms after pressure conditions were increased during of approximately 1.2 s. This indicated that a significant expelled reached. approximately 0.79. This was expected as the result of the non-optimized fraction before of the optimum propellant mass conditions indeed efficiency combustion pressure from optimum. External chamber wall temperatures did not exceed 1800 °F. Summary where the is obtained. efficiency maximum After this maximum, the as the chamber decreases and decreases the conditions move away Electrolysis an attractive over the decades, but has not yet been used for propulsion has been option for satellite recognized and spacecraft as in-space vapor thruster technology the electrolysis propulsion propellants, without the need for a pressurization system, pumps, or compressors. The gaseous propellant tanks can be sized for the largest burn missions. Recent feed electrolysis, advances propellant in water tankage, chamber materials, warrants renewed and fuel cell consideration for propulsion option. An electrolysis system would generate GH2/GO2 required for the mission, with the bulk propellant stored as water until needed. of the Electrolysis propulsion would provide performance than the established propulsion options and at the same thrust levels. 1.5 s, which combustionPDF Image | Electrolysis Spacecraft Propulsion Applications
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