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work on propellant tankage 6's has greatly improved the performance factor. State of the art extensivelyJ TM These non-optimized thrusters have achieved Isp's up to 360 s at stochiometric mixture ratio. Most recently, 2200-N, GH2/GO2 thrusters were developed for the X-33, the performance factors tanks factor burst generally assumed to be pressurized in flight, this safety factor is conservative for tanks that are not pressurized when humans, launch vehicles, or should volumes enable and 4 5 million-cm tanks for large storable bipropellants, resulting in greatly improved performance. 17 As the result of an space from station propulsion 110 to 220 N) system (thrust have also been levels tested NASA TM-113157 5 are smaller 4 miilion-cm tanks), with other spacecraft are at risk. The for large a safety (lower for (maximum pressure) expected of 1.5. operating Because pressure / tanks are technology demonstrator vehicle Reusable Launch Vehicle. 9 for the factor available have a However, aggressive performance factors are feasible using thin bladder-liners overwrapped with T1000 carbon fiber composite. Prototype bladder-lined tanks of modest size have recently been fabricated which achieved 4 million-cm using thick end domes and two heavy stainless steel bosses sized for automotive applications. Reducing the mass of the bosses and end domes is aggressive compared to space qualified pressure performance factor of vessels which 2 million-cm. million-cm tanks for tank volumes (which modest intensive are close stochiometric development program, these thrusters available. For volumes. result contained Thus, justified even for small volumes, if only the mass increment of turning structural members into pressure vessels is considered as tank weight. This results in a significant weight reduction as compared to the use of conventional tankage. Thrusters For the current study, a I-N GH2/GO2 thruster was build into the testbed. This thruster consisted of an ignitor, an injector, a chamber, a throat, and a 23.3:1 area ratio nozzle. Small GH2/GO2 thrusters have been developed and tested over the last three decades. 13Flight type thrusters built for satellite electrolysis propulsion concepts (thrust levels from 0.5 to 22 N) have been tested extensively. 43'14 A 22-N thruster demonstrated over 69,000 firings with a total of 4 hours burn time without noticeable degradation, achieving an Isp of 355 s. In the same program, a 0.5-N thruster demonstrated over 150,000 firings and 10 hours total burn time, with a performance of 331 s. These tests showed that for these small thrusters, optimal ignition was achieved at higher chamber pressures (>160 kPa), driving optimal designs to operate at higher tank and electrolysis pressures. Thrusters built for potential application as the to being commercially GH2/GOz, Ir/Re with an additional in Small low generally are readily performance factors) within required structural members. factors are oxide may coated coating for increased oxidation-resistance Several 22-N, oxide- have been tested on aggressive performance be a better option. Ir/Re thrusters performance commercially of the best state-of-the-art Ir/Re rockets have allowed the of fuel-film cooling for 6 temperature over chamber material. virtual elimination In all of the past work, fuel-film used for thermal and oxidation thruster walls. The presence of such a fuel-film reduced thruster performance. In order to maximize thruster performance in the highly oxidizing combustion environment of a stochiometric GH2/GOz thruster materials, such as iridium-coated rhenium (lr/Re) may be needed. This material provides a 700 K increase in operating GH2/GO2 up to a mixture ratio of 17.18 Leveraging the results of advanced thruster materials research and redesigning thrusters to operate with radiative cooling alone, can increase specific impulse by a significant margin (projected Isp > 380 s) while at the same time operating in an oxidizing environment. The additional performance that from GH2/GO2 systems is could higher than from using the same storable materials. propellant systems One major difference between GH2/GO2 and established chemical thrusters is the need for an ignition source. Incorporation of an or ignition power stringent rockets. source may increase complexity requirements and may pulsing requirements of Spark ignition has been used extensively in previous GH2/GO2 thruster programs and is the baseline for the X-33 thruster. Alternative ignition sources, including laser, resonance, and catalytic ignition have also been investigated for GHffGO2" 19 Ignition systems are being investigated under technology programs for upgrade of the Shuttle Orbiter RCS and manned lunar/Mars spacecraft, both of which will probably use oxygen/hydrocarbon propellants. not meet some low the thrust cooling was protection thruster, advanced be obtainedPDF Image | Electrolysis Spacecraft Propulsion Applications
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