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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Design, weight, cost, and tooling concepts for cost optimized insulation system for 260 in. dia solid rocket motor
Under the SET Program Task 4 - Regional Turboprop/Turbofan Engine Advanced Combustor Study, a total of ten low-emissions combustion system concepts were evaluated analytically for three different gas turbine engine geometries and three different levels of oxides of nitrogen (NOx) reduction technology, using an existing AlliedSignal three-dimensional (3-D) Computational Fluid Dynamics (CFD) code to predict Landing and Takeoff (LTO) engine cycle emission values. A list of potential Barrier Technologies to the successful implementation of these low-NOx combustor designs was created and assessed. A trade study was performed that ranked each of the ten study configurations on the basis of a number of manufacturing and durability factors, in addition to emissions levels. The results of the trade study identified three basic NOx-emissions reduction concepts that could be incorporated in proposed follow-on combustor technology development programs aimed at demonstrating low-NOx combustor hardware. These concepts are: high-flow swirlers and primary orifices, fuel-preparation cans, and double-dome swirlers.
This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge configuration; and life cycle testing.
The primary challenge of the X-33 CAN is to build and test a prototype LO2 and LH2 densification ground support equipment (GSE) unit, and perform tank thermodynamic testing within the 15 month phase 1 period. The LO2 and LH2 propellant densification system will be scaled for the IPTD LO2 and LH2 tank configurations. The IPTD tanks were selected for the propellant technology demonstration because of the potential benefits to the phase 1 plan: tanks will be built in time to support thermodynamic testing; minimum cost; minimum schedule risk; future testing at MSFC will build on phase 1 data base; and densification system will be available to support X-33 and RLV engine test at IPTD. The objective of the task 1 effort is to define the preliminary requirements of the propellant densification GSE and tank recirculation system. The key densification system design parameters to be established in Task 1 are: recirculation flow rate; heat exchanger inlet temperature; heat exchanger outlet temperature; maximum heat rejection rate; vent flow rate (GN2 and GH2); densification time; and tank pressure level.
Dual mode hydrazine system for attitude control and maneuvering propulsion of interplanetary space probe
Aerodynamic commercial aircraft engine design for minimum noise level
Weightless containerless processing for space, electromagnetic position control, force measurements and techniques, and hydrodynamics
Vacuum jacketed components on launch umbilical tower service arms at Launch Complex 39 - low heat cryogenic rotary joint
Melting and handling of molten material in weightless space environment
Analyzing thermal control system for B2H6/OF2 propulsion module
Test description and results for high Mach number transonic fan stages to determine range and distortion tolerance for various rotor tip casing treatment configurations
Tabulations of blade element and circumferential distortion flow data in tests of high Mach number transonic fan stages with various rotor tip casing treatment configurations
Two flight prototype solid propellant pulsed plasma microthruster propulsion systems for the SMS satellite were fabricated, assembled and tested. The propulsion system is a completely self contained system requiring only three electrical inputs to operate: a 29.4 volt power source, a 28 volt enable signal and a 50 millsec long command fire signal that can be applied at any rate from 50 ppm to 110 ppm. The thrust level can be varied over a range 2.2 to 1 at constant impulse bit amplitude. By controlling the duration of the 28 volt enable either steady state thrust or a series of discrete impulse bits can be generated. A new technique of capacitor charging was implemented to reduce high voltage stress on energy storage capacitors.
The author has identified the following significant results. Although it is readily recognized that there is a need for ground truth to provide adequate guidance for remote sensing data interpretation, it is noted that, in terms of radar remote sensing, this ground truth is often inadequate. It is necessary to make basic electrical and physical measurements of the surface and to some depth below it. A brief outline is presented of a ground truth scheme which uses measurements of the dielectric constant. Two portable instruments were designed specifically for this purpose; these were: (1) a Q-meter for measurement of dielectric constant and loss tangent; and (2) an instrument to measure electrical properties of the two operating frequencies of the imaging radar. Although extensive data are lacking, several general cases of radar-earth surface and interaction are described; also, examples of radar imagery and some data on ice and snow are presented. It is concluded that the next logical step is to begin to quantify the radar ground truth in preparation for machine interpretation and automatic data processing of the radar imagery.
There are no author-identified significant results in this report.
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Waste monomethyl hydrazine scrubber liquor, consisting of aqueous solutions containing small amounts of CH4, Cl2, CH3Cl, CH2Cl2, and CHCl3 as well as large amounts of CH3OH is scheduled to be dumped in stabilization ponds along with nitrate and nitrite salt solutions obtained as waste liquors from the N2O4 scrubbers. The wastes are investigated as to the hazardous materials generated by such combinations of items as described as well as the finite lifetime of such materials in the stabilization ponds. The gas liquid chromatograph was used in the investigation. A series of experiments designed to convert nitrate and nitrite salts to the environmentally innocuous N2O and N2 using solar energy is reported. Results indicate that this solar conversion is feasible.
Methods of accelerated and abbreviated testing were developed and applied to solar cell encapsulants. These encapsulants must provide protection for as long as 20 years outdoors at different locations within the United States. Consequently, encapsulants were exposed for increasing periods of time to the inherent climatic variables of temperature, humidity, and solar flux. Property changes in the encapsulants were observed. The goal was to predict long term behavior of encapsulants based upon experimental data obtained over relatively short test periods.