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Performance and operating characteristics are presented for a solar powered aircraft intended to remain aloft for long periods. The critical technologies which limit the performance are identified. By using the techniques presented, the effects of variation in the system parameters are studied. Practical design consideration are discussed.
Performance evaluation of Brayton space power system 400-hertz inverters
Predicted and observed calorimetric exhaust efficiencies of capacitor powered coaxial plasma guns
Predicted and observed calorimetric exhaust efficiencies of capacitor powered coaxial plasma guns
MOS/silicon-on-sapphire (SOS) technology is shown to permit the realization of large-scale integrated arrays that combine the best features of monolithic bipolar and MOS technologies. The perfect isolation and reduced capacitance of SOS technology make possible static MOS circuits with nearly an order-of-magnitude improvement in speed and dynamic power dissipation over their monolithic counterparts. CMOS/SOS memory arrays have been fabricated with speeds comparable to TTL bipolar memory arrays even when operated at TTL compatible levels. Quiescent power dissipation of two described arrays is typically less than 1 microwatt/bit. The SOS/MOS technology is compatible with both aluminum and self-aligned silicon-gate processing.
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The autorotative performance of an assumed helicopter was studied to determine the effect of inoperative jet units located at the rotor-blade tip on the helicopter rate of descent. For a representative ramjet design, the effect of the jet drag is to increase the minimum rate of descent of the helicopter from about 1,OO feet per minute to 3,700 feet per minute when the rotor is operating at a tip speed of approximately 600 feet per second. The effect is less if the rotor operates at lower tip speeds, but the rotor kinetic energy and the stall margin available for the landing maneuver are then reduced. Power-off rates of descent of pulse-jet helicopters would be expected to be less than those of ramjet. helicopters because pulse jets of current design appear to have greater ratios of net power-on thrust to power-off, drag than currently designed rain jets. Iii order to obtain greater accuracy in studies of autorotative performance, calculations in'volving high power-off rates of descent should include the weight-supporting effect of the fuselage parasite-drag force and the fact that the rotor thrust does not equal the weight of the helicopter.
The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.
The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.
The Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) has been providing power to the Mars Science Laboratory (MSL) rover, Curiosity, for surface operations since 2012. The Jet Propulsion Laboratory (JPL) has employed the Life Performance Prediction Model (LPPM) to generate predictions and outputs based upon flight data from the rover and experimental data from the thermoelectric (TE) couples. In order to make these predictions, LPPM requires various spacecraft inputs relevant to RTG performance such as fin root temperature, load voltage, general purpose heat source (GPHS) inventory, etc., combined with couple data such as thermoelectric properties, interface degradation, and sublimation properties. With tens of thousands of hours of flight data to date, LPPM can not only track performance to date, but also make predictions on the power output of the MMRTG through the 17-year end-of-design-life (EODL). These predictions are useful for mission planners and surface operators who rely on the MMRTG for power in order to perform the critical science necessary during the mission. This paper will showcase the power data extracted from the flight electronics, comparing it to the predictions and outputs generated from LPPM.