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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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Inducing losses in the vortex chamber of vortex valves and vortex amplifiers resolves the problem of unstable operation caused by a sufficiently large positive feedback. Induced losses also reduce pressure gain and throttling range of vortex pressure amplifier.
Utilizing small throttling capability about high thrust region of 92.5 % for lunar module powered descent
Rocket engine prototype for Mariner Mars 71 spacecraft, describing performance test results and response, thermal and throttling characteristics
Combustion driven shock tunnel applied to tailored interface operating conditions, controlling mass flow into tube by throttling plates
Valve throttles transfer of vapor between heat input surface and heat rejection surface of heat pipe to control rate of transfer of energy. Valve is operated by control signals which sense temperatures or pressures.
Evaluation of throttling injector concepts applicable to advanced cryogenic engines
Liquid propellant rocket engine throttling injector design based on exhaust gas counter flow and heat transfer for inflowing propellant vaporization and combustion stability tests
High-speed motion pictures were taken of conditions in the injector liquid-oxygen cavity of an RL-10 rocket engine during throttled engine operation. Photographs were taken during operation of the engine in the chugging region as the helium gas was injected to stabilize combustion, during operation at rated thrust, and during transition into chugging conditions as the gas injection was discontinued. Results of the investigation indicate that, during chugging rocket operation of the RL-10 engine, a high population of fairly large bubbles formed and collapsed within the liquid-oxygen cavity at the same frequency as the chamber pressure oscillations. When gaseous helium was injected into the liquid-oxygen cavity, a fog rapidly spread over the entire field of view, and the system immediately became stable. The injection of gaseous helium at rated conditions produced a very slight increase in engine performance but not enough to produce a net gain in a typical mission payload with the extra equipment needed. The inherent low-frequency system instability associated with the fuel system at low thrust levels was reduced by injecting either gaseous helium or hydrogen. Complete stabilization was achieved in some cases, and a reduction in the severity of the oscillations in others. This was apparently due to the ·anchoring of the phase change front to the location of the gas injection.
Application of vortex amplifier principles to throttling of monopropellant thrusters is described. Integration of flow control, catalyst bed, and thruster nozzle is accomplished by design of thruster. Diagrams of thrusters are provided to show details of construction and operation.
The development of a digital autopilot program for a transport aircraft and the evaluation of that system's performance on a transport aircraft simulator is discussed. The digital autopilot includes three axis attitude stabilization, automatic throttle control and flight path guidance functions with emphasis on the mode progression from descent into the terminal area through automatic landing. The study effort involved a sequence of tasks starting with the definition of detailed system block diagrams of control laws followed by a flow charting and programming phase and concluding with performance verification using the transport aircraft simulation. The autopilot control laws were programmed in FORTRAN 4 in order to isolate the design process from requirements peculiar to an individual computer.
The evaluation and understanding of maximum propellant utilization, with mercury used as the propellant are presented. The primary-electron region in the ion chamber of a bombardment thruster is analyzed at maximum utilization. The results of this analysis, as well as experimental data from a range of ion-chamber configurations, show a nearly constant loss rate for unionized propellant at maximum utilization over a wide range of total propellant flow rate. The discharge loss level of 1000 eV/ion was used as a definition of maximum utilization, but the exact level of this definition has no effect on the qualitative results and little effect on the quantitative results. There are obvious design applications for the results of this investigation, but the results are particularly significant whenever efficient throttled operation is required.
Analytical and experimental phases of the subject investigation are described. The analytical program for the single jet determines the terminal shock location, the jet boundary, the interface profile, the bow shock profile, the shear layer growth and the dead air region pressure. The experimental program described was conducted over the range from free stream Mach 0.4 to 2.0 at angles-of-attack up to 18 deg and at thrusting coefficients up to C sub T = T/q sub infinity A sub m = 30. Variables investigated included aeroshell angle, number of nozzles, engine thrust, size of nozzles, nozzle throttling and gas composition. The influence of these variables on the aeroshell stability, drag, and loads was determined by integrating pressure measurements on the aeroshell. The total system forces consist of components due to pure thrust and components due to pressure on the aeroshell arising from the jet-free stream interaction. Shadowgraphs provided flow field geometries which proved to be within 10% of those predicted analytically.
Variable magnetic baffles were used to allow a change in baffle geometry to improve thruster performance and stability. Test results are presented. Thruster performance characteristics were analyzed to show how magnetic baffle field strength can improve performance at a fixed beam current, as well as increase the range of ion beam current for throttling.
Throttleable monopropellant hydrazine catalytic reactors of a size applicable to a planetary landing vehicle were designed, fabricated, and tested. An experimental evaluation of two 2670-N reactor designs was conducted. The steady state and dynamic characteristics of the thruster/valve combinations were determined. The results of the testing, including the engine characteristic velocity, smoothness of combustion, insensitivity to heat sterilization, and response during various simulated duty cycles are presented and discussed. No problems of a fundamental nature were encountered as a result of rapid dynamic throttling of these large hydrazine reactors.
Analyses were performed to determine the structural loads caused by maneuvers of the docked Apollo command and service module/lunar module (CSM/LM). Results of CSM/LM docked interface loads analyses and service-propulsion-system engine support structure load analyses are compared with the structural allowable loads of the CSM/LM docked interface and the service-propulsion-system engine thrust mount, respectively, for different phases of spaceflight operations. An analysis also was performed to investigate the loads that resulted from the failure case in which the LM descent propulsion system is started in the full-throttle position.
An analytical study has shown that, with proper selection of feedback gains, an automatic control system can reduce excursions in altitude of a jet transport flying in turbulence without increasing structural loads. The control system uses feedback of attitude-angle and pitch-rate signals to the elevator and uses feedback of altitude and altitude-rate signals to the throttle.
Study of the multimission capability of a solar electric propulsion spacecraft. The study includes such factors as spacecraft constraints, payload variations with launch opportunity, propellant on-loading and off-loading requirements, and limitations in thrusting time, throttling range and steering. The study determined that a single spacecraft design, with P sub zero = 12 kW and I sub s = 3000 sec launched by a Titan IIIC or Titan IIID/Centaur can be used to conduct at least five diverse missions: Eros rendezvous, 1 AU out-of-ecliptic, Mercury orbiter, Jupiter flyby/probe, and Jupiter orbiter. Flight time and injection velocity vary as needed to achieve the desired payload capability. The mission analysis approach and results are presented.
Thruster performance and stability over a range of beam currents is a function of the hollow cathode propellant flow which depends in part on the baffle geometry. Variable magnetic baffles have been used to allow a change in baffle geometry to improve thruster performance and stability. Test results of a 30-cm thruster with magnetic baffle are presented. Thruster performance characteristics are analyzed to show how the magnetic baffle field strength can improve performance at a fixed beam current, as well as increase the range of ion beam current for throttling.