Nasa engineering models of the mars atmos- phere for entry vehicle design
Mars atmosphere models for entry vehicle design
MARS ATMOSPHERE↗
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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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Engineering models of the Martian atmosphere and surface
Mars surface and atmosphere models for landing module design
LANDING MODULE↗
Design and construction of an engineering model solid cryogen refrigerator for infrared detector cooling at 50 deg K
Design and construction of prototype spacecraft cryogenic refrigerator for infrared detector cooling
Caren, R. P.↗
The NASA space vehicle design criteria program - Engineering models of the environment.
Environmental criteria for NASA space vehicle design program, discussing monographs and environment models
Charak, M. T.↗
Engineering models of the Venus atmosphere based on an interpretation of recent space vehicle observations of Venus.
Venusian atmosphere model developed from satellite and earth based radar data for environmental criteria in spacecraft design and mission planning
Schiffer, R. A.↗
Design documentation and test hardware for an engineering model of a Serpentuator Final report
Design, fabrication, and testing of Serpentuator model for intervehicular performance
Loyd, J. R., III↗
Lunar antenna mast - Engineering model Final report
Design and acceptance of erectable and retractable lunar antenna
Crawford, R. F.↗
Development of an engineering model traveling wave tube amplifier for space communication systems
A design has been made of a 100 watt traveling-wave tube amplifier for use in space communication applications. The features of very high overall efficiency and heat rejection of waste heat at low thermal densities were predominant in the design concept. The design concept was proven by building a series of tubes, operating at efficiencies up to 50%. These tubes utilized heat pipe cooling and heat distribution such that 150 watts of waste heat was rejected at a density of less than 1.5 watts per square inch. A power supply to convert a 28 volt primary line of the needs of the TWT was built and operated at 85% efficiency.
Eallonardo, C. M.↗
Pulsed plasma solid propellant microthruster for the synchronous meteorological satellite. Task 4: Engineering model fabrication and test report
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.
Guman, W. J.↗
Design and fabrication of engineering model fiber-optics detector
The design and fabrication of an annular ring detector consisting of optical fibers terminated with photodetectors is described. The maximum width of each concentric ring has to be small enough to permit the resolution of a Ronchi ruling transform with a dot spacing of 150 microns. A minimum of 100 concentric rings covering a circular area of 2.54 cm diameter also is necessary. A fiber-optic array consisting of approximately 89,000 fibers of 76 microns diameter was fabricated to meet the above requirements. The fibers within a circular area of 2.5 cm diameter were sorted into 168 adjacent rings concentric with the center fiber. The response characteristics of several photodetectors were measured, and the data used to compare their linearity of response and dynamic range. Also, coupling loss measurements were made for three different methods of terminating the optical fibers with a photodetector.
Mcsweeney, A.↗
Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model
Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 percent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 percent for an 85 percent efficient fan stage.
Wesoky, H. L.↗
Low-speed wind tunnel tests of a 50.8-centimeter (20-in.) 1.15-pressure-ratio fan engine model
At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.
Wesoky, H. L.↗
Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model
Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 per cent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 per cent for an 85 per cent efficient fan stage. For comparison, a similar 1.55 pressure ratio fan system would have a propulsive efficiency of 62 per cent.
Wesoky, H. L.↗
A study of the use of remote sensing data in hydrologic engineering models
There are no author-identified significant results in this report.
Algazi, V. R.↗
A computerized engineering model for evaporative water cooling towers
Explore the source record for details and available documents.
Park, J. E.↗
Computing the Power-Density Spectrum for an Engineering Model
Computer program for calculating of power-density spectrum (PDS) from data base generated by Advanced Continuous Simulation Language (ACSL) uses algorithm that employs fast Fourier transform (FFT) to calculate PDS of variable. Accomplished by first estimating autocovariance function of variable and then taking FFT of smoothed autocovariance function to obtain PDS. Fast-Fourier-transform technique conserves computer resources.
Dunn, H. J.↗
Endurance test of a 30-CM-diameter engineering model ion thruster. Task 12: Investigation of thin-film erosion monitors for ion thrusters
An investigation of short term measurement techniques for predicting the wearout of ion thrusters resulting from sputter erosion damage is described. The previously established laminar thin film techniques to provide high precision erosion rate data. However, the erosion rates obtained using this technique are generally substantially higher than those obtained during long term endurance tests (by virtue of the as deposited nature of the thin films), so that the results must be interpreted in a relative sense. Absolute measurements can be performed using a new masked substrate arrangement which was developed during this study. This new technique provides a means for estimating the lifetimes of critical discharge chamber components based on direct measurements of sputter erosion depths obtained during short duration (10 hour) tests. The method enables the effects on lifetime of thruster design and operating parameters to be inferred without the investment of the time and capital required to conduct long term (1000 hour) endurance tests. Results obtained using the direct measurement technique are shown to agree with sputter erosion depths calculated for the plasma conditions of the test and also with lifetest results. The direct measurement approach is shown to be applicable to both mercury and argon discharge plasma environments and should be useful in estimating the lifetimes of inert gas and extended performance mercury ion thrusters presently under development.
Beattie, J. R.↗
Characterization of 8-cm engineering model thruster
Development of 8 cm ion thruster technology which was conducted in support of the Ion Auxiliary Propulsion System (IAPS) flight contract (Contract NAS3-21055) is discussed. The work included characterization of thruster performance, stability, and control; a study of the effects of cathode aging; environmental qualification testing; and cyclic lifetesting of especially critical thruster components.
Williamson, W. S.↗