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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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At least 1,063 records · Page 59
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Isotopic composition of nitrogen - Implications for the past history of Mars' atmosphere

Models are presented for the past history of nitrogen on Mars, based on Viking measurements showing that the atmosphere is enriched in N-15. The enrichment is attributed to selective escape, with fast atoms formed in the exosphere by electron impact dissociation of N2 and by dissociative recombination of N2(+). The initial partial pressure of N2 should have been at least as large as several millibars and could have been as large as 30 millibars if surface processes were to represent an important sink for atmospheric HNO2 and HNO3.

Mcelroy, M. B.↗
NASA NTRS1977

Experimental determination of sound and high-speed flow interaction

An experimental facility for measuring the attenuation of sound at various frequencies and modes in a finite duct with high-speed flow (with or without an axial pressure gradient) is described. The facility consists of an anechoic chamber transonic compressor with inlets of different area variations (or axial gradients); sound of selected frequencies and modes is produced with eight circumferential acoustic drivers. The experimental results indicate that high Mach number inlets can increase noise propagation. In addition, an inlet with a constant axial gradient is found to have a better acoustic and aerodynamic performance than inlets of the same area ratio with steep gradients near the throat or exit.

Lumsdaine, E.↗
NASA NTRS1983

Additional flow field studies of the GA(W)-1 airfoil with 30-percent chord Fowler flap including slot-gap variations and cove shape modifications

Experimental measurements were made to determine the effects of slot gap opening and flap cove shape on flap and airfoil flow fields. Test model was the GA(W)-1 airfoil with 0.30c Fowler flap deflected 35 degrees. Tests were conducted with optimum, wide and narrow gaps, and with three cove shapes. Three test angles were selected, corresponding to pre-stall and post-stall conditions. Reynolds number was 2,200,000 and Mach number was 0.13. Force, surface pressure, total pressure, and split-film turbulence measurements were made. Results were compared with theory for those parameters for which theoretical values were available.

Wentz, W. H., Jr.↗
NASA NTRS1985

Numerical solution of strong radiation gasdynamic interactions in a hydrogen-seedant mixture

Theoretical and experimental results are provided for the absorption of laser energy by a gas flow. A two-dimensional numerical model described the temperature field, fluid dynamics, variable absorptivity and gas dynamics. The model was tested using a hydrogen-seedant mixture injected with 10.6 micron laser radiation. The laser mix was selected to allow reduced temperature gradients and energy source terms. It was found possible, by varying parameters such as the laser power and flash pressure, to precisely position the peak absorptance region in higher speed flows. More refined models are being tested to achieve accurate absorptance predictions in flows below Mach 0.0001.

Merkle, C. L.↗
NASA NTRS1986

Atmospheric environment for space shuttle (STS-51L) launch

A summary is given of selected atmospheric conditions observed near Space Shuttle STS-51L launch time on January 28, 1986, at Kennedy Space Center, Florida. Values of ambient pressure, temperature, moisture, ground winds, visual observations (cloud), and winds aloft are included. The sequence of pre-launch Jimsphere measured vertical wind profiles is given. The final atmospheric tape, which consists of wind and thermodynamic parameters versus altitude, for STS-51L vehicle ascent has been constructed. The STS-51L ascent atmospheric data tape has been constructed by Marshall Space Flight Center's Atmospheric Sciences Division to provide an internally consistent data set for use in post flight performance assessments.

Jasper, G. L.↗
NASA NTRS1989

Development of the improved helicopter icing spray system (IHISS)

Boeing Helicopters has been awarded a contract by the U.S. Army to design, fabricate and test a replacement for the existing Helicopter Icing Spray System (HISS). The Improved Hiss (IHISS), capable of deployment from any CH-47D helicopter, will include new icing spray nozzles and pneumatic pressure source, and a significantly larger water tank and spray boom. Results are presented for extensive bench and icing tunnel test programs used to select and modify an improved spray nozzle and validate spray boom aerodynamic characteristics. The resulting system will provide a significantly larger icing cloud with droplet characteristics closely matching natural icing conditions.

Peterson, Andrew A.↗
NASA NTRS1992

Atmospheric environment for Space Shuttle Atlantis (STS-39) launch

A summary is presented of selected atmospheric conditions observed near space shuttle Atlantis STS-39 launch time on 28 April 1991, at Kennedy Space Center, FL. Values of ambient pressure, temperature, moisture, ground winds, visual observations (cloud), and winds aloft are included. The sequence of prelaunch Jimsphere measured vertical wind profiles is given. The final atmospheric tape, which consists of wind and thermodynamic parameters versus altitude, for STS-39 vehicle ascent was constructed. The STS-39 ascent atmospheric data tape was constructed by NASA-Marshall to provide an internally consistent data set for use in postflight performance assessments and represents the best estimate of the launch environment to the 400,000 ft altitude that was traversed by the STS-39 vehicle.

Jasper, G. L.↗
NASA NTRS1992

Atmospheric environment for Space Shuttle Columbia (STS-40) launch

A summary of selected atmospheric conditions observed near the Space Shuttle Columbia (STS-40) launch time on 5 Jun. 1991, at KSC is presented. Values of ambient pressure, temperature, moisture, ground winds, visual observation (cloud), and winds aloft are included. The sequence of prelaunch Jimsphere-measured vertical wind profiles is given. The final atmospheric tape, which consists of wind and thermodynamic parameters versus altitude, for STS-40 vehicle ascent was constructed. The STS-40 ascent atmospheric data tape was constructed by MSFC's Earth Science and Applications Division to provide an internally consistent data set for use in postflight performance assessments and represents the best estimate of the launch environment to the 400,000-ft altitude that was traversed by the STS-40 vehicle.

Jasper, G. L.↗
NASA NTRS1992

Space suits and life support systems for the exploration of Mars

The requirements and technologies needed for space suits to be used for the manned exploration of Mars are examined. Alternative concepts are proposed for both the space suit and the portable life support system (collectively called the Extravehicular Mobility Unit, or EMU) needed for Mars exploration. EMU system requirements are outlined. It is pointed out that the most fundamental difference between a Mars EMU and those that preceded it is that the design of a Mars EMU must be driven by science and permanent habitability requirements, while all prior EMU designs have been driven by engineering requirements. The EMU weight issues are discussed, and the system mass and mobility concerns are addressed, along with the backpack-to-body-weight ratio. The challenges of thermal and cosmic radiation protection, micrometeorite protection, and EMU system and crew heat rejection are dealt with briefly, as well as the physiological issues of pressure regulation and bacterial or contaminant isolation. A mathematical model is then presented for evaluation of candidate EMU designs and for concept optimization and selection. Lead technology issues are also discussed.

Kuznetz, Lawrence H.↗
NASA NTRS1992

A wall interference assessment/correction system

A Wall Signature method originally developed by Hackett has been selected to be adapted for the Ames 12-ft Wind Tunnel WIAC system in the project. This method uses limited measurements of the static pressure at the wall, in conjunction with the solid wall boundary condition, to determine the strength and distribution of singularities representing the test article. The singularities are used in turn for estimating blockage wall interference. The lifting interference will be treated separately by representing in a horseshoe vortex system for the model's lifting effects. The development and implementation of a working prototype will be completed, delivered and documented with a software manual. The WIAC code will be validated by conducting numerically simulated experiments rather than actual wind tunnel experiments. The simulations will be used to generate both free-air and confined wind-tunnel flow fields for each of the test articles over a range of test configurations. Specifically, the pressure signature at the test section wall will be computed for the tunnel case to provide the simulated 'measured' data. These data will serve as the input for the WIAC method--Wall Signature method. The performance of the WIAC method then may be evaluated by comparing the corrected data with those of the free-air simulation.

Lo, Ching F.↗
NASA NTRS1994

A wall interference assessment/correction system

A Wall Signature method originally developed by Hackett has been selected to be adapted for the Ames 12-ft Wind Tunnel WIAC system in the project. This method uses limited measurements of the static pressure at the wall, in conjunction with the solid wall boundary condition, to determine the strength and distribution of singularities representing the test article. The singularities are used in turn for estimating wall interference at the model location. The development and implementation of a working prototype will be completed, delivered and documented with a software manual. The WIAC code will be validated by conducting numerically simulated experiments rather than actual wind tunnel experiments. The simulations will be used to generate both free-air and confined wind-tunnel flow fields for each of the test articles over a range of test configurations. Specifically, the pressure signature at the test section wall will be computed for the tunnel case to provide the simulated 'measured' data. These data will serve as the input for the WIAC method--Wall Signature method. The performance of the WIAC method then may be evaluated by comparing the corrected parameters with those for the free-air simulation. The following two additional tasks are included: (1) On-line wall interference calculation: The developed wall signature method (modified Hackett's method) for Ames 12-ft Tunnel will be the pre-computed coefficients which facilitate the on-line calculation of wall interference, and (2) Support system effects estimation: The effects on the wall pressure measurements due to the presence of the model support systems will be evaluated.

Lo, C. F.↗
NASA NTRS1996

Application of an Upwind High Resolution Finite-Differencing Scheme and Multigrid Method in Steady-State Incompressible Flow Simulations

The analysis and design of a submarine propulsor requires the ability to predict the characteristics of both laminar and turbulent flows to a higher degree of accuracy. This report presents results of certain benchmark computations based on an upwind, high-resolution, finite-differencing Navier-Stokes solver. The purpose of the computations is to evaluate the ability, the accuracy and the performance of the solver in the simulation of detailed features of viscous flows. Features of interest include flow separation and reattachment, surface pressure and skin friction distributions. Those features are particularly relevant to the propulsor analysis. Test cases with a wide range of Reynolds numbers are selected; therefore, the effects of the convective and the diffusive terms of the solver can be evaluated separately. Test cases include flows over bluff bodies, such as circular cylinders and spheres, at various low Reynolds numbers, flows over a flat plate with and without turbulence effects, and turbulent flows over axisymmetric bodies with and without propulsor effects. Finally, to enhance the iterative solution procedure, a full approximation scheme V-cycle multigrid method is implemented. Preliminary results indicate that the method significantly reduces the computational effort.

NAVIER STOKES EQUATIONS↗
NASA NTRS2017

Europa Propulsion Valve Seat Material Testing

The Europa mission and spacecraft design presented unique challenges for selection of valve seat materials that met the fluid compatibility requirements, and combined fluid compatibility and high radiation exposure level requirements. The Europa spacecraft pressurization system valves will be exposed to fully saturated propellant vapor for the duration of the mission. The effects of Nitrogen Tetroxide (NTO) and Monomethylhydrazine (MMH) propellant vapors on heritage valve seat materials, such as Vespel SP-1 and Polychlorotrifluoroethylene (PCTFE), were evaluated to determine if an alternate material is required. In liquid system applications, Teflon is the only available compatible valve seat material. Radiation exposure data for Teflon in an air or vacuum environment has been previously documented. Radiation exposure data for Teflon in an oxidizer environment such as NTO, was not available, and it was unknown whether the effects would be similar to those on air-exposed samples. Material testing was conducted by Marshall Space Flight Center (MSFC) and White Sands Test Facility (WSTF) to determine the effects of propellant vapor on heritage seat materials for pressurization valve applications, and the effects of combined radiation and NTO propellant exposure on Teflon. The results indicated that changes in heritage pressurization valve seat materials' properties rendered them unsuitable for the Europa application. The combined radiation and NTO exposure testing of Teflon produced results equivalent to combined radiation and air exposure results.

Addona, Brad M.↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results for a Blunt Mars Entry Vehicle Concept during Supersonic Retropropulsion

Simulations of unsteady supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle were performed using the OVERFLOW Computational Fluid Dynamics (CFD) solver. High-fidelity flow solver techniques, including Detached Eddy Simulation (DES) turbulence modeling and Adaptive Mesh Refinement (AMR), were employed to obtain improved realism in CFD predictions. Simulation conditions and geometry configurations were designed to match specific runs in the Descent System Study (DSS) wind tunnel testing (WTT) campaign. The accuracy of each simulation is assessed by direct comparison to experimental data. Comparisons of computational predictions of the SRP flowfield and bow shock shape to experimental schlieren imaging show reasonable prediction of mean shock shape, with approximately 10% similarity in shock standoff distance for selected conditions, as well as similarity in local, time-varying fluctuations of the shock-plume interaction. Comparisons of discrete measurements of surface pressure coefficient (Cp) indicate CFD accuracy within approximately 10% of the experiment across the majority of the model heatshield, with larger variations at some of the heatshield edge locations with stronger flow unsteadiness. Simulated unsteadiness of these chaotic flows, which were highly dynamic and multi-modal, was shown to be within 20-40% of experimentally-measured pressure standard deviation (SD) for the majority of the sampled locations.

Supersonic Retropropulsion↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results for a Blunt Mars Entry Vehicle Concept during Supersonic Retropropulsion

Simulations of unsteady supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle were performed using the OVERFLOW Computational Fluid Dynamics (CFD) solver. High-fidelity flow solver techniques, including Detached Eddy Simulation (DES) turbulence modeling and Adaptive Mesh Refinement (AMR), were employed to obtain improved realism in CFD predictions. Simulation conditions and geometry configurations were designed to match specific runs in the Descent System Study (DSS) wind tunnel testing (WTT) campaign. The accuracy of each simulation is assessed by direct comparison to experimental data. Comparisons of computational predictions of the SRP flowfield and bow shock shape to experimental schlieren imaging show reasonable prediction of mean shock shape, with approximately 10% similarity in shock standoff distance for selected conditions, as well as similarity in local, time-varying fluctuations of the shock-plume interaction. Comparisons of discrete measurements of surface pressure coefficient (Cp) indicate CFD accuracy within approximately 10% of the experiment across the majority of the model heatshield, with larger variations at some of the heatshield edge locations with stronger flow unsteadiness. Simulated unsteadiness of these chaotic flows, which were highly dynamic and multi-modal, was shown to be within 20-40% of experimentally-measured pressure standard deviation (SD) for the majority of the sampled locations.

Supersonic Retropropulsion↗
NASA NTRS2001

Droplet Combustion in a Slow Convective Flow

The objective of the present flight experiment definition study is to investigate the effects of slow forced convective flows on the dynamics of isolated single droplet combustion and is designed to complement the quiescent, microgravity droplet combustion experiments (DCE-1 and DCE-2) of Williams and Dryer. The fuels selected for this study are the same as those of DCE, namely, a sooting alkane fuel (heptane) and a non-sooting alcohol (methanol), and imposed flow rates are chosen between 0 and 20 cm/s with varying ambient oxygen concentrations and pressures. Within this velocity range, both accelerating and decelerating flow effects will also be investigated. Two different approaches to generate the forced flow are currently under development in ground-based facilities; the first is a flow tunnel concept where the forced flow is imposed against a stationary droplet, and in the second a tethered droplet is translated at a specified velocity in a quiescent ambient medium. Depending upon the engineering feasibility a selection will be made between these two approaches so that the experiment can be accommodated in the Multiple Droplet Combustion Apparatus (MDCA) currently being designed for the International Space Station. Recently, we have finished designing and fabricating the experimental rigs using both the above mentioned concepts. The flow tunnel concept is implemented in a 2.2 second drop tower rig. Preliminary experiments have been carried out using heptane and methanol in air at atmospheric pressure. The translating droplet apparatus is scheduled to be tested in the 5 second drop facility in the near future. This report presents some of the experimental results obtained for heptane.

Nayagam, V.↗
NASA NTRS1975

A Three-Dimensional Solution of Flows over Wings with Leading-Edge Vortex Separation. Part 1: Engineering Document

A method of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings with leading-edge vortex separation in incompressible flow is presented. The method employs an inviscid flow model in which the wing and the rolled-up vortex sheets are represented by piecewise, continuous quadratic doublet sheet distributions. The Kutta condition is imposed on all wing edges. Computed results are compared with experimental data and with the predictions of the leading-edge suction analogy for a selected number of wing planforms over a wide range of angle of attack. These comparisons show the method to be very promising, capable of producing not only force predictions, but also accurate predictions of detailed surface pressure distributions, loads, and moments.

Brune, G. W.↗
NASA NTRS1973

Evaluation of coated columbium alloy heat shields for space shuttle thermal protection system application. Volume 2, phase 2: Subsize heat shield and small size TPS evaluation

Initially a trade study was conducted of seven heat shield configurations. These were evaluated for structural reliability, fabricability, weight, inspectability, and refurbishability. Two concepts, a tee-stiffened and an open corrugation, were selected as offering the most potential for system success. Fourteen subsize heat shields of a full scale section were fabricated from C-129Y and Cb-752 and silicide coated with R-512E. These subsize panels were subjected to a simulated flight profile representing temperature, local surface pressures, and applied pressure differential loads. All corrugated panels of both alloys sustained 100 cycles without structural or coating failure. All Cb-752/R-512E panels performed well with one panel being successfully repaired after 66 cycles and completing 100 cycles. As a result of this evaluating the Cb-752/R-512E system was selected for hardware application during the subsequent phases. In addition, the tee-stiffened configuration was selected for further development and application in Phase III. This selection was based on an overall assessment of relative weight, cost, and structural performance of the tee-stiffened and open corrugation TPS.

Black, W. E.↗