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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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Spectroscopic experiments with gas-driven shock tube - technical developments, redundant measurements of plasma state, and atomic line strengths for neutral carbon
Shock tube tests of Langmuir probe for determining electron density and temperature surrounding reentry vehicle
Interaction of glow front in electrodynamic shock tube with transverse magnetic field
Spectroscopy of light atoms and ions and transition probability determinations using gas-driven shock tube
Production of carbide singlets during carbon compound decomposition in shock tube
Atmospheric argon effect on hypersonic stagnation point convective heat transfer, using arc-heated shock tube simulating flight velocities up to 34,000 fps
Radiation sources from high temperature equilibrium air noting measurements, equipment and results
Equilibrium air total radiation mechanism, vacuum UV radiation and relation to hypervelocity entry studied, using shock tube blunt model test flow
Thin film thermal radiation detector noting construction, calibration, performance, error sources and advantages of pyroelectric over thin film gauges
High temperature air plasma total radiant intensity measurements in shock tube
Ignition kinetics of carbon monoxide-oxygen reaction
JPL free piston shock tube design, calibration and performance
Drag measurements of cones in rarefied flow regime extended to higher and lower cone semivertex angles
Shock tube simulation of rocket exhaust plumes and effects on vehicles in space environment
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The design of a supersonic combustion engine testbed (SCET) aircraft is presented. The hypersonic waverider will utilize both supersonic combustion ramjet (SCRAMjet) and turbofan-ramjet engines. The waverider concept, system integration, electrical power, weight analysis, cockpit, landing skids, and configuration modeling are addressed in the configuration considerations. The subsonic, supersonic and hypersonic aerodynamics are presented along with the aerodynamic stability and landing analysis of the aircraft. The propulsion design considerations include: engine selection, turbofan ramjet inlets, SCRAMjet inlets and the SCRAMjet diffuser. The cooling requirements and system are covered along with the topics of materials and the hydrogen fuel tanks and insulation system. A cost analysis is presented and the appendices include: information about the subsonic wind tunnel test, shock expansion calculations, and an aerodynamic heat flux program.
Numerical results obtained with the direct simulation Monte Carlo (DSMC) method are presented for Mach 15.6 nitrogen flow about a 70-deg spherically blunted cone at zero incidence. This flow condition is one of several generated in the Large Energy National Shock (LENS) tunnel during tests of a 15.24 cm diameter model with an afterbody sting. The freestream Knudsen number, based on model diameter, is 0.0023. The focus of the DSMC calculations is to characterize the near wake flow under conditions where rarefaction effects may influence afterbody aerothermal loads. This report provides information concerning computational details along with flowfield and surface quantities. Calculations show that the flow enveloping the test model is in thermal nonequilibrium and a sizable vortex develops in the near wake. Along the model baseplane the heating rates are about 0.6 percent of the forebody stagnation value while the maximum heating along the sting is about 4.2 percent of the forebody stagnation value. Comparison of a Navier-Stokes solution with the present calculations show good agreement for surface heating, pressure, and skin friction results.
A series of experiments to measure pressure and heating for code validation involving hypersonic, laminar, separated flows was conducted at the Calspan-University at Buffalo Research Center (CUBRC) in the Large Energy National Shock (LENS) tunnel. The experimental data serves as a focus for a code validation session but are not available to the authors until the conclusion of this session. The first set of experiments considered here involve Mach 9.5 and Mach 11.3 N, flow over a hollow cylinder-flare with 30 deg flare angle at several Reynolds numbers sustaining laminar, separated flow. Truncated and extended flare configurations are considered. The second set of experiments, at similar conditions, involves flow over a sharp, double cone with fore-cone angle of 25 deg and aft-cone angle of 55 deg. Both sets of experiments involve 30 deg compressions. Location of the separation point in the numerical simulation is extremely sensitive to the level of grid refinement in the numerical predictions. The numerical simulations also show a significant influence of Reynolds number on extent of separation. Flow unsteadiness was easily introduced into the double cone simulations using aggressive relaxation parameters that normally promote convergence.