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SHOCK STRUCTURE AND THE LEADING EDGE PROBLEM
Rarefied gas dynamics - low density flow near leading edge of flat plate in hypersonic stream
Fluctuating surface pressure characteristics on slender cones in subsonic, supersonic, and hypersonic Mach-number flow
A program is presented to determine the flow-induced unsteady aerodynamic loads on space-shuttle orbiter configurations in the early reentry phase. Experiments on 8 degree half-angle cone/flat-base configurations were conducted at free-stream Mach numbers of 6 and 15 and Reynolds numbers corresponding to shuttle reentry. Data are available at free-stream Mach-numbers of 0.67, 2.5, 3, 4, 6, 15, and 22. It is shown that empirical expressions could be developed for the Mach-number dependence of the overall fluctuation pressure under an attached turbulent boundary layer and under a separated base-flow region.
A similarity solution of the Navier-Stokes equations with wall catalysis and slip for hypersonic, low Reynolds number flow over spheres
The slip conditions for a multicomponent mixture with diffusion, wall-catalyzed atom recombination and thermal radiation are derived. The more realistic multicomponent species slip conditions are shown to be necessary for accurate merged shock layer solutions on a sphere. These slip conditions are used in a first-order similarity solution of the Navier-Stokes equations with nonequilibrium chemistry for the merged shock layer. Results of this quick numerical solution are compared with a time dependent solution around the sphere and with measured arc jet results at low Reynolds numbers. The similarity solution, unlike the time dependent solution, shows smooth radial profiles of the pressure and smooth variations of velocity slip, skin friction, temperature slip and heat transfer around the body. The present first-order similarity solution is valid up to 25 deg from the stagnation point and takes less than 1% of the computer time to run a time dependent scheme. The smaller stand-off distance obtained from the similarity solution is supported by experimental data. The measured heat flux is closer to the similarity solution than the time dependent method at the stagnation point and shows the proper variation with circumferential angle up to at least 40 deg.
Hypersonic merged-layer flow on a sphere
With the objective to achieve the desired reduction in velocity by aerodynamic forces, an aeroassisted orbital transfer vehicle (AOTV) is expected to fly in the higher region of the atmosphere for a sustained period of time. This aeroassist maneuver will occur in the transitional regime, if the vehicle is designed for a low ballistic coefficient. In this case, a merged-layer (ML) is formed around the space vehicle. The flow characteristics of the vehicle can be studied by making use of the full Navier-Stokes (NS) equations, taking into account slip and temperature jump conditions. The present investigation has the objective to obtain a two-term series solution of the full Navier-Stokes equations with surface slip and temperature jump conditions for the ML flow on a sphere. Attention is given to the mathematical formulation of the problem, the numerical method of integration, and the results.
Hypersonic merged-layer flow on a sphere
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Rarefaction effects for hypersonic re-entry flow
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Recent improvements in the nonequilibrium VSL scheme for hypersonic blunt-body flows
The nonequilibrium viscous shock-layer (VSL) solution scheme is revisited to improve its solution accuracy in the stagnation point region and also to minimize and control the errors in the conservation of elemental mass. The stagnation-point solution is improved by using a second-order expansion for the normal velocity and the elemental mass conservation is improved by directly imposing the element conservation equations as solution constraints. These modifications are such that the general structure and computational efficiency of the nonequilibrium VSL scheme is not affected. This revised nonequilibrium VSL scheme is used to study the Mach 20 flow over a 7-deg sphere-cone vehicle under zero and 20-deg angle-of-attack conditons. Comparisons are made with the corresponding predictions of Navier-Stokes and Parabolized Navier-Stokes solution schemes. The results of these tests show that the nonequilibrium blunt-body VSL scheme is indeed an accurate, fast and extremely efficient means for generating the blunt-body flowfield over spherical nosetips at small-to-large angles of attack.
Computation of laminar hypersonic compression-corner flows
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Laser-spectroscopic measurement techniques for hypersonic, turbulent wind tunnel flows
A review is given of the nature, present status, and capabilities of two laser spectroscopic methods for the simultaneous measurement of temperature, density, and their fluctuations owing to turbulence in high speed wind tunnel flows. One method is based on the two frequency excitation of nitric oxide seeded into a nitrogen flow, using tunable dye lasers. The second, more recent method relies on the excitation of oxygen in air flows using a tunable, ArF excimer laser. Signal are obtained from both the laser induced fluorescence and from Raman scattering of the same laser pulse. Measurements are demonstrated in the turbulent boundary layer of a Mach-2 channel flow.
A system to measure flow moisture content in hypersonic wind tunnels
The technique and equipment is described which is used for obtaining data on the moisture content in two NASA Langley Hypersonic Wind Tunnels. A detailed description of the sampling system and its operation is presented along with the moisture analyzer used. The procedure used for converting dew point to parts of water per million by volume (ppmv) is included with graphs that show tunnel moisture content at various pressures.
Numerical study of the flow establishment time in hypersonic shock tunnels
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Equations, Tables, and Figures for Use in the Analysis of Helium Flow at Supersonic and Hypersonic Speeds
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Computational Aerothermal Modeling and Simulation of Hypersonic, Low-density Flows for the ADEPT Drag Modulated Aerocapture Mission
Drag Modulated Aerocapture (DMA) for Venus orbiter missions relies on the deployable aeroshell ADEPT to perform orbit insertion Delta-V
Apparatus and method for generating large mass flow of high temperature air at hypersonic speeds
High temperature, high mass air flow and a high Reynolds number test air flow in the Mach number 8-10 regime of adequate test flow duration is attained by pressurizing a ceramic-lined storage tank with air to a pressure of about 100 to 200 atmospheres. The air is heated to temperatures of 7,000 to 8,000 R prior to introduction into the tank by passing the air over an electric arc heater means. The air cools to 5,500 to 6,000 R while in the tank. A decomposable gas such as nitrous oxide or a combustible gas such as propane is injected into the tank after pressurization and the heated pressurized air in the tank is rapidly released through a Mach number 8-10 nozzle. The injected gas medium upon contact with the heated pressurized air effects an exothermic reaction which maintains the pressure and temperature of the pressurized air during the rapid release.
A survey of leeside flow and heat transfer on delta planform configurations
The dominant feature of hypersonic leeside flow fields is the presence of coiled vortices which are generated as the leeside boundary layer responds to the inviscid flow field. Depending on geometric and stream parameters, the vortices can be completely maintained within the boundary layer or can lie outside which is the more familiar case of rolled-up vortices associated with separated flow. Three main areas are covered: (1) fundamental leeside flow phenomena associated with inviscid flows, separation and vortex phenomena involving viscid-inviscid interactions on sharp-edged delta wings and cones; (2) the application of these fundamental concepts to delta-wing bodies such as the space shuttle; and (3) the opportunity and a proposed approach to use early shuttle orbiter entry flights to obtain leeside data.
Viscous shock layer solutions for turbulent flow of radiating gas mixtures in chemical equilibrium
The viscous shock layer equations for hypersonic laminar and turbulent flows of radiating or nonradiating gas mixtures in chemical equilibrium are presented for two-dimensional and axially symmetric flow fields. Solutions are obtained using an implicit finite difference scheme and results are presented for hypersonic flow over spherically blunted cone configurations at free stream conditions representative of entry into the atmosphere of Venus. These data are compared with solutions obtained using other methods of analysis.
An Analysis of the Applicability of the Hypersonic Similarity Law to the Study of Flow About Bodies of Revolution at Zero Angle of Attack
The hypersonic similarity law as derived by Tsien has been investigated by comparing the pressure distributions along bodies of revolution at zero angle of attack. In making these comparisons, particular attention was given to determining the limits of Mach number and fineness ratio for which the similarity law applies. For the purpose of this investigation, pressure distributions determined by the method of characteristics for ogive cylinders for values of Mach numbers and fineness ratios varying from 1.5 to 12 were compared. Pressures on various cones and on cone cylinders were also compared in this study. The pressure distributions presented demonstrate that the hypersonic similarity law is applicable over a wider range of values of Mach numbers and fineness ratios than might be expected from the assumptions made in the derivation. This is significant since within the range of applicability of the law a single pressure distribution exists for all similarly shaped bodies for which the ratio of free-stream Mach number to fineness ratio is constant. Charts are presented for rapid determination of pressure distributions over ogive cylinders for any combination of Mach number and fineness ratio within defined limits.