Parametric analysis of radiative structure in aerobrake shock layers
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Greendyke, Robert B.↗
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Enhancements to viscous-shock-layer technique
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Gupta, Roop N.↗
Computation of nonequilibrium shock layers for the comet sample return vehicles
This paper contains computational results for the flow around the body of the 'Rosetta' spacecraft. The flow is computed for the predicted trajectory point of maximum heating, using a Navier-Stokes solver coupled to the chemical kinetics. The calculations are performed for catalytic and non-catalytic boundary conditions at the wall, and with a systematic grid sensitivity study. The calculations verify some of the assumptions made in an early study, and predict the location of peak convective heating.
Cambier, Jean-Luc↗
Spectral measurements of shock layer radiation in an arc-jet wind tunnel
Measurements were made of the radiating gas cap of a blunt body in an NASA Ames 20 MW arcjet wind tunnel. The test gas was air. Spectra of the flux incident on a small aperture centered at the stagnation region were obtained. A helium-cooled MgF window transmitted flux into an evacuated collimating system that focused the aperture onto the entrance slit of a spectrometer. Data were obtained with films and by photomultipliers. The range covered was 120 nm to 1000 nm and the resolution was 0.05 nm to 0.5 nm. This paper presents preliminary results from the experiment. Representative spectral records from 200 nm to 1000 nm are shown. The spectra show the atomic lines from oxygen and nitrogen in the IR, as well as the molecular systems of NO, N2, N2(+), and CN. Copper, as a contaminant, and carbon are tentatively identified.
Palumbo, Giuseppe↗
VUV shock layer radiation in an arc-jet wind tunnel experiment
Measurements were made of the radiating gas cap of a blunt body in a NASA-Ames 20 MW arc-jet wind tunnel. The test gas was air. Spectra of the flux incident on a small aperture centered at the stagnation region were obtained. A helium-cooled, magnesium fluoride window transmitted the flux into an evacuated collimating system that focused the aperture onto the entrance slit of a spectrometer. Data were obtained with films and by photomultipliers. The spectral ranges covered were the vacuum ultraviolet, VUV, (120 nm to 200 nm) and the ultraviolet to near infrared (200 nm to 900 nm) with resolutions from 0.05 nm to 0.5 nm. This paper presents the preliminary VUV results from the experiment. Results from the 200 nm to 900 nm spectral range have been presented elsewhere. Representative spectral records from 120 nm to 200 nm are shown. The intense atomic oxygen and nitrogen lines which are of concern to hypersonic flight are measured. Carbon lines are are also seen. These results will be used to help develop and validate aerothermodynamic computational models of arc-jet wind tunnel performance and help to assess the importance of VUV heating to entering spacecraft.
Craig, Roger A.↗
Validation of High-Speed Turbulent Boundary Layer and Shock-Boundary Layer Interaction Computations with the OVERFLOW Code
The capability of the OVERFLOW code to accurately compute high-speed turbulent boundary layers and turbulent shock-boundary layer interactions is being evaluated. Configurations being investigated include a Mach 2.87 flat plate to compare experimental velocity profiles and boundary layer growth, a Mach 6 flat plate to compare experimental surface heat transfer,a direct numerical simulation (DNS) at Mach 2.25 for turbulent quantities, and several Mach 3 compression ramps to compare computations of shock-boundary layer interactions to experimental laser doppler velocimetry (LDV) data and hot-wire data. The present paper describes outlines the study and presents preliminary results for two of the flat plate cases and two small-angle compression corner test cases.
Oliver, A. B.↗
Numerical computation of three-dimensional blunt body flow fields with an impinging shock
A time-marching finite-difference method was used to solve the compressible Navier-Stokes equations for the three-dimensional wing-leading-edge shock impingement problem. The bow shock was treated as a discontinuity across which the exact shock jump conditions were applied. All interior shock layer detail such as shear layers, shock waves, jets, and the wall boundary layer were automatically captured in the solution. The impinging shock was introduced by discontinuously changing the freestream conditions across the intersection line at the bow shock. A special storage-saving procedure for sweeping through the finite-difference mesh was developed which reduces the required amount of computer storage by at least a factor of two without sacrificing the execution time. Numerical results are presented for infinite cylinder blunt body cases as well as the three-dimensional shock impingement case. The numerical results are compared with existing experimental and theoretical results.
Holst, T. L.↗
Numerical computation of two-dimensional viscous blunt body flows with an impinging shock
Two-dimensional, viscous, blunt body flows with an impinging shock wave are computed using a time-dependent, finite-difference method to solve the complete set of Navier-Stokes equations. The bow shock wave is treated as a discontinuity, while all interior shock layer detail such as shear layers, shock waves, jets, and the wall boundary layer are automatically captured in the solution. Numerical results are presented for cases in which shock waves of different strengths are allowed to impinge on the flow field surrounding a circular cylinder resulting in different shock interference patterns. The two-dimensional results are compared qualitatively with existing three-dimensional experiments.
Tannehil, J. C.↗
Numerical computation of viscous blunt body flows with a planar impinging shock
Two- and three-dimensional, viscous blunt body flows with planar impinging shocks are computed using an explicit, time-dependent, finite-difference method to solve the complete set of Navier-Stokes equations. The bow shock is treated as a discontinuity, while all interior shock layer detail such as shear layers, shock waves, jets and the wall boundary layer are automatically captured in the solution. Numerical results are presented for cases in which planar shock waves of different strengths and orientations are allowed to impinge on the flow field surroundings an infinite cylinder resulting in two- and three-dimensional shock interference patterns. The numerical results are compared with experiment.
Holst, T. L.↗
Numerical solution of the interaction of a shock wave with a laminar boundary layer
Shock wave interaction with laminar boundary layer on flat plate using modified Lax-Wendroff difference technique
Maccormack, R. W.↗
An experimental study of massive blowing from a nosetip during Jovian entry
The fluid dynamic structure and stability of the boundary layer and shock layer ahead of a highly blowing heat shield in an environment simulating entry into the Jovian atmosphere were investigated. The experimental studies were conducted in the Calspan 96 in shock tunnel with a freestream of a H2/He mixture at Mach numbers above 10. Measurements of heat transfer and pressure were made on the porous model surface, and high-speed schlieren photography was used to examine the structure and stability of the shock layer. Results are presented which show how the heat transfer and shock layer stability vary for gases of different molecular weights for a range of Reynolds numbers and mass-addition rates.
Holden, M. S.↗
A Comparison of EAST Shock-Tube Radiation Measurements with a New Air Radiation Model
This paper presents a comparison between the recent EAST shock tube radiation measurements (Grinstead et al., AIAA 2008-1244) and the HARA radiation model. The equilibrium and nonequilibrium radiation measurements are studied for conditions relevant to lunar-return shock-layers; specifically shock velocities ranging from 9 to 11 kilometers per second at initial pressures of 0.1 and 0.3 Torr. The simulated shock-tube flow is assumed one-dimensional and is calculated using the LAURA code, while a detailed nonequilibrium radiation prediction is obtained in an uncoupled manner from the HARA code. The measured and predicted intensities are separated into several spectral ranges to isolate significant spectral features, mainly strong atomic line multiplets. The equations and physical data required for the prediction of these strong atomic lines are reviewed and their uncertainties identified. The 700-1020 nm wavelength range, which accounts for roughly 30% of the radiative flux to a peak-heating lunar return shock-layer, is studied in detail and the measurements and predictions are shown to agree within 15% in equilibrium. The plus or minus 1.5% uncertainty on the measured shock velocity is shown to cause up to a plus or minus 30% difference in the predicted radiation. This band of predictions contains the measured values in almost all cases. For the highly nonequilibrium 0.1 Torr cases, the nonequilibrium radiation peaks are under-predicted by about half. This under-prediction is considered acceptable when compared to the order-of-magnitude over-prediction obtained using a Boltzmann population of electronic states. The reasonable comparison in the nonequilibrium regions provides validation for both the non-Boltzmann modeling in HARA and the thermochemical nonequilibrium modeling in LAURA. The N2 (+)(1-) and N2(2+) molecular band systems are studied in the 290 480 nm wavelength range for both equilibrium and nonequilibrium regimes. The non-Boltzmann rate models for these systems, which have significant uncertainties, are tuned to improve the comparison with measurements.
Johnston, Christopher O.↗
Tangential blowing for control of strong normal shock - Boundary layer interactions on inlet ramps
The use of tangential blowing from a row of holes in an aft facing step is found to provide good control of the ramp boundary layer, normal shock interaction on a fixed geometry inlet over a wide range of inlet mass flow ratios. Ramp Mach numbers of 1.36 and 1.96 are investigated. The blowing geometry is found to have a significant effect on system performance at the highest Mach number. The use of high-temperature air in the blowing system, however, has only a slight effect on performance. The required blowing rates are significantly high for the most severe test conditions. In addition, the required blowing coefficient is found to be proportional to the normal shock pressure rise.
Schwendemann, M. F.↗
Influence of precursor heating on viscous flow around a Jovian entry body
The influence of changes in precursor region flow properties (resulting from the absorption of radiation from the shock layer) on the entire shock layer flow phenomena is investigated. Axially symmetric case is considered for both the preheating zone (precursor region) and shock layer. The flow in the shock layer is considered to be viscous with chemical equilibrium but radiative nonequilibrium. Realistic thermo-physical and spectral models are employed and results are obtained by implicit finite difference and iterative procedures. The results indicate that precursor heating increases the radiative heating of the body by a maximum of 7.5% for a 116 KM entry conditions.
Tiwari, S. N.↗
Influence of precursor heating on viscous flow around a Jovian entry body
The influence of changes in precursor region flow properties (resulting from the absorption of radiation from the shock layer) on the entire shock layer flow phenomena was investigated. The axially symmetric case is considered for both the preheating zone (precursor region) and shock layer. The flow in the shock layer is assumed to be viscous with chemical equilibrium but radiative nonequilibrium. Realistic thermophysical and spectral models are employed, and results are obtained by implicit finite difference and iterative procedures. The results indicate that precursor heating increases the radiative heating of the body by a maximum of 7.5 percent for 116 km entry conditions.
Tiwari, S. N.↗
Prediction of electron density in the boundary layer on entry vehicles with ablation
A rather general computer code has been developed for the numerical solution of the laminar boundary layer and thin shock layer (stagnation point only) equations for a multi-component gas mixture with finite reaction rates. The purpose of this paper is to indicate the capabilities of this computer program by presenting solutions for pure air flows and flows with carbon ablation. At the body surface the effects of oxidation and sublimation of carbon are taken into account with the following species included in the gas model: 0 2 , N 2 , O, N, NO, NO + , C 1 , C 2 , C 3 , CN, CO, and CO 2 . Results for the boundary layer along a 10° half-angle hyperboloid with ablation of carbon are compared to pure air results. Shock layer results for several wall temperatures and altitudes are presented and compared to pure air results. Other types of problems that can be solved with this computer code are indicated.
F. G. Blottner↗
An experimental study of three-dimensional shock wave/boundary layer interactions generated by sharp fins
The interaction between a turbulent boundary layer and a shock wave generated by a sharp fin with leading edge sweepback was investigated. The incoming flow was at Mach 2.96 and at a unit Reynolds number of 63 x 10 to the 6th power 0.1 m. The approximate incoming boundary layer thickness was either 4 mm or 17 mm. The fins used were at 5 deg, 9 deg and 15 deg incidence and had leading edge sweepback from 0 deg to 65 deg. The tests consisted of surface kerosene lampblack streak visualization, surface pressure measurements, shock wave shape determination by shadowgraphs, and localized vapor screen visualization. The upstream influence lengths of the fin interactions were correlated using viscous and inviscid flow parameters. The parameters affecting the surface features close to the fin and way from the fin were also identified. Essentially, the surface features in the farfield were found to be conical.
Lu, F. K.↗
Porosity effect on supercritical airfoil drag reduction by shock wave/boundary layer control
An investigation of the passive shock wave/boundary layer control for reducing the drag of 14 percent-thick supercritical airfoil was conducted in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel at transonic Mach numbers. Various porous surfaces with a cavity beneath it was positioned on the area of the airfoil, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a uniform porosity surface the normal shock wave for solid surface was changed to a lambda shock wave, and the wake impact pressure data indicated an appreciable drag reduction at transonic Mach numbers. For a free stream Mach number of 0.81 the profile drag coefficient for the airfoil top surface with uniform porosity was 46 percent lower than for the solid surface airfoil.
Nagamatsu, H. T.↗