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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Viscous shock-layer analysis of hypersonic flows over long slender bodies

Numerical solutions from the time-steady viscous shock-layer equations are presented for the hypersonic laminar and turbulent flow of a perfect gas over long slender bodies. These results are obtained from a spatial-marching implicit finite-difference technique. Detailed comparisons have been made with other predictions and experimental data to assess the accuracy of the present numerical technique, especially for slender-body flows. Results from the present method show that coupling the normal momentum and continuity equations and the use of the Vigneron pressure condition in the subsonic nose region give quite accurate and stable results. These results compare (with some exceptions) favorably with those obtained using PNS and other VSL methods.

Gupta, R. N.↗
NASA NTRSDate not supplied

Towards a Third-Order Accurate, Second-Derivative-Free, Shock-Capturing Finite-Volume Method for Hypersonic Flows on Tetrahedral Grids

In this paper, we report progress in the development of a third-order accurate, second-derivative-free, shock-capturing finite-volume solver for three-dimensional unstructured grids. The method is economical in the sense that the computation and storage of second derivatives are not required for third-order accuracy. It is based on point-valued numerical solutions stored at cells, gradients computed and stored at nodes, and an efficient projected-derivative formula that eliminates the need for second derivatives in a quadratic solution interpolation. The projected-derivative formula is also used to eliminate second derivatives from a high-order flux quadrature formula, so that it can be implemented conveniently in the form of a numerical flux at a face center plus a correction term. Similarly, a high-order source quadrature formula can also be implemented in the form of a cell-center point evaluation plus a similar correction term. These features make it relatively straightforward to extend an existing second-order finite-volume code to third-order. This paper reports progress of implementing the method in the NASA VULCAN-CFD code and discusses the implementation of a high-order accurate limiter for shock capturing.

Weighted Least-Squares↗
NASA NTRS2003

Portable Fluorescence Imaging System for Hypersonic Flow Facilities

A portable fluorescence imaging system has been developed for use in NASA Langley s hypersonic wind tunnels. The system has been applied to a small-scale free jet flow. Two-dimensional images were taken of the flow out of a nozzle into a low-pressure test section using the portable planar laser-induced fluorescence system. Images were taken from the center of the jet at various test section pressures, showing the formation of a barrel shock at low pressures, transitioning to a turbulent jet at high pressures. A spanwise scan through the jet at constant pressure reveals the three-dimensional structure of the flow. Future capabilities of the system for making measurements in large-scale hypersonic wind tunnel facilities are discussed.

Wilkes, J. A.↗
NASA NTRS1988

On the instability of hypersonic flow past a wedge

The instability of a compressible flow past a wedge is investigated in the hypersonic limit. Particular attention is given to the Tollmien-Schlichting waves governed by triple-deck theory though some discussion of inviscid modes is given. It is shown that the attached shock has a significant effect on the growth rates of Tollmien-Schlichting waves. Moreover, the presence of the shock allows for more than one unstable Tollmien-Schlichting wave. Indeed, an infinite discrete spectrum of unstable waves is induced by the shock, but these modes are unstable over relatively small but high frequency ranges. The shock is shown to have little effect on the inviscid modes considered by previous authors and an asymptotic description of inviscid modes in the hypersonic limit is given.

Cowley, Stephen↗
NASA NTRS1990

On the instability of hypersonic flow past a wedge

The instability of a compressible flow past a wedge is investigated in the hypersonic limit. Particular attention is paid to Tollmien-Schlichting waves governed by triple-deck theory though some discussion of inviscid modes is given. It is shown that the attached shock can have a significant effect on the growth rates of Tollmien-Schlichting waves. Moreover, the presence of the shock allows for more than one unstable Tollmien-Schlichting wave. Indeed an infinite discrete spectrum of unstable waves is induced by the shock, but these modes are unstable over relatively small but high frequency ranges. The shock is shown to have little effect on the inviscid modes considered by previous authors and an asymptotic description of inviscid modes in the hypersonic limit is given.

Cowley, Stephen↗
NASA NTRS1992

Hypersonic flow separation in shock wave boundary layer interactions

An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed.

Hamed, A.↗
NASA NTRS1989

The computation of thermo-chemical nonequilibrium hypersonic flows

Several conceptual designs for vehicles that would fly in the atmosphere at hypersonic speeds have been developed recently. For the proposed flight conditions the air in the shock layer that envelops the body is at a sufficiently high temperature to cause chemical reaction, vibrational excitation, and ionization. However, these processes occur at finite rates which, when coupled with large convection speeds, cause the gas to be removed from thermo-chemical equilibrium. This non-ideal behavior affects the aerothermal loading on the vehicle and has ramifications in its design. A numerical method to solve the equations that describe these types of flows in 2-D was developed. The state of the gas is represented with seven chemical species, a separate vibrational temperature for each diatomic species, an electron translational temperature, and a mass-average translational-rotational temperature for the heavy particles. The equations for this gas model are solved numerically in a fully coupled fashion using an implicit finite volume time-marching technique. Gauss-Seidel line-relaxation is used to reduce the cost of the solution and flux-dependent differencing is employed to maintain stability. The numerical method was tested against several experiments. The calculated bow shock wave detachment on a sphere and two cones was compared to those measured in ground testing facilities. The computed peak electron number density on a sphere-cone was compared to that measured in a flight test. In each case the results from the numerical method were in excellent agreement with experiment. The technique was used to predict the aerothermal loads on an Aeroassisted Orbital Transfer Vehicle including radiative heating. These results indicate that the current physical model of high temperature air is appropriate and that the numerical algorithm is capable of treating this class of flows.

Candler, Graham↗
NASA NTRS2007

Fluorescence Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Most of the models did not survive repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2-inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various configurations were studied including different sting placements relative to the models, different model orientations and attachment angles, and different NO seeding methods. The angle of attack of the models was also varied and the location of the laser sheet was scanned to provide three-dimensional flowfield information. Virtual Diagnostics Interface technology, developed at NASA Langley, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images. Lessons learned and recommendations for future experiments are discussed.

Alderfer, D. W.↗
NASA NTRS2008

Fluorescence Imaging and Streamline Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.

Danehy, Paul M.↗
NASA NTRSDate not supplied

Investigation of Improved Wall Heat Flux Behavior Observed When Computing Hypersonic Flow Using a Face-averaged Nodal-gradient Approach On Tetrahedral Grids

We investigate the origin of a significant reduction in computed numerical “noise” in the spatial variation of computed wall heat flux observed when using a node-centered weighted least squares(WLSQ) gradient approach when solving the Reynolds-averaged Navier-Stokes (RANS) equations for hypersonic boundary layer flows on highly stretched tetrahedral grids. This investigation is conducted using a 2nd-order, cell-centered, finite-volume, discretization which we employ to contrast results obtained using two WLSQ gradient approaches: a conventional 3-D, cell-centered, node-neighbor, WLSQ gradient (NN-CCG) approach and a 3-D, face-averaged, node-centered gradient (F-ANG)approach. The F-ANG approach, is shown to produce significantly less “noise” in the spatial variation of wall heat flux, than the NN-CCG approach and we demonstrate the cause of this “noise” via a series of numerical experiments where we compute hypersonic laminar and turbulent boundary layer flows.We further investigate the level of grid resolution required for a tetrahedral grid to approach the fidelity of a grid resolved hexahedral grid solution obtained using an equivalent structured grid method.

Jeffery A. White↗
NASA NTRS1987

Three dimensional hypersonic flow simulations with the CSCM implicit upwind Navier-Stokes method

The Bardina and Lombard (1985) three-dimensional CSCM Navier-Stokes method is presently extended to the simulation of complex hypersonic reentry vehicle external flows at angle of attack. The robust stability of the method derives from the combination of conservative implicit upwind flux difference splitting with a three-dimensional diagonally-dominant approximate factorization and relaxation scheme and characteristic-based implicit boundary approximations. The method's efficiency derives from an implicit symmetric Gauss-Seidel 'method of planes' relaxation scheme with alternating directional space marching sweeps along the flow coordinate direction.

Bardina, Jorge↗
NASA NTRS1990

Transition to turbulence in hypersonic flow

An outline of the project and recent progress toward the study of transition in hypersonic boundary layers is given. Aspects of the numerical method and the results of test computations are presented. At present the laminar flows over a flat plate have been computed for M less than or equal to 5.

Vandervegt, J. J. W.↗
NASA NTRS1977

Spatially resolved excitation temperature measurements in a hypersonic flow using the hook method

Following an analysis of the hook equations for variation of refractivity along a spectrograph slit a series of experiments is reported, in which CO2 was used as the test gas, and the flow variations were spatially resolved. The hooked spectra were used to measure the ground state number densities of the impurity atom of CrI present in the flow. The excitation temperature of the CrI 1-eV level was obtained from the separation of the hooks of the CrI 425.4-nm transition resonance line and simultaneously the 434.4-nm CrI eV lower level line, from the shock to the body of the hypersonic shock tunnel flow. This excitation temperature is compared with theoretical calculations of the CO2 gas temperature. Some further applications of the hook type methods to gas diagnostics are discussed.

Sandeman, R. J.↗
NASA NTRSDate not supplied

Dust Particle Aeroheating Calculations for Mars Entry Hypersonic Flows

The purpose of the current research effort is to predict particle trajectories in a hypersonic flowfield while accounting for all relevant particle-fluid and particle-particle interactions. The resulting particle solver is titled DUst Simulation & Tracking (DUST) and works in conjunction with the US3D CFD flow solver. The key elements underpinning the current work are an efficient mesh-localiztion algorithm, time-driven hard-sphere model, point-to-point MPI framework, coarse-graining using computational parcels, and high-order Adams-Bashforth time stepping. The coupled US3D-DUST framework will be applied to conduct a multi-physics examination of dust-laden flows around the the Schiarparelli capsule.

Dusty Flows↗
NASA NTRSDate not supplied

Dust Particle Aeroheating Calculations for Mars Entry Hypersonic Flows

The purpose of the current research effort is to predict particle trajectories in a hypersonic flowfield while accounting for all relevant particle-fluid and particle-particle interactions. The resulting particle solver is titled DUst Simulation & Tracking (DUST) and works in conjunction with the US3D CFD flow solver. The key elements underpinning the current work are an efficient mesh-localiztion algorithm, time-driven hard-sphere model, point-to-point MPI framework, coarse-graining using computational parcels, and high-order Adams-Bashforth time stepping. The coupled US3D-DUST framework will be applied to conduct a multi-physics examination of dust-laden flows around the the Mars 2020 capsule.

Dusty Flows↗
NASA NTRSDate not supplied

A Comparison of NO Laser-Induced Fluorescence Models at Conditions Relevant to Supersonic and Hypersonic Flows

Planar laser-induced fluorescence (PLIF) of the nitric oxide (NO) molecule has been widely used in wind tunnel facilities for flow visualization, velocity, and temperature measurements. The experimental PLIF measurements are often compared with synthetic PLIF images using computationally derived temperatures, pressures, velocities, and species mole fractions. This approach is commonly referred to as computational flow imaging (CFI). In the present work, we compare signal intensity from PLIF models with experimental PLIF measurements obtained within a low pressure gas cell system at pressures and NO mole fractions relevant to supersonic and hypersonic flowfields. Experimental measurements were compared to several different laser induced-fluorescence models reported in the literature including LIFBASE, LINUS, and a NASA two-level model. The experimental measurements agreed well with all of the models at lower pressures and lower NO mole fractions; the fluorescence there is linear with both of these parameters. However, at higher pressures and mole fractions, the signal becomes nonlinear with respect to these parameters as self-quenching limits the signal and absorption further limits the signal. In fact, for the experimental path length of the experiment, the combination of high pressure and high NO mole fraction causes the experimental results to deviate significantly from the predicted results that neglect absorption of the incident laser sheet. The LINUS model, which allows absorption to be calculated, provided results that agreed better with the experimental measurements. Since supersonic and hypersonic flowfields may contain a region of the flow with high pressures and measurements in large-scale facilities often include a long path length, neglecting absorption may have a significantly negative effect on the CFI comparison to experimental PLIF images. As a result, PLIF models that account for absorption should be included in computational flow imaging approaches for laser induced fluorescence.

laser-induced fluorescence↗