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At least 91 records · Page 5

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NASA NTRS1987

Navier-Stokes simulation of a hypersonic generic wing/fuselage

An unsteady thin-layer Navier-Stokes code is used to calculate a generic wing/fuselage configuration at a Mach number of 25 and freestream conditions corresponding to an altitude of 220,000 feet. Calculations were performed with the assumptions of a perfect gas and with chemical equilibrium, and the boundary layer was assumed to be turbulent and to have a surface temperature prescribed at 1255 K. Results for the two different gas assumptions were compared in terms of distributions of pressure, density, temperature, Mach number, ratio of specific heat, and heat transfer. Numerical problems arising in the calculations were identified.

Wai, John C.↗

NASA NTRS1988

Experimental investigation of the subsonic high-altitude operation of the NASA Lewis 10- by 10-foot supersonic wind tunnel

An experimental investigation was conducted in the NASA Lewis 10- by 10-Foot Supersonic Wind Tunnel during subsonic tunnel operation in the aerodynamic cycle to determine the test section flow characteristics near the Advanced Turboprop Project propeller model plane of rotation. The investigation used an eight-probe pitot static flow survey rake to measure total and static pressures at two locations in the wind tunnel: the test section and the bellmouth section (upstream of the two-dimensional flexible-wall nozzle). A cone angularity probe was used to measure any flow angularity in the test section. The evaluation was conducted at tunnel Mach numbers from 0.10 to 0.35 and at three operating altitudes from 2,000 to 50,000 ft. which correspond to tunnel reference total pressures from 1960 to 245 psfa, respectively. The results of this experimental investigation indicate a total-pressure loss area in the center of the test section and a static-pressure gradient from the test section centerline to the wall. These total and static pressure differences were observed at all tunnel operating altitudes and diminished at lower tunnel velocities. The total-pressure loss area was also found in the bellmouth section, which indicates that the loss mechanism is not the tunnel flexible-wall nozzle. The flow in the test section is essentially axial since very small flow angles were measured. The results also indicate that a correction to the tunnel total and static pressures must be applied in order to determine accurate freestream conditions at the test section centerline.

Hughes, Christopher E.↗

NASA NTRS1988

Nonequilibrium thermal radiation for an aeroassist flight experiment vehicle

The direct-simulation Monte Carlo method incorporating a dissociating and ionizing gas model for air with thermal radiation is used to characterize the hypersonic flow about an axisymmetric representation of an aeroassist flight experiment (AFE) vehicle, whose freestream conditions correspond to selected points along the entry, aerobraking, and exit phases of the trajectory. Calculations for two trajectory conditions indicate that the radiative heating of the AFE forebody is lower than the convective heating, but becomes significant as the maximum convective heating rate condition is approached.

Moss, James N.↗

NASA NTRS1988

Direct simulation of AFE forebody and wake flow with thermal radiation

Calculated results for the flowfield structure and surface quantities are presented for an axisymmetric representation of an aeroassist flight experiment vehicle. The direct simulation Monte Carlo (DSMC) method is used to perform the calculations, since the flow is highly nonequilibrium about the vehicle during both the compression and expansion phases. The body configuration is an elliptically blunt nose followed by a skirt with a circular radius and an afterbody. Freestream conditions correspond to a single point along the entry trajectory at an altitude of 90 km and a velocity of 9.9 km/s. The calculations account for nonequilibrium in the translational and internal modes, dissociation, ionization, and thermal radiation. The degree of dissociation is large, but the maximum ionization is only about 2 percent by mole fraction. The bluntforebody flow experiences a high degree of thermal nonequilibrium in which the translational temperature is generally greater than the internal temperature. However, as the flow expands about the aerobrake skirt and afterbody, the internal temperature is generally greater than the translational temperature.

Moss, James N.↗

NASA NTRS1988

Overview of CFD methods and comparison with flight aerothermal data

An overview of previously published aerothermal investigations which demonstrate the capabilities of detailed computational fluid dynamics and engineering codes to predict the aerothermal environment about an entry vehicle is presented. The overview consists of a brief discussion of the computational methods and experimental data and includes comparisons between the computed results and data. The overview focuses primarily on analyses of flight data since these data provide the unique capability to assess the real-gas chemistry options in the codes. The computed results are based on a series of codes which are employed by the Aerothermodynamics Branch of the Space System Division at the Langley Research Center. The flight data, which were measured on the Reentry F, the Space Shuttle, and the Fire II vehicles, represent a wide range of vehicle configurations and freestream conditions. Also, results of one recent set of ground tests are included since the tests provide data on a model of a pending flight project. The comparisons of the predicted results and data demonstrate the adequacy of the present computational fluid dynamics capabilities and indicate the potential to predict the aerothermal environment about future flight vehicles.

Sutton, Kenneth↗

NASA NTRS1989

High speed corner and gap-seal computations using an LU-SGS scheme

The hybrid Lower-Upper Symmetric Gauss-Seidel (LU-SGS) algorithm was added to a widely used series of 2D/3D Euler/Navier-Stokes solvers and was demonstrated for a particular class of high-speed flows. A limited study was conducted to compare the hybrid LU-SGS for approximate Newton iteration and diagonalized Beam-Warming (DBW) schemes on a work and convergence history basis. The hybrid LU-SGS algorithm is more efficient and easier to implement than the DBW scheme originally present in the code for the cases considered. The code was validated for the hypersonic flow through two mutually perpendicular flat plates and then used to investigate the flow field in and around a simplified scramjet module gap seal configuration. Due to the similarities, the gap seal flow was compared to hypersonic corner flow at the same freestream conditions and Reynolds number.

Coirier, William J.↗

NASA NTRS1989

Effect of nose bluntness on flow field over slender bodies in hypersonic flows

A parametric study has been conducted to determine the effects of nose bluntness on the enire flowfield over slender bodies under different hypersonic freestream conditions. The analysis is carried out for air under perfect- and equilibrium-gas assumptions. The analyses range from a few simplified approaches to the solution of the complete Navier-Stokes equations. Specific results obtained for spherically blunted cones and ogives demonstrate that there are significant differences in flowfield and surface quantities between sharp and blunted bodies. Depending upon the flow conditions and geometry, the differences are found to persist as far as 260 nose radii downstream.

Singh, D. J.↗

NASA NTRS1989

High speed corner and gap-seal computations using an LU-SGS scheme

Coirier, William J.↗

NASA NTRS1990

Numerical investigation of the origin of vortex asymmetry

The flow about a slender body of revolution placed at incidence to an oncoming stream is numerically investigated for angles of attack ranging from 20 to 80 degrees and a Reynolds number of 200,000, based on freestream conditions and maximum body diameter. Navier-Stokes computations are applied to investigate the phenomena governing the onset of vortex asymmetry. Time-accurate solutions are given for an ogive-cylinder body with and without a space-fixed time-invariant disturbance added near the tip. At a 40-degree angle of attack the flow is found to be steady but becomes highly asymmetric when the disturbance is added. The level of asymmetry is dependent upon the size and location of the disturbance. It is suggested that the origin of the asymmetry is a convective-type instability of the originally symmetric flow. For higher angles of attack (60 and 80 degrees) the flow around the cylindrical part becomes unsteady, and vortex shedding is observed. It is suggested that the origin of flow unsteadiness and vortex shedding in the wake is an absolute-type instability of the originally steady flow.

Degani, David↗

NASA NTRS1990

Numerical simulation of an F-16A at angle of attack

The transonic flowfield around an F-16A fighter configuration at a moderate incidence angle is simulated by solving the Navier-Stokes equations on a single-block grid. The numerical solution matches experimental freestream conditions with a mach number of 0.85, 16 degrees angle of attack, and a characteristic Reynolds number of 12.75 million. MacCormack's explicit algorithm is used in conjuction with a local time step and consecutive mesh refinement procedure to accelerate numerical convergence. The Baldwin-Lomax algebraic model provides turbulent closure. Computed surface pressure distributions and the aircraft lift coefficient compare favorably with wind tunnel data. The drag coefficient in the simulation overpredicts the experimental value by 8 percent.

Huband, G. W.↗

NASA NTRS1989

Direct simulation of AFE forebody and wake flow with thermal radiation

Calculated results for the flowfield structure and surface quantities are presented for an axisymmetric representation of an aeroassist flight experiment vehicle. The direct simulation Monte Carlo (DSMC) method is used to perform the calculations, since the flow is highly nonequilibrium about the vehicle during both the compression and expansion phases. The body configuration is an elliptically blunt nose followed by a skirt with a circular radius and an afterbody. Freestream conditions correspond to a single point along the entry trajectory at an altitude of 90 km and a velocity of 9.9 km/s. The calculations account for nonequilibrium in the translational and internal modes, dissociation, ionization, and thermal radiation. The degree of dissociation is large, but the maximum ionization is only about 2 percent by mole fraction. The blunt forebody flow experiences a high degree of thermal nonequilibrium in which the translational temperature is generally greater than the internal temperature. However, as the flow expands about the aerobrake skirt and afterbody, the internal temperature is generally greater than the translational temperature.

Moss, James N.↗

NASA NTRS1991

A hardened CARS system utilized for temperature measurements in a supersonic combustor

A coherent anti-Stokes Raman scattering (CARS) system has been hardened for utilization in a NASA Langley supersonic combustion test cell that can obtain temperature cross sections of the flow at three locations. This system is remotely operated and environmentally protected. Measurements were obtained in a scramjet combustor model consisting of a rearward-facing step, followed by an expansion duct. The freestream conditions were Mach 2, with static pressure that ranged from 0.8 to 1.9 atm, and a static temperature of about 800 K. Data for two different flow conditions were obtained that provided a comparison between reacting and nonreacting mixing of injected hydrogen fuel with the combustion-heated supersonic stream.

Antcliff, Richard R.↗

NASA NTRS1991

Unstructured Euler flow solutions using hexahedral cell refinement

An attempt is made to extend grid refinement into three dimensions by using unstructured hexahedral grids. The flow solver is developed using the TIGER (topologically Independent Grid, Euler Refinement) as the starting point. The program uses an unstructured hexahedral mesh and a modified version of the Jameson four-stage, finite-volume Runge-Kutta algorithm for integration of the Euler equations. The unstructured mesh allows for local refinement appropriate for each freestream condition, thereby concentrating mesh cells in the regions of greatest interest. This increases the computational efficiency because the refinement is not required to extend throughout the entire flow field.

Melton, John E.↗

NASA NTRS1990

DSMC calculations for the delta wing

Results are reported from three-dimensional direct simulation Monte Carlo (DSMC) computations, using a variable-hard-sphere molecular model, of hypersonic flow on a delta wing. The body-fitted grid is made up of deformed hexahedral cells divided into six tetrahedral subcells with well defined triangular faces; the simulation is carried out for 9000 time steps using 150,000 molecules. The uniform freestream conditions include M = 20.2, T = 13.32 K, rho = 0.00001729 kg/cu m, and T(wall) = 620 K, corresponding to lambda = 0.00153 m and Re = 14,000. The results are presented in graphs and briefly discussed. It is found that, as the flow expands supersonically around the leading edge, an attached leeside flow develops around the wing, and the near-surface density distribution has a maximum downstream from the stagnation point. Coefficients calculated include C(H) = 0.067, C(DP) = 0.178, C(DF) = 0.110, C(L) = 0.714, and C(D) = 1.089. The calculations required 56 h of CPU time on the NASA Langley Voyager CRAY-2 supercomputer.

Celenligil, M. Cevdet↗

NASA NTRS1991

Interactive design of hypersonic waverider geometries

The paper deals with an inverse design code utilizing the method of oscillating cones; the code integrated into an interactive graphics software package allows manipulation of both the exit-plane shock profile and leading edge of the vehicle. Another interactive feature of the system is the ability to vary freestream conditions and reevaluate the governing conditions. The development of the oscillating cones is shown on five classes each of which is chosen to demonstrate an aspect of improved design flexibility over previous studies. Results are evaluated using a robust flow solver, insuring that the shock shapes specified in the design process are recovered. It is pointed out that the expanded range of waverider geometries that may be generated using the oscillating cones technique may provide insight into visually oriented optimization parameters such as volumetric efficiency and practical planform.

Center, K. B.↗

NASA NTRS1991

Studies of aerothermal loads generated in regions of shock/shock interaction in hypersonic flow

Experimental studies were conducted to examine the aerothermal characteristics of shock/shock/boundary layer interaction regions generated by single and multiple incident shocks. The presented experimental studies were conducted over a Mach number range from 6 to 19 for a range of Reynolds numbers to obtain both laminar and turbulent interaction regions. Detailed heat transfer and pressure measurements were made for a range of interaction types and incident shock strengths over a transverse cylinder, with emphasis on the 3 and 4 type interaction regions. The measurements were compared with the simple Edney, Keyes, and Hains models for a range of interaction configurations and freestream conditions. The complex flowfields and aerothermal loads generated by multiple-shock impingement, while not generating as large peak loads, provide important test cases for code prediction. The detailed heat transfer and pressure measurements proved a good basis for evaluating the accuracy of simple prediction methods and detailed numerical solutions for laminar and transitional regions or shock/shock interactions.

Holden, Michael S.↗

NASA NTRS1992

Monte Carlo simulation of a near-continuum shock-shock interaction problem

A complex shock interaction is calculated with direct simulation Monte Carlo (DSMC). The calculation is performed for the near-continuum flow produced when an incident shock impinges on the bow shock of a 0.1 in. radius cowl lip for freestream conditions of approximately Mach 15 and 35 km altitude. Solutions are presented both for a full finite-rate chemistry calculation and for a case with chemical reactions suppressed. In each case, both the undisturbed flow about the cowl lip and the full shock interaction flowfields are calculated. Good agreement has been obtained between the no-chemistry simulation of the undisturbed flow and a perfect gas solution obtained with the viscous shock-layer method. Large differences in calculated surface properties when different chemical models are used demonstrate the necessity of adequately representing the chemistry when making surface property predictions. Preliminary grid refinement studies make it possible to estimate the accuracy of the solutions.

Carlson, Ann B.↗

NASA NTRS1991

Application of the DSMC method to hypersonic flow about a delta wing

The direct simulation Monte Carlo method was applied to a hypersonic rarefied flow about a delta wing at angle of incidence. Three-dimensional calculations were performed using a single-species gas model that accounts for energy exchange with the internal modes. Computed results were compared with the available experimental data obtained in a nitrogen wind-tunnel experiment. Discrepancies were shown to exist between the experimental and calculated results. These differences may be caused by the inadequately defined freestream conditions.

Celenligil, M. C.↗