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At least 73 records · Page 4

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NASA NTRS1972

Results of a study of Mach number and Reynolds number effects on the lee side vortex flow field characteristics of an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees. Volume 2: Appendix

A listing of the data reduction computer program output for the analysis of the lee side vortex flow about an ogive-cylinder-frustum-cylinder is presented. A computed local flow property for the flow field location array at a specific body station and freestream condition is listed. Significant local flow properties are computed. The local circulation strengths in the radial and circumferential directions are analyzed.

Foley, J. E.↗

NASA NTRS1975

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 were obtained using an implicit finite-difference scheme and results are presented for hypersonic flow over spherically-blunted cone configurations at freestream conditions representative of entry into the atmosphere of Venus. These data are compared with solutions obtained using other methods of analysis.

Anderson, E. C.↗

NASA NTRS1976

Results of an investigation of jet plume effects on a 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 8 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82C), volume 1

The primary test objective was to define the base pressure environment of the first and second stage mated vehicle in a supersonic flow field from Mach 2.60 through 3.50 with simulated rocket engine exhaust plumes. The secondary objective was to obtain the pressure environment of the Orbiter at various vent port locations at these same freestream conditions. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed Orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of test equipment and tested configurations are shown.

Hawthorne, P. J.↗

NASA NTRS1976

Results of an investigation of jet plume effects on an 0.010-scale model (75-OTS) of the space shuttle integrated vehicle in the 9 x 7-foot leg of the NASA/Ames unitary wind tunnel (IA82B), volume 1

The base pressure environment was investigated for the first and second stage mated vehicle in a supersonic flow field from Mach 1.55 through 2.20 with simulated rocket engine exhaust plumes. The pressure environment was investigated for the orbiter at various vent port locations at these same freestream conditions. The Mach number environment around the base of the model with rocket plumes simulated was examined. Data were obtained at angles of attack from -4 deg through +4 deg at zero yaw, and at yaw angles from -4 deg through +4 deg at zero angle of attack, with rocket plume sizes varying from smaller than nominal to much greater than nominal. Failed orbiter engine data were also obtained. Elevon hinge moments and wing panel load data were obtained during all runs. Photographs of the tested configurations are shown.

Hawthorne, P. J.↗

NASA NTRS1978

Low speed test of a high-bypass-ratio propulsion system with an asymmetric inlet designed for a tilt-nacelle V/STOL airplane

A large scale model of a lift/cruise fan inlet designed for a tilt nacelle V/STOL airplane was tested with a high bypass ratio turbofan. Testing was conducted at low freestream velocities with inlet angles of attack ranging from 0 deg to 120 deg. The operating limits for the nacelle were found to be related to inlet boundary layer separation. Small separations originating in the inlet diffuser cause little or no performance degradation. However, at sufficiently severe freestream conditions the separation changes abruptly to a lip separation. This change is associated with a significant reduction in nacelle net thrust as well as a sharp increase in fan blade vibratory stresses. Consequently, the onset of lip separation is regarded as the nacelle operating limit. The test verified that the asymmetric inlet design will provide high performance and stable operation at the design forward speed and angle of attack conditions. At some of these, however, operation near the lower end of the design inlet airflow range is not feasible due to the occurrence of lip separation.

Syberg, J.↗

NASA NTRS1978

Effects of small angle of attack on the radiating viscous shock layer solutions for Jovian entry

Results are presented for the laminar and turbulent flows of a reacting mixture of hydrogen-helium in chemical equilibrium over blunted cones of large half angles at small angles of attack. The freestream conditions correspond to a typical Jovian entry trajectory point. Radiative heat flux is calculated by using a realistic step-function model for the absorption coefficient. An investigation is also made to determine whether a blunted cone at small angles of attack can be replaced by equivalent bodies at zero angle of attack for the purpose of calculating the shock shapes, surface pressure, skin friction and heating rates.

Kumar, A.↗

NASA NTRS1978

Space shuttle afterbody aerodynamics/plume simulation data summary

A series of parametric wind tunnel tests was conducted to provide a base for developing a simulation of afterbody/base aerodynamics for multibody/multibase rocket-powered vehicles (such as Space Shuttle) which use unheated air as the simulant gas in development wind tunnel tests. The tests described were parameterized on external configuration, nozzle internal configuration, base geometry, propulsion gas type, and freestream Mach number (0.5 to 3.5). The tests were conducted over a 4-year period. Presented in this report are the data and pertinent reference information necessary to perform an analysis which would lead to a simulation procedure. The type of data obtained during the tests described herein include model base afterbody, and nozzle internal surface static pressure distributions, model chamber pressure and temperature, and freestream conditions. Also included is a brief description of simulation procedures that were used by the Space Shuttle program.

Blackwell, K. L.↗

NASA NTRS1978

Convective heat flux in a laser-heated thruster

An analysis is performed to estimate the convective heating to the wall in a laser-heated thruster on the basis of a solution of the laminar boundary-layer equations with variable transport properties. A local similiarity approximation is used, and it is assumed that the gas phase is in equilibrium. For the thruster described by Wu (1976), the temperature and pressure distributions along the nozzle are obtained from the core calculation. The similarity solutions and heat flux are obtained from the freestream conditions of the boundary layer, in order to determine if it is necessary to couple the boundary losses directly to the core calculation. In addition, the effects of mass injection on the convective heat transfer across the boundary layer with large density-viscosity product gradient are examined.

Wu, P. K. S.↗

NASA NTRS1980

Flowfield calculations past an entry probe with ablated nose shape

The effects of ablated nose shapes on the flowfield solutions are studied, using a time-dependent finite-difference method developed by Kumar, et al. (1979). Solutions are obtained for the laminar flow of a radiating mixture of H-He in chemical equilibrium past a blunt axisymmetric body at zero angle of attack. The freestream conditions correspond to a point on a typical Jovian entry trajectory, and the initial probe shape is a 45-deg half-angle spherically blunted cone. It is found that as nose bluntness increases, the following occur: in the nose region, shock standoff distances and radiative heating rates increase substantially; surface pressure level increases, but convective heating rates decrease.

Kumar, A.↗

NASA NTRS1981

Forebody drag reduction

It is shown that the mechanisms of forebody drag reduction by means of either a spike or a forward-facing jet are similar, with the maximum achievable drag reduction being of the same order. Because the jet may be a relatively cool gas, however, the forward facing jet has the additional capability of reducing the aerodynamic heating that is so severe at high Mach numbers. By means of the correlation presented, jet ejection parameters may be chosen to achieve maximum permissible forebody drag reduction. The correlation method uses a momentum coefficient that characterizes jet efflux and freestream conditions.

Pamadi, B. N.↗

NASA NTRS1983

Hypersonic flows over biconics using a variable-effective-gamma, Parabolized-Navier-Stokes code

Hypersonic flows over straight and bent biconics are calculated for a range of freestream conditions in which the gas behind the shock is treated as either perfect or real. The Parabolized-Navier-Stokes (PNS) equations form the basis of the approximation scheme. Good comparisons with experimental data for pressures, forces and moments, heat transfer, and oil-flow patterns serve to validate the perfect-gas version of the code. Circumferential velocity vector plots further aid in the interpretation of leeside oil-flow patterns. A variable-effective-gamma (VEG) option is implemented for the real-gas calculations. Gamma, now defined as the ratio of enthalpy to internal energy, is determined from a locally valid linear relation in enthalpy and pressure at every mesh point which in turn is calculated from a benchmark equilibrium code. The VEG option is easily incorporated into a host PNS code because it uses the underlying perfect-gas structure of the code. Comparisons to experimental data for heat transfer and shock shape in high enthalpy air have been obtained using the VEG option.

Gnoffo, P. A.↗

NASA NTRS1984

Effect of geometry on airfoil icing characteristics

A droplet trajectory computer code is used to predict the water droplet impingement characteristics of several low- and medium-speed airfoils. The maximum impingement efficiency, total collection efficiency, and limits of impingement are analyzed as functions of the airfoil geometry and freestream conditions. The airfoil geometry is represented by leading edge radius, maximum thickness, maximum camber, and angle of attack. The analysis shows that the primary effects are an increase in maximum impingement efficiency with a decrease in leading edge radius, a reduction in total collection efficiency for thicker airfoils, and a change in the limits of impingement for airfoils of different maximum camber.

Bragg, M. B.↗

NASA NTRS1984

Flowfield computation of an aerobrake orbital transfer vehicle

Viscous real gas and ideal gas supersonic flowfields over the forebody of a aeroassisted orbital transfer vehicle are determined using a unsteady factored implicit algorithm. Air in chemical equilibrium is considered and its local thermodynamic properties are computed by an equilibrium composition method. Numerical solutions are obtained for both real and ideal gases at a Mach number of 30 and at angles of attack up to 20 degrees. Shock stand-off distances and surface pressure distributions are presented for the gas models with and without viscous effects. For the freestream conditions considered, viscous effects dominate the flow.

Balakrishnan, A.↗

NASA NTRS1985

Numerical solutions of Navier-Stokes equations for a Butler wing

The flow field is simulated on the surface of a given delta wing (Butler wing) at zero incident in a uniform stream. The simulation is done by integrating a set of flow field equations. This set of equations governs the unsteady, viscous, compressible, heat conducting flow of an ideal gas. The equations are written in curvilinear coordinates so that the wing surface is represented accurately. These equations are solved by the finite difference method, and results obtained for high-speed freestream conditions are compared with theoretical and experimental results. In this study, the Navier-Stokes equations are solved numerically. These equations are unsteady, compressible, viscous, and three-dimensional without neglecting any terms. The time dependency of the governing equations allows the solution to progress naturally for an arbitrary initial initial guess to an asymptotic steady state, if one exists. The equations are transformed from physical coordinates to the computational coordinates, allowing the solution of the governing equations in a rectangular parallel-piped domain. The equations are solved by the MacCormack time-split technique which is vectorized and programmed to run on the CDC VPS 32 computer.

Abolhassani, J. S.↗

NASA NTRS1986

Direct simulation of hypersonic flows over blunt slender bodies

Results of a numerical study of low-density hypersonic flow about cylindrically blunted wedges and spherically blunted cones with body half angles of 0, 5, and 10 deg are presented. Most of the transitional flow regime encountered during entry between the free molecule and continuum regimes is simulated for a reentry velocity of 7.5 km/s by including freestream conditions of 70 to 100 km. The bodies are at zero angle of incidence and have diffuse and finite catalytic surfaces. Translational, thermodynamic, and chemical nonequilibrium effects are considered in the numerical simulation by utilizing the direct simulation Monte Carlo (DSMC) method. The numerical simulations show that noncontinuum effects such as surface temperature jump, and velocity slip are evident for all cases considered. The onset of chemical dissociation occurs at a simulated altitude of 96 km for the two-dimensional configurations. Comparisons between the DSMC and continuum viscous shock-layer calculations highlight the significant difference in flowfield structure predicted by the two methods.

Moss, J. N.↗

NASA NTRS1986

An efficient algorithm for solution of the unsteady transonic small-disturbance equation

A time accurate approximate factorization (AF) algorithm is formulated for solution of the three dimensional unsteady transonic small-disturbance equation. The AF algorithm consists of a time linearization procedure coupled with a Newton iteration technique. Superior stability characteristics of the new algorithm are demonstrated through applications to steady and oscillatory flows at subsonic and supersonic freestream conditions for an F-5 fighter wing. For steady flow calculations, the size of the time step is cycled to achieve rapid convergence. For unsteady flow calculations, the AF algorithm is sufficiently robust to allow the step size to be selected based on accuracy rather than on stability considerations. Therefore, accurate solutions are obtained in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods.

Batina, John T.↗

NASA NTRS1987

An efficient algorithm for solution of the unsteady transonic small-disturbance equation

A time accurate approximation factorization (AF) algorithm is formulated for solution of the three-dimensional unsteady transonic small-disturbance equation. The AF algorithm consists of a time linearization procedure coupled with a Newton iteration technique. Superior stability characteristics of the new algorithm are demonstrated through applications to steady and oscillatory flows at subsonic and supersonic freestream conditions for an F-5 fighter wing. For steady flow calculations, the size of the time step is cycled to achieve rapid convergence. For unsteady flow calculations, the AF algorithm is sufficiently robust to allow the step size to be selected based on accuracy rather than on stability considerations. Therefore, accurate solutions are obtained in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods.

Batina, John T.↗

NASA NTRS1987

Direct simulation of stagnation streamline flow for hypersonic reentry

The direct simulation Monte Carlo (DSMC) method is applied to simulate one-dimensional flow along the stagnation streamline. The freestream conditions considered are those encountered by the nose region of the Space Shuttle Orbiter during the hypersonic reentry. The range of altitudes (75 to 92 km) considered in the present calculations covers continuum to the less rarefied portion of the transition flow regime. The calculations account for thermal as well as the chemical nonequilibrium effects. The attention is focused on the flow structure along the stagnation streamline for different specified shock locations at a given altitude. The effect of shock location on the stagnation-point gas composition and heat transfer is analyzed for specific altitude; consequently, these results are appropriate for hemispherical bodies with different nose radii. Finally, comparison is made between the present DSMC results and the viscous shock-layer calculations, which show good agreement at the lowest altitude.

Dogra, Virendra K.↗