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Buning, P. G.

Publications and source records attributed to Buning, P. G..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Source Term Model for an Array of Vortex Generator Vanes

A source term model was developed for numerical simulations of an array of vortex generators. The source term models the side force created by a vortex generator being modeled. The model is obtained by introducing a side force to the momentum and energy equations that can adjust its strength automatically based on a local flow. The model was tested and calibrated by comparing data from numerical simulations and experiments of a single low-profile vortex generator vane, which is only a fraction of the boundary layer thickness, over a flat plate. The source term model allowed a grid reduction of about seventy percent when compared with the numerical simulations performed on a fully gridded vortex generator without adversely affecting the development and capture of the vortex created. The source term model was able to predict the shape and size of the stream wise vorticity and velocity contours very well when compared with both numerical simulations and experimental data.

Buning, P. G.↗
NASA NTRS1996

Computed Flow About The Integrated Space Shuttle, Revisited

Report discusses numerical simulations of flow of air about integrated space shuttle (complete space shuttle assembly including orbiter, solid rocket boosters, and external tank) in ascent. Updated version of report described in "Computed Flow About The Integrated Space Shuttle" (ARC-12685). Goal of studies to improve understanding of, and ability to predict how integrated space shuttle performs during both nominal and aborted ascent under various conditions.

Buning, P. G.↗
NASA NTRS1995

Inverse Design of Nacelles Using Multi-Block Navier Stokes Codes

The objective of this work is to reshape a nacelle to achieve a specified nacelle pressure distribution. The nacelle may be either isolated or installed on an airplane. There are no restrictions on the attitude (toe, incidence, and roll) and position of the nacelle. The design algorithm is coupled to two different multi-block 3-D Navier Stokes flow solvers. The coupling between design and analysis is automated to the point where the design proceeds with minimal user input.

Naik, D. A.↗
NASA NTRS1993

Numerical simulation of the Space Shuttle Launch Vehicle flowfield with real gas solid rocket plume effects

A numerical simulation of the external ascent flowfield of the Space Shuttle Launch Vehicle (SSLV) was carried out at the freestream Mach number 1.25, the angle of attack -5.1 deg, and the flight Reynolds number 3.25 x 10 exp 6/ft. The simulation is an extension of the solution by Kandula et al. (1991) and incorporates variable gamma effects with a high fidelity grid appropriate for a flight Reynolds number flow solution. Three-dimensional SSLV steady-state simulations with either perfect-gas or variable-gamma-gas Redesigned-Solid-Rocket-Motor (RSRM) plumes are computed on a 14-grid overlapping body-conforming grid system, and the influence of the RSRM exhaust plumes on the SSLV component pressure distributions and integrated loads is examined.

Slotnick, J. P.↗
NASA NTRS1993

Evaluation of Baldwin-Barth turbulence model with thin-layer Navier-Stokes computation of an axisymmetric afterbody-exhaust jet flowfield

Axisymmetric steady flowfield over a steep circular arc afterbody with a choked exhaust jet in a subsonic freestream is numerically investigated using overflow code, developed by NASA Ames. The accuracy of Baldwin-Barth one-equation turbulence model for the computation of separated flow and plume shear layers is assessed by a comparison of predictions with available measurements from NASA Langley. The results show that the Baldwin-Barth turbulence model performs better than the Baldwin-Lomax algebraic model in predicting the surface pressure distribution in the separated flow region of the afterbody. In the jet mixing layer, the Baldwin-Barth model considerably unpredicts the growth of the shear layer, as does the Baldwin-Lomax model. It is shown that the one-equation turbulence model needs improvement for the prediction of separated flow due to smooth adverse pressure gradient, and mixing layer growth.

Kandula, M.↗
NASA NTRS1993

Aerodynamics of powered missile separation from F/A-18 aircraft

A 3D dynamic 'chimera' algorithm that solves the thin-layer Navier-Stokes equations over multiple moving bodies was modified to numerically simulate the aerodynamics, missile dynamics, and missile plume interactions of a missile separating from a generic wing and from an F/A-18 aircraft in transonic flow. The missile is mounted below the wing for missile separation from the wing and on the F/A-18 fuselage at the engine inlet side for missile separation from aircraft. Static and powered missile separation cases are considered to examine the influence of the missile and plume on the wing and F/A-18 fuselage and engine inlet. The aircraft and missile are at two degrees angle of attack, Reynolds number of 10 million, freestream Mach number of 1.05 and plume Mach number of 3.0. The computational results show the details of the flow field.

Ahmad, J. U.↗
NASA NTRS1991

Computed Flow About The Integrated Space Shuttle

Report discusses numerical simulations of flow of air about integrated Space Shuttle in ascent. Goal: to improve understanding of, and ability to predict, how integrated Space Shuttle will perform during both nominal and aborted ascent under various conditions. These and other numerical simulations intended both to supplement wind-tunnel data, corrupted to some extent by scaling and wall-interference effects, and to compensate for scarcity of valid flight data.

Buning, P. G.↗
NASA NTRS1989

Numerical aerodynamic simulation of the space shuttle ascent environment

After the STS 51-L accident, an extensive review of the Space Shuttle Orbiter's ascent aerodynamic loads uncovered several questionable areas that required further analysis. The insight gained by comparing the Shuttle ascent CFD numerical simulations, obtained by the NASA Ames Space Shuttle Flow Simulation Group, to the current IVBC-3 aerodynamic loads database was instrumental in resolving uncertainties on the Orbiter payload bay doors and fuselage. Initial confidence in the numerical simulations was gained by comparing them with the limited flight data that had been obtained during the Orbiter Flight Test (OFT) program. Current CFD results exist for Mach numbers 0.6, 0.9, 1.05, 1.55, 2.0, and 2.5. Since the pre STS-1 wind tunnel test program (IA-105) often yields considerable differences when compared to STS-5 flight data, the M(sub infinity) = 1.05 transonic case is the most investigated. The IA308 mated-vehicle hot gas plume wind tunnel test, recently completed at AEDC 16T (transonic) and Lewis (hypersonic), is also used to compare with the computation where applicable.

Slotnick, Jeff P.↗
NASA NTRS1988

Numerical simulation of the integrated space shuttle vehicle in ascent

A simulation of the flow about the integrated space shuttle vehicle in ascent mode has been undertaken for various flight conditions using the Chimera composite grid discretization approach. Overset body-conforming grids were used to represent each geometric component, and an implicit approximately factored finite-difference procedure was used to solve the three-dimensional thin-layer Navier-Stokes equations. The computational results have been compared with both wind tunnel and flight test data. Although relatively good agreement is obtained with the experimental data, further refinement and evaluation of numerical error is under way.

Buning, P. G.↗
NASA NTRS1986

Navier-Stokes computations for exotic airfoils

A hyperbolic grid generator with improvements for handling sharp corners and concave surfaces combined with a Navier-Stokes flow solver is applied to complex two dimensional airfoil configurations. Steady separated flow about an iced leading edge of an airfoil is presented. Unsteady viscous separated flows past an airfoil at two angles of attack with a spoiler deployed at 60 deg are compared with experiment. The spoiler computations are performed with two different topological maps of the physical domain to the computational domain. Graphical techniques for the static and unsteady display of the flow fields are discussed.

Barth, T. J.↗
NASA NTRS1985

Accelerating an iterative process by explicit annihilation

A slowly convergent stationary iterative process can be accelerated by explicitly annihilating (i.e., eliminating) the dominant eigenvector component of the error. The dominant eigenvalue or complex pair of eigenvalues can be estimated from the solution during the iteration. The corresponding eigenvector or complex pair of eigenvectors can then be annihilated by applying an explicit Richardson process over the basic iterative method. This can be done entirely in real arithmetic by analytically combining the complex conjugate annihilation steps. The technique is applied to an implicit algorithm for the calculation of two dimensional steady transonic flow over a circular cylinder using the equations of compressible inviscid gas dynamics. This demonstrates the use of explicit annihilation on a nonlinear problem.

Jespersen, D. C.↗
NASA NTRS1985

Developments in the simulation of compressible inviscid and viscous flow on supercomputers

In anticipation of future supercomputers, finite difference codes are rapidly being extended to simulate three-dimensional compressible flow about complex configurations. Some of these developments are reviewed. The importance of computational flow visualization and diagnostic methods to three-dimensional flow simulation is also briefly discussed.

Steger, J. L.↗
NASA NTRS1985

Navier-Stokes computations for exotic airfoils

An efficient hyperbolic grid generator with improvements for handling sharp corners and concave surfaces is combined with an efficient and accurate Navier-Stokes flow solver. This combination is applied to some rather complex two-dimensional airfoil configurations. Steady separated flow about an iced leading edge of an airfoil is presented. Unsteady viscous separated flows past an airfoil at two angles of attack with a spoiler deployed at 60 degrees are presented and compared with experiment. The spoiler computations are performed with two different topological maps of the physical domain to the computational domain. Innovative graphical techniques for both the static and unsteady display of the flow fields are presented and discussed.

Barth, T. J.↗
NASA NTRS1985

Graphics and flow visualization in computational fluid dynamics

Techniques for displaying two- and three-dimensional flowfield solutions are described. Several methods of illustrating flow structure are addressed including particle tracing, simulated oil flow, and shock finding. These are incorporated into an interactive graphics program for CFD flowfields, called PLOT3D. Emphasis is made on the difficulty in visualizing three-dimensional flow features, and the importance of color, fast 3D image manipulation, and dynamic movie play-back in displaying such flows. The need for advanced algorithms to identify shock waves, vortices, and separation lines is pointed out. It is likely that the supercomputer will be needed for this process because of the size of 3D and/or unsteady CFD databases.

Buning, P. G.↗
NASA NTRS1985

A 3-D chimera grid embedding technique

A three-dimensional (3-D) chimera grid-embedding technique is described. The technique simplifies the construction of computational grids about complex geometries. The method subdivides the physical domain into regions which can accommodate easily generated grids. Communication among the grids is accomplished by interpolation of the dependent variables at grid boundaries. The procedures for constructing the composite mesh and the associated data structures are described. The method is demonstrated by solution of the Euler equations for the transonic flow about a wing/body, wing/body/tail, and a configuration of three ellipsoidal bodies.

Benek, J. A.↗
NASA NTRS1985

Developments in the simulation of compressible inviscid and viscous flow on supercomputers

In anticipation of future supercomputers, finite difference codes are rapidly being extended to simulate three-dimensional compressible flow about complex configurations. Some of these developments are reviewed. The importance of computational flow visualization and diagnostic methods to three-dimensional flow simulation is also briefly discussed.

Steger, J. L.↗
NASA NTRS1984

Viscous computation of a space shuttle flow field

A procedure is presented, as well as some results, to calculate the flow over the winged orbiter. This necessitates the use of two computer codes. A parabolized marching Navier-Stokes code is used to obtain the solution up to the bow shock-wing shock interaction region and for the region after the interaction. An unsteady Navier-Stokes code is to be used in the region of the shock interaction. Only resuls for the marching code are presented. For the flow conditions calculated, M infinity = 7.9, alpha = 25 deg, T(wall) = 540 R, Re(L) = 60728 per inch, laminar or turbulent, the PNS code was marched up to an X/L = 0.7 which is where the bow shock-wing shock interaction region occurs.

Chaussee, D. S.↗