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Weilmuenster, K. J.

Publications and source records attributed to Weilmuenster, K. J..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Summary Report of the Orbital X-34 Wing Static Aeroelastic Study

This report documents the results of a computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid aerodynamic characteristics taking into account the wing structural flexibility. This was a joint exercise between LaRC and SDRC of California. SDRC modeled the structural details of the wing, and provided the structural deformation for a given pressure distribution on its surfaces. This study was done for a Mach number of 1.35 and an angle of attack of 9 deg.; the freestream dynamic pressure was assumed to be 607 lb/sq ft. Only the wing and the body were simulated in the CFD computations. Two wing configurations were examined. The first had the elevons in the undeflected position and the second had the elevons deflected 20 deg. up. The results indicated that with elevon undeflected, the wing twists by about 1.5 deg. resulting in a reduction in the angle of attack at the wing tip to by 1.5 deg. The maximum vertical deflection of the wing is about 3.71 inches at the wing tip. For the wing with the undeflected elevons, the effect of this wing deformation is to reduce the normal force coefficient (C(sub N)) by 0.012 and introduce a noise up pitching moment coefficient (C(sub m)) of 0.042.

Prabhn, Ramadas K.↗
NASA NTRS1993

HL-20 computational fluid dynamics analysis

The essential elements of a computational fluid dynamics analysis of the HL-20/personnel launch system aerothermal environment at hypersonic speeds including surface definition, grid generation, solution techniques, and visual representation of results are presented. Examples of solution technique validation through comparison with data from ground-based facilities are presented, along with results from computations at flight conditions. Computations at flight points indicate that real-gas effects have little or no effect on vehicle aerodynamics and, at these conditions, results from approximate techniques for determining surface heating are comparable with those obtained from Navier-Stokes solutions.

Weilmuenster, K. J.↗
NASA NTRS1993

A multiblock analysis for shuttle orbiter re-entry heating from Mach 24 to Mach 12

A multiblock, laminar heating analysis for the shuttle orbiter at three trajectory points ranging from Mach 24.3 to Mach 12.86 on re-entry is described. The analysis is performed using the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) with both a seven species chemical nonequilibrium model and an equilibrium model. A finite-catalytic-wall model appropriate for shuttle tiles at a radiative equilibrium wall temperature is applied. Computed heating levels are generally in good agreement with the flight data though a few rather large discrepancies remain unexplained. The multiblock relaxation strategy partitions the flowfield into manageable blocks requiring a fraction of the computational resources (time and memory) required by a full domain approach. In hot, the computational cost for a solution at even a single trajectory point would be prohibitively expensive at the given resolution without the multiblock approach. Converged blocks are reassembled to enable a fully coupled converged solution over the entire vehicle, starting from a nearly converged initial condition.

Gnoffo, Peter A.↗
NASA NTRS1993

Navier-Stokes simulations of the Shuttle Orbiter aerodynamic characteristics with emphasis on pitch trim and bodyflap

An analysis of the longitudinal aerodynamics of the Shuttle Orbiter in the hypersonic flight regime is made through the use of computational fluid dynamics (CFD). Particular attention is given to establishing the cause of the 'pitching moment anomaly' which occurred on the Orbiter's first flight and to computing the aerodynamics of a complete Orbiter configuration at flight conditions. Data from ground based facilities as well as Orbiter flight data are used to validate the computed results. Analysis shows that the 'pitching moment anomaly' is a real gas chemistry effect which cannot be simulated in ground-based facilities. Computed flight aerodynamics for the Orbiter are within 5 percent of the measured flight values and trim bodyflap deflections are predicted to within 10 percent.

Weilmuenster, K. J.↗
NASA NTRS1992

High angle-of-attack inviscid Shuttle Orbiter computation

Results are presented of a comparison between the pressure distributions predicted by the perfect-gas computational fluid dynamics and the Shuttle Orbiter wind-tunnel data for high angles of attack, using the LAURA (for Langley Aerothermodynamic Upwind Relaxation Algorithm) as applied to the wind-tunnel condition to predict the flow over the vehicle. It is shown that the calculated pressures compare well with the wind tunnel data for both the windward and the leeward sides, indicating that the salient inviscid flow features were properly modeled.

Kleb, William L.↗
NASA NTRS1992

Solution strategies and heat transfer calculations for three-dimensional configurations at hypersonic speeds

A procedure which reduces the memory requirements for computing the viscous flow over a modified Orbiter geometry at a hypersonic flight condition is presented. The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) code which incorporates a thermochemical nonequilibrium chemistry model, a finite rate catalytic wall boundary condition and wall temperature distribution based on radiation equilibrium is used in this study. In addition, the effect of choice of 'min mod' function, eigenvalue limiter and grid density on surface heating is investigated. The surface heating from a flowfield calculation at Mach number 22, altitude of 230,000 ft and 40 deg angle of attack is compared with flight data from three Orbiter flights.

Weilmuenster, K. J.↗
NASA NTRS1992

Characteristics of the Shuttle Orbiter leeside flow during a reentry condition

A study of the leeside flow characteristics of the Shuttle Orbiter is presented for a reentry flight condition. The flow is computed using a point-implicit, finite-volume scheme known as the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). LAURA is a second-order accurate, laminar Navier-Stokes solver, incorporating finite-rate chemistry with a radiative equilibrium wall temperature distribution and finite-rate wall catalysis. The resulting computational solution is analyzed in terms of salient flow features and the surface quantities are compared with flight data.

Kleb, William L.↗
NASA NTRS1991

Comparison of heating rate calculations with experimental data on a modified Shuttle orbiter at Mach 6

Heating-rate calculations from an 'engineering' code and a 'benchmark' Navier-Stokes code are compared with experimental wind tunnel data obtained on a 'modified' 0.0075-scale Shuttle orbiter at a Mach number of 6 and a freestream Reynolds number of 1.9 x 10 to the 6th/m. Comparisons are presented along the windward symmetry plane, in a circumferential direction around the body, and along the wing leading edge at angles of attack of 30 and 40 deg. Both codes provide accurate predictions of the heating rate over most of the body, but because the run time for the engineering code is relatively short, it is an ideal tool for both parametric and design studies.

Hamilton, H. H., II↗
NASA NTRS1991

Gridding strategies and associated results for winged entry vehicles

Two volume grid strategies based on similar software are presented for a geometrically complex entry vehicle. It is shown that quick relatively simple grid generation techniques can be used to construct computational grids about entry vehicles with nontrivial geometric complexity. Particular attention is given to both single- and dual-block grid topologies for a lifting body and solution generated on each grid. Results indicate that the grid topology can have an effect on the nature of the solution.

Weilmuenster, K. J.↗
NASA NTRS1990

Surface definition and grid generation about an Assured Crew Return Vehicle (ACRV) for Space Station Freedom

The surface definition and grid generation about an Assured Crew Return Vehicle (ACRV) for the Space Station Freedom are described. The purpose of the surface definition and grid generation is to provide the necessary geometry information for CFD calculations about the vehicle. There are two salient features in this description. The first is that the numerical model representing the ACRV configuration is obtained from the measurement of an existing wind tunnel model. The method for smoothing the measured data and obtaining the numerical model is described. The second feature is the description of the algebraic grid generation method and software to compute volume grids about the ACRV. The methods and software allow rapid computation of volume grids for a wide range of flow conditions.

Smith, R. E.↗
NASA NTRS1987

Simplified curve fits for the thermodynamic properties of equilibrium air

New, improved curve fits for the thermodynamic properties of equilibrium air have been developed. The curve fits are for pressure, speed of sound, temperature, entropy, enthalpy, density, and internal energy. These curve fits can be readily incorporated into new or existing computational fluid dynamics codes if real gas effects are desired. The curve fits are constructed from Grabau-type transition functions to model the thermodynamic surfaces in a piecewise manner. The accuracies and continuity of these curve fits are substantially improved over those of previous curve fits. These improvements are due to the incorporation of a small number of additional terms in the approximating polynomials and careful choices of the transition functions. The ranges of validity of the new curve fits are temperatures up to 25 000 K and densities from 10 to the -7 to 10 to the 3d power amagats.

Srinivasan, S.↗
NASA NTRS1986

Calculation of convective heating on proposed aeroassist flight experiment vehicle

Convective heating calculations of a NASA aeroassist flight experiment (AFE) configuration consisting of a blunted, elliptic cone raked off at the base, and fitted with a skirt-type afterbody with a generous corner radius, are presented. The HALIS inviscid flowfield code is used to supply edge properties to an approximate three-dimensional boundary-layer method to calculate the surface heating. Calculations at Mach number 9.86 and approximate Reynolds number of 500,000, for angles of attack from -10 to 10 degrees, are in general good agreement with Langley 31-inch Mach 10 Tunnel experimental data. The results validate the application of the present overall AFE configuration heating pattern heating contours to aeroassist orbital transfer vehicle heating distribution determinations.

Hamilton, H. H., II↗
NASA NTRS1986

A comparison of computed and measured aerodynamic characteristics of a proposed aeroassist flight experiment configuration

The use of experimental or computed data to evaluate the performance of aeroassist orbital transfer vehicles (AOTVs) is discussed. Aerodynamic and surface pressure data are derived from computed flowfield solutions (the HALIS inviscid flowfield code), the Newtonian theory, and ground-based, wind tunnel studies (hypersonic He tunnel, 31-inch Mach 10 tunnel, and a hypersonic CF4 tunnel). The wind tunnel and HALIS models were tested at angles-of-attack that ranged from -10 to 10 deg. The effects of the ellipticity of the nose, the radius of the circular arc, and the angle through which the arc passes on the aerodynamic characteristics of the vehicles are examined. It is observed that the HALIS and CF4 tunnel generated surface pressure on the AOTVs produced the most useful aerodynamic data. Good correlation is obtained for the experimental and computational data; however, the Newtonian results do not correspond to the tunnel/HALIS data.

Weilmuenster, K. J.↗
NASA NTRS1986

Simplified curve fits for the thermodynamic properties of equilibrium air

New improved curve fits for the thermodynamic properties of equilibrium air were developed. The curve fits are for p = p(e,rho), a = a(e,rho), T = T(e,rho), s = s(e,rho), T = T(p,rho), h = h(p,rho), rho = rho(p,s), e = e(p,s) and a = a(p,s). These curve fits can be readily incorporated into new or existing Computational Fluid Dynamics (CFD) codes if real-gas effects are desired. The curve fits were constructed using Grabau-type transition functions to model the thermodynamic surfaces in a piecewise manner. The accuracies and continuity of these curve fits are substantially improved over those of previous curve fits appearing in NASA CR-2470. These improvements were due to the incorporation of a small number of additional terms in the approximating polynomials and careful choices of the transition functions. The ranges of validity of the new curve fits are temperatures up to 25,000 K and densities from 10 to the minus 7th to 100 amagats (rho/rho sub 0).

Srinivasan, S.↗
NASA NTRS1986

A comparison of computed and experimental surface pressure and heating on 70 deg sphere cones at angles of attack to 20 deg

A method for handling the flow fields around 70 deg spherical cones at angles of attack up to 20 deg is described. Surface pressure and heating results are compared with existing wind tunnel data. Computed surface pressures are found to be in excellent agreement with experimental data, while the surface heating from approximate and finite difference forms of the boundary layer equations are in good agreement with the data. The effect of bluntness ratio on surface pressures and heating is very small. Body corner radius, while not having a significant impact on body aerodynamics, produced large changes in surface pressure and heating distribution.

Weilmuenster, K. J.↗
NASA NTRS1985

Experimental aerodynamic coefficients on a Shuttle-like vehicle at Mach 6 and 10 and comparison to prediction

Aerodynamic coefficients measured for an Orbiter-like configuration are presented for a range of angle of attack from 15 deg to 45 deg at Mach 6 and 10. The low-gamma aspect of a real gas on aerodynamic characteristics and shock detachment distance were simulated by testing this configuration in Mach 6 air (freestream gamma equal to 1.4) and Mach 6 CF4 (freestream gamma equal to 1.17). The effects of Mach number, Reynolds number, and gamma on these measurements are examined and comparisons made to an inviscid flowfield computer code known as HALIS. Pitching-moment coefficients measured in CF4 revealed a nose-up increment when compared with measurements in perfect air, indicative of real-gas effects. In general, the HALIS code accurately predicted the measured aerodynamic coefficients in air and CF4.

Micol, J. R.↗