The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA Ames 3.5-ft Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pilot-pressure flowfield surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at NASA Ames demonstrate the capabilities of this code. HE advanced computational fluid dynamics (CFD) com- puter codes being developed for use in the design of such hypersonic aircraft as the National Aero-Space Plane (NASP) and other hypersonic vehicles require comparisons of the com- putational results with a broad spectrum of experimental data to fully assess the validity of the codes and to develop confi- dence in the numerical simulation procedures. This is particu- larly true for complex flowfields with such features as bound- ary-layer transition and turbulence, rapid flow expansions, and leeside flow with the attendant flow separation and vor- tices. Validated codes for such flowfields will be critical to the development of the NASP and other hypersonic vehicles. Therefore, the objective of the present investigation is to es- tablish a benchmark experimental data base for a generic hy- personic vehicle shape for validation and/or calibration of advanced CFD computer codes. This is being done by con- ducting a comprehensive test program for a generic all-body hypersonic aircraft model in the NASA Ames 3.5-ft Hyper- sonic Wind Tunnel to obtain pertinent surface and flowfield data over a broad range of test conditions. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pitot-pressure flowfield surveys will be presented in this, paper. Of particular signifi- cance, comparisons of the experimental results with computa- tional results from the NASA Ames UPS code (an upwind parabolized Navier-Stokes solver) will be shown to demon- strate the capabilities of this code. Some comparisons of the data with computations from approximate inviscid methods will also be given.

Flow over an All-Body Hypersonic Aircraft: Experiment and Computation

The objective of the present investigation is to establish a benchmark experimental data base for a generic hypersonic vehicle shape for validation and/or calibration of advanced computational fluid dynamics computer codes. This paper includes results from the comprehensive test program conducted in the NASA/Ames 3.5-foot Hypersonic Wind Tunnel for a generic all-body hypersonic aircraft model. Experimental and computational results on flow visualization, surface pressures, surface convective heat transfer, and pitot-pressure flow-field surveys are presented. Comparisons of the experimental results with computational results from an upwind parabolized Navier-Stokes code developed at Ames demonstrate the capabilities of this code.

Experimental and computational surface and flow-field results for an all-body hypersonic aircraft

The aerodynamic influence coefficients calculated using an existing linear theory program were used to modify the pressures calculated using impact theory. Application of the combined approach to several wing-alone configurations shows that the combined approach gives improved predictions of the local pressure and loadings over either linear theory alone or impact theory alone. The approach not only removes most of the short-comings of the individual methods, as applied in the Mach 4 to 8 range, but also provides the basis for an inverse design procedure applicable to high speed configurations.

https://ntrs.nasa.gov/api/citations/19790005845/downloads/19790005845.pdf

Feasibility of combining linear theory and impact theory methods for the analysis and design of high speed configurations

An experimental investigation was conducted to determine the effects of several variations in body shape on the aerodynamic characteristics of an all-body hypersonic aircraft configuration. The basic configuration had a delta planform with an elliptic cone forebody and an afterbody of elliptic cross section terminating in a straight-line trailing edge. Variations in body shape included the ratio of maximum cross-sectional to body planform area, body leading-edge sweep, and forebody length ratio. In addition, the effects of a thin wing mounted on one of the bodies was investigated, and the aerodynamic characteristics of just the forebodies of two of the configurations were determined. The models had no stabilizing surfaces or propulsion system packages. Ranges of angle of attack (-4 deg to +15 deg) and angle of sideslip (-4 deg to +8 deg) were investigated. Of the four complete bodies, the configuration with the lowest ratio of cross-sectional to body planform area had the highest maximum lift-drag ratio and the greatest level of longitudinal stability at most Mach numbers. All the configurations had positive longitudinal stability near maximum lift-drag ratio at most Mach numbers. With exception of the lowest subsonic Mach numbers, changes in body sweep angle and in forebody length ratio had only minor effects on maximum lift-drag ratio.

https://ntrs.nasa.gov/api/citations/19720016346/downloads/19720016346.pdf

Effects of body shape on the aerodynamic characteristics of an all-body hypersonic aircraft configuration at Mach numbers from 0.65 to 10.6

N UMERICAL simulation of the flowfield of a practical configuration poses severe difficulty due to complex gridgeneration procedures. A Cartesian grid with near-wall extruded grids [1], chimera or overset grids [2], grid-free methods [3], and a combination of the aforementioned methods [4] is used to solve flow past complex configurations. The grid-free methods operate on a distribution of points in the domain and require a set of supporting nodes around each point to evaluate the spatial derivatives of the governing fluid equations. The point distribution can be obtained from structured, unstructured, Cartesian, hybrid, or overlapped meshes, or a randomdistribution of points. In recent years, quite a few grid-free methods were proposed in the field of compressible fluid flow. Among them, the least-squares kinetic upwind method developed by Deshpande et al. [5] has received much attention of researchers and been applied to number of complex flight vehicle configurations. Recently, the method was successfully applied to a store separation dynamics problem using a chimera cloud of points [6]. Batina [7] developed a gridlessmethod that used the least-squares method with the unbiased support of points for the discretization of spatial derivatives. Artificial viscosity is used to stabilize the solutions and applied to inviscid and laminar flows. Lohner et al. [8] developed a finite point method, in which an upwind scheme was used to stabilize the solutions, and applied it to inviscid compressible flows. In the finite point method, the direction of upwinding is based on coefficients of the least-squares discretization, which is purely geometric. In recent years, the grid-freemethod has been successfully applied to simulate turbulent flow past complex flight vehicle configurations [9,10]. These methods used either overset grids or extruded layers of points near the wall, along with Cartesian grids in the offbody region, to get the distribution of points; and neighbors are obtained using search algorithms guided by grid information, which is therefore known as the semimeshless method. Lohner et al. [8] developed the advancing point generationmethod to generate a cloud of points, and the neighbors of those points were obtained using a local Delaunay triangulation. The applications, so far, are limited to subsonic and transonic flows. In the present work, a grid-free Euler and Navier–Stokes (GEANS) solver has been developed using a gridless method [7] with upwind fluxes for flow stabilization, and it has been validated for hypersonic flows at higher angles of attack. One of the main drawbacks of the grid-free methods is the lack of conservation. Katz and Jameson [11] enforced conservation by modifying weights in the least-squares discretization. However, such modified weights may become negative for certain distribution of points that leads to nonpositive solutions. Chiu et al. [12] proposed a method of generating meshless coefficients with conservation constraints at the discrete level; however, thismethodwas complex to implement for three-dimensional (3-D) problems with an anisotropic distribution of points. In the present work, high-speed flows are simulated without enforcing the conservation property. A detail experimental results [13] for an all-body hypersonic aircraft is available for comparison of aerodynamic forces and moments in addition to local flowfields. The flowfield around the geometry is very complex, which involves strong compressions in the windward side and strong expansions in the leeward side, with flow separation and vortices at a hypersonic Mach number. Furthermore, highaspect-ratio grid cells are required to resolve the very fine details of the flowfield. The simulation of such flowfields requires a robust flow solver that can handle both strong oblique shock wave and high expansion regions, as well as be accurate enough to resolve the boundary layer. Therefore, the aforementioned configuration is considered for validating the grid-free Euler and Navier–Stokes solver [14] at hypersonic speed and the results are compared with the experimental values. The geometry considered for validations in the present work is simpler and amenable for generation of simple structured grids. Therefore, structured grids are generated around the body to get a distribution of points and supporting nodes are obtained using the structured grid adjacency relation.

Development and Validation of a Grid-Free Viscous Solver

Aerodynamic characteristics of a model designed to represent an all body, hypersonic cruise aircraft are presented for Mach numbers from 0.65 to 10.6. The configuration had a delta planform with an elliptic cone forebody and an afterbody of elliptic cross section. Detailed effects of varying angle of attack (-2 to +15 deg), angle of sideslip (-2 to +8 deg), Mach number, and configuration buildup were considered. In addition, the effectiveness of horizontal tail, vertical tail, and canard stabilizing and control surfaces was investigated. The results indicate that all configurations were longitudinally stable near maximum lift drag ratio. The configurations with vertical tails were directionally stable at all angles of attack. Trim penalties were small at hypersonic speeds for a center of gravity location representative of the airplane, but because of the large rearward travel of the aerodynamic center, trim penalties were severe at transonic Mach numbers.

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W. P. Nelms

Papers

Aerodynamic characteristics of an all-body hypersonic aircraft configuration at Mach numbers from 0.65 to 10.6