Scott D. Thomas

AERODYNAMIC SHAPE OPTIMIZATION USING UNSTRUCTURED GRID METHODS

Two unstructured-grid Euler-based CFD codes, AIRPLANE and FLOWCART were coupled to a gradientbased quasi-Newton finite-difference optimization algorithm. These two optimization techniques were developed to provide detailed aerodynamic shape optimization methods for complete configurations with efficient grid generation methods. AIRPLANE utilizes a tetrahedral mesh whereas FLOWCART uses a hexahedral Cartesian mesh. Several codes were developed to facilitate aerodynamic shape optimization. These include developing surface grid perturbation methods with thickness constraint and overall surface overlap evaluations, used with both Euler codes, and adding a multigrid capability and a variety of mesh movement techniques to the AIRPLANE method. The AIRPLANE multigridding approach was proven to be accurate and effective, with typical speedup ratios of 3 to 5. The mesh movement techniques were effective in reducing the grid generation wall clock time by 70%. Detailed results of the AIRPLANE-based optimization technique are presented. The performance gains resulting from optimization are verified by computations with FLOWCART and OVERFLOW, and comparisons with experimental data on the baseline and optimized configurations. The low speed computational results of the baseline and optimized models were incorporated into an approach and landing simulation database. The controllability and handling qualities were good to excellent based on a piloted simulation in the NASA Ames Vertical Motion Simulator (VMS). FLOWCARTbased optimization was validated by comparing its gradients and design solutions with AIRPLANE’s on two separate optimization problems with identical design variables and objective functions.

Euler/experiment correlations of sonic boom pressure signatures

The ability of inviscid computational fluid dynamics (CFD) codes to compute sonic boom pressure signatures is examined using three different codes that solve the Euler equations of fluid flow on structured hexahedral and unstructured tetrahedral grids. The results of these Euler codes were evaluated by comparing the computed pressure signatures with near-field experimental data. The computational pressure signatures were determined at distances of one body length or less below the configuration in the plane of symmetry and extrapolated to experimental distances using the waveform parameter method. The extrapolated CFD pressure signatures gave acceptable correlations with experimental data, provided that fine grids were used between the surface and the spatial location of the pressure signature. I. Introduction T HE feasibility of a low-boom supersonic transport is again being investigated. A commercial transport is more efficient when supersonic flight is maintained for the entire cruise portion of the mission. However, sonic boom noise may limit the extent of overland supersonic flight over populated areas. A major research effort has begun to design supersonic transport configurations which exhibit acceptable sonic boom characteristics. As part of the design process, computational methods have been developed for analyzing sonic boom characteristics. Over the past two decades, CFD codes have become capable of computing the flowfield about realistic high-speed civil transport configurations. The simultaneous increase in computational resources is the primary reason that use of these codes has become practical. The use of CFD codes for sonic boom prediction and configuration design offers the potential for making wind-tunnel testing more productive. However, the accuracy of CFD codes for sonic boom prediction needs to be evaluated before application to low-boom design. The use of Euler flow solvers coupled with extrapolation methods for the computation of sonic boom pressure signatures is a new application of CFD. For instance, Siclari and Darden1 applied a supersonic Euler code to two low-boom concepts, designed for Mach 2 and Mach 3. The near-field CFD results were extrapolated for comparison with extrapolated wind-tunnel data, corresponding to aircraft at cruise altitude, on and off ground track. The extrapolation method was similar to the one used in the present report. Also, Page and Plotkin2 calculated the flow about a cone-wing model at Mach 2.01, and extrapolated the near-field pressures on a cylinder around the model for comparison with near-field wind-tunnel data. The approach described below validates the concept of using a planar extrapolation method with an Euler flow solver by comparing extrapolated near-field pressure signatures with experimental data for a range of configurations and flow conditions. The original paper3 contains additional information, particularly about the grid systems which were used.

Refined Tetrahedral Meshes with Mach Cone Aligned Prisms for Sonic Boom Analysis

A tetrahedral mesh generation method for acquiring accurate sonic boom pressure signatures several body lengths from an aircraft model has been developed. The method serves as a tool for aerodynamicists to efficiently create useful meshes for sonic boom analysis. The procedure includes generating a refined near-field grid with a cylindrically shaped boundary that encompasses the model just beyond its surface and a prismatic mesh from the cylindrical boundary to the far field. Projecting the boundary in the radial direction and forming prisms between neighboring layer faces creates the prism mesh. Each prism is subdivided into three tetrahedra resulting in a mesh comprised entirely of tetrahedral cells. The prism structure permits radial stretching and mesh alignment with the Mach cone around the aircraft model for accurate on- and offtrack signatures. Computational results for four models compared with experimental data validate this methodology.

An edge-based solution-adaptive method applied to the AIRPLANE code

ABSTRACTComputational methods to solve large-scale re-alistic problems in fluid flow can be made more effi-cient and cost effective by using them in conjunctionwith dynamic mesh adaption procedures that per-form simultaneous coarsening and refinement to cap-ture flow features of interest. This work couples thetetrahedral mesh adaption scheme, called 3D.TAG,with the AIRPLANE code to solve complete aircraftconfiguration problems in transonic and supersonicflow regimes. Results indicate that the near-fieldsonic boom pressure signature of a cone-cylinder isimproved, the oblique and normal shocks are bet-ter resolved on a transonic wing, and the bow shockahead of an unstarted inlet is better defined.INTRODUCTIONTraditional computational methods can bemade more efficient and cost effective by redistribut-ing the available mesh points to capture flowfieldphenomena of interest. Such adaptive proceduresevolve with the solution and provide a robust andreliable methodology. Highly localized regions ofmesh refinement are required in order to accuratelycapture shock waves, contact discontinuities, andshear layers. This provides the aerodynamicist withthe opportunity to obtain flow solutions on adaptedmeshes that are comparable to those obtained onglobally-fine grids.Two types of solution-adaptive grid strategiesare commonly used with unstructured-grid methods.

Swing-Wing Inline-Fuselage Transport Design Studies at Supersonic Flight Conditions

www.nasa.gov AERONAUTICS RESEARCH MISSION DIRECTOR ATE Swing-wing transports—with a wing that may be swept and then returned to its original position during flight—offer improved aerodynamic performance across all flight regimes, shorter field lengths, and steeper and quieter climb. is conceptual transport aircraft would spend much of its time in supersonic cruise, so careful shaping to enhance performance (lift to drag ratio) and reduce drag will improve range and safety, and reduce weight and fuel burn.

Fast viscous correction method for full-potential transonic wing analysis

An analysis of the transonic flowfield around a three-dimensional wing is carried out using a strip method. Attention is given to the boundary layer growth in the streamwise direction. A viscous correction technique is defined for the TWING code for solving the full potential equations. A viscous ramp at the base of a shock is superimposed on the boundary layer displacement thickness generated by an integral boundary layer method. A relationship is then obtained between the effective displacement thickness and a vertical component of the surface velocity, a transpirational boundary condition. The viscous correction is found to be unnecessary in weak shock conditions but gives a better shock position and pressure distribution in a strong shock condition when compared with data from an ONERA M6 airfoil and the Hinson and Burdges (1980) Wing A.