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A directional-split modular user-friendly grid-point distribution code is applied to several test problems. Grid-point spacing is determined by user-specified constants denoting maximum and minimum grid spacings and relating the relative influence of smoothness and orthogonality. Estimates of truncation error, in terms of flowfield gradients and/or geometric features, are used to determine the point distribution.
The marker and cell technique is applied to three idealized hydrodynamic situations of interest to the Space Shuttle External Tank (ET). These applications illustrate the versatility of this technique as well as its usefulness in solving similar problems in the aerospace industry. The codes used are the simplified marker and cell (SMAC) code and the variable grid simplified marker and cell code named ERIE. The examples focus on the dropout height in the LH2 tank, the slosh behavior of propellant in the bottom dome of the ET during thrust termination, both of which are required to accurately determine the amount of usable propellants, and the effect on liquid surface due to LO2 tank wall motions occurring during Space Shuttle lift-off. Comparisons with experimental results and linear theory are also presented.
A directional-split, modular, user-friendly grid point distribution code is applied to several test problems. The code is self-adaptive in the sense that grid point spacing is determined by user-specified constants denoting maximum and minimum grid spacings and constants relating the relative influence of smoothness and orthogonality. Estimates of truncation error, in terms of flow-field gradients and/or geometric features, are used to determine the point distribution. Points are redistributed along grid lines in a specified direction in an elliptic manner over a user-specified subdomain, while orthogonality and smoothness are controlled in a parabolic (marching) manner in the remaining directions. Multidirectional adaption is achieved by sequential application of the method in each coordinate direction. The flow-field solution is redistributed onto the newly distributed grid points after each unidirectional adaption by a simple one-dimensional interpolation scheme. For time-accurate schemes such interpolation is not necessary and time-dependent metrics are carried in the fluid dynamic equations to account for grid movement.
FUN3D, an unstructured grid Navier-Stokes CFD code is capable of overset grid simu- lations, but does not have an internal method for assembling an overset grid system from a group of component grids. FUN3D currently relies on the third party codes Suggar++ and DiRTlib to perform domain assembly and provide intergrid connectivity for overset simulations. Rotorcraft simulations with moving, deforming blades require domain assembly and mesh deformation at each time step. For these simulations, the three primary drivers of computational cost for each time step are: deforming the mesh, performing domain assembly, and performing subiterations of the flow solver. FUN3D exhibits strong and weak scalability for the flow solver subiterations and the mesh deformation. However, FUN3D is currently hardwired directly to the serial version of Suggar++, which has a fixed cost for a given mesh system. Therefore, domain assembly begins to dominate the total cost of each time step as grid systems become larger. An integrated method for parallel domain assembly is presented that addresses scalability for large grid systems. Restructuring within FUN3D to accommodate integrated domain assembly is also discussed, which could enable use of the parallel Suggar++ library.
Grid modification methods have been under development at NASA to enable better predictions of low boom pressure signatures from supersonic aircraft. As part of this effort, two new codes, Stretched and Sheared Grid - Modified (SSG) and Boom Grid (BG), have been developed in the past year. The CFD results from these codes have been compared with ones from the earlier grid modification codes Stretched and Sheared Grid (SSGRID) and Mach Cone Aligned Prism (MCAP) and also with the available experimental results. NASA's unstructured grid suite of software TetrUSS and the automatic sourcing code AUTOSRC were used for base grid generation and flow solutions. The BG method has been evaluated on three wind tunnel models. Pressure signatures have been obtained up to two body lengths below a Gulfstream aircraft wind tunnel model. Good agreement with the wind tunnel results have been obtained for both on-track and off-track (up to 53 degrees) cases. On-track pressure signatures up to ten body lengths below a Straight Line Segmented Leading Edge (SLSLE) wind tunnel model have been extracted. Good agreement with the wind tunnel results have been obtained. Pressure signatures have been obtained at 1.5 body lengths below a Lockheed Martin aircraft wind tunnel model. Good agreement with the wind tunnel results have been obtained for both on-track and off-track (up to 40 degrees) cases. Grid sensitivity studies have been carried out to investigate any grid size related issues. Methods have been evaluated for fully turbulent, mixed laminar/turbulent and fully laminar flow conditions.
The objective of this work was to provide forecasters with a tool to indicate the warm season climatological probability of one or more lightning strikes within a circle at a site within a specified time interval. This paper described the AMU work conducted in developing flow regime based climatologies of lightning probabilities for the SLF and seven airports in the NWS MLB CWA in east-central Florida. The paper also described the GUI developed by the AMU that is used to display the data for the operational forecasters. There were challenges working with gridded lightning data as well as the code that accompanied the gridded data. The AMU modified the provided code to be able to produce the climatologies of lightning probabilities based on eight flow regimes for 5-, 10-, 20-, and 30-n mi circles centered on eight sites in 1-, 3-, and 6-hour increments.
A parametric cranked delta planform study has been conducted in the Langley 14- by 22-Foot Subsonic Tunnel with the following objectives: (1) to evaluate the vortex flap design methodology for cranked delta wings, (2) to determine the influence of leading-edge sweep and the outboard wing on vortex flap effectiveness, (3) to evaluate novel flow control concepts, and (4) to validate unstructured grid Euler computer code predictions with modeled vortex and trailing-edge flaps. Two families of cranked delta planforms were investigated. One family had constant aspect ratio, while the other had a constant nondimensional semispan location of the leading-edge break. The inboard leading-edge sweep of the planforms was varied between 68 deg., 71 deg., and 74 deg., while outboard leading-edge sweep was varied between 48 deg. and 61 deg. Vortex flaps for the different planforms were designed by an analytical vortex flap design method. The results indicate that the effectiveness of the vortex flaps was only slightly influenced by the variations in the parametric planforms. The unstructured grid Euler computer code was successfully used to model the configurations with vortex flaps. The vortex trap concept was successfully demonstrated.
Engineering codes are needed which can calculate convective heating rates accurately and expeditiously on the surfaces of high-speed vehicles. One code which has proven to meet these needs is the Langley Approximate Three-Dimensional Convective Heating (LATCH) code. It uses the axisymmetric analogue in an integral boundary-layer method to calculate laminar and turbulent heating rates along inviscid surface streamlines. It requires the solution of the inviscid flow field to provide the surface properties needed to calculate the streamlines and streamline metrics. The LATCH code has been used with inviscid codes which calculated the flow field on structured grids, Several more recent inviscid codes calculate flow field properties on unstructured grids. The present research develops a method to calculate inviscid surface streamlines, the streamline metrics, and heating rates using the properties calculated from inviscid flow fields on unstructured grids. Mr. Chris Riley, prior to his departure from NASA LaRC, developed a preliminary code in the C language, called "UNLATCH", to accomplish these goals. No publication was made on his research. The present research extends and improves on the code developed by Riley. Particular attention is devoted to the stagnation region, and the method is intended for programming in the FORTRAN 90 language.
Data from the "Turbulence Modeling Resource" website for turbulent flow over an NACA-0012 airfoil is analyzed to determine the convergence behavior of three second-order CFD (Computational Fluid Dynamics) codes: CFL3D (Computational Fluids Lab 3 Dimensional flow solver), FUN3D (Fully Unstructured Navier-stokes flow solver), and TAU (German Aerospace Center (DLR) 2 dimensional code for unstructured hybrid grids solving the Reynolds-Averaged Navier-Stokes equations or the Euler equations). The convergence of both integrated properties and pointwise data are examined. Several different methods for estimating errors and computing convergence rates are compared. A high-order extension to the Richardson extrapolation is developed that improves the accuracy of the mesh limit values and provides a quantitative estimate of the threshold of the asymptotic regime. The coefficient of total drag exhibits second-order convergence for all three codes, and convergence is monotone over a sequence of 7 grids. Other force coefficients are not so well behaved. The convergence rates of the viscous component of drag on the three nest grids ranges from 3:0 for CFL3D to 1:0 for FUN3D. The three codes are converging to similar but not identical solutions. The largest differences between the codes are in the coefficient of lift for which the difference between CFL3D and FUN3D is greater than 10 (sup minus 4). The best agreement occurs in the viscous component of drag, which is the only force component for which all three codes are converging towards each other at a rate of second-order. The agreement between the two unstructured grid codes is good with all properties except lift converging towards common values at a rate of second-order. No one code was universally better than the other. The TAU code has the lowest error in total drag, FUN3D has the lowest error in lift, and CFL3D has the lowest error in the viscous component of drag.
In June 2001, the first AIAA Drag Prediction Workshop was held to evaluate results obtained from extensive N-Version testing of a series of RANS CFD codes. The geometry used for the computations was the DLR-F4 wing-body combination which resembles a medium-range subsonic transport. The cases reported include the design cruise point, drag polars at eight Mach numbers, and drag rise at three values of lift. Although comparisons of the code-to-code medians with available experimental data were similar to those obtained in previous studies, the code-to-code scatter was more than an order-of-magnitude larger than expected and far larger than desired for design and for experimental validation. The second Drag Prediction Workshop was held in June 2003 with emphasis on the determination of installed pylon-nacelle drag increments and on grid refinement studies. The geometry used was the DLR-F6 wing-body-pylon-nacelle combination for which the design cruise point and the cases run were similar to the first workshop except for additional runs on coarse and fine grids to complement the runs on medium grids. The code-to-code scatter was significantly reduced for the wing-body configuration compared to the first workshop, although still much larger than desired. However, the grid refinement studies showed no sign$cant improvement in code-to-code scatter with increasing grid refinement.
The advancing front technique is being used to develop a code to generate grids around complex 3-D configurations for use in computing the invisid flow solutions by the Euler equations. By the advancing front technique points are introduced concurrently with the connectivity information so that a separate library is not required. The generation of a 3-D grid is accomplished in several steps. First the boundaries of the domain to be gridded must be described by two-, three- or four-sided surface patches. Next, a background mesh is required to control the grid spacing and stretching throughout the domain. This coarse tetrahedral grid is not required to conform to any of the boundaries. Next, each of the patches is mapped to 2-D, triangulated by the advancing front technique and mapped back to 3-D. These triangles form the initial front for the generation of the final tetrahedral mesh.
A static and dynamic Chimera overset-grid capability is added to an established NASA tetrahedral unstructured parallel Navier-Stokes flow solver, USM3D. Modifications to the solver primarily consist of a few strategic calls to the Donor interpolation Receptor Transaction library (DiRTlib) to facilitate communication of solution information between various grids. The assembly of multiple overlapping grids into a single-zone composite grid is performed by the Structured, Unstructured and Generalized Grid AssembleR (SUGGAR) code. Several test cases are presented to verify the implementation, assess overset-grid solution accuracy and convergence relative to single-grid solutions, and demonstrate the prescribed relative grid motion capability.
QX MAN is a grid and solution file manipulation program written primarily for the PARC code and the GRIDGEN family of grid generation codes. QX MAN combines many of the features frequently encountered in grid generation, grid refinement, the setting-up of initial conditions, and post processing. QX MAN allows the user to manipulate single block and multi-block grids (and their accompanying solution files) by splitting, concatenating, rotating, translating, re-scaling, and stripping or adding points. In addition, QX MAN can be used to generate an initial solution file for the PARC code. The code was written to provide several formats for input and output in order for it to be useful in a broad spectrum of applications.
A discussion of the strengths and weaknesses of overset composite grid and solution technology is given, along with a sampling of current work in the area. Major trends are identified, and the observation is made that generalized and hybridized overset methods provide a natural framework for combining disparate mesh types and physics models. Because of this, the author concludes that overset methods will be the foundation for the general purpose computational fluid dynamics programs of the future.
This paper describes and analyzes a series of nearly 90 CFD test cases performed as a contribution to the second Drag Prediction Workshop, held in association with the AIAA in June 2003. Two configurations are included: DLR-F6 wing-body and wing-body-nacelle-pylon. The ability of CFD to predict the drag, lift, and pitching moment from experiment-including the "delta" arising from the addition of the nacelle and pylon-is assessed. In general, at a fixed angle of attack CFD overpredicts lift, but predicts the delta C (sub L) reasonably well. At low lift levels (C (sub L) less than 0.3)), delta C (sub D) is 20-30 drag counts (30-45%) high. At the target lift coefficient of C(sub L) = 0.5, delta C (sub D) is overpredicted by between 11-16 counts. However, the primary contribution of this paper is mot so much the assessment of CFD against experiment, but rather a detailed assessment and analysis of CFD variation. The series of test cases are designed to determine the sensitivity/variability of CFD to a variety of factors, including grid, turbulence model, transition code, and viscous model. Using medium-level grids (6-11 million points) at the target lift coefficient, the maximum variation in drag due to different grids is 5-11 drag counts, due to code is 5-10 counts, due to turbulence model is 7-15 counts, due to transition is 10-11 counts, and due to viscous model is 4-5 counts. Other specific variations are described in the paper.
The subsonic flow field was numerically simulated around several High Speed Research Reference H configurations at various pitch and yaw angles. A sequence of structured-viscous grids were generated; the first grid modeled the wing-body high-lift geometry, and the second grid incorporated the nacelles and the horizontal tail. The third grid modeled the full-span geometry for sideslip calculations, and was obtained by mirroring a coarser version of the second grid. The CFL3D code, a Reynolds averaged, thin-layer Navier-Stokes flow solver for structural grids, was used for the flow solver and modeled the free-air Reference H high-lift configuration at wind tunnel conditions of Mach number 0.24 and Reynolds number of 1.4 x 10(exp 5) per in. Pitch sweeps were performed at angles of attack from 6 deg to 15 deg. Sideslip angle sweeps at 0 deg <= Beta <= +18 deg were performed at an angle of attack of 8 deg. The lateral and longitudinal performance characteristics were well predicted and very good force and moment comparisons were obtained. A very complex multiple vortical system develops at the higher angles of attack, and detailed postprocessing of the solutions provided a comprehensive three-dimensional understanding of the flow which helps to correlate and interpret the wind tunnel data.
Grid generation issues relating to the simulation of the X33 aerothermal environment using the GASP code are explored. Required grid densities and normal grid stretching are discussed with regards to predicting the fluid dynamic and heating environments with the desired accuracy. The generation of volume grids is explored and includes discussions of structured grid generation packages such as GRIDGEN, GRIDPRO and HYPGEN. Volume grid manipulation techniques for obtaining desired outer boundary and grid clustering using the OUTBOUND code are examined. The generation of the surface grid with the required surface grid with the required surface grid topology is also discussed. Utilizing grids without singular axes is explored as a method of avoiding numerical difficulties at the singular line.
The development of a high speed propeller noise prediction code at Langley Research Center is described. The code utilizes two recent acoustic formulations in the time domain for subsonic and supersonic sources. The structure and capabilities of the code are discussed. Grid size study for accuracy and speed of execution on a computer is also presented. The code is tested against an earlier Langley code. Considerable increase in accuracy and speed of execution are observed. Some examples of noise prediction of a high speed propeller for which acoustic test data are available are given. A brisk derivation of formulations used is given in an appendix.