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In the current globalization, we can see many innovations being introduced or implemented in every aspect of field that are considered to be existed. Every country is aiming to develop its power over all the aspects that considered for comparison with other countries in order to stand at same level of competition with others. One such power considered by all countries to develop every possible way to have a healthy competition is the military power which involves basically innovations of fast moving aircraft having a high lift coefficient and low drag coefficient. Such an aircraft having the high lift and low drag coefficient is TEJAS (HAL) developed by country India on which the purpose of paper mainly sustains. The paper mainly focuses on steady-state flow analysis over aircraft TEJAS using the computer aided modelling techniques and also the comparison of the results obtained from the modelled techniques. The paper also outlines the designing of the structural model of the TEJAS in a modelling software, creation of a finite computational domain, segmentation of this domain into discrete intervals, applying boundary conditions such as velocity in order to obtain plots and desired results determining the coefficient of pressure, lift and drag coefficient, velocity magnitude etc. This paper also aims in creating awareness to the future students about the techniques involved and knowledge required for developing a designed modelled. This paper also highlights the use of CFD techniques involved for the purpose of fluid flow simulation of the aircraft especially performing the meshing techniques, pre and post processing techniques and finally the evaluation of the simulation. Finally this paper can be seen as source by future generation students in gaining knowledge about design, analysis and simulation of the structured model on various conditions, about the field of aerospace engineering and new innovations being developed and also about the career involved when the above fields were chosen foe specialization purposes

Implicit finite differences schemes are often the preferred numerical schemes in computational fluid dynamics, requiring less stringent stability bounds than the explicit schemes. Each iteration in an implicit scheme, however, involves global data dependencies in the form of second and higher order recurrences. Efficient parallel implementations of such iterative methods, therefore, are considerably more difficult and non-intuitive. In this paper, we consider the parallelization of the implicit schemes that are used for solving the Euler and the thin layer Navier-Stokes equations and that require inversions of large linear systems in the form of block tri-diagonal and/or block penta- diagonal matrices. We focus our attention on three-dimensional cases an present schemes that minimize the total execution time. We describe partitioning and scheduling schemes for alleviating the effects of the global data dependencies. An analysis of the communication and the computation aspects of these methods is presented. The effect of the boundary conditions on the parallel schemes is also discussed. The ARC-3D code, developed at NASA Ames, is used as an example application. Performance of the proposed methods is verified on the Victor multiprocessor system which is a message passing architecture developed at the IBM, T.J. Watson Research Center. (KR)

The Design Analysis Branch (NE-Ml) at the Kennedy Space Center has not had the ability to accurately couple Rigid Body Dynamics (RBD) and Computational Fluid Dynamics (CFD). OVERFLOW-D is a flow solver that has been developed by NASA to have the capability to analyze and simulate dynamic motions with up to six Degrees of Freedom (6-DOF). Two simulations were prepared over the course of the internship to demonstrate 6DOF motion of rigid bodies under aerodynamic loading. The geometries in the simulations were based on a conceptual Space Launch System (SLS). The first simulation that was prepared and computed was the motion of a Solid Rocket Booster (SRB) as it separates from its core stage. To reduce computational time during the development of the simulation, only half of the physical domain with respect to the symmetry plane was simulated. Then a full solution was prepared and computed. The second simulation was a model of the SLS as it departs from a launch pad under a 20 knot crosswind. This simulation was reduced to Two Dimensions (2D) to reduce both preparation and computation time. By allowing 2-DOF for translations and 1-DOF for rotation, the simulation predicted unrealistic rotation. The simulation was then constrained to only allow translations.

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from

Some recent developments in acceleration of convergence methods for vector sequences are reviewed. The methods considered are the minimal polynomial extrapolation, the reduced rank extrapolation, and the modified minimal polynomial extrapolation. The vector sequences to be accelerated are those that are obtained from the iterative solution of linear or nonlinear systems of equations. The convergence and stability properties of these methods as well as different ways of numerical implementation are discussed in detail. Based on the convergence and stability results, strategies that are useful in practical applications are suggested. Two applications to computational fluid mechanics involving the three dimensional Euler equations for ducted and external flows are considered. The numerical results demonstrate the usefulness of the methods in accelerating the convergence of the time marching techniques in the solution of steady state problems.

The aerospace plane design challenge is presented in the form of the view-graphs. The following topics are included: the CFD design technology development; CFD validation vs. measurable fluid dynamics validation; and discussion of results.

This paper describes and discusses the textbook, Fundamentals of Computational Fluid Dynamics by Lomax, Pulliam, and Zingg, which is intended for a graduate level first course in computational fluid dynamics. This textbook emphasizes fundamental concepts in developing, analyzing, and understanding numerical methods for the partial differential equations governing the physics of fluid flow. Its underlying philosophy is that the theory of linear algebra and the attendant eigenanalysis of linear systems provides a mathematical framework to describe and unify most numerical methods in common use in the field of fluid dynamics. Two linear model equations, the linear convection and diffusion equations, are used to illustrate concepts throughout. Emphasis is on the semi-discrete approach, in which the governing partial differential equations (PDE's) are reduced to systems of ordinary differential equations (ODE's) through a discretization of the spatial derivatives. The ordinary differential equations are then reduced to ordinary difference equations (O(Delta)E's) using a time-marching method. This methodology, using the progression from PDE through ODE's to O(Delta)E's, together with the use of the eigensystems of tridiagonal matrices and the theory of O(Delta)E's, gives the book its distinctiveness and provides a sound basis for a deep understanding of fundamental concepts in computational fluid dynamics.

In this paper we describe the parallelization of the multi-zone code versions of the NAS Parallel Benchmarks employing multi-level OpenMP parallelism. For our study we use the NanosCompiler, which supports nesting of OpenMP directives and provides clauses to control the grouping of threads, load balancing, and synchronization. We report the benchmark results, compare the timings with those of different hybrid parallelization paradigms and discuss OpenMP implementation issues which effect the performance of multi-level parallel applications.

The current compute environment that most researchers are using for the calculation of 3D unsteady Computational Fluid Dynamic (CFD) results is a super-computer class machine. The Massively Parallel Processors (MPP's) such as the 160 node IBM SP2 at NAS and clusters of workstations acting as a single MPP (like NAS's SGI Power-Challenge array and the J90 cluster) provide the required computation bandwidth for CFD calculations of transient problems. If we follow the traditional computational analysis steps for CFD (and we wish to construct an interactive visualizer) we need to be aware of the following: (1) Disk space requirements. A single snap-shot must contain at least the values (primitive variables) stored at the appropriate locations within the mesh. For most simple 3D Euler solvers that means 5 floating point words. Navier-Stokes solutions with turbulence models may contain 7 state-variables. (2) Disk speed vs. Computational speeds. The time required to read the complete solution of a saved time frame from disk is now longer than the compute time for a set number of iterations from an explicit solver. Depending, on the hardware and solver an iteration of an implicit code may also take less time than reading the solution from disk. If one examines the performance improvements in the last decade or two, it is easy to see that depending on disk performance (vs. CPU improvement) may not be the best method for enhancing interactivity. (3) Cluster and Parallel Machine I/O problems. Disk access time is much worse within current parallel machines and cluster of workstations that are acting in concert to solve a single problem. In this case we are not trying to read the volume of data, but are running the solver and the solver outputs the solution. These traditional network interfaces must be used for the file system. (4) Numerics of particle traces. Most visualization tools can work upon a single snap shot of the data but some visualization tools for transient problems require dealing with time.

This paper describes an experiment in which a large-scale scientific application development for tightly-coupled parallel machines is adapted to the distributed execution environment of the Information Power Grid (IPG). A brief overview of the IPG and a description of the computational fluid dynamics (CFD) algorithm are given. The Globus metacomputing toolkit is used as the enabling device for the geographically-distributed computation. Modifications related to latency hiding and Load balancing were required for an efficient implementation of the CFD application in the IPG environment. Performance results on a pair of SGI Origin 2000 machines indicate that real scientific applications can be effectively implemented on the IPG; however, a significant amount of continued effort is required to make such an environment useful and accessible to scientists and engineers.

The hierarchy of codes; boundary layers and turbulence models; uniqueness and artificial viscosity; efficiency of solution algorithms; data display and analysis; and cost of experimental validation are briefly examined.

Information on computational fluid dynamics (CFD) research and applications carried out at the NASA Langley Research Center is given in viewgraph form. The Langley CFD strategy, the five-year plan in CFD and flow physics, 3-block grid topology, the effect of a patching algorithm, F-18 surface flow, entropy and vorticity effects that improve accuracy of unsteady transonic small disturbance theory, and the effects of reduced frequency on first harmonic components of unsteady pressures due to airfoil pitching are among the topics covered.

Three dimensional, steady-state, laminar, fully viscous Navier-Stokes simulations were used to predict the behavior of incompressible liquids that were undergoing steady spin and steady precession at a fixed precession angle. These numerical simulations can predict steady viscous and pressure moments. These moments tend to increase the precession angle and reduce the spin rate of the container system. For a completely filled cylinder, liquid-induced roll and side (yaw) moments were computed as functions of cylinder height to diameter. Keywords: Finite difference, Incompressible flow, Liquid-filled projectile, Liquid moment, Low reynolds number, Rotating liquids.

This Technical Memorandum presents an algorithm for deforming a two- or three-dimensional aerodynamic grid given deflections on a physical boundary. This algorithm is to be used in coupling a computational fluid dynamics code with a computational structural mechanics code. The deformation algorithm was tested successfully on both two- and three-dimensional grids. The method sufficiently maintalns grid quality for smooth deflections of realistic orders of magnitude. The method is computationally efficient: the time required to deform the grid is small compared to the time required to solve the fluid dynamics.

The project was concerned with the development of high performance computer codes for problems in transonic aerodynamics. A practical rule to calculate the wave drag for solutions of the Euler equations was developed from an entropy equality. Symmetric shockless airfoils were analyzed for which uniqueness fails in the transonic case not just for potential flow, but also for the Euler equations.

The project was concerned with the development of high performance computer codes for problems in transonic aerodynamics. A practical rule to calculate the wave drag for solutions of the Euler equations was developed from an entropy equality. Symmetric shockless airfoils were analyzed for which uniqueness fails in the transonic case not just for potential flow, but also for the Euler equations.

The project was concerned with the development of high performance computer codes for problems in transonic aerodynamics. A practical rule to calculate the wave drag for solutions of the Euler equations was developed from an entropy equality. Symmetric shockless airfoils were analyzed for which uniqueness fails in the transonic case not just for potential flow, but also for the Euler equations.

Motivated by progressive climate-change influence on ice degradation in caves, in this paper we present a novel methodology to investigate the link between air dynamics and ice melting. Specifically, we use surveys available for the Leupa ice cave (LIC), located in the Canin-Kanin group in the southeastern Alps and a general purpose computational fluid dynamics model (CFD). Detailed numerical simulations are evaluated on the basis of well-established approaches that consider domain, grid, boundary-conditions, turbulence closure models, buoyancy effects, porous media properties and verification with measured data. External atmospheric conditions are the main trigger for internal circulation but morphology and thermal characteristics of ice and bedrock induce a dynamical process of heat exchange ultimately responsible for ice melting. This process is generally poorly documented in real conditions. Using CFD analyses we show that both in summer and winter, warm and cold air currents within the cave are “disturbed” by several vortices and stagnation zones which locally modify the energy balance. To account for this we introduce a macroscopic physical model based on energy balance between ice surfaces and the inner ice cave airflow to determine the heat exchanged between ice and air. Using this model, a prediction of ice thickness decay over time is obtained. In the case of LIC a reduction of initial 4 cm per year is first obtained with projection of a much faster increase. The methodology is general and easily extendable to other sites, proving to be a powerful method to estimate ice evolution in caves induced by external and internal forcing.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

A study was conducted to evaluate the accuracy of a commercial computational fluid dynamics (CFD) code (CFDRC-ACE+) for predicting incompressible air jet flows with simple geometries. Specifically, the axis- symmetric and two-dimensional heated air-jets were simulated using a standard k- epsilon turbulence model. These CFD predictions were directly compared to an extensive compilation of experimental data from archive literature. The round jet results indicated that the code over-predicted the velocity-spreading rate by 24% and the temperature spreading rate by 29%

Current objectives at NASA Johnson Space Center ate directed at future upgrade and replacement of the U.S. Space Shuttle's, currently toxic, Reaction Control System thrusters with dual mode thrusters that use nontoxic propellants. Experimentation to determine any performance advantages obtained using a dual mode thruster has not been performed by NASA. A computational fluid dynamics analysis is performed to evaluate the internal flow characteristics of this thruster under low thrust mode, torch igniter only, conditions. Several computational models, both two- and three-dimensional, are constructed to simulate the internal, steady-state flow characteristics. Comparison is made with current data on a similar type of flow (highly underexpanded free-jet flow) to show the appearance of barrel shocks and Mach disks. Regions of stagnate flow where heat transfer to chamber surfaces will be high and engine thrust performance are predicted based on computational data. Two different flow solvers, one using a finite volume method and the other using a finite difference method, are used to predict the engine's performance. A comparison of the two flow solvers is given based on their relative performance to compute solutions to this problem

Current objectives at NASA Johnson Space Center ate directed at future upgrade and replacement of the U.S. Space Shuttle's, currently toxic, Reaction Control System thrusters with dual mode thrusters that use nontoxic propellants. Experimentation to determine any performance advantages obtained using a dual mode thruster has not been performed by NASA. A computational fluid dynamics analysis is performed to evaluate the internal flow characteristics of this thruster under low thrust mode, torch igniter only, conditions. Several computational models, both two- and three-dimensional, are constructed to simulate the internal, steady-state flow characteristics. Comparison is made with current data on a similar type of flow (highly underexpanded free-jet flow) to show the appearance of barrel shocks and Mach disks. Regions of stagnate flow where heat transfer to chamber surfaces will be high and engine thrust performance are predicted based on computational data. Two different flow solvers, one using a finite volume method and the other using a finite difference method, are used to predict the engine's performance. A comparison of the two flow solvers is given based on their relative performance to compute solutions to this problem

NASA/Marshall Space Flight Center (MSFC) has an ongoing effort to transfer to industry the technologies developed at MSFC for rocket propulsion systems. The Technology Utilization (TU) Office at MSFC promotes these efforts and accepts requests for assistance from industry. One such solicitation involves a request from North American Marine Jet, Inc. (NAMJ) for assistance in the design of a water-jet-drive system to fill a gap in NAMJ's product line. NAMJ provided MSFC with a baseline axial flow impeller design as well as the relevant working parameters (rpm, flow rate, etc.). This baseline design was analyzed using CFD, and significant deficiencies identified. Four additional analyses were performed involving MSFC changes to the geometric and operational parameters of the baseline case. Subsequently, the impeller was redesigned by NAMJ and analyzed by MSFC. This new configuration performs significantly better than the baseline design. Similar cooperative activities are planned for the design of the jet-drive inlet.

A system and method for generating fluid flow parameter data for use in aerodynamic heating analysis. Computational fluid dynamics data is generated for a number of points in an area on a surface to be analyzed. Sub-areas corresponding to areas of the surface for which an aerodynamic heating analysis is to be performed are identified. A computer system automatically determines a sub-set of the number of points corresponding to each of the number of sub-areas and determines a value for each of the number of sub-areas using the data for the sub-set of points corresponding to each of the number of sub-areas. The value is determined as an average of the data for the sub-set of points corresponding to each of the number of sub-areas. The resulting parameter values then may be used to perform an aerodynamic heating analysis.

A computational method for the analysis of longitudinal-mode liquid rocket combustion instability has been developed based on the unsteady, quasi-one-dimensional Euler equations where the combustion process source terms were introduced through the incorporation of a two-zone, linearized representation: (1) A two-parameter collapsed combustion zone at the injector face, and (2) a two-parameter distributed combustion zone based on a Lagrangian treatment of the propellant spray. The unsteady Euler equations in inhomogeneous form retain full hyperbolicity and are integrated implicitly in time using second-order, high-resolution, characteristic-based, flux-differencing spatial discretization with Roe-averaging of the Jacobian matrix. This method was initially validated against an analytical solution for nonreacting, isentropic duct acoustics with specified admittances at the inflow and outflow boundaries. For small amplitude perturbations, numerical predictions for the amplification coefficient and oscillation period were found to compare favorably with predictions from linearized small-disturbance theory as long as the grid exceeded a critical density (100 nodes/wavelength). The numerical methodology was then exercised on a generic combustor configuration using both collapsed and distributed combustion zone models with a short nozzle admittance approximation for the outflow boundary. In these cases, the response parameters were varied to determine stability limits defining resonant coupling onset.

For long-duration in-space storage of cryogenic propellants, an axial jet mixer is one concept for controlling tank pressure and reducing thermal stratification. Extensive ground-test data from the 1960s to the present exist for tank diameters of 10 ft or less. The design of axial jet mixers for tanks on the order of 30 ft diameter, such as those planned for the Ares V Earth Departure Stage (EDS) LH2 tank, will require scaling of available experimental data from much smaller tanks, as well designing for microgravity effects. This study will assess the ability for Computational Fluid Dynamics (CFD) to handle a change of scale of this magnitude by performing simulations of existing ground-based axial jet mixing experiments at two tank sizes differing by a factor of ten. Simulations of several axial jet configurations for an Ares V scale EDS LH2 tank during low Earth orbit (LEO) coast are evaluated and selected results are also presented. Data from jet mixing experiments performed in the 1960s by General Dynamics with water at two tank sizes (1 and 10 ft diameter) are used to evaluate CFD accuracy. Jet nozzle diameters ranged from 0.032 to 0.25 in. for the 1 ft diameter tank experiments and from 0.625 to 0.875 in. for the 10 ft diameter tank experiments. Thermally stratified layers were created in both tanks prior to turning on the jet mixer. Jet mixer efficiency was determined by monitoring the temperatures on thermocouple rakes in the tanks to time when the stratified layer was mixed out. Dye was frequently injected into the stratified tank and its penetration recorded. There were no velocities or turbulence quantities available in the experimental data. A commercially available, time accurate, multi-dimensional CFD code with free surface tracking (FLOW-3D from Flow Science, Inc.) is used for the simulations presented. Comparisons are made between computed temperatures at various axial locations in the tank at different times and those observed experimentally. The affect of various modeling parameters on the agreement obtained are assessed.

Launch vehicles frequently experience a reduced stability margin through the transonic Mach number range. This reduced stability margin can be caused by the aerodynamic undamping one of the lower-frequency flexible or rigid body modes. Analysis of the behavior of a flexible vehicle is routinely performed with quasi-steady aerodynamic line loads derived from steady rigid aerodynamics. However, a quasi-steady aeroelastic stability analysis can be unconservative at the critical Mach numbers, where experiment or unsteady computational aeroelastic analysis show a reduced or even negative aerodynamic damping.Amethod of enhancing the quasi-steady aeroelastic stability analysis of a launch vehicle with unsteady aerodynamics is developed that uses unsteady computational fluid dynamics to compute the response of selected lower-frequency modes. The response is contained in a time history of the vehicle line loads. A proper orthogonal decomposition of the unsteady aerodynamic line-load response is used to reduce the scale of data volume and system identification is used to derive the aerodynamic stiffness, damping, and mass matrices. The results are compared with the damping and frequency computed from unsteady computational aeroelasticity and from a quasi-steady analysis. The results show that incorporating unsteady aerodynamics in this way brings the enhanced quasi-steady aeroelastic stability analysis into close agreement with the unsteady computational aeroelastic results.

A study was undertaken to investigate and analyze the flow field results produced by various computational solvers for a projectile of interest to the U.S. Army. Computational fluid dynamics (CFD) techniques were used to obtain numerical solutions for the flow field of a kinetic energy projectile with the original and a modified (experimental) sabot. Computed results were obtained at Mach 4.5 and a 0 deg angle of attack. Qualitative flow field features showed the pressure on the surface of the model as well as pressures in the flow field. The surface pressure data on the projectile were extracted from the solution files and compared. In all cases, the results were comparable. These results show the predictive capabilities of CFD techniques in the analysis of supersonic flow over projectiles with sabots. They also provide an insight into the software capabilities of several of the many tools available to research scientists in the field of CFD.

Probabilistic methods have been implemented with a computational fluid dynamics simulation of fluid flow and heat transfer in a gas turbine engine. The simulation models the temperatures in the T-56 series III 1-2 spacer. Three input quantities are treated as random variables. A random sampling method is used to generate the input values from a probability density function. The implemented Monte Carlo method uses a large number of samples of the input variables to calculate results repeatedly for the output variables (i.e. the predicted temperatures). Statistics of the predicted temperatures, such as the mean and variance, are then calculated. The variation in the predicted temperature at one point in the turbine of the T56 engine is seen to be more sensitive to variability in some parameters than in others.

An algorithm is presented to simulate fluid dynamics on a three qubit type II quantum computer: a lattice of small quantum computers that communicate classical information. The algorithm presented is called a three qubit factorized quantum lattice gas algorithm. It is modeled after classical lattice gas algorithms which move virtual particles along an imaginary lattice and change the particles' momentums using collision rules when they meet at a lattice node. Instead of moving particles, the quantum algorithm presented here moves probabilities, which interact via a unitary collision operator. Probabilities are determined using ensemble measurement and are moved with classical communications channels. The lattice node spacing is defined to be a microscopic scale length. A mesoscopic governing equation for the lattice is derived for the most general three qubit collision operator which preserves particle number. In the continuum limit of the lattice, a governing macroscopic partial differential equation-the diffusion equation-is derived for a particular collision operator using a Chapman-Enskog expansion. A numerical simulation of the algorithm is carried out on a conventional desktop computer and compared to the analytic solution of the diffusion equation. The simulation agrees very well with the known solution.

Image L88-6867 is available as an electronic file from the photo lab. See URL. -- Color graphic CAD CFD model of the National Aerospace Plane.

This report describes a computational study undertaken to investigate the performance of the CFD++ flow solver for prediction of nonlinear aerodynamics of a complex finned missile using structured hexahedral and unstructured tetrahedral grids. A monoplane fixed-wing missile with elliptic cross sections provided a geometrically complex model. Numerical solutions were obtained for this configuration at supersonic speed for various roll orientations, angles of attack, and jaw angles. Steady-state solutions were obtained using a three-dimensional Reynolds-Averaged Navier-Stokes solver with a two-equation turbulence model. Numerical results show the qualitative features of the flow fields at various cross-sectional and streamwise positions along the computational model of the missile. Aerodynamic coefficients were extracted from the computed solutions and found to match well with the available experimental data for these configurations. These numerical results show the effectiveness of using computational fluid dynamics techniques to produce an accurate prediction of the aerodynamics of geometrically complete configurations.

The three-dimensional, unsteady computations of an Unmanned Underwater vehicle with flapping fins were carried out. Several parametric studies were performed varying the amplitude and frequency of oscillation of the fin and the angle of attack of the fin at the root. The objective of these computations was to investigate the importance of these parameters on the fluid dynamics of force production in order to propel the vehicle at a constant speed of 3 kt. An unstructured grid-based unsteady Navier-Stokes solver with automatic adaptive remeshing was used to compute the flow about the vehicle through several complete cycles of fin oscillation for each of the cases studied. The computations show that the angle of attack of 20 of the root section of the fin is near optimum. The vehicle is capable of sustaining a 3 kt current and maintaining position with the fins flapping at a frequency of 2 Hz and an amplitude of 11 4%. As the frequency of oscillation is increased, the net thrust produced increases, but the vehicle will be subjected to large excursions in the normal force. As the fin is made rigid, there is a substantial penalty in lift during the upstroke.

The three-dimensional, unsteady computations of an Unmanned Underwater vehicle with flapping fins were carried out. Several parametric studies were performed varying the amplitude and frequency of oscillation of the fin and the angle of attack of the fin at the root. The objective of these computations was to investigate the importance of these parameters on the fluid dynamics of force production in order to propel the vehicle at a constant speed of 3 kt. An unstructured grid-based unsteady Navier-Stokes solver with automatic adaptive remeshing was used to compute the flow about the vehicle through several complete cycles of fin oscillation for each of the cases studied. The computations show that the angle of attack of 20 of the root section of the fin is near optimum. The vehicle is capable of sustaining a 3 kt current and maintaining position with the fins flapping at a frequency of 2 Hz and an amplitude of 11 4%. As the frequency of oscillation is increased, the net thrust produced increases, but the vehicle will be subjected to large excursions in the normal force. As the fin is made rigid, there is a substantial penalty in lift during the upstroke.

The Air Force SEEK EAGLE Office (AFSEO), Eglin Air Force Base (AFB), FL, is the United States Air Force (USAF) authority for weapons certification efforts. AFSEO performs test and evaluation for aircraft/store compatibility certification and uses Computational Fluid Dynamics (CFD) to support this process. Determining the flow about an aircraft/store combination can be extremely difficult. Complicated geometry such as pylons, launchers, and internal weapons bays can create severe acoustic and aerothermodynamic environments, which are challenging to numerically simulate. The additional challenge of rapidly and accurately simulating the trajectory of a store separation in a high-volume simulation environment is beyond the capabilities of most CFD programs. The USAF requirement for numerous, simultaneous and quick-reaction solutions for a wide variety of stores and aircraft can only be accomplished through application of parallel high-performance computing resources that meet the significant computational and memory demands of the various cases.

A summary of the 2nd Japan-US Symposium on Finite Element Methods in Large Scale Computational Fluid Dynamics, conducted Mar. 14-16, 1994 at Chuo University in Tokyo, Japan is presented. Abstracts of all symposium presentations are included. This report is based upon information collected via symposium attendance, review of the symposium proceedings, and a post symposium conversation with Professor Mutsuto Kawahara, Department of Civil Engineering, Chuo University, who served as the Japanese co-chair of the symposium organizing committee. (AN)

A summary of the 2nd Japan-US Symposium on Finite Element Methods in Large Scale Computational Fluid Dynamics, conducted Mar. 14-16, 1994 at Chuo University in Tokyo, Japan is presented. Abstracts of all symposium presentations are included. This report is based upon information collected via symposium attendance, review of the symposium proceedings, and a post symposium conversation with Professor Mutsuto Kawahara, Department of Civil Engineering, Chuo University, who served as the Japanese co-chair of the symposium organizing committee. (AN)

A summary of the 2nd Japan-US Symposium on Finite Element Methods in Large Scale Computational Fluid Dynamics, conducted Mar. 14-16, 1994 at Chuo University in Tokyo, Japan is presented. Abstracts of all symposium presentations are included. This report is based upon information collected via symposium attendance, review of the symposium proceedings, and a post symposium conversation with Professor Mutsuto Kawahara, Department of Civil Engineering, Chuo University, who served as the Japanese co-chair of the symposium organizing committee. (AN)

Channel flows have three components of velocity, 1component in flow direction and 2components in spanwise of Channel. Because of heterogeneity of velocity undulations, in the channel section the vortices are formed that are nominated “Secondary Current Cells”. These cells cause the lateral shear stress in the channel walls and introduce the erosion and scour of the channel walls. Heterogeneity of the turbulency is caused due to boundary conditions of the bed, side wall and free surface, the aspect ratio and the geometry of the channel. The mechanism of secondary currents can be expressed by longitudinal vorticity equation. In this research, the flow in trapezoidal channels has been modeled using the “Computational Fluid Dynamics” (CFD) with “ANSYS CFX” soft ware, and additional to investigation of sufficiency of CFD in modeling of secondary flow cells, depth averaged velocity and boundary shear stress distributions are perused. For this purpose, the experimental model of Tominaga, and analytical model of Shiono and knight has been used to verify the numerical results. The numerical results are in agreement with the experimental and analytical models.

A system and method for generating fluid flow parameter data for use in aerodynamic heating analysis. Computational fluid dynamics data is generated for a number of points in an area on a surface to be analyzed. Sub-areas corresponding to areas of the surface for which an aerodynamic heating analysis is to be performed are identified. A computer system automatically determines a sub-set of the number of points corresponding to each of the number of sub-areas and determines a value for each of the number of sub-areas using the data for the sub-set of points corresponding to each of the number of sub-areas. The value is determined as an average of the data for the sub-set of points corresponding to each of the number of sub-areas. The resulting parameter values then may be used to perform an aerodynamic heating analysis.

Computational fluid dynamics (CFD) is a  branch of fluid mechanics  that uses numerical analysis and data structures to analyze and solve problems that involve fluid flows.

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Using the commercially available FLUENT 3-D flow field solver, this research effort investigated vortex breakdown over a delta wing at high angle of attack (a) in preparation for investigation of active control of vortex breakdown using steady, along- core blowing A flat delta-shaped half-wing with sharp leading edge and sweep angle of 600 was modeled at a 180 in a wind tunnel at Mach 0,04 and Reynolds number of 3,4 x 10(sub 5). A hybrid (combination of structured and unstructured) numerical mesh was generated to accommodate blowing ports on the wing surface. Results for cases without and with along-core blowing included comparison of various turbulence models for predicting both flow field physics and quantitative flow characteristics, FLUENT turbulence models included Spalart-Allmaras (S-A), Renormalization Group k-e, Reynolds Stress (RSM), and Large Eddy Simulation (LES), as well as comparison with laminar and inviscid models. Mesh independence was also investigated, and solutions were compared with experimentally determined results and theoretical prediction, These research results show that, excepting the LES model for which the computational mesh was insufficiently refined and which was not extensively investigated, none of the turbulence models above, as implemented with the given numerical grid, generated a solution which was suitably comparable to the experimental data. Much more work is required to find a suitable combination of numerical grid and turbulence model.

This collection of papers was presented at the Computational Fluid Dynamics (CFD) Conference held at Ames Research Center in California on March 12 through 14, 1991. It is an overview of CFD activities at NASA Lewis Research Center. The main thrust of computational work at Lewis is aimed at propulsion systems. Specific issues related to propulsion CFD and associated modeling will also be presented. Examples of results obtained with the most recent algorithm development will also be presented.

Computational Fluid Dynamics (CFD) has been used in recent applications to affect subcomponent designs in liquid propulsion rocket engines. This paper elucidates three such applications for turbine stage, pump stage, and combustor chamber geometries. Details of these applications include the development of a high turning airfoil for a gas generator (GG) powered, liquid oxygen (LOX) turbopump, single-stage turbine using CFD as an integral part of the design process. CFD application to pump stage design has emphasized analysis of inducers, impellers, and diffuser/volute sections. Improvements in pump stage impeller discharge flow uniformity have been seen through CFD optimization on coarse grid models. In the area of combustor design, recent CFD analysis of a film cooled ablating combustion chamber has been used to quantify the interaction between film cooling rate, chamber wall contraction angle, and geometry and their effects of these quantities on local wall temperature. The results are currently guiding combustion chamber design and coolant flow rate for an upcoming subcomponent test. Critical aspects of successful integration of CFD into the design cycle includes a close-coupling of CFD and design organizations, quick turnaround of parametric analyses once a baseline CFD benchmark has been established, and the use of CFD methodology and approaches that address pertinent design issues. In this latter area, some problem details can be simplified while retaining key physical aspects to maintain analytical integrity.