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At least 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Demonstration of an automated CFD system for three-dimensional flow simulations

In this paper the capabilities of an automated CFD system which is currently available at NLR are demonstrated. Transonic flow around the AS28G wing/body configuration and hypersonic flow through a generic three-dimensional mixed-compression airbreathing inlet are simulated. An assessment of the level of automation of the current CFD-system is made. The problem-turnaround time lies within the order of a week for both applications.

Vanderburg, J. W.↗
NASA NTRS2007

Three-Dimensional Navier-Stokes Calculations Using the Modified Space-Time CESE Method

The space-time conservation element solution element (CESE) method is modified to address the robustness issues of high-aspect-ratio, viscous, near-wall meshes. In this new approach, the dependent variable gradients are evaluated using element edges and the corresponding neighboring solution elements while keeping the original flux integration procedure intact. As such, the excellent flux conservation property is retained and the new edge-based gradients evaluation significantly improves the robustness for high-aspect ratio meshes frequently encountered in three-dimensional, Navier-Stokes calculations. The order of accuracy of the proposed method is demonstrated for oblique acoustic wave propagation, shock-wave interaction, and hypersonic flows over a blunt body. The confirmed second-order convergence along with the enhanced robustness in handling hypersonic blunt body flow calculations makes the proposed approach a very competitive CFD framework for 3D Navier-Stokes simulations.

Chang, Chau-lyan↗
NASA NTRS1987

Application of an upwind algorithm to the three-dimensional parabolized Navier-Stokes equations

A new computer code for the solution of the three-dimensional parabolized Navier-Stokes equations has been developed. The code employs a state-of-the-art upwind algorithm to capture strong shock waves. The algorithm is implicit, uses finite volumes, and is second-order accurate in the crossflow directions. The new code is validated through application to laminar hypersonic flows past two simple body shapes: a circular cone of 10 deg half-angle, and a generic all-body hypersonic vehicle. Cone flow solutions were computed at angles of attack of 12, 20, and 24 deg and results are in agreement with experimental data. Results are also presented for the flow past the all-body vehicle at angles of incidence of 0 and 10 deg.

Lawrence, Scott L.↗
NASA NTRS1948

On similarity rules for transonic flows

A method used by Tsien to derive similarity rules for hypersonic flows is utilized to derive Von Karman's similarity rules for transonic flows. At the lower limit of the transonic region of flow the theory yields a formula for the critical stream Mach numbers of a given family of symmetrical profiles. It is further shown that this formula can also be obtained by means of the Prandtl-Glauert small-perturbation method. Investigation of the behavior of the similarity parameter in the region where the thickness coefficient approaches zero and the critical stream Mach number approaches unity shows that it possesses a limiting value characteristic of the prescribed family of shapes.

Kaplan, Carl↗
NASA NTRS1989

The solution of the Navier-Stokes equations using Gauss-Seidel line relaxation

The Navier-Stokes equations in an implicit flux-split difference formulation are solved numerically using a Gauss-Seidel line-relaxation procedure. Particular attention is given to the selection of flux-vector splitting method and flux splitting in boundary layers. Results for sample problems involving (1) turbulent supersonic flow over a cone and (2) the viscous hypersonic flow of a chemically reacting gas in thermal nonequilibrium past a blunted cone are presented in extensive graphs and briefly characterized. The present flux-split procedures are shown to provide accurate shear-layer calculations.

Maccormack, Robert W.↗
NASA NTRS1991

A numerical investigation of turbulent baseflow

The turbulent base flow of a cone with a sting in hypersonic flow is investigated using a time-dependent Navier-Stokes code called UWIN. A PNS solver provides the upstream flow conditions which drive the development of the base flow. The computational method is described along with its application to the base flow problem. The grid generation and boundary conditions are described. The results and the summary conclude this note.

Chaussee, Denny S.↗
NASA NTRS1992

Three-dimensional boundary layer calculation by a characteristic method

A numerical method for solving the three-dimensional boundary layer equations for bodies of arbitrary shape is presented. In laminar flows, the application domain extends from incompressible to hypersonic flows with the assumption of chemical equilibrium. For turbulent boundary layers, the application domain is limited by the validity of the mixing length model used. In order to respect the hyperbolic nature of the equations reduced to first order partial derivative terms, the momentum equations are discretized along the local streamlines using of the osculator tangent plane at each node of the body fitted coordinate system. With this original approach, it is possible to overcome the use of the generalized coordinates, and therefore, it is not necessary to impose an extra hypothesis about the regularity of the mesh in which the boundary conditions are given. By doing so, it is possible to limit, and sometimes to suppress, the pre-treatment of the data coming from an inviscid calculation. Although the proposed scheme is only semi-implicit, the method remains numerically very efficient.

Houdeville, R.↗
NASA NTRS1993

A time-accurate finite volume method valid at all flow velocities

A finite volume method to solve the Navier-Stokes equations at all flow velocities (e.g., incompressible, subsonic, transonic, supersonic and hypersonic flows) is presented. The numerical method is based on a finite volume method that incorporates a pressure-staggered mesh and an incremental pressure equation for the conservation of mass. Comparison of three generally accepted time-advancing schemes, i.e., Simplified Marker-and-Cell (SMAC), Pressure-Implicit-Splitting of Operators (PISO), and Iterative-Time-Advancing (ITA) scheme, are made by solving a lid-driven polar cavity flow and self-sustained oscillatory flows over circular and square cylinders. Calculated results show that the ITA is the most stable numerically and yields the most accurate results. The SMAC is the most efficient computationally and is as stable as the ITA. It is shown that the PISO is the most weakly convergent and it exhibits an undesirable strong dependence on the time-step size. The degenerated numerical results obtained using the PISO are attributed to its second corrector step that cause the numerical results to deviate further from a divergence free velocity field. The accurate numerical results obtained using the ITA is attributed to its capability to resolve the nonlinearity of the Navier-Stokes equations. The present numerical method that incorporates the ITA is used to solve an unsteady transitional flow over an oscillating airfoil and a chemically reacting flow of hydrogen in a vitiated supersonic airstream. The turbulence fields in these flow cases are described using multiple-time-scale turbulence equations. For the unsteady transitional over an oscillating airfoil, the fluid flow is described using ensemble-averaged Navier-Stokes equations defined on the Lagrangian-Eulerian coordinates. It is shown that the numerical method successfully predicts the large dynamic stall vortex (DSV) and the trailing edge vortex (TEV) that are periodically generated by the oscillating airfoil. The calculated streaklines are in very good comparison with the experimentally obtained smoke picture. The calculated turbulent viscosity contours show that the transition from laminar to turbulent state and the relaminarization occur widely in space as well as in time. The ensemble-averaged velocity profiles are also in good agreement with the measured data and the good comparison indicates that the numerical method as well as the multipletime-scale turbulence equations successfully predict the unsteady transitional turbulence field. The chemical reactions for the hydrogen in the vitiated supersonic airstream are described using 9 chemical species and 48 reaction-steps. Consider that a fast chemistry can not be used to describe the fine details (such as the instability) of chemically reacting flows while a reduced chemical kinetics can not be used confidently due to the uncertainty contained in the reaction mechanisms. However, the use of a detailed finite rate chemistry may make it difficult to obtain a fully converged solution due to the coupling between the large number of flow, turbulence, and chemical equations. The numerical results obtained in the present study are in good agreement with the measured data. The good comparison is attributed to the numerical method that can yield strongly converged results for the reacting flow and to the use of the multiple-time-scale turbulence equations that can accurately describe the mixing of the fuel and the oxidant.

Kim, S.-W.↗
NASA NTRS1998

Computational Analysis of a Laminar Jet in a Cross-Flow

A jet in cross-flow (JIFC) consists of a jet exhausting at a large angle into a freestream flow. It is a flow field which is relevant to a wide variety of technologies and applications. Despite the nearly 65 years of JIFC research there are few results available for laminar hypersonic flows, a combination which will be encountered by re-entry and high altitude vehicles over some portion of their flight path. This research consists of developing a numerical model to investigate the interaction of a normal sonic jet exhausting into a hypersonic cross-flow. The model was validated by comparing experimental measurements with corresponding numerical results generated by the model.

Jeffries, Damon K.↗
NASA NTRS1990

Shock tunnel studies of scramjet phenomena, supplement 5

A series of reports are presented on SCRAMjet studies, shock tunnel studies, and expansion tube studies. The SCRAMjet studies include: (1) Investigation of a Supersonic Combustion Layer; (2) Wall Injected SCRAMjet Experiments; (3) Supersonic Combustion with Transvers, Circular, Wall Jets; (4) Dissociated Test Gas Effects on SCRAMjet Combustors; (5) Use of Silane as a Fuel Additive for Hypersonic Thrust Production, (6) Pressure-length Correlations in Supersonic Combustion; (7) Hot Hydrogen Injection Technique for Shock Tunnels; (8) Heat Release - Wave Interaction Phenomena in Hypersonic Flows; (9) A Study of the Wave Drag in Hypersonic SCRAMjets; (10) Parametric Study of Thrust Production in the Two Dimensional SCRAMjet; (11) The Design of a Mass Spectrometer for use in Hypersonic Impulse Facilities; and (12) Development of a Skin Friction Gauge for use in an Impulse Facility. The shock tunnel studies include: (1) Hypervelocity flow in Axisymmetric Nozzles; (2) Shock Tunnel Development; and (3) Real Gas Efects in Hypervelocity Flows over an Inclined Cone. The expansion tube studies include: (1) Investigation of Flow Characteristics in TQ Expansion Tube; and (2) Disturbances in the Driver Gas of a Shock Tube.

Casey, R.↗
NASA NTRS2004

Evaluation of a CFD Method for Aerodynamic Database Development using the Hyper-X Stack Configuration

A computational fluid dynamic (CFD) study is performed on the Hyper-X (X-43A) Launch Vehicle stack configuration in support of the aerodynamic database generation in the transonic to hypersonic flow regime. The main aim of the study is the evaluation of a CFD method that can be used to support aerodynamic database development for similar future configurations. The CFD method uses the NASA Langley Research Center developed TetrUSS software, which is based on tetrahedral, unstructured grids. The Navier-Stokes computational method is first evaluated against a set of wind tunnel test data to gain confidence in the code s application to hypersonic Mach number flows. The evaluation includes comparison of the longitudinal stability derivatives on the complete stack configuration (which includes the X-43A/Hyper-X Research Vehicle, the launch vehicle and an adapter connecting the two), detailed surface pressure distributions at selected locations on the stack body and component (rudder, elevons) forces and moments. The CFD method is further used to predict the stack aerodynamic performance at flow conditions where no experimental data is available as well as for component loads for mechanical design and aero-elastic analyses. An excellent match between the computed and the test data over a range of flow conditions provides a computational tool that may be used for future similar hypersonic configurations with confidence.

Parikh, Paresh↗
NASA NTRS1972

On the structure of hypersonic turbulent boundary layers.

Extensive hot-wire auto- and cross-correlation measurements obtained in a fully developed compressible turbulent boundary layer are presented. A tentative mechanism of turbulence production and growth in hypersonic flow suggested by these measurements is developed. This flow model is consistent with previous observations in incompressible flows. Detailed measurements of the mean properties of the hypersonic turbulent boundary layer are also presented and compared with results from various transformation and finite-difference prediction methods. It is shown that none of the theories predict all the properties of the hypersonic turbulent boundary layer and that additional measurements are needed to provide more adequate physics of turbulent processes for use in the various theories.

Owen, F. K.↗
NASA NTRSDate not supplied

Understanding the Role of CFD, Including RANS, DES, LES and DNS, in Hypersonic Propulsion Design and Analysis

Computational fluid dynamics (CFD) has matured to the point where it is becoming ubiquitous in the design and analysis of aerospace vehicles and their propulsion systems. The current state-of-the-art of CFD in hypersonic propulsion design and analysis as well as a historical perspective of its development over the last 50 years is presented. The current state-of-the-art for the design of finite rate chemically reacting turbulent hypersonic flow is shown to currently be Reynolds averaged simulations (RAS) where the Reynolds averaged Navier-Stokes (RAS) equations are solved to obtain steady state solutions. An overview of the physical modeling techniques employed by RAS for the simulation of hypersonic chemically reacting flow are presented. However, scale resolving simulation techniques, such as detached eddy simulation (DES), large eddy simulations (LES), hybrid RAS/LES and direct numerical simulations (DNS), are becoming becoming powerful tools for point design analysis. One example of such a scale resolving technique, hybrid RAS/LES, will be briefly described as well as how it can be used to help address model form uncertainties in RAS.

Hypersonic↗
NASA NTRS1988

Experimental program for real gas flow code validation at NASA Ames Research Center

The experimental program for validating real gas hypersonic flow codes at NASA Ames is described. Ground based test facilities used include ballistic ranges, shock tubes and shock tunnels, arcjet facilities and heated air hypersonic wind tunnels. Also included are large scale computer systems for kinetic theory simulations and benchmark code solutions. Flight tests consist of the Aeroassist Flight Experiment, the Space Shuttle, Project Fire 2, and planetary probes such as Galileo, Pioneer Venus and PAET.

Deiwert, George S.↗
NASA NTRS1989

Experimental program for real gas flow code validation at NASA Ames Research Center

The experimental program for validating real gas hypersonic flow codes at NASA Ames Rsearch Center is described. Ground-based test facilities used include ballistic ranges, shock tubes and shock tunnels, arc jet facilities and heated-air hypersonic wind tunnels. Also included are large-scale computer systems for kinetic theory simulations and benchmark code solutions. Flight tests consist of the Aeroassist Flight Experiment, the Space Shuttle, Project Fire 2, and planetary probes such as Galileo, Pioneer Venus, and PAET.

Deiwert, George S.↗
NASA NTRS1987

Engineering analysis of slender-body aerodynamics using Sychev similarity parameters

The similarity parameters deduced by Sychev for inviscid hypersonic flow over slender bodies are reviewed and used to correlate flowfield, surface-pressure, normal-force and center-of-pressure data for supersonic flow over thin slender wings and smooth slender bodies at low-to-high angles of attack. Although Sychev expected similarity to hold only for hypersonic freestream and cross flows, it is demonstrated empirically that similarity holds for any value of cross-flow Mach number if the axial flow component is supersonic. It is also shown for thin wings that similarity holds for much larger values of aspect ratio than Sychev supposed. A one-term power-law expression is found to fit all of the normal-force correlations which suggests that it may be possible to develop a simple semi-empirical method for estimating the aerodynamic characteristics of arbitrary slender airframes.

Hemsch, Michael J.↗