Search NASA⌕ Search

SEARCH · Search NASA

Results for “Ramp compression”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Mach 6.5 air induction system design for the Beta 2 two-stage-to-orbit booster vehicle

A preliminary, two-dimensional, mixed compression air induction system is designed for the Beta II Two Stage to Orbit booster vehicle to minimize installation losses and efficiently deliver the required airflow. Design concepts, such as an external isentropic compression ramp and a bypass system were developed and evaluated for performance benefits. The design was optimized by maximizing installed propulsion/vehicle system performance. The resulting system design operating characteristics and performance are presented. The air induction system design has significantly lower transonic drag than similar designs and only requires about 1/3 of the bleed extraction. In addition, the design efficiently provides the integrated system required airflow, while maintaining adequate levels of total pressure recovery. The excellent performance of this highly integrated air induction system is essential for the successful completion of the Beta II booster vehicle mission.

Midea, Anthony C.↗
NASA NTRS1992

Comparison between computational and experimental data for a hypersonic laser propelled vehicle

A 3-D axisymmetric hypersonic engine inlet was investigated using PARC2D, an ideal gas Computational Fluid Dynamics code. The code was used to predict the results of tests conducted in the Rensselaer Polytechnic Institute Hypersonic Shock Tunnel which measured surface and pitot pressures, and shock positions (through Schlieren photography) at freestream Mach numbers of 10, 13, and 15. A strong viscous/shock interaction was observed in both the experiment and the CFD results, due to the model's parabolic compression ramp. Good agreement was found between the experimental results and the CFD solution both for surface pressures and shock positions. Agreement between pitot pressures was less reliable.

Messitt, D. G.↗
NASA NTRS1993

Supersonic boundary-layer flow turbulence modeling

Baldwin-Lomax and kappa-epsilon turbulence models were modified for use in Navier-Stokes numerical computations of Mach 2.9 supersonic turbulent boundary layer flows along compression ramps. The computational results of Reynolds shear stress profiles were compared with experimental data. The Baldwin-Lomax model was modified to account for the Reynolds shear stress amplification within the flow field. A hybrid kappa-epsilon model with viscous sublayer turbulence treatment was constructed to predict the Reynolds shear stress profiles within the entire flow field. These modified turbulence models were effective for the computations of the surface pressure and the skin friction factor variations along an 8 deg ramp surface. The hybrid kappa-epsilon model could improve the predictions of the Reynolds shear stress profile and the skin friction factor near the corner of a 16 deg ramp.

Wang, Chi-Rong↗
NASA NTRS1992

An effective multigrid method for high-speed flows

The use is considered of a multigrid method with central differencing to solve the Navier-Stokes equations for high speed flows. The time dependent form of the equations is integrated with a Runge-Kutta scheme accelerated by local time stepping and variable coefficient implicit residual smoothing. Of particular importance are the details of the numerical dissipation formulation, especially the switch between the second and fourth difference terms. Solutions are given for 2-D laminar flow over a circular cylinder and a 15 deg compression ramp.

Swanson, R. C.↗
NASA NTRS1992

An implicit multigrid scheme for hypersonic strong-interaction flowfields

An implicit multigrid method for the compressible Navier-Stokes equations is applied to a series of two-dimensional, steady, laminar hypersonic flows over a compression ramp, including both attached and separated flow conditions. The algorithm uses upwind-biased differencing for the convective and pressure terms and central differencing for the shear-stress and heat-transfer terms. An implicit spatially factored scheme is used to advance the solution in terms on a given mesh, and a V-cycle multigrid strategy is used to accelerate convergence. The multigrid algorithm provided an order of magnitude decrease in the computational time required to obtain a converged position of the separation location and enabled a substantially grid-independent result to be obtained for each of the flows considered.

Thomas, James L.↗
NASA NTRS1993

Aspects of a high-resolution scheme for the Navier-Stokes equations

In this paper we emphasize the importance of the form of the numerical dissipation model in computing accurate viscous flow solutions. A high-resolution scheme for viscous flows based on three-point central differencing and a matrix dissipation is considered. The various components of this scheme, including 'entropy fix', limiter function, and boundary-point dissipation are discussed. By analyzing boundary-point dissipation stencils, we confirm that with the matrix dissipation model the normal numerical dissipation terms in the streamwise momentum equation are independent of the Reynolds number. Such independence is not achieved with a scalar dissipation form. The accuracy of the central-difference scheme, with and without matrix dissipation, and the flux-difference split scheme of Roe, which is classified as a high-resolution scheme, is compared. For this comparison, three high Reynolds number laminar flows are considered. Solutions of the Navier-Stokes equations are obtained for low-speed flow over a flat plate, transonic flow over an airfoil with transition near the leading edge, and hypersonic flow over a compression ramp. The emphasis of the comparison is primarily on the details of the viscous flows. The necessity of the high-resolution property is revealed.

Swanson, R. C.↗
NASA NTRS1993

Aerodynamic heating in the vicinity of hypersonic, axisymmetric, shock-wave boundary-layer interactions

Aerodynamic heating rates are calculated from time-dependent temperature measurements in the vicinity of shock-wave boundary-layer interactions due to conical compression ramps on an axisymmetric body. The data were acquired at the Ohio State University Aeronautical and Astronautical Research Laboratory and at the Air Force Flight Dynamics Laboratory at Mach numbers of 6 and 10. The model is a cylindrical body with a 10 deg conical nose. Conical ramps with half-angles of 10, 20, 25, 30, and 35 deg serve as shock-wave generators. Flowfield surveys are made in the vicinity of the ramp vertices, separation points, and reattachment points. Experimental results quantify temperature response and the resulting heat transfer rates as a function of ramp angle, Reynolds number and freestream Mach number. The temperature responses within the flowfield appear to be steady-state for all angles and all Reynolds numbers, and hence, the heat transfer rates appear to be steady-state.

Flanagan, M. J., Jr.↗
NASA NTRS1952

Analytical and Experimental Investigation of Inlet-engine Matching for Turbojet-powered Aircraft at Mach Numbers up to 2.0

An analysis of inlet-turbojet-engine matching for a range of Mach numbers up to 2.0 indicates large performance penalties when fixed-geometry inlets are used. Use of variable-geometry inlets, however, nearly eliminates th The analysis was confirmed experimentally by investigating at Mach numbers of 0, 0.63, and 1.5 to 2.0 two single oblique-shock-type inlets of different compression-ramp angles, which simulated a variable-geometry configuration. The experimental investigation indicated that total-pressure recoveries comparable withose attainable with well designed nose inlets were obtained with the side inlets when all the boundary layer ahead of the inlets was removed. Serious drag penalties resulted at a Mach number of 2.0 from the use of blunt-cowl leading edges. However, sharp-lip inlets produced large losses in thrust for the take-off condition. These thrust penalties which are associated with the the low-speed operation of the sharp-lip inlet designs can probably be avoided without impairing the supersonic performance of the inlet by the use of auxiliary inlets or blow-in doors.

Esenwein, Fred T↗
NASA NTRS1956

Performance of variable two-dimensional inlet designed for engine-inlet matching I : performance at design Mach number of 3.07

The inlet was investigated with both a two-oblique-shock and an isentropic compression ramp. Provisions for a variety of bypass systems were incorporated and none decreased pressure recovery by more than about 1 percent. By utilizing throat boundary-layer control, critical pressure recoveries of 64 and 70 percent were obtained with the two-shock and isentropic ramps, respectively. Diffuser-exit air distortion was about 5 percent at critical operation Theoretical cowl pressure drag was equal to about 10 percent of the net thrust of an assumed engine.

Beheim, M A↗
NASA NTRS1993

Unsteady Phenomena in Shock Wave/Boundary Layer Interaction

A brief review is given of the unsteadiness of shock wave/turbulent boundary layer interaction. The focus is on interactions generated by swept and unswept compression ramps, by flares, steps and incident shock waves, by cylinders and blunt fins, and by glancing shock waves. The effects of Mach number, Reynolds number, and separated flow scale are discussed as are the physical causes of the unsteadiness. The implications that the unsteadiness has for interpreting time-average surface and flowfield data, and for comparisons of such experimental data with computation, is also briefly discussed. Finally, some suggestions for future work are given. It is clear that there are large gaps in the data base and that many aspects of such phenomena are poorly understood. Much work remains to be done.

Dolling, D. S.↗
NASA NTRS1994

An Interactive, Design and Educational Tool for Supersonic External-Compression Inlets

A workstation-based interactive design tool called VU-INLET was developed for the inviscid flow in rectangular, supersonic, external-compression inlets. VU-INLET solves for the flow conditions from free stream, through the supersonic compression ramps, across the terminal normal shock region and the subsonic diffuser to the engine face. It calculates the shock locations, the capture streamtube, and the additive drag of the inlet. The inlet geometry can be modified using a graphical user interface and the new flow conditions recalculated interactively. Free stream conditions and engine airflow can also be interactively varied and off-design performance evaluated. Flow results from VU-INLET can be saved to a file for a permanent record, and a series of help screens make the simulator easy to learn and use. This paper will detail the underlying assumptions of the models and the numerical methods used in the simulator.

Benson, Thomas J.↗
NASA NTRS1994

Computational Analysis in Support of the SSTO Flowpath Test

A synergistic approach of combining computational methods and experimental measurements is used in the analysis of a hypersonic inlet. There are four major focal points within this study which examine the boundary layer growth on a compression ramp upstream of the cowl lip of a scramjet inlet. Initially, the boundary layer growth on the NASP Concept Demonstrator Engine (CDE) is examined. The follow-up study determines the optimum diverter height required by the SSTO Flowpath test to best duplicate the CDE results. These flow field computations are then compared to the experimental measurements and the mass average Mach number is determined for this inlet.

Duncan, Beverly S.↗
NASA NTRS1994

Fluctuating wall pressures near separation in highly swept turbulent interactions

Fluctuating wall pressures have been measured in Mach 5 interactions generated by sharp, upswept fins at angles of attack of 16 to 28 deg. The results show that rms pressure distributions, like the mean, can be collapsed in conical coordinates. The wall pressure signal near separation is intermittent and is qualitatively similar to that measured in unswept interactions and other swept flows. However, the dominant separation shock frequencies in the swept flows are up to an order of magnitude higher than those in unswept interactions under identical incoming flow conditions. In light of the present observations, it appears that an earlier remark, by Gilson and Dolling, that separation is characterized by a shuddering compression system, in contrast to a translating separation shock is erroneous due to a combination of weak shock strength and inadequate spatial resolution. Furthermore, although the maximum rms pressure near separation increases with increasing interaction strength, as in swept compression ramp flows, comparison of data from the two flow types indicates that the appropriate correlating parameter is the interaction sweepback angle.

Schmisseur, J. D.↗
NASA NTRS1993

Direct Monte Carlo Simulations of Hypersonic Viscous Interactions Including Separation

Results of calculations obtained using the direct simulation Monte Carlo method for Mach 25 flow over a control surface are presented. The numerical simulations are for a 35-deg compression ramp at a low-density wind-tunnel test condition. Calculations obtained using both two- and three-dimensional solutions are reviewed, and a qualitative comparison is made with the oil flow pictures highlight separation and three-dimensional flow structure.

Moss, James N.↗
NASA NTRS1997

Annual Research Briefs

This report contains the 1997 annual progress reports of the research fellows and students supported by the Center for Turbulence Research (CTR). Titles include: Invariant modeling in large-eddy simulation of turbulence; Validation of large-eddy simulation in a plain asymmetric diffuser; Progress in large-eddy simulation of trailing-edge turbulence and aeronautics; Resolution requirements in large-eddy simulations of shear flows; A general theory of discrete filtering for LES in complex geometry; On the use of discrete filters for large eddy simulation; Wall models in large eddy simulation of separated flow; Perspectives for ensemble average LES; Anisotropic grid-based formulas for subgrid-scale models; Some modeling requirements for wall models in large eddy simulation; Numerical simulation of 3D turbulent boundary layers using the V2F model; Accurate modeling of impinging jet heat transfer; Application of turbulence models to high-lift airfoils; Advances in structure-based turbulence modeling; Incorporating realistic chemistry into direct numerical simulations of turbulent non-premixed combustion; Effects of small-scale structure on turbulent mixing; Turbulent premixed combustion in the laminar flamelet and the thin reaction zone regime; Large eddy simulation of combustion instabilities in turbulent premixed burners; On the generation of vorticity at a free-surface; Active control of turbulent channel flow; A generalized framework for robust control in fluid mechanics; Combined immersed-boundary/B-spline methods for simulations of flow in complex geometries; and DNS of shock boundary-layer interaction - preliminary results for compression ramp flow.

Spinks, Debra↗
NASA NTRS2001

Hypersonic Boundary Layer Instability Over a Corner

A boundary-layer transition study over a compression corner was conducted under a hypersonic flow condition. Due to the discontinuities in boundary layer flow, the full Navier-Stokes equations were solved to simulate the development of disturbance in the boundary layer. A linear stability analysis and PSE method were used to get the initial disturbance for parallel and non-parallel flow respectively. A 2-D code was developed to solve the full Navier-stokes by using WENO(weighted essentially non-oscillating) scheme. The given numerical results show the evolution of the linear disturbance for the most amplified disturbance in supersonic and hypersonic flow over a compression ramp. The nonlinear computations also determined the minimal amplitudes necessary to cause transition at a designed location.

Balakumar, Ponnampalam↗
NASA NTRS2004

Temporal Large-Eddy Simulation

In 1999, Stolz and Adams unveiled a subgrid-scale model for LES based upon approximately inverting (defiltering) the spatial grid-filter operator and termed .the approximate deconvolution model (ADM). Subsequently, the utility and accuracy of the ADM were demonstrated in a posteriori analyses of flows as diverse as incompressible plane-channel flow and supersonic compression-ramp flow. In a prelude to the current paper, a parameterized temporal ADM (TADM) was developed and demonstrated in both a priori and a posteriori analyses for forced, viscous Burger's flow. The development of a time-filtered variant of the ADM was motivated-primarily by the desire for a unifying theoretical and computational context to encompass direct numerical simulation (DNS), large-eddy simulation (LES), and Reynolds averaged Navier-Stokes simulation (RANS). The resultant methodology was termed temporal LES (TLES). To permit exploration of the parameter space, however, previous analyses of the TADM were restricted to Burger's flow, and it has remained to demonstrate the TADM and TLES methodology for three-dimensional flow. For several reasons, plane-channel flow presents an ideal test case for the TADM. Among these reasons, channel flow is anisotropic, yet it lends itself to highly efficient and accurate spectral numerical methods. Moreover, channel-flow has been investigated extensively by DNS, and a highly accurate data base of Moser et.al. exists. In the present paper, we develop a fully anisotropic TADM model and demonstrate its utility in simulating incompressible plane-channel flow at nominal values of Re(sub tau) = 180 and Re(sub tau) = 590 by the TLES method. The TADM model is shown to perform nearly as well as the ADM at equivalent resolution, thereby establishing TLES as a viable alternative to LES. Moreover, as the current model is suboptimal is some respects, there is considerable room to improve TLES.

Pruett, C. D.↗