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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.

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All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Three-dimensional, shock-on-shock interaction problem

The unsteady, three-dimensional flowfield resulting from the interaction of a plane shock with a cone-shaped vehicle traveling supersonically is determined, using a second-order, shock-capturing, finite-difference approach. The time-dependent, inviscid gasdynamic equations are transformed to include the self-similar property of the flow, to align various coordinate surfaces with known shock waves, and to cluster points in the vicinity of the intersection of the transmitted incident shock and the surface of the vehicle. The governing partial differential equations in conservation-law form are then solved iteratively using MacCormack's (1969) algorithm.

Kutler, P.↗
NASA NTRS1974

On the shock-on-shock interaction problem

The unsteady, inviscid flowfield that results when a supersonic vehicle strikes a planar oblique shock wave, though difficult to simulate experimentally, is quite easy to model and compute numerically. The complicated flowfield, which contains multiple shock wave interactions, is determined using a second-order, shock-capturing, finite-difference approach which solves the time-dependent Euler equations under a self-similar transformation. A series of numerical results for a simple two-dimensional wedge is presented which describes the entire disturbed region, including the wave structure, and shows good agreement with the available two- and three-dimensional experimental data.-

Kutler, P.↗
NASA NTRS2012

Micro-Ramp Flow Control for Oblique Shock Interactions: Comparisons of Computational and Experimental Data

Computational fluid dynamics was used to study the effectiveness of micro-ramp vortex generators to control oblique shock boundary layer interactions. Simulations were based on experiments previously conducted in the 15- by 15-cm supersonic wind tunnel at the NASA Glenn Research Center. Four micro-ramp geometries were tested at Mach 2.0 varying the height, chord length, and spanwise spacing between micro-ramps. The overall flow field was examined. Additionally, key parameters such as boundary-layer displacement thickness, momentum thickness and incompressible shape factor were also examined. The computational results predicted the effects of the microramps well, including the trends for the impact that the devices had on the shock boundary layer interaction. However, computing the shock boundary layer interaction itself proved to be problematic since the calculations predicted more pronounced adverse effects on the boundary layer due to the shock than were seen in the experiment.

Hirt, Stephanie M.↗
NASA NTRS2010

Micro-Ramp Flow Control for Oblique Shock Interactions: Comparisons of Computational and Experimental Data

Computational fluid dynamics was used to study the effectiveness of micro-ramp vortex generators to control oblique shock boundary layer interactions. Simulations were based on experiments previously conducted in the 15 x 15 cm supersonic wind tunnel at NASA Glenn Research Center. Four micro-ramp geometries were tested at Mach 2.0 varying the height, chord length, and spanwise spacing between micro-ramps. The overall flow field was examined. Additionally, key parameters such as boundary-layer displacement thickness, momentum thickness and incompressible shape factor were also examined. The computational results predicted the effects of the micro-ramps well, including the trends for the impact that the devices had on the shock boundary layer interaction. However, computing the shock boundary layer interaction itself proved to be problematic since the calculations predicted more pronounced adverse effects on the boundary layer due to the shock than were seen in the experiment.

Hirt, Stefanie M.↗
NASA NTRS1991

Two and three-dimensional shock-shock interactions on the blunt leading edges of the hypersonic inlets

The effect of shock impingement on the blunt leading edges of the top and sidewall compression type inlet of a scramjet engine is studied numerically. The impinging shock is caused by the vehicle forebody. The interaction of this forebody shock with the inlet leading edge shock results in a very complex flowfield containing local regions of high pressure and intense heating. This complex flowfield in calculated by solving the Navier-Stokes equations using a finite volume flux splitting technique due to van Leer. To resolve the finer details of the flow structure as well as to predict the surface heat transfer accurately, adaptive grid technique is used in the analysis. Results of the present numerical study are compared with available experimental results.

Singh, D. J.↗
NASA NTRS1975

Mean flow field and surface heating produced by unequal shock interactions at hypersonic speeds

Mean velocity profiles were measured in a free shear layer produced by the interaction of two unequal strength shock waves at hypersonic free-stream Mach numbers. Measurements were made over a unit Reynolds number range of 3,770,000 per meter to 17,400,000 per meter based on the flow on the high velocity side of the shear layer. The variation in measured spreading parameters with Mach number for the fully developed flows is consistent with the trend of the available zero velocity ratio data when the Mach numbers for the data given in this study are taken to be characteristic Mach numbers based on the velocity difference across the mixing layer. Surface measurements in the shear-layer attachment region of the blunt-body model indicate peak local heating and static pressure consistent with other published data. Transition Reynolds numbers were found to be significantly lower than those found in previous data.

Birch, S. F.↗
NASA NTRS1988

Viscous hypersonic shock-on-shock interaction on blunt cowl lips

The effects of impinging ramp shocks on a two-dimensional viscous inlet cowl of the National Aerospace Plane (NASP) or blunt body flow fields for various impingement locations are numerically simulated by a recently developed second-order implicit total variation diminishing (TVD) algorithm. The results demonstrate that accurate computation of heat transfer rate is crucially dependent on adequate normal grid resolution at the wall. The various numerical results are compared with recent experimental data. Due to the complex flow patterns that occur for the shock-on-shock flows, adaptive grid procedures are utilized.

Klopfer, G. H.↗
NASA NTRS2007

Application of Micro-ramp Flow Control Devices to an Oblique Shock Interaction

Tests are planned in the 15cm x 15cm supersonic wind tunnel at NASA Glenn to demonstrate the applicability of micro-ramp flow control to the management of shock wave boundary layer interactions. These tests will be used as a database for computational fluid dynamics (CFD) validation and Design of Experiments (DoE) design information. Micro-ramps show potential for mechanically simple and fail-safe boundary layer control.

Hirt, Stefanie↗
NASA NTRS2018

Computational and Experimental Study of Plume and Shock Interaction Effects on Sonic Boom in the NASA Ames 9x7 Supersonic Wind Tunnel

A wind tunnel test was performed in the NASA Ames 9x7 Supersonic Wind Tunnel focusing on the shock waves traveling through and interacting with an exhaust nozzle plume. This experimental study was conducted to develop and validate the CFD capability required to accurately include nozzle flow with impinging shock effects on near field and ground‐propagated sonic boom signatures. The model was made to be generic, and included a simple nozzle shape, two different aft decks, and a few generic horizontal tails. High pressure air was pumped through a nozzle at various nozzle pressure ratios (NPR) to represent the engine plume in flight. The three different aft body representations each created a different shock wave signature that passed through the plume. An aft deck configuration, where part of the aircraft shields the nozzle plume, was also tested. Retroreflective Background-Oriented Schlieren (RBOS) was used to obtain schlieren images of the flow field around the model and behind the model. This study compares wind tunnel data and numerical simulations conducted by the NASA Tetrahedral Unstructured Software System CFD code, USM3D.

Winski, Courtney S.↗
NASA NTRS1986

Plasma waves in the shock interaction regions at Comet Giacobini-Zinner

The nature of the comet-solar wind interaction is studied by analyzing the detailed evolution of the plasma wave spectra of Comet Giacobini-Zinner across the interaction region. Electron heat fluxes and associated electron plasma waves, steepened low-frequency wave packets, and density fluctuations observed upstream of Giacobini-Zinner shocks are also found upstream of quasi-parallel bow shocks. Downstream, the pulsations in the cometary magnetic field magnitude, in addition to the large density spikes, are usually also found downstream of quasi-parallel bow shocks. Other similarities to interplanetary shocks and terrestrial bow shocks are described.

Kennel, C. F.↗
NASA NTRS1992

Internal shock interactions in propulsion/airframe integrated three-dimensional sidewall compression scramjet inlets

The advantages and design requirements of propulsion/airframe integration for high Mach number flight are studied in terms of the 3D sidewall compression scramjet inlet. The present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. The source of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.

Holland, Scott D.↗
NASA NTRS2016

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions

NASA and Industry are performing vehicle studies of configurations with low sonic boom pressure signatures. The computational analyses of modern configuration designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty in the aft signatures with often greater boundary layer effects and nozzle jet pressures. Wind tunnel testing at significantly lower Reynolds numbers than in flight and without inlet and nozzle jet pressures make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel from Mach 1.6 to 2.0 will be used to assess the effects of shocks from components passing through nozzle jet plumes on the sonic boom pressure signature and provide datasets for comparison with CFD codes. A large number of high-fidelity numerical simulations of wind tunnel test models with a variety of shock generators that simulate horizontal tails and aft decks have been studied to provide suitable models for sonic boom pressure measurements using a minimally intrusive pressure rail in the wind tunnel. The computational results are presented and the evolution of candidate wind tunnel models is summarized and discussed in this paper.

Wind↗
NASA NTRS1968

Jet-shock interactions.

Finite difference scheme from method of characteristics to investigate flow field and intersecting exhaust plumes from nozzles

Mourer, J.↗
NASA NTRS1972

Hypersonic shock-interaction phenomena applicable to space shuttle configurations

The convective heat transfer distribution for space shuttle configurations is discussed. The viscous/inviscid interactions associated with the complex three dimensional flow fields are examined. Two basic conditions are considered as follows: (1) models consisting of basic elemental combinations and (2) models of specific flight vehicles. The test facilities and test programs used to obtain data on the fuselage flow field and the wing flow field are described.

Bertin, J. J.↗
NASA NTRS2004

DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments

The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolution, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Moss, James N.↗
NASA NTRS2004

DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments

The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolutions, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Moss, James N.↗
NASA NTRS1993

Supersonic shock wave/vortex interaction

Although shock wave/vortex interaction is a basic and important fluid dynamics problem, very little research has been conducted on this topic. Therefore, a detailed experimental study of the interaction between a supersonic streamwise turbulent vortex and a shock wave was carried out at the Penn State Gas Dynamics Laboratory. A vortex is produced by replaceable swirl vanes located upstream of the throat of various converging-diverging nozzles. The supersonic vortex is then injected into either a coflowing supersonic stream or ambient air. The structure of the isolated vortex is investigated in a supersonic wind tunnel using miniature, fast-response, five-hole and total temperature probes and in a free jet using laser Doppler velocimetry. The cases tested have unit Reynolds numbers in excess of 25 million per meter, axial Mach numbers ranging from 2.5 to 4.0, and peak tangential Mach numbers from 0 (i.e., a pure jet) to about 0.7. The results show that the typical supersonic wake-like vortex consists of a non-isentropic, rotational core, where the reduced circulation distribution is self similar, and an outer isentropic, irrotational region. The vortex core is also a region of significant turbulent fluctuations. Radial profiles of turbulent kinetic energy and axial-tangential Reynolds stress are presented. The interactions between the vortex and both oblique and normal shock waves are investigated using nonintrusive optical diagnostics (i.e. schlieren, planar laser scattering, and laser Doppler velocimetry). Of the various types, two Mach 2.5 overexpanded-nozzle Mach disc interactions are examined in detail. Below a certain vortex strength, a 'weak' interaction exists in which the normal shock is perturbed locally into an unsteady 'bubble' shock near the vortex axis, but vortex breakdown (i.e., a stagnation point) does not occur. For stronger vortices, a random unsteady 'strong' interaction results that causes vortex breakdown. The vortex core reforms downstream of the rear stagnation point, and the reduced circulation distribution once again becomes self-similar in this region. A-new model of this interaction is proposed. Finally, a curve defining the approximate limits of supersonic vortex breakdown is presented.

Settles, G. S.↗