Numerical study of shock-wave/boundary layer interactions in premixed hydrogen-air hypersonic flows

NASA Technical Reports Server (NTRS)

Yungster, Shaye

1990-01-01

A computational study of shock wave/boundary layer interactions involving premixed combustible gases, and the resulting combustion processes is presented. The analysis is carried out using a new fully implicit, total variation diminishing (TVD) code developed for solving the fully coupled Reynolds-averaged Navier-Stokes equations and species continuity equations in an efficient manner. To accelerate the convergence of the basic iterative procedure, this code is combined with vector extrapolation methods. The chemical nonequilibrium processes are simulated by means of a finite-rate chemistry model for hydrogen-air combustion. Several validation test cases are presented and the results compared with experimental data or with other computational results. The code is then applied to study shock wave/boundary layer interactions in a ram accelerator configuration. Results indicate a new combustion mechanism in which a shock wave induces combustion in the boundary layer, which then propagates outwards and downstream. At higher Mach numbers, spontaneous ignition in part of the boundary layer is observed, which eventually extends along the entire boundary layer at still higher values of the Mach number.

Real Gas Scale Effects on Hypersonic Laminar Boundary-Layer Parameters Including Effects of Entropy-Layer Swallowing

DTIC Science & Technology

1975-12-01

crossed the essentially normal portion of the bow shock is swallowed by the boundary layer. The flow along the edge of the boundary layer on the aft...portions hf the body will then have passed through an oblique part of the bow shock and will be in a different state than had it passed through a normal...determination of the local edge flow conditions may be improvedby taking into con- sideration the inclination of the bow shock where the local flow stream- line

Correlation of heat transfer for the zero pressure gradient hypersonic laminar boundary layer for several gases

NASA Technical Reports Server (NTRS)

Cook, W. J.

1973-01-01

A theoretical study of heat transfer for zero pressure gradient hypersonic laminar boundary layers for various gases with particular application to the flows produced in an expansion tube facility was conducted. A correlation based on results obtained from solutions to the governing equations for five gases was formulated. Particular attention was directed toward the laminar boundary layer shock tube splitter plates in carbon dioxide flows generated by high speed shock waves. Computer analysis of the splitter plate boundary layer flow provided information that is useful in interpreting experimental data obtained in shock tube gas radiation studies.

Interferometric data for a shock-wave/boundary-layer interaction

NASA Technical Reports Server (NTRS)

Dunagan, Stephen E.; Brown, James L.; Miles, John B.

1986-01-01

An experimental study of the axisymmetric shock-wave / boundary-layer strong interaction flow generated in the vicinity of a cylinder-cone intersection was conducted. The study data are useful in the documentation and understanding of compressible turbulent strong interaction flows, and are part of a more general effort to improve turbulence modeling for compressible two- and three-dimensional strong viscous/inviscid interactions. The nominal free stream Mach number was 2.85. Tunnel total pressures of 1.7 and 3.4 atm provided Reynolds number values of 18 x 10(6) and 36 x 10(6) based on model length. Three cone angles were studied giving negligible, incipient, and large scale flow separation. The initial cylinder boundary layer upstream of the interaction had a thickness of 1.0 cm. The subsonic layer of the cylinder boundary layer was quite thin, and in all cases, the shock wave penetrated a significant portion of the boundary layer. Owing to the thickness of the cylinder boundary layer, considerable structural detail was resolved for the three shock-wave / boundary-layer interaction cases considered. The primary emphasis was on the application of the holographic interferometry technique. The density field was deduced from an interferometric analysis based on the Able transform. Supporting data were obtained using a 2-D laser velocimeter, as well as mean wall pressure and oil flow measurements. The attached flow case was observed to be steady, while the separated cases exhibited shock unsteadiness. Comparisons with Navier-Stokes computations using a two-equation turbulence model are presented.

Transonic shock-induced dynamics of a flexible wing with a thick circular-arc airfoil

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Rivera, Jose A., Jr.

1991-01-01

Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.

Control of shock-wave boundary-layer interactions by bleed in supersonic mixed compression inlets

NASA Technical Reports Server (NTRS)

Fukuda, M. K.; Reshotko, E.; Hingst, W. R.

1975-01-01

An experimental investigation has been conducted to determine the effect of bleed region geometry and bleed rate on shock wave-boundary layer interactions in an axisymmetric, mixed-compression inlet at a Mach number of 2.5. The full realizable reduction in transformed form factor is obtained by bleeding off about half the incident boundary layer mass flow. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise. Slanted holes are more effective than normal holes. Two different bleed hole sizes were tested without detectable difference in performance.

Laser interferometer/Preston tube skin-friction comparison in shock/boundary-layer interaction

NASA Technical Reports Server (NTRS)

Kim, K.-S.; Lee, Y.; Settles, G. S.

1991-01-01

An evaluation is conducted of the accuracy of the 'Preston tube' surface pitot-pressure skin friction measurement method relative to the already proven laser interferometer skin-friction meter in a swept shock wave/turbulent boundary-layer interaction. The Preston tube was used to estimate the total shear-stress distribution in a fin-generated swept shock-wave/turbulent boundary-layer interaction. The Keener-Hopkins calibration method using the isentropic relation to calculate the Preston-tube Mach number produces the best results.

Flow and Turbulence Modeling and Computation of Shock Buffet Onset for Conventional and Supercritical Airfoils

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

1998-01-01

Flow and turbulence models applied to the problem of shock buffet onset are studied. The accuracy of the interactive boundary layer and the thin-layer Navier-Stokes equations solved with recent upwind techniques using similar transport field equation turbulence models is assessed for standard steady test cases, including conditions having significant shock separation. The two methods are found to compare well in the shock buffet onset region of a supercritical airfoil that involves strong trailing-edge separation. A computational analysis using the interactive-boundary layer has revealed a Reynolds scaling effect in the shock buffet onset of the supercritical airfoil, which compares well with experiment. The methods are next applied to a conventional airfoil. Steady shock-separated computations of the conventional airfoil with the two methods compare well with experiment. Although the interactive boundary layer computations in the shock buffet region compare well with experiment for the conventional airfoil, the thin-layer Navier-Stokes computations do not. These findings are discussed in connection with possible mechanisms important in the onset of shock buffet and the constraints imposed by current numerical modeling techniques.

Understanding Micro-Ramp Control for Shock Boundary Layer Interactions

DTIC Science & Technology

2008-02-07

micro-ramps on a supersonic boundary layer at M=3.0 was investigated using monotone integrated Large Eddy Simulations (MILES) and Reynolds Averaged Navier... Supersonic boundary layer flow with micro-ramp and no shock wave 3.2 SBLI with no micro-ramp 3.3 SBLI with micro-ramp 3.4 Micro-ramp size and location IV . C...ramps on a supersonic boundary layer at M=3.0 was investigated using monotone integrated Large Eddy Simulations (MILES) and Reynolds Averaged Navier

Experimental investigation of a two-dimensional shock-turbulent boundary layer interaction with bleed

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Tanji, F. T.

1983-01-01

The two-dimensional interaction of an oblique shock wave with a turbulent boundary layer that included the effect of bleed was examined experimentally using a shock generator mounted across a supersonic wind tunnel The studies were performed at Mach numbers 2.5 and 2.0 and unit Reynolds number of approximately 2.0 x 10 to the 7th/meter. The study includes surface oil flow visualization, wall static pressure distributions and boundary layer pitot pressure profiles. In addition, the variation of the local bleed rates were measured. The results show the effect of the bleed on the boundary layer as well as the effect of the flow conditions on the local bleed rate.

Control and reduction of unsteady pressure loads in separated shock wave turbulent boundary layer interaction

NASA Technical Reports Server (NTRS)

Dolling, David S.; Barter, John W.

1995-01-01

The focus was on developing means of controlling and reducing unsteady pressure loads in separated shock wave turbulent boundary layer interactions. Section 1 describes how vortex generators can be used to effectively reduce loads in compression ramp interaction, while Section 2 focuses on the effects of 'boundary-layer separators' on the same interaction.

Evolution of wave patterns and temperature field in shock-tube flow

NASA Astrophysics Data System (ADS)

Kiverin, A. D.; Yakovenko, I. S.

2018-05-01

The paper is devoted to the numerical analysis of wave patterns behind a shock wave propagating in a tube filled with a gaseous mixture. It is shown that the flow inside the boundary layer behind the shock wave is unstable, and the way the instability develops fully corresponds to the solution obtained for the boundary layer over a flat plate. Vortical perturbations inside the boundary layer determine the nonuniformity of the temperature field. In turn, exactly these nonuniformities define the way the ignition kernels arise in the combustible mixture after the reflected shock interaction with the boundary layer. In particular, the temperature nonuniformity determines the spatial limitations of probable ignition kernel position relative to the end wall and side walls of the tube. In the case of low-intensity incident shocks the ignition could start not farther than the point of first interaction between the reflected shock wave and roller vortices formed in the process of boundary layer development. Proposed physical mechanisms are formulated in general terms and can be used for interpretation of the experimental data in any systems with a delayed exothermal reaction start. It is also shown that contact surface thickening occurs due to its interaction with Tollmien-Schlichting waves. This conclusion is of importance for understanding the features of ignition in shock tubes operating in the over-tailored regime.

A new method for designing shock-free transonic configurations

NASA Technical Reports Server (NTRS)

Sobieczky, H.; Fung, K. Y.; Seebass, A. R.; Yu, N. J.

1978-01-01

A method for the design of shock free supercritical airfoils, wings, and three dimensional configurations is described. Results illustrating the procedure in two and three dimensions are given. They include modifications to part of the upper surface of an NACA 64A410 airfoil that will maintain shock free flow over a range of Mach numbers for a fixed lift coefficient, and the modifications required on part of the upper surface of a swept wing with an NACA 64A410 root section to achieve shock free flow. While the results are given for inviscid flow, the same procedures can be employed iteratively with a boundary layer calculation in order to achieve shock free viscous designs. With a shock free pressure field the boundary layer calculation will be reliable and not complicated by the difficulties of shock wave boundary layer interaction.

Wind-US Code Contributions to the First AIAA Shock Boundary Layer Interaction Prediction Workshop

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Vyas, Manan A.; Yoder, Dennis A.

2013-01-01

This report discusses the computations of a set of shock wave/turbulent boundary layer interaction (SWTBLI) test cases using the Wind-US code, as part of the 2010 American Institute of Aeronautics and Astronautics (AIAA) shock/boundary layer interaction workshop. The experiments involve supersonic flows in wind tunnels with a shock generator that directs an oblique shock wave toward the boundary layer along one of the walls of the wind tunnel. The Wind-US calculations utilized structured grid computations performed in Reynolds-averaged Navier-Stokes mode. Four turbulence models were investigated: the Spalart-Allmaras one-equation model, the Menter Baseline and Shear Stress Transport k-omega two-equation models, and an explicit algebraic stress k-omega formulation. Effects of grid resolution and upwinding scheme were also considered. The results from the CFD calculations are compared to particle image velocimetry (PIV) data from the experiments. As expected, turbulence model effects dominated the accuracy of the solutions with upwinding scheme selection indicating minimal effects.

Experimental Investigation of Normal Shock Boundary-Layer Interaction with Hybrid Flow Control

NASA Technical Reports Server (NTRS)

Vyas, Manan A.; Hirt, Stefanie M.; Anderson, Bernhard H.

2012-01-01

Hybrid flow control, a combination of micro-ramps and micro-jets, was experimentally investigated in the 15x15 cm Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center. Full factorial, a design of experiments (DOE) method, was used to develop a test matrix with variables such as inter-ramp spacing, ramp height and chord length, and micro-jet injection flow ratio. A total of 17 configurations were tested with various parameters to meet the DOE criteria. In addition to boundary-layer measurements, oil flow visualization was used to qualitatively understand shock induced flow separation characteristics. The flow visualization showed the normal shock location, size of the separation, path of the downstream moving counter-rotating vortices, and corner flow effects. The results show that hybrid flow control demonstrates promise in reducing the size of shock boundary-layer interactions and resulting flow separation by means of energizing the boundary layer.

Weak incident shock interactions with Mach 8 laminar boundary layers. [of flat plate

NASA Technical Reports Server (NTRS)

Kaufman, L. G., II; Johnson, C. B.

1974-01-01

Weak shock-wave interactions with boundary layers on a flat plate were investigated experimentally in Mach 8 variable-density tunnel for plate-length Reynolds numbers. The undisturbed boundary layers were laminar over the entire plate length. Pressure and heat-transfer distributions were obtained for wedge-generated incident shock waves that resulted in pressure rises ranging from 1.36 to 4.46 (both nonseparated and separated boundary-layer flows). The resulting heat-transfer amplifications ranged from 1.45 to 14. The distributions followed established trends for nonseparated flows, for incipient separation, and for laminar free-interaction pressure rises. The experimental results corroborated established trends for the extent of the pressure rise and for certain peak heat-transfer correlations.

Prediction and measurement of heat transfer rates for the shock-induced unsteady laminar boundary layer on a flat plate

NASA Technical Reports Server (NTRS)

Cook, W. J.

1972-01-01

The unsteady laminar boundary layer induced by the flow-initiating shock wave passing over a flat plate mounted in a shock tube was theoretically and experimentally studied in terms of heat transfer rates to the plate for shock speeds ranging from 1.695 to 7.34 km/sec. The theory presented by Cook and Chapman for the shock-induced unsteady boundary layer on a plate is reviewed with emphasis on unsteady heat transfer. A method of measuring time-dependent heat-transfer rates using thin-film heat-flux gages and an associated data reduction technique are outlined in detail. Particular consideration is given to heat-flux measurement in short-duration ionized shocktube flows. Experimental unsteady plate heat transfer rates obtained in both air and nitrogen using thin-film heat-flux gages generally agree well with theoretical predictions. The experimental results indicate that the theory continues to predict the unsteady boundary layer behavior after the shock wave leaves the trailing edge of the plate even though the theory is strictly applicable only for the time interval in which the shock remains on the plate.

Effect of Pulsed Plasma Jets on the Recovering Boundary Layer Downstream of a Reflected Shock Interaction

NASA Astrophysics Data System (ADS)

Greene, Benton; Clemens, Noel; Magari, Patrick; Micka, Daniel; Ueckermann, Mattheus

2015-11-01

Shock-induced turbulent boundary layer separation can have many detrimental effects in supersonic inlets including flow distortion and instability, structural fatigue, poor pressure recovery, and unstart. The current study investigates the effect of pulsed plasma jets on the recovering boundary layer downstream of a reflected shock wave-boundary layer interaction. The effects of pitch and skew angle of the jet as well as the heating parameter and discharge time scale are tested using several pulsing frequencies. In addition, the effect of the plasma jets on the undisturbed boundary layer at 6 mm and 11 mm downstream of the jets is measured. A pitot-static pressure probe is used to measure the velocity profile of the boundary layer 35 mm downstream of the plasma jets, and the degree of boundary layer distortion is compared between the different models and run conditions. Additionally, the effect of each actuator configuration on the shape of the mean separated region is investigated using surface oil flow visualization. Previous studies with lower energy showed a weak effect on the downstream boundary layer. The current investigation will attempt to increase this effect using a higher-energy discharge. Funded by AFRL through and SBIR in collaboration with Creare, LLC.

On the mechanism of flow evolution in shock-tube experiments

NASA Astrophysics Data System (ADS)

Kiverin, Alexey; Yakovenko, Ivan

2018-02-01

The paper studies numerically the flow development behind the shock wave propagating inside the tube. The detailed analysis of the flow patterns behind the shock wave allows determination of the gas-dynamical origins of the temperature non-uniformities responsible for the subsequent localized start of chemical reactions in the test mixture. In particular, it is shown that the temperature field structure is determined mainly by the mechanisms of boundary layer instability development. The kinetic energy dissipation related to the flow deceleration inside boundary layer results in local heating of the test gas. At the same time, the heat losses to the tube wall lead to the cooling of the gas. Therefore the temperature stratification takes place on the scales of the boundary layer. As soon as the shock wave reflected from the end-wall of the tube interacts with the developed boundary layer the localized hot regions arise at a certain distance from the end wall. The position of these hot regions is associated with the zones of shock wave interaction with roller vortices at the margin between the boundary layer and the bulk flow. Formulated mechanism of the temperature field evolution can be used to explain the peculiarities of non-steady shock-induced ignition of combustible mixtures with moderate ignition delay times, where the ignition starts inside localized kernels at distance from the end wall.

Experimental and numerical investigation of the effect of distributed suction on oblique shock wave/turbulent boundary layer interaction. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Benhachmi, Driss; Greber, Isaac; Hingst, Warren R.

1988-01-01

A combined experimental and numerical study of the interaction of an incident oblique shock wave with a turbulent boundary layer on a rough plate and on a porous plate with suction is presented. The experimental phase involved the acquisition of mean data upstream of, within, and downstream of the interaction region at Mach numbers 2.5 and 3.0. Data were taken at unit Reynolds numbers of 1.66 E7 and 1.85 E7 m respectively, and for flow deflection angles of 0, 4, 6 and 8 degs. Measured data include wall static pressure, pitot pressure profiles, and local bleed distributions on the porous plate. On the rough plate, with no suction, the boundary layer profiles were modified near the wall, but not separated for the 4 deg flow deflection angle. For the higher deflection angles of 6 and 8 degs, the boundary layer was separated. Suction increases the strength of the incident shock required to separate the turbulent boundary layer; for all shock strengths tested, separation is completely eliminated. The pitot pressure profiles are affected throughout the whole boundary layer; they are fuller than the ones obtained on the rough plate. It is also found that the combination of suction and roughness introduces spatial perturbations.

Interaction of two glancing, crossing shock waves with a turbulent boundary-layer at various Mach numbers

NASA Technical Reports Server (NTRS)

Hingst, Warren R.; Williams, Kevin E.

1991-01-01

A preliminary experimental investigation was conducted to study two crossing, glancing shock waves of equal strengths, interacting with the boundary-layer developed on a supersonic wind tunnel wall. This study was performed at several Mach numbers between 2.5 and 4.0. The shock waves were created by fins (shock generators), spanning the tunnel test section, that were set at angles varying from 4 to 12 degrees. The data acquired are wall static pressure measurements, and qualitative information in the form of oil flow and schlieren visualizations. The principle aim is two-fold. First, a fundamental understanding of the physics underlying this flow phenomena is desired. Also, a comprehensive data set is needed for computational fluid dynamic code validation. Results indicate that for small shock generator angles, the boundary-layer remains attached throughout the flow field. However, with increasing shock strengths (increasing generator angles), boundary layer separation does occur and becomes progressively more severe as the generator angles are increased further. The location of the separation, which starts well downstream of the shock crossing point, moves upstream as shock strengths are increased. At the highest generator angles, the separation appears to begin coincident with the generator leading edges and engulfs most of the area between the generators. This phenomena occurs very near the 'unstart' limit for the generators. The wall pressures at the lower generator angles are nominally consistent with the flow geometries (i.e. shock patterns) although significantly affected by the boundary-layer upstream influence. As separation occurs, the wall pressures exhibit a gradient that is mainly axial in direction in the vicinity of the separation. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.

Pressure-sensing performance of upright cylinders in a Mach 10 boundary-layer

NASA Technical Reports Server (NTRS)

Johnson, Steven; Murphy, Kelly

1994-01-01

An experimental research program to provide basic knowledge of the pressure-sensing performance of upright, flushported cylinders in a hypersonic boundary layer is described. Three upright cylinders of 0.25-, 0.5- and l.0-in. diameters and a conventional rake were placed in the test section sidewall boundary layer of the 31 Inch Mach 10 Wind Tunnel at NASA Langley Research Center, Hampton, Virginia. Boundary-layer pressures from these cylinders were compared to those measured with a conventional rake. A boundary-layer thickness-to-cylinder-diameter ratio of 8 proved sufficient to accurately measure an overall pressure profile and ascertain the boundary-layer thickness. Effects of Reynolds number, flow angularity, and shock wave impingement on pressure measurement were also investigated. Although Reynolds number effects were negligible at the conditions studied, flow angularity above 10 deg significantly affects the measured pressures. Shock wave impingement was used to investigate orifice-to-orifice pressure crosstalk. No crosstalk was measured. The lower pressure measured above the oblique shock wave impingement showed no influence of the higher pressure generated at the lower port locations.

Shock-induced microdeformations in quartz and other mineralogical indications of an impact event at the Cretaceous-Tertiary boundary

USGS Publications Warehouse

Bohor, B.F.

1990-01-01

The event terminating the Cretaceous period and the Mesozoic era caused massive extinctions of flora and fauna worldwide. Theories of the nature of this event can be classed as endogenic (volcanic, climatic, etc.) or exogenic (extraterrestrial causes). Mineralogical evidence from the boundary clays and claystones strongly favor the impact of an extraterrestrial body as the cause of this event. Nonmarine KT boundary claystones are comprised of two separate layers-an upper layer composed of high-angle ejecta material (shocked quartz, altered glass and spinel) and a basal kaolinitic layer containing spherules, clasts, and altered glass, together with some shocked grains. Recognition of this dual-layered nature of the boundary clay is important for the determination of the timing and processes involved in the impact event and in the assignment and interpretation of geochemical signatures. Multiple sets of shock-induced microdeformations (planar features) in quartz grains separated from KT boundary clays provide compelling evidence of an impact event. This mineralogical manifestation of shock metamorphism is associated worldwide with a large positive anomaly of iridium in these boundary clays, which has also been considered indicative of the impact of a large extraterrestrial body. Global distributions of maximum sizes of shocked quartz grains from the boundary clays and the mineralogy of the ejecta components favor an impact on or near the North American continent. Spinel crystals (magnesioferrite) occur in the boundary clays as micrometer-sized octahedra or skeletal forms. Their composition differs from that of spinels found in terrestrial oceanic basalts. Magnesioferrite crystals are restricted to the high-angle ejecta layer of the boundary clays and their small size and skeletal morphology suggest that they are condensation products of a vaporized bolide. Hollow spherules ranging up to 1 mm in size are ubiquitously associated with the boundary clays. In nonmarine sections, where a high-angle ejecta layer and an underlying kaolinitic layer can be distinguished, the spherules are found only in the kaolinitic layer. The morphologies and surface features of these spherules suggest that they are original forms, and not secondary growths or algal bodies. These impact spherules closely resemble microtektites in size and shape. All of these features of the boundary clay are uniquely associated with impact, and cannot have been formed by volcanic or other terrestrial processes. ?? 1990.

Investigations of Compression Shocks and Boundary Layers in Gases Moving at High Speed

NASA Technical Reports Server (NTRS)

Ackeret, J.; Feldmann, F.; Rott, N.

1947-01-01

The mutual influences of compression shocks and friction boundary layers were investigated by means of high speed wind tunnels.Schlieren optics provided a clear picture of the flow phenomena and were used for determining the location of the compression shocks, measurement of shock angles, and also for Mach angles. Pressure measurement and humidity measurements were also taken into consideration.Results along with a mathematical model are described.

Comparison of three large-eddy simulations of shock-induced turbulent separation bubbles

NASA Astrophysics Data System (ADS)

Touber, Emile; Sandham, Neil D.

2009-12-01

Three different large-eddy simulation investigations of the interaction between an impinging oblique shock and a supersonic turbulent boundary layer are presented. All simulations made use of the same inflow technique, specifically aimed at avoiding possible low-frequency interferences with the shock/boundary-layer interaction system. All simulations were run on relatively wide computational domains and integrated over times greater than twenty five times the period of the most commonly reported low-frequency shock-oscillation, making comparisons at both time-averaged and low-frequency-dynamic levels possible. The results confirm previous experimental results which suggested a simple linear relation between the interaction length and the oblique-shock strength if scaled using the boundary-layer thickness and wall-shear stress. All the tested cases show evidences of significant low-frequency shock motions. At the wall, energetic low-frequency pressure fluctuations are observed, mainly in the initial part of interaction.

Numerical and Experimental Investigation of Multiple Shock Wave/Turbulent Boundary Layer Interactions in a Rectangular Duct

DTIC Science & Technology

1988-01-06

the bottom % kall followin,, the interaction. At 6Wuh = 0.35 the shock train would not stay attached to a single wall long enough for the surface...Interaction of a Shock Wave with a Laminar Boundary Layer," Lecture Notes in Physics, Vol. 8, Springer-Verlag, 1971 , pp. 151-163. 51 MacCormack, R. W

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. I - Pressure distribution

NASA Technical Reports Server (NTRS)

Messiter, A. F.

1980-01-01

Asymptotic solutions are derived for the pressure distribution in the interaction of a weak normal shock wave with a turbulent boundary layer. The undisturbed boundary layer is characterized by the law of the wall and the law of the wake for compressible flow. In the limiting case considered, for 'high' transonic speeds, the sonic line is very close to the wall. Comparisons with experiment are shown, with corrections included for the effect of longitudinal wall curvature and for the boundary-layer displacement effect in a circular pipe.

Comparison of theoretical and experimental boundary-layer development in a Mach 2.5 mixed-compression inlet

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Towne, C. E.

1974-01-01

An analytical investigation was made of the boundary layer flow in an axisymmetric Mach 2.5 mixed compression inlet, and the results were compared with experimental measurements. The inlet tests were conducted in the Lewis 10- by 10-foot supersonic wind tunnel at a unit Reynolds number of 8.2 million/m. The inlet incorporated porous bleed regions for boundary layer control, and the effect of this bleed was taken into account in the analysis. The experimental boundary layer data were analyzed by using similarity laws from which the skin friction coefficient was obtained. The boundary layer analysis included predictions of laminar and turbulent boundary layer growth, transition, and the effects of the shock boundary layer interactions. In addition, the surface static pressures were compared with those obtained from an inviscid characteristics program. The results of investigation showed that the analytical techniques gave satisfactory predictions of the boundary layer flow except in regions that were badly distorted by the terminal shock.

Numerical solution to the glancing sidewall oblique shock wave/turbulent boundary layer interaction in three dimension

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Benson, T. J.

1983-01-01

A supersonic three-dimensional viscous forward-marching computer design code called PEPSIS is used to obtain a numerical solution of the three-dimensional problem of the interaction of a glancing sidewall oblique shock wave and a turbulent boundary layer. Very good results are obtained for a test case that was run to investigate the use of the wall-function boundary-condition approximation for a highly complex three-dimensional shock-boundary layer interaction. Two additional test cases (coarse mesh and medium mesh) are run to examine the question of near-wall resolution when no-slip boundary conditions are applied. A comparison with experimental data shows that the PEPSIS code gives excellent results in general and is practical for three-dimensional supersonic inlet calculations.

Numerical solution to the glancing sidewall oblique shock wave/turbulent boundary layer interaction in three-dimension

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Benson, T. J.

1983-01-01

A supersonic three-dimensional viscous forward-marching computer design code called PEPSIS is used to obtain a numerical solution of the three-dimensional problem of the interaction of a glancing sidewall oblique shock wave and a turbulent boundary layer. Very good results are obtained for a test case that was run to investigate the use of the wall-function boundary-condition approximation for a highly complex three-dimensional shock-boundary layer interaction. Two additional test cases (coarse mesh and medium mesh) are run to examine the question of near-wall resolution when no-slip boundary conditions are applied. A comparison with experimental data shows that the PEPSIS code gives excellent results in general and is practical for three-dimensional supersonic inlet calculations.

Self-sustained oscillations of a shock wave interacting with a boundary layer on a supercritical airfoil

NASA Technical Reports Server (NTRS)

Ventres, C. S.; Howe, M. S.

1984-01-01

A theory is proposed of the self-sustaining oscillations of a weak shock on an airfoi in steady, transonic flow. The interaction of the shock with the boundary layer on the airfoil produces displacement thickness fluctuations which convect downstream and generate sound by interaction with the trailing edge. A feedback loop is established when this sound impinges on the shock wave, resulting in the production of further fluctuations in the displacement thickness. The details are worked out for an idealized mean boundary layer velocity profile, but strong support for the basic hypotheses of the theory is provided by a comparison with recent experiments involving the generation of acoustic 'tone bursts' by a supercritical airfoil section.

Self-sustained oscillations of a shock wave interacting with a boundary layer on a supercritical airfoil

NASA Technical Reports Server (NTRS)

Ventres, C. S.; Howe, M. S.

1983-01-01

A theory is proposed of the self-sustaining oscillations of a weak shock on an airfoil in steady, transonic flow. The interaction of the shock with the boundary layer on the airfoil produces displacement thickness fluctuations which convect downstream and generate sound by interaction with the trailing edge. A feedback loop is established when this sound impinges on the shock wave, resulting in the production of further fluctuations in the displacement thickness. The details are worked out for an idealized mean boundary layer velocity profile, but strong support for the basic hypotheses of the theory is provided by a comparison with recent experiments involving the generation of acoustic "tone bursts' by a supercritical airfoil section.

Assessment of Turbulent Shock-Boundary Layer Interaction Computations Using the OVERFLOW Code

NASA Technical Reports Server (NTRS)

Oliver, A. B.; Lillard, R. P.; Schwing, A. M.; Blaisdell, G> A.; Lyrintzis, A. S.

2007-01-01

The performance of two popular turbulence models, the Spalart-Allmaras model and Menter s SST model, and one relatively new model, Olsen & Coakley s Lag model, are evaluated using the OVERFLOWcode. Turbulent shock-boundary layer interaction predictions are evaluated with three different experimental datasets: a series of 2D compression ramps at Mach 2.87, a series of 2D compression ramps at Mach 2.94, and an axisymmetric coneflare at Mach 11. The experimental datasets include flows with no separation, moderate separation, and significant separation, and use several different experimental measurement techniques (including laser doppler velocimetry (LDV), pitot-probe measurement, inclined hot-wire probe measurement, preston tube skin friction measurement, and surface pressure measurement). Additionally, the OVERFLOW solutions are compared to the solutions of a second CFD code, DPLR. The predictions for weak shock-boundary layer interactions are in reasonable agreement with the experimental data. For strong shock-boundary layer interactions, all of the turbulence models overpredict the separation size and fail to predict the correct skin friction recovery distribution. In most cases, surface pressure predictions show too much upstream influence, however including the tunnel side-wall boundary layers in the computation improves the separation predictions.

Interactive calculation procedures for mixed compression inlets

NASA Technical Reports Server (NTRS)

Reshotko, Eli

1983-01-01

The proper design of engine nacelle installations for supersonic aircraft depends on a sophisticated understanding of the interactions between the boundary layers and the bounding external flows. The successful operation of mixed external-internal compression inlets depends significantly on the ability to closely control the operation of the internal compression portion of the inlet. This portion of the inlet is one where compression is achieved by multiple reflection of oblique shock waves and weak compression waves in a converging internal flow passage. However weak these shocks and waves may seem gas-dynamically, they are of sufficient strength to separate a laminar boundary layer and generally even strong enough for separation or incipient separation of the turbulent boundary layers. An understanding was developed of the viscous-inviscid interactions and of the shock wave boundary layer interactions and reflections.

CFL3D Contribution to the AIAA Supersonic Shock Boundary Layer Interaction Workshop

NASA Technical Reports Server (NTRS)

Rumsey, Christopher L.

2010-01-01

This paper documents the CFL3D contribution to the AIAA Supersonic Shock Boundary Layer Interaction Workshop, held in Orlando, Florida in January 2010. CFL3D is a Reynolds-averaged Navier-Stokes code. Four shock boundary layer interaction cases are computed using a one-equation turbulence model widely used for other aerodynamic problems of interest. Two of the cases have experimental data available at the workshop, and two of the cases do not. The effect of grid, flux scheme, and thin-layer approximation are investigated. Comparisons are made to the available experimental data. All four cases exhibit strong three-dimensional behavior in and near the interaction regions, resulting from influences of the tunnel side-walls.

Effects of shock on hypersonic boundary layer stability

NASA Astrophysics Data System (ADS)

Pinna, F.; Rambaud, P.

2013-06-01

The design of hypersonic vehicles requires the estimate of the laminar to turbulent transition location for an accurate sizing of the thermal protection system. Linear stability theory is a fast scientific way to study the problem. Recent improvements in computational capabilities allow computing the flow around a full vehicle instead of using only simplified boundary layer equations. In this paper, the effect of the shock is studied on a mean flow provided by steady Computational Fluid Dynamics (CFD) computations and simplified boundary layer calculations.

Experimental Study of Fillets to Reduce Corner Effects in an Oblique Shock-Wave/Boundary Layer Interaction

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.

2015-01-01

A test was conducted in the 15 cm x 15 cm supersonic wind tunnel at NASA Glenn Research Center that focused on corner effects of an oblique shock-wave/boundary-layer interaction. In an attempt to control the interaction in the corner region, eight corner fillet configurations were tested. Three parameters were considered for the fillet configurations: the radius, the fillet length, and the taper length from the square corner to the fillet radius. Fillets effectively reduced the boundary-layer thickness in the corner; however, there was an associated penalty in the form of increased boundary-layer thickness at the tunnel centerline. Larger fillet radii caused greater reductions in boundary-layer thickness along the corner bisector. To a lesser, but measureable, extent, shorter fillet lengths resulted in thinner corner boundary layers. Overall, of the configurations tested, the largest radius resulted in the best combination of control in the corner, evidenced by a reduction in boundary-layer thickness, coupled with minimal impacts at the tunnel centerline.

An experimental investigation of shock wave-turbulent boundary layer interactions with and without boundary layer suction: A data summary report

NASA Technical Reports Server (NTRS)

Sun, C. C.; Childs, M. E.

1977-01-01

Tabulated data from a series of experimental studies of the interaction of a shock wave with a turbulent boundary layer in axisymmetric flow configurations is presented. The studies were conducted at the walls of circular wind tunnels and on the cylindrical centerbody of an annular flow channel. Detailed pitot pressure profiles and wall static pressure profiles upstream of, within and downstream of the interaction region are given. Results are presented for flows at nominal freestream Mach Numbers of 2, 3 and 4. For studies at the tunnel sidewalls, the shock waves were produced by conical shock generators mounted on the centerline of the wind tunnel at zero angle of attack. The annular ring generator was used to produce the shock wave at the centerbody of the annular flow channel. The effects of boundary layer bleed were examined in the investigation. Both bleed rate and bleed location were studied. Most of the bleed studies were conducted with bleed holes drilled normal to the wall surface but the effects of slot suction were also examined. A summary of the principal results and conclusions is given.

Shock wave boundary layer interaction on suction side of compressor profile in single passage test section

NASA Astrophysics Data System (ADS)

Flaszynski, Pawel; Doerffer, Piotr; Szwaba, Ryszard; Kaczynski, Piotr; Piotrowicz, Michal

2015-11-01

The shock wave boundary layer interaction on the suction side of transonic compressor blade is one of the main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to investigate the flow structure on the suction side of a profile, a design of a generic test section in linear transonic wind tunnel was proposed. The experimental and numerical results for the flow structure investigations are shown for the flow conditions as the existing ones on the suction side of the compressor profile. Near the sidewalls the suction slots are applied for the corner flow structure control. It allows to control the Axial Velocity Density Ratio (AVDR), important parameter for compressor cascade investigations. Numerical results for Explicit Algebraic Reynolds Stress Model with transition modeling are compared with oil flow visualization, schlieren and Pressure Sensitive Paint. Boundary layer transition location is detected by Temperature Sensitive Paint.

Experimental research on crossing shock wave boundary layer interactions

NASA Astrophysics Data System (ADS)

Settles, G. S.; Garrison, T. J.

1994-10-01

An experimental research effort of the Penn State Gas Dynamics Laboratory on the subject of crossing shock wave boundary layer interactions is reported. This three year study was supported by AFOSR Grant 89-0315. A variety of experimental techniques were employed to study the above phenomena including planar laser scattering flowfield visualization, kerosene lampblack surface flow visualization, laser-interferometer skin friction surveys, wall static pressure measurements, and flowfield five-hole probe surveys. For a model configuration producing two intersecting shock waves, measurements were made for a range of oblique shock strengths at freestream Mach numbers of 3.0 and 3.85. Additionally, measurements were made at Mach 3.85 for a configuration producing three intersecting waves. The combined experimental dataset was used to formulate the first detailed flowfield models of the crossing-shock and triple-shock wave/boundary layer interactions. The structure of these interactions was found to be similar over a broad range of interaction strengths and is dominated by a large, separated, viscous flow region.

An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

NASA Technical Reports Server (NTRS)

Hamed, Awatef

1995-01-01

The design of high efficiency supersonic inlets is a complex task involving the optimization of a number of performance parameters such as pressure recovery, spillage, drag, and exit distortion profile, over the flight Mach number range. Computational techniques must be capable of accurately simulating the physics of shock/boundary layer interactions, secondary corner flows, flow separation, and bleed if they are to be useful in the design. In particular, bleed and flow separation, play an important role in inlet unstart, and the associated pressure oscillations. Numerical simulations were conducted to investigate some of the basic physical phenomena associated with bleed in oblique shock wave boundary layer interactions that affect the inlet performance.

An Experimental Study into the Scaling of an Unswept-Sharp-Fin-Generated Shock/Turbulent Boundary Layer Interaction.

DTIC Science & Technology

1983-01-01

Influence Scaling of 2D and 3D Shock/Turbulent ioundary Layer Interactions at Compression Corners." AIM Paper 81-334, January 1981. 5. Kubota, H...generating 3D shock wave/boundary layer interactions 2 Unswept sharp fin interaction and coordinate system 3 Cobra probe measurements of Peake (4) at Mach 4...were made by two Druck 50 PSI transducers, each in- stalled in a computer-controlled 48-port Model 48J4 Scani- valve and referenced to vacuum. A 250

Transonic Shock-Wave/Boundary-Layer Interactions on an Oscillating Airfoil

NASA Technical Reports Server (NTRS)

Davis, Sanford S.; Malcolm, Gerald N.

1980-01-01

Unsteady aerodynamic loads were measured on an oscillating NACA 64A010 airfoil In the NASA Ames 11 by 11 ft Transonic Wind Tunnel. Data are presented to show the effect of the unsteady shock-wave/boundary-layer interaction on the fundamental frequency lift, moment, and pressure distributions. The data show that weak shock waves induce an unsteady pressure distribution that can be predicted quite well, while stronger shock waves cause complex frequency-dependent distributions due to flow separation. An experimental test of the principles of linearity and superposition showed that they hold for weak shock waves while flows with stronger shock waves cannot be superimposed.

Three-dimensional separation for interaction of shock waves with turbulent boundary layers

NASA Technical Reports Server (NTRS)

Goldberg, T. J.

1973-01-01

For the interaction of shock waves with turbulent boundary layers, obtained experimental three-dimensional separation results and correlations with earlier two-dimensional and three-dimensional data are presented. It is shown that separation occurs much earlier for turbulent three-dimensional than for two-dimensional flow at hypersonic speeds.

Effects of non-adiabatic walls on shock/boundary-layer interaction using direct numerical simulations

NASA Astrophysics Data System (ADS)

Volpiani, Pedro S.; Bernardini, Matteo; Larsson, Johan

2017-11-01

The influence of wall thermal conditions on the properties of an impinging shock wave interacting with a turbulent supersonic boundary layer is a research topic that still remains underexplored. In the present study, direct numerical simulations (DNS) are employed to investigate the flow properties of a shock wave interacting with a turbulent boundary layer at free-stream Mach number M∞ = 2.28 with distinct wall thermal conditions and shock strengths. Instantaneous and mean flow fields, wall quantities and the low-frequency unsteadiness are analyzed. While heating contributes to increase the extent of the interaction zone, wall cooling turns out to be a good candidate for flow control. The distribution of the Stanton number shows a good agreement with prior experimental studies and confirms the strong heat transfer and complex pattern within the interaction region. Numerical results indicate that the changes in the interaction length are mainly linked to the incoming boundary layer as suggested in previous studies (Souverein et al., 2013 and Jaunet et al., 2014). This work was supported by the Air Force Office of Scientific Research, Grant FA95501610385.

A time-accurate high-resolution TVD scheme for solving the Navier-Stokes equations

NASA Technical Reports Server (NTRS)

Kim, Hyun Dae; Liu, Nan-Suey

1992-01-01

A total variation diminishing (TVD) scheme has been developed and incorporated into an existing time-accurate high-resolution Navier-Stokes code. The accuracy and the robustness of the resulting solution procedure have been assessed by performing many calculations in four different areas: shock tube flows, regular shock reflection, supersonic boundary layer, and shock boundary layer interactions. These numerical results compare well with corresponding exact solutions or experimental data.

Flight Demonstration of a Shock Location Sensor Using Constant Voltage Hot-Film Anemometry

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Sarma, Garimella R.; Mangalam, Siva M.

1997-01-01

Flight tests have demonstrated the effectiveness of an array of hot-film sensors using constant voltage anemometry to determine shock position on a wing or aircraft surface at transonic speeds. Flights were conducted at the NASA Dryden Flight Research Center using the F-15B aircraft and Flight Test Fixture (FTF). A modified NACA 0021 airfoil was attached to the side of the FTF, and its upper surface was instrumented to correlate shock position with pressure and hot-film sensors. In the vicinity of the shock-induced pressure rise, test results consistently showed the presence of a minimum voltage in the hot-film anemometer outputs. Comparing these results with previous investigations indicate that hot-film anemometry can identify the location of the shock-induced boundary layer separation. The flow separation occurred slightly forward of the shock- induced pressure rise for a laminar boundary layer and slightly aft of the start of the pressure rise when the boundary layer was tripped near the airfoil leading edge. Both minimum mean output and phase reversal analyses were used to identify the shock location.

An experimental study of fluctuating pressure loads beneath swept shock/boundary-layer interactions

NASA Technical Reports Server (NTRS)

Settles, Gary S.

1991-01-01

A database is established on the fluctuating pressure loads produced on aerodynamic surfaces beneath 3-D shock wave/boundary layer interactions. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins. Fin angles from 5 to 25 deg at freestream Mach numbers between 2.5 and 4 produce a variety of interaction strengths from weak to very strong. Miniature Kulite pressure transducers mounted in the flat plate were used to measure interaction-induced wall pressure fluctuations. These data will be correlated with proposed new optical data on the fluctuations of the interaction structure, especially that of the lambda-shock system and its associated high-speed jet impingement.

A search for shocked quartz grains in the Allerød-Younger Dryas boundary layer

NASA Astrophysics Data System (ADS)

Hoesel, Annelies; Hoek, Wim Z.; Pennock, Gillian M.; Kaiser, Knut; Plümper, Oliver; Jankowski, Michal; Hamers, Maartje F.; Schlaak, Norbert; Küster, Mathias; Andronikov, Alexander V.; Drury, Martyn R.

2015-03-01

The Younger Dryas impact hypothesis suggests that multiple airbursts or extraterrestrial impacts occurring at the end of the Allerød interstadial resulted in the Younger Dryas cold period. So far, no reproducible, diagnostic evidence has, however, been reported. Quartz grains containing planar deformation features (known as shocked quartz grains), are considered a reliable indicator for the occurrence of an extraterrestrial impact when found in a geological setting. Although alleged shocked quartz grains have been reported at a possible Allerød-Younger Dryas boundary layer in Venezuela, the identification of shocked quartz in this layer is ambiguous. To test whether shocked quartz is indeed present in the proposed impact layer, we investigated the quartz fraction of multiple Allerød-Younger Dryas boundary layers from Europe and North America, where proposed impact markers have been reported. Grains were analyzed using a combination of light and electron microscopy techniques. All samples contained a variable amount of quartz grains with (sub)planar microstructures, often tectonic deformation lamellae. A total of one quartz grain containing planar deformation features was found in our samples. This shocked quartz grain comes from the Usselo palaeosol at Geldrop Aalsterhut, the Netherlands. Scanning electron microscopy cathodoluminescence imaging and transmission electron microscopy imaging, however, show that the planar deformation features in this grain are healed and thus likely to be older than the Allerød-Younger Dryas boundary. We suggest that this grain was possibly eroded from an older crater or distal ejecta layer and later redeposited in the European sandbelt. The single shocked quartz grain at this moment thus cannot be used to support the Younger Dryas impact hypothesis.

Nonequilibrium viscous flow over Jovian entry probes at high altitudes

NASA Technical Reports Server (NTRS)

Kumar, A.; Szema, K. Y.; Tiwari, S. N.

1979-01-01

The viscous chemical nonequilibrium flow around a Jovian entry body is investigated at high altitudes using two different methods. First method is only for the stagnation region and integrates the full Navier-Stokes equations from the body surface to the freestream. The second method uses viscous shock layer equations between the body surface and the shock. Due to low Reynolds numbers, both methods use surface slip boundary conditions and the second method also uses shock slip boundary conditions. The results of the two methods are compared at the stagnation point. It is found that the entire shock layer is under chemical nonequilibrium at higher altitudes and that the slip boundary conditions are important at these altitudes.

Computation of three-dimensional shock wave and boundary-layer interactions

NASA Technical Reports Server (NTRS)

Hung, C. M.

1985-01-01

Computations of the impingement of an oblique shock wave on a cylinder and a supersonic flow past a blunt fin mounted on a plate are used to study three dimensional shock wave and boundary layer interaction. In the impingement case, the problem of imposing a planar impinging shock as an outer boundary condition is discussed and the details of particle traces in windward and leeward symmetry planes and near the body surface are presented. In the blunt fin case, differences between two dimensional and three dimensional separation are discussed, and the existence of an unique high speed, low pressure region under the separated spiral vortex core is demonstrated. The accessibility of three dimensional separation is discussed.

Assessment of CFD capability for prediction of hypersonic shock interactions

NASA Astrophysics Data System (ADS)

Knight, Doyle; Longo, José; Drikakis, Dimitris; Gaitonde, Datta; Lani, Andrea; Nompelis, Ioannis; Reimann, Bodo; Walpot, Louis

2012-01-01

The aerothermodynamic loadings associated with shock wave boundary layer interactions (shock interactions) must be carefully considered in the design of hypersonic air vehicles. The capability of Computational Fluid Dynamics (CFD) software to accurately predict hypersonic shock wave laminar boundary layer interactions is examined. A series of independent computations performed by researchers in the US and Europe are presented for two generic configurations (double cone and cylinder) and compared with experimental data. The results illustrate the current capabilities and limitations of modern CFD methods for these flows.

Flowfield analysis for successive oblique shock wave-turbulent boundary layer interactions

NASA Technical Reports Server (NTRS)

Sun, C. C.; Childs, M. E.

1976-01-01

A computation procedure is described for predicting the flowfields which develop when successive interactions between oblique shock waves and a turbulent boundary layer occur. Such interactions may occur, for example, in engine inlets for supersonic aircraft. Computations are carried out for axisymmetric internal flows at M 3.82 and 2.82. The effect of boundary layer bleed is considered for the M 2.82 flow. A control volume analysis is used to predict changes in the flow field across the interactions. Two bleed flow models have been considered. A turbulent boundary layer program is used to compute changes in the boundary layer between the interactions. The results given are for flows with two shock wave interactions and for bleed at the second interaction site. In principle the method described may be extended to account for additional interactions. The predicted results are compared with measured results and are shown to be in good agreement when the bleed flow rate is low (on the order of 3% of the boundary layer mass flow), or when there is no bleed. As the bleed flow rate is increased, differences between the predicted and measured results become larger. Shortcomings of the bleed flow models at higher bleed flow rates are discussed.

Conical similarity of shock/boundary layer interactions generated by swept fins

NASA Technical Reports Server (NTRS)

Lu, F. K.; Settles, G. S.

1983-01-01

A parametric experimental study has been made of the class of 3D shock wave/turbulent boundary layer interactions generated by swept-leading-edge fins. The fin sweepback angles ranged from 0 to 65 deg at angles of attack of 5, 9, and 15 deg. Two equilibrium 2D turbulent boundary layers with a free-stream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All the resulting interactions were found to possess conical symmetry of surface pressures and skin friction lines beyond an initial inception zone. Further, these interactions revealed a simple similarity based on inviscid shock strength irrespective of fin sweepback or angle of attack.

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. II - Wall shear stress

NASA Technical Reports Server (NTRS)

Liou, M. S.; Adamson, T. C., Jr.

1980-01-01

Asymptotic methods are used to calculate the shear stress at the wall for the interaction between a normal shock wave and a turbulent boundary layer on a flat plate. A mixing length model is used for the eddy viscosity. The shock wave is taken to be strong enough that the sonic line is deep in the boundary layer and the upstream influence is thus very small. It is shown that unlike the result found for laminar flow an asymptotic criterion for separation is not found; however, conditions for incipient separation are computed numerically using the derived solution for the shear stress at the wall. Results are compared with available experimental measurements.

Interference heating from interactions of shock waves with turbulent boundary layers at Mach 6

NASA Technical Reports Server (NTRS)

Johnson, C. B.; Kaufman, L. G., II

1974-01-01

An experimental investigation of interference heating resulting from interactions of shock waves and turbulent boundary layers was conducted. Pressure and heat-transfer distributions were measured on a flat plate in the free stream and on the wall of the test section of the Langley Mach 6 high Reynolds number tunnel for Reynolds numbers ranging from 2 million to 400 million. Various incident shock strengths were obtained by varying a wedge-shock generator angle (from 10 deg to 15 deg) and by placing a spherical-shock generator at different vertical positions above the instrumented flat plate and tunnel wall. The largest heating-rate amplification factors obtained for completely turbulent boundary layers were 22.1 for the flat plate and 11.6 for the tunnel wall experiments. Maximum heating correlated with peak pressures using a power law with a 0.85 exponent. Measured pressure distributions were compared with those calculated using turbulent free-interaction pressure rise theories, and separation lengths were compared with values calculated by using different methods.

A computer program for the calculation of the flow field including boundary layer effects for mixed-compression inlets at angle of attack

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.

1982-01-01

A computer program was developed which is capable of calculating the flow field in the supersonic portion of a mixed compression aircraft inlet operating at angle of attack. The supersonic core flow is computed using a second-order three dimensional method-of-characteristics algorithm. The bow shock and the internal shock train are treated discretely using a three dimensional shock fitting procedure. The boundary layer flows are computed using a second-order implicit finite difference method. The shock wave-boundary layer interaction is computed using an integral formulation. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data listings, are provided.

Effects of cavity size on the control of transonic internal flow around a biconvex circular arc airfoil

NASA Astrophysics Data System (ADS)

Rahman, M. Mostaqur; Hasan, A. B. M. Toufique; Rabbi, M. S.

2017-06-01

In transonic flow conditions, self-sustained shock wave oscillation on biconvex airfoils is initiated by the complex shock wave boundary layer interaction which is frequently observed in several modern internal aeronautical applications such as inturbine cascades, compressor blades, butterfly valves, fans, nozzles, diffusers and so on. Shock wave boundary layer interaction often generates serious problems such as unsteady boundary layer separation, self-excited shock waveoscillation with large pressure fluctuations, buffeting excitations, aeroacoustic noise, nonsynchronous vibration, high cycle fatigue failure and intense drag rise. Recently, the control of the self-excited shock oscillation around an airfoil using passive control techniques is getting intense interest. Among the passive means, control using open cavity has found promising. In this study, the effect of cavity size on the control of self-sustained shock oscillation was investigated numerically. The present computations are validated with available experimental results. The results showed that the average root mean square (RMS) of pressure oscillation around the airfoil with open cavity has reduced significantly when compared to airfoil without cavity (clean airfoil).

Budget of Turbulent Kinetic Energy in a Shock Wave Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Vyas, Manan A.; Waindim, Mbu; Gaitonde, Datta V.

2016-01-01

Implicit large-eddy simulation (ILES) of a shock wave/boundary-layer interaction (SBLI) was performed. Quantities present in the exact equation of the turbulent kinetic energy transport were accumulated and used to calculate terms like production, dissipation, molecular diffusion, and turbulent transport. The present results for a turbulent boundary layer were validated by comparison with direct numerical simulation data. It was found that a longer development domain was necessary for the boundary layer to reach an equilibrium state and a finer mesh resolution would improve the predictions. In spite of these findings, trends of the present budget match closely with that of the direct numerical simulation. Budgets for the SBLI region are presented at key axial stations. These budgets showed interesting dynamics as the incoming boundary layer transforms and the terms of the turbulent kinetic energy budget change behavior within the interaction region.

Color surface-flow visualization of fin-generated shock wave boundary-layer interactions

NASA Technical Reports Server (NTRS)

Lu, F. K.; Settles, G. S.

1990-01-01

Kerosene-lampblack mixtures with addition of a ground colored chalk were used in an experiment on visualizing surface flows of swept shock boundary-layer interactions. The results show that contrasting colors intensify the visualization of different regions of the interaction surface, and help the eye in following the fine streaks to locate the upstream influence. The study confirms observations of the separation occurring at shock strength below accepted values. The superiority of the reported technique over the previous monochrome technique is demonstrated.

Color surface-flow visualization of fin-generated shock wave boundary-layer interactions

NASA Astrophysics Data System (ADS)

Lu, F. K.; Settles, G. S.

1990-03-01

Kerosene-lampblack mixtures with addition of a ground colored chalk were used in an experiment on visualizing surface flows of swept shock boundary-layer interactions. The results show that contrasting colors intensify the visualization of different regions of the interaction surface, and help the eye in following the fine streaks to locate the upstream influence. The study confirms observations of the separation occurring at shock strength below accepted values. The superiority of the reported technique over the previous monochrome technique is demonstrated.

Numerical simulation of supersonic flow using a new analytical bleed boundary condition

NASA Technical Reports Server (NTRS)

Harloff, G. J.; Smith, G. E.

1995-01-01

A new analytical bleed boundary condition is used to compute flowfields for a strong oblique shock wave/boundary layer interaction with a baseline and three bleed rates at a freestream Mach number of 2.47 with an 8 deg shock generator. The computational results are compared to experimental Pitot pressure profiles and wall static pressures through the interaction region. An algebraic turbulence model is employed for the bleed and baseline cases, and a one equation model is also used for the baseline case where the boundary layer is separated.

CFD application to supersonic/hypersonic inlet airframe integration. [computational fluid dynamics (CFD)

NASA Technical Reports Server (NTRS)

Benson, Thomas J.

1988-01-01

Supersonic external compression inlets are introduced, and the computational fluid dynamics (CFD) codes and tests needed to study flow associated with these inlets are outlined. Normal shock wave turbulent boundary layer interaction is discussed. Boundary layer control is considered. Glancing sidewall shock interaction is treated. The CFD validation of hypersonic inlet configurations is explained. Scramjet inlet modules are shown.

Static Performance of a Fixed-Geometry Exhaust Nozzle Incorporating Porous Cavities for Shock-Boundary Layer Interaction Control

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Hunter, Craig A.

1999-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a fixed-geometry exhaust nozzle incorporating porous cavities for shock-boundary layer interaction control. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and 27 porous configurations. For the porous configurations, the effects of percent open porosity, hole diameter, and cavity depth were determined. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at over-expanded conditions. Porous configurations were capable of controlling off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. The ability of the porous nozzle concept to alternately alleviate separation or encourage stable separation of exhaust flow through shock-boundary layer interaction control offers tremendous off-design performance benefits for fixed-geometry nozzle installations. In addition, the ability to encourage separation on one divergent flap while alleviating it on the other makes it possible to generate thrust vectoring using a fixed-geometry nozzle.

Swept shock/boundary layer interaction experiments in support of CFD code validation

NASA Technical Reports Server (NTRS)

Settles, G. S.; Lee, Y.

1992-01-01

Research on the topic of shock wave/turbulent boundary-layer interaction was carried out during the past three years at the Penn State Gas Dynamics Laboratory. This report describes the experimental research program which provides basic knowledge and establishes new data on heat transfer in swept shock wave/boundary-layer interactions. An equilibrium turbulent boundary-layer on a flat plate is subjected to impingement by swept planar shock waves generated by a sharp fin. Five different interactions with fin angle ranging from 10 deg to 20 deg at freestream Mach numbers of 3.0 and 4.0 produce a variety of interaction strengths from weak to very strong. A foil heater generates a uniform heat flux over the flat plate surface, and miniature thin-film-resistance sensors mounted on it are used to measure the local surface temperature. The heat convection equation is then solved for the heat transfer distribution within an interaction, yielding a total uncertainty of about +/- 10 percent. These experimental data are compared with the results of numerical Navier-Stokes solutions which employ a k-epsilon turbulence model. Finally, a simplified form of the peak heat transfer correlation for fin interactions is suggested.

Detailed Study of Moderately Swept Impinging Oblique Shock/Boundary-Layer Interaction in Mach 2.3 Flow

NASA Astrophysics Data System (ADS)

Threadgill, James; Doerhmann, Adam; Little, Jesse

2017-11-01

A detailed experimental investigation of an impinging oblique Shock/Boundary Layer Interaction (SBLI) with 30° sweep in Mach 2.3 flow has been conducted. Despite its non-dimensional form, this canonical SBLI configuration has attracted little attention and remains poorly understood. Using a 12 .5° shock generator mounted in the freestream over a turbulent boundary layer, the interaction has been characterized with oil flow visualization, fast-response pressure transducers, and particle image velocimetry. Velocity vectors are used to extract the 3D interaction structure. These data are compared to wall pressure measurements and surface skin-friction streamlines. A local collapse of data normal to separation indicates a swept equivalence to Free Interaction Theory, albeit at a lower angle of sweep than imposed by the shock generator. Conditions at reattachment align with the imposed shock. Low-frequency shock motion near separation is observed, analogous to unswept SBLIs, with significant correlations that indicate spanwise traveling ripples in the shock foot. However, the magnitude of wall-pressure unsteadiness in this location is lower and shifted to higher frequencies than observed in equivalent unswept SBLI counterparts. Supported by the Air Force Office of Scientific Research (FA9550-15-1-0430).

An experimental study of a three-dimensional shock wave/turbulent boundary-layer interaction at a hypersonic Mach number

NASA Technical Reports Server (NTRS)

Kussoy, M. I.; Horstman, K. C.; Kim, K.-S.

1991-01-01

Experimental data for a series of three-dimensional shock-wave/turbulent-boundary-layer interaction flows at Mach 8.2 are presented. The test bodies, composed of sharp fins fastened to a flat-plate test surface, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure, heat-transfer, and skin-friction distributions, as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data were obtained for the purpose of validating computational models of these hypersonic interactions.

Documentation of Two- and Three-Dimensional Hypersonic Shock Wave/Turbulent Boundary Layer Interaction Flows

NASA Technical Reports Server (NTRS)

Kussoy, Marvin I.; Horstman, Clifford C.

1989-01-01

Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.

Control of a shock wave-boundary layer interaction using localized arc filament plasma actuators

NASA Astrophysics Data System (ADS)

Webb, Nathan Joseph

Supersonic flight is currently possible, but expensive. Inexpensive supersonic travel will require increased efficiency of high-speed air entrainment, an integral part of air-breathing propulsion systems. Although mixed compression inlet geometry can significantly improve entrainment efficiency, numerous Shock Wave-Boundary Layer Interactions (SWBLIs) are generated in this configuration. The boundary layer must therefore develop through multiple regions of adverse pressure gradient, causing it to thicken, and, in severe cases, separate. The associated increase in unsteadiness can have adverse effects on downstream engine hardware. The most severe consequence of these interactions is the increased aerodynamic blockage generated by the thickened boundary layer. If the increase is sufficient, it can choke the flow, causing inlet unstart, and resulting in a loss of thrust and high transient forces on the engine, airframe, and aircraft occupants. The potentially severe consequences associated with SWBLIs require flow control to ensure proper operation. Traditionally, boundary layer bleed has been used to control the interaction. Although this method is effective, it has inherent efficiency penalties. Localized Arc Filament Plasma Actuators (LAFPAs) are designed to generate perturbations for flow control. Natural flow instabilities act to amplify certain perturbations, allowing the LAFPAs to control the flow with minimal power input. LAFPAs also have the flexibility to maintain control over a variety of operating conditions. This work seeks to examine the effectiveness of LAFPAs as a separation control method for an oblique, impinging SWBLI. The low frequency unsteadiness in the reflected shock was thought to be the natural manifestation of a Kelvin-Helmholtz instability in the shear layer above the separation region. The LAFPAs were therefore placed upstream of the interaction to allow their perturbations to convect to the receptivity region (near the shear layer origin/separation line). Streamwise PIV measurements did not show that the boundary layer or separation region were energized by the actuation. The primary effect of the LAFPAs was the displacement of the reflected shock upstream. Jaunet et al. (2012) observed a similar shift in the reflected shock when they heated the wall beneath the boundary layer. A significantly greater power deposition was used in that work, and significantly larger shock displacements were observed. Although the LAFPAs output significantly less power (albeit in an unsteady, highly localized fashion), a parametric sweep strongly pointed to heating as the primary control mechanism. Further investigation and analysis showed that the near-wall heating of the flow by the plasma was the primary control mechanism of the LAFPAs, despite the small power input. The reflected shock was displaced by an increase in the separation region size, which was caused by the degradation of the upstream boundary layer. The LAFPAs degrade the upstream boundary layer through a variety of heating associated mechanisms: 1) Decreasing the density increases the mass flow deficit, 2) The altered skin-friction coefficient acts to retard the flow and make the velocity profile less full, and 3) The heating moves the sonic line further from the wall. Other mechanisms may also play a role.

Parallelization of the Flow Field Dependent Variation Scheme for Solving the Triple Shock/Boundary Layer Interaction Problem

NASA Technical Reports Server (NTRS)

Schunk, Richard Gregory; Chung, T. J.

2001-01-01

A parallelized version of the Flowfield Dependent Variation (FDV) Method is developed to analyze a problem of current research interest, the flowfield resulting from a triple shock/boundary layer interaction. Such flowfields are often encountered in the inlets of high speed air-breathing vehicles including the NASA Hyper-X research vehicle. In order to resolve the complex shock structure and to provide adequate resolution for boundary layer computations of the convective heat transfer from surfaces inside the inlet, models containing over 500,000 nodes are needed. Efficient parallelization of the computation is essential to achieving results in a timely manner. Results from a parallelization scheme, based upon multi-threading, as implemented on multiple processor supercomputers and workstations is presented.

Calculation of oblique-shock-wave laminar-boundary-layer interaction on a flat plate

NASA Technical Reports Server (NTRS)

Goldberg, U.; Reshotko, E.

1980-01-01

A finite difference solution to the problem of the interaction between an impinging oblique shock wave and the laminar boundary layer on a flat plate is presented. The boundary layer equations coupled with the Prandtl-Meyer relation for the external flow are used to calculate the flow field. A method for the calculation of the separated flow region is presented and discussed. Comparisons between this theory and the experimental results of other investigators show fairly good agreement. Results are presented for the case of a cooled wall with an oncoming flow at Mach number 2.0 without and with suction. The results show that a small amount of suction greatly reduces the extent of the separated region in the vicinity of the shock impingement location.

Internal hypersonic flow. [in thin shock layer

NASA Technical Reports Server (NTRS)

Lin, T. C.; Rubin, S. G.

1974-01-01

An approach for studying hypersonic internal flow with the aid of a thin-shock-layer approximation is discussed, giving attention to a comparison of thin-shock-layer results with the data obtained on the basis of the imposition theory or a finite-difference integration of the Euler equations. Relations in the case of strong interaction are considered together with questions of pressure distribution and aspects of the boundary-layer solution.

Shock Train/Boundary-Layer Interaction in Rectangular Scramjet Isolators

NASA Astrophysics Data System (ADS)

Geerts, Jonathan Simon

Numerous studies of the dual-mode scramjet isolator, a critical component in preventing inlet unstart and/or vehicle loss by containing a collection of flow disturbances called a shock train, have been performed since the dual-mode propulsion cycle was introduced in the 1960s. Low momentum corner flow and other three-dimensional effects inherent to rectangular isolators have, however, been largely ignored in experimental studies of the boundary layer separation driven isolator shock train dynamics. Furthermore, the use of two dimensional diagnostic techniques in past works, be it single-perspective line-of-sight schlieren/shadowgraphy or single axis wall pressure measurements, have been unable to resolve the three-dimensional flow features inside the rectangular isolator. These flow characteristics need to be thoroughly understood if robust dual-mode scramjet designs are to be fielded. The work presented in this thesis is focused on experimentally analyzing shock train/boundary layer interactions from multiple perspectives in aspect ratio 1.0, 3.0, and 6.0 rectangular isolators with inflow Mach numbers ranging from 2.4 to 2.7. Secondary steady-state Computational Fluid Dynamics studies are performed to compare to the experimental results and to provide additional perspectives of the flow field. Specific issues that remain unresolved after decades of isolator shock train studies that are addressed in this work include the three-dimensional formation of the isolator shock train front, the spatial and temporal low momentum corner flow separation scales, the transient behavior of shock train/boundary layer interaction at specific coordinates along the isolator's lateral axis, and effects of the rectangular geometry on semi-empirical relations for shock train length prediction. (Abstract shortened by ProQuest.).

A numerical solution of the Navier-Stokes equations for chemically nonequilibrium, merged stagnation shock layers on spheres and two-dimensional cylinders in air

NASA Technical Reports Server (NTRS)

Johnston, K. D.; Hendricks, W. L.

1978-01-01

Results of solving the Navier-Stokes equations for chemically nonequilibrium, merged stagnation shock layers on spheres and two-dimensional cylinders are presented. The effects of wall catalysis and slip are also examined. The thin shock layer assumption is not made, and the thick viscous shock is allowed to develop within the computational domain. The results show good comparison with existing data. Due to the more pronounced merging of shock layer and boundary layer for the sphere, the heating rates for spheres become higher than those for cylinders as the altitude is increased.

Pseudo-shock waves and their interactions in high-speed intakes

NASA Astrophysics Data System (ADS)

Gnani, F.; Zare-Behtash, H.; Kontis, K.

2016-04-01

In an air-breathing engine the flow deceleration from supersonic to subsonic conditions takes places inside the isolator through a gradual compression consisting of a series of shock waves. The wave system, referred to as a pseudo-shock wave or shock train, establishes the combustion chamber entrance conditions, and therefore influences the performance of the entire propulsion system. The characteristics of the pseudo-shock depend on a number of variables which make this flow phenomenon particularly challenging to be analysed. Difficulties in experimentally obtaining accurate flow quantities at high speeds and discrepancies of numerical approaches with measured data have been readily reported. Understanding the flow physics in the presence of the interaction of numerous shock waves with the boundary layer in internal flows is essential to developing methods and control strategies. To counteract the negative effects of shock wave/boundary layer interactions, which are responsible for the engine unstart process, multiple flow control methodologies have been proposed. Improved analytical models, advanced experimental methodologies and numerical simulations have allowed a more in-depth analysis of the flow physics. The present paper aims to bring together the main results, on the shock train structure and its associated phenomena inside isolators, studied using the aforementioned tools. Several promising flow control techniques that have more recently been applied to manipulate the shock wave/boundary layer interaction are also examined in this review.

Calculation of heat transfer on shuttle type configurations including the effects of variable entropy at boundary layer edge

NASA Technical Reports Server (NTRS)

Dejarnette, F. R.

1972-01-01

A relatively simple method is presented for including the effect of variable entropy at the boundary-layer edge in a heat transfer method developed previously. For each inviscid surface streamline an approximate shockwave shape is calculated using a modified form of Maslen's method for inviscid axisymmetric flows. The entropy for the streamline at the edge of the boundary layer is determined by equating the mass flux through the shock wave to that inside the boundary layer. Approximations used in this technique allow the heating rates along each inviscid surface streamline to be calculated independent of the other streamlines. The shock standoff distances computed by the present method are found to compare well with those computed by Maslen's asymmetric method. Heating rates are presented for blunted circular and elliptical cones and a typical space shuttle orbiter at angles of attack. Variable entropy effects are found to increase heating rates downstream of the nose significantly higher than those computed using normal-shock entropy, and turbulent heating rates increased more than laminar rates. Effects of Reynolds number and angles of attack are also shown.

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

NASA Technical Reports Server (NTRS)

Suder, Kenneth L.

1996-01-01

A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions. The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row. The blockage is evaluated in the core flow area as well as in the casing endwall region. Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed. Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed. In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis. The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer. The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer. However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss.

Effects of boundary layer on flame propagation generated by forced ignition behind an incident shock wave

NASA Astrophysics Data System (ADS)

Ishihara, S.; Tamura, S.; Ishii, K.; Kataoka, H.

2016-09-01

To study the effects of the boundary layer on the deflagration to detonation transition (DDT) process, the mixture behind an incident shock wave was ignited using laser breakdown. Ignition timing was controlled so that the interaction of the resulting flame with a laminar or turbulent boundary layer could be examined. In the case of the interaction with a laminar boundary layer, wrinkling of the flame was observed after the flame reached the corner of the channel. On the other hand, interaction with the turbulent boundary layer distorted the flame front and increased the spreading rate of the flame followed by prompt DDT. The inner structure of the turbulent boundary layer plays an important role in the DDT process. The region that distorted the flame within the turbulent boundary layer was found to be the intermediate region 0.01< y/δ < 0.4, where y is the distance from the wall and δ is the boundary layer thickness. The flame disturbance by the turbulent motions is followed by the flame interaction with the inner layer near the wall, which in turn generates a secondary-ignition kernel that produced a spherical accelerating flame, which ultimately led to the onset of detonation. After the flame reached the intermediate region, the time required for DDT was independent of the ignition position. The effect of the boundary layer on the propagating flame, thus, became relatively small after the accelerating flame was generated.

Effect of non-equilibrium flow chemistry and surface catalysis on surface heating to AFE

NASA Technical Reports Server (NTRS)

Stewart, David A.; Henline, William D.; Chen, Yih-Kanq

1991-01-01

The effect of nonequilibrium flow chemistry on the surface temperature distribution over the forebody heat shield on the Aeroassisted Flight Experiment (AFE) vehicle was investigated using a reacting boundary-layer code. Computations were performed by using boundary-layer-edge properties determined from global iterations between the boundary-layer code and flow field solutions from a viscous shock layer (VSL) and a full Navier-Stokes solution. Surface temperature distribution over the AFE heat shield was calculated for two flight conditions during a nominal AFE trajectory. This study indicates that the surface temperature distribution is sensitive to the nonequilibrium chemistry in the shock layer. Heating distributions over the AFE forebody calculated using nonequilibrium edge properties were similar to values calculated using the VSL program.

An LDA (Laser-Doppler Anemometry) investigation of three-dimensional normal shock wave boundary-layer interactions

NASA Technical Reports Server (NTRS)

Chriss, R. M.; Hingst, W. R.; Strazisar, A. J.; Keith, T. G., Jr.

1989-01-01

Nonintrusive measurements were made of a normal shock wave/boundary layer interaction. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made in the vicinity of the shock wave. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. Laser Doppler Anemometry, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The results define the flow field structure in detail for each case.

A new Cretaceous-Tertiary boundary locality in the western powder River basin, Wyoming: biological and geological implications

USGS Publications Warehouse

Nichols, D.J.; Brown, J.L.; Attrep, M.; Orth, C.J.

1992-01-01

A newly discovered Cretaceous-Tertiary (K-T) boundary locality in the western Powder River basin, Wyoming, is characterized by a palynologically defined extinction horizon, a fern-spore abundance anomaly, a strong iridium anomaly, and shock-metamorphosed quartz grains. Detailed microstratigraphic analyses show that about one third of the palynoflora (mostly angiosperm pollen) disappeared abruptly, placing the K-T boundary within a distinctive, 1- to 2-cm-thick claystone layer. Shocked quartz grains are concentrated at the top of this layer, and although fern-spore and iridium concentrations are high in this layer, they reach their maximum concentrations in a 2-cm-thick carbonaceous claystone that overlies the boundary claystone layer. The evidence supports the theory that the K-T boundary event was associated with the impact of an extraterrestrial body or bodies. Palynological analyses of samples from the K-T boundary interval document extensive changes in the flora that resulted from the boundary event. The palynologically and geochemically defined K-T boundary provides a unique time-line of use in regional basin analysis. ?? 1992.

On the boundary conditions on a shock wave for hypersonic flow around a descent vehicle

NASA Astrophysics Data System (ADS)

Golomazov, M. M.; Ivankov, A. A.

2013-12-01

Stationary hypersonic flow around a descent vehicle is examined by considering equilibrium and nonequilibrium reactions. We study how physical-chemical processes and shock wave conditions for gas species influence the shock-layer structure. It is shown that conservation conditions of species on the shock wave cause high-temperature and concentration gradients in the shock layer when we calculate spacecraft deceleration trajectory in the atmosphere at 75 km altitude.

Lunar and Planetary Science XXXV: Effects of Impacts: Shock and Awe

NASA Technical Reports Server (NTRS)

Kyte, F. T.; Koeberl, C.

2004-01-01

This document discusses the following topics: Zircon as a Shock Indicator in Impactites of Drill Core Yaxcopoil-1, Chicxulub Impact Structure, Mexico; Experimental Investigation of Shock Effects in a Metapelitic Granulite; Experimental Reproduction of Shock Veins in Single-Crystal Minerals; Post-Shock Crystal-Plastic Processes in Quartz from Crystalline Target Rocks of the Charlevoix Impact Structure; Shock Reequilibration of Fluid Inclusions; How Does Tektite Glass Lose Its Water?; Assessing the Role of Anhydrite in the KT Mass Extinction: Hints from Shock-loading Experiments; A Mineralogical and Geochemical Study of the Nonmarine Permian/Triassic Boundary in the Southern Karoo Basin, South Africa; Extraterrestrial Chromium in the Permian-Triassic Boundary at Graphite Peak, Antarctica; Magnetic Fe,Si,Al-rich Impact Spherules from the P-T Boundary Layer at Graphite Peak, Antarctica; A Newly Recognized Late Archean Impact Spherule Layer in the Reivilo Formation, Griqualand West Basin, South Africa; Initial Cr-Isotopic and Iridium Measurements of Concentrates from Late Eocene Cpx-Spherule Deposits; An Ordinary Chondrite Impactor Composition for the Bosumtwi Impact Structure, Ghana, West Africa: Discussion of Siderophile Element Contents and Os and Cr Isotope Data.

An experimental study of the turbulent boundary layer on a transport wing in subsonic and transonic flow

NASA Technical Reports Server (NTRS)

Spaid, Frank W.; Roos, Frederick W.; Hicks, Raymond M.

1990-01-01

The upper surface boundary layer on a transport wing model was extensively surveyed with miniature yaw probes at a subsonic and a transonic cruise condition. Additional data were obtained at a second transonic test condition, for which a separated region was present at mid-semispan, aft of mid-chord. Significant variation in flow direction with distance from the surface was observed near the trailing edge except at the wing root and tip. The data collected at the transonic cruise condition show boundary layer growth associated with shock wave/boundary layer interaction, followed by recovery of the boundary layer downstream of the shock. Measurements of fluctuating surface pressure and wingtip acceleration were also obtained. The influence of flow field unsteadiness on the boundary layer data is discussed. Comparisons among the data and predictions from a variety of computational methods are presented. The computed predictions are in reasonable agreement with the experimental data in the outboard regions where 3-D effects are moderate and adverse pressure gradients are mild. In the more highly loaded mid-span region near the trailing edge, displacement thickness growth was significantly underpredicted, except when unrealistically severe adverse pressure gradients associated with inviscid calculations were used to perform boundary layer calculations.

Effects of Nose Bluntness on Hypersonic Boundary-Layer Receptivity and Stability Over Cones

NASA Technical Reports Server (NTRS)

Kara, Kursat; Balakumar, Ponnampalam; Kandil, Osama A.

2011-01-01

The receptivity to freestream acoustic disturbances and the stability properties of hypersonic boundary layers are numerically investigated for boundary-layer flows over a 5 straight cone at a freestream Mach number of 6.0. To compute the shock and the interaction of the shock with the instability waves, the Navier-Stokes equations in axisymmetric coordinates were solved. In the governing equations, inviscid and viscous flux vectors are discretized using a fifth-order accurate weighted-essentially-non-oscillatory scheme. A third-order accurate total-variation-diminishing Runge-Kutta scheme is employed for time integration. After the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. The appearance of instability waves near the nose region and the receptivity of the boundary layer with respect to slow mode acoustic waves are investigated. Computations confirm the stabilizing effect of nose bluntness and the role of the entropy layer in the delay of boundary-layer transition. The current solutions, compared with experimental observations and other computational results, exhibit good agreement.

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

NASA Technical Reports Server (NTRS)

Hirt, Stefanie; Anderson, Bernhard

2007-01-01

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.

Structural characteristics of the shock-induced boundary layer separation extended to the leading edge

NASA Astrophysics Data System (ADS)

Tao, Y.; Liu, W. D.; Fan, X. Q.; Zhao, Y. L.

2017-07-01

For a better understanding of the local unstart of supersonic/hypersonic inlet, a series of experiments has been conducted to investigate the shock-induced boundary layer separation extended to the leading edge. Using the nanoparticle-based planar laser scattering, we recorded the fine structures of these interactions under different conditions and paid more attention to their structural characteristics. According to their features, these interactions could be divided into four types. Specifically, Type A wave pattern is similar to the classic shock wave/turbulent boundary layer interaction, and Type B wave configuration consists of an overall Mach reflection above the large scale separation bubble. Due to the gradual decrease in the size of the separation bubble, the separation bubble was replaced by several vortices (Type C wave pattern). Besides, for Type D wave configuration which exists in the local unstart inlet, there appears to be some flow spillage around the leading edge.

Swept shock/boundary-layer interactions: Scaling laws, flowfield structure, and experimental methods

NASA Technical Reports Server (NTRS)

Settles, Gary S.

1993-01-01

A general review is given of several decades of research on the scaling laws and flowfield structures of swept shock wave/turbulent boundary layer interactions. Attention is further restricted to the experimental study and physical understanding of the steady-state aspects of these flows. The interaction produced by a sharp, upright fin mounted on a flat plate is taken as an archetype. An overall framework of quasiconical symmetry describing such interactions is first developed. Boundary-layer separation, the interaction footprint, Mach number scaling, and Reynolds number scaling are then considered, followed by a discussion of the quasiconical similarity of interactions produced by geometrically-dissimilar shock generators. The detailed structure of these interaction flowfields is next reviewed, and is illustrated by both qualitative visualizations and quantitative flow images in the quasiconical framework. Finally, the experimental techniques used to investigate such flows are reviewed, with emphasis on modern non-intrusive optical flow diagnostics.

Preliminary study of the interactions caused by crossing shock waves and a turbulent boundary layer

NASA Technical Reports Server (NTRS)

Ketchum, A. C.; Bogdonoff, S. M.; Fernando, E. M.; Batcho, P. F.

1989-01-01

The subject research, the first phase of an extended study of the interaction of crossing shock waves with a turbulent boundary layer, has revealed the complexity of the resulting flow. Detailed surface visualization and mean wall static pressure distributions show little resemblance to the inviscid flow approximation, and the exploratory high frequency measurements show that the flow downstream of the theoretical inviscid shock crossing position has a significant unsteady characteristic. Further developments of the (unsteady) high frequency measurements are required to fully characterize the unsteadiness and the requirements to include this component in flowfield modeling.

Molecular Diagnostics of Diffusive Boundary Layers

NASA Astrophysics Data System (ADS)

Rawlings, J. M. C.; Hartquist, T. W.

1997-10-01

We have examined the chemistry in thin (<~0.01 pc) boundary layers between dark star-forming cores and warm, shocked T Tauri winds on the assumption that turbulence-driven diffusion occurs within them. The results indicate that emissions from C+, CH, OH, H2O and the J = 6 --> 5 transition of CO, among others, may serve as diagnostics of the boundary layers.

Numerical investigations of shock wave interaction with laminar boundary layer on compressor profile

NASA Astrophysics Data System (ADS)

Piotrowicz, M.; Flaszyński, P.

2016-10-01

The investigation of shockwave boundary layer interaction on suction side of transonic compressor blade is one of main objectives of TFAST project (Transition Location Effect on Shock Wave Boundary Layer Interaction). In order to look more closely into the flow structure on suction side of a profile, a design of generic test section in linear transonic wind tunnel was proposed. The experimental and numerical results of flow structure on a suction side of the compressor profile investigations are presented. The numerical simulations are carried out for EARSM (Explicit Algebraic Reynolds Stress Model) turbulence model with transition model. The result are compared with oil flow visualisation, schlieren pictures, Pressure Sensitive Paint (PSP) and static pressure.

Inlets, ducts, and nozzles

NASA Technical Reports Server (NTRS)

Abbott, John M.; Anderson, Bernhard H.; Rice, Edward J.

1990-01-01

The internal fluid mechanics research program in inlets, ducts, and nozzles consists of a balanced effort between the development of computational tools (both parabolized Navier-Stokes and full Navier-Stokes) and the conduct of experimental research. The experiments are designed to better understand the fluid flow physics, to develop new or improved flow models, and to provide benchmark quality data sets for validation of the computational methods. The inlet, duct, and nozzle research program is described according to three major classifications of flow phenomena: (1) highly 3-D flow fields; (2) shock-boundary-layer interactions; and (3) shear layer control. Specific examples of current and future elements of the research program are described for each of these phenomenon. In particular, the highly 3-D flow field phenomenon is highlighted by describing the computational and experimental research program in transition ducts having a round-to-rectangular area variation. In the case of shock-boundary-layer interactions, the specific details of research for normal shock-boundary-layer interactions are described. For shear layer control, research in vortex generators and the use of aerodynamic excitation for enhancement of the jet mixing process are described.

A documentation of two- and three-dimensional shock-separated turbulent boundary layers

NASA Technical Reports Server (NTRS)

Brown, J. D.; Brown, J. L.; Kussoy, M. I.

1988-01-01

A shock-related separation of a turbulent boundary layer has been studied and documented. The flow was that of an axisymmetric turbulent boundary layer over a 5.02-cm-diam cylinder that was aligned with the wind tunnel axis. The boundary layer was compressed by a 30 deg half-angle conical flare, with the cone axis inclined at an angle alpha to the cylinder axis. Nominal test conditions were P sub tau equals 1.7 atm and M sub infinity equals 2.85. Measurements were confined to the upper-symmetry, phi equals 0 deg, plane. Data are presented for the cases of alpha equal to 0. 5. and 10 deg and include mean surface pressures, streamwise and normal mean velocities, kinematic turbulent stresses and kinetic energies, as well as reverse-flow intermittencies. All data are given in tabular form; pressures, streamwise velocities, turbulent shear stresses, and kinetic energies are also presented graphically.

A viscous flow study of shock-boundary layer interaction, radial transport, and wake development in a transonic compressor

NASA Technical Reports Server (NTRS)

Hah, Chunill; Reid, Lonnie

1991-01-01

A numerical study based on the 3D Reynolds-averaged Navier-Stokes equation has been conducted to investigate the detailed flow physics inside a transonic compressor. 3D shock structure, shock-boundary layer interaction, flow separation, radial mixing, and wake development are all investigated at design and off-design conditions. Experimental data based on laser anemometer measurements are used to assess the overall quality of the numerical solution. An additional experimental study to investigate end-wall flow with a hot-film was conducted, and these results are compared with the numerical results. Detailed comparison with experimental data indicates that the overall features of the 3D shock structure, the shock-boundary layer interaction, and the wake development are all calculated very well in the numerical solution. The numerical results are further analyzed to examine the radial mixing phenomena in the transonic compressor. A thin sheet of particles is injected in the numerical solution upstream of the compressor. The movement of particles is traced with a 3D plotting package. This numerical survey of tracer concentration reveals the fundamental mechanisms of radial transport in this transonic compressor.

Heat transfer, velocity-temperature correlation, and turbulent shear stress from Navier-Stokes computations of shock wave/turbulent boundary layer interaction flows

NASA Technical Reports Server (NTRS)

Wang, C. R.; Hingst, W. R.; Porro, A. R.

1991-01-01

The properties of 2-D shock wave/turbulent boundary layer interaction flows were calculated by using a compressible turbulent Navier-Stokes numerical computational code. Interaction flows caused by oblique shock wave impingement on the turbulent boundary layer flow were considered. The oblique shock waves were induced with shock generators at angles of attack less than 10 degs in supersonic flows. The surface temperatures were kept at near-adiabatic (ratio of wall static temperature to free stream total temperature) and cold wall (ratio of wall static temperature to free stream total temperature) conditions. The computational results were studied for the surface heat transfer, velocity temperature correlation, and turbulent shear stress in the interaction flow fields. Comparisons of the computational results with existing measurements indicated that (1) the surface heat transfer rates and surface pressures could be correlated with Holden's relationship, (2) the mean flow streamwise velocity components and static temperatures could be correlated with Crocco's relationship if flow separation did not occur, and (3) the Baldwin-Lomax turbulence model should be modified for turbulent shear stress computations in the interaction flows.

A database of aerothermal measurements in hypersonic flow for CFD validation

NASA Technical Reports Server (NTRS)

Holden, M. S.; Moselle, J. R.

1992-01-01

This paper presents an experimental database selected and compiled from aerothermal measurements obtained on basic model configurations on which fundamental flow phenomena could be most easily examined. The experimental studies were conducted in hypersonic flows in 48-inch, 96-inch, and 6-foot shock tunnels. A special computer program was constructed to provide easy access to the measurements in the database as well as the means to plot the measurements and compare them with imported data. The database contains tabulations of model configurations, freestream conditions, and measurements of heat transfer, pressure, and skin friction for each of the studies selected for inclusion. The first segment contains measurements in laminar flow emphasizing shock-wave boundary-layer interaction. In the second segment, measurements in transitional flows over flat plates and cones are given. The third segment comprises measurements in regions of shock-wave/turbulent-boundary-layer interactions. Studies of the effects of surface roughness of nosetips and conical afterbodies are presented in the fourth segment of the database. Detailed measurements in regions of shock/shock boundary layer interaction are contained in the fifth segment. Measurements in regions of wall jet and transpiration cooling are presented in the final two segments.

Compressible Boundary Layer Investigation for Ramjet/scramjet Inlets and Nozzles

NASA Astrophysics Data System (ADS)

Goldfeld, M. A.; Starov, A. V.; Semenova, Yu. V.

2005-02-01

The results of experimental investigation of a turbulent boundary layer on compression and expansion surfaces are presented. They include the study of the shock wave and/or expansion fan action upon the boundary layer, boundary layer separation and its relaxation. Complex events of paired interactions and the flow on compression convex-concave surfaces were studied [M. Goldfeld, 1993]. The possibility and conditions of the boundary layer relaminarization behind the expansion fan and its effect on the relaxation length are presented. Different model configurations for wide range conditions were investigated. Comparison of results for different interactions was carried out.

Three-Dimensional Navier-Stokes Simulations with Two-Equation Turbulence Models of Intersecting Shock-Waves/Turbulent Boundary Layer at Mach 8.3

NASA Technical Reports Server (NTRS)

Bardina, J. E.; Coakley, T. J.

1994-01-01

An investigation of the numerical simulation with two-equation turbulence models of a three-dimensional hypersonic intersecting (SWTBL) shock-wave/turbulent boundary layer interaction flow is presented. The flows are solved with an efficient implicit upwind flux-difference split Reynolds-averaged Navier-Stokes code. Numerical results are compared with experimental data for a flow at Mach 8.28 and Reynolds number 5.3x10(exp 6) with crossing shock-waves and expansion fans generated by two lateral 15 fins located on top of a cold-wall plate. This experiment belongs to the hypersonic database for modeling validation. Simulations show the development of two primary counter-rotating cross-flow vortices and secondary turbulent structures under the main vortices and in each corner singularity inside the turbulent boundary layer. A significant loss of total pressure is produced by the complex interaction between the main vortices and the uplifted jet stream of the boundary layer. The overall agreement between computational and experimental data is generally good. The turbulence modeling corrections show improvements in the predictions of surface heat transfer distribution and an increase in the strength of the cross-flow vortices. Accurate predictions of the outflow flowfield is found to require accurate modeling of the laminar/turbulent boundary layers on the fin walls.

The effect of varying Mach number on crossing, glancing shocks/turbulent boundary-layer interactions

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Williams, K. E.

1991-01-01

Two crossing side-wall shocks interacting with a supersonic tunnel wall boundary layer have been investigated over a Mach number range of 2.5 to 4.0. The investigation included a range of equal shock strengths produced by shock generators at angles from 4.0 to 12.0 degrees. Results of flow visualization show that the interaction is unseparated at the low shock generator angles. With increasing shock strength, the flow begins to form a separated region that grows in size and moves forward and eventually the model unstarts. The wall static pressures show a symmetrical compression that merges on the centerline upstream of the inviscid shock locations and becomes more 1D downstream. The region of the 1D pressure gradient moves upstream with increasing shock strengths until it coincides with the leading edge of the shock generators at the limit before model unstart. At the limiting conditions the wall pressure gradients are primarily in the axial direction throughout.

Simulation of hypersonic shock wave - laminar boundary layer interactions

NASA Astrophysics Data System (ADS)

Kianvashrad, N.; Knight, D.

2017-06-01

The capability of the Navier-Stokes equations with a perfect gas model for simulation of hypersonic shock wave - laminar boundary layer interactions is assessed. The configuration is a hollow cylinder flare. The experimental data were obtained by Calspan-University of Buffalo (CUBRC) for total enthalpies ranging from 5.07 to 21.85 MJ/kg. Comparison of the computed and experimental surface pressure and heat transfer is performed and the computed §ow¦eld structure is analyzed.

Shock Tunnel Operation and Correlation of Boundary Layer Transition on a Cone in Hypervelocity Flow

DTIC Science & Technology

2013-07-01

conditions from the ideal reflected-shock pressure to measured reservoir pressure using an isentropic expan- sion. Furthermore, the 1-D nozzle computation...does not account for boundary layer growth within the nozzle , off-design operation conditions that lead to flow nonuni- formity, or vibration...translation nonequilibrium and freezing within the nozzle , which is significant for the N2 cases. For the uncertainties that can be quantified, we have combined

Separation attenuation in swept shock wave-boundary-layer interactions using different microvortex generator geometries

NASA Astrophysics Data System (ADS)

Martis, R. R.; Misra, A.

2017-09-01

A numerical study is conducted to determine the effectiveness of six different microvortex generator geometries in controlling swept shock wave/boundary-layer interactions. The geometries considered are base ramp, base ramp with declining angle of 45°, blunt ramp, split ramp, thick vanes, and ramped vanes. Microvortex generators with a gap were found to be better suited for delaying the separation. Thick vanes showed the largest delay in separation among the devices studied.

Planar Reflection of Gaseous Detonations

NASA Astrophysics Data System (ADS)

Damazo, Jason Scott

Pipes containing flammable gaseous mixtures may be subjected to internal detonation. When the detonation normally impinges on a closed end, a reflected shock wave is created to bring the flow back to rest. This study built on the work of Karnesky (2010) and examined deformation of thin-walled stainless steel tubes subjected to internal reflected gaseous detonations. A ripple pattern was observed in the tube wall for certain fill pressures, and a criterion was developed that predicted when the ripple pattern would form. A two-dimensional finite element analysis was performed using Johnson-Cook material properties; the pressure loading created by reflected gaseous detonations was accounted for with a previously developed pressure model. The residual plastic strain between experiments and computations was in good agreement. During the examination of detonation-driven deformation, discrepancies were discovered in our understanding of reflected gaseous detonation behavior. Previous models did not accurately describe the nature of the reflected shock wave, which motivated further experiments in a detonation tube with optical access. Pressure sensors and schlieren images were used to examine reflected shock behavior, and it was determined that the discrepancies were related to the reaction zone thickness extant behind the detonation front. During these experiments reflected shock bifurcation did not appear to occur, but the unfocused visualization system made certainty impossible. This prompted construction of a focused schlieren system that investigated possible shock wave-boundary layer interaction, and heat-flux gauges analyzed the boundary layer behind the detonation front. Using these data with an analytical boundary layer solution, it was determined that the strong thermal boundary layer present behind the detonation front inhibits the development of reflected shock wave bifurcation.

On Parametric Sensitivity of Reynolds-Averaged Navier-Stokes SST Turbulence Model: 2D Hypersonic Shock-Wave Boundary Layer Interactions

NASA Technical Reports Server (NTRS)

Brown, James L.

2014-01-01

Examined is sensitivity of separation extent, wall pressure and heating to variation of primary input flow parameters, such as Mach and Reynolds numbers and shock strength, for 2D and Axisymmetric Hypersonic Shock Wave Turbulent Boundary Layer interactions obtained by Navier-Stokes methods using the SST turbulence model. Baseline parametric sensitivity response is provided in part by comparison with vetted experiments, and in part through updated correlations based on free interaction theory concepts. A recent database compilation of hypersonic 2D shock-wave/turbulent boundary layer experiments extensively used in a prior related uncertainty analysis provides the foundation for this updated correlation approach, as well as for more conventional validation. The primary CFD method for this work is DPLR, one of NASA's real-gas aerothermodynamic production RANS codes. Comparisons are also made with CFL3D, one of NASA's mature perfect-gas RANS codes. Deficiencies in predicted separation response of RANS/SST solutions to parametric variations of test conditions are summarized, along with recommendations as to future turbulence approach.

Separation control in a hypersonic shock wave / turbulent boundary-layer interaction

NASA Astrophysics Data System (ADS)

Schreyer, Anne-Marie; Bermejo-Moreno, Ivan; Kim, Jeonglae; Urzay, Javier

2016-11-01

Hypersonic vehicles play a key role for affordable access to space. The associated flow fields are strongly affected by shock wave/turbulent boundary-layer interactions, and the inherent separation causes flow distortion and low-frequency unsteadiness. Microramp sub-boundary layer vortex generators are a promising means to control separation and diminish associated detrimental effects. We investigate the effect of a microramp on the low-frequency unsteadiness in a fully separated interaction. A large eddy simulation of a 33 ∘ -compression-ramp interaction was performed for an inflow Mach number of 7.2 and a Reynolds number based on momentum thickness of Reθ = 3500 , matching the experiment of Schreyer et al. (2011). For the control case, we introduced a counter-rotating vortex pair, as induced by a single microramp, into the boundary layer through the inflow conditions. We applied a dynamic mode decomposition (DMD) on both cases to identify coherent structures that are responsible for the dynamic behavior. Based on the DMD, we discuss the reduction of the separation zone and the stabilization of the shock motion achieved by the microramp, and contribute to the description of the governing mechanisms. Pursued during the 2016 CTR Summer Program at Stanford University.

Flow-around modes for a rhomboid wing with a stall vortex in the shock layer

NASA Astrophysics Data System (ADS)

Zubin, M. A.; Maximov, F. A.; Ostapenko, N. A.

2017-12-01

The results of theoretical and experimental investigation of an asymmetrical hypersonic flow around a V-shaped wing with the opening angle larger than π on the modes with attached shockwaves on forward edges, when the stall flow is implemented on the leeward wing cantilever behind the kink point of the cross contour. In this case, a vortex of nonviscous nature is formed in which the velocities on the sphere exceeding the speed of sound and resulting in the occurrence of pressure shocks with an intensity sufficient for the separation of the turbulent boundary layer take place in the reverse flow according to the calculations within the framework of the ideal gas. It is experimentally established that a separation boundary layer can exist in the reverse flow, and its structure is subject to the laws inherent to the reverse flow in the separation region of the turbulent boundary layer arising in the supersonic conic flow under the action of a shockwave incident to the boundary layer.

Numerical and experimental investigation of VG flow control for a low-boom inlet

NASA Astrophysics Data System (ADS)

Rybalko, Michael

The application of vortex generators (VGs) for shock/boundary layer interaction flow control in a novel external compression, axisymmetric, low-boom concept inlet was studied using numerical and experimental methods. The low-boom inlet design features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. This allows reduced external gas dynamic waves at high mass flow rates but suffers from flow separation near the throat and a large hub-side boundary layer at the Aerodynamic Interface Plane (AIP), which marks the inflow to the jet engine turbo-machinery. Supersonic VGs were investigated to reduce the shock-induced flow separation near the throat while subsonic VGs were investigated to reduce boundary layer radial distortion at the AIP. To guide large-scale inlet experiments, Reynolds-Averaged Navier-Stokes (RANS) simulations using three-dimensional, structured, chimera (overset) grids and the WIND-US code were conducted. Flow control cases included conventional and novel types of vortex generators at positions both upstream of the terminating normal shock (supersonic VGs) and downstream (subsonic VGs). The performance parameters included incompressible axisymmetric shape factor, post-shock separation area, inlet pressure recovery, and mass flow ratio. The design of experiments (DOE) methodology was used to select device size and location, analyze the resulting data, and determine the optimal choice of device geometry. Based on the above studies, a test matrix of supersonic and subsonic VGs was adapted for a large-scale inlet test to be conducted at the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). Comparisons of RANS simulations with data from the Fall 2010 8'x6' inlet test showed that predicted VG performance trends and case rankings for both supersonic and subsonic devices were consistent with experimental results. For example, experimental surface oil flow visualization revealed a significant post-shock separation bubble with flow recirculation for the baseline (no VG) case that was substantially broken up in the micro-ramp VG case, consistent with simulations. Furthermore, the predicted subsonic VG performance with respect to a reduction in radial distortion (quantified in terms of axisymmetric incompressible shape factor) was found to be consistent with boundary layer rake measurements. To investigate the unsteady turbulent flow features associated with the shock-induced flow separation and the hub-side boundary layer, a detached eddy simulation (DES) approach using the WIND-US code was employed to model the baseline inlet flow field. This approach yielded improved agreement with experimental data for time-averaged diffuser stagnation pressure profiles and allowed insight into the pressure fluctuations and turbulent kinetic energy distributions which may be present at the AIP. In addition, streamwise shock position statistics were obtained and compared with experimental Schlieren results. The predicted shock oscillations were much weaker than those seen experimentally (by a factor of four), which indicates that the mechanism for the experimental shock oscillations was not captured. In addition, the novel supersonic vortex generator geometries were investigated experimentally (prior to the large-scale inlet 8'x6' wind tunnel tests) in an inlet-relevant flow field containing a Mach 1.4 normal shock wave followed by a subsonic diffuser. A parametric study of device height and distance upstream of the normal shock was undertaken for split-ramp and ramped-vane geometries. Flow field diagnostics included high-speed Schlieren, oil flow visualization, and Pitot-static pressure measurements. Parameters including flow separation, pressure recovery, centerline incompressible boundary layer shape factor, and shock stability were analyzed and compared to the baseline uncontrolled case. While all vortex generators tested eliminated centerline flow separation, the presence of VGs also increased the significant three-dimensionality of the flow via increased side-wall interaction. The stronger streamwise vorticity generated by ramped-vanes also yielded improved pressure recovery and fuller boundary layer velocity profiles within the subsonic diffuser. (Abstract shortened by UMI.)

Collisionless slow shocks in magnetotail reconnection

NASA Astrophysics Data System (ADS)

Cremer, Michael; Scholer, Manfred

The kinetic structure of collisionless slow shocks in the magnetotail is studied by solving the Riemann problem of the collapse of a current sheet with a normal magnetic field component using 2-D hybrid simulations. The collapse results in a current layer with a hot isotropic distribution and backstreaming ions in a boundary layer. The lobe plasma outside and within the boundary layer exhibits a large perpendicular to parallel temperature anisotropy. Waves in both regions propagate parallel to the magnetic field. In a second experiment a spatially limited high density beam is injected into a low beta background plasma and the subsequent wave excitation is studied. A model for slow shocks bounding the reconnection layer in the magnetotail is proposed where backstreaming ions first excite obliquely propagating waves by the electromagnetic ion/ion cyclotron instability, which lead to perpendicular heating. The T⊥/T∥ temperature anisotropy subsequently excites parallel propagating Alfvén ion cyclotron waves, which are convected into the slow shock and are refracted in the downstream region.

A computational study on oblique shock wave-turbulent boundary layer interaction

NASA Astrophysics Data System (ADS)

Joy, Md. Saddam Hossain; Rahman, Saeedur; Hasan, A. B. M. Toufique; Ali, M.; Mitsutake, Y.; Matsuo, S.; Setoguchi, T.

2016-07-01

A numerical computation of an oblique shock wave incident on a turbulent boundary layer was performed for free stream flow of air at M∞ = 2.0 and Re1 = 10.5×106 m-1. The oblique shock wave was generated from a 8° wedge. Reynolds averaged Navier-Stokes (RANS) simulation with k-ω SST turbulence model was first utilized for two dimensional (2D) steady case. The results were compared with the experiment at the same flow conditions. Further, to capture the unsteadiness, a 2D Large Eddy Simulation (LES) with sub-grid scale model WMLES was performed which showed the unsteady effects. The frequency of the shock oscillation was computed and was found to be comparable with that of experimental measurement.

Shock tube and shock wave research; Proceedings of the Eleventh International Symposium, University of Washington, Seattle, Wash., July 11-14, 1977

NASA Technical Reports Server (NTRS)

Ahlborn, B. (Editor); Hertzberg, A.; Russell, D.

1978-01-01

Papers are presented on the applications of shock-wave technology to the study of hydrodynamics, the use of the pressure-wave machine for charging diesel engines, and measurements of the heat-transfer rate in gas-turbine components. Consideration is given to shock propagation along 90-degree bends, the explosive dissemination of liquids, and rotational and vibrational relaxation behind weak shock waves in water vapor. Shock phenomena associated with expansion flows are described and stratospheric-related research using the shock tube is outlined. Attention is given to shock-wave ignition of magnesium powders, Mach reflection and boundary layers, and transition in the shock-induced unsteady boundary layer on a flat plate. Shock-tube measurements of induction and post-induction rates for low-Btu gas mixtures are presented and shock-initiated ignition in COS-N2O-Ar mixtures is described. Cluster growth rates in supersaturated lead vapor are presented and a study of laser-induced plasma motion in a solenoidal magnetic field is reviewed.

An experimental investigation of compressible three-dimensional boundary layer flow in annular diffusers

NASA Technical Reports Server (NTRS)

Om, Deepak; Childs, Morris E.

1987-01-01

An experimental study is described in which detailed wall pressure measurements have been obtained for compressible three-dimensional unseparated boundary layer flow in annular diffusers with and without normal shock waves. Detailed mean flow-field data were also obtained for the diffuser flow without a shock wave. Two diffuser flows with shock waves were investigated. In one case, the normal shock existed over the complete annulus whereas in the second case, the shock existed over a part of the annulus. The data obtained can be used to validate computational codes for predicting such flow fields. The details of the flow field without the shock wave show flow reversal in the circumferential direction on both inner and outer surfaces. However, there is a lag in the flow reversal between the inner nad the outer surfaces. This is an interesting feature of this flow and should be a good test for the computational codes.

Direct measurements of wall shear stress by buried wire gages in a shock-wave boundary-layer interaction region

NASA Technical Reports Server (NTRS)

Murthy, V. S.; Rose, W. C.

1977-01-01

Detailed measurements of wall shear stress (skin friction) were made with specially developed buried wire gages in the interaction regions of a Mach 2.9 turbulent boundary layer with externally generated shocks. Separation and reattachment points inferred by these measurements support the findings of earlier experiments which used a surface oil flow technique and pitot profile measurements. The measurements further indicate that the boundary layer tends to attain significantly higher skin-friction values downstream of the interaction region as compared to upstream. Comparisons between measured wall shear stress and published results of some theoretical calculation schemes show that the general, but not detailed, behavior is predicted well by such schemes.

Numerical solution for the interaction of shock wave with laminar boundary layer in two-dimensional flow on a flat plate. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Landau, U.

1984-01-01

The finite difference computation method was investigated for solving problems of interaction between a shock wave and a laminar boundary layer, through solution of the complete Navier-Stokes equations. This method provided excellent solutions, was simple to perform and needed a relatively short solution time. A large number of runs for various flow conditions could be carried out from which the interaction characteristics and principal factors that influence interaction could be studied.

The behaviour of turbulence anisotropy through shock waves and expansions

NASA Technical Reports Server (NTRS)

Minh, H. H.; Kollmann, W.; Vandromme, D.

1985-01-01

A second order closure has been implemented in an implicit Navier-Stokes solver to study the behavior of the Reynolds stresses under the influence of severe pressure gradients. In the boundary layer zone, the strongly sheared character of the mean flow dominates the turbulence generation mechanisms. However, the pressure gradients play also a very important role for these processes, but at different locations within the boundary layer. This aspect may be emphasized by the analysis of turbulence anisotropy through shock waves and expansions.

Flap effectiveness appraisal for winged re-entry vehicles

NASA Astrophysics Data System (ADS)

de Rosa, Donato; Pezzella, Giuseppe; Donelli, Raffaele S.; Viviani, Antonio

2016-05-01

The interactions between shock waves and boundary layer are commonplace in hypersonic aerodynamics. They represent a very challenging design issue for hypersonic vehicle. A typical example of shock wave boundary layer interaction is the flowfield past aerodynamic surfaces during control. As a consequence, such flow interaction phenomena influence both vehicle aerodynamics and aerothermodynamics. In this framework, the present research effort describes the numerical activity performed to simulate the flowfield past a deflected flap in hypersonic flowfield conditions for a winged re-entry vehicle.

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Davis, David O.

2015-01-01

Preliminary results of an experimental investigation of a Mach 2.5 two-dimensional axisymmetric shock-wave/boundary-layer interaction (SWBLI) are presented. The purpose of the investigation is to create a SWBLI dataset specifically for CFD validation purposes. Presented herein are the details of the facility and preliminary measurements characterizing the facility and interaction region. The results will serve to define the region of interest where more detailed mean and turbulence measurements will be made.

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Davis, David Owen

2015-01-01

Preliminary results of an experimental investigation of a Mach 2.5 two-dimensional axisymmetric shock-wave/ boundary-layer interaction (SWBLI) are presented. The purpose of the investigation is to create a SWBLI dataset specifically for CFD validation purposes. Presented herein are the details of the facility and preliminary measurements characterizing the facility and interaction region. These results will serve to define the region of interest where more detailed mean and turbulence measurements will be made.

Swept shock/boundary layer interaction experiments in support of CFD code validation

NASA Technical Reports Server (NTRS)

Settles, G. S.; Lee, Y.

1990-01-01

Research on the topic of shock wave/turbulent boundary layer interaction was carried out. Skin friction and surface pressure measurements in fin-induced, swept interactions were conducted, and heat transfer measurements in the same flows are planned. The skin friction data for a strong interaction case (Mach 4, fin-angles equal 16 and 20 degrees) were obtained, and their comparison with computational results was published. Surface pressure data for weak-to-strong fin interactions were also obtained.

Budget of Turbulent Kinetic Energy in a Shock Wave Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Vyas, Manan; Waindim, Mbu; Gaitonde, Datta

2016-01-01

Implicit large-eddy simulation (ILES) of a shock wave boundary-layer interaction (SBLI) was performed. Quantities present in the exact equation of the turbulent kinetic energy (TKE) transport were accumulated. These quantities will be used to calculate the components of TKE-like production, dissipation, transport, and dilatation. Correlations of these terms will be presented to study the growth and interaction between various terms. A comparison with its RANS (Reynolds-Averaged Navier-Stokes) counterpart will also be presented.

Analysis of the interaction of a weak normal shock wave with a turbulent boundary layer

NASA Technical Reports Server (NTRS)

Melnik, R. E.; Grossman, B.

1974-01-01

The method of matched asymptotic expansions is used to analyze the interaction of a normal shock wave with an unseparated turbulent boundary layer on a flat surface at transonic speeds. The theory leads to a three-layer description of the interaction in the double limit of Reynolds number approaching infinity and Mach number approaching unity. The interaction involves an outer, inviscid rotational layer, a constant shear-stress wall layer, and a blending region between them. The pressure distribution is obtained from a numerical solution of the outer-layer equations by a mixed-flow relaxation procedure. An analytic solution for the skin friction is determined from the inner-layer equations. The significance of the mathematical model is discussed with reference to existing experimental data.

Effect of aspect ratio on sidewall boundary-layer influence in two-dimensional airfoil testing

NASA Technical Reports Server (NTRS)

Murthy, A. V.

1986-01-01

The effect of sidewall boundary layers in airfoil testing in two-dimensional wind tunnels is investigated. The non-linear crossflow velocity variation induced because of the changes in the sidewall boundary-layer thickness is represented by the flow between a wavy wall and a straight wall. Using this flow model, a correction for the sidewall boundary-layer effects is derived in terms of the undisturbed sidewall boundary-layer properties, the test Mach number and the airfoil aspect ratio. Application of the proposed correction to available experimental data showed good correlation for the shock location and pressure distribution on airfoils.

Computation of the shock-wave boundary layer interaction with flow separation

NASA Technical Reports Server (NTRS)

Ardonceau, P.; Alziary, T.; Aymer, D.

1980-01-01

The boundary layer concept is used to describe the flow near the wall. The external flow is approximated by a pressure displacement relationship (tangent wedge in linearized supersonic flow). The boundary layer equations are solved in finite difference form and the question of the presence and unicity of the solution is considered for the direct problem (assumed pressure) or converse problem (assumed displacement thickness, friction ratio). The coupling algorithm presented implicitly processes the downstream boundary condition necessary to correctly define the interacting boundary layer problem. The algorithm uses a Newton linearization technique to provide a fast convergence.

An eddy-viscosity treatment of the unsteady turbulent boundary layer on a flat plate in an expansion tube

NASA Technical Reports Server (NTRS)

Gupta, R. N.; Trimpi, R. L.

1974-01-01

An analysis is presented for the relaxation of a turbulent boundary layer on a semiinfinite flat plate after passage of a shock wave and a trailing driver gas-driven gas interface. The problem has special application to expansion tube flows. The flow-governing equations have been transformed into the Lamcrocco variables. The numerical results indicate that a fully turbulent boundary layer relaxes faster to the final steady-state values of heat transfer and skin-friction than a fully laminar boundary layer.

New diagnostic technique for the study of turbulent boundary-layer separation

NASA Technical Reports Server (NTRS)

Horstman, C. C.; Owen, F. K.

1974-01-01

Description of a diagnostic technique for determining the unsteady character of turbulent boundary-layer separation. The technique uses thin platinum films mounted flush with the model surface. Voltages from these films provide measurements related to the flow character above the film. For illustration, results obtained by this technique are presented for the interaction of a hypersonic shock wave and a turbulent boundary layer, with and without separation.

Characterization of the Shear Layer in a Mach 3 Shock/Turbulent Boundary Layer Interaction

NASA Astrophysics Data System (ADS)

Helm, Clara; Priebe, Stephan; Li, Justine; Dupont, Pierre; Martin, Pino

2013-11-01

The unsteady motion of fully separated shock and turbulent boundary layers interactions (STBLIs) is characterized by an energized low-frequency motion that is two orders of magnitude lower than that of the incoming turbulence. In addition, the spectra shows significant energy content at frequency that is between the characteristic low frequency and the higher frequency motions of the incoming turbulence. The intermediate frequency content is hypothesized to be associated with the existence of Kelvin-Helmholtz type structures, which form in the shear layer downstream of the separation shock and are shed near the reattachment point downstream of the interaction. The current research is concerned with investigating the origins of the intermediate frequencies, and how they may be related to or possibly influence the low-frequency unsteadiness. Specifically, LES data of a Mach 3 STBLI over a 24o ramp are used to estimate convection velocities within the shear layer downstream of the shock. In addition, Brown and Thomas type correlations are used to estimate time and length scales of the eddies in the shear layer. This work is supported by the Air Force Office of Scientific Research under grant AF/9550-10-1-0164.

Bifurcation parameters of a reflected shock wave in cylindrical channels of different roughnesses

NASA Astrophysics Data System (ADS)

Penyazkov, O.; Skilandz, A.

2018-03-01

To investigate the effect of bifurcation on the induction time in cylindrical shock tubes used for chemical kinetic experiments, one should know the parameters of the bifurcation structure of a reflected shock wave. The dynamics and parameters of the shock wave bifurcation, which are caused by reflected shock wave-boundary layer interactions, are studied experimentally in argon, in air, and in a hydrogen-nitrogen mixture for Mach numbers M = 1.3-3.5 in a 76-mm-diameter shock tube without any ramp. Measurements were taken at a constant gas density behind the reflected shock wave. Over a wide range of experimental conditions, we studied the axial projection of the oblique shock wave and the pressure distribution in the vicinity of the triple Mach configuration at 50, 150, and 250 mm from the endwall, using side-wall schlieren and pressure measurements. Experiments on a polished shock tube and a shock tube with a surface roughness of 20 {μ }m Ra were carried out. The surface roughness was used for initiating small-scale turbulence in the boundary layer behind the incident shock wave. The effect of small-scale turbulence on the homogenization of the transition zone from the laminar to turbulent boundary layer along the shock tube perimeter was assessed, assuming its influence on a subsequent stabilization of the bifurcation structure size versus incident shock wave Mach number, as well as local flow parameters behind the reflected shock wave. The influence of surface roughness on the bifurcation development and pressure fluctuations near the wall, as well as on the Mach number, at which the bifurcation first develops, was analyzed. It was found that even small additional surface roughness can lead to an overshoot in pressure growth by a factor of two, but it can stabilize the bifurcation structure along the shock tube perimeter.

An experimental study of three-dimensional shock wave/boundary layer interactions generated by sharp fins

NASA Technical Reports Server (NTRS)

Lu, F. K.; Settles, G. S.; Bogdonoff, S. M.

1983-01-01

The interaction between a turbulent boundary layer and a shock wave generated by a sharp fin with leading edge sweepback was investigated. The incoming flow was at Mach 2.96 and at a unit Reynolds number of 63 x 10 to the 6th power 0.1 m. The approximate incoming boundary layer thickness was either 4 mm or 17 mm. The fins used were at 5 deg, 9 deg and 15 deg incidence and had leading edge sweepback from 0 deg to 65 deg. The tests consisted of surface kerosene lampblack streak visualization, surface pressure measurements, shock wave shape determination by shadowgraphs, and localized vapor screen visualization. The upstream influence lengths of the fin interactions were correlated using viscous and inviscid flow parameters. The parameters affecting the surface features close to the fin and way from the fin were also identified. Essentially, the surface features in the farfield were found to be conical.

Fuselage Structure Response to Boundary Layer, Tonal Sound, and Jet Noise

NASA Technical Reports Server (NTRS)

Maestrello, L.

2004-01-01

Experiments have been conducted to study the response of curved aluminum and graphite-epoxy fuselage structures to flow and sound loads from turbulent boundary layer, tonal sound, and jet noise. Both structures were the same size. The aluminum structure was reinforced with tear stoppers, while the graphite-epoxy structure was not. The graphite-epoxy structure weighed half as much as the aluminum structure. Spatiotemporal intermittence and chaotic behavior of the structural response was observed, as jet noise and tonal sound interacted with the turbulent boundary layer. The fundamental tone distributed energy to other components via wave interaction with the turbulent boundary layer. The added broadband sound from the jet, with or without a shock, influenced the responses over a wider range of frequencies. Instantaneous spatial correlation indicates small localized spatiotemporal regions of convected waves, while uncorrelated patterns dominate the larger portion of the space. By modifying the geometry of the tear stoppers between panels and frame, the transmitted and reflected waves of the aluminum panels were significantly reduced. The response level of the graphite-epoxy structure was higher, but the noise transmitted was nearly equal to that of the aluminum structure. The fundamental shock mode is between 80 deg and 150 deg and the first harmonic is between 20 deg and 80 deg for the underexpanded supersonic jet impinging on the turbulent boundary layer influencing the structural response. The response of the graphite-epoxy structure due to the fundamental mode of the shock impingement was stabilized by an externally fixed oscillator.

Effective Boundary Conditions for Continuum Method of Investigation of Rarefied Gas Flow over Blunt Body

NASA Astrophysics Data System (ADS)

Brykina, I. G.; Rogov, B. V.; Semenov, I. L.; Tirskiy, G. A.

2011-05-01

Super- and hypersonic rarefied gas flow over blunt bodies is investigated by using asymptotically correct viscous shock layer (VSL) model with effective boundary conditions and thin viscous shock layer model. Correct shock and wall conditions for VSL are proposed with taking into account terms due to the curvature which are significant at low Reynolds number. These conditions improve original Davis's VSL model [1]. Numerical calculation of Krook equation [2] is carried out to verify continuum results. Continuum numerical and asymptotic solutions are compared with kinetic solution, free-molecule flow solution and with DSMC solutions [3, 4, 5] over a wide range of free-stream Knudsen number Kn∞. It is shown that taking into account terms with shock and surface curvatures have a pronounced effect on skin friction and heat-transfer in transitional flow regime. Using the asymptotically correct VSL model with effective boundary conditions significantly extends the range of its applicability to higher Kn∞ numbers.

Flowfield dynamics in blunt fin-induced shock wave/turbulent boundary layer interactions

NASA Technical Reports Server (NTRS)

Dolling, David S.; Brusniak, Leon

1994-01-01

Fluctuating wall pressure measurements have been made on centerline upstream of a blunt fin in a Mach 5 turbulent boundary layer. By examining the ensemble averaged wall pressure distributions for different separation shock foot positions, it has been shown that local fluctuating wall pressure measurements are due to a distinct pressure distribution, Rho(sub i), which undergoes a stretching and flattening effect as its upstream boundary translates aperiodically between the upstream influence and separation lines. The locations of the maxima and minima in the wall pressure standard deviation can be accurately predicted using this distribution, providing quantitative confirmation of the model. This model also explains the observed cross-correlations and ensemble average measurements within the interaction. Using the Rho(sub i) model, wall pressure signals from under the separated flow region were used to reproduce the position-time history of the separation shock foot. Further, the negative time delay peak in the cross-correlation between the predicted and actual shock foot histories suggests that the separated region fluctuations precede shock foot motion. The unsteady behavior of the primary horseshoe vortex and its relation to the unsteady separation shock are described.

Computation of Viscous-Inviscid Interactions

DTIC Science & Technology

1981-02-01

porte stir 11Epaisseur de d~placement Pizr) at non sur la direction angu- laire de Il’couiement f9"(X,) . Doe le cas incompressible, par example, !a...into the boundary layer. The diffraction of the shock wave by the nonuniform flow in the boundary layer leads to significant normal pressrre gradients...deivative. This equ . wemr’s the propagation of small disturbances in the nonuniform flow in the boundary layer. Within this model, disturbances

Unsteady separation in sharp fin-induced shock wave/turbulent boundary layer interaction at Mach 5

NASA Technical Reports Server (NTRS)

Schmisseur, J. D.; Dolling, D. S.

1992-01-01

Fluctuating wall-pressure measurements are made in shock-wave/turbulent-boundary-layer interactions generated by sharp/unswept fins at angles of attack of 16, 18, 20, 22, 24, 26, and 28 degrees at Mach 5. The experiment was conducted under approximately adiabatic wall temperature conditions. The mean and rms pressure distributions can be collapsed in conical coordinates. The wall-pressure signal near separation is intermittent for all angles of attack (16-28 deg) and is qualitatively similar to that measured in unswept flows. However, the shock frequencies are higher - about 5 kHz compared to 0.5-1 kHz. Over the range of sweepbacks examined, from 25-55 deg, the spectral content of the fluctuating pressures does not change. Thus, the increase in separation-shock frequency from 1 to 5 kHz occurs at lower interaction sweepback and is not a continuous process with increasing sweepback. Power spectra at the position of maximum rms in the intermittent region for interactions in different incoming boundary layers have the same center frequency. The maximum rms in the intermittent region correlates with interaction sweepback, not with overall inviscid pressure rise.

Shock wave/turbulent boundary layer interaction in the flow field of a tri-dimension wind tunnel

NASA Technical Reports Server (NTRS)

Benay, R.; Pot, T.

1986-01-01

The first results of a thorough experimental analysis of a strong three-dimensional shock-wave/turbulent boundary-layer interaction occurring in a three dimensional transonic channel are presented. The aim of this experiment is to help in the physical understanding of a complex field, including several separations, and to provide a well documented case to test computational methods. The flowfield has been probed in many points by means of a three-component laser Doppler velocimeter. The results presented relate only to the mean velocity field. They clearly show the formation in the flow of a strong vortical motion resulting from the shock wave interaction.

Effect of particle momentum transfer on an oblique-shock-wave/laminar-boundary-layer interaction

NASA Astrophysics Data System (ADS)

Teh, E.-J.; Johansen, C. T.

2016-11-01

Numerical simulations of solid particles seeded into a supersonic flow containing an oblique shock wave reflection were performed. The momentum transfer mechanism between solid and gas phases in the shock-wave/boundary-layer interaction was studied by varying the particle size and mass loading. It was discovered that solid particles were capable of significant modulation of the flow field, including suppression of flow separation. The particle size controlled the rate of momentum transfer while the particle mass loading controlled the magnitude of momentum transfer. The seeding of micro- and nano-sized particles upstream of a supersonic/hypersonic air-breathing propulsion system is proposed as a flow control concept.

An analysis of the relaxation of laminar boundary layer on a flat plate after passage of an interface with application to expansion-tube flows

NASA Technical Reports Server (NTRS)

Gupta, R. N.

1972-01-01

The relaxation of the accelerating-gas boundary layer to the test-gas boundary layer over a flat plate in an expansion tube is analyzed. Several combinations of test gas and acceleration gas are considered. The problem is treated in two conically similar limits: (1) when the time lag between the arrival of the shock and the interface at the leading edge of the plate is very large, and (2) when this lag is negligible. The time-dependent laminar-boundary-layer equations of a binary mixture of perfect gases are taken as the flow-governing equations. This coupled set of differential equations, written in terms of the Lam-Crocco variables, has been solved by a line-relaxation finite-difference techniques. The results presented include the Stanton number and the local skin-friction coefficient as functions of shock Mach number and the nondimensional distance-time variable. The results indicate that more than 95 percent of the test-gas boundary layer exists over a length, measured from the leading edge of the plate, equal to about three-tenths of the distance traversed by the interface in the free stream.

Skin friction measurements by laser interferometry in swept shock wave/turbulent boundary-layer interactions

NASA Technical Reports Server (NTRS)

Kim, Kwang-Soo; Settles, Gary S.

1988-01-01

The laser interferometric skin friction meter was used to measure wall shear stress distributions in two interactions of fin-generated swept shock waves with turbulent boundary layers. The basic research configuration was an unswept sharp-leading-edge fin of variable angle mounted on a flatplate. The results indicate that such measurements are practical in high-speed interacting flows, and that a repeatability of + or - 6 percent or better is possible. Marked increases in wall shear were observed in both swept interactions tested.

Sensitivity of shock boundary-layer interactions to weak geometric perturbations

NASA Astrophysics Data System (ADS)

Kim, Ji Hoon; Eaton, John K.

2016-11-01

Shock-boundary layer interactions can be sensitive to small changes in the inlet flow and boundary conditions. Robust computational models must capture this sensitivity, and validation of such models requires a suitable experimental database with well-defined inlet and boundary conditions. To that end, the purpose of this experiment is to systematically document the effects of small geometric perturbations on a SBLI flow to investigate the flow physics and establish an experimental dataset tailored for CFD validation. The facility used is a Mach 2.1, continuous operation wind tunnel. The SBLI is generated using a compression wedge; the region of interest is the resulting reflected shock SBLI. The geometric perturbations, which are small spanwise rectangular prisms, are introduced ahead of the compression ramp on the opposite wall. PIV is used to study the SBLI for 40 different perturbation geometries. Results show that the dominant effect of the perturbations is a global shift of the SBLI itself. In addition, the bumps introduce weaker shocks of varying strength and angles, depending on the bump height and location. Various scalar validation metrics, including a measure of shock unsteadiness, and their uncertainties are also computed to better facilitate CFD validation. Ji Hoon Kim is supported by an OTR Stanford Graduate Fellowship.

Kinetic theory and turbulent discontinuities. [shock tube flow

NASA Technical Reports Server (NTRS)

Johnson, J. A., III; I, L.; Li, Y.; Ramaian, R.; Santigo, J. P.

1981-01-01

Shock tube discontinuities were used to test and extend a kinetic theory of turbulence. In shock wave and contact surface fluctuations, coherent phenomena were found which provide new support for the microscopic nonempirical approach to turbulent systems, especially those with boundary layer-like instabilities.

An entropy and viscosity corrected potential method for rotor performance prediction

NASA Technical Reports Server (NTRS)

Bridgeman, John O.; Strawn, Roger C.; Caradonna, Francis X.

1988-01-01

An unsteady Full-Potential Rotor code (FPR) has been enhanced with modifications directed at improving its drag prediction capability. The shock generated entropy has been included to provide solutions comparable to the Euler equations. A weakly interacted integral boundary layer has also been coupled to FPR in order to estimate skin-friction drag. Pressure distributions, shock positions, and drag comparisons are made with various data sets derived from two-dimensional airfoil, hovering, and advancing high speed rotor tests. In all these comparisons, the effect of the nonisentropic modification improves (i.e., weakens) the shock strength and wave drag. In addition, the boundary layer method yields reasonable estimates of skin-friction drag. Airfoil drag and hover torque data comparisons are excellent, as are predicted shock strength and positions for a high speed advancing rotor.

Simulation and stability analysis of oblique shock-wave/boundary-layer interactions at Mach 5.92

NASA Astrophysics Data System (ADS)

Hildebrand, Nathaniel; Dwivedi, Anubhav; Nichols, Joseph W.; Jovanović, Mihailo R.; Candler, Graham V.

2018-01-01

We investigate flow instability created by an oblique shock wave impinging on a Mach 5.92 laminar boundary layer at a transitional Reynolds number. The adverse pressure gradient of the oblique shock causes the boundary layer to separate from the wall, resulting in the formation of a recirculation bubble. For sufficiently large oblique shock angles, the recirculation bubble is unstable to three-dimensional perturbations and the flow bifurcates from its original laminar state. We utilize direct numerical simulation (DNS) and global stability analysis to show that this first occurs at a critical shock angle of θ =12 .9∘ . At bifurcation, the least-stable global mode is nonoscillatory and it takes place at a spanwise wave number β =0.25 , in good agreement with DNS results. Examination of the critical global mode reveals that it originates from an interaction between small spanwise corrugations at the base of the incident shock, streamwise vortices inside the recirculation bubble, and spanwise modulation of the bubble strength. The global mode drives the formation of long streamwise streaks downstream of the bubble. While the streaks may be amplified by either the lift-up effect or by Görtler instability, we show that centrifugal instability plays no role in the upstream self-sustaining mechanism of the global mode. We employ an adjoint solver to corroborate our physical interpretation by showing that the critical global mode is most sensitive to base flow modifications that are entirely contained inside the recirculation bubble.

Temperature determination of shock layer using spectroscopic techniques

NASA Technical Reports Server (NTRS)

Akundi, Murty A.

1989-01-01

Shock layer temperature profiles are obtained through analysis of radiation from shock layers produced by a blunt body inserted in an arc jet flow. Spectral measurements of N2(+) have been made at 0.5 inch, 1.0 inch, and 1.4 inches from the blunt body. A technique is developed to measure the vibrational and rotational temperatures of N2(+). Temperature profiles from the radiation layers show a high temperature near the shock front and decreasing temperature near the boundary layer. Precise temperature measurements could not be made using this technique due to the limited resolution. Use of a high resolution grating will help to make a more accurate temperature determination. Laser induced fluorescence technique is much better since it gives the scope for selective excitation and a better spacial resolution.

Transonic Shock Oscillations and Wing Flutter Calculated with an Interactive Boundary Layer Coupling Method

NASA Technical Reports Server (NTRS)

Edwards, John W.

1996-01-01

A viscous-inviscid interactive coupling method is used for the computation of unsteady transonic flows involving separation and reattachment. A lag-entrainment integral boundary layer method is used with the transonic small disturbance potential equation in the CAP-TSDV (Computational Aeroelasticity Program - Transonic Small Disturbance) code. Efficient and robust computations of steady and unsteady separated flows, including steady separation bubbles and self-excited shock-induced oscillations are presented. The buffet onset boundary for the NACA 0012 airfoil is accurately predicted and shown computationally to be a Hopf bifurcation. Shock-induced oscillations are also presented for the 18 percent circular arc airfoil. The oscillation onset boundaries and frequencies are accurately predicted, as is the experimentally observed hysteresis of the oscillations with Mach number. This latter stability boundary is identified as a jump phenomenon. Transonic wing flutter boundaries are also shown for a thin swept wing and for a typical business jet wing, illustrating viscous effects on flutter and the effect of separation onset on the wing response at flutter. Calculations for both wings show limit cycle oscillations at transonic speeds in the vicinity of minimum flutter speed indices.

Tangential blowing for control of strong normal shock - Boundary layer interactions on inlet ramps

NASA Technical Reports Server (NTRS)

Schwendemann, M. F.; Sanders, B. W.

1982-01-01

The use of tangential blowing from a row of holes in an aft facing step is found to provide good control of the ramp boundary layer, normal shock interaction on a fixed geometry inlet over a wide range of inlet mass flow ratios. Ramp Mach numbers of 1.36 and 1.96 are investigated. The blowing geometry is found to have a significant effect on system performance at the highest Mach number. The use of high-temperature air in the blowing system, however, has only a slight effect on performance. The required blowing rates are significantly high for the most severe test conditions. In addition, the required blowing coefficient is found to be proportional to the normal shock pressure rise.

The effects of vortex structure and vortex translation on the tropical cyclone boundary layer wind field

NASA Astrophysics Data System (ADS)

Williams, Gabriel J.

2015-03-01

The effects of vortex translation and radial vortex structure in the distribution of boundary layer winds in the inner core of mature tropical cyclones are examined using a high-resolution slab model and a multilevel model. It is shown that the structure and magnitude of the wind field (and the corresponding secondary circulation) depends sensitively on the radial gradient of the gradient wind field above the boundary layer. Furthermore, it is shown that vortex translation creates low wave number asymmetries in the wind field that rotate anticyclonically with height. A budget analysis of the steady state wind field for both models was also performed in this study. Although the agradient force drives the evolution of the boundary layer wind field for both models, it is shown that the manner in which the boundary layer flow responds to this force differs between the two model representations. In particular, the inner core boundary layer flow in the slab model is dominated by the effects of horizontal advection and horizontal diffusion, leading to the development of shock structures in the model. Conversely, the inner core boundary layer flow in the multilevel model is primarily influenced by the effects of vertical advection and vertical diffusion, which eliminates shock structures in this model. These results further indicate that special care is required to ensure that qualitative applications from slab models are not unduly affected by the neglect of vertical advection. This article was corrected on 31 MAR 2015. See the end of the full text for details.

Computation of airfoil buffet boundaries

NASA Technical Reports Server (NTRS)

Levy, L. L., Jr.; Bailey, H. E.

1981-01-01

The ILLIAC IV computer has been programmed with an implicit, finite-difference code for solving the thin layer compressible Navier-Stokes equation. Results presented for the case of the buffet boundaries of a conventional and a supercritical airfoil section at high Reynolds numbers are found to be in agreement with experimentally determined buffet boundaries, especially at the higher freestream Mach numbers and lower lift coefficients where the onset of unsteady flows is associated with shock wave-induced boundary layer separation.

Swept Impinging Oblique Shock/Boundary-Layer Interactions

NASA Astrophysics Data System (ADS)

Little, Jesse; Threadgill, James; Stab, Ilona

2016-11-01

Oblique shock waves impinging on boundary layers are common flow features associated with high-speed flows around complex body geometries and through internal channel flows. The increasingly three-dimensional surface geometries of modern vehicles has led to a prevalence of complex shock/boundary-layer interactions. Sweep has been observed to vary the interaction structure, unsteadinesses, and similarity scalings. Sharp-fins and highly-swept ramps have been noted to induce a quasi-conical development of the interaction, in contrast to a quasi-cylindrical scaling observed in low-sweep interactions. However, swept impinging oblique shock cases have largely been overlooked, with evidence of only cylindrical similarities observed in hypersonic conditions. Flow deflection beyond the maximum turning angle has been proposed as the mechanism for conical interaction development but such behavior has not been established for the present configuration. This study examines the effect of sweep on the interaction induced by a 12.5° generator in Mach 2.3 flow using oil-flow, Schlieren and PIV. Results document the development of similarity scalings at various angles of sweep, and highlight the difficulty in replicating a quasi-infinite span conditions in a moderately sized wind tun Supported by the Air Force Office of Scientific Research (FA9550-15-1-0430) and Raytheon Missile Systems.

Role of secondary flows on flow separation induced by shock/boundary layer interaction in supersonic inlets

NASA Astrophysics Data System (ADS)

Morajkar, Rohan

Flow separation in the scramjet air intakes is one of the reasons of failure of these engines which rely on shock waves to achieve flow compression. The shock waves interact with the boundary layers (Shock/ Boundary Layer Interaction or SBLI) on the intake walls inducing adverse pressure gradients causing flow separation. In this experimental study we investigate the role of secondary flows associated with the corners of ducted flows and identify the mechanisms by which they affect flow separation induced by a shock wave interacting with the boundary layers developing along supersonic inlets. The coupling between flow three-dimensionality, shock waves and secondary flows is in fact a key aspect that limits the performance and control of supersonic inlets. The study is conducted at the University of Michigan Glass Supersonic Wind Tunnel (GSWT). This facility replicates some of the features of the three-dimensional (3D) flow-field in a low aspect ratio supersonic inlet. The study uses stereoscopic particle image velocimetry (SPIV) to measure the three-component (3C) velocity field on several orthogonal planes, and thus allows us to identify the length scales of separation, its locations and statistical properties. Furthermore, these measurements allow us to extract the 3D structure of the underlying vortical features, which are important in determining the overall structure of separated regions and their dynamics. The measurements and tools developed are used to study flow fields of three cases: (1) Moderately strong SBLI (Mach 2.75 with 6° deflection), (2) weak SBLI (Mach 2.75 with 4.6° deflection) and (3) secondary corner flows in empty channels. In the configuration of the initial work (moderately strong SBLI), the shock wave system interacts with the boundary layers on the sidewall and the floor of the duct (inlet), thus generating both a swept-shock and an incident-shock interactions. Furthermore, the swept-shock interaction taking place on the sidewalls interacts with the secondary flows in the corners of the tunnel, which are prone to separation. This interaction causes major flow separation on the sidewall as fluid is swept from the sidewall. Flow separation on the floor should be expected given the strength of the SBLI (moderately strong case), but it is instead not observed in the mean flow fields. Our hypothesis is that interacting secondary flows are one of the factors responsible for the sidewall separation and directing the incoming flow towards the center-plane to stabilize and energize the flow on the center of the duct, thus preventing or at least reducing, flow separation on the floor. The secondary flows in an empty tunnel are then investigated to study their evolution and effects on the primary flow field to identify potential separation sites. The results from the empty tunnel experiments are then used to predict locations of flow separations in the moderately strong and weak SBLIs. The predictions were found to be in agreement with the observations.

Shock Generation and Control Using DBD Plasma Actuators

NASA Technical Reports Server (NTRS)

Patel, Mehul P.; Cain, Alan B.; Nelson, Christopher C.; Corke, Thomas C.; Matlis, Eric H.

2012-01-01

This report is the final report of a NASA Phase I SBIR contract, with some revisions to remove company proprietary data. The Shock Boundary Layer Interaction (SBLI) phenomena in a supersonic inlet involve mutual interaction of oblique shocks with boundary layers, forcing the boundary layer to separate from the inlet wall. To improve the inlet efficiency, it is desired to prevent or delay shock-induced boundary layer separation. In this effort, Innovative Technology Applications Company (ITAC), LLC and the University of Notre Dame (UND) jointly investigated the use of dielectric-barrier-discharge (DBD) plasma actuators for control of SBLI in a supersonic inlet. The research investigated the potential for DBD plasma actuators to suppress flow separation caused by a shock in a turbulent boundary layer. The research involved both numerical and experimental investigations of plasma flow control for a few different SBLI configurations: (a) a 12 wedge flow test case at Mach 1.5 (numerical and experimental), (b) an impinging shock test case at Mach 1.5 using an airfoil as a shock generator (numerical and experimental), and (c) a Mach 2.0 nozzle flow case in a simulated 15 X 15 cm wind tunnel with a shock generator (numerical). Numerical studies were performed for all three test cases to examine the feasibility of plasma flow control concepts. These results were used to guide the wind tunnel experiments conducted on the Mach 1.5 12 degree wedge flow (case a) and the Mach 1.5 impinging shock test case (case b) which were at similar flow conditions as the corresponding numerical studies to obtain experimental evidence of plasma control effects for SBLI control. The experiments also generated data that were used in validating the numerical studies for the baseline cases (without plasma actuators). The experiments were conducted in a Mach 1.5 test section in the University of Notre Dame Hessert Laboratory. The simulation results from cases a and b indicated that multiple spanwise actuators in series and at a voltage of 75 kVp-p could fully suppress the flow separation downstream of the shock. The simulation results from case c showed that the streamwise plasma actuators are highly effective in creating pairs of counter-rotating vortices, much like the mechanical vortex generators, and could also potentially have beneficial effects for SBLI control. However, to achieve these effects, the positioning and the quantity of the DBD actuators used must be optimized. The wind tunnel experiments mapped the baseline flow with good agreement to the numerical simulations. The experimental results were conducted with spanwise actuators for cases a and b, but were limited by the inability to generate a sufficiently high voltage due to arcing in the wind-tunnel test-section. The static pressure in the tunnel was lower than the static pressure in an inlet at flight conditions, promoting arching and degrading the actuator performance.

Flow-field measurements in the windward surface shock layer of space shuttle orbiter configurations at Mach number 8. [wind tunnel tests of scale models

NASA Technical Reports Server (NTRS)

Martindale, W. R.; Carter, L. D.

1975-01-01

Pitot pressure and total-temperature measurements were made in the windward surface shock layer of two 0.0175-scale space shuttle orbiter models at simulated re-entry conditions. Corresponding surface static pressure measurements were also made. Flow properties at the edge of the model boundary layer were derived from these measurements and compared with values calculated using conventional methods.

CFD on hypersonic flow geometries with aeroheating

NASA Astrophysics Data System (ADS)

Sohail, Muhammad Amjad; Chao, Yan; Hui, Zhang Hui; Ullah, Rizwan

2012-11-01

The hypersonic flowfield around a blunted cone and cone-flare exhibits some of the major features of the flows around space vehicles, e.g. a detached bow shock in the stagnation region and the oblique shock wave/boundary layer interaction at the cone-flare junction. The shock wave/boundary layer interaction can produce a region of separated flow. This phenomenon may occur, for example, at the upstream-facing corner formed by a deflected control surface on a hypersonic entry vehicle, where the length of separation has implications for control effectiveness. Computational fluid-dynamics results are presented to show the flowfield around a blunted cone and cone-flare configurations in hypersonic flow with separation. This problem is of particular interest since it features most of the aspects of the hypersonic flow around planetary entry vehicles. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface pressure and heat flux with aeroheating. Indeed, flow separation is induced by the shock wave boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is a particularly delicate task for numerical codes. Laminar flow and turbulent computations have been carried out using a full Navier-Stokes solver, with freestream conditions provided by the experimental data obtained at Mach 6, 8, and 16.34 wind tunnel. The numerical results are compared with the measured pressure and surface heat flux distributions in the wind tunnel and a good agreement is found, especially on the length of the recirculation region and location of shock waves. The critical physics of entropy layer, boundary layers, boundary layers and shock wave interaction and flow behind shock are properly captured and elaborated.. Hypersonic flows are characterized by high Mach number and high total enthalpy. An elevated temperature often results in thermo-chemical reactions in the gas, which play a major role in aero thermodynamic characterization of high-speed aerospace vehicles. Computational simulation of such flows, therefore, needs to account for a range of physical phenomena. Further, the numerical challenges involved in resolving strong gradients and discontinuities add to the complexity of computational fluid dynamics (CFD) simulation. In this article, physical modeling and numerical methodology-related issues involved in hypersonic flow simulation are highlighted. State-of-the-art CFD challenges are discussed in the context of many prominent applications of hypersonic flows. In the first part of paper, hypersonic flow is simulated and aerodynamics characteristics are calculated. Then aero heating with chemical reactions are added in the simulations and in the end part heat transfer with turbulence modeling is simulated. Results are compared with available data.

Convective response of a wall-mounted hot-film sensor in a shock tube

NASA Technical Reports Server (NTRS)

Roberts, A. Sidney, Jr.; Ortgies, Kelly R.; Gartenberg, Ehud; Carraway, Debra L.

1991-01-01

Shock tube experiments were performed in order to determine the response of a single hot-film element of a sensor array to transiently induced flow behind weak normal shock waves. The experiments attempt to isolate the response due only to the change in convective heat transfer at the hot-film surface mounted on the wall of the shock tube. The experiments are described, the results being correlated with transient boundary layer theory and compared with an independent set of experimental results. One of the findings indicates that the change in the air properties (temperature and pressure) precedes the air mass transport, causing an ambiguity in the sensor response to the development of the velocity boundary layer. Also, a transient, local heat transfer coefficient is formulated to be used as a forcing function in an hot-film instrument model and simulation which remains under investigation.

Investigation of passive shock wave-boundary layer control for transonic airfoil drag reduction

NASA Technical Reports Server (NTRS)

Nagamatsu, H. T.; Brower, W. B., Jr.; Bahi, L.; Ross, J.

1982-01-01

The passive drag control concept, consisting of a porous surface with a cavity beneath it, was investigated with a 12-percent-thick circular arc and a 14-percent-thick supercritical airfoil mounted on the test section bottom wall. The porous surface was positioned in the shock wave/boundary layer interaction region. The flow circulating through the porous surface, from the downstream to the upstream of the terminating shock wave location, produced a lambda shock wave system and a pressure decrease in the downstream region minimizing the flow separation. The wake impact pressure data show an appreciably drag reduction with the porous surface at transonic speeds. To determine the optimum size of porosity and cavity, tunnel tests were conducted with different airfoil porosities, cavities and flow Mach numbers. A higher drag reduction was obtained by the 2.5 percent porosity and the 1/4-inch deep cavity.

Effect of shock interactions on mixing layer between co-flowing supersonic flows in a confined duct

NASA Astrophysics Data System (ADS)

Rao, S. M. V.; Asano, S.; Imani, I.; Saito, T.

2018-03-01

Experiments are conducted to observe the effect of shock interactions on a mixing layer generated between two supersonic streams of Mach number M _{1} = 1.76 and M _{2} = 1.36 in a confined duct. The development of this mixing layer within the duct is observed using high-speed schlieren and static pressure measurements. Two-dimensional, compressible Reynolds averaged Navier-Stokes equations are solved using the k-ω SST turbulence model in Fluent. Further, adverse pressure gradients are imposed by placing inserts of small (<7% of duct height) but finite (> boundary layer thickness) thickness on the walls of the test section. The unmatched pressures cause the mixing layer to bend and lead to the formation of shock structures that interact with the mixing layer. The mixing layer growth rate is found to increase after the shock interaction (nearly doubles). The strongest shock is observed when a wedge insert is placed in the M _{2} flow. This shock interacts with the mixing layer exciting flow modes that produce sinusoidal flapping structures which enhance the mixing layer growth rate to the maximum (by 1.75 times). Shock fluctuations are characterized, and it is observed that the maximum amplitude occurs when a wedge insert is placed in the M _{2} flow.

Shock induced Richtmyer-Meshkov instability in the presence of a wall boundary layer

NASA Astrophysics Data System (ADS)

Jourdan, G.; Billiotte, M.; Houas, L.

1996-06-01

An experimental investigation on gaseous mixing zones originated from the Richtmyer-Meshkov instability has been undertaken in a square cross section shock tube. Mass concentration fields, of one of the two mixing constituents, have been determined within the mixing zone when the shock wave passes from the heavy gas to the light one, from one gas to an other of close density, and from the light gas to the heavy one. Results have been obtained before and after the coming back of the reflected shock wave. The diagnostic method is based on the infrared absorption of one of the two constituents of the mixing zone. It is shown that the mixing zone is strongly deformed by the wall boundary layer. The consequence is the presence of strong gradients of concentration in the direction perpendicular to the shock wave propagation. Finally, it is pointed out that the mixing goes more homogeneous when the Atwood number tends to zero.

Current topics in shock waves; Proceedings of the International Symposium on Shock Waves and Shock Tubes, 17th, Lehigh University, Bethlehem, PA, July 17-21, 1989

NASA Astrophysics Data System (ADS)

Kim, Yong W.

Various papers on shock waves are presented. The general topics addressed include: shock formation, focusing, and implosion; shock reflection and diffraction; turbulence; laser-produced plasmas and waves; ionization and shock-plasma interaction; chemical kinetics, pyrolysis, and soot formation; experimental facilities, techniques, and applications; ignition of detonation and combustion; particle entrainment and shock propagation through particle suspension; boundary layers and blast simulation; computational methods and numerical simulation.

Pyrolysis Gas Flow in Thermally Ablating Media Using Time-Implicit Discontinuous Galerkin Methods

DTIC Science & Technology

2011-01-01

Aeronautics and Astronautics 2 the dissociated and ionized gas species (present in the shock layer, which is between the bow shock and boundary layer... wind tunnel experiment was conducted in [20] with a carbon-phenolic sample that was exposed to a heat flux of 1400 W/cm 2 . Experiment results were...type of problems [7-10]. In work by Persson and Peraire, they have been applied to various problems of viscous flows, shocks , turbulent flows and

Numerical studies of porous airfoils in transonic flow. Ph.D. Thesis. Final Report, 1 Jun. 1985 - 31 Aug. 1986

NASA Technical Reports Server (NTRS)

Chow, C. Y.

1986-01-01

A numerical tool is constructed to examine the effects of a porous surface on transonic airfoil performance and to help understand the flow structure of passive shockwave/boundary layer interactions. The porous region is located near the shock with a cavity underneath it. This study is composed of two parts. Solved in the first part, with an inviscid-flow approach, is the transonic full-potential equation associated with transpiration boundary conditions which are obtained from porosity modeling. The numerical results indicate that a porous airfoil has a wave drag lower than that of a solid airfoil. The observed lambda-shock structure in the wind-tunnel testing can be predicted. Furthermore, the lift could be increased with an appropriate porosity distribution. In the second part of this work, the modified version of either an interactive boundary layer (IBL) algorithm or a thin-layer Navier-Stokes (TLNS) algorithm is used to study the outer flow, while a stream-function formulation is used to model the inner flow in the shallow cavity. The coupling procedure at the porous surface is based on Darcy's law and the assumption of a constant total pressure in the cavity. In addition, a modified Baldwin-Lomax turbulence model is used to describe the transpired turbulent boundary layer in the TLNS approach, while the Cebeci turbulence model is used in the IBL approach. According to the present analysis, a porous surface can reduce the wave drag appreciably, but can also increase the viscous losses. As has been observed experimentally, the numerical results indicate that the total drag is reduced at higher Mach numbers and increased at lower Mach numbers when the angles of attack are small. Furthermore, the streamline pattern of passive shock/boundary layer interaction are revealed.

Relaxation of the accelerating-gas boundary layer to the test-gas boundary layer on a flat plate in an expansion tube

NASA Technical Reports Server (NTRS)

Gupta, R. N.; Trimpi, R. L.

1973-01-01

An analytic investigation of the relaxation of the accelerating-gas boundary layer to the test-gas boundary layer over a flat plate mounted in an expansion tube has been conducted. In this treatment, nitrogen has been considered as the test gas and helium as the accelerating gas. The problem is analyzed in two conically similar limits: (1) when the time lag between the arrival of the shock and the interface at the leading edge of the plate is very large, and (2) when this time lag is negligible. The transient laminar boundary-layer equations of a perfect binary-gas mixture are taken as the flow governing equations. These coupled equations have been solved numerically by Gauss-Seidel line-relaxation method. The results predict the transient behavior as well as the time required for an all-helium accelerating-gas boundary layer to relax to an all-nitrogen boundary layer.

Progress in hypersonic turbulence modeling

NASA Technical Reports Server (NTRS)

Wilcox, David C.

1991-01-01

A compressibility modification is developed for k-omega (Wilcox, 1988) and k-epsilon (Jones and Launder, 1972) models, that is similar to those of Sarkar et al. (1989) and Zeman (1990). Results of the perturbation solution for the compressible wall layer demonstrate why the Sarkar and Zeman terms yield inaccurate skin friction for the flat-plate boundary layer. A new compressibility term is developed which permits accurate predictions of the compressible mixing layer, flat-plate boundary layer, and shock separated flows.

Laser skin friction measurements and CFD comparison of weak-to-strong swept shock/boundary-layer interactions

NASA Technical Reports Server (NTRS)

Kim, K.-S.; Lee, Y.; Alvi, F. S.; Settles, G. S.; Horstman, C. C.

1990-01-01

A joint experimental and computational study of skin friction in weak-to-strong swept shock wave/turbulent boundary-layer interactions has been carried out. A planar shock wave is generated by a sharp fin at angles of attack alpha = 10 deg and 16 deg at M(infinity) = 3 and 16 and 20 deg at M(infinity) = 4. Measurements are made using the Laser Interferometer Skin Friction meter, which optically detects the rate of thinning of an oil film applied to the test surface. The results show a systematic rise in the peak c(f) at the rear part of the interaction, where the separated flow atttaches. For the stronget case studied, this peak is an order of magnitude higher than the incoming freestream c(f)level.

Experimental studies of shock-wave/wall-jet interaction in hypersonic flow

NASA Technical Reports Server (NTRS)

Holden, Michael S.; Rodriguez, Kathleen

1994-01-01

Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not significantly reduced by even the largest levels of film cooling. For the case studies in the absence of film cooling, the interaction regions were unseparated. However, adding film cooling resulted in regions of boundary layer separation induced in the film cooling layer -- the size of which regions first increased and then decreased with increased film cooling. Surprisingly, the size of the separated regions and the magnitude of the recompression heating were not strongly influenced by the thickness of the cooling film, nor by the point of shock impingement relative to the exit plane of the nozzles. The lip thickness was found to have little effect on cooling effectiveness. Measurements with and in the absence of shock interaction were compared with the results of earlier experimental studies and correlated in terms of the major parameters controlling these flows.

Experimental studies of shock-wave/wall-jet interaction in hypersonic flow, part A

NASA Technical Reports Server (NTRS)

Holden, Michael S.; Rodriguez, Kathleen

1994-01-01

Experimental studies have been conducted to examine slot film cooling effectiveness and the interaction between the cooling film and an incident planar shock wave in turbulent hypersonic flow. The experimental studies were conducted in the 48-inch shock tunnel at Calspan at a freestream Mach number of close to 6.4 and at a Reynolds number of 35 x 10(exp 6) based on the length of the model at the injection point. The Mach 2.3 planar wall jet was generated from 40 transverse nozzles (with heights of both 0.080 inch and 0.120 inch), producing a film that extended the full width of the model. The nozzles were operated at pressures and velocities close to matching the freestream, as well as at conditions where the nozzle flows were over- and under-expanded. A two-dimensional shock generator was used to generate oblique shocks that deflected the flow through total turnings of 11, 16, and 21 degrees; the flows impinged downstream of the nozzle exits. Detailed measurements of heat transfer and pressure were made both ahead and downstream of the injection station, with the greatest concentration of measurements in the regions of shock-wave/boundary layer interaction. The major objectives of these experimental studies were to explore the effectiveness of film cooling in the presence of regions of shock-wave/boundary layer interaction and, more specifically, to determine how boundary layer separation and the large recompression heating rates were modified by film cooling. Detailed distributions of heat transfer and pressure were obtained in the incident-shock/wall-jet interaction region for a series of shock strengths and impingement positions for each of the two nozzle heights. Measurements were also made to examine the effects of nozzle lip thickness on cooling effectiveness. The major conclusion from these studies was that the effect of the cooling film could be readily dispersed by relatively weak incident shocks, so the peak heating in the recompression region was not significantly reduced by even the largest levels of film cooling. For the case studies in the absence of film cooling, the interaction regions were unseparated. However, adding film cooling resulted in regions of boundary layer separation induced in the film cooling layer, the size of which regions first increased and then decreased with increased film cooling. Surprisingly, the size of the separated regions and the magnitude of the recompression heating were not strongly influenced by the thickness of the cooling film, nor by the point of shock impingement relative to the exit plane of the nozzles. The lip thickness was found to have little effect on cooling effectiveness. Measurements with and in the absence of shock interaction were compared with the results of earlier experimental studies and correlated in terms of the major parameters controlling these flows.

Discussion of flight experiments with an entry research vehicle

NASA Technical Reports Server (NTRS)

Potter, J. L.

1985-01-01

The focus of interest is the maneuvering flight of advanced entry vehicles operating at altitudes above 50 km and at velocities of 5 to 8 km/s. Information resulting in more accurate aerodynamic analysis is sought and measurement techniques that appear to be applicable are identified. Measurements discussed include: shock layer or boundary layer profiles of velocity, temperature, species mass fractions, and other gas properties associated with aerodynamic heating; surface energy transfer process; nonequilibrium flow processes and pressure distribution; separated, vortic leeside flow of nonequilibrium fluid; boundary layer transition on highly swept configurations; and shock and surface slip and gas/surface interaction. Further study should focus on evolving measurement techniques, installation requirements, and on identification of the portions of flights where successful results seem probable.

Instability of a Supersonic Boundary-Layer with Localized Roughness

NASA Technical Reports Server (NTRS)

Marxen, Olaf; Iaccarino, Gianluca; Shaqfeh, Eric S. G.

2010-01-01

A localized 3-D roughness causes boundary-layer separation and (weak) shocks. Most importantly, streamwise vortices occur which induce streamwise (low U, high T) streaks. Immersed boundary method (volume force) suitable to represent roughness element in DNS. Favorable comparison between bi-global stability theory and DNS for a "y-mode" Outlook: Understand the flow physics (investigate "z-modes" in DNS through sinuous spanwise forcing, study origin of the beat in DNS).

Aeroheating Measurement of Apollo Shaped Capsule with Boundary Layer Trip in the Free-piston Shock Tunnel HIEST

NASA Technical Reports Server (NTRS)

Hideyuki, TANNO; Tomoyuki, KOMURO; Kazuo, SATO; Katsuhiro, ITOH; Lillard, Randolph P.; Olejniczak, Joseph

2013-01-01

An aeroheating measurement test campaign of an Apollo capsule model with laminar and turbulent boundary layer was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. A 250mm-diameter 6.4%-scaled Apollo CM capsule model made of SUS-304 stainless steel was applied in this study. To measure heat flux distribution, the model was equipped with 88 miniature co-axial Chromel-Constantan thermocouples on the heat shield surface of the model. In order to promote boundary layer transition, a boundary layer trip insert with 13 "pizza-box" isolated roughness elements, which have 1.27mm square, were placed at 17mm below of the model geometric center. Three boundary layer trip inserts with roughness height of k=0.3mm, 0.6mm and 0.8mm were used to identify the appropriate height to induce transition. Heat flux records with or without roughness elements were obtained for model angles of attack 28º under stagnation enthalpy between H(sub 0)=3.5MJ/kg to 21MJ/kg and stagnation pressure between P(sub 0)=14MPa to 60MPa. Under the condition above, Reynolds number based on the model diameter was varied from 0.2 to 1.3 million. With roughness elements, boundary layer became fully turbulent less than H(sub 0)=9MJ/kg condition. However, boundary layer was still laminar over H(sub 0)=13MJ/kg condition even with the highest roughness elements. An additional experiment was also performed to correct unexpected heat flux augmentation observed over H(sub 0)=9MJ/kg condition.

Correlation parameters for the study of leeside heating on a lifting body at hypersonic speeds

NASA Technical Reports Server (NTRS)

Vidal, R. J.

1974-01-01

Leeside heating was studied with the aim of gaining some insight into: (1) the magnitude of the leeside heating rates and (2) the methods to be used to extrapolate wind tunnel leeside heating rates to the full scale flight condition. This study was based on existing experimental data obtained in a hypersonic shock tunnel on lifting body configurations that are typical of shuttle orbiter vehicles. Heat transfer was first measured on the windward side to determine the boundary layer type. Then the leeside heating was investigated with the classified boundary layer. Correlation data are given on the windward turbulent boundary layer, the windward laminar boundary layer, and the leeside surfaces.

Interactive boundary-layer calculations of a transonic wing flow

NASA Technical Reports Server (NTRS)

Kaups, Kalle; Cebeci, Tuncer; Mehta, Unmeel

1989-01-01

Results obtained from iterative solutions of inviscid and boundary-layer equations are presented and compared with experimental values. The calculated results were obtained with an Euler code and a transonic potential code in order to furnish solutions for the inviscid flow; they were interacted with solutions of two-dimensional boundary-layer equations having a strip-theory approximation. Euler code results are found to be in better agreement with the experimental data than with the full potential code, especially in the presence of shock waves, (with the sole exception of the near-tip region).

Numerical Study of Boundary-Layer in Aerodynamics

NASA Technical Reports Server (NTRS)

Shih, Tom I-P.

1997-01-01

The accomplishments made in the following three tasks are described: (1) The first task was to study shock-wave boundary-layer interactions with bleed - this study is relevant to boundary-layer control in external and mixed-compression inlets of supersonic aircraft; (2) The second task was to test RAAKE, a code developed for computing turbulence quantities; and (3) The third task was to compute flow around the Ames ER-2 aircraft that has been retrofitted with containers over its wings and fuselage. The appendices include two reports submitted to AIAA for publication.

Numerical study of heterogeneous mean temperature and shock wave in a resonator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Yano, Takeru

2015-10-28

When a frequency of gas oscillation in an acoustic resonator is sufficiently close to one of resonant frequencies of the resonator, the amplitude of gas oscillation becomes large and hence the nonlinear effect manifests itself. Then, if the dissipation effects due to viscosity and thermal conductivity of the gas are sufficiently small, the gas oscillation may evolve into the acoustic shock wave, in the so-called consonant resonators. At the shock front, the kinetic energy of gas oscillation is converted into heat by the dissipation process inside the shock layer, and therefore the temperature of the gas in the resonator rises.more » Since the acoustic shock wave travels in the resonator repeatedly over and over again, the temperature rise becomes noticeable in due course of time even if the shock wave is weak. We numerically study the gas oscillation with shock wave in a resonator of square cross section by solving the initial and boundary value problem of the system of three-dimensional Navier-Stokes equations with a finite difference method. In this case, the heat conduction across the boundary layer on the wall of resonator causes a spatially heterogeneous distribution of mean (time-averaged) gas temperature.« less

A numerical study of shock wave reflections on low density foam

NASA Astrophysics Data System (ADS)

Baer, M. R.

1992-06-01

A continuum mixture theory is used to describe shock wave reflections on low density open-cell polyurethane foam. Numerical simulations are compared to the shock tube experiments of Skews (1991) and detailed wave fields are shown of a shock wave interacting with a layer of foam adjacent to a rigid wall boundary. These comparisons demonstrate that a continuum mixture theory describes well the shock interactions with low density foam.

Calibration of Axisymmetric and Quasi-1D Solvers for High Enthalpy Nozzles

NASA Technical Reports Server (NTRS)

Papadopoulos, P. E.; Gochberg, L. A.; Tokarcik-Polsky, S.; Venkatapathy, E.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1994-01-01

The proposed paper will present a numerical investigation of the flow characteristics and boundary layer development in the nozzles of high enthalpy shock tunnel facilities used for hypersonic propulsion testing. The computed flow will be validated against existing experimental data. Pitot pressure data obtained at the entrance of the test cabin will be used to validate the numerical simulations. It is necessary to accurately model the facility nozzles in order to characterize the test article flow conditions. Initially the axisymmetric nozzle flow will be computed using a Navier Stokes solver for a range of reservoir conditions. The calculated solutions will be compared and calibrated against available experimental data from the DLR HEG piston-driven shock tunnel and the 16-inch shock tunnel at NASA Ames Research Center. The Reynolds number is assumed to be high enough at the throat that the boundary layer flow is assumed turbulent at this point downstream. The real gas affects will be examined. In high Mach number facilities the boundary layer is thick. Attempts will be made to correlate the boundary layer displacement thickness. The displacement thickness correlation will be used to calibrate the quasi-1D codes NENZF and LSENS in order to provide fast and efficient tools of characterizing the facility nozzles. The calibrated quasi-1D codes will be implemented to study the effects of chemistry and the flow condition variations at the test section due to small variations in the driver gas conditions.

The effects of micro-vortex generators on normal shock wave/boundary layer interactions

NASA Astrophysics Data System (ADS)

Herges, Thomas G.

Shock wave/boundary-layer interactions (SWBLIs) are complex flow phenomena that are important in the design and performance of internal supersonic and transonic flow fields such as engine inlets. This investigation was undertaken to study the effects of passive flow control devices on normal shock wave/boundary layer interactions in an effort to gain insight into the physics that govern these complex interactions. The work concentrates on analyzing the effects of vortex generators (VGs) as a flow control method by contributing a greater understanding of the flowfield generated by these devices and characterizing their effects on the SWBLI. The vortex generators are utilized with the goal of improving boundary layer health (i.e., reducing/increasing the boundary-layer incompressible shape factor/skin friction coefficient) through a SWBLI, increasing pressure recovery, and reducing flow distortion at the aerodynamic interface plane while adding minimal drag to the system. The investigation encompasses experiments in both small-scale and large-scale inlet testing, allowing multiple test beds for improving the characterization and understanding of vortex generators. Small-scale facility experiments implemented instantaneous schlieren photography, surface oil-flow visualization, pressure-sensitive paint, and particle image velocimetry to characterize the effects of an array of microramps on a normal shock wave/boundary-layer interaction. These diagnostics measured the time-averaged and instantaneous flow organization in the vicinity of the microramps and SWBLI. The results reveal that a microramp produces a complex vortex structure in its wake with two primary counter-rotating vortices surrounded by a train of Kelvin- Helmholtz (K-H) vortices. A streamwise velocity deficit is observed in the region of the primary vortices in addition to an induced upwash/downwash which persists through the normal shock with reduced strength. The microramp flow control also increased the spanwise-averaged skin-friction coefficient and reduced the spanwise-averaged incompressible shape factor, thereby improving the health of the boundary layer. The velocity in the near-wall region appears to be the best indicator of microramp effectiveness at controlling SWBLIs. Continued analysis of additional micro-vortex generator designs in the small-scale facility revealed reduced separation within a subsonic diffuser downstream of the normal shock wave/boundary layer interaction. The resulting attached flow within the diffuser from the micro-vortex generator control devices reduces shock wave position and pressure RMS fluctuations within the diffuser along with increased pressure recovery through the shock and at the entrance of the diffuser. The largest effect was observed by the micro-vortex generators that produce the strongest streamwise vortices. High-speed pressure measurements also indicated that the vortex generators shift the energy of the pressure fluctuations to higher frequencies. Implementation of micro-vortex generators into a large-scale, supersonic, axisymmetric, relaxed-compression inlet have been investigated with the use of a unique and novel flow-visualization measurement system designed and successfully used for the analysis of both upstream micro-VGs (MVGs) and downstream VGs utilizing surface oil-flow visualization and pressure-sensitive paint measurements. The inlet centerbody and downstream diffuser vortex-generator regions were imaged during wind-tunnel testing internally through the inlet cowl with the diagnostic system attached to the cowl. Surface-flow visualization revealed separated regions along the inlet centerbody for large mass-flow rates without vortex generators. Upstream vortex generators did reduce separation in the subsonic diffuser, and a unique perspective of the flowfield produced by the downstream vortex generators was obtained. In addition, pressure distributions on the inlet centerbody and vortex generators were measured with pressure-sensitive paint. At low mass-flow ratios the onset of buzz occurs in the large-scale low-boom inlet. Inlet buzz and how it is affected by vortex generators was characterized using shock tracking through high-speed schlieren imaging and pressure fluctuation measurements. The analysis revealed a dominant low frequency oscillation at 21.0 Hz for the single-stream inlet, corresponding with the duration of one buzz cycle. Pressure oscillations prior to the onset of buzz were not detected, leaving the location where the shock wave triggers large separation on the compression spike as the best indicator for the onset of buzz. The driving mechanism for a buzz cycle has been confirmed as the rate of depressurization and repressurization of the inlet as the buzz cycle fluctuates between an effectively unstarted (blocked) inlet and supercritical operation (choked flow), respectively. High-frequency shock position oscillations/pulsations (spike buzz) were also observed throughout portions of the inlet buzz cycle. The primary effect of the VGs was to trigger buzz at a higher mass-flow ratio.

Supersonic and hypersonic shock/boundary-layer interaction database

NASA Technical Reports Server (NTRS)

Settles, Gary S.; Dodson, Lori J.

1994-01-01

An assessment is given of existing shock wave/tubulent boundary-layer interaction experiments having sufficient quality to guide turbulence modeling and code validation efforts. Although the focus of this work is hypersonic, experiments at Mach numbers as low as 3 were considered. The principal means of identifying candidate studies was a computerized search of the AIAA Aerospace Database. Several hundred candidate studies were examined and over 100 of these were subjected to a rigorous set of acceptance criteria for inclusion in the data-base. Nineteen experiments were found to meet these criteria, of which only seven were in the hypersonic regime (M is greater than 5).

Skin-Friction Measurements in a 3-D, Supersonic Shock-Wave/Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Wideman, J. K.; Brown, J. L.; Miles, J. B.; Ozcan, O.

1994-01-01

The experimental documentation of a three-dimensional shock-wave/boundary-layer interaction in a nominal Mach 3 cylinder, aligned with the free-stream flow, and 20 deg. half-angle conical flare offset 1.27 cm from the cylinder centerline. Surface oil flow, laser light sheet illumination, and schlieren were used to document the flow topology. The data includes surface-pressure and skin-friction measurements. A laser interferometric skin friction data. Included in the skin-friction data are measurements within separated regions and three-dimensional measurements in highly-swept regions. The skin-friction data will be particularly valuable in turbulence modeling and computational fluid dynamics validation.

Effects of Nose Bluntness on Stability of Hypersonic Boundary Layers over Blunt Cone

NASA Technical Reports Server (NTRS)

Kara, K.; Balakumar, P.; Kandil, O. A.

2007-01-01

Receptivity and stability of hypersonic boundary layers are numerically investigated for boundary layer flows over a 5-degree straight cone at a free-stream Mach number of 6.0. To compute the shock and the interaction of shock with the instability waves, we solve the Navier-Stokes equations in axisymmetric coordinates. The governing equations are solved using the 5th-order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. After the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. Generation of instability waves from leading edge region and receptivity of boundary layer to slow acoustic waves are investigated. Computations are performed for a cone with nose radii of 0.001, 0.05 and 0.10 inches that give Reynolds numbers based on the nose radii ranging from 650 to 130,000. The linear stability results showed that the bluntness has a strong stabilizing effect on the stability of axisymmetric boundary layers. The transition Reynolds number for a cone with the nose Reynolds number of 65,000 is increased by a factor of 1.82 compared to that for a sharp cone. The receptivity coefficient for a sharp cone is about 4.23 and it is very small, approx.10(exp -3), for large bluntness.

Receptivity of Hypersonic Boundary Layers Due to Acoustic Disturbances over Blunt Cone

NASA Technical Reports Server (NTRS)

Kara, K.; Balakumar, P.; Kandil, O. A.

2007-01-01

The transition process induced by the interaction of acoustic disturbances in the free-stream with boundary layers over a 5-degree straight cone and a wedge with blunt tips is numerically investigated at a free-stream Mach number of 6.0. To compute the shock and the interaction of shock with the instability waves the Navier-Stokes equations are solved in axisymmetric coordinates. The governing equations are solved using the 5th -order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. After the mean flow field is computed, acoustic disturbances are introduced at the outer boundary of the computational domain and unsteady simulations are performed. Generation and evolution of instability waves and the receptivity of boundary layer to slow and fast acoustic waves are investigated. The mean flow data are compared with the experimental results. The results show that the instability waves are generated near the leading edge and the non-parallel effects are stronger near the nose region for the flow over the cone than that over a wedge. It is also found that the boundary layer is much more receptive to slow acoustic wave (by almost a factor of 67) as compared to the fast wave.

Stability and modal analysis of shock/boundary layer interactions

NASA Astrophysics Data System (ADS)

Nichols, Joseph W.; Larsson, Johan; Bernardini, Matteo; Pirozzoli, Sergio

2017-02-01

The dynamics of oblique shock wave/turbulent boundary layer interactions is analyzed by mining a large-eddy simulation (LES) database for various strengths of the incoming shock. The flow dynamics is first analyzed by means of dynamic mode decomposition (DMD), which highlights the simultaneous occurrence of two types of flow modes, namely a low-frequency type associated with breathing motion of the separation bubble, accompanied by flapping motion of the reflected shock, and a high-frequency type associated with the propagation of instability waves past the interaction zone. Global linear stability analysis performed on the mean LES flow fields yields a single unstable zero-frequency mode, plus a variety of marginally stable low-frequency modes whose stability margin decreases with the strength of the interaction. The least stable linear modes are grouped into two classes, one of which bears striking resemblance to the breathing mode recovered from DMD and another class associated with revolving motion within the separation bubble. The results of the modal and linear stability analysis support the notion that low-frequency dynamics is intrinsic to the interaction zone, but some continuous forcing from the upstream boundary layer may be required to keep the system near a limit cycle. This can be modeled as a weakly damped oscillator with forcing, as in the early empirical model by Plotkin (AIAA J 13:1036-1040, 1975).

Theoretical Calculation of Viscous-Inviscid Transonic Flows.

DTIC Science & Technology

1980-08-01

Taylor Naval Ship Research and Development Center Aviation and Surface Effects Department / (See reverse side) Bethesda, Maryland 20084 ! CONTROLLING...Interactions ... .......... ... 46 18 -ffect of Boundary Layer on Blade Surface Pressures in a Transonic Fan Rotor Tip Section Cascade...complicated by the viscous effect . The strong viscous-inviscid interaction caused by the shock wave thickens the boundary layer rapidly, and the flow eventually

The dynamics and control of fluctuating pressure loads in the reattachment region of a supersonic free shear layer

NASA Technical Reports Server (NTRS)

Smits, A. J.

1990-01-01

The primary aim is to investigate the mechanisms which cause the unsteady wall-pressure fluctuations in shock wave turbulent shear layer interactions. The secondary aim is to find means to reduce the magnitude of the fluctuating pressure loads by controlling the unsteady shock motion. The particular flow proposed for study is the unsteady shock wave interaction formed in the reattachment zone of a separated supersonic flow. Similar flows are encountered in many practical situations, and they are associated with high levels of fluctuating wall pressure. Wall pressure fluctuations were measured in the reattachment region of the supersonic free shear layer. The free shear layer was formed by the separation of a Mach 2.9 turbulent boundary layer from a backward facing step. Reattachment occurred on a 20 deg ramp. By adjusting the position of the ramp, the base pressure was set equal to the freestream pressure, and the free shear layer formed in the absence of a separation shock. An array of flush-mounted, miniature, high-frequency pressure transducers was used to make multichannel measurements of the fluctuating wall pressure in the vicinity of the reattachment region. Contrary to previous observations of this flow, the reattachment region was found to be highly unsteady, and the pressure fluctuations were found to be significant. The overall behavior of the wall pressure loading is similar in scale and magnitude to the unsteadiness of the wall pressure field in compression ramp flows at the same Mach number. Rayleigh scattering was used to visualize the instantaneous shock structure in the streamwise and spanwise direction. Spanwise wrinkles on the order of half the boundary layer thickness were observed.

Investigation of Materials for Boundary Layer Control in a Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Braafladt, Alexander; Lucero, John M.; Hirt, Stefanie M.

2013-01-01

During operation of the NASA Glenn Research Center 15- by 15-Centimeter Supersonic Wind Tunnel (SWT), a significant, undesirable corner flow separation is created by the three-dimensional interaction of the wall and floor boundary layers in the tunnel corners following an oblique-shock/ boundary-layer interaction. A method to minimize this effect was conceived by connecting the wall and floor boundary layers with a radius of curvature in the corners. The results and observations of a trade study to determine the effectiveness of candidate materials for creating the radius of curvature in the SWT are presented. The experiments in the study focus on the formation of corner fillets of four different radii of curvature, 6.35 mm (0.25 in.), 9.525 mm (0.375 in.), 12.7 mm (0.5 in.), and 15.875 mm (0.625 in.), based on the observed boundary layer thickness of 11.43 mm (0.45 in.). Tests were performed on ten candidate materials to determine shrinkage, surface roughness, cure time, ease of application and removal, adhesion, eccentricity, formability, and repeatability. Of the ten materials, the four materials which exhibited characteristics most promising for effective use were the heavy body and regular type dental impression materials, the basic sculpting epoxy, and the polyurethane sealant. Of these, the particular material which was most effective, the heavy body dental impression material, was tested in the SWT in Mach 2 flow, and was observed to satisfy all requirements for use in creating the corner fillets in the upcoming experiments on shock-wave/boundary-layer interaction.

A normal shock-wave turbulent boundary-layer interaction at transonic speeds

NASA Technical Reports Server (NTRS)

Mateer, G. G.; Brosh, A.; Viegas, J. R.

1976-01-01

Experimental results, including surveys of the mean and fluctuating flow, and measurements of surface pressure, skin friction, and separation length, are compared with solutions to the Navier-Stokes equations utilizing various algebraic eddy viscosity models to describe the Reynolds shear stresses. The experimental data, obtained at a free-stream Mach number of 1.5 and Reynolds numbers between 10 million and 80 million, show that a separated zone forms near the foot of the shock and that its length is proportional to the initial boundary-layer thickness; that a supersonic region forms downstream of the shock; and that the shear stress increases significantly through the interaction and subsequently decays downstream. The computations adequately represent the qualitative features of the flow field throughout the interaction but quantitatively underpredict the extent of separation and the downstream level of skin friction.

A Combined CFD/Characteristic Method for Prediction and Design of Hypersonic Inlet with Nose Bluntness

NASA Astrophysics Data System (ADS)

Gao, Wenzhi; Li, Zhufei; Yang, Jiming

Leading edge bluntness is widely used in hypersonic inlet design for thermal protection[1]. Detailed research of leading edge bluntness on hypersonic inlet has been concentrated on shock shape correlation[2], boundary layer flow[3], inlet performance[4], etc. It is well known that blunted noses cause detached bow shocks which generate subsonic regions around the noses and entropy layers in the flowfield.

The efficient simulation of separated three-dimensional viscous flows using the boundary-layer equations

NASA Technical Reports Server (NTRS)

Van Dalsem, W. R.; Steger, J. L.

1985-01-01

A simple and computationally efficient algorithm for solving the unsteady three-dimensional boundary-layer equations in the time-accurate or relaxation mode is presented. Results of the new algorithm are shown to be in quantitative agreement with detailed experimental data for flow over a swept infinite wing. The separated flow over a 6:1 ellipsoid at angle of attack, and the transonic flow over a finite-wing with shock-induced 'mushroom' separation are also computed and compared with available experimental data. It is concluded that complex, separated, three-dimensional viscous layers can be economically and routinely computed using a time-relaxation boundary-layer algorithm.

Experimental investigation of the micro-ramp based shock wave and turbulent boundary layer interaction control

NASA Astrophysics Data System (ADS)

Bo, Wang; Weidong, Liu; Yuxin, Zhao; Xiaoqiang, Fan; Chao, Wang

2012-05-01

Using a nanoparticle-based planar laser-scattering technique and supersonic particle image velocimetry, we investigated the effects of micro-ramp control on incident shockwave and boundary-layer interaction (SWBLI) in a low-noise supersonic wind-tunnel with Mach number 2.7 and Reynolds number Rθ = 5845. High spatiotemporal resolution wake structures downstream of the micro-ramps were detected, while a complex evolution process containing a streamwise counter-rotating vortex pair and large-scale hairpin-like vortices with Strouhal number Stδ of about 0.5-0.65 was revealed. The large-scale structures could survive while passing through the SWBLI region. Reflected shockwaves are clearly seen to be distorted accompanied by high-frequency fluctuations. Micro-ramp applications have a distinct influence on flow patterns of the SWBLI field that vary depending on spanwise locations. Both the shock foot and separation line exhibit undulations corresponding with modifications of the velocity distribution of the incoming boundary layer. Moreover, by energizing parts of the boundary flow, the micro-ramp is able to dampen the separation.

Mineralogic evidence for an impact event at the cretaceous-tertiary boundary

USGS Publications Warehouse

Bohor, B.F.; Foord, E.E.; Modreski, P.J.; Triplehorn, Don M.

1984-01-01

A thin claystone layer found in nonmarine rocks at the palynological Cretaceous-Tertiary boundary in eastern Montana contains an anomalously high value of iridium. The nonclay fraction is mostly quartz with minor feldspar, and some of these grains display planar features. These planar features are related to specific crystallographic directions in the quartz lattice. The shocked quartz grains also exhibit asterism and have lowered refractive indices. All these mineralogical features are characteristic of shock metamorphism and are compelling evidence that the shocked grains are the product of a high velocity impact between a large extraterrestrial body and the earth. The shocked minerals represent silicic target material injected into the stratosphere by the impact of the projectile.

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

NASA Technical Reports Server (NTRS)

Moss, James N.; Bird, Graeme A.

2004-01-01

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.

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

NASA Technical Reports Server (NTRS)

Moss, James N.; Bird, Graeme A.

2004-01-01

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.

A review of the use of vortex generators for mitigating shock-induced separation

NASA Astrophysics Data System (ADS)

Titchener, Neil; Babinsky, Holger

2015-09-01

This article reviews research into the potential of vortex generators to mitigate shock-induced separation. Studies ranging from those conducted in the early post-war era to those performed recently are discussed. On the basis of the investigations described in this report, it is clear that vortex generators can alleviate shock-induced boundary layer separation. Yet, it will be shown that their potential and efficiency varies considerably in practical applications. Much more success is reported in transonic test cases compared to separation induced in purely supersonic interactions. Under a variety of flow conditions, the best performance is achieved with vortex generators with a height of roughly half the boundary layer thickness and a shape similar to a swept vane. Notwithstanding this, vortex generator performance is not as consistent as it is in low-speed applications. Further work is required before vortex generators can be implemented into the design process for eliminating shock-induced separation on transonic wings and in supersonic inlets.

Some observations on boundary conditions for numerical conservation laws

NASA Technical Reports Server (NTRS)

Kamowitz, David

1988-01-01

Four choices of outflow boundary conditions are considered for numerical conservation laws. All four methods are stable for linear problems, for which examples are presented where either a boundary layer forms or the numerical scheme, together with the boundary condition, is unstable due to the formation of a reflected shock. A simple heuristic argument is presented for determining the suitability of the boundary condition.

The shocking properties of supersonic flows: Dependence of the thermal overstability on M, α, and Tc / T0

NASA Astrophysics Data System (ADS)

Pittard, J. M.; Dobson, M. S.; Durisen, R. H.; Dyson, J. E.; Hartquist, T. W.; O'Brien, J. T.

2005-07-01

We present hydrodynamical calculations of radiative shocks with low Mach numbers and find that the well-known global overstability can occur if the temperature exponent (α) of the cooling is sufficiently negative. We find that the stability of radiative shocks increases with decreasing Mach number, with the result that M=2 shocks require α ⪉ -1.2 in order to be overstable. Such values occur within a limited temperature range of many cooling curves. We observe that Mach numbers of order 100 are needed before the strong shock limit of α_cr ≈ 0.4 is reached, and we discover that the frequency of oscillation of the fundamental mode also has a strong Mach number dependence. We find that feedback between the cooling region and the cold dense layer (CDL) further downstream is a function of Mach number, with stronger feedback and oscillation of the boundary between the CDL and the cooling region occuring at lower Mach numbers. This feedback can be quantified in terms of the reflection coefficient of sound waves, and in those cases where the cooling layer completely disappears at the end of each oscillation cycle, the initial velocity of the waves driven into the upstream pre-shock flow and into the downstream CDL, and the velocity of the the boundary between the CDL and the cooling layer, can be understood in terms of the solution to the Riemann problem. An interesting finding is that the stability properties of low Mach number shocks can be dramatically altered if the shocked gas is able to cool to temperatures less than the pre-shock value (i.e. when χ < 1, where χ is the ratio of the temperature of the cold dense layer to the pre-shock temperature). In such circumstances, low Mach number shocks have values of α_cr which are comparable to values obtained for higher Mach number shocks when χ = 1. For instance, α_cr=-0.1 when M=2 and χ=0.1, comparable to that when M=10 and χ=1. Thus, it is probable that low Mach number astrophysical shocks will be overstable in a variety of situations. We also explore the effect of different assumptions for the initial hydrodynamic set up and the type of boundary condition imposed downstream, and find that the properties of low Mach number shocks are relatively insensitive to these issues. The results of this work are relevant to astrophysical shocks with low Mach numbers, such as supernova remnants (SNRs) immersed in a hot interstellar medium (e.g., within a starburst region), and shocks in molecular clouds, where time-dependent chemistry can lead to overstability.

Thin-layer approximation and algebraic model for separated turbulent flows

NASA Technical Reports Server (NTRS)

Baldwin, B.; Lomax, H.

1978-01-01

An algebraic turbulence model for two- and three-dimensional separated flows is specified that avoids the necessity for finding the edge of the boundary layer. Properties of the model are determined and comparisons made with experiment for an incident shock on a flat plate, separated flow over a compression corner, and transonic flow over an airfoil. Separation and reattachment points from numerical Navier-Stokes solutions agree with experiment within one boundary-layer thickness. Use of law-of-the-wall boundary conditions does not alter the predictions significantly. Applications of the model to other cases are contained in companion papers.

The lagRST Model: A Turbulence Model for Non-Equilibrium Flows

NASA Technical Reports Server (NTRS)

Lillard, Randolph P.; Oliver, A. Brandon; Olsen, Michael E.; Blaisdell, Gregory A.; Lyrintzis, Anastasios S.

2011-01-01

This study presents a new class of turbulence model designed for wall bounded, high Reynolds number flows with separation. The model addresses deficiencies seen in the modeling of nonequilibrium turbulent flows. These flows generally have variable adverse pressure gradients which cause the turbulent quantities to react at a finite rate to changes in the mean flow quantities. This "lag" in the response of the turbulent quantities can t be modeled by most standard turbulence models, which are designed to model equilibrium turbulent boundary layers. The model presented uses a standard 2-equation model as the baseline for turbulent equilibrium calculations, but adds transport equations to account directly for non-equilibrium effects in the Reynolds Stress Tensor (RST) that are seen in large pressure gradients involving shock waves and separation. Comparisons are made to several standard turbulence modeling validation cases, including an incompressible boundary layer (both neutral and adverse pressure gradients), an incompressible mixing layer and a transonic bump flow. In addition, a hypersonic Shock Wave Turbulent Boundary Layer Interaction with separation is assessed along with a transonic capsule flow. Results show a substantial improvement over the baseline models for transonic separated flows. The results are mixed for the SWTBLI flows assessed. Separation predictions are not as good as the baseline models, but the over prediction of the peak heat flux downstream of the reattachment shock that plagues many models is reduced.

A method for calculating aerodynamic heating on sounding rocket tangent ogive noses.

NASA Technical Reports Server (NTRS)

Wing, L. D.

1973-01-01

A method is presented for calculating the aerodynamic heating and shear stresses at the wall for tangent ogive noses that are slender enough to maintain an attached nose shock through that portion of flight during which heat transfer from the boundary layer to the wall is significant. The lower entropy of the attached nose shock combined with the inclusion of the streamwise pressure gradient yields a reasonable estimate of the actual flow conditions. Both laminar and turbulent boundary layers are examined and an approximation of the effects of (up to) moderate angles-of-attack is included in the analysis. The analytical method has been programmed in FORTRAN IV for an IBM 360/91 computer.

A method for calculating aerodynamic heating on sounding rocket tangent ogive noses

NASA Technical Reports Server (NTRS)

Wing, L. D.

1972-01-01

A method is presented for calculating the aerodynamic heating and shear stresses at the wall for tangent ogive noses that are slender enough to maintain an attached nose shock through that portion of flight during which heat transfer from the boundary layer to the wall is significant. The lower entropy of the attached nose shock combined with the inclusion of the streamwise pressure gradient yields a reasonable estimate of the actual flow conditions. Both laminar and turbulent boundary layers are examined and an approximation of the effects of (up to) moderate angles-of-attack is included in the analysis. The analytical method has been programmed in FORTRAN 4 for an IBM 360/91 computer.

Magnetic Fe, Si, Al-Rich Impact Spherules from the P-T Boundary Layer at Graphite Peak, Antarctica

NASA Technical Reports Server (NTRS)

Petaev, M. I.; Jacobsen, S. B.; Basu, A. R.; Becker, L.

2004-01-01

The geological boundary between Triassic and Permian strata coincides with the greatest life extinction in the Earth's history. Although the cause of the extinction is still the subject of intense debates, recent discoveries in the P-T boundary layer of shocked quartz grains, fullerenes with the extraterrestrial noble gases, Fe metal nuggets, and chondritic meteorite fragments all point to a powerful collision of Earth with a celestial body in the late Permian. Here we report the discovery of magnetic Fe, Si, Al-rich impact spherules which accompany the chondritic meteorite fragments in some samples from the P-T boundary layer at Graphite Peak, Antarctica.

Aero-optics overview. [laser applications

NASA Technical Reports Server (NTRS)

Gilbert, K. G.

1980-01-01

Various aero-optical phenomena are discussed with reference to their effect on airborne high energy lasers. Major emphasis is placed on: compressibility effects induced in the surrounding flow field; viscous effects which manifests themselves as aircraft boundary layers or shear layers; inviscid flow fields surrounding the aircraft due to airflow around protuberance such as laser turret assemblies; and shocks, established whenever local flow exceeds Mach one. The significant physical parameters affecting the interaction of a laser beam with a turbulent boundary layer are also described.

Effects of nonequilibrium ablation chemistry on Viking radio blackout.

NASA Technical Reports Server (NTRS)

Evans, J. S.; Schexnayder, C. J., Jr.; Grose, W. L.

1973-01-01

The length of the entry blackout period during descent of the Viking Lander into the Mars atmosphere is predicted from calculated profiles of electron density in the shock layer over the aeroshell. Nonequilibrium chemistry plays a key role in the calculation, both in the inviscid flow and in the boundary layer. This is especially true in the boundary layer contaminated with ablation material, for which nonequilibrium chemistry predicts electron densities two decades lower than the same case calculated with equilibrium chemistry.

Dynamics of the outgoing turbulent boundary layer in a Mach 5 unswept compression ramp interaction

NASA Technical Reports Server (NTRS)

Gramann, Richard A.; Dolling, David S.

1990-01-01

Wall pressure fluctuations have been measured under the unsteady separation shock and on the ramp face in an unswept Mach 5 compression ramp interaction. The freestream Reynolds number was 51.0 x 10 to the 6th/m, and the incoming turbulent boundary layer developed on the tunnel floor under approximately adiabatic wall temperature conditions. Standard data-acquisition methods, as well as real-time and posttest conditional sampling techniques were used. The results show that the mean and rms pressure levels are strong functions of separation shock position. At all stations on the ramp, from the corner to where the pressure reaches the theoretical inviscid value, the pressure signals have two dominant components: a low frequency component characteristic of the global unsteadiness, which correlates with the separation shock motion, and a higher frequency component associated with turbulence. The former is the major contributor to the overall signal variance.

Study of shock-induced combustion using an implicit TVD scheme

NASA Technical Reports Server (NTRS)

Yungster, Shayne

1992-01-01

The supersonic combustion flowfields associated with various hypersonic propulsion systems, such as the ram accelerator, the oblique detonation wave engine, and the scramjet, are being investigated using a new computational fluid dynamics (CFD) code. The code solves the fully coupled Reynolds-averaged Navier-Stokes equations and species continuity equations in an efficient manner. It employs an iterative method and a second order differencing scheme to improve computational efficiency. The code is currently being applied to study shock wave/boundary layer interactions in premixed combustible gases, and to investigate the ram accelerator concept. Results obtained for a ram accelerator configuration indicate a new combustion mechanism in which a shock wave induces combustion in the boundary layer, which then propagates outward and downstream. The combustion process creates a high pressure region over the back of the projectile resulting in a net positive thrust forward.

Stability of hypersonic boundary-layer flows with chemistry

NASA Technical Reports Server (NTRS)

Reed, Helen L.; Stuckert, Gregory K.; Haynes, Timothy S.

1993-01-01

The effects of nonequilibrium chemistry and three dimensionality on the stability characteristics of hypersonic flows are discussed. In two-dimensional (2-D) and axisymmetric flows, the inclusion of chemistry causes a shift of the second mode of Mack to lower frequencies. This is found to be due to the increase in size of the region of relative supersonic flow because of the lower speeds of sound in the relatively cooler boundary layers. Although this shift in frequency is present in both the equilibrium and nonequilibrium air results, the equilibrium approximation predicts modes which are not observed in the nonequilibrium calculations (for the flight conditions considered). These modes are superpositions of incoming and outgoing unstable disturbances which travel supersonically relative to the boundary-layer edge velocity. Such solutions are possible because of the finite shock stand-off distance. Their corresponding wall-normal profiles exhibit an oscillatory behavior in the inviscid region between the boundary-layer edge and the bow shock. For the examination of three-dimensional (3-D) effects, a rotating cone is used as a model of a swept wing. An increase of stagnation temperature is found to be only slightly stabilizing. The correlation of transition location (N = 9) with parameters describing the crossflow profile is discussed. Transition location does not correlate with the traditional crossflow Reynolds number. A new parameter that appears to correlate for boundary-layer flow was found. A verification with experiments on a yawed cone is provided.

Multi-fidelity numerical simulations of shock/turbulent-boundary layer interaction with uncertainty quantification

NASA Astrophysics Data System (ADS)

Bermejo-Moreno, Ivan; Campo, Laura; Larsson, Johan; Emory, Mike; Bodart, Julien; Palacios, Francisco; Iaccarino, Gianluca; Eaton, John

2013-11-01

We study the interaction between an oblique shock wave and the turbulent boundary layers inside a nearly-square duct by combining wall-modeled LES, 2D and 3D RANS simulations, targeting the experiment of Campo, Helmer & Eaton, 2012 (nominal conditions: M = 2 . 05 , Reθ = 6 , 500). A primary objective is to quantify the effect of aleatory and epistemic uncertainties on the STBLI. Aleatory uncertainties considered include the inflow conditions (Mach number of the incoming air stream and thickness of the boundary layers) and perturbations of the duct geometry upstream of the interaction. The epistemic uncertainty under consideration focuses on the RANS turbulence model form by injecting perturbations in the Reynolds stress anisotropy in regions of the flow where the model assumptions (in particular, the Boussinesq eddy-viscosity hypothesis) may be invalid. These perturbations are then propagated through the flow solver into the solution. The uncertainty quantification (UQ) analysis is done through 2D and 3D RANS simulations, assessing the importance of the three-dimensional effects imposed by the nearly-square duct geometry. Wall-modeled LES are used to verify elements of the UQ methodology and to explore the flow features and physics of the STBLI for multiple shock strengths. Financial support from the United States Department of Energy under the PSAAP program is gratefully acknowledged.

Control of a Normal Shock Boundary Layer Interaction with Ramped Vanes of Various Sizes

NASA Astrophysics Data System (ADS)

Lee, Sang; Loth, Eric

2017-11-01

A novel vortex generator design positioned upstream of a normal shock and a subsequent diffuser was investigated using large eddy simulations. In particular, ``ramped-vane'' flow control devices with three difference heights relative to the incoming boundary layer thickness (0.34 δ 0.52 δ and 0.75 δ were placed in a supersonic boundary layer with a freestream Mach number of 1.3 and a Reynolds number of 2,400 based on momentum thickness. These devices are similar to subsonic vanes but are designed to be more mechanically robust while having low wave drag. The devices generated strong streamwise vortices that entrained high momentum fluid to the near-wall region and increased turbulent mixing. The devices also decreased shock-induced flow separation, which resulted in a higher downstream skin friction in the diffuser. In general, the largest ramped-vane (0.75 δ) produced the largest reductions in flow separation, shape factor and overall unsteadiness. However, the medium-sized ramped vane (0.52 δ) was able to also reduce both the separation area and the diffuser displacement thickness. The smallest device (0.34 δ) had a weak impact of the flow in the diffuser, though a 10% reduction in the shape factor was achieved.

Experimental, Theoretical, and Computational Investigation of Separated Nozzle Flows

NASA Technical Reports Server (NTRS)

Hunter, Craig A.

2004-01-01

A detailed experimental, theoretical, and computational study of separated nozzle flows has been conducted. Experimental testing was performed at the NASA Langley 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a sub-scale, non-axisymmetric, two-dimensional, convergent- divergent nozzle. In addition, two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and algebraic Reynolds stress modeling. For reference, experimental and computational results were compared with theoretical predictions based on one-dimensional gas dynamics and an approximate integral momentum boundary layer method. Experimental results from this study indicate that off-design overexpanded nozzle flow was dominated by shock induced boundary layer separation, which was divided into two distinct flow regimes; three- dimensional separation with partial reattachment, and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results indicate that with controlled separation, the entire overexpanded range of nozzle performance would be within 10% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical exhaust system, this may allow a fixed geometry nozzle to cover an entire flight envelope. The computational simulation was in excellent agreement with experimental data over most of the test range, and did a good job of modeling internal flow and thrust performance. An exception occurred at low nozzle pressure ratios, where the two-dimensional computational model was inconsistent with the three-dimensional separation observed in the experiment. In general, the computation captured the physics of the shock boundary layer interaction and shock induced boundary layer separation in the nozzle, though there were some differences in shock structure compared to experiment. Though minor, these differences could be important for studies involving flow control or thrust vectoring of separated nozzles. Combined with other observations, this indicates that more detailed, three-dimensional computational modeling needs to be conducted to more realistically simulate shock-separated nozzle flows.

Measurements in Regions of Shock Wave/Turbulent Boundary Layer Interaction from Mach 3 to 10 for Open and Blind Code Evaluation/Validation

DTIC Science & Technology

2013-03-01

rakes containing pitot pressure probes, stagnation heat transfer gauges on hemispherical cylinders, total temperature measurements with vented...defined configurations in both “true temperature” and “cold” supersonic and hypersonic flows with boundary and flow conditions to provide the basis for...conducted over the past 50 years to provide data on both wedge- and shock- induced turbulent separated regions in supersonic and hypersonic flows suffer from

Physics and Chemistry of MW Laser-induced Discharge in Gas Flows and Plasma Jets

DTIC Science & Technology

2007-12-01

with the large scaled flow pulsations . In 3.3 the results of numerical modeling of a thin low-density heated channel of limited length – shock layer...in Fig. 3.2.13. The red points correspond to the values of time moments for Fig. 3.2.11, 12. Mechanism of heated area boundary pulsations ...Mechanism of heated area boundary pulsations is analogical to described above mechanism of the bow shock position pulsations and is connected with

Uncertainty Assessments of 2D and Axisymmetric Hypersonic Shock Wave - Turbulent Boundary Layer Interaction Simulations at Compression Corners

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Berry, Scott A.; VanNorman, John W.

2011-01-01

This paper is one of a series of five papers in a special session organized by the NASA Fundamental Aeronautics Program that addresses uncertainty assessments for CFD simulations in hypersonic flow. Simulations of a shock emanating from a compression corner and interacting with a fully developed turbulent boundary layer are evaluated herein. Mission relevant conditions at Mach 7 and Mach 14 are defined for a pre-compression ramp of a scramjet powered vehicle. Three compression angles are defined, the smallest to avoid separation losses and the largest to force a separated flow engaging more complicated flow physics. The Baldwin-Lomax and the Cebeci-Smith algebraic models, the one-equation Spalart-Allmaras model with the Catrix-Aupoix compressibility modification and two-equation models including Menter SST, Wilcox k-omega 98, and Wilcox k-omega 06 turbulence models are evaluated. Each model is fully defined herein to preclude any ambiguity regarding model implementation. Comparisons are made to existing experimental data and Van Driest theory to provide preliminary assessment of model form uncertainty. A set of coarse grained uncertainty metrics are defined to capture essential differences among turbulence models. Except for the inability of algebraic models to converge for some separated flows there is no clearly superior model as judged by these metrics. A preliminary metric for the numerical component of uncertainty in shock-turbulent-boundary-layer interactions at compression corners sufficiently steep to cause separation is defined as 55%. This value is a median of differences with experimental data averaged for peak pressure and heating and for extent of separation captured in new, grid-converged solutions presented here. This value is consistent with existing results in a literature review of hypersonic shock-turbulent-boundary-layer interactions by Roy and Blottner and with more recent computations of MacLean.

Formation of a cavitation cluster in the vicinity of a quasi-empty rupture

NASA Astrophysics Data System (ADS)

Bol'shakova, E. S.; Kedrinskiy, V. K.

2017-09-01

The presentation deals with one of the experimental and numerical models of a quasi-empty rupture in the magma melt. This rupture is formed in the liquid layer of a distilled cavitating fluid under shock loading within the framework of the problem formulation with a small electromagnetic hydrodynamic shock tube. It is demonstrated that the rupture is shaped as a spherical segment, which retains its topology during the entire process of its evolution and collapsing. The dynamic behavior of the quasi-empty rupture is analyzed, and the growth of cavitating nuclei in the form of the boundary layer near the entire rupture interface is found. It is shown that rupture implosion is accompanied by the transformation of the bubble boundary layer to a cavitating cluster, which takes the form of a ring-shaped vortex floating upward to the free surface of the liquid layer. A p-κ mathematical model is formulated, and calculations are performed to investigate the implosion of a quasi-empty spherical cavity in the cavitating liquid, generation of a shock wave by this cavity, and dynamics of the bubble density growth in the cavitating cluster by five orders of magnitude.

Boundary Layer Effects on Unsteady Airloads.

DTIC Science & Technology

1981-02-01

Magnus have shown by a "viscous ramp" behind the shock, whose inclination and height can be deduced from measured shock values such that calculated shock...sat- isfactory treatment of the shock). See YanglII -6 for these specific results. " Magnus 1 1 -7 (solution of the complete, nonlinear, inviscid...34, AFFDL-TR-78-202, December 1978. 111-7 R. J. Magnus and H. Yoshihara, "Calculations of Transonic Flow Over an Oscillating Airfoil", AIAA Paper 75-98

Letter: Transient interaction between plasma jet and supersonic compression ramp flow

NASA Astrophysics Data System (ADS)

Huang, He-Xia; Tan, Hui-Jun; Sun, Shu; Zhang, Yu-Chao; Cheng, Lin

2018-04-01

The rapid flow evolution between a plasma jet and a 20° compression ramp flow is captured by a high-speed schlieren system at Mach 2.0. Several interesting flow phenomena are observed for the first time. The pulsed jet, which generates strong perturbations, forces the crossflow boundary layer to separate and forms a forward moving shock. A typical shock-on-shock interaction occurs when the precursor shock intersects with the original shock. The interaction is initially regular, and then it transforms into an irregular one with a Mach stem connecting the precursor shock and original ramp shock.

Towards a Viscous Wall Model for Immersed Boundary Methods

NASA Technical Reports Server (NTRS)

Brehm, Christoph; Barad, Michael F.; Kiris, Cetin C.

2016-01-01

Immersed boundary methods are frequently employed for simulating flows at low Reynolds numbers or for applications where viscous boundary layer effects can be neglected. The primary shortcoming of Cartesian mesh immersed boundary methods is the inability of efficiently resolving thin turbulent boundary layers in high-Reynolds number flow application. The inefficiency of resolving the thin boundary is associated with the use of constant aspect ratio Cartesian grid cells. Conventional CFD approaches can efficiently resolve the large wall normal gradients by utilizing large aspect ratio cells near the wall. This paper presents different approaches for immersed boundary methods to account for the viscous boundary layer interaction with the flow-field away from the walls. Different wall modeling approaches proposed in previous research studies are addressed and compared to a new integral boundary layer based approach. In contrast to common wall-modeling approaches that usually only utilize local flow information, the integral boundary layer based approach keeps the streamwise history of the boundary layer. This allows the method to remain effective at much larger y+ values than local wall modeling approaches. After a theoretical discussion of the different approaches, the method is applied to increasingly more challenging flow fields including fully attached, separated, and shock-induced separated (laminar and turbulent) flows.

A Experimental Study of Fluctuating Pressure Loads Beneath Swept Shock Wave/boundary Layer Interactions

NASA Astrophysics Data System (ADS)

Garg, Sanjay

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath three-dimensional shock wave/boundary layer interactions is described. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins. Fin angles from 10 ^circ to 20^circ at freestream Mach numbers of 3 and 4 produce a variety of interaction strengths from weak to very strong. Miniature pressure transducers flush-mounted in the flat plate have been used to measure interaction-induced wall pressure fluctuations. The distributions of properties of the pressure fluctuations, such as their rms level, amplitude distribution and power spectra, are also determined. Measurements have been made for the first time in the aft regions of these interactions, revealing fluctuating pressure levels as high as 155 dB, which places them in the category of significant aeroacoustic load generators. The fluctuations near the foot of the fin are dominated by low frequency (0-5 kHz) components, and are caused by a previously unrecognized random motion of the primary attachment line. This phenomenon is probably intimately linked to the unsteadiness of the separation shock at the start of the interaction. The characteristics of the pressure fluctuations are explained in light of the features of the interaction flowfield. In particular, physical mechanisms responsible for the generation of high levels of surface pressure fluctuations are proposed based on the results of the study. The unsteadiness of the flowfield of the surface is also examined via a novel, non-intrusive optical technique. Results show that the entire shock structure generated by the interaction undergoes relatively low-frequency oscillations.

Effects of Hybrid Flow Control on a Normal Shock Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.; Vyas, Manan A.

2013-01-01

Hybrid flow control, a combination of micro-ramps and steady micro-jets, was experimentally investigated in the 15x15 cm Supersonic Wind Tunnel at the NASA Glenn Research Center. A central composite design of experiments method, was used to develop response surfaces for boundary-layer thickness and reversed-flow thickness, with factor variables of inter-ramp spacing, ramp height and chord length, and flow injection ratio. Boundary-layer measurements and wall static pressure data were used to understand flow separation characteristics. A limited number of profiles were measured in the corners of the tunnel to aid in understanding the three-dimensional characteristics of the flowfield.

Numerical investigation of the boundary layer separation in chemical oxygen iodine laser

NASA Astrophysics Data System (ADS)

Huai, Ying; Jia, Shuqin; Wu, Kenan; Jin, Yuqi; Sang, Fengting

2017-11-01

Large eddy simulation is carried out to model the flow process in a supersonic chemical oxygen iodine laser. Unlike the common approaches relying on the tensor representation theory only, the model in the present work is an explicit anisotropy-resolving algebraic Subgrid-scale scalar flux formulation. With an accuracy in capturing the unsteady flow behaviours in the laser. Boundary layer separation initiated by the adverse pressure gradient is identified using Large Eddy Simulation. To quantify the influences of flow boundary layer on the laser performance, the fluid computations coupled with a physical optics loaded cavity model is developed. It has been found that boundary layer separation has a profound effect on the laser outputs due to the introduced shock waves. The F factor of the output beam decreases to 10% of the original one when the boundary transit into turbulence for the setup depicted in the paper. Because the pressure is always greater on the downstream of the boundary layer, there will always be a tendency of boundary separation in the laser. The results inspire designs of the laser to apply positive/passive control methods avoiding the boundary layer perturbation.

Computational and experimental investigation of two-dimensional scramjet inlets and hypersonic flow over a sharp flat plate

NASA Astrophysics Data System (ADS)

Messitt, Donald G.

1999-11-01

The WIND code was employed to compute the hypersonic flow in the shock wave boundary layer merged region near the leading edge of a sharp flat plate. Solutions were obtained at Mach numbers from 9.86 to 15.0 and free stream Reynolds numbers of 3,467 to 346,700 in-1 (1.365 · 105 to 1.365 · 107 m-1) for perfect gas conditions. The numerical results indicated a merged shock wave and viscous layer near the leading edge. The merged region grew in size with increasing free stream Mach number, proportional to Minfinity 2/Reinfinity. Profiles of the static pressure in the merged region indicated a strong normal pressure gradient (∂p/∂y). The normal pressure gradient has been neglected in previous analyses which used the boundary layer equations. The shock wave near the leading edge was thick, as has been experimentally observed. Computed shock wave locations and surface pressures agreed well within experimental error for values of the rarefaction parameter, chi/M infinity2 < 0.3. A preliminary analysis using kinetic theory indicated that rarefied flow effects became important above this value. In particular, the WIND solution agreed well in the transition region between the merged flow, which was predicted well by the theory of Li and Nagamatsu, and the downstream region where the strong interaction theory applied. Additional computations with the NPARC code, WIND's predecessor, demonstrated the ability of the code to compute hypersonic inlet flows at free stream Mach numbers up to 20. Good qualitative agreement with measured pressure data indicated that the code captured the important physical features of the shock wave - boundary layer interactions. The computed surface and pitot pressures fell within the combined experimental and numerical error bounds for most points. The calculations demonstrated the need for extremely fine grids when computing hypersonic interaction flows.

High enthalpy hypersonic boundary layer flow

NASA Technical Reports Server (NTRS)

Yanow, G.

1972-01-01

A theoretical and experimental study of an ionizing laminar boundary layer formed by a very high enthalpy flow (in excess of 12 eV per atom or 7000 cal/gm) with allowance for the presence of helium driver gas is described. The theoretical investigation has shown that the use of variable transport properties and their respective derivatives is very important in the solution of equilibrium boundary layer equations of high enthalpy flow. The effect of low level helium contamination on the surface heat transfer rate is minimal. The variation of ionization is much smaller in a chemically frozen boundary layer solution than in an equilibrium boundary layer calculation and consequently, the variation of the transport properties in the case of the former was not essential in the integration. The experiments have been conducted in a free piston shock tunnel, and a detailed study of its nozzle operation, including the effects of low levels of helium driver gas contamination has been made. Neither the extreme solutions of an equilibrium nor of a frozen boundary layer will adequately predict surface heat transfer rate in very high enthalpy flows.

Relaxation of an unsteady turbulent boundary layer on a flat plate in an expansion tube

NASA Technical Reports Server (NTRS)

Gurta, R. N.; Trimpi, R. L.

1974-01-01

An analysis is presented for the relaxation of a turbulent boundary layer on a semi-infinite flat plate after passage of a shock wave and a trailing driver gas-driven gas interface. The problem has special application to expansion-tube flows. The flow-governing equations have been transformed into the Crocco variables, and a time-similar solution is presented in terms of the dimensionless distance-time variable alpha and the dimensionless velocity variable beta. An eddy-viscosity model, similar to that of time-steady boundary layers, is applied to the inner and outer regions of the boundary layer. A turbulent Prandtl number equal to the molecular Prandtl number is used to relate the turbulent heat flux to the eddy viscosity. The numerical results, obtained by using the Gauss-Seidel line-relaxation method, indicate that a fully turbulent boundary layer relaxes faster to the final steady-state values of heat transfer and skin friction than a laminar boundary layer. The results also give a fairly good estimate of the local skin friction and heat transfer for near steady-flow conditions.

Monte Carlo simulations of thermal comptonization process in a two-component advective flow around a neutron star

NASA Astrophysics Data System (ADS)

Bhattacharjee, Ayan; Chakrabarti, Sandip K.

2017-12-01

We explore spectral properties of a two-component advective flow around a neutron star. We compute the effects of thermal Comptonization of soft photons emitted from a Keplerian disc and the boundary layer of the neutron star by the post-shock region of a sub-Keplerian flow, formed due to the centrifugal barrier. The shock location Xs is also the inner edge of the Keplerian disc. We compute a series of realistic spectra assuming a set of electron temperatures of the post-shock region TCE, the temperature of the Normal BOundary Layer (NBOL) TNS of the neutron star and the shock location Xs. These parameters depend on the disc and halo accretion rates (\\dot{m}d and \\dot{m}h, respectively) that control the resultant spectra. We find that the spectrum becomes harder when \\dot{m}_h is increased. The spectrum is controlled strongly by TNS due to its proximity to the Comptonizing cloud since photons emitted from the NBOL cool down the post-shock region very effectively. We also show the evidence of spectral hardening as the inclination angle of the disc is increased.

Performance of Low Dissipative High Order Shock-Capturing Schemes for Shock-Turbulence Interactions

NASA Technical Reports Server (NTRS)

Sandham, N. D.; Yee, H. C.

1998-01-01

Accurate and efficient direct numerical simulation of turbulence in the presence of shock waves represents a significant challenge for numerical methods. The objective of this paper is to evaluate the performance of high order compact and non-compact central spatial differencing employing total variation diminishing (TVD) shock-capturing dissipations as characteristic based filters for two model problems combining shock wave and shear layer phenomena. A vortex pairing model evaluates the ability of the schemes to cope with shear layer instability and eddy shock waves, while a shock wave impingement on a spatially-evolving mixing layer model studies the accuracy of computation of vortices passing through a sequence of shock and expansion waves. A drastic increase in accuracy is observed if a suitable artificial compression formulation is applied to the TVD dissipations. With this modification to the filter step the fourth-order non-compact scheme shows improved results in comparison to second-order methods, while retaining the good shock resolution of the basic TVD scheme. For this characteristic based filter approach, however, the benefits of compact schemes or schemes with higher than fourth order are not sufficient to justify the higher complexity near the boundary and/or the additional computational cost.

Turbulence measurements in hypersonic shock-wave boundary-layer interaction flows

NASA Technical Reports Server (NTRS)

Mikulla, V.; Horstman, C. C.

1976-01-01

Turbulent intensity and Reynolds shear stress measurements are presented for two nonadiabatic hypersonic shock-wave boundary-layer interaction flows, one with and one without separation. These measurements were obtained using a new hot-wire probe specially designed for heated flows. Comparison of the separated and attached flows shows a significant increase above equilibrium values in the turbulent intensity and shear stress downstream of the interaction region for the attached case, while for the separated case, the turbulent fluxes remain close to equilibrium values. This effect results in substantial differences in turbulence lifetime for the two flows. We propose that these differences are due to a coupling between the turbulent energy and separation bubble unsteadiness, a hypothesis supported by the statistical properties of the turbulent fluctuations.

Investigation of Heat Transfer to a Flat Plate in a Shock Tube.

DTIC Science & Technology

1987-12-01

2 Objectives and Scope . . . . . .. .. .. .... 5 11. Theory ............... ....... 7 Shock Tube Principles........... 7 Boundary Layer Theory ...in *excess of theory , but the rounded edge flat plate exhibited data which matched or was less than what theory predicted for each Mach number tested...normal shock advancing along an infinite flat plate. For x< Ugt there is a region of interaction between the downstream influence of the leading edge

Numerical investigation of over expanded flow behavior in a single expansion ramp nozzle

NASA Astrophysics Data System (ADS)

Mousavi, Seyed Mahmood; Pourabidi, Reza; Goshtasbi-Rad, Ebrahim

2018-05-01

The single expansion ramp nozzle is severely over-expanded when the vehicle is at low speed, which hinders its ability to provide optimal configurations for combined cycle engines. The over-expansion leads to flow separation as a result of shock wave/boundary-layer interaction. Flow separation, and the presence of shocks themselves, result in a performance loss in the single expansion ramp nozzle, leading to reduced thrust and increased pressure losses. In the present work, the unsteady two dimensional compressible flow in an over expanded single expansion ramp nozzle has been investigated using finite volume code. To achieve this purpose, the Reynolds stress turbulence model and full multigrid initialization, in addition to the Smirnov's method for examining the errors accumulation, have been employed and the results are compared with available experimental data. The results show that the numerical code is capable of predicting the experimental data with high accuracy. Afterward, the effect of discontinuity jump in wall temperature as well as the length of straight ramp on flow behavior have been studied. It is concluded that variations in wall temperature and length of straight ramp change the shock wave boundary layer interaction, shock structure, shock strength as well as the distance between Lambda shocks.

HIFiRE-5 Boundary Layer Transition and HIFiRE-1 Shock Boundary Layer Interaction

DTIC Science & Technology

2015-10-01

pressure calcium-fluoride window and a new wind tunnel model consisting of a steel forebody and a PEEK plastic shell. This model contained a large...wind tunnel data. These models were then applied to the supersonic 2 Approved for public release; distribution unlimited. ascent phase of the...the tuned RANS models predicted upstream influence well. These results lend confidence to the approach of calibrating RANS models with wind tunnel

Optimization of bump and blowing to control the flow through a transonic compressor blade cascade

NASA Astrophysics Data System (ADS)

Mazaheri, K.; Khatibirad, S.

2018-03-01

Shock control bump (SCB) and blowing are two flow control methods, used here to improve the aerodynamic performance of transonic compressors. Both methods are applied to a NASA rotor 67 blade section and are optimized to minimize the total pressure loss. A continuous adjoint algorithm is used for multi-point optimization of a SCB to improve the aerodynamic performance of the rotor blade section, for a range of operational conditions around its design point. A multi-point and two single-point optimizations are performed in the design and off-design conditions. It is shown that the single-point optimized shapes have the best performance for their respective operating conditions, but the multi-point one has an overall better performance over the whole operating range. An analysis is given regarding how similarly both single- and multi-point optimized SCBs change the wave structure between blade sections resulting in a more favorable flow pattern. Interactions of the SCB with the boundary layer and the wave structure, and its effects on the separation regions are also studied. We have also introduced the concept of blowing for control of shock wave and boundary-layer interaction. A geometrical model is introduced, and the geometrical and physical parameters of blowing are optimized at the design point. The performance improvements of blowing are compared with the SCB. The physical interactions of SCB with the boundary layer and the shock wave are analyzed. The effects of SCB on the wave structure in the flow domain outside the boundary-layer region are investigated. It is shown that the effects of the blowing mechanism are very similar to the SCB.

A Study of Supersonic Compression-Corner Interactions using Hybrid LES/RANS Models

DTIC Science & Technology

2014-01-20

Mach 2.5 shock / boundary layer interaction in a wind tunnel (experiments conducted at Cambridge University [15]) as a means of assessing methods... wind tunnel . The shock impinges upon the bottom surface of the wind tunnel , creating a region of shock -separated flow. The structure of the SBLI... waves into a shock wave (Figure 19, X = 0.1016 and X = 0.1278 stations) are also not well-predicted. The hot-wire measurements may not be as

Flow of supersonic jets across flat plates: Implications for ground-level flow from volcanic blasts

NASA Astrophysics Data System (ADS)

Orescanin, Mara M.; Prisco, David; Austin, Joanna M.; Kieffer, Susan W.

2014-04-01

We report on laboratory experiments examining the interaction of a jet from an overpressurized reservoir with a canonical ground surface to simulate lateral blasts at volcanoes such as the 1980 blast at Mount St. Helens. These benchmark experiments test the application of supersonic jet models to simulate the flow of volcanic jets over a lateral topography. The internal shock structure of the free jet is modified such that the Mach disk shock is elevated above the surface. In elevation view, the width of the shock is reduced in comparison with a free jet, while in map view the dimensions are comparable. The distance of the Mach disk shock from the vent is in good agreement with free jet data and can be predicted with existing theory. The internal shock structures can interact with and penetrate the boundary layer. In the shock-boundary layer interaction, an oblique shock foot is present in the schlieren images and a distinctive ground signature is evident in surface measurements. The location of the oblique shock foot and the surface demarcation are closely correlated with the Mach disk shock location during reservoir depletion, and therefore, estimates of a ground signature in a zone devastated by a blast can be based on the calculated shock location from free jet theory. These experiments, combined with scaling arguments, suggest that the imprint of the Mach disk shock on the ground should be within the range of 4-9 km at Mount St. Helens depending on assumed reservoir pressure and vent dimensions.

The LENS Facilities and Experimental Studies to Evaluate the Modeling of Boundary Layer Transition, Shock/Boundary Layer Interaction, Real Gas, Radiation and Plasma Phenomena in Contemporary CFD Codes

DTIC Science & Technology

2010-04-01

Layer Interaction, Real Gas, Radiation and Plasma Phenomena in Contemporary CFD Codes Michael S. Holden, PhD CUBRC , Inc. 4455 Genesee Street Buffalo...NUMBER 5f. WORK UNIT NUMBER 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) CUBRC , Inc. 4455 Genesee Street Buffalo, NY 14225, USA 8. PERFORMING...HyFly Navy EMRG Reentry-F Slide 2 X-43 HIFiRE-2 Figure 17: Transition in Hypervelocity Flows: CUBRC Focus – Fully Duplicated Ground Test

Supersonic flow visualization of a nacelle in close proximity to a simulated wing

NASA Technical Reports Server (NTRS)

Biber, Kasim; Ellis, David R.

1993-01-01

A flow visualization study was made in the 9 x 9 inch supersonic wind tunnel at Wichita State University to examine shock and boundary layer flow interaction for a nacelle in close proximity to the lower surface of a simulated wing. The test matrix included variations of angle of attack from -2 degrees to +4 degrees, nacelle-wing gap from 0.5 to 3-nacelle inlet diameter (0.12 inch), and Reynolds number based on nacelle length (1.164 inch) from 1.16 x 10(exp 6) to 1.45 x 10(exp 6) at a nominal Mach number of 2. Schlieren pictures of wing and nacelle flowfield were recorded by a video camera during each tunnel run. Results show that the nacelle inlet shock wave remains attached to the inlet lip and its impingement does not significantly affect the wing boundary layer. At the nacelle trailing edge location, the wing boundary layer thickness is approximately one nacelle inlet diameter at alpha = 0 degrees and it decreases with increase of angle of attack.

An approximate viscous shock layer technique for calculating chemically reacting hypersonic flows about blunt-nosed bodies

NASA Technical Reports Server (NTRS)

Cheatwood, F. Mcneil; Dejarnette, Fred R.

1991-01-01

An approximate axisymmetric method was developed which can reliably calculate fully viscous hypersonic flows over blunt nosed bodies. By substituting Maslen's second order pressure expression for the normal momentum equation, a simplified form of the viscous shock layer (VSL) equations is obtained. This approach can solve both the subsonic and supersonic regions of the shock layer without a starting solution for the shock shape. The approach is applicable to perfect gas, equilibrium, and nonequilibrium flowfields. Since the method is fully viscous, the problems associated with a boundary layer solution with an inviscid layer solution are avoided. This procedure is significantly faster than the parabolized Navier-Stokes (PNS) or VSL solvers and would be useful in a preliminary design environment. Problems associated with a previously developed approximate VSL technique are addressed before extending the method to nonequilibrium calculations. Perfect gas (laminar and turbulent), equilibrium, and nonequilibrium solutions were generated for airflows over several analytic body shapes. Surface heat transfer, skin friction, and pressure predictions are comparable to VSL results. In addition, computed heating rates are in good agreement with experimental data. The present technique generates its own shock shape as part of its solution, and therefore could be used to provide more accurate initial shock shapes for higher order procedures which require starting solutions.

Mitigation of Adverse Effects Caused by Shock Wave Boundary Layer Interactions Through Optimal Wall Shaping

NASA Technical Reports Server (NTRS)

Liou, May-Fun; Lee, Byung Joon

2013-01-01

It is known that the adverse effects of shock wave boundary layer interactions in high speed inlets include reduced total pressure recovery and highly distorted flow at the aerodynamic interface plane (AIP). This paper presents a design method for flow control which creates perturbations in geometry. These perturbations are tailored to change the flow structures in order to minimize shock wave boundary layer interactions (SWBLI) inside supersonic inlets. Optimizing the shape of two dimensional micro-size bumps is shown to be a very effective flow control method for two-dimensional SWBLI. In investigating the three dimensional SWBLI, a square duct is employed as a baseline. To investigate the mechanism whereby the geometric elements of the baseline, i.e. the bottom wall, the sidewall and the corner, exert influence on the flow's aerodynamic characteristics, each element is studied and optimized separately. It is found that arrays of micro-size bumps on the bottom wall of the duct have little effect in improving total pressure recovery though they are useful in suppressing the incipient separation in three-dimensional problems. Shaping sidewall geometry is effective in re-distributing flow on the side wall and results in a less distorted flow at the exit. Subsequently, a near 50% reduction in distortion is achieved. A simple change in corner geometry resulted in a 2.4% improvement in total pressure recovery.

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. Part 2: Wall shear stress

NASA Technical Reports Server (NTRS)

Liou, M. S.; Adamson, T. C., Jr.

1979-01-01

An analysis is presented of the flow in the two inner layers, the Reynolds stress sublayer and the wall layer. Included is the calculation of the shear stress at the wall in the interaction region. The limit processes considered are those used for an inviscid flow.

Shock wave propagation in layered planetary embryos

NASA Astrophysics Data System (ADS)

Arkani-Hamed, Jafar; Ivanov, Boris A.

2014-05-01

The propagation of impact-induced shock wave inside a planetary embryo is investigated using the Hugoniot equations and a new scaling law, governing the particle velocity variations along a shock ray inside a spherical body. The scaling law is adopted to determine the impact heating of a growing embryo in its early stage when it is an undifferentiated and uniform body. The new scaling law, similar to other existing scaling laws, is not suitable for a large differentiated embryo consisting of a silicate mantle overlying an iron core. An algorithm is developed in this study on the basis of the ray theory in a spherically symmetric body which relates the shock parameters at the top of the core to those at the base of the mantle, thus enabling the adoption of scaling laws to estimate the impact heating of both the mantle and the core. The algorithm is applied to two embryo models: a simple two-layered model with a uniform mantle overlying a uniform core, and a model where the pre-shock density and acoustic velocity of the embryo are radially dependent. The former illustrates details of the particle velocity, shock pressure, and temperature increase behind the shock front in a 2D axisymmetric geometry. The latter provides a means to compare the results with those obtained by a hydrocode simulation. The agreement between the results of the two techniques in revealing the effects of the core-mantle boundary on the shock wave transmission across the boundary is encouraging.

Volumetric imaging of supersonic boundary layers using filtered Rayleigh scattering background suppression

NASA Technical Reports Server (NTRS)

Forkey, Joseph N.; Lempert, Walter R.; Bogdonoff, Seymour M.; Miles, Richard B.; Russell, G.

1995-01-01

We demonstrate the use of Filtererd Rayleigh Scattering and a 3D reconstruction technique to interrogate the highly three dimensional flow field inside of a supersonic inlet model. A 3 inch by 3 inch by 2.5 inch volume is reconstructed yielding 3D visualizations of the crossing shock waves and of the boundary layer. In this paper we discuss the details of the techniques used, and present the reconstructured 3D images.

Computational Fluid Dynamics (CFD) Simulations of a Finned Projectile with Microflaps for Flow Control

DTIC Science & Technology

2016-04-01

fields associated with these control mechanisms for US Army weapons are complex, involving 3-dimensional (3-D) shock- boundary layer interactions...distribution over the rear finned section and thus produce control forces and moments. Dykes et al.6 used a flat - plate fin interaction design of...cells—tetrahedrals, triangular prisms, and pyramids—were used in the mesh. Grid points shown in Fig. 3a were clustered in the boundary layer region

Flow analysis for the nacelle of an advanced ducted propeller at high angle-of-attack and at cruise with boundary layer control

NASA Technical Reports Server (NTRS)

Hwang, D. P.; Boldman, D. R.; Hughes, C. E.

1994-01-01

An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

Flow field predictions for a slab delta wing at incidence

NASA Technical Reports Server (NTRS)

Conti, R. J.; Thomas, P. D.; Chou, Y. S.

1972-01-01

Theoretical results are presented for the structure of the hypersonic flow field of a blunt slab delta wing at moderately high angle of attack. Special attention is devoted to the interaction between the boundary layer and the inviscid entropy layer. The results are compared with experimental data. The three-dimensional inviscid flow is computed numerically by a marching finite difference method. Attention is concentrated on the windward side of the delta wing, where detailed comparisons are made with the data for shock shape and surface pressure distributions. Surface streamlines are generated, and used in the boundary layer analysis. The three-dimensional laminar boundary layer is computed numerically using a specially-developed technique based on small cross-flow in streamline coordinates. In the rear sections of the wing the boundary layer decreases drastically in the spanwise direction, so that it is still submerged in the entropy layer at the centerline, but surpasses it near the leading edge. Predicted heat transfer distributions are compared with experimental data.

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. Part 1: Pressure distribution. Part 2: Wall shear stress. Part 3: Simplified formulas for the prediction of surface pressures and skin friction

NASA Technical Reports Server (NTRS)

Adamson, T. C., Jr.; Liou, M. S.; Messiter, A. F.

1980-01-01

An asymptotic description is derived for the interaction between a shock wave and a turbulent boundary layer in transonic flow, for a particular limiting case. The dimensionless difference between the external flow velocity and critical sound speed is taken to be much smaller than one, but large in comparison with the dimensionless friction velocity. The basic results are derived for a flat plate, and corrections for longitudinal wall curvature and for flow in a circular pipe are also shown. Solutions are given for the wall pressure distribution and the shape of the shock wave. Solutions for the wall shear stress are obtained, and a criterion for incipient separation is derived. Simplified solutions for both the wall pressure and skin friction distributions in the interaction region are given. These results are presented in a form suitable for use in computer programs.

An experimental and numerical investigation of shock-wave induced turbulent boundary-layer separation at hypersonic speeds

NASA Technical Reports Server (NTRS)

Marvin, J. G.; Horstman, C. C.; Rubesin, M. W.; Coakley, T. J.; Kussoy, M. I.

1975-01-01

An experiment designed to test and guide computations of the interaction of an impinging shock wave with a turbulent boundary layer is described. Detailed mean flow-field and surface data are presented for two shock strengths which resulted in attached and separated flows, respectively. Numerical computations, employing the complete time-averaged Navier-Stokes equations along with algebraic eddy-viscosity and turbulent Prandtl number models to describe shear stress and heat flux, are used to illustrate the dependence of the computations on the particulars of the turbulence models. Models appropriate for zero-pressure-gradient flows predicted the overall features of the flow fields, but were deficient in predicting many of the details of the interaction regions. Improvements to the turbulence model parameters were sought through a combination of detailed data analysis and computer simulations which tested the sensitivity of the solutions to model parameter changes. Computer simulations using these improvements are presented and discussed.

Simulation of ideal-gas flow by nitrogen and other selected gases at cryogenic temperatures. [transonic flow in cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Hall, R. M.; Adcock, J. B.

1981-01-01

The real gas behavior of nitrogen, the gas normally used in transonic cryogenic tunnels, is reported for the following flow processes: isentropic expansion, normal shocks, boundary layers, and interactions between shock waves and boundary layers. The only difference in predicted pressure ratio between nitrogen and an ideal gas which may limit the minimum operating temperature of transonic cryogenic wind tunnels occur at total pressures approaching 9 atm and total temperatures 10 K below the corresponding saturation temperature. These pressure differences approach 1 percent for both isentropic expansions and normal shocks. Alternative cryogenic test gases were also analyzed. Differences between air and an ideal diatomic gas are similar in magnitude to those for nitrogen and should present no difficulty. However, differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. It is concluded that helium and cryogenic hydrogen would not approximate the compressible flow of an ideal diatomic gas.

Turbulent Output-Based Anisotropic Adaptation

NASA Technical Reports Server (NTRS)

Park, Michael A.; Carlson, Jan-Renee

2010-01-01

Controlling discretization error is a remaining challenge for computational fluid dynamics simulation. Grid adaptation is applied to reduce estimated discretization error in drag or pressure integral output functions. To enable application to high O(10(exp 7)) Reynolds number turbulent flows, a hybrid approach is utilized that freezes the near-wall boundary layer grids and adapts the grid away from the no slip boundaries. The hybrid approach is not applicable to problems with under resolved initial boundary layer grids, but is a powerful technique for problems with important off-body anisotropic features. Supersonic nozzle plume, turbulent flat plate, and shock-boundary layer interaction examples are presented with comparisons to experimental measurements of pressure and velocity. Adapted grids are produced that resolve off-body features in locations that are not known a priori.

Turbulence measurements in hypersonic shock-wave boundary-layer interaction flows

NASA Technical Reports Server (NTRS)

Mikulla, V.; Horstman, C. C.

1976-01-01

Turbulent intensity and Reynolds shear stress measurements are presented for two nonadiabatic hypersonic shock-wave boundary-layer interaction flows, one with and one without separation. These measurements were obtained using a new hot-wire probe specially designed for heated flows. Comparison of the separated and attached flows shows a significant increase above equilibrium values in the turbulent intensity and shear stress downstream of the interaction region for the attached case, while for the separated case, the turbulent fluxes remain close to equilibrium values. This effect results in substantial differences in turbulence lifetimes for the two flows. It is proposed that these differences are due to a coupling between the turbulent energy and separation bubble unsteadiness, a hypothesis supported by the statistical properties of the turbulent fluctuations.

A comparison of hypersonic vehicle flight and prediction results

NASA Technical Reports Server (NTRS)

Iliff, Kenneth W.; Shafer, Mary F.

1995-01-01

Aerodynamic and aerothermodynamic comparisons between flight and ground test for four hypersonic vehicles are discussed. The four vehicles are the X-15, the Reentry F, the Sandia Energetic Reentry Vehicle Experiment (SWERVE), and the Space Shuttle. The comparisons are taken from papers published by researchers active in the various programs. Aerodynamic comparisons include reaction control jet interaction on the Space Shuttle. Various forms of heating including catalytic, boundary layer, shock interaction and interference, and vortex impingement are compared. Predictions were significantly exceeded for the heating caused by vortex impingement (on the Space Shuttle OMS pods) and for heating caused by shock interaction and interference on the X-15 and the Space Shuttle. Predictions of boundary-layer state were in error on the X-15, the SWERVE, and the Space Shuttle vehicles.

Physics of magnetospheric boundary layers

NASA Technical Reports Server (NTRS)

Cairns, Iver H.

1995-01-01

This final report was concerned with the ideas that: (1) magnetospheric boundary layers link disparate regions of the magnetosphere-solar wind system together; and (2) global behavior of the magnetosphere can be understood only by understanding its internal linking mechanisms and those with the solar wind. The research project involved simultaneous research on the global-, meso-, and micro-scale physics of the magnetosphere and its boundary layers, which included the bow shock, the magnetosheath, the plasma sheet boundary layer, and the ionosphere. Analytic, numerical, and simulation projects were performed on these subjects, as well as comparisons of theoretical results with observational data. Other related activity included in the research included: (1) prediction of geomagnetic activity; (2) global MHD (magnetohydrodynamic) simulations; (3) Alfven resonance heating; and (4) Critical Ionization Velocity (CIV) effect. In the appendixes are list of personnel involved, list of papers published; and reprints or photocopies of papers produced for this report.

An experimental study of the sources of fluctuating pressure loads beneath swept shock/boundary-layer interactions

NASA Technical Reports Server (NTRS)

Settles, G. S.; Garg, S.

1993-01-01

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath three dimensional shock wave/boundary layer interactions is described. Such loads constitute a fundamental problem of critical concern to future supersonic and hypersonic flight vehicles. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins at angle of attack. Fin angles from 10 to 20 deg at freestream Mach numbers of 3 and 4 produce a variety of interaction strengths from weak to very strong. Miniature Kulite pressure transducers flush-mounted in the flat plate are used to measure interaction-induced wall pressure fluctuations. The distributions of properties of the pressure fluctuations, such as their ring levels, amplitude distributions, and power spectra, are also determined. Measurements were made for the first time in the aft regions of these interactions, revealing fluctuating pressure levels as high as 160 dB. These fluctuations are dominated by low frequency (0-5 kHz) signals. The maximum ring levels in the interactions show an increasing trend with increasing interaction strength. On the other hand, the maximum ring levels in the forward portion of the interactions decrease linearly with increasing interaction sweep back. These ring pressure distributions and spectra are correlated with the features of the interaction flowfield. The unsteadiness of the off-surface flowfield is studied using a new, non-intrusive technique based on the shadow graph method. The results indicate that the entire lambda-shock structure generated by the interaction undergoes relatively low-frequency oscillations. Some regions where particularly strong fluctuations are generated were identified. Fluctuating pressure measurements are also made along the line of symmetry of an axisymmetric jet impinging upon a flat plate at an angle. This flow was chosen as a simple analog to the impinging jet region found in the rear portion of the shock wave/boundary layer interactions under study. It is found that a sharp peak in ring pressure level exists at or near the mean stagnation point. It is suggested that the phenomena responsible for this peak may be active in the swept interactions as well, and may cause the extremely high fluctuating pressures observed in the impinging jet region in the present experimental program.

Reynolds-Stress Budgets in an Impinging Shock Wave/Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Vyas, Manan A.; Yoder, Dennis A.; Gaitonde, Datta V.

2018-01-01

Implicit large-eddy simulation (ILES) of a shock wave/boundary-layer interaction (SBLI) was performed. Comparisons with experimental data showed a sensitivity of the current prediction to the modeling of the sidewalls. This was found to be common among various computational studies in the literature where periodic boundary conditions were used in the spanwise direction, as was the case in the present work. Thus, although the experiment was quasi-two-dimensional, the present simulation was determined to be two-dimensional. Quantities present in the exact equation of the Reynolds-stress transport, i.e., production, molecular diffusion, turbulent transport, pressure diffusion, pressure strain, dissipation, and turbulent mass flux were calculated. Reynolds-stress budgets were compared with past large-eddy simulation and direct numerical simulation datasets in the undisturbed portion of the turbulent boundary layer to validate the current approach. The budgets in SBLI showed the growth in the production term for the primary normal stress and energy transfer mechanism was led by the pressure strain term in the secondary normal stresses. The pressure diffusion term, commonly assumed as negligible by turbulence model developers, was shown to be small but non-zero in the normal stress budgets, however it played a key role in the primary shear stress budget.

Shock Boundary Layer Interaction Flow Control with Micro Vortex Generators

DTIC Science & Technology

2011-05-01

Pitot rake ( p̄02p01 ) u = time-averaged streamwise velocity ufs = time-averaged freestream streamwise velocity u∗ = √ τw ρw = wall-shear velocity w...upstream of the normal shock-wave 2 = station 2, at the Pitot rake location I. Introduction With the exception of the scramjet, all current air-breathing...to this.7 1 shock holder near-normal shock μVGs 123 143 14 hole Pitot rake 6o x vg variable φ cylinder mounted on the centre-line 380 M ∞ =1.4

3D Plenoptic PIV Measurements of a Shock Wave Boundary Layer Interaction

NASA Astrophysics Data System (ADS)

Thurow, Brian; Bolton, Johnathan; Arora, Nishul; Alvi, Farrukh

2016-11-01

Plenoptic particle image velocimetry (PIV) is a relatively new technique that uses the computational refocusing capability of a single plenoptic camera and volume illumination with a double-pulsed light source to measure the instantaneous 3D/3C velocity field of a flow field seeded with particles. In this work, plenoptic PIV is used to perform volumetric velocity field measurements of a shock-wave turbulent boundary layer interaction (SBLI). Experiments were performed in a Mach 2.0 flow with the SBLI produced by an unswept fin at 15°angle of attack. The measurement volume was 38 x 25 x 32 mm3 and illuminated with a 400 mJ/pulse Nd:YAG laser with 1.7 microsecond inter-pulse time. Conventional planar PIV measurements along two planes within the volume are used for comparison. 3D visualizations of the fin generated shock and subsequent SBLI are presented. The growth of the shock foot and separation region with increasing distance from the fin tip is observed and agrees with observations made using planar PIV. Instantaneous images depict 3D fluctuations in the position of the shock foot from one image to the next. The authors acknowledge the support of the Air Force Office of Scientific Research.

Detection and Characterization of Boundary-Layer Transition in Flight at Supersonic Conditions Using Infrared Thermography

NASA Technical Reports Server (NTRS)

Banks, Daniel W.

2008-01-01

Infrared thermography is a powerful tool for investigating fluid mechanics on flight vehicles. (Can be used to visualize and characterize transition, shock impingement, separation etc.). Updated onboard F-15 based system was used to visualize supersonic boundary layer transition test article. (Tollmien-Schlichting and cross-flow dominant flow fields). Digital Recording improves image quality and analysis capability. (Allows accurate quantitative (temperature) measurements, Greater enhancement through image processing allows analysis of smaller scale phenomena).

Numerical solution of the hypersonic viscous-shock-layer equations for laminar, transitional, and turbulent flows of a perfect gas over blunt axially symmetric bodies

NASA Technical Reports Server (NTRS)

Anderson, E. C.; Moss, J. N.

1975-01-01

The viscous shock layer equations applicable to hypersonic laminar, transitional, and turbulent flows of a perfect gas over two-dimensional plane or axially symmetric blunt bodies are presented. The equations are solved by means of an implicit finite difference scheme, and the results are compared with a turbulent boundary layer analysis. The agreement between the two solution procedures is satisfactory for the region of flow where streamline swallowing effects are negligible. For the downstream regions, where streamline swallowing effects are present, the expected differences in the two solution procedures are evident.

Experimental studies of transpiration cooling with shock interaction in hypersonic flow, part B

NASA Technical Reports Server (NTRS)

Holden, Michael S.

1994-01-01

This report describes the result of experimental studies conducted to examine the effects of the impingement of an oblique shock on the flowfield and surface characteristics of a transpiration-cooled wall in turbulent hypersonic flow. The principal objective of this work was to determine whether the interaction between the oblique shock and the low-momentum region of the transpiration-cooled boundary layer created a highly distorted flowfield and resulted in a significant reduction in the cooling effectiveness of the transpiration-cooled surface. As a part of this program, we also sought to determine the effectiveness of transpiration cooling with nitrogen and helium injectants for a wide range of blowing rates under constant-pressure conditions in the absence of shock interaction. This experimental program was conducted in the Calspan 48-Inch Shock Tunnel at nominal Mach numbers of 6 and 8, for a Reynolds number of 7.5 x 10(exp 6). For these test conditions, we obtained fully turbulent boundary layers upstream of the interaction regions over the transpiration-cooled segment of the flat plate. The experimental program was conducted in two phases. In the first phase, we examined the effects of mass-addition level and coolant properties on the cooling effectiveness of transpiration-cooled surfaces in the absence of shock interaction. In the second phase of the program, we examined the effects of oblique shock impingement on the flowfield and surface characteristics of a transpiration-cooled surface. The studies were conducted for a range of shock strengths with nitrogen and helium coolants to examine how the distribution of heat transfer and pressure and the characteristics of the flowfield in the interaction region varied with shock strength and the level of mass addition from the transpiration-cooled section of the model. The effects of the distribution of the blowing rate along the interaction regions were also examined for a range of blowing rates through the transpiration-cooled panels. The regions of shockwave/boundary layer interaction examined in these studies were induced by oblique shocks generated with a sharp, flat plate, inclined to the freestream at angles of 5 degrees, 7.5 degrees, and 10 degrees. It was found that, in the absence of an incident shock, transpiration cooling was a very effective method for reducing both the heat transfer and the skin friction loads on the surface. The helium coolant was found to be significantly more effective than nitrogen, because of its low molecular weight and high specific heat. The studies of shock-wave/transpiration-cooled surface interaction demonstrated that the interaction region between the incident shock and the low-momentum transpiration-cooled boundary layer did not result in a significant increase in the size of attached or separated interaction regions, and did not result in significant flowfield distortions above the interaction region. The increase in heating downstream of the shock-impingement point could easily be reduced to the values without shock impingement by a relatively small increase in the transpiration cooling in this region. Surprisingly, this increase in cooling rate did not result in a significant increase in size of the region ahead of the incident shock or create a significantly enlarged interaction region with a resultant increase in the distortion level in the inviscid flow. Thus, transpiration cooling appears to be a very effective technique to cool the internal surfaces of scramjet engines, where shocks in the engine would induce large local increases in wall heating and create viscous/inviscid interactions that could significantly disturb the smooth flow through the combustor. However, if hydrogen is used as the coolant, burning upstream of shock impingement might result in localized hot spots. Clearly, further research is needed in this area.

Fluid dynamic modeling and numerical simulation of low-density hypersonic flow

NASA Astrophysics Data System (ADS)

Cheng, H. K.; Wong, Eric Y.

1988-06-01

The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.

Fluid dynamic modeling and numerical simulation of low-density hypersonic flow

NASA Technical Reports Server (NTRS)

Cheng, H. K.; Wong, Eric Y.

1988-01-01

The concept of a viscous shock-layer and several related versions of continuum theories/methods are examined for their adequacy as a viable framework to study flow physics and aerothermodynamics of relevance to sustained hypersonic flights. Considering the flat plate at angle of attack, or the wedge, as a generic example for the major aerodynamic component of a hypersonic vehicle, the relative importance of the molecular-transport effects behind the shock (in the form of the 'shock slip') and the wall-slip effects are studied. In the flow regime where the shock-transition-zone thickness remains small compared to the shock radius of curvature, a quasi-one-dimensional shock structure under the Burnett/thirteen-moment approximation, as well as particulate/collisional models, can be consistently developed. The fully viscous version of the shock-layer model is shown to provide the crucial boundary condition downstream the shock in this case. The gas-kinetic basis of the continuum description for the flow behind the bow shock, and certain features affecting the non-equilibrium flow chemistry, are also discussed.

Numerical Investigation of Flow in an Over-Expanded Nozzle with Porous Surfaces

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa; Abdol-Hamid, K. S.; Hunter, Craig A.

2005-01-01

A new porous condition has been implemented in the PAB3D solver for simulating the flow over porous surfaces. The newly-added boundary condition is utilized to compute the flow field of a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. The flow fields for a baseline nozzle (no porosity) and for a nozzle with porous surfaces (10% porosity ratio) are computed for NPR varying from 2.01 to 9.54. Computational model results indicate that the over-expanded nozzle flow was dominated by shock-induced boundary-layer separation. Porous configurations were capable of controlling off-design separation in the nozzle by encouraging stable separation of the exhaust flow. Computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented and discussed. Computed results are in excellent agreement with experimental data.

Numerical Investigation of Flow in an Over-expanded Nozzle with Porous Surfaces

NASA Technical Reports Server (NTRS)

Abdol-Hamid, Khaled S.; Elmilingui, Alaa A.; Hunter, Craig A.

2006-01-01

A new porous condition has been implemented in the PAB3D solver for simulating the flow over porous surfaces. The newly-added boundary condition is utilized to compute the flow field of a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. The flow fields for a baseline nozzle (no porosity) and for a nozzle with porous surfaces (10% porosity ratio) are computed for NPR varying from 2.01 to 9.54. Computational model results indicate that the over-expanded nozzle flow is dominated by shock-induced boundary-layer separation. Porous configurations are capable of controlling off-design separation in the nozzle by encouraging stable separation of the exhaust flow. Computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented and discussed. Computed results are in excellent agreement with experimental data.

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

Davis, David O.

2015-01-01

Experimental investigations of specific flow phenomena, e.g., Shock Wave Boundary-Layer Interactions (SWBLI), provide great insight to the flow behavior but often lack the necessary details to be useful as CFD validation experiments. Reasons include: 1.Undefined boundary conditions Inconsistent results 2.Undocumented 3D effects (CL only measurements) 3.Lack of uncertainty analysis While there are a number of good subsonic experimental investigations that are sufficiently documented to be considered test cases for CFD and turbulence model validation, the number of supersonic and hypersonic cases is much less. This was highlighted by Settles and Dodsons [1] comprehensive review of available supersonic and hypersonic experimental studies. In all, several hundred studies were considered for their database.Of these, over a hundred were subjected to rigorous acceptance criteria. Based on their criteria, only 19 (12 supersonic, 7 hypersonic) were considered of sufficient quality to be used for validation purposes. Aeschliman and Oberkampf [2] recognized the need to develop a specific methodology for experimental studies intended specifically for validation purposes.

Application of the adjoint optimisation of shock control bump for ONERA-M6 wing

NASA Astrophysics Data System (ADS)

Nejati, A.; Mazaheri, K.

2017-11-01

This article is devoted to the numerical investigation of the shock wave/boundary layer interaction (SWBLI) as the main factor influencing the aerodynamic performance of transonic bumped airfoils and wings. The numerical analysis is conducted for the ONERA-M6 wing through a shock control bump (SCB) shape optimisation process using the adjoint optimisation method. SWBLI is analyzed for both clean and bumped airfoils and wings, and it is shown how the modified wave structure originating from upstream of the SCB reduces the wave drag, by improving the boundary layer velocity profile downstream of the shock wave. The numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm are used to find the optimum location and shape of the SCB for the ONERA-M6 airfoil and wing. Two different geometrical models are introduced for the 3D SCB, one with linear variations, and another with periodic variations. Both configurations result in drag reduction and improvement in the aerodynamic efficiency, but the periodic model is more effective. Although the three-dimensional flow structure involves much more complexities, the overall results are shown to be similar to the two-dimensional case.

Physics of magnetospheric boundary layers

NASA Technical Reports Server (NTRS)

Cairns, I. H.

1993-01-01

The central ideas of this grant are that the magnetospheric boundary layers link disparate regions of the magnetosphere together, and the global behavior of the magnetosphere can be understood only by understanding the linking mechanisms. Accordingly the present grant includes simultaneous research on the global, meso-, and micro-scale physics of the magnetosphere and its boundary layers. These boundary layers include the bow shock, magnetosheath, the plasma sheet boundary layer, and the ionosphere. Analytic, numerical and simulation projects have been performed on these subjects, as well as comparison of theoretical results with observational data. Very good progress has been made, with four papers published or in press and two additional papers submitted for publication during the six month period 1 June - 30 November 1993. At least two projects are currently being written up. In addition, members of the group have given papers at scientific meetings. The further structure of this report is as follows: section two contains brief accounts of research completed during the last six months, while section three describes the research projects intended for the grant's final period.

A comparison of measured and predicted test flow in an expansion tube with air and oxygen test gases

NASA Technical Reports Server (NTRS)

Aaggard, K. V.; Goad, W. K.

1975-01-01

Simultaneous time-resolved measurements of temperature, density, pitot pressure, and wall pressure in both air and O2 test gases were obtained in the Langley pilot model expansion tube. These tests show nonequilibrium chemical and vibrational relaxation significantly affect the test-flow condition. The use of an electromagnetic device to preopen the secondary diaphragm before the arrival of the primary shock wave resulted in an improvement in the agreement between the measured pitot pressure and the value inferred from measured density and interface velocity. Boundary-layer splitter plates used to reduce the wall boundary layer show that this disagreement in the measured and inferred pitot pressures is not a result of boundary-layer effects.

Direct numerical simulations of mack-mode damping on porous coated cones

NASA Astrophysics Data System (ADS)

Lüdeke, H.; Wartemann, V.

2013-06-01

The flow field over a 3 degree blunt cone is investigated with respect to a hypersonic stability analysis of the boundary-layer flow at Mach 6 with porous as well as smooth walls by comparing local direct numerical simulations (DNS) and linear stability theory (LST) data. The original boundary-layer profile is generated by a finite volume solver, using shock capturing techniques to generate an axisymmetric flow field. Local boundary-layer profiles are extracted from this flow field and hypersonic Mack-modes are superimposed for cone-walls with and without a porous surface used as a passive transition-reduction device. Special care is taken of curvature effects of the wall on the mode development over smooth and porous walls.

On the instability of hypersonic flow past a flat plate

NASA Technical Reports Server (NTRS)

Blackaby, Nicholas; Cowley, Stephen; Hall, Philip

1990-01-01

The instability of hypersonic boundary-layer flows over flat plates is considered. The viscosity of the fluid is taken to be governed by Sutherland's law, which gives a much more accurate representation of the temperature dependence of fluid viscosity at hypersonic speeds than Chapman's approximate linear law; although at lower speeds the temperature variation of the mean state is less pronounced so that the Chapman law can be used with some confidence. Attention is focussed on the so-called (vorticity) mode of instability of the viscous hypersonic boundary layer. This is thought to be the fastest growing inviscid disturbance at hypersonic speeds; it is also believed to have an asymptotically larger growth rate than any viscous or centrifugal instability. As a starting point the instability of the hypersonic boundary layer which exists far downstream from the leading edge of the plate is investigated. In this regime the shock that is attached to the leading edge of the plate plays no role, so that the basic boundary layer is non-interactive. It is shown that the vorticity mode of instability of this flow operates on a significantly different lengthscale than that obtained if a Chapman viscosity law is assumed. In particular, it is found that the growth rate predicted by a linear viscosity law overestimates the size of the growth rate by O(M(exp 2). Next, the development of the vorticity mode as the wavenumber decreases is described, and it is shown that acoustic modes emerge when the wavenumber has decreased from it's O(1) initial value to O(M (exp -3/2). Finally, the inviscid instability of the boundary layer near the leading edge in the interaction zone is discussed and particular attention is focussed on the strong interaction region which occurs sufficiently close to the leading edge. It is found that the vorticity mode in this regime is again unstable, and that it is concentrated in the transition layer at the edge of the boundary layer where the temperature adjusts from its large, O(M(exp 2), value in the viscous boundary layer, to its O(1) free stream value. The existence of the shock indirectly, but significantly, influences the instability problem by modifying the basic flow structure in this layer.

Factors influencing flow steadiness in laminar boundary layer shock interactions

NASA Astrophysics Data System (ADS)

Tumuklu, Ozgur; Levin, Deborah A.; Gimelshein, Sergey F.; Austin, Joanna M.

2016-11-01

The Direct Simulation Monte Carlo method has been used to model laminar shock wave boundary interactions of hypersonic flow over a 30/55-deg double-wedge and "tick-shaped" model configurations studied in the Hypervelocity Expansion Tube facility and T-ADFA free-piston shock tunnel, respectively. The impact of thermochemical effects on these interactions by changing the chemical composition from nitrogen to air as well as argon for a stagnation enthalpy of 8.0 MJ/kg flow are investigated using the 2-D wedge model. The simulations are found to reproduce many of the classic features related to Edney Type V strong shock interactions that include the attached, oblique shock formed over the first wedge, the detached bow shock from the second wedge, the separation zone, and the separation and reattachment shocks that cause complex features such as the triple point for both cases. However, results of a reacting air flow case indicate that the size of the separation length, and the movement of the triple point toward to the leading edge is much less than the nitrogen case.

Tangent-ogive nose cones

NASA Technical Reports Server (NTRS)

Wing, L. D.

1976-01-01

Program calculates aerodynamic heating and shear stresses at wall for tangent-ogive noses that are slender enough to maintain an attached nose shock during portion of flight when heat transfer from boundary layer to wall is significant.

An experimental/computational study of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5 - Experimental results

NASA Technical Reports Server (NTRS)

Rodi, Patrick E.; Dolling, David S.

1992-01-01

A combined experimental/computational study has been performed of sharp fin induced shock wave/turbulent boundary layer interactions at Mach 5. The current paper focuses on the experiments and analysis of the results. The experimental data include mean surface heat transfer, mean surface pressure distributions and surface flow visualization for fin angles of attack of 6, 8, 10, 12, 14 and 16-degrees at Mach 5 under a moderately cooled wall condition. Comparisons between the results and correlations developed earlier show that Scuderi's correlation for the upstream influence angle (recast in a conical form) is superior to other such correlations in predicting the current results, that normal Mach number based correlations for peak pressure heat transfer are adequate and that the initial heat transfer peak can be predicted using pressure-interaction theory.

Assessment of Computational Fluid Dynamics (CFD) Models for Shock Boundary-Layer Interaction

NASA Technical Reports Server (NTRS)

DeBonis, James R.; Oberkampf, William L.; Wolf, Richard T.; Orkwis, Paul D.; Turner, Mark G.; Babinsky, Holger

2011-01-01

A workshop on the computational fluid dynamics (CFD) prediction of shock boundary-layer interactions (SBLIs) was held at the 48th AIAA Aerospace Sciences Meeting. As part of the workshop numerous CFD analysts submitted solutions to four experimentally measured SBLIs. This paper describes the assessment of the CFD predictions. The assessment includes an uncertainty analysis of the experimental data, the definition of an error metric and the application of that metric to the CFD solutions. The CFD solutions provided very similar levels of error and in general it was difficult to discern clear trends in the data. For the Reynolds Averaged Navier-Stokes methods the choice of turbulence model appeared to be the largest factor in solution accuracy. Large-eddy simulation methods produced error levels similar to RANS methods but provided superior predictions of normal stresses.

Effect of a finite ionization rate on the radiative heating of outer planet atmospheric entry probes

NASA Technical Reports Server (NTRS)

Nelson, H. F.

1982-01-01

The influence of finite rate ionization in the inviscid gas just behind the stagnation shock wave on the radiative heating of probes entering the hydrogen-helium atmosphere of the major plants was investigated. Two opposing conclusions were reached as to how the ionization rate assumption affects the radiative transfer. Hydrogen-helium shock waves with a cold nonblowing wall boundary condition at the probe heat shield are emphasized. The study is limited to the stagnation shock layer.

A linear shock cell model for jets of arbitrary exit geometry

NASA Technical Reports Server (NTRS)

Morris, P. J.; Bhat, T. R. S.; Chen, G.

1989-01-01

The shock cell structures of single supersonic non-ideally expanded jets with arbitrary exit geometry are studied. Both vortex sheets and realistic mean profiles are considered for the jet shear layer. The boundary element method is used to predict the shock spacing and screech tones in a vortex sheet model of a single jet. This formulation enables the calculations to be performed only on the vortex sheet. This permits the efficient and convenient study of complicated jet geometries. Results are given for circular, elliptic and rectangular jets and the results are compared with analysis and experiment. The agreement between the predictions and measurements is very good but depends on the assumptions made to predict the geometry of the fully expanded jet. A finite diffference technique is used to examine the effect of finite mixing layer thickness for a single jet. The finite thickness of the mixing layer is found to decrease the shock spacing by approximately 20 percent over the length of the jet potential core.

Boundary-layer electron profiles for entry of a blunts slender body at high altitude

NASA Technical Reports Server (NTRS)

Evans, J. S.; Schexnayder, C. J., Jr.; Huber, P. W.

1973-01-01

New calculations of boundary-layer electron concentration profiles for entry of a blunt-nosed slender body into the earth's atmosphere are compared with previous calculations in which ambipolar diffusion was neglected. The old and new results agree in those flight regimes where ambipolar diffusion is unimportant, but large differences are noted in both peak electron concentration and profile shape at the higher altitudes, where diffusion effects are greatest. The new results are also compared with flight-measured profiles and with calculated profiles for a viscous-shock-layer theory which was recently reported in the literature. The boundary-layer results and the data agree in most respects. Differences which occur between predicted results and the data in the outer parts of the profile are discussed in terms of the effects of aerodynamic heating of the probes.

Application of Chimera Navier-Stokes Code for High Speed Flows

NASA Technical Reports Server (NTRS)

Ajmani, Kumud

1997-01-01

The primary task for this year was performed in support of the "Trailblazer" project. The purpose of the task was to perform an extensive CFD study of the shock boundary-layer interaction between the engine-diverters and the primary body surfaces of the Trailblazer vehicle. Information gathered from this study would be used to determine the effectiveness of the diverters in preventing the boundary-layer coming off of the vehicle forebody from entering the main engines. The PEGSUS code was used to define the "holes" and "boundaries" for each grid. Two sets of CFD calculations were performed.Extensive post-processing of the results was performed.

On the interaction between the shock wave attached to a wedge and freestream disturbances

NASA Technical Reports Server (NTRS)

Duck, Peter W.; Lasseigne, D. Glenn; Hussaini, M. Y.

1993-01-01

A study of the interaction of small amplitude, unsteady, freestream disturbances with a shock wave induced by a wedge in supersonic flow is presented. These disturbances may be acoustic waves, vorticity waves, or entropy waves (or indeed a combination of all three). Their interactions then generate behind the shock disturbances of all three classes, an aspect that is investigated in some detail, our motivation being to investigate possible mechanisms for boundary-layer receptivity, caused through the amplification and modification of freestream turbulence through the shock-body coupling. Also, the possibility of enhanced mixing owing to additional vorticity produced by the shock-body coupling is investigated.

Aerodynamic heating effects on wall-modeled large-eddy simulations of high-speed flows

NASA Astrophysics Data System (ADS)

Yang, Xiang; Urzay, Javier; Moin, Parviz

2017-11-01

Aerospace vehicles flying at high speeds are subject to increased wall-heating rates because of strong aerodynamic heating in the near-wall region. In wall-modeled large-eddy simulations (WMLES), this near-wall region is typically not resolved by the computational grid. As a result, the effects of aerodynamic heating need to be modeled using an LES wall model. In this investigation, WMLES of transitional and fully turbulent high-speed flows are conducted to address this issue. In particular, an equilibrium wall model is employed in high-speed turbulent Couette flows subject to different combinations of thermal boundary conditions and grid sizes, and in transitional hypersonic boundary layers interacting with incident shock waves. Specifically, the WMLES of the Couette-flow configuration demonstrate that the shear-stress and heat-flux predictions made by the wall model show only a small sensitivity to the grid resolution even in the most adverse case where aerodynamic heating prevails near the wall and generates a sharp temperature peak there. In the WMLES of shock-induced transition in boundary layers, the wall model is tested against DNS and experiments, and it is shown to capture the post-transition aerodynamic heating and the overall heat transfer rate around the shock-impingement zone. This work is supported by AFOSR.

Theoretical and experimental studies of reentry plasmas

NASA Technical Reports Server (NTRS)

Dunn, M. G.; Kang, S.

1973-01-01

A viscous shock-layer analysis was developed and used to calculate nonequilibrium-flow species distributions in the plasma layer of the RAM vehicle. The theoretical electron-density results obtained are in good agreement with those measured in flight. A circular-aperture flush-mounted antenna was used to obtain a comparison between theoretical and experimental antenna admittance in the presence of ionized boundary layers of low collision frequency. The electron-temperature and electron-density distributions in the boundary layer were independently measured. The antenna admittance was measured using a four-probe microwave reflectometer and these measured values were found to be in good agreement with those predicted. Measurements were also performed with another type of circular-aperture antenna and good agreement was obtained between the calculations and the experimental results. A theoretical analysis has been completed which permits calculation of the nonequilibrium, viscous shock-layer flow field for a sphere-cone body. Results are presented for two different bodies at several different altitudes illustrating the influences of bluntness and chemical nonequilibrium on several gas dynamic parameters of interest. Plane-wave transmission coefficients were calculated for an approximate space-shuttle body using a typical trajectory.

Thermodynamic evaluation of transonic compressor rotors using the finite volume approach

NASA Technical Reports Server (NTRS)

Nicholson, S.; Moore, J.

1986-01-01

A method was developed which calculates two-dimensional, transonic, viscous flow in ducts. The finite volume, time marching formulation is used to obtain steady flow solutions of the Reynolds-averaged form of the Navier Stokes equations. The entire calculation is performed in the physical domain. The method is currently limited to the calculation of attached flows. The features of the current method can be summarized as follows. Control volumes are chosen so that smoothing of flow properties, typically required for stability, is now needed. Different time steps are used in the different governing equations to improve the convergence speed of the viscous calculations. A new pressure interpolation scheme is introduced which improves the shock capturing ability of the method. A multi-volume method for pressure changes in the boundary layer allows calculations which use very long and thin control volumes. A special discretization technique is also used to stabilize these calculations. A special formulation of the energy equation is used to provide improved transient behavior of solutions which use the full energy equation. The method is then compared with a wide variety of test cases. The freestream Mach numbers range from 0.075 to 2.8 in the calculations. Transonic viscous flow in a converging diverging nozzle is calculated with the method; the Mach number upstream of the shock is approximately 1.25. The agreement between the calculated and measured shock strength and total pressure losses is good. Essentially incompressible turbulent boundary layer flow in a adverse pressure gradient is calculated and the computed distribution of mean velocity and shear stress are in good agreement with the measurements. At the other end of the Mach number range, a flat plate turbulent boundary layer with a freestream Mach number of 2.8 is calculated using the full energy equation; the computed total temperature distribution and recovery factor agree well with the measurements when a variable Prandtl number is used through the boundary layer.

Large Eddy Simulation of Supersonic Inlet Flows

DTIC Science & Technology

1998-04-01

shock/turbulence interaction in order to identify and explain factors important in shock/boundary layer interaction. Direct numerical simulation of a... factors : increase in the adverse pressure rise (due to pm2 increasing while pcl decreases) and decrease in streamwise momentum flux (due to pc...momentum flux. Both factors make the vortex more susceptible to breakdown. This implies that if the free-stream pressure rise exceeds the axial

Experimental studies of hypersonic shock-wave boundary-layer interactions

NASA Technical Reports Server (NTRS)

Lu, Frank K.

1992-01-01

Two classes of shock-wave boundary-layer interactions were studied experimentally in a shock tunnel in which a low Reynolds number, turbulent flow at Mach 8 was developed on a cold, flat test surface. The two classes of interactions were: (1) a swept interaction generated by a wedge ('fin') mounted perpendicularly on the flat plate; and (2) a two-dimensional, unseparated interaction induced by a shock impinging near an expansion corner. The swept interaction, with wedge angles of 5-20 degrees, was separated and there was also indication that the strongest interactions prossessed secondary separation zones. The interaction spread out extensively from the inviscid shock location although no indication of quasi-conical symmetry was evident. The surface pressure from the upstream influence to the inviscid shock was relatively low compared to the inviscid downstream value but it rose rapidly past the inviscid shock location. However, the surface pressure did not reach the downstream inviscid value and reasons were proposed for this anomalous behavior compared to strongly separated, supersonic interactions. The second class of interactions involved weak shocks impinging near small expansion corners. As a prelude to studying this interaction, a hypersonic similarity parameter was identified for the pure, expansion corner flow. The expansion corner severely damped out surface pressure fluctuations. When a shock impinged upstream of the corner, no significant changes to the surface pressure were found as compared to the case when the shock impinged on a flat plate. But, when the shock impinged downstream of the corner, a close coupling existed between the two wave systems, unlike the supersonic case. This close coupling modified the upstream influence. Regardless of whether the shock impinged ahead or behind the corner, the downstream region was affected by the close coupling between the shock and the expansion. Not only was the mean pressure distribution modified but the unsteadiness in the surface pressure was reduced compared to the flat-plate case.

New links between the Chicxulub impact structure and the Cretaceous/Tertiary boundary

USGS Publications Warehouse

Sharpton, V.L.; Dalrymple, G.B.; Marin, L.E.; Ryder, G.; Schuraytz, B.C.; Urrutia-Fucugauchi, J.

1992-01-01

THE 200-km-diameter Chicxulub structure1-3 in northern Yucatan, Mexico has emerged as the prime candidate for the Cretaceous/Tertiary (K/T) boundary impact crater3-6. Concentric geophysical anomalies associated with enigmatic occurrences of Upper Cretaceous breccias and andesitic rocks led Penfield and Camargo1 to suspect that this structure was a buried impact basin. More recently, the discovery of shocked quartz grains in a Chicxulub breccia3, and chemical similarities between Chicxulub rocks and K/T tektite-like glasses3-6 have been advanced as evidence that the Chicxulub structure is a K/T impact site. Here we present evidence from core samples that Chicxulub is indeed a K/T source crater, and can apparently account for all the evidence of impact distributed globally at the K/T boundary without the need for simultaneous multiple impacts or comet showers. Shocked breccia clasts found in the cores are similar to shocked lithic fragments found worldwide in the K/T boundary ejecta layer7,8. The Chicxulub melt rocks that we studied contain anomalously high levels of iridium (up to 13.5 parts per 109), also consistent with the indium-enriched K/T boundary layer9. Our best estimate of the crystallization age of these melt rocks, as determined by 40Ar/39Ar analyses, is 65.2??0.4 (1??) Myr, in good agreement with the mean plateau age of 64.98 ?? 0.05 Myr recently reported10. Furthermore, these melt rocks acquired a remanent magnetization indicating that they cooled during an episode of reversed geomagnetic polarity. The only such episode consistent with 40Ar/39Ar constraints is chron 29R, which includes the K/T boundary.

A Source-Term Based Boundary Layer Bleed/Effusion Model for Passive Shock Control

NASA Technical Reports Server (NTRS)

Baurle, Robert A.; Norris, Andrew T.

2011-01-01

A modeling framework for boundary layer effusion has been developed based on the use of source (or sink) terms instead of the usual practice of specifying bleed directly as a boundary condition. This framework allows the surface boundary condition (i.e. isothermal wall, adiabatic wall, slip wall, etc.) to remain unaltered in the presence of bleed. This approach also lends itself to easily permit the addition of empirical models for second order effects that are not easily accounted for by simply defining effective transpiration values. Two effusion models formulated for supersonic flows have been implemented into this framework; the Doerffer/Bohning law and the Slater formulation. These models were applied to unit problems that contain key aspects of the flow physics applicable to bleed systems designed for hypersonic air-breathing propulsion systems. The ability of each model to predict bulk bleed properties was assessed, as well as the response of the boundary layer as it passes through and downstream of a porous bleed system. The model assessment was performed with and without the presence of shock waves. Three-dimensional CFD simulations that included the geometric details of the porous plate bleed systems were also carried out to supplement the experimental data, and provide additional insights into the bleed flow physics. Overall, both bleed formulations fared well for the tests performed in this study. However, the sample of test problems considered in this effort was not large enough to permit a comprehensive validation of the models.

Analysis of Unsteady Simulations to Inform Turbulence Modeling

NASA Technical Reports Server (NTRS)

Vyas, Manan; Waindim, Mbu; Gaitonde, Datta

2016-01-01

In this work, budgets of the turbulent kinetic energy are presented for a two-dimensional shock wave boundary-layer interaction (SBLI). The work should be of interest to the SBLI research and turbulence modeling community.

Dinosaurs, spherules, and the “magic” layer: A new K-T boundary clay site in Wyoming

NASA Astrophysics Data System (ADS)

Bohor, Bruce F.; Triplehorn, Don M.; Nichols, Douglas J.; Millard, Hugh T., Jr.

1987-10-01

A new Cretaceous-Tertiary (K-T) boundary clay site has been found along Dogie Creek in Wyoming in the drainage of Lance Creek—the type area of the Lance Formation of latest Cretaceous age. The boundary clay was discovered in the uppermost part of the Lance Formation, 4 7 cm beneath the lowest lignite in the Paleocene Fort Union Formation and approximately 1 m above a fragmented dinosaur bone. The boundary clay consists of a basal kaolinitic claystone layer as much as 3 cm thick containing hollow goyazite spherules, overlain by a 2 3 mm smectitic layer (the “magic” layer) containing both shock-metamorphosed minerals and an iridium anomaly of 21 ppb. A palynological break coincides with the base of the claystone layer; numerous Late Cretaceous palynomorph species terminate at this boundary. The paleontological significance of this new boundary site lies in its close association with the well-studied assemblage of dinosaurs and other vertebrates and flora within the type area of the Lance Formation. The spherules at the Dogie Creek site are extremely well preserved by virtue of their replacement by the mineral goyazite. This preservation should facilitate the resolution of the origin of the spherules and of their host layer.

Aerothermodynamic environment of a Titan aerocapture vehicle

NASA Technical Reports Server (NTRS)

Tiwari, S. N.; Chow, H.

1982-01-01

The extent of convective and radiative heating for a Titan aerocapture vehicle is investigated. The flow in the shock layer is assumed to be axisymmetric, steady, viscous, and compressible. It is further assumed that the gas is in chemical and local thermodynamic equilibrium and tangent slab approximation is used for the radiative transport. The effect of the slip boundary conditions on the body surface and at the shock wave are included in the analysis of high-altitude entry conditions. The implicit finite difference techniques is used to solve the viscous shock-layer equations for a 45 degree sphere cone at zero angle of attack. Different compositions for the Titan atmosphere are assumed, and results are obtained for the entry conditions specified by the Jet Propulsion Laboratory.

Towards Perfectly Absorbing Boundary Conditions for Euler Equations

NASA Technical Reports Server (NTRS)

Hayder, M. Ehtesham; Hu, Fang Q.; Hussaini, M. Yousuff

1997-01-01

In this paper, we examine the effectiveness of absorbing layers as non-reflecting computational boundaries for the Euler equations. The absorbing-layer equations are simply obtained by splitting the governing equations in the coordinate directions and introducing absorption coefficients in each split equation. This methodology is similar to that used by Berenger for the numerical solutions of Maxwell's equations. Specifically, we apply this methodology to three physical problems shock-vortex interactions, a plane free shear flow and an axisymmetric jet- with emphasis on acoustic wave propagation. Our numerical results indicate that the use of absorbing layers effectively minimizes numerical reflection in all three problems considered.

Bubble dynamics in a compressible liquid in contact with a rigid boundary

PubMed Central

Wang, Qianxi; Liu, Wenke; Zhang, A. M.; Sui, Yi

2015-01-01

A bubble initiated near a rigid boundary may be almost in contact with the boundary because of its expansion and migration to the boundary, where a thin layer of water forms between the bubble and the boundary thereafter. This phenomenon is modelled using the weakly compressible theory coupled with the boundary integral method. The wall effects are modelled using the imaging method. The numerical instabilities caused by the near contact of the bubble surface with the boundary are handled by removing a thin layer of water between them and joining the bubble surface with its image to the boundary. Our computations correlate well with experiments for both the first and second cycles of oscillation. The time history of the energy of a bubble system follows a step function, reducing rapidly and significantly because of emission of shock waves at inception of a bubble and at the end of collapse but remaining approximately constant for the rest of the time. The bubble starts being in near contact with the boundary during the first cycle of oscillation when the dimensionless stand-off distance γ = s/Rm < 1, where s is the distance of the initial bubble centre from the boundary and Rm is the maximum bubble radius. This leads to (i) the direct impact of a high-speed liquid jet on the boundary once it penetrates through the bubble, (ii) the direct contact of the bubble at high temperature and high pressure with the boundary, and (iii) the direct impingement of shock waves on the boundary once emitted. These phenomena have clear potential to damage the boundary, which are believed to be part of the mechanisms of cavitation damage. PMID:26442148

Experimental Investigation of the Effects of Viscosity on the Drag and Base Pressure of Bodies of Revolution at a Mach Number 1.5

NASA Technical Reports Server (NTRS)

Chapman, Dean R; Perkins, Edward W

1951-01-01

Models were tested to evaluate effects of Reynolds number for both laminar and turbulent boundary layers. Principal geometric variables investigated were afterbody shape and length-diameter ratio. Force tests and base-pressure measurements were made. Schlieren photographs were used to analyze the effects of viscosity on flow separation and shock-wave configuration and to verify the condition of the boundary layer as deduced from the force tests. The results are discussed and compared with theoretical calculations.

Validation of a three-dimensional viscous analysis of axisymmetric supersonic inlet flow fields

NASA Technical Reports Server (NTRS)

Benson, T. J.; Anderson, B. H.

1983-01-01

A three-dimensional viscous marching analysis for supersonic inlets was developed. To verify this analysis several benchmark axisymmetric test configurations were studied and are compared to experimental data. Detailed two-dimensional results for shock-boundary layer interactions are presented for flows with and without boundary layer bleed. Three dimensional calculations of a cone at angle of attack and a full inlet at attack are also discussed and evaluated. Results of the calculations demonstrate the code's ability to predict complex flow fields and establish guidelines for future calculations using similar codes.

A review of high-speed, convective, heat-transfer computation methods

NASA Technical Reports Server (NTRS)

Tauber, Michael E.

1989-01-01

The objective of this report is to provide useful engineering formulations and to instill a modest degree of physical understanding of the phenomena governing convective aerodynamic heating at high flight speeds. Some physical insight is not only essential to the application of the information presented here, but also to the effective use of computer codes which may be available to the reader. A discussion is given of cold-wall, laminar boundary layer heating. A brief presentation of the complex boundary layer transition phenomenon follows. Next, cold-wall turbulent boundary layer heating is discussed. This topic is followed by a brief coverage of separated flow-region and shock-interaction heating. A review of heat protection methods follows, including the influence of mass addition on laminar and turbulent boundary layers. Also discussed are a discussion of finite-difference computer codes and a comparison of some results from these codes. An extensive list of references is also provided from sources such as the various AIAA journals and NASA reports which are available in the open literature.

A review of high-speed, convective, heat-transfer computation methods

NASA Technical Reports Server (NTRS)

Tauber, Michael E.

1989-01-01

The objective is to provide useful engineering formulations and to instill a modest degree of physical understanding of the phenomena governing convective aerodynamic heating at high flight speeds. Some physical insight is not only essential to the application of the information presented here, but also to the effective use of computer codes which may be available to the reader. Given first is a discussion of cold-wall, laminar boundary layer heating. A brief presentation of the complex boundary layer transition phenomenon follows. Next, cold-wall turbulent boundary layer heating is discussed. This topic is followed by a brief coverage of separated flow-region and shock-interaction heating. A review of heat protection methods follows, including the influence of mass addition on laminar and turbulent boundary layers. Next is a discussion of finite-difference computer codes and a comparison of some results from these codes. An extensive list of references is also provided from sources such as the various AIAA journals and NASA reports which are available in the open literature.

Comparison of secondary flows and boundary-layer accumulations in several turbine nozzles

NASA Technical Reports Server (NTRS)

Kofskey, Milton G; Allen, Hubert W; Herzig, Howard Z

1953-01-01

An investigation was made of losses and secondary flows in three different turbine nozzle configurations in annular cascade. Appreciable outer shroud loss cores (passage vortices) were found to exist at the discharge of blades which had thickened suction surface boundary layers near the outer shroud. Blade designs having thinner boundary layers did not show such outer shroud loss cores, but indicated greater inward radial flow of low momentum air, in the wake loss is to this extent an indication of the presence or absence of radial flow. The blade wake was a combination of profile loss and low momentum air from the outer shroud, and the magnitude of the wake loss is to this extent an indication of the presence or absence of radial flow. At a high Mach number, shock-boundary-layer thickening on the blade suction surfaces provided an additional radial flow path for low momentum air, which resulted in large inner shroud loss regions accompanied by large deviations from design values of discharge angle. (author)

Shock-darkening in ordinary chondrites: Determination of the pressure-temperature conditions by shock physics mesoscale modeling

NASA Astrophysics Data System (ADS)

Moreau, J.; Kohout, T.; Wünnemann, K.

2017-11-01

We determined the shock-darkening pressure range in ordinary chondrites using the iSALE shock physics code. We simulated planar shock waves on a mesoscale in a sample layer at different nominal pressures. Iron and troilite grains were resolved in a porous olivine matrix in the sample layer. We used equations of state (Tillotson EoS and ANEOS) and basic strength and thermal properties to describe the material phases. We used Lagrangian tracers to record the peak shock pressures in each material unit. The post-shock temperatures (and the fractions of the tracers experiencing temperatures above the melting point) for each material were estimated after the passage of the shock wave and after the reflections of the shock at grain boundaries in the heterogeneous materials. The results showed that shock-darkening, associated with troilite melt and the onset of olivine melt, happened between 40 and 50 GPa with 52 GPa being the pressure at which all tracers in the troilite material reach the melting point. We demonstrate the difficulties of shock heating in iron and also the importance of porosity. Material impedances, grain shapes, and the porosity models available in the iSALE code are discussed. We also discuss possible not-shock-related triggers for iron melt.

Structure of the reconnection layer and the associated slow shocks: Two-dimensional simulations of a Riemann problem

NASA Astrophysics Data System (ADS)

Cremer, Michael; Scholer, Manfred

2000-12-01

The kinetic structure of the reconnection layer in the magnetotail is investigated by two-dimensional hybrid simulations. As a proxy, the solution of the Riemann problem of the collapse of a current sheet with a normal magnetic field component is considered for two cases of the plasma beta (particle to magnetic field pressure): β=0.02 and β=0.002. The collapse results in an expanding layer of compressed and heated plasma, which is accelerated up to the Alfvén speed vA. The boundary layer separating this hot reconnection like layer from the cold lobe plasma is characterized by a beam of back-streaming ions with a field-aligned bulk speed of ~=2vA relative to the cold lobe ion population at rest. As a consequence, obliquely propagating waves are excited via the electromagnetic ion/ion cyclotron instability, which led to perpendicular heating of the ions in the boundary layer as well as further outside the layer in the lobe. In both regions, waves are found which propagate almost parallel to the magnetic field and which are identified as Alfvén ion cyclotron (AIC) waves. These waves are excited by the temperature anisotropy instability. The temperature anisotropy increases with decreasing plasma beta. Thus the anisotropy threshold of the instability is exceeded even in the case of a rather small beta value. The AIC waves, when convected downstream of what can be defined as the the slow shock, make an important contribution to the ion thermalization process. More detailed information on the dissipation process in the slow shocks is gained by analyzing individual ion trajectories.

Effects of bleed-hole geometry and plenum pressure on three-dimensional shock-wave/boundary-layer/bleed interactions

NASA Technical Reports Server (NTRS)

Chyu, Wei J.; Rimlinger, Mark J.; Shih, Tom I.-P.

1993-01-01

A numerical study was performed to investigate 3D shock-wave/boundary-layer interactions on a flat plate with bleed through one or more circular holes that vent into a plenum. This study was focused on how bleed-hole geometry and pressure ratio across bleed holes affect the bleed rate and the physics of the flow in the vicinity of the holes. The aspects of the bleed-hole geometry investigated include angle of bleed hole and the number of bleed holes. The plenum/freestream pressure ratios investigated range from 0.3 to 1.7. This study is based on the ensemble-averaged, 'full compressible' Navier-Stokes (N-S) equations closed by the Baldwin-Lomax algebraic turbulence model. Solutions to the ensemble-averaged N-S equations were obtained by an implicit finite-volume method using the partially-split, two-factored algorithm of Steger on an overlapping Chimera grid.

Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Numbers

NASA Technical Reports Server (NTRS)

Xiao, X.; Edwards, J. R.; Hassan, H. A.; Gaffney, R. L., Jr.

2007-01-01

A new turbulence model suited for calculating the turbulent Prandtl number as part of the solution is presented. The model is based on a set of two equations: one governing the variance of the enthalpy and the other governing its dissipation rate. These equations were derived from the exact energy equation and thus take into consideration compressibility and dissipation terms. The model is used to study two cases involving shock wave/boundary layer interaction at Mach 9.22 and Mach 5.0. In general, heat transfer prediction showed great improvement over traditional turbulence models where the turbulent Prandtl number is assumed constant. It is concluded that using a model that calculates the turbulent Prandtl number as part of the solution is the key to bridging the gap between theory and experiment for flows dominated by shock wave/boundary layer interactions.

Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Numbers

NASA Technical Reports Server (NTRS)

Gaffney, R. L., Jr.; Xiao, X.; Edwards, J. R.; Hassan, H. A.

2005-01-01

A new turbulence model suited for calculating the turbulent Prandtl number as part of the solution is presented. The model is based on a set of two equations: one governing the variance of the enthalpy and the other governing its dissipation rate. These equations were derived from the exact energy equation and thus take into consideration compressibility and dissipation terms. The model is used to study two cases involving shock wave/boundary layer interaction at Mach 9.22 and Mach 5.0. In general, heat transfer prediction showed great improvement over traditional turbulence models where the turbulent Prandtl number is assumed constant. It is concluded that using a model that calculates the turbulent Prandtl number as part of the solution is the key to bridging the gap between theory and experiment for flows dominated by shock wave/boundary layer interactions.

Investigation to optimize the passive shock wave-boundary layer control for supercritical airfoil drag reduction

NASA Technical Reports Server (NTRS)

Nagamatsu, H. T.; Ficarra, R.; Orozco, R.

1983-01-01

The optimization of passive shock wave/boundary layer control for supercritical airfoil drag reduction was investigated in a 3 in. x 15.4 in. Transonic Blowdown Wind Tunnel. A 14% thick supercritical airfoil was tested with 0%, 1.42% and 2.8% porosities at Mach numbers of .70 to .83. The 1.42% case incorporated a linear increase in porosity with the flow direction while the 2.8% case was uniform porosity. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid-surface airfoils. While the results show that linear 1.42% porosity actually led to a slight increase in drag it was found that the uniform 2.8% porosity can lead to a drag reduction of 46% at M = .81.

Interaction of aerodynamic noise with laminar boundary layers in supersonic wind tunnels

NASA Technical Reports Server (NTRS)

Schopper, M. R.

1984-01-01

The interaction between incoming aerodynamic noise and the supersonic laminar boundary layer is studied. The noise field is modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is staffed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. Some consideration is given to the possible appreciable thickness of the weak shock fronts. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level. The situation regarding experimental evidence of the caustic region is reviewed.

Numerical simulations of the flow in the HYPULSE expansion tube

NASA Technical Reports Server (NTRS)

Wilson, Gregory J.; Sussman, Myles A.; Bakos, Robert J.

1995-01-01

Axisymmetric numerical simulations with finite-rate chemistry are presented for two operating conditions in the HYPULSE expansion tube. The operating gas for these two cases is nitrogen and the computations are compared to experimental data. One test condition is at a total enthalpy of 15.2 MJ/Kg and a relatively low static pressure of 2 kPa. This case is characterized by a laminar boundary layer and significant chemical nonequilibrium in the acceleration gas. The second test condition is at a total enthalpy of 10.2 MJ/Kg and a static pressure of 38 kPa and is characterized by a turbulent boundary layer. For both cases, the time-varying test gas pressure predicted by the simulations is in good agreement with experimental data. The computations are also found to be in good agreement with Mirels' correlations for shock tube flow. It is shown that the nonuniformity of the test gas observed in the HYPULSE expansion tube is strongly linked to the boundary layer thickness. The turbulent flow investigated has a larger boundary layer and greater test gas nonuniformity. In order to investigate possibilities of improving expansion tube flow quality by reducing the boundary layer thickness, parametric studies showing the effect of density and turbulent transition point on the test conditions are also presented. Although an increase in the expansion tube operating pressure level would reduce the boundary layer thickness, the simulations indicate that the reduction would be less than what is predicted by flat plate boundary layer correlations.

An efficient solution technique for shockwave-boundary layer interactions with flow separation and slot suction effects

NASA Technical Reports Server (NTRS)

Edwards, Jack R.; Mcrae, D. Scott

1991-01-01

An efficient method for computing two-dimensional compressible Navier-Stokes flow fields is presented. The solution algorithm is a fully-implicit approximate factorization technique based on an unsymmetric line Gauss-Seidel splitting of the equation system Jacobian matrix. Convergence characteristics are improved by the addition of acceleration techniques based on Shamanskii's method for nonlinear equations and Broyden's quasi-Newton update. Characteristic-based differencing of the equations is provided by means of Van Leer's flux vector splitting. In this investigation, emphasis is placed on the fast and accurate computation of shock-wave-boundary layer interactions with and without slot suction effects. In the latter context, a set of numerical boundary conditions for simulating the transpiration flow in an open slot is devised. Both laminar and turbulent cases are considered, with turbulent closure provided by a modified Cebeci-Smith algebraic model. Comparisons with computational and experimental data sets are presented for a variety of interactions, and a fully-coupled simulation of a plenum chamber/inlet flowfield with shock interaction and suction is also shown and discussed.

Structural Response Prediction: Full-field, Dynamic Pressure and Displacement Measurements of a Panel Excited by Shock Boundary-layer Interaction

DTIC Science & Technology

2015-02-01

research cell14. The RC-19 facility is a continuous flow wind tunnel designed to study the mechanisms that govern the mixing and combustion process... angle of 39° from the tunnel bottom wall. The shock generator can translate 170 mm in the flow direction to allow for the shock wave to impinge from...approximate absolute pressure of 20.5 kPa. A series of “ wind -off” images for PSP were collected at that time. The tunnel was then started by setting the

Experimental Investigation of the Application of Microramp Flow Control to an Oblique Shock Interaction

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.; Anderson, Bernhard H.

2009-01-01

The effectiveness of microramp flow control devices in controlling an oblique shock interaction was tested in the 15- by 15-Centimeter Supersonic Wind Tunnel at NASA Glenn Research Center. Fifteen microramp geometries were tested varying the height, chord length, and spacing between ramps. Measurements of the boundary layer properties downstream of the shock reflection were analyzed using design of experiments methods. Results from main effects, D-optimal, full factorial, and central composite designs were compared. The designs provided consistent results for a single variable optimization.

Numerical Study of Boundary Layer Interaction with Shocks: Method Improvement and Test Computation

NASA Technical Reports Server (NTRS)

Adams, N. A.

1995-01-01

The objective is the development of a high-order and high-resolution method for the direct numerical simulation of shock turbulent-boundary-layer interaction. Details concerning the spatial discretization of the convective terms can be found in Adams and Shariff (1995). The computer code based on this method as introduced in Adams (1994) was formulated in Cartesian coordinates and thus has been limited to simple rectangular domains. For more general two-dimensional geometries, as a compression corner, an extension to generalized coordinates is necessary. To keep the requirements or limitations for grid generation low, the extended formulation should allow for non-orthogonal grids. Still, for simplicity and cost efficiency, periodicity can be assumed in one cross-flow direction. For easy vectorization, the compact-ENO coupling algorithm as used in Adams (1994) treated whole planes normal to the derivative direction with the ENO scheme whenever at least one point of this plane satisfied the detection criterion. This is apparently too restrictive for more general geometries and more complex shock patterns. Here we introduce a localized compact-ENO coupling algorithm, which is efficient as long as the overall number of grid points treated by the ENO scheme is small compared to the total number of grid points. Validation and test computations with the final code are performed to assess the efficiency and suitability of the computer code for the problems of interest. We define a set of parameters where a direct numerical simulation of a turbulent boundary layer along a compression corner with reasonably fine resolution is affordable.

Proximal Cretaceous-Tertiary boundary impact deposits in the Caribbean

NASA Technical Reports Server (NTRS)

Hildebrand, Alan R.; Boynton, Willam V.

1990-01-01

Trace element, isotopic, and mineralogic studies indicate that the proposed impact at the Cretaceous-Tertiary boundary occurred in an ocean basin, although a minor component of continental material is required. The size and abundance of shocked minerals and the restricted geographic occurrence of the ejecta layer and impact-wave deposits suggest an impact between the Americas. Coarse boundary sediments at sites 151 and 153 in the Colombian Basin and 5- to 450-meter-thick boundary sediments in Cuba may be deposits of a giant wave produced by a nearby oceanic impact.

Overview of NASA GRC's Efforts In SWBLI

NASA Technical Reports Server (NTRS)

Long-Davis, Mary Jo

2017-01-01

Overview of NASA Efforts (related to SWBLI research) An overview of NASA's restructured ARMD Program and the resulting new projects. Areas of research pertaining to shock wave boundary layer interaction are highlighted. Plans and status for specific tasks are presented.

Experimental studies on flow visualization and velocity field of compression ramp with different incoming boundary layers

NASA Astrophysics Data System (ADS)

Wu, Yu; Yi, Shi-He; He, Lin; Chen, Zhi; Zhu, Yang-Zhu

2014-11-01

Experimental studies which focus on flow visualization and the velocity field of a supersonic laminar/turbulent flow over a compression ramp were carried out in a Mach 3.0 wind tunnel. Fine flow structures and velocity field structures were obtained via NPLS (nanoparticle-tracer planar laser scattering) and PIV (particle image velocimetry) techniques, time-averaged flow structures were researched, and spatiotemporal evolutions of transient flow structures were analyzed. The flow visualization results indicated that when the ramp angles were 25°, a typical separation occurred in the laminar flow, some typical flow structures such as shock induced by the boundary layer, separation shock, reversed flow and reattachment shock were visible clearly. While a certain extent separation occurred in turbulent flow, the separation region was much smaller. When the ramp angles were 28°, laminar flow separated further, and the separation region expanded evidently, flow structures in the separation region were complex. While a typical separation occurred in turbulent flow, reversed flow structures were significant, flow structures in the separation region were relatively simple. The experimental results of velocity field were corresponding to flow visualization, and the velocity field structures of both compression ramp flows agreed with the flow structures well. There were three layered structures in the U component velocity, and the V component velocity appeared like an oblique “v”. Some differences between these two compression ramp flows can be observed in the velocity profiles of the shear layer and the shearing intensity.

Version 2 of the Protuberance Correlations for the Shuttle-Orbiter Boundary Layer Transition Tool

NASA Technical Reports Server (NTRS)

King, Rudolph A.; Kegerise, Michael A.; Berry, Scott A.

2009-01-01

Orbiter-specific transition data, acquired in four ground-based facilities (LaRC 20-Inch Mach 6 Air Tunnel, LaRC 31-Inch Mach 10 Air Tunnel, LaRC 20-Inch Mach 6 CF4 Tunnel, and CUBRC LENS-I Shock Tunnel) with three wind tunnel model scales (0.75, 0.90, and 1.8%) and from Orbiter historical flight data, have been analyzed to improve a pre-existing engineering tool for reentry transition prediction on the windward side of the Orbiter. Boundary layer transition (BLT) engineering correlations for transition induced by isolated protuberances are presented using a laminar Navier-Stokes (N-S) database to provide the relevant boundary-layer properties. It is demonstrated that the earlier version of the BLT correlation that had been developed using parameters derived from an engineering boundary-layer code has improved data collapse when developed with the N-S database. Of the new correlations examined, the proposed correlation 5, based on boundary-layer edge and wall properties, was found to provide the best overall correlation metrics when the entire database is employed. The second independent correlation (proposed correlation 7) selected is based on properties within the boundary layer at the protuberance height. The Aeroheating Panel selected a process to derive the recommended coefficients for Version 2 of the BLT Tool. The assumptions and limitations of the recommended protuberance BLT Tool V.2 are presented.

Turbulence Modeling Validation, Testing, and Development

NASA Technical Reports Server (NTRS)

Bardina, J. E.; Huang, P. G.; Coakley, T. J.

1997-01-01

The primary objective of this work is to provide accurate numerical solutions for selected flow fields and to compare and evaluate the performance of selected turbulence models with experimental results. Four popular turbulence models have been tested and validated against experimental data often turbulent flows. The models are: (1) the two-equation k-epsilon model of Wilcox, (2) the two-equation k-epsilon model of Launder and Sharma, (3) the two-equation k-omega/k-epsilon SST model of Menter, and (4) the one-equation model of Spalart and Allmaras. The flows investigated are five free shear flows consisting of a mixing layer, a round jet, a plane jet, a plane wake, and a compressible mixing layer; and five boundary layer flows consisting of an incompressible flat plate, a Mach 5 adiabatic flat plate, a separated boundary layer, an axisymmetric shock-wave/boundary layer interaction, and an RAE 2822 transonic airfoil. The experimental data for these flows are well established and have been extensively used in model developments. The results are shown in the following four sections: Part A describes the equations of motion and boundary conditions; Part B describes the model equations, constants, parameters, boundary conditions, and numerical implementation; and Parts C and D describe the experimental data and the performance of the models in the free-shear flows and the boundary layer flows, respectively.

Fires at the K/T boundary - Carbon at the Sumbar, Turkmenia, site

NASA Technical Reports Server (NTRS)

Wolbach, Wendy S.; Anders, Edward; Nazarov, Michael A.

1990-01-01

Results are reported on carbon analysis and on C and Ir correlations in samples from the marine K-T boundary site SM-4 at the Sumbar River in Turkmenia (USSR), which has the largest known Ir anomaly (580 ng/cq cm). In addition, the boundary clay is thick, and is undisturbed by bioturbation. Kerogen and delta-C-13 elemental carbon in the boundary clay were resolved using a Cr2O7(2-) oxidation method of Wolbach and Anders (1989). It was found that Ir and shocked quartz, both representing impact ejecta, rise sharply at the boundary, peak in the basal layer, and then decline. On the other hand, soot and total elemental C show a similar spike in the basal layer but then rise rather than fall, peking at 7 cm. Results indicate that fires at the SM-4 K-T boundary site started before the basal layer had settled, implying that ignition and spreading of major fires became possible at the time of or very soon after the meteorite impact.

Dilatation-dissipation corrections for advanced turbulence models

NASA Technical Reports Server (NTRS)

Wilcox, David C.

1992-01-01

This paper analyzes dilatation-dissipation based compressibility corrections for advanced turbulence models. Numerical computations verify that the dilatation-dissipation corrections devised by Sarkar and Zeman greatly improve both the k-omega and k-epsilon model predicted effect of Mach number on spreading rate. However, computations with the k-gamma model also show that the Sarkar/Zeman terms cause an undesired reduction in skin friction for the compressible flat-plate boundary layer. A perturbation solution for the compressible wall layer shows that the Sarkar and Zeman terms reduce the effective von Karman constant in the law of the wall. This is the source of the inaccurate k-gamma model skin-friction predictions for the flat-plate boundary layer. The perturbation solution also shows that the k-epsilon model has an inherent flaw for compressible boundary layers that is not compensated for by the dilatation-dissipation corrections. A compressibility modification for k-gamma and k-epsilon models is proposed that is similar to those of Sarkar and Zeman. The new compressibility term permits accurate predictions for the compressible mixing layer, flat-plate boundary layer, and a shock separated flow with the same values for all closure coefficients.

Stabilizing Effect of Sweep on Low-Frequency STBLI Unsteadiness

NASA Astrophysics Data System (ADS)

Adler, Michael; Gaitonde, Datta

2017-11-01

A Large-Eddy Simulation database is generated to examine unsteady shock/turbulent boundary-layer-interaction (STBLI) mechanisms in a Mach 2 swept-compression-corner. Such interactions exhibit open separation, with separation relief from the sweep, and lack the closed mean recirculation found in spanwise-homogeneous STBLIs. We find that the swept interaction lacks the low-frequency coherent shock unsteadiness, two-decades below incoming turbulent boundary layer scales, that is a principal feature of comparable closed separation STBLIs. Rather, the prominent unsteady content is a mid-frequency regime that develops in the separated shear layer and scales weakly with the local separation length. Additionally, a linear perturbation analysis of the unsteady flow indicates that the feedback pathway (associated with an absolute instability in spanwise-homogeneous interactions) is absent in swept-compression-corner interactions. This suggests that 1) the linear oscillator is an essential component of low-frequency unsteadiness in interactions with closed separation. 2) Low-frequency control efforts should be focused on disrupting this oscillator. 3) Introduction of 3D effects constitute one mechanism to disrupt the oscillator.

Turbulence modeling for hypersonic flows

NASA Technical Reports Server (NTRS)

Marvin, J. G.; Coakley, T. J.

1989-01-01

Turbulence modeling for high speed compressible flows is described and discussed. Starting with the compressible Navier-Stokes equations, methods of statistical averaging are described by means of which the Reynolds-averaged Navier-Stokes equations are developed. Unknown averages in these equations are approximated using various closure concepts. Zero-, one-, and two-equation eddy viscosity models, algebraic stress models and Reynolds stress transport models are discussed. Computations of supersonic and hypersonic flows obtained using several of the models are discussed and compared with experimental results. Specific examples include attached boundary layer flows, shock wave boundary layer interactions and compressible shear layers. From these examples, conclusions regarding the status of modeling and recommendations for future studies are discussed.

Numerical Prediction Methods (Reynolds-Averaged Navier-Stokes Simulations of Transonic Separated Flows)

NASA Technical Reports Server (NTRS)

Mehta, Unmeel; Lomax, Harvard

1981-01-01

During the past five years, numerous pioneering archival publications have appeared that have presented computer solutions of the mass-weighted, time-averaged Navier-Stokes equations for transonic problems pertinent to the aircraft industry. These solutions have been pathfinders of developments that could evolve into a major new technological capability, namely the computational Navier-Stokes technology, for the aircraft industry. So far these simulations have demonstrated that computational techniques, and computer capabilities have advanced to the point where it is possible to solve forms of the Navier-Stokes equations for transonic research problems. At present there are two major shortcomings of the technology: limited computer speed and memory, and difficulties in turbulence modelling and in computation of complex three-dimensional geometries. These limitations and difficulties are the pacing items of the continuing developments, although the one item that will most likely turn out to be the most crucial to the progress of this technology is turbulence modelling. The objective of this presentation is to discuss the state of the art of this technology and suggest possible future areas of research. We now discuss some of the flow conditions for which the Navier-Stokes equations appear to be required. On an airfoil there are four different types of interaction of a shock wave with a boundary layer: (1) shock-boundary-layer interaction with no separation, (2) shock-induced turbulent separation with immediate reattachment (we refer to this as a shock-induced separation bubble), (3) shock-induced turbulent separation without reattachment, and (4) shock-induced separation bubble with trailing edge separation.

Direct numerical simulation of shockwave and turbulent boundary layer interactions

NASA Astrophysics Data System (ADS)

Wu, Minwei

Direct numerical simulations (DNS) of a shockwave/turbulent boundary layer interaction (STBLI) at Mach number 3 and Reynolds number based on the momentum thickness of 2300 are performed. A 4th-order accurate, bandwidth-optimized weighted-essentially-non-oscillatory (WENO) scheme is used and the method is found to be too dissipative for the STBLI simulation due to the over-adaptation properties of this original WENO scheme. In turn, a relative limiter is introduced to mitigate the problem. Tests on the Shu-Osher problem show that the modified WENO scheme decreases the numerical dissipation significantly. By utilizing a combination of the relative limiter and the absolute limiter described by Jiang & Shu [32], the DNS results are improved further. The DNS data agree well with the reference experiments of Bookey et al. [10] in the size of the separation bubble, the separation and reattachment point, the mean wall-pressure distribution, and the velocity profiles both upstream and downstream of the interaction region. The DNS data show that velocity profiles change dramatically along the streamwise direction. Downstream of the interaction, the velocity profiles show a characteristic "dip" in the logarithmic region, as shown by the experiments of Smits & Muck [66] at much higher Reynolds number. In the separation region, the velocity profiles are found to resemble those of a laminar flow, yet the flow does not fully relaminarize. The mass-flux turbulence intensity is amplified by a factor of about 5 throughout the interaction, which is consistent with that found in higher Reynolds experiments of Selig et al. [52]. All Reynolds stress components are greatly amplified by the interaction. Assuming that the ow is still two dimensional downstream of the interaction, the components rhou'u', rhov'v', rho w'w', and rho u'w' are amplified by factors of 6, 6, 12, and 24, respectively, where u is the streamwise and w is the wall-normal velocity. However, analyses of the turbulence structure show that the ow is not uniform in the spanwise direction downstream of the interaction. A pair of counter-rotating vortices is observed in streamwise-wall-normal planes in the mean ow downstream of the ramp corner. Taking the three-dimensionality into account, the amplification factors of the Reynolds stresses are greatly decreased. The component rhou'w' is amplified by a factor of about 10, which is comparable to that found in the experiments of Smits & Muck [66]. Strong Reynolds analogy (SRA) relations are also studied using the DNS data. The SRA is found to hold in the incoming boundary layer of the DNS. However, inside and downstream of the interaction region, the SRA relations are not satisfied. From the DNS analyses, the shock motion is characterized by a low frequency component (of order 0.01Uinfinity/delta). In addition, the motion of the shock is found to have two aspects: a spanwise wrinkling motion and a streamwise oscillatory motion. The spanwise wrinkling is observed to be a local feature with high frequencies (of order Uinfinity /delta). Two-point correlations reveal that the spanwise wrinkling is closely related to the low momentum motions in the incoming boundary layer as they convect through the shock. The low frequency shock motion is found to be a streamwise oscillation motion. Conditional statistics show that there is no significant difference in the mean properties of the incoming boundary layer when the shock is at an upstream or downstream location. However, analyses of the unsteadiness of the separation bubble reveal that the low frequency shock motion is driven by the downstream flow.

Influence of shock waves from plasma actuators on transonic and supersonic airflow

NASA Astrophysics Data System (ADS)

Mursenkova, I. V.; Znamenskaya, I. A.; Lutsky, A. E.

2018-03-01

This paper presents experimental and numerical investigations of high-current sliding surface discharges of nanosecond duration and their effect on high-speed flow as plasma actuators in a shock tube. This study deals with the effectiveness of a sliding surface discharge at low and medium air pressure. Results cover the electrical characteristics of the discharge and optical visualization of the discharge and high-speed post-discharge flow. A sliding surface discharge is first studied in quiescent air conditions and then in high-speed flow, being initiated in the boundary layer at a transverse flow velocity of 50-950 m s-1 behind a flat shock wave in air of density 0.04-0.45 kg m-3. The discharge is powered by a pulse voltage of 25-30 kV and the electric current is ~0.5 kA. Shadow imaging and particle image velocimetry (PIV) are used to measure the flow field parameters after the pulse surface discharge. Shadow imaging reveals shock waves originating from the channels of the discharge configurations. PIV is used to measure the velocity field resulting from the discharge in quiescent air and to determine the homogeneity of energy release along the sliding discharge channel. Semicylindrical shock waves from the channels of the sliding discharge have an initial velocity of more than 600 m s-1. The shock-wave configuration floats in the flow along the streamlined surface. Numerical simulation based on the equations of hydrodynamics matched with the experiment showed that 25%-50% of the discharge energy is instantly transformed into heat energy in a high-speed airflow, leading to the formation of shock waves. This energy is comparable to the flow enthalpy and can result in significant modification of the boundary layer and the entire flow.

Experimental Investigation of Three-Dimensional Shock Wave Turbulent Boundary Layer Interaction: An Exploratory Study of Blunt Fin-Induced Flows.

DTIC Science & Technology

1980-03-01

distributions could be obtained. The pressure tappings were sampled using two computer controlled 48 port Model 48J4 Scanivalves equipped with Druck ...the boundary layer becomes turbulent, the upstream in- fluence drops to between 2 and 3D . 3.2 Pressure Distributions Off the Plane of Symmetry 3.2.1...upstream influence varies between 0.3 cm (0.12") and 7.6 cm (3.0"), a ratio of about 25, yet in terms of D , Iu lies between 2 and 3D . The figure shows

EIDOSCOPE: particle acceleration at plasma boundaries

NASA Astrophysics Data System (ADS)

Vaivads, A.; Andersson, G.; Bale, S. D.; Cully, C. M.; De Keyser, J.; Fujimoto, M.; Grahn, S.; Haaland, S.; Ji, H.; Khotyaintsev, Yu. V.; Lazarian, A.; Lavraud, B.; Mann, I. R.; Nakamura, R.; Nakamura, T. K. M.; Narita, Y.; Retinò, A.; Sahraoui, F.; Schekochihin, A.; Schwartz, S. J.; Shinohara, I.; Sorriso-Valvo, L.

2012-04-01

We describe the mission concept of how ESA can make a major contribution to the Japanese Canadian multi-spacecraft mission SCOPE by adding one cost-effective spacecraft EIDO (Electron and Ion Dynamics Observatory), which has a comprehensive and optimized plasma payload to address the physics of particle acceleration. The combined mission EIDOSCOPE will distinguish amongst and quantify the governing processes of particle acceleration at several important plasma boundaries and their associated boundary layers: collisionless shocks, plasma jet fronts, thin current sheets and turbulent boundary layers. Particle acceleration and associated cross-scale coupling is one of the key outstanding topics to be addressed in the Plasma Universe. The very important science questions that only the combined EIDOSCOPE mission will be able to tackle are: 1) Quantitatively, what are the processes and efficiencies with which both electrons and ions are selectively injected and subsequently accelerated by collisionless shocks? 2) How does small-scale electron and ion acceleration at jet fronts due to kinetic processes couple simultaneously to large scale acceleration due to fluid (MHD) mechanisms? 3) How does multi-scale coupling govern acceleration mechanisms at electron, ion and fluid scales in thin current sheets? 4) How do particle acceleration processes inside turbulent boundary layers depend on turbulence properties at ion/electron scales? EIDO particle instruments are capable of resolving full 3D particle distribution functions in both thermal and suprathermal regimes and at high enough temporal resolution to resolve the relevant scales even in very dynamic plasma processes. The EIDO spin axis is designed to be sun-pointing, allowing EIDO to carry out the most sensitive electric field measurements ever accomplished in the outer magnetosphere. Combined with a nearby SCOPE Far Daughter satellite, EIDO will form a second pair (in addition to SCOPE Mother-Near Daughter) of closely separated satellites that provides the unique capability to measure the 3D electric field with high accuracy and sensitivity. All EIDO instrumentation are state-of-the-art technology with heritage from many recent missions. The EIDOSCOPE orbit will be close to equatorial with apogee 25-30 RE and perigee 8-10 RE. In the course of one year the orbit will cross all the major plasma boundaries in the outer magnetosphere; bow shock, magnetopause and magnetotail current sheets, jet fronts and turbulent boundary layers. EIDO offers excellent cost/benefits for ESA, as for only a fraction of an M-class mission cost ESA can become an integral part of a major multi-agency L-class level mission that addresses outstanding science questions for the benefit of the European science community.

Shock loading predictions from application of indicial theory to shock-turbulence interactions

NASA Technical Reports Server (NTRS)

Keefe, Laurence R.; Nixon, David

1991-01-01

A sequence of steps that permits prediction of some of the characteristics of the pressure field beneath a fluctuating shock wave from knowledge of the oncoming turbulent boundary layer is presented. The theory first predicts the power spectrum and pdf of the position and velocity of the shock wave, which are then used to obtain the shock frequency distribution, and the pdf of the pressure field, as a function of position within the interaction region. To test the validity of the crucial assumption of linearity, the indicial response of a normal shock is calculated from numerical simulation. This indicial response, after being fit by a simple relaxation model, is used to predict the shock position and velocity spectra, along with the shock passage frequency distribution. The low frequency portion of the shock spectra, where most of the energy is concentrated, is satisfactorily predicted by this method.

Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

NASA Technical Reports Server (NTRS)

Bose, Deepak; Brown, James L.; Prabhu, Dinesh K.; Gnoffo, Peter; Johnston, Christopher O.; Hollis, Brian

2011-01-01

The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

Real-gas effects 1: Simulation of ideal gas flow by cryogenic nitrogen and other selected gases

NASA Technical Reports Server (NTRS)

Hall, R. M.

1980-01-01

The thermodynamic properties of nitrogen gas do not thermodynamically approximate an ideal, diatomic gas at cryogenic temperatures. Choice of a suitable equation of state to model its behavior is discussed and the equation of Beattie and Bridgeman is selected as best meeting the needs for cryogenic wind tunnel use. The real gas behavior of nitrogen gas is compared to an ideal, diatomic gas for the following flow processes: isentropic expansion; normal shocks; boundary layers; and shock wave boundary layer interactions. The only differences in predicted pressure ratio between nitrogen and an ideal gas that may limit the minimum operating temperatures of transonic cryogenic wind tunnels seem to occur at total pressures approaching 9atmospheres and total temperatures 10 K below the corresponding saturation temperature, where the differences approach 1 percent for both isentropic expansions and normal shocks. Several alternative cryogenic test gases - air, helium, and hydrogen - are also analyzed. Differences in air from an ideal, diatomic gas are similar in magnitude to those of nitrogen. Differences for helium and hydrogen are over an order of magnitude greater than those for nitrogen or air. Helium and hydrogen do not approximate the compressible flow of an ideal, diatomic gas.

Grid adaption using Chimera composite overlapping meshes

NASA Technical Reports Server (NTRS)

Kao, Kai-Hsiung; Liou, Meng-Sing; Chow, Chuen-Yen

1993-01-01

The objective of this paper is to perform grid adaptation using composite over-lapping meshes in regions of large gradient to capture the salient features accurately during computation. The Chimera grid scheme, a multiple overset mesh technique, is used in combination with a Navier-Stokes solver. The numerical solution is first converged to a steady state based on an initial coarse mesh. Solution-adaptive enhancement is then performed by using a secondary fine grid system which oversets on top of the base grid in the high-gradient region, but without requiring the mesh boundaries to join in any special way. Communications through boundary interfaces between those separated grids are carried out using tri-linear interpolation. Applications to the Euler equations for shock reflections and to a shock wave/boundary layer interaction problem are tested. With the present method, the salient features are well resolved.

Grid adaptation using chimera composite overlapping meshes

NASA Technical Reports Server (NTRS)

Kao, Kai-Hsiung; Liou, Meng-Sing; Chow, Chuen-Yen

1994-01-01

The objective of this paper is to perform grid adaptation using composite overlapping meshes in regions of large gradient to accurately capture the salient features during computation. The chimera grid scheme, a multiple overset mesh technique, is used in combination with a Navier-Stokes solver. The numerical solution is first converged to a steady state based on an initial coarse mesh. Solution-adaptive enhancement is then performed by using a secondary fine grid system which oversets on top of the base grid in the high-gradient region, but without requiring the mesh boundaries to join in any special way. Communications through boundary interfaces between those separated grids are carried out using trilinear interpolation. Application to the Euler equations for shock reflections and to shock wave/boundary layer interaction problem are tested. With the present method, the salient features are well-resolved.

Grid adaptation using Chimera composite overlapping meshes

NASA Technical Reports Server (NTRS)

Kao, Kai-Hsiung; Liou, Meng-Sing; Chow, Chuen-Yen

1993-01-01

The objective of this paper is to perform grid adaptation using composite over-lapping meshes in regions of large gradient to capture the salient features accurately during computation. The Chimera grid scheme, a multiple overset mesh technique, is used in combination with a Navier-Stokes solver. The numerical solution is first converged to a steady state based on an initial coarse mesh. Solution-adaptive enhancement is then performed by using a secondary fine grid system which oversets on top of the base grid in the high-gradient region, but without requiring the mesh boundaries to join in any special way. Communications through boundary interfaces between those separated grids are carried out using tri-linear interpolation. Applications to the Euler equations for shock reflections and to a shock wave/boundary layer interaction problem are tested. With the present method, the salient features are well resolved.

Fluorescence Visualization of Hypersonic Flow Past Triangular and Rectangular Boundary-layer Trips

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Garcia, A. P.; Borg, Stephen E.; Dyakonov, Artem A.; Berry, Scott A.; Inman, Jennifer A.; Alderfer, David W.

2007-01-01

Planar laser-induced fluorescence (PLIF) flow visualization has been used to investigate the hypersonic flow of air over surface protrusions that are sized to force laminar-to-turbulent boundary layer transition. These trips were selected to simulate protruding Space Shuttle Orbiter heat shield gap-filler material. Experiments were performed in the NASA Langley Research Center 31-Inch Mach 10 Air Wind Tunnel, which is an electrically-heated, blowdown facility. Two-mm high by 8-mm wide triangular and rectangular trips were attached to a flat plate and were oriented at an angle of 45 degrees with respect to the oncoming flow. Upstream of these trips, nitric oxide (NO) was seeded into the boundary layer. PLIF visualization of this NO allowed observation of both laminar and turbulent boundary layer flow downstream of the trips for varying flow conditions as the flat plate angle of attack was varied. By varying the angle of attack, the Mach number above the boundary layer was varied between 4.2 and 9.8, according to analytical oblique-shock calculations. Computational Fluid Dynamics (CFD) simulations of the flowfield with a laminar boundary layer were also performed to better understand the flow environment. The PLIF images of the tripped boundary layer flow were compared to a case with no trip for which the flow remained laminar over the entire angle-of-attack range studied. Qualitative agreement is found between the present observed transition measurements and a previous experimental roughness-induced transition database determined by other means, which is used by the shuttle return-to-flight program.

Observations of the magnetopause current layer: Cases with no boundary layer and tests of recent models

NASA Technical Reports Server (NTRS)

Eastman, Timothy E.

1995-01-01

Evidence for the probable existence of magnetospheric boundary layers was first presented by Hones, et al. (1972), based on VELA satellite plasma observations (no magnetic field measurements were obtained). This magnetotail boundary layer is now known to be the tailward extension of the high-latitude boundary layer or plasma mantle (first uniquely identified using HEOS 2 plasma and field observations by Rosenbauer et al., 1975) and the low-latitude boundary layer (first uniquely identified using IMP 6 plasma and field observations by Eastman et al., 1976). The magnetospheric boundary layer is the region of magnetosheath-like plasma located Earthward of, but generally contiguous with the magnetopause. This boundary layer is typically identified by comparing low-energy (less than 10 keV) ion spectra across the magnetopause. Low-energy electron measurements are also useful for identifying the boundary layer because the shocked solar wind or magnetosheath has a characteristic spectral signature for electrons as well. However, there are magnetopause crossings where low-energy electrons might suggest a depletion layer outside the magnetopause even though the traditional field-rotation signature indicates that this same region is a boundary layer Earthward of the current layer. Our analyses avoided crossings which exhibit such ambiguities. Pristine magnetopause crossings are magnetopause crossings for which the current layer is well defined and for which there is no adjoining magnetospheric boundary layer as defined above. Although most magnetopause models to date apply to such crossings, few comparisons between such theory and observations of pristine magnetopause crossings have been made because most crossings have an associated magnetospheric boundary layer which significantly affects the applicable boundary conditions for the magnetopause current layer. Furthermore, almost no observational studies of magnetopause microstructure have been done even though key theoretical issues have been discussed for over two decades. This is because plasma instruments deployed prior to the ISEE and AMPTE missions did not have the required time resolution and most ISEE investigations to-date have focused on tests of MHD plasma models, especially reconnection. More recently, many phenomenological and theoretical models have been developed to explain the existence and characteristics of the magnetospheric boundary layers with only limited success to date. The cases with no boundary layer treated in this study provide a contrary set of conditions to those observed with a boundary layer. For the measured parameters of such cases, a successful boundary layer model should predict no plasma penetration across the magnetopause. Thus, this research project provides the first direct observational tests of magnetopause models using pristine magnetopause crossings and provides important new results on magnetopause microstructure and associated kinetic processes.

An immersed boundary formulation for simulating high-speed compressible viscous flows with moving solids

NASA Astrophysics Data System (ADS)

Qu, Yegao; Shi, Ruchao; Batra, Romesh C.

2018-02-01

We present a robust sharp-interface immersed boundary method for numerically studying high speed flows of compressible and viscous fluids interacting with arbitrarily shaped either stationary or moving rigid solids. The Navier-Stokes equations are discretized on a rectangular Cartesian grid based on a low-diffusion flux splitting method for inviscid fluxes and conservative high-order central-difference schemes for the viscous components. Discontinuities such as those introduced by shock waves and contact surfaces are captured by using a high-resolution weighted essentially non-oscillatory (WENO) scheme. Ghost cells in the vicinity of the fluid-solid interface are introduced to satisfy boundary conditions on the interface. Values of variables in the ghost cells are found by using a constrained moving least squares method (CMLS) that eliminates numerical instabilities encountered in the conventional MLS formulation. The solution of the fluid flow and the solid motion equations is advanced in time by using the third-order Runge-Kutta and the implicit Newmark integration schemes, respectively. The performance of the proposed method has been assessed by computing results for the following four problems: shock-boundary layer interaction, supersonic viscous flows past a rigid cylinder, moving piston in a shock tube and lifting off from a flat surface of circular, rectangular and elliptic cylinders triggered by shock waves, and comparing computed results with those available in the literature.

Adaptive wing and flow control technology

NASA Astrophysics Data System (ADS)

Stanewsky, E.

2001-10-01

The development of the boundary layer and the interaction of the boundary layer with the outer “inviscid” flow field, exacerbated at high speed by the occurrence of shock waves, essentially determine the performance boundaries of high-speed flight. Furthermore, flight and freestream conditions may change considerably during an aircraft mission while the aircraft itself is only designed for multiple but fixed design points thus impairing overall performance. Consequently, flow and boundary layer control and adaptive wing technology may have revolutionary new benefits for take-off, landing and cruise operating conditions for many aircraft by enabling real-time effective geometry optimization relative to the flight conditions. In this paper we will consider various conventional and novel means of boundary layer and flow control applied to moderate-to-large aspect ratio wings, delta wings and bodies with the specific objectives of drag reduction, lift enhancement, separation suppression and the improvement of air-vehicle control effectiveness. In addition, adaptive wing concepts of varying complexity and corresponding aerodynamic performance gains will be discussed, also giving some examples of possible structural realizations. Furthermore, penalties associated with the implementation of control and adaptation mechanisms into actual aircraft will be addressed. Note that the present contribution is rather application oriented.

Aerodynamic Interaction between Delta Wing and Hemisphere-Cylinder in Supersonic Flow

NASA Astrophysics Data System (ADS)

Nishino, Atsuhiro; Ishikawa, Takahumi; Nakamura, Yoshiaki

As future space vehicles, Reusable Launch Vehicle (RLV) needs to be developed, where there are two kinds of RLV: Single Stage To Orbit (SSTO) and Two Stage To Orbit (TSTO). In the latter case, the shock/shock interaction and shock/boundary layer interaction play a key role. In the present study, we focus on the supersonic flow field with aerodynamic interaction between a delta wing and a hemisphere-cylinder, which imitate a TSTO, where the clearance, h, between the delta wing and hemisphere-cylinder is a key parameter. As a result, complicated flow patterns were made clear, including separation bubbles.

β-distribution for Reynolds stress and turbulent heat flux in relaxation turbulent boundary layer of compression ramp

NASA Astrophysics Data System (ADS)

Hu, YanChao; Bi, WeiTao; Li, ShiYao; She, ZhenSu

2017-12-01

A challenge in the study of turbulent boundary layers (TBLs) is to understand the non-equilibrium relaxation process after sep-aration and reattachment due to shock-wave/boundary-layer interaction. The classical boundary layer theory cannot deal with the strong adverse pressure gradient, and hence, the computational modeling of this process remains inaccurate. Here, we report the direct numerical simulation results of the relaxation TBL behind a compression ramp, which reveal the presence of intense large-scale eddies, with significantly enhanced Reynolds stress and turbulent heat flux. A crucial finding is that the wall-normal profiles of the excess Reynolds stress and turbulent heat flux obey a β-distribution, which is a product of two power laws with respect to the wall-normal distances from the wall and from the boundary layer edge. In addition, the streamwise decays of the excess Reynolds stress and turbulent heat flux also exhibit power laws with respect to the streamwise distance from the corner of the compression ramp. These results suggest that the relaxation TBL obeys the dilation symmetry, which is a specific form of self-organization in this complex non-equilibrium flow. The β-distribution yields important hints for the development of a turbulence model.

Airfoil shape for flight at subsonic speeds. [design analysis and aerodynamic characteristics of the GAW-1 airfoil

NASA Technical Reports Server (NTRS)

Whitcomb, R. T. (Inventor)

1976-01-01

An airfoil is examined that has an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency. Diagrams illustrating supersonic flow and shock waves over the airfoil are shown.

Penetration Characteristics of Air, Carbon Dioxide and Helium Transverse Sonic Jets in Mach 5 Cross Flow.

PubMed

Erdem, Erinc; Kontis, Konstantinos; Saravanan, Selvaraj

2014-12-08

An experimental investigation of sonic air, CO 2 and Helium transverse jets in Mach 5 cross flow was carried out over a flat plate. The jet to freestream momentum flux ratio, J, was kept the same for all gases. The unsteady flow topology was examined using high speed schlieren visualisation and PIV. Schlieren visualisation provided information regarding oscillating jet shear layer structures and bow shock, Mach disc and barrel shocks. Two-component PIV measurements at the centreline, provided information regarding jet penetration trajectories. Barrel shocks and Mach disc forming the jet boundary were visualised/quantified also jet penetration boundaries were determined. Even though J is kept the same for all gases, the penetration patterns were found to be remarkably different both at the nearfield and the farfield. Air and CO 2 jet resulted similar nearfield and farfield penetration pattern whereas Helium jet spread minimal in the nearfield.

Penetration Characteristics of Air, Carbon Dioxide and Helium Transverse Sonic Jets in Mach 5 Cross Flow

PubMed Central

Erdem, Erinc; Kontis, Konstantinos; Saravanan, Selvaraj

2014-01-01

An experimental investigation of sonic air, CO2 and Helium transverse jets in Mach 5 cross flow was carried out over a flat plate. The jet to freestream momentum flux ratio, J, was kept the same for all gases. The unsteady flow topology was examined using high speed schlieren visualisation and PIV. Schlieren visualisation provided information regarding oscillating jet shear layer structures and bow shock, Mach disc and barrel shocks. Two-component PIV measurements at the centreline, provided information regarding jet penetration trajectories. Barrel shocks and Mach disc forming the jet boundary were visualised/quantified also jet penetration boundaries were determined. Even though J is kept the same for all gases, the penetration patterns were found to be remarkably different both at the nearfield and the farfield. Air and CO2 jet resulted similar nearfield and farfield penetration pattern whereas Helium jet spread minimal in the nearfield. PMID:25494348

Supersonic Wind Tunnel Tests of a Half-axisymmetric 12 Deg-spike Inlet to a Rocket-based Combined-cycle Propulsion System

NASA Technical Reports Server (NTRS)

DeBonis, J. R.; Trefny, C. J.

2001-01-01

Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.

An evaluation of three two-dimensional computational fluid dynamics codes including low Reynolds numbers and transonic Mach numbers

NASA Technical Reports Server (NTRS)

Hicks, Raymond M.; Cliff, Susan E.

1991-01-01

Full-potential, Euler, and Navier-Stokes computational fluid dynamics (CFD) codes were evaluated for use in analyzing the flow field about airfoils sections operating at Mach numbers from 0.20 to 0.60 and Reynolds numbers from 500,000 to 2,000,000. The potential code (LBAUER) includes weakly coupled integral boundary layer equations for laminar and turbulent flow with simple transition and separation models. The Navier-Stokes code (ARC2D) uses the thin-layer formulation of the Reynolds-averaged equations with an algebraic turbulence model. The Euler code (ISES) includes strongly coupled integral boundary layer equations and advanced transition and separation calculations with the capability to model laminar separation bubbles and limited zones of turbulent separation. The best experiment/CFD correlation was obtained with the Euler code because its boundary layer equations model the physics of the flow better than the other two codes. An unusual reversal of boundary layer separation with increasing angle of attack, following initial shock formation on the upper surface of the airfoil, was found in the experiment data. This phenomenon was not predicted by the CFD codes evaluated.

Downstream boundary effects on the frequency of self-excited oscillations in transonic diffuser flows

NASA Astrophysics Data System (ADS)

Hsieh, T.

1986-10-01

Investigation of downstream boundary effects on the frequency of self-excited oscillations in two-dimensional, separated transonic diffuser flows were conducted numerically by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equation with two equation turbulence models. It was found that the flow fields are very sensitive to the location of the downstream boundary. Extension of the diffuser downstream boundary significantly reduces the frequency and amplitude of oscillations for pressure, velocity, and shock. The existence of a suction slot in the experimental setpup obscures the physical downstream boundary and therefore presents a difficulty for quantitative comparisons between computation and experiment.

Investigation of Turbulent Boundary-Layer Separation Using Laser Velocimetry

NASA Technical Reports Server (NTRS)

Modarress, D.; Johnson, D. A.

1979-01-01

Boundary-layer measurements realized by laser velocimetry are presented for a Much 2.9, two-dimensional, shock-wave/turbulent boundary-layer interaction containing an extensive region of separated flow. Mean velocity and turbulent intensity profiles were obtained from upstream of the interaction zone to downstream of the mean reattachment point. The superiority of the laser velocimeter technique over pressure sensors in turbulent separated flows is demonstrated by a comparison of the laser velocimeter data with results obtained from local pilot and static pressure measurements for the same flow conditions. The locations of the mean separation and reattachment points as deduced from the mean velocity measurements are compared to oil-now visualization results. Representative velocity probability density functions obtained in the separated now region are also presented. Critical to the success of this investigation were: the use of Bragg cell frequency shifting and artificial seeding of the now with submicron light-scattering particles.

On the boundary flow using pulsed nanosecond DBD plasma actuators

NASA Astrophysics Data System (ADS)

Zhao, Zi-Jie; Cui, Y. D.; Li, Jiun-Ming; Zheng, Jian-Guo; Khoo, B. C.

2018-05-01

Our previous studies in quiescent air environment [Z. J. Zhao et al., AIAA J. 53(5) (2015) 1336; J. G. Zheng et al., Phys. Fluids 26(3) (2014) 036102] reveal experimentally and numerically that the shock wave generated by the nanosecond pulsed plasma is fundamentally a microblast wave. The shock-induced burst perturbations (overpressure and induced velocity) are found to be restricted to a very narrow region (about 1 mm) behind the shock front and last only for a few microseconds. These results indicate that the pulsed nanosecond dielectric barrier discharge (DBD) plasma actuator has stronger local effects in time and spatial domain. In this paper, we further investigate the effects of pulsed plasma on the boundary layer flow over a flat plate. The present investigation reveals that the nanosecond pulsed plasma actuator generates intense perturbations and tends to promote the laminar boundary over a flat plate to turbulent flow. The heat effect after the pulsed plasma discharge was observed in the external flow, lasting a few milliseconds for a single pulse and reaching a quasi-stable state for multi-pulses.

Acoustic loads prediction on jet aircraft

NASA Technical Reports Server (NTRS)

Reddy, N. N.

1992-01-01

A nearfield aircraft noise prediction computer program is presented for the F-22 aircraft. The dominant sources of noise are jet turbulent mixing noise, jet broadband shock noise, and fluctuating pressure under the turbulent boundary layer. All results from this investigation are presented in viewgraph format.

Acoustic loads prediction on jet aircraft

NASA Astrophysics Data System (ADS)

Reddy, N. N.

1992-07-01

A nearfield aircraft noise prediction computer program is presented for the F-22 aircraft. The dominant sources of noise are jet turbulent mixing noise, jet broadband shock noise, and fluctuating pressure under the turbulent boundary layer. All results from this investigation are presented in viewgraph format.

Boundary-Layer Transition on a Slender Cone in Hypervelocity Flow with Real Gas Effects

NASA Astrophysics Data System (ADS)

Jewell, Joseph Stephen

The laminar to turbulent transition process in boundary layer flows in thermochemical nonequilibrium at high enthalpy is measured and characterized. Experiments are performed in the T5 Hypervelocity Reflected Shock Tunnel at Caltech, using a 1 m length 5-degree half angle axisymmetric cone instrumented with 80 fast-response annular thermocouples, complemented by boundary layer stability computations using the STABL software suite. A new mixing tank is added to the shock tube fill apparatus for premixed freestream gas experiments, and a new cleaning procedure results in more consistent transition measurements. Transition location is nondimensionalized using a scaling with the boundary layer thickness, which is correlated with the acoustic properties of the boundary layer, and compared with parabolized stability equation (PSE) analysis. In these nondimensionalized terms, transition delay with increasing CO2 concentration is observed: tests in 100% and 50% CO2, by mass, transition up to 25% and 15% later, respectively, than air experiments. These results are consistent with previous work indicating that CO2 molecules at elevated temperatures absorb acoustic instabilities in the MHz range, which is the expected frequency of the Mack second-mode instability at these conditions, and also consistent with predictions from PSE analysis. A strong unit Reynolds number effect is observed, which is believed to arise from tunnel noise. NTr for air from 5.4 to 13.2 is computed, substantially higher than previously reported for noisy facilities. Time- and spatially-resolved heat transfer traces are used to track the propagation of turbulent spots, and convection rates at 90%, 76%, and 63% of the boundary layer edge velocity, respectively, are observed for the leading edge, centroid, and trailing edge of the spots. A model constructed with these spot propagation parameters is used to infer spot generation rates from measured transition onset to completion distance. Finally, a novel method to control transition location with boundary layer gas injection is investigated. An appropriate porous-metal injector section for the cone is designed and fabricated, and the efficacy of injected CO2 for delaying transition is gauged at various mass flow rates, and compared with both no injection and chemically inert argon injection cases. While CO2 injection seems to delay transition, and argon injection seems to promote it, the experimental results are inconclusive and matching computations do not predict a reduction in N factor from any CO2 injection condition computed.

On the quasi-conical flowfield structure of the swept shock wave-turbulent boundary layer interaction

NASA Technical Reports Server (NTRS)

Knight, Doyle D.; Badekas, Dias

1991-01-01

The swept oblique shock-wave/turbulent-boundary-layer interaction generated by a 20-deg sharp fin at Mach 4 and Reynolds number 21,000 is investigated via a series of computations using both conical and three-dimensional Reynolds-averaged Navier-Stokes equations with turbulence incorporated through the algebraic turbulent eddy viscosity model of Baldwin-Lomax. Results are compared with known experimental data, and it is concluded that the computed three-dimensional flowfield is quasi-conical (in agreement with the experimental data), the computed three-dimensional and conical surface pressure and surface flow direction are in good agreement with the experiment, and the three-dimensional and conical flows significantly underpredict the peak experimental skin friction. It is pointed out that most of the features of the conical flowfield model in the experiment are observed in the conical computation which also describes the complete conical streamline pattern not included in the model of the experiment.

Investigation of corner shock boundary layer interactions to understand inlet unstart

NASA Astrophysics Data System (ADS)

Funderburk, Morgan

2015-11-01

Inlet unstart is a detrimental phenomenon in dual-mode ramjet/scramjet engines that causes severe loss of thrust, large transient structural load, and potentially a loss of the aircraft. In order to analyze the effects that the corner shock boundary layer interaction (SBLI) has on initiating and perpetuating inlet unstart, a qualitative and quantitative investigation into mean and dynamic features of corner SBLI at various Mach numbers is made. Surface streakline visualization showed that the corner SBLI is highly three-dimensional with a dominant presence of corner separation vortex. Further, the peak r.m.s. pressure was located at the periphery of corner separation vortex, suggesting that the unsteady loading is caused by the corner vortex. Power spectral densities of wall-pressure fluctuations in the peak r.m.s. location were analyzed in order to characterize the dominant frequencies of oscillation of the flow structures and to unravel the dynamic interactions between them in order to expand the operating margin of future hypersonic air breathing vehicles.

Simulation of interior ballistics flows in a shock tube

NASA Astrophysics Data System (ADS)

Seiler, F.

1983-07-01

The flow in front of and behind a projectile was investigated in a interior ballistics shock tube simulator. Flow patterns and heat flow were examined for flows with and without gas leakage. The boundary layers behind the piston can clearly be shown by differential interferograms. The dependence of the heat flow into the measuring tube wall on the base form is smaller than the signal perturbations. Flow patterns show no appreciable effect of gas leakage on the flow behind the piston; strong flow effects arise in front of the piston. The same effects are shown by heat flow measurements. In case of gas leakage heat flows into the tube wall before the piston reaches the wall. In the slit between piston and wall a maximum heat flow is found. High temperature gradients, due to the fact that hot gases come closer to the tube wall than in the boundary layer flow behind the piston, lead to high thermal loading of the wall materials which can cause cracks.

Application of an Engineering Inviscid-Boundary Layer Method to Slender Three-Dimensional Vehicle Forebodies

NASA Technical Reports Server (NTRS)

Riley, Christopher J.

1993-01-01

An engineering inviscid-boundary layer method has been modified for application to slender three-dimensional (3-D) forebodies which are characteristic of transatmospheric vehicles. An improved shock description in the nose region has been added to the inviscid technique which allows the calculation of a wider range of body geometries. The modified engineering method is applied to the perfect gas solution over a slender 3-D configuration at angle of attack. The method predicts surface pressures and laminar heating rates on the windward side of the vehicle that compare favorably with numerical solutions of the thin-layer Navier-Stokes equations. These improvements extend the 3-D capabilities of the engineering method and significantly increase its design applications.

Understanding the Flow Physics of Shock Boundary-Layer Interactions Using CFD and Numerical Analyses

NASA Technical Reports Server (NTRS)

Friedlander, David J.

2013-01-01

Computational fluid dynamic (CFD) analyses of the University of Michigan (UM) Shock/Boundary-Layer Interaction (SBLI) experiments were performed as an extension of the CFD SBLI Workshop held at the 48th AIAA Aerospace Sciences Meeting in 2010. In particular, the UM Mach 2.75 Glass Tunnel with a semi-spanning 7.75deg wedge was analyzed in attempts to explore key physics pertinent to SBLI's, including thermodynamic and viscous boundary conditions as well as turbulence modeling. Most of the analyses were 3D CFD simulations using the OVERFLOW flow solver, with additional quasi-1D simulations performed with an in house MATLAB code interfacing with the NIST REFPROP code to explore perfect verses non-ideal air. A fundamental exploration pertaining to the effects of particle image velocimetry (PIV) on post-processing data is also shown. Results from the CFD simulations showed an improvement in agreement with experimental data with key contributions including adding a laminar zone upstream of the wedge and the necessity of mimicking PIV particle lag for comparisons. Results from the quasi-1D simulation showed that there was little difference between perfect and non-ideal air for the configuration presented.

Turbulent dusty boundary layer in an ANFO surface-burst explosion

NASA Astrophysics Data System (ADS)

Kuhl, A. L.; Ferguson, R. E.; Chien, K. Y.; Collins, J. P.

1992-01-01

This paper describes the results of numerical simulations of the dusty, turbulent boundary layer created by a surface burst explosion. The blast wave was generated by the detonation of a 600-T hemisphere of ANFO, similar to those used in large-scale field tests. The surface was assumed to be ideally noncratering but contained an initial loose layer of dust. The dust-air mixture in this fluidized bed was modeled as a dense gas (i.e., an equilibrium model, valid for very small-diameter dust particles). The evolution of the flow was calculated by a high-order Godunov code that solves the nonsteady conservation laws. Shock interactions with dense layer generated vorticity near the wall, a result that is similar to viscous, no-slip effects found in clean flows. The resulting wall shear layer was unstable, and rolled up into large-scale rotational structures. These structures entrained dense material from the wall layer and created a chaotically striated flow. The boundary layer grew due to merging of the large-scale structures and due to local entrainment of the dense material from the fluidized bed. The chaotic flow was averaged along similarity lines (i.e., lines of constant values of x = r/Rs and y = z/Rs where R(sub s) = ct(exp alpha)) to establish the mean-flow profiles and the r.m.s. fluctuating-flow profiles of the boundary layer.

A Study of the Unstable Modes in High Mach Number Gaseous Jets and Shear Layers

NASA Astrophysics Data System (ADS)

Bassett, Gene Marcel

1993-01-01

Instabilities affecting the propagation of supersonic gaseous jets have been studied using high resolution computer simulations with the Piecewise-Parabolic-Method (PPM). These results are discussed in relation to jets from galactic nuclei. These studies involve a detailed treatment of a single section of a very long jet, approximating the dynamics by using periodic boundary conditions. Shear layer simulations have explored the effects of shear layers on the growth of nonlinear instabilities. Convergence of the numerical approximations has been tested by comparing jet simulations with different grid resolutions. The effects of initial conditions and geometry on the dominant disruptive instabilities have also been explored. Simulations of shear layers with a variety of thicknesses, Mach numbers and densities perturbed by incident sound waves imply that the time for the excited kink modes to grow large in amplitude and disrupt the shear layer is taug = (546 +/- 24) (M/4)^{1.7 } (Apert/0.02) ^{-0.4} delta/c, where M is the jet Mach number, delta is the half-width of the shear layer, and A_ {pert} is the perturbation amplitude. For simulations of periodic jets, the initial velocity perturbations set up zig-zag shock patterns inside the jet. In each case a single zig-zag shock pattern (an odd mode) or a double zig-zag shock pattern (an even mode) grows to dominate the flow. The dominant kink instability responsible for these shock patterns moves approximately at the linear resonance velocity, nu_ {mode} = cextnu_ {relative}/(cjet + c_ {ext}). For high resolution simulations (those with 150 or more computational zones across the jet width), the even mode dominates if the even penetration is higher in amplitude initially than the odd perturbation. For low resolution simulations, the odd mode dominates even for a stronger even mode perturbation. In high resolution simulations the jet boundary rolls up and large amounts of external gas are entrained into the jet. In low resolution simulations this entrainment process is impeded by numerical viscosity. The three-dimensional jet simulations behave similarly to two-dimensional jet runs with the same grid resolutions.

Hypersonic aerodynamic characteristics of a family of power-law, wing body configurations

NASA Technical Reports Server (NTRS)

Townsend, J. C.

1973-01-01

The configurations analyzed are half-axisymmetric, power-law bodies surmounted by thin, flat wings. The wing planform matches the body shock-wave shape. Analytic solutions of the hypersonic small disturbance equations form a basis for calculating the longitudinal aerodynamic characteristics. Boundary-layer displacement effects on the body and the wing upper surface are approximated. Skin friction is estimated by using compressible, laminar boundary-layer solutions. Good agreement was obtained with available experimental data for which the basic theoretical assumptions were satisfied. The method is used to estimate the effects of power-law, fineness ratio, and Mach number variations at full-scale conditions. The computer program is included.

Separation control by vortex generator devices in a transonic channel flow

NASA Astrophysics Data System (ADS)

Bur, Reynald; Coponet, Didier; Carpels, Yves

2009-12-01

An experimental study was conducted in a transonic channel to control by mechanical vortex generator devices the strong interaction between a shock wave and a separated turbulent boundary layer. Control devices—co-rotating and counter-rotating vane-type vortex generators—were implemented upstream of the shock foot region and tested both on a steady shock wave and on a forced shock oscillation configurations. The spanwise spacing of vortex generator devices along the channel appeared to be an important parameter to control the flow separation region. When the distance between each device is decreased, the vortices merging is more efficient to reduce the separation. Their placement upstream of the shock wave is determinant to ensure that vortices have mixed momentum all spanwise long before they reach the separation line, so as to avoid separation cells. Then, vortex generators slightly reduced the amplitude of the forced shock wave oscillation by delaying the upstream displacement of the leading shock.

Oscillations in a Forward-Facing Cavity Measured Using Laser-Differential Interferometry in a Hypersonic Quiet Tunnel

DTIC Science & Technology

2007-12-11

and behind the normal 5 Figure 2.1: Diagram of noise radiated from boundary layer transition on the nozzle wall shock of several pitot probes and a...are lower than those measured with the 0.067 in.-dia. pitot probe because the pressure fluctuations in the subsonic region behind the normal shock are...in.-dia. pitot probe before repolishing nozzle . . . . . . . . . . . . . . . . . . 16 2.10 Freestream noise level in the PQFLT measured with 1 in

Multidomain approach for calculating compressible flows

NASA Technical Reports Server (NTRS)

Cambier, L.; Chazzi, W.; Veuillot, J. P.; Viviand, H.

1982-01-01

A multidomain approach for calculating compressible flows by using unsteady or pseudo-unsteady methods is presented. This approach is based on a general technique of connecting together two domains in which hyperbolic systems (that may differ) are solved with the aid of compatibility relations associated with these systems. Some examples of this approach's application to calculating transonic flows in ideal fluids are shown, particularly the adjustment of shock waves. The approach is then applied to treating a shock/boundary layer interaction problem in a transonic channel.

Transient hot-film sensor response in a shock tube

NASA Technical Reports Server (NTRS)

Roberts, A. S., Jr.; Ortgies, K. R.; Gartenberg, E.

1989-01-01

Shock tube experiments were performed to determine the response of a hot-film sensor, mounted flush on the side wall of a shock tube, to unsteady flow behind a normal shock wave. The present experiments attempt to isolate the response of the anemometer due only to the change in convective heat transfer at the hot-film surface. The experiments, performed at low supersonic shock speeds in air, are described along with the data acquisition procedure. The change in convective heat transfer is deduced from the data and the results are compared with those from transient boundary layer theory and another set of experimental results. Finally, a transient local heat transfer coefficient is formulated for use as the forcing function in a hot-film sensor instrument model simulation.

Mechanism analysis of Magnetohydrodynamic heat shield system and optimization of externally applied magnetic field

NASA Astrophysics Data System (ADS)

Li, Kai; Liu, Jun; Liu, Weiqiang

2017-04-01

As a novel thermal protection technique for hypersonic vehicles, Magnetohydrodynamic (MHD) heat shield system has been proved to be of great intrinsic value in the hypersonic field. In order to analyze the thermal protection mechanisms of such a system, a physical model is constructed for analyzing the effect of the Lorentz force components in the counter and normal directions. With a series of numerical simulations, the dominating Lorentz force components are analyzed for the MHD heat flux mitigation in different regions of a typical reentry vehicle. Then, a novel magnetic field with variable included angle between magnetic induction line and streamline is designed, which significantly improves the performance of MHD thermal protection in the stagnation and shoulder areas. After that, the relationships between MHD shock control and MHD thermal protection are investigated, based on which the magnetic field above is secondarily optimized obtaining better performances of both shock control and thermal protection. Results show that the MHD thermal protection is mainly determined by the Lorentz force's effect on the boundary layer. From the stagnation to the shoulder region, the flow deceleration effect of the counter-flow component is weakened while the flow deflection effect of the normal component is enhanced. Moreover, there is no obviously positive correlation between the MHD shock control and thermal protection. But once a good Lorentz force's effect on the boundary layer is guaranteed, the thermal protection performance can be further improved with an enlarged shock stand-off distance by strengthening the counter-flow Lorentz force right after shock.

A sharp interface Cartesian grid method for viscous simulation of shocked particle-laden flows

NASA Astrophysics Data System (ADS)

Das, Pratik; Sen, Oishik; Jacobs, Gustaaf; Udaykumar, H. S.

2017-09-01

A Cartesian grid-based sharp interface method is presented for viscous simulations of shocked particle-laden flows. The moving solid-fluid interfaces are represented using level sets. A moving least-squares reconstruction is developed to apply the no-slip boundary condition at solid-fluid interfaces and to supply viscous stresses to the fluid. The algorithms developed in this paper are benchmarked against similarity solutions for the boundary layer over a fixed flat plate and against numerical solutions for moving interface problems such as shock-induced lift-off of a cylinder in a channel. The framework is extended to 3D and applied to calculate low Reynolds number steady supersonic flow over a sphere. Viscous simulation of the interaction of a particle cloud with an incident planar shock is demonstrated; the average drag on the particles and the vorticity field in the cloud are compared to the inviscid case to elucidate the effects of viscosity on momentum transfer between the particle and fluid phases. The methods developed will be useful for obtaining accurate momentum and heat transfer closure models for macro-scale shocked particulate flow applications such as blast waves and dust explosions.

Studies of aerothermal loads generated in regions of shock/shock interaction in hypersonic flow

NASA Technical Reports Server (NTRS)

Holden, Michael S.; Moselle, John R.; Lee, Jinho

1991-01-01

Experimental studies were conducted to examine the aerothermal characteristics of shock/shock/boundary layer interaction regions generated by single and multiple incident shocks. The presented experimental studies were conducted over a Mach number range from 6 to 19 for a range of Reynolds numbers to obtain both laminar and turbulent interaction regions. Detailed heat transfer and pressure measurements were made for a range of interaction types and incident shock strengths over a transverse cylinder, with emphasis on the 3 and 4 type interaction regions. The measurements were compared with the simple Edney, Keyes, and Hains models for a range of interaction configurations and freestream conditions. The complex flowfields and aerothermal loads generated by multiple-shock impingement, while not generating as large peak loads, provide important test cases for code prediction. The detailed heat transfer and pressure measurements proved a good basis for evaluating the accuracy of simple prediction methods and detailed numerical solutions for laminar and transitional regions or shock/shock interactions.

Stagnation-point heat-transfer rate predictions at aeroassist flight conditions

NASA Technical Reports Server (NTRS)

Gupta, Roop N.; Jones, Jim J.; Rochelle, William C.

1992-01-01

The results are presented for the stagnation-point heat-transfer rates used in the design process of the Aeroassist Flight Experiment (AFE) vehicle over its entire aeropass trajectory. The prediction methods used in this investigation demonstrate the application of computational fluid dynamics (CFD) techniques to a wide range of flight conditions and their usefulness in a design process. The heating rates were computed by a viscous-shock-layer (VSL) code at the lower altitudes and by a Navier-Stokes (N-S) code for the higher altitude cases. For both methods, finite-rate chemically reacting gas was considered, and a temperature-dependent wall-catalysis model was used. The wall temperature for each case was assumed to be radiative equilibrium temperature, based on total heating. The radiative heating was estimated by using a correlation equation. Wall slip was included in the N-S calculation method, and this method implicitly accounts for shock slip. The N-S/VSL combination of projection methods was established by comparison with the published benchmark flow-field code LAURA results at lower altitudes, and the direct simulation Monte Carlo results at higher altitude cases. To obtain the design heating rate over the entire forward face of the vehicle, a boundary-layer method (BLIMP code) that employs reacting chemistry and surface catalysis was used. The ratio of the VSL or N-S method prediction to that obtained from the boundary-layer method code at the stagnation point is used to define an adjustment factor, which accounts for the errors involved in using the boundary-layer method.

Airfoil shape for flight at subsonic speeds

DOEpatents

Whitcomb, Richard T.

1976-01-01

An airfoil having an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency.

Assessment of CFD Modeling Capability for Hypersonic Shock Wave Boundary Layer Interactions

DTIC Science & Technology

2015-11-30

preparation. 8 Point of Contact Dr. John Spyropoulos NAVAIR 22195 Elmer Road Building 106, Floor 2, Room 209B Patuxent River, MD 20670 Email...Dynamics, Krieger Publish- ing Company, Malabar, Florida, 1965. [15] Millikan, R. and White, D., "Systematics of Vibrational Relaxation," Journal of

Use of Vortex Generators to Reduce Distortion for Mach 1.6 Streamline-Traced Supersonic Inlets

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank; Slater, John W.; Trefny, Chuck

2016-01-01

Reduce the total pressure distortion at the engine-fan face due to low-momentum flow caused by the interaction of an external terminal shock at the turbulent boundary layer along a streamline-traced external-compression (STEX) inlet for Mach 1.6.

Some Effects of Leading-Edge Sweep on Boundary-Layer Transition at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Chapman, Gray T.

1961-01-01

The effects of crossflow and shock strength on transition of the laminar boundary layer behind a swept leading edge have been investigated analytically and with the aid of available experimental data. An approximate method of determining the crossflow Reynolds number on a leading edge of circular cross section at supersonic speeds is presented. The applicability of the critical crossflow criterion described by Owen and Randall for transition on swept wings in subsonic flow was examined for the case of supersonic flow over swept circular cylinders. A wide range of applicability of the subsonic critical values is indicated. The corresponding magnitude of crossflow velocity necessary to cause instability on the surface of a swept wing at supersonic speeds was also calculated and found to be small. The effects of shock strength on transition caused by Tollmien-Schlichting type of instability are discussed briefly. Changes in local Reynolds number, due to shock strength, were found analytically to have considerably more effect on transition caused by Tollmien-Schlichting instability than on transition caused by crossflow instability. Changes in the mechanism controlling transition from Tollmien-Schlichting instability to crossflow instability were found to be possible as a wing is swept back and to result in large reductions in the length of laminar flow.

EASI - EQUILIBRIUM AIR SHOCK INTERFERENCE

NASA Technical Reports Server (NTRS)

Glass, C. E.

1994-01-01

New research on hypersonic vehicles, such as the National Aero-Space Plane (NASP), has raised concerns about the effects of shock-wave interference on various structural components of the craft. State-of-the-art aerothermal analysis software is inadequate to predict local flow and heat flux in areas of extremely high heat transfer, such as the surface impingement of an Edney-type supersonic jet. EASI revives and updates older computational methods for calculating inviscid flow field and maximum heating from shock wave interference. The program expands these methods to solve problems involving the six shock-wave interference patterns on a two-dimensional cylindrical leading edge with an equilibrium chemically reacting gas mixture (representing, for example, the scramjet cowl of the NASP). The inclusion of gas chemistry allows for a more accurate prediction of the maximum pressure and heating loads by accounting for the effects of high temperature on the air mixture. Caloric imperfections and specie dissociation of high-temperature air cause shock-wave angles, flow deflection angles, and thermodynamic properties to differ from those calculated by a calorically perfect gas model. EASI contains pressure- and temperature-dependent thermodynamic and transport properties to determine heating rates, and uses either a calorically perfect air model or an 11-specie, 7-reaction reacting air model at equilibrium with temperatures up to 15,000 K for the inviscid flowfield calculations. EASI solves the flow field and the associated maximum surface pressure and heat flux for the six common types of shock wave interference. Depending on the type of interference, the program solves for shock-wave/boundary-layer interaction, expansion-fan/boundary-layer interaction, attaching shear layer or supersonic jet impingement. Heat flux predictions require a knowledge (from experimental data or relevant calculations) of a pertinent length scale of the interaction. Output files contain flow-field information for the various shock-wave interference patterns and their associated maximum surface pressure and heat flux predictions. EASI is written in FORTRAN 77 for a DEC VAX 8500 series computer using the VAX/VMS operating system, and requires 75K of memory. The program is available on a 9-track 1600 BPI magnetic tape in DEC VAX BACKUP format. EASI was developed in 1989. DEC, VAX, and VMS are registered trademarks of the Digital Equipment Corporation.

A computational fluid dynamics simulation of the hypersonic flight of the Pegasus(TM) vehicle using an artificial viscosity model and a nonlinear filtering method. M.S. Thesis

NASA Technical Reports Server (NTRS)

Mendoza, John Cadiz

1995-01-01

The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.

Transient Growth Analysis of Compressible Boundary Layers with Parabolized Stability Equations

NASA Technical Reports Server (NTRS)

Paredes, Pedro; Choudhari, Meelan M.; Li, Fei; Chang, Chau-Lyan

2016-01-01

The linear form of parabolized linear stability equations (PSE) is used in a variational approach to extend the previous body of results for the optimal, non-modal disturbance growth in boundary layer flows. This methodology includes the non-parallel effects associated with the spatial development of boundary layer flows. As noted in literature, the optimal initial disturbances correspond to steady counter-rotating stream-wise vortices, which subsequently lead to the formation of stream-wise-elongated structures, i.e., streaks, via a lift-up effect. The parameter space for optimal growth is extended to the hypersonic Mach number regime without any high enthalpy effects, and the effect of wall cooling is studied with particular emphasis on the role of the initial disturbance location and the value of the span-wise wavenumber that leads to the maximum energy growth up to a specified location. Unlike previous predictions that used a basic state obtained from a self-similar solution to the boundary layer equations, mean flow solutions based on the full Navier-Stokes (NS) equations are used in select cases to help account for the viscous-inviscid interaction near the leading edge of the plate and also for the weak shock wave emanating from that region. These differences in the base flow lead to an increasing reduction with Mach number in the magnitude of optimal growth relative to the predictions based on self-similar mean-flow approximation. Finally, the maximum optimal energy gain for the favorable pressure gradient boundary layer near a planar stagnation point is found to be substantially weaker than that in a zero pressure gradient Blasius boundary layer.

Shock Wave Propagation in Layered Planetary Interiors: Revisited

NASA Astrophysics Data System (ADS)

Arkani-Hamed, J.; Monteux, J.

2017-12-01

The end of the terrestrial planet accretion is characterized by numerous large impacts. About 90% of the mass of a large planet is accreted while the core mantle separation is occurring, because of the accretionary and the short-lived radio-isotope heating. The characteristics of the shockwave propagation, hence the existing scaling laws are poorly known within the layered planets. Here, we use iSALE-2D hydrocode simulations to calculate shock pressure in a differentiated Mars type body for impact velocities of 5-20 km/s, and impactor sizes of 100-400 km. We use two different rheologies for the target interior, an inviscid model ("no-stress model") and a pressure and damage-dependent strength model ("elaborated model"). To better characterize the shock pressure within the whole mantle as a function of distance from the impact site, we propose the following distribution: (1) a near field zone larger than the isobaric core that extends to 7-15 times the projectile radius into the target, where the peak shock pressure decays exponentially with increasing distance, (2) a far field zone where the pressure decays with distance following a power law. The shock pressure decreases more rapidly with distance in the near field for the elaborated model than for the no-stress model because of the influence of acoustic fluidization and damage. However to better illustrate the influence of the rheology on the shock propagation, we use the same expressions to fit the shock pressure with distance for both models. At the core-mantle boundary, CMB, the peak shock pressure jumps as the shock wave enters the core. We derived the boundary condition at CMB for the peak shock pressure. It is less sensitive to the impact velocity or the impactor size, but strongly depends on the rheology of the planet's mantle. Because of the lower shock wave velocity in the core compared to that in the mantle, the refracted shockwave propagates toward the symmetry axis of the planet, and the shock pressure in the core decreases following a second power law. In this study, we express the output obtained from iSALE hydrocodes by scaling laws to illustrate the influence of the ray angle relative to the axis of symmetry, the target rheology, the impactor size and the impact velocity. We use these shock-pressure scaling laws to determine the impact heating of terrestrial planets.

Experimental Investigation of Unsteady Shock Wave Turbulent Boundary Layer Interactions About a Blunt Fin

NASA Technical Reports Server (NTRS)

Barnhart, Paul J.; Greber, Isaac

1997-01-01

A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from the dynamic pressure transducers were found to collapse to self-similar distributions as a function of the distance perpendicular to the separation line. The standard deviation of the pressure signals showed initial peaked distribution, with the maximum standard deviation point corresponding to the location of the separation line at Mach number 3.0 to 5.0. At Mach 2.0 the maximum standard deviation point was found to occur significantly upstream of the separation line. The intermittency distributions of the separation shock wave motion were found to be self-similar profiles for all Mach numbers. The intermittent region length was found to increase with Mach number and decrease with interaction sweepback angle. For Mach numbers 3.0 to 5.0 the separation line was found to correspond to high intermittencies or equivalently to the downstream locus of the separation shock wave motion. The Mach 2.0 tests, however, showed that the intermittent region occurs significantly upstream of the separation line. Power spectral densities measured in the intermittent regions were found to have self-similar frequency distributions when compared as functions of a Strouhal number for all Mach numbers and interaction sweepback angles. The maximum zero-crossing frequencies were found to correspond with the peak frequencies in the power spectra measured in the intermittent region.

Investigating a 65-Ma-Old Smoking Gun: Deep Drilling of the Chicxulub Impact Structure

NASA Astrophysics Data System (ADS)

Dressler, B. O.; Sharpton, V. L.; Morgan, J.; Buffler, R.; Moran, D.; Smit, J.; Stöffler, D.; Urrutia, J.

The Phanerozoic paleontological record is marked by several biological extinction events. One of them, at the Cretaceous/Tertiary (K/T) boundary, was responsible for the demise of about 5% of genera and 75% of species, including the dinosaurs. These drastic and abrupt changes in the development of life on Earth puzzled paleontologists in the past. Many a cause was put forward to account for them, amongst them climate changes, disease, or overspecialization. About 20 years ago, Alvarez et al. [1980] discovered a high iridium concentration in an Italian K/T boundary clay layer. They proposed that the iridium was derived from an extra-terrestrial impact 65 Ma ago and that the impact was the cause for the K/T boundary extinctions. The iridium layer was subsequently found at K/T boundary locations worldwide. Further evidence for a K/T impact came from the discovery of shocked quartz, nano-diamonds, glass spherules, and nickel-rich spinels in microkrystites in the iridium-rich layer. There was evidence for an impact event, but no crater.

Turbulence Modeling for Shock Wave/Turbulent Boundary Layer Interactions

NASA Technical Reports Server (NTRS)

Lillard, Randolph P.

2011-01-01

Accurate aerodynamic computational predictions are essential for the safety of space vehicles, but these computations are of limited accuracy when large pressure gradients are present in the flow. The goal of the current project is to improve the state of compressible turbulence modeling for high speed flows with shock wave / turbulent boundary layer interactions (SWTBLI). Emphasis will be placed on models that can accurately predict the separated region caused by the SWTBLI. These flows are classified as nonequilibrium boundary layers because of the very large and variable adverse pressure gradients caused by the shock waves. The lag model was designed to model these nonequilibrium flows by incorporating history effects. Standard one- and two-equation models (Spalart Allmaras and SST) and the lag model will be run and compared to a new lag model. This new model, the Reynolds stress tensor lag model (lagRST), will be assessed against multiple wind tunnel tests and correlations. The basis of the lag and lagRST models are to preserve the accuracy of the standard turbulence models in equilibrium turbulence, when the Reynolds stresses are linearly related to the mean strain rates, but create a lag between mean strain rate effects and turbulence when nonequilibrium effects become important, such as in large pressure gradients. The affect this lag has on the results for SWBLI and massively separated flows will be determined. These computations will be done with a modified version of the OVERFLOW code. This code solves the RANS equations on overset grids. It was used for this study for its ability to input very complex geometries into the flow solver, such as the Space Shuttle in the full stack configuration. The model was successfully implemented within two versions of the OVERFLOW code. Results show a substantial improvement over the baseline models for transonic separated flows. The results are mixed for the SWBLI assessed. Separation predictions are not as good as the baseline models, but the over prediction of the peak heat flux downstream of the reattachment shock that plagues many models is reduced.

Mach 6 flow field surveys beneath the forebody of an airbreathing missile

NASA Technical Reports Server (NTRS)

Johnson, P. J.; Hunt, J. L.

1986-01-01

Wall static, local stream static, and pitot pressure surveys were made on the windward side of a hypersonic airbreathing missile at full-scale length Reynolds numbers. In the inviscid part of the flow field, the experimental massflow ratios agreed with trends predicted by a three-dimensional method-of-characteristics solution. At a longitudinal station 3.5 diameters downstrea of the nose, the boundary layer was transitional or turbulent at zero incidence but became laminar as the angle of attack increased. The bell-shaped distribution of the boundary layer across the width of the body affected the mass flow distribution out to the bow shock and decreased the mass flow available the engine inlet.

Status of holographic interferometry at Wright Patterson Air Force Base

NASA Technical Reports Server (NTRS)

Seibert, George

1987-01-01

At Wright Patterson AFB, holographic interferometry has been used for nearly 15 years in a variety of supersonic and hypersonic wind tunnels. Specifically, holographic interferometry was used to study boundary layers, shock boundary layer interaction, and general flow diagnostics. Although a considerable amount of quantitative work was done, the difficulty of reducing data severely restricted this. In the future, it is of interest to use holographic interferometry in conjunction with laser Doppler velocimetry to do more complete diagnostics. Also, there is an interest to do particle field diagnostics in the combustion research facility. Finally, there are efforts in nondestructive testing where automated fringe readout and analysis would be extremely helpful.

Grid adaption based on modified anisotropic diffusion equations formulated in the parametic domain

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hagmeijer, R.

1994-11-01

A new grid-adaption algorithm for problems in computational fluid dynamics is presented. The basic equations are derived from a variational problem formulated in the parametric domain of the mapping that defines the existing grid. Modification of the basic equations provides desirable properties in boundary layers. The resulting modified anisotropic diffusion equations are solved for the computational coordinates as functions of the parametric coordinates and these functions are numerically inverted. Numerical examples show that the algorithm is robust, that shocks and boundary layers are well-resolved on the adapted grid, and that the flow solution becomes a globally smooth function of themore » computational coordinates.« less

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

NASA Technical Reports Server (NTRS)

Jones, Daniel R.; Allgood, Daniel C.; Saunders, Grady P.

2016-01-01

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Investigation of the effects of pressure gradient, temperature and wall temperature ratio on the stagnation point heat transfer for circular cylinders and gas turbine vanes

NASA Technical Reports Server (NTRS)

Nagamatsu, H. T.; Duffy, R. E.

1984-01-01

Low and high pressure shock tubes were designed and constructed for the purpose of obtaining heat transfer data over a temperature range of 390 to 2500 K, pressures of 0.3 to 42 atm, and Mach numbers of 0.15 to 1.5 with and without pressure gradient. A square test section with adjustable top and bottom walls was constructed to produce the favorable and adverse pressure gradient over the flat plate with heat gages. A water cooled gas turbine nozzle cascade which is attached to the high pressure shock tube was obtained to measuse the heat flux over pressure and suction surfaces. Thin-film platinum heat gages with a response time of a few microseconds were developed and used to measure the heat flux for laminar, transition, and turbulent boundary layers. The laminar boundary heat flux on the shock tube wall agreed with Mirel's flat plate theory. Stagnation point heat transfer for circular cylinders at low temperature compared with the theoretical prediction, but for a gas temperature of 922 K the heat fluxes were higher than the predicted values. Preliminary flat plate heat transfer data were measured for laminar, transition, and turbulent boundary layers with and without pressure gradients for free-stream temperatures of 350 to 2575 K and flow Mach numbers of 0.11 to 1.9. The experimental heat flux data were correlated with the laminar and turbulent theories and the agreement was good at low temperatures which was not the case for higher temperatures.

Similar solutions for viscous hypersonic flow over a slender three-fourths-power body of revolution

NASA Technical Reports Server (NTRS)

Lin, Chin-Shun

1987-01-01

For hypersonic flow with a shock wave, there is a similar solution consistent throughout the viscous and inviscid layers along a very slender three-fourths-power body of revolution The strong pressure interaction problem can then be treated by the method of similarity. Numerical calculations are performed in the viscous region with the edge pressure distribution known from the inviscid similar solutions. The compressible laminar boundary-layer equations are transformed into a system of ordinary differential equations. The resulting two-point boundary value problem is then solved by the Runge-Kutta method with a modified Newton's method for the corresponding boundary conditions. The effects of wall temperature, mass bleeding, and body transverse curvature are investigated. The induced pressure, displacement thickness, skin friction, and heat transfer due to the previously mentioned parameters are estimated and analyzed.

Experimental plasma studies

NASA Technical Reports Server (NTRS)

Dunn, M. G.

1972-01-01

The rate coefficients for the reactions C(+) + e(-) + e(-) yields C + e(-) and CO(+) + e(-) yields C + O were measured over the electron temperature range of approximately 1500 deg K to 7000 deg K. The measurements were performed in CO that had expanded from equilibrium reservoir conditions of 7060 deg K at 17.3 atm pressure and from 6260 deg K at 10.0 atm pressure. Two RAM flight probes were used to measure electron density and electron temperature in the expanding flow of a shock tunnel. Experiments were performed in the inviscid flow with both probes and in the nozzle-wall boundary layer with the constant bias-voltage probe. The distributions of electron density and electron temperature were independently measured using voltage-swept thin-wire probes. Thin-wire Langmuir probes were also used to measure the electron-density and electron-temperature distributions in the boundary layer of a sharp flat plate located on the nozzle centerline. Admittance measurements were performed with the RAM C and RAM C-C S-band antennas in the presence of an ionized boundary layer.

Vortex Generators in a Two-Dimensional, External-Compression Supersonic Inlet

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank K.; Slater, John W.

2016-01-01

Vortex generators within a two-dimensional, external-compression supersonic inlet for Mach 1.6 were investigated to determine their ability to increase total pressure recovery, reduce total pressure distortion, and improve the boundary layer. The vortex generators studied included vanes and ramps. The geometric factors of the vortex generators studied included height, length, spacing, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated through the computational solution of the steady-state Reynolds-averaged Navier-Stokes equations on multi-block, structured grids. The vortex generators were simulated by either gridding the geometry of the vortex generators or modeling the vortices generated by the vortex generators. The inlet performance was characterized by the inlet total pressure recovery, total pressure distortion, and incompressible shape factor of the boundary-layer at the engine face. The results suggested that downstream vanes reduced the distortion and improved the boundary layer. The height of the vortex generators had the greatest effect of the geometric factors.

An evaluation of four single element airfoil analytic methods

NASA Technical Reports Server (NTRS)

Freuler, R. J.; Gregorek, G. M.

1979-01-01

A comparison of four computer codes for the analysis of two-dimensional single element airfoil sections is presented for three classes of section geometries. Two of the computer codes utilize vortex singularities methods to obtain the potential flow solution. The other two codes solve the full inviscid potential flow equation using finite differencing techniques, allowing results to be obtained for transonic flow about an airfoil including weak shocks. Each program incorporates boundary layer routines for computing the boundary layer displacement thickness and boundary layer effects on aerodynamic coefficients. Computational results are given for a symmetrical section represented by an NACA 0012 profile, a conventional section illustrated by an NACA 65A413 profile, and a supercritical type section for general aviation applications typified by a NASA LS(1)-0413 section. The four codes are compared and contrasted in the areas of method of approach, range of applicability, agreement among each other and with experiment, individual advantages and disadvantages, computer run times and memory requirements, and operational idiosyncrasies.

Calculation of the flow field including boundary layer effects for supersonic mixed compression inlets at angles of attack

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.

1982-01-01

The flow field in supersonic mixed compression aircraft inlets at angle of attack is calculated. A zonal modeling technique is employed to obtain the solution which divides the flow field into different computational regions. The computational regions consist of a supersonic core flow, boundary layer flows adjacent to both the forebody/centerbody and cowl contours, and flow in the shock wave boundary layer interaction regions. The zonal modeling analysis is described and some computational results are presented. The governing equations for the supersonic core flow form a hyperbolic system of partial differential equations. The equations for the characteristic surfaces and the compatibility equations applicable along these surfaces are derived. The characteristic surfaces are the stream surfaces, which are surfaces composed of streamlines, and the wave surfaces, which are surfaces tangent to a Mach conoid. The compatibility equations are expressed as directional derivatives along streamlines and bicharacteristics, which are the lines of tangency between a wave surface and a Mach conoid.

Synchroneity of the K-T oceanic mass extinction and meteorite impact: Blake Nose, western North Atlantic

USGS Publications Warehouse

Norris, R.D.; Huber, B.T.; Self-Trail, J.

1999-01-01

A 10-cm-thick layer of green spherules occurs precisely at the biostratigraphic boundary between the Cretaceous and Paleogene (K-T boundary) at Ocean Drilling Program Site 1049 (lat 30??08???N, long 76??06???W). The spherulitic layer contains abundant rock fragments (chalk, limestone, dolomite, chert, mica books, and schist) as well as shocked quartz, abundant large Cretaceous planktic foraminifera, and rounded clasts of clay as long as 4 mm interpreted as altered tektite glass probably derived from the Chicxulub impact structure. Most of the Cretaceous foraminifera present above the spherule layer are not survivors since small specimens are conspicuously rare compared to large individuals. Instead, the Cretaceous taxa in Paleocene sediments are thought to be reworked. The first Paleocene planktic foraminifera and calcareous nannofossil species are recorded immediately above the spherule bed, the upper part of which contains an iridium anomaly. Hence, deposition of the impact ejecta exactly coincided with the biostratigraphic K-T boundary and demonstrates that the impact event was synchronous with the evolutionary turnover in the oceans. These results are consistent with a reanalysis of the biostratigraphy of the K-T boundary stratotype, which argues that shallow-marine K-T boundary sections are not biostratigraphically more complete than deep-sea K-T boundary sites.

Hybrid Large Eddy Simulation / Reynolds Averaged Navier-Stokes Modeling in Directed Energy Applications

NASA Astrophysics Data System (ADS)

Zilberter, Ilya Alexandrovich

In this work, a hybrid Large Eddy Simulation / Reynolds-Averaged Navier Stokes (LES/RANS) turbulence model is applied to simulate two flows relevant to directed energy applications. The flow solver blends the Menter Baseline turbulence closure near solid boundaries with a Lenormand-type subgrid model in the free-stream with a blending function that employs the ratio of estimated inner and outer turbulent length scales. A Mach 2.2 mixing nozzle/diffuser system representative of a gas laser is simulated under a range of exit pressures to assess the ability of the model to predict the dynamics of the shock train. The simulation captures the location of the shock train responsible for pressure recovery but under-predicts the rate of pressure increase. Predicted turbulence production at the wall is found to be highly sensitive to the behavior of the RANS turbulence model. A Mach 2.3, high-Reynolds number, three-dimensional cavity flow is also simulated in order to compute the wavefront aberrations of an optical beam passing thorough the cavity. The cavity geometry is modeled using an immersed boundary method, and an auxiliary flat plate simulation is performed to replicate the effects of the wind-tunnel boundary layer on the computed optical path difference. Pressure spectra extracted on the cavity walls agree with empirical predictions based on Rossiter's formula. Proper orthogonal modes of the wavefront aberrations in a beam originating from the cavity center agree well with experimental data despite uncertainty about in flow turbulence levels and boundary layer thicknesses over the wind tunnel window. Dynamic mode decomposition of a planar wavefront spanning the cavity reveals that wavefront distortions are driven by shear layer oscillations at the Rossiter frequencies; these disturbances create eddy shocklets that propagate into the free-stream, creating additional optical wavefront distortion.

Special Course on Shock-Wave/Boundary-Layer Interactions in Supersonic and Hypersonic Flows (Interactions entre Ondes de Choc et Couches Limites dans les Ecoulements Supersoniques et Hpersoniques).

DTIC Science & Technology

1993-08-01

the station X 125.5. Ftrther as the experimental results catch up with and overtake downstream, the difference between the theoretical and the...showing up in the conical region of the flow but not log Re. in the inception zone. Its spanwise extent grows with increasing shock strength (Fig. 6e) but...compressibilitN on the turbulent Unfortunately, few systematic interaction experiments nature of the interaction up to at least Mach 4. have been

Radiative Viscous Shock Layer Analysis of Fire, Apollo, and PAET Flight Data

NASA Technical Reports Server (NTRS)

Balakrishnan, A.; Park, Chul; Green, Michael J.

1986-01-01

Equilibrium, radiating viscous shock layer solutions are obtained for a number of trajectory points of the Fire II, Apollo 4, and PAET experimental flight vehicles. Convective heating rates calculated by a benchmark code agree well with two engineering correlations, except at high altitudes corresponding to low densities. Calculated radiation intensities are compared with the flight radiometer data and with inviscid flow results. Differences as great as 70% are observed between measured data and the viscous calculations. Because of boundary-layer absorption, viscous effects reduce the intensity to the wall by as much as 30% compared with inviscid intensities. Preliminary chemical and thermal nonequilibrium flow calculations along a stagnation streamline for a PAET trajectory predict an enhancement to the radiation owing to the chemical relaxation. Stagnation point solutions are also presented for future aeroassisted orbital transfer vehicle geometries with nose radii of 0.3-15 m.

Radiative viscous-shock-layer analysis of Fire, Apollo, and PAET flight data

NASA Technical Reports Server (NTRS)

Balakrishnan, A.; Park, C.; Green, M. J.

1985-01-01

Equilibrium, radiating viscous-shock-layer solutions are obtained for a number of trajectory points of the Fire II, Apollo 4, and PAET experimental flight vehicles. Convective heating rates calculated by a benchmark code agree well, except at high altitudes corresponding to low densities, with two engineering correlations. Calculated radiation intensities are compared with the flight radiometer data and with inviscid flow results. Differences as great as 70 percent are observed between measured data and the viscous calculations. Viscous effects reduce the intensity toward the wall, because of boundary-layer absorption, by as much as 30 percent, compared with inviscid intensities. Preliminary chemical and thermal nonequilibrium flow calculations along a stagnation streamline for a PAET trajectory predict enhancement of radiation owing to chemical relaxation. Stagnation point solutions are also presented for future air-assisted orbital transfer vehicle geometries with nose radii ranging from 0.3 to 15 m.

HIFiRE-1 Turbulent Shock Boundary Layer Interaction - Flight Data and Computations

NASA Technical Reports Server (NTRS)

Kimmel, Roger L.; Prabhu, Dinesh

2015-01-01

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory (AFRL) and Australian Defence Science and Technology Organisation (DSTO). This flight contained a cylinder-flare induced shock boundary layer interaction (SBLI). Computations of the interaction were conducted for a number of times during the ascent. The DPLR code used for predictions was calibrated against ground test data prior to exercising the code at flight conditions. Generally, the computations predicted the upstream influence and interaction pressures very well. Plateau pressures on the cylinder were predicted well at all conditions. Although the experimental heat transfer showed a large amount of scatter, especially at low heating levels, the measured heat transfer agreed well with computations. The primary discrepancy between the experiment and computation occurred in the pressures measured on the flare during second stage burn. Measured pressures exhibited large overshoots late in the second stage burn, the mechanism of which is unknown. The good agreement between flight measurements and CFD helps validate the philosophy of calibrating CFD against ground test, prior to exercising it at flight conditions.

Blunt-Body Aerothermodynamic Database from High-Enthalpy CO2 Testing in an Expansion Tunnel

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Prabhu, Dinesh K.; Maclean, Matthew; Dufrene, Aaron

2016-01-01

An extensive database of heating, pressure, and flow field measurements on a 70-deg sphere-cone blunt body geometry in high-enthalpy, CO2 flow has been generated through testing in an expansion tunnel. This database is intended to support development and validation of computational tools and methods to be employed in the design of future Mars missions. The test was conducted in an expansion tunnel in order to avoid uncertainties in the definition of free stream conditions noted in previous studies performed in reflected shock tunnels. Data were obtained across a wide range of test velocity/density conditions that produced various physical phenomena of interest, including laminar and transitional/turbulent boundary layers, non-reacting to completely dissociated post-shock gas composition and shock-layer radiation. Flow field computations were performed at the test conditions and comparisons were made with the experimental data. Based on these comparisons, it is recommended that computational uncertainties on surface heating and pressure, for laminar, reacting-gas environments can be reduced to +/-10% and +/-5%, respectively. However, for flows with turbulence and shock-layer radiation, there were not sufficient validation-quality data obtained in this study to make any conclusions with respect to uncertainties, which highlights the need for further research in these areas.

DSMC Computations for Regions of Shock/Shock and Shock/Boundary Layer Interaction

NASA Technical Reports Server (NTRS)

Moss, James N.

2001-01-01

This paper presents the results of a numerical study of hypersonic interacting flows at flow conditions that include those for which experiments have been conducted in the Calspan-University of Buffalo Research Center (CUBRC) Large Energy National Shock (LENS) tunnel and the ONERA R5Ch low-density wind tunnel. The computations are made with the direct simulation Monte Carlo (DSMC) method of Bird. The focus is on Mach 9.3 to 11.4 flows about flared axisymmetric configurations, both hollow cylinder flares and double cones. The results presented highlight the sensitivity of the calculations to grid resolution, provide results concerning the conditions for incipient separation, and provide information concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Role of Turbulent Prandtl Number on Heat Flux at Hypersonic Mach Number

NASA Technical Reports Server (NTRS)

Xiao, X.; Edwards, J. R.; Hassan, H. A.

2004-01-01

Present simulation of turbulent flows involving shock wave/boundary layer interaction invariably overestimates heat flux by almost a factor of two. One possible reason for such a performance is a result of the fact that the turbulence models employed make use of Morkovin's hypothesis. This hypothesis is valid for non-hypersonic Mach numbers and moderate rates of heat transfer. At hypersonic Mach numbers, high rates of heat transfer exist in regions where shock wave/boundary layer interactions are important. As a result, one should not expect traditional turbulence models to yield accurate results. The goal of this investigation is to explore the role of a variable Prandtl number formulation in predicting heat flux in flows dominated by strong shock wave/boundary layer interactions. The intended applications involve external flows in the absence of combustion such as those encountered in supersonic inlets. This can be achieved by adding equations for the temperature variance and its dissipation rate. Such equations can be derived from the exact Navier-Stokes equations. Traditionally, modeled equations are based on the low speed energy equation where the pressure gradient term and the term responsible for energy dissipation are ignored. It is clear that such assumptions are not valid for hypersonic flows. The approach used here is based on the procedure used in deriving the k-zeta model, in which the exact equations that governed k, the variance of velocity, and zeta, the variance of vorticity, were derived and modeled. For the variable turbulent Prandtl number, the exact equations that govern the temperature variance and its dissipation rate are derived and modeled term by term. The resulting set of equations are free of damping and wall functions and are coordinate-system independent. Moreover, modeled correlations are tensorially consistent and invariant under Galilean transformation. The final set of equations will be given in the paper.

Investigation of viscous/inviscid interaction in transonic flow over airfoils with suction

NASA Technical Reports Server (NTRS)

Vemuru, C. S.; Tiwari, S. N.

1988-01-01

The viscous/inviscid interaction over transonic airfoils with and without suction is studied. The streamline angle at the edge of the boundary layer is used to couple the viscous and inviscid flows. The potential flow equations are solved for the inviscid flow field. In the shock region, the Euler equations are solved using the method of integral relations. For this, the potential flow solution is used as the initial and boundary conditions. An integral method is used to solve the laminar boundary-layer equations. Since both methods are integral methods, a continuous interaction is allowed between the outer inviscid flow region and the inner viscous flow region. To avoid the Goldstein singularity near the separation point the laminar boundary-layer equations are derived in an inverse form to obtain solution for the flows with small separations. The displacement thickness distribution is specified instead of the usual pressure distribution to solve the boundry-layer equations. The Euler equations are solved for the inviscid flow using the finite volume technique and the coupling is achieved by a surface transpiration model. A method is developed to apply a minimum amount of suction that is required to have an attached flow on the airfoil. The turbulent boundary layer equations are derived using the bi-logarithmic wall law for mass transfer. The results are found to be in good agreement with available experimental data and with the results of other computational methods.

A 3-D Navier-Stokes CFD study of turbojet/ramjet nozzle plume interactions at Mach 3.0 and comparison with data

NASA Technical Reports Server (NTRS)

Chang, Ing; Hunter, Louis G.

1995-01-01

Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually consist of a single integrated turboramjet or as in this study, a turbojet housed in an upper bay with a separate ramjet housed in a lower bay. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. One nozzle concept under consideration has a common nozzle, where the plumes from the turbojet and ramjet interact with one another as they expand to ambient conditions. In this paper, the two plumes interact at the end of a common 2-D cowl, when they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. At this condition, the turbojet engine operated at a higher NPR than the ramjet, where the turbojet overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data used to compare with the PARC3D code with k-kl two equation turbulence model. The 2-D and 3-D centerline CFD solutions are in good agreement, but as the CFD solutions approach the outer sidewall, a slight variance occurs. The outer wall boundary layers are thin and do not present much of an interaction, however, where the confluence interaction shocks interact with the thin boundary layer on the outer wall, strong vortices run down each shock causing substantial disturbances in the boundary layer. These disturbances amplify somewhat as they propagate downstream axially from the confluence point. The nozzle coefficient (CFG) is reduced 1/2 percent as a result of this sidewall interaction, from 0.9850 to 0.9807. This three-dimensional reduction is in better agreement with the experimental value of 0.9790.

A method for the direct numerical simulation of hypersonic boundary-layer instability with finite-rate chemistry

DOE Office of Scientific and Technical Information (OSTI.GOV)

Marxen, Olaf, E-mail: olaf.marxen@vki.ac.be; Aeronautics and Aerospace Department, von Karman Institute for Fluid Dynamics, Chaussée de Waterloo, 72, 1640 Rhode-St-Genèse; Magin, Thierry E.

2013-12-15

A new numerical method is presented here that allows to consider chemically reacting gases during the direct numerical simulation of a hypersonic fluid flow. The method comprises the direct coupling of a solver for the fluid mechanical model and a library providing the physio-chemical model. The numerical method for the fluid mechanical model integrates the compressible Navier–Stokes equations using an explicit time advancement scheme and high-order finite differences. This Navier–Stokes code can be applied to the investigation of laminar-turbulent transition and boundary-layer instability. The numerical method for the physio-chemical model provides thermodynamic and transport properties for different gases as wellmore » as chemical production rates, while here we exclusively consider a five species air mixture. The new method is verified for a number of test cases at Mach 10, including the one-dimensional high-temperature flow downstream of a normal shock, a hypersonic chemical reacting boundary layer in local thermodynamic equilibrium and a hypersonic reacting boundary layer with finite-rate chemistry. We are able to confirm that the diffusion flux plays an important role for a high-temperature boundary layer in local thermodynamic equilibrium. Moreover, we demonstrate that the flow for a case previously considered as a benchmark for the investigation of non-equilibrium chemistry can be regarded as frozen. Finally, the new method is applied to investigate the effect of finite-rate chemistry on boundary layer instability by considering the downstream evolution of a small-amplitude wave and comparing results with those obtained for a frozen gas as well as a gas in local thermodynamic equilibrium.« less

Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Kim, Hyun D.; Frate, Franco C.

2001-01-01

A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

In-Flight Boundary-Layer Transition on a Large Flat Plate at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Banks, Daniel W.; Fredericks, Michael Alan; Tracy, Richard R.; Matisheck, Jason R.; Vanecek, Neal D.

2012-01-01

A flight experiment was conducted to investigate the pressure distribution, local flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.0. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. The tests used a F-15B testbed aircraft with a bottom centerline mounted test fixture. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating for future laminar flow flight tests employing infrared thermography. Boundary-layer transition was captured using an onboard infrared imaging system. The infrared imagery was captured in both analog and digital formats. Surface pressures were measured with electronically scanned pressure modules connected to 60 surface-mounted pressure orifices. The local flow field was measured with five 5-hole conical probes mounted near the leading edge of the test fixture. Flow field measurements revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration. The infrared imaging system was able to capture shock wave impingement on the surface of the flat plate in addition to indicating laminar-to-turbulent boundary-layer transition.

An experimental investigation of the impingement of a planar shock wave on an axisymmetric body at Mach 3

NASA Technical Reports Server (NTRS)

Brosh, A.; Kussoy, M. I.

1983-01-01

An experimental study of the flow caused by a planar shock wave impinging obliquely on a cylinder is presented. The complex three dimensional shock wave and boundary layer interaction occurring in practical problems, such as the shock wave impingement from the shuttle nose on an external fuel tank, and store carriage interference on a supersonic tactical aircraft were investigated. A data base for numerical computations of complex flows was also investigated. The experimental techniques included pressure measurements and oil flow patterns on the surface of the cylinder, and shadowgraphs and total and static pressure surveys on the leeward and windward planes of symmetry. The complete data is presented in tabular form. The results reveal a highly complex flow field with two separation zones, regions of high crossflow, and multiple reflected shocks and expansion fans.

Planar Reflection of Detonations Waves

NASA Astrophysics Data System (ADS)

Damazo, Jason; Shepherd, Joseph

2012-11-01

An experimental study examining normally reflected gaseous detonation waves is undertaken so that the physics of reflected detonations may be understood. Focused schlieren visualization is used to describe the boundary layer development behind the incident detonation wave and the nature of the reflected shock wave. Reflected shock wave bifurcation-which has received extensive study as it pertains to shock tube performance-is predicted by classical bifurcation theory, but is not observed in the present study for undiluted hydrogen-oxygen and ethylene-oxygen detonation waves. Pressure and thermocouple gauges are installed in the floor of the detonation tube so as to examine both the wall pressure and heat flux. From the pressure results, we observe an inconsistency between the measured reflected shock speed and the measured reflected shock strength with one dimensional flow predictions confirming earlier experiments performed in our laboratory. This research is sponsored by the DHS through the University of Rhode Island, Center of Excellence for Explosives Detection.

Properties of the Equatorial Magnetotail Flanks ˜50-200 RE Downtail

NASA Astrophysics Data System (ADS)

Artemyev, A. V.; Angelopoulos, V.; Runov, A.; Wang, C.-P.; Zelenyi, L. M.

2017-12-01

In space, thin boundaries separating plasmas with different properties serve as a free energy source for various plasma instabilities and determine the global dynamics of large-scale systems. In planetary magnetopauses and shock waves, classical examples of such boundaries, the magnetic field makes a significant contribution to the pressure balance and plasma dynamics. The configuration and properties of such boundaries have been well investigated and modeled. However, much less is known about boundaries that form between demagnetized plasmas where the magnetic field is not important for pressure balance. The most accessible example of such a plasma boundary is the equatorial boundary layer of the Earth's distant magnetotail. Rather, limited measurements since its first encounter in the late 1970s by the International Sun-Earth Explorer-3 spacecraft revealed the basic properties of this boundary, but its statistical properties and structure have not been studied to date. In this study, we use Geotail and Acceleration, Reconnection, Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) missions to investigate the equatorial boundary layer from lunar orbit (˜55 Earth radii, RE, downtail) to as far downtail as ˜200 RE. Although the magnetic field has almost no effect on the structure of the boundary layer, the layer separates well the hot, rarefied plasma sheet from dense cold magnetosheath plasmas. We suggest that the most important role in plasma separation is played by polarization electric fields, which modify the efficiency of magnetosheath ion penetration into the plasma sheet. We also show that the total energies (bulk flow plus thermal) of plasma sheet ions and magnetosheath ions are very similar; that is, magnetosheath ion thermalization (e.g., via ion scattering by magnetic field fluctuations) is sufficient to produce hot plasma sheet ions without any additional acceleration.

Shock-Wave/Boundary-Layer Interactions in Hypersonic Low Density Flows

NASA Technical Reports Server (NTRS)

Moss, James N.; Olejniczak, Joseph

2004-01-01

Results of numerical simulations of Mach 10 air flow over a hollow cylinder-flare and a double-cone are presented where viscous effects are significant. The flow phenomena include shock-shock and shock- boundary-layer interactions with accompanying flow separation, recirculation, and reattachment. The purpose of this study is to promote an understanding of the fundamental gas dynamics resulting from such complex interactions and to clarify the requirements for meaningful simulations of such flows when using the direct simulation Monte Carlo (DSMC) method. Particular emphasis is placed on the sensitivity of computed results to grid resolution. Comparisons of the DSMC results for the hollow cylinder-flare (30 deg.) configuration are made with the results of experimental measurements conducted in the ONERA RSCh wind tunnel for heating, pressure, and the extent of separation. Agreement between computations and measurements for various quantities is good except that for pressure. For the same flow conditions, the double- cone geometry (25 deg.- 65 deg.) produces much stronger interactions, and these interactions are investigated numerically using both DSMC and Navier-Stokes codes. For the double-cone computations, a two orders of magnitude variation in free-stream density (with Reynolds numbers from 247 to 24,7 19) is investigated using both computational methods. For this range of flow conditions, the computational results are in qualitative agreement for the extent of separation with the DSMC method always predicting a smaller separation region. Results from the Navier-Stokes calculations suggest that the flow for the highest density double-cone case may be unsteady; however, the DSMC solution does not show evidence of unsteadiness.

Viscous Forces in Velocity Boundary Layers around Planetary Ionospheres.

PubMed

Pérez-De-Tejada

1999-11-01

A discussion is presented to examine the role of viscous forces in the transport of solar wind momentum to the ionospheric plasma of weakly magnetized planets (Venus and Mars). Observational data are used to make a comparison of the Reynolds and Maxwell stresses that are operative in the interaction of the solar wind with local plasma (planetary ionospheres). Measurements show the presence of a velocity boundary layer formed around the flanks of the ionosphere where the shocked solar wind has reached super-Alfvénic speeds. It is found that the Reynolds stresses in the solar wind at that region can be larger than the Maxwell stresses and thus are necessary in the local acceleration of the ionospheric plasma. From an order-of-magnitude calculation of the Reynolds stresses, it is possible to derive values of the kinematic viscosity and the Reynolds number that are suitable to the gyrotropic motion of the solar wind particles across the boundary layer. The value of the kinematic viscosity is comparable to those inferred from studies of the transport of solar wind momentum to the earth's magnetosphere and thus suggest a common property of the solar wind around planetary obstacles. Similar conditions could also be applicable to velocity boundary layers formed in other plasma interaction problems in astrophysics.

Approximate convective heating equations for hypersonic flows

NASA Technical Reports Server (NTRS)

Zoby, E. V.; Moss, J. N.; Sutton, K.

1979-01-01

Laminar and turbulent heating-rate equations appropriate for engineering predictions of the convective heating rates about blunt reentry spacecraft at hypersonic conditions are developed. The approximate methods are applicable to both nonreacting and reacting gas mixtures for either constant or variable-entropy edge conditions. A procedure which accounts for variable-entropy effects and is not based on mass balancing is presented. Results of the approximate heating methods are in good agreement with existing experimental results as well as boundary-layer and viscous-shock-layer solutions.

Observation of a brine layer on an ice surface with an environmental scanning electron microscope at higher pressures and temperatures.

PubMed

Krausko, Ján; Runštuk, Jiří; Neděla, Vilém; Klán, Petr; Heger, Dominik

2014-05-20

Observation of a uranyl-salt brine layer on an ice surface using backscattered electron detection and ice surface morphology using secondary-electron detection under equilibrium conditions was facilitated using an environmental scanning electron microscope (ESEM) at temperatures above 250 K and pressures of hundreds of Pa. The micrographs of a brine layer over ice grains prepared by either slow or shock freezing provided a complementary picture of the contaminated ice grain boundaries. Fluorescence spectroscopy of the uranyl ions in the brine layer confirmed that the species exists predominately in the solvated state under experimental conditions of ESEM.

Experimental investigation of tangential blowing for control of the strong shock boundary layer interaction on inlet ramps

NASA Technical Reports Server (NTRS)

Schwendemann, M. F.

1981-01-01

A 0.165-scale isolated inlet model was tested in the NASA Lewis Research Center 8-ft by 6-ft Supersonic Wind Tunnel. Ramp boundary layer control was provided by tangential blowing from a row of holes in an aft-facing step set into the ramp surface. Testing was performed at Mach numbers from 1.36 to 1.96 using both cold and heated air in the blowing system. Stable inlet flow was achieved at all Mach numbers. Blowing hole geometry was found to be significant at 1.96M. Blowing air temperature was found to have only a small effect on system performance. High blowing levels were required at the most severe test conditions.

An experimental and computational investigation of the flow field about a transonic airfoil in supercritical flow with turbulent boundary-layer separation

NASA Technical Reports Server (NTRS)

Rubesin, M. W.; Okuno, A. F.; Levy, L. L., Jr.; Mcdevitt, J. B.; Seegmiller, H. L.

1976-01-01

A combined experimental and computational research program is described for testing and guiding turbulence modeling within regions of separation induced by shock waves incident in turbulent boundary layers. Specifically, studies are made of the separated flow the rear portion of an 18%-thick circular-arc airfoil at zero angle of attack in high Reynolds number supercritical flow. The measurements include distributions of surface static pressure and local skin friction. The instruments employed include highfrequency response pressure cells and a large array of surface hot-wire skin-friction gages. Computations at the experimental flow conditions are made using time-dependent solutions of ensemble-averaged Navier-Stokes equations, plus additional equations for the turbulence modeling.

Computational Modeling and Validation for Hypersonic Inlets

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.

1996-01-01

Hypersonic inlet research activity at NASA is reviewed. The basis for the paper is the experimental tests performed with three inlets: the NASA Lewis Research Center Mach 5, the McDonnell Douglas Mach 12, and the NASA Langley Mach 18. Both three-dimensional PNS and NS codes have been used to compute the flow within the three inlets. Modeling assumptions in the codes involve the turbulence model, the nature of the boundary layer, shock wave-boundary layer interaction, and the flow spilled to the outside of the inlet. Use of the codes and the experimental data are helping to develop a clearer understanding of the inlet flow physics and to focus on the modeling improvements required in order to arrive at validated codes.

Effect of back-pressure forcing on shock train structures in rectangular channels

NASA Astrophysics Data System (ADS)

Gnani, F.; Zare-Behtash, H.; White, C.; Kontis, K.

2018-04-01

The deceleration of a supersonic flow to the subsonic regime inside a high-speed engine occurs through a series of shock waves, known as a shock train. The generation of such a flow structure is due to the interaction between the shock waves and the boundary layer inside a long and narrow duct. The understanding of the physics governing the shock train is vital for the improvement of the design of high-speed engines and the development of flow control strategies. The present paper analyses the sensitivity of the shock train configuration to a back-pressure variation. The complex characteristics of the shock train at an inflow Mach number M = 2 in a channel of constant height are investigated with two-dimensional RANS equations closed by the Wilcox k-ω turbulence model. Under a sinusoidal back-pressure variation, the simulated results indicate that the shock train executes a motion around its mean position that deviates from a perfect sinusoidal profile with variation in oscillation amplitude, frequency, and whether the pressure is first increased or decreased.

Performance predictions for an SSME configuration with an enlarged throat

NASA Technical Reports Server (NTRS)

Nickerson, G. R.; Dang, L. D.

1985-01-01

The Two Dimensional Kinetics (TDK) computer program that was recently developed for NASA was used to predict the performance of a Large Throat Configuration of the Space Shuttle Main Engine (SSME). Calculations indicate that the current design SSME contains a shock wave that is induced by the nozzle wall shape. In the Large Throat design an even stronger shock wave is predicted. Because of the presence of this shock wave, earlier performance predictions that have neglected shock wave effects have been questioned. The JANNAF thrust chamber performance prediction procedures given in a reference were applied. The analysis includes the effects of two dimensional reacting flow with a shock wave. The effects of the boundary layer with a regenatively cooled wall are also included. A Purdue computer program was used to compute axially symmetric supersonic nozzle flows with an induced shock, but is restricted to flows with a constant ratio of specific heats. Thus, the TDK program was also run with ths assumption and the results of the two programs were compared.

Experimental Investigation of 'Transonic Resonance' with Convergent-Divergent Nozzles

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.; Dahl, M. D.; Bencic, T. J.; Zaman, Khairul (Technical Monitor)

2001-01-01

Convergent-divergent nozzles, when run at pressure ratios lower than the design value, often undergo a flow resonance accompanied by the emission of acoustic tones. The phenomenon, different in characteristics from conventional 'screech' tones, has been studied experimentally. Unlike screech, the frequency increases with increasing supply pressure. There is a 'staging' behavior; 'odd harmonic' stages resonate at lower pressures while the fundamental occurs in a range of higher pressures corresponding to a fully expanded Mach number (M(sub j)) around unity. The frequency (f(sub N)) variation with M(sub j) depends on the half angle-of-divergence (theta) of the nozzle. At smaller theta, the slope of f(sub N) versus M(sub j) curve becomes steeper. The resonance involves standing waves and is driven by unsteady shock/boundary layer interaction. The distance between the foot of the shock and the nozzle exit imposes the lengthscale (L'). The fundamental corresponds to a quarterwave resonance, the next stage at a lower supply pressure corresponds to a three-quarter-wave resonance, and so on. The principal trends in the frequency variation are explained simply from the characteristic variation of the length-scale L'. Based on the data, correlation equations are provided for the prediction of f(sub N). A striking feature is that tripping of the boundary layer near the nozzle's throat tends to suppress the resonance. In a practical nozzle a tendency for the occurrence of the phenomenon is thought to be a source of 'internal noise'; thus, there is a potential for noise benefit simply by appropriate boundary layer tripping near the nozzle's throat.

Solute transport by flow yields geometric shocks in shape evolution

NASA Astrophysics Data System (ADS)

Huang, Jinzi (Mac); Davies Wykes, Megan; Hajjar, George; Ristroph, Leif; Shelley, Michael

2017-11-01

Geological processes such as erosion and dissolution of surfaces often lead to striking shapes with strikingly sharp features. We present observations of such features forming in dissolution under gravity. In our experiment, a dissolving body with initially smooth surface evolves into an increasingly sharp needle shape. A mathematical model of its shape dynamics, derived from a boundary layer theory, predicts that a geometric shock forms at the tip of dissolved body, with the tip curvature becoming infinite in finite time. We further discuss the model's application to similar processes, such as flow driven erosion which can yield corners.

Ultrasonic absorption characteristics of porous carbon-carbon ceramics with random microstructure for passive hypersonic boundary layer transition control

NASA Astrophysics Data System (ADS)

Wagner, Alexander; Hannemann, Klaus; Kuhn, Markus

2014-06-01

Preceding studies in the high enthalpy shock tunnel Göttingen of the German Aerospace Center (DLR) revealed that carbon fibre reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities leading to a delay of boundary layer transition onset. To assess the ultrasonic absorption properties of the material, a test rig was set up to measure the reflection coefficient at ambient pressures ranging from 0.1 × 105 to 1 × 105 Pa. For the first time, broadband ultrasonic sound transducers with resonance frequencies of up to 370 kHz were applied to directly cover the frequency range of interest with respect to the second-mode instabilities observed in previous experiments. The reflection of ultrasonic waves from three flat plate test samples with a porous layer thickness between 5 and 30 mm was investigated and compared to an ideally reflecting surface. C/C was found to absorb up to 19 % of the acoustic power transmitted towards the material. The absorption characteristics were investigated theoretically by means of the quasi-homogeneous absorber theory. The experimental results were found to be in good agreement with the theory.

Optimal Growth in Hypersonic Boundary Layers

NASA Technical Reports Server (NTRS)

Paredes, Pedro; Choudhari, Meelan M.; Li, Fei; Chang, Chau-Lyan

2016-01-01

The linear form of the parabolized linear stability equations is used in a variational approach to extend the previous body of results for the optimal, nonmodal disturbance growth in boundary-layer flows. This paper investigates the optimal growth characteristics in the hypersonic Mach number regime without any high-enthalpy effects. The influence of wall cooling is studied, with particular emphasis on the role of the initial disturbance location and the value of the spanwise wave number that leads to the maximum energy growth up to a specified location. Unlike previous predictions that used a basic state obtained from a self-similar solution to the boundary-layer equations, mean flow solutions based on the full Navier-Stokes equations are used in select cases to help account for the viscous- inviscid interaction near the leading edge of the plate and for the weak shock wave emanating from that region. Using the full Navier-Stokes mean flow is shown to result in further reduction with Mach number in the magnitude of optimal growth relative to the predictions based on the self-similar approximation to the base flow.

Dynamic non-equilibrium wall-modeling for large eddy simulation at high Reynolds numbers

NASA Astrophysics Data System (ADS)

Kawai, Soshi; Larsson, Johan

2013-01-01

A dynamic non-equilibrium wall-model for large-eddy simulation at arbitrarily high Reynolds numbers is proposed and validated on equilibrium boundary layers and a non-equilibrium shock/boundary-layer interaction problem. The proposed method builds on the prior non-equilibrium wall-models of Balaras et al. [AIAA J. 34, 1111-1119 (1996)], 10.2514/3.13200 and Wang and Moin [Phys. Fluids 14, 2043-2051 (2002)], 10.1063/1.1476668: the failure of these wall-models to accurately predict the skin friction in equilibrium boundary layers is shown and analyzed, and an improved wall-model that solves this issue is proposed. The improvement stems directly from reasoning about how the turbulence length scale changes with wall distance in the inertial sublayer, the grid resolution, and the resolution-characteristics of numerical methods. The proposed model yields accurate resolved turbulence, both in terms of structure and statistics for both the equilibrium and non-equilibrium flows without the use of ad hoc corrections. Crucially, the model accurately predicts the skin friction, something that existing non-equilibrium wall-models fail to do robustly.

Transient interaction between a reaction control jet and a hypersonic crossflow

NASA Astrophysics Data System (ADS)

Miller, Warrick A.; Medwell, Paul R.; Doolan, Con J.; Kim, Minkwan

2018-04-01

This paper presents a numerical study that focuses on the transient interaction between a reaction control jet and a hypersonic crossflow with a laminar boundary layer. The aim is to better understand the underlying physical mechanisms affecting the resulting surface pressure and control force. Implicit large-eddy simulations were performed with a round, sonic, perfect air jet issuing normal to a Mach 5 crossflow over a flat plate with a laminar boundary layer, at a jet-to-crossflow momentum ratio of 5.3 and a pressure ratio of 251. The pressure distribution induced on the flat plate is unsteady and is influenced by vortex structures that form around the jet. A horseshoe vortex structure forms upstream and consists of six vortices: two quasi-steady vortices and two co-rotating vortex pairs that periodically coalesce. Shear-layer vortices shed periodically and cause localised high pressure regions that convect downstream with constant velocity. A longitudinal counter-rotating vortex pair is present downstream of the jet and is formed from a series of trailing vortices which rotate about a common axis. Shear-layer vortex shedding causes periodic deformation of barrel and bow shocks. This changes the location of boundary layer separation which also affects the normal force on the plate.

A Comparison of Active and Passive Methods for Control of Hypersonic Boundary Layers on Airbreathing Configurations

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Nowak, Robert J.

2003-01-01

Active and passive methods for control of hypersonic boundary layers have been experimentally examined in NASA Langley Research Center wind tunnels on a Hyper-X model. Several configurations for forcing transition using passive discrete roughness elements and active mass addition, or blowing, methods were compared in two hypersonic facilities, the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air tunnels. Heat transfer distributions, obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the nominal Mach 7 flight trajectory of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For the passive roughness examination, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The prior passive roughness study resulted in a swept ramp configuration being selected for the Mach 7 flight vehicle that was scaled to be roughly 0.6 of the calculated boundary layer thickness. For the active jet blowing study, the blowing manifold pressure was systematically varied for each configuration, while monitoring the mass flow, to determine the jet penetration height with schlieren and transition movement with the phosphor system for comparison to the passive results. All the blowing concepts tested were adequate for providing transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model static pressure or higher.

Permanent-Change Thermal Paints for Hypersonic Flight-Test

DTIC Science & Technology

2010-09-24

thermochromic liquid crystals (Ireland et al. 1999, Ireland & Jones 2000), and temperature sensitive paints (Liu & Sullivan 2005), thermal paints are...surfaces and fin-fuselage junctions, shock boundary layer interactions, scramjet combustion chambers and around control thrusters. Thermal paints can...use of discrete wired sensors may also not practical on some locations on a hypersonic vehicle. Thermochromic liquid crystals Coatings of

Wall-temperature effects on the aerodynamics of a hydrogen-fueled transport concept in Mach 8 blowdown and shock tunnels

NASA Technical Reports Server (NTRS)

Penland, J. A.; Marcum, D. C., Jr.; Stack, S. H.

1983-01-01

Results are presented from two separate tests on the same blended wing-body hydrogen fueled transport model at a Mach number of about 8 and a range of Reynolds numbers (based on theoretical body length) of 0.597 x 10 to the 6th power to about 156.22 x 10 to the 6th power. Tests were made in conventional hypersonic blowdown tunnel and a hypersonic shock tunnel at angles of attack of -2 deg to about 8 deg, with an extensive study made at a constant angle of attack of 3 deg. The model boundary-layer flow varied from laminar at the lower Reynolds numbers to predominantly turbulent at the higher Reynolds numbers. Model wall temperatures and stream static temperatures varied widely between the two tests, particularly at the lower Reynolds numbers. These temperature differences resulted in marked variations of the axial-force coefficients between the two tests, due in part to the effects of induced pressure and viscous interaction variations. The normal-force coefficient was essentially independent of Reynolds number. Analysis of results utilized current theoretical computer programs and basic boundary-layer theory.

Laser interferometer skin-friction measurements of crossing-shock-wave/turbulent-boundary-layer interactions

NASA Technical Reports Server (NTRS)

Garrison, T. J.; Settles, G. S.; Narayanswami, N.; Knight, D. D.

1994-01-01

Wall shear stress measurements beneath crossing-shock-wave/turbulent boundary-layer interactions have been made for three interactions of different strengths. The interactions are generated by two sharp fins at symetric angles of attack mounted on a flat plate. The shear stress measurements were made for fin angles of 7 and 11 deg at Mach 3 and 15 deg at Mach 3.85. The measurements were made using a laser interferometer skin-friction meter, a device that determines the wall shear by optically measuring the time rate of thinning of an oil film placed on the test model surface. Results of the measurements reveal high skin-friction coefficients in the vicinity of the fin/plate junction and the presence of quasi-two-dimensional flow separation on the interaction center line. Additionally, two Navier-Stokes computations, one using a Baldwin-Lomax turbulence model and one using a k-epsilon model, are compared with the experimental results for the Mach 3.85, 15-deg interaction case. Although the k-epsilon model did a reasonable job of predicting the overall trend in portions of the skin-friction distribution, neither computation fully captured the physics of the near-surface flow in this complex interaction.

Experimental Database with Baseline CFD Solutions: 2-D and Axisymmetric Hypersonic Shock-Wave/Turbulent-Boundary-Layer Interactions

NASA Technical Reports Server (NTRS)

Marvin, Joseph G.; Brown, James L.; Gnoffo, Peter A.

2013-01-01

A database compilation of hypersonic shock-wave/turbulent boundary layer experiments is provided. The experiments selected for the database are either 2D or axisymmetric, and include both compression corner and impinging type SWTBL interactions. The strength of the interactions range from attached to incipient separation to fully separated flows. The experiments were chosen based on criterion to ensure quality of the datasets, to be relevant to NASA's missions and to be useful for validation and uncertainty assessment of CFD Navier-Stokes predictive methods, both now and in the future. An emphasis on datasets selected was on surface pressures and surface heating throughout the interaction, but include some wall shear stress distributions and flowfield profiles. Included, for selected cases, are example CFD grids and setup information, along with surface pressure and wall heating results from simulations using current NASA real-gas Navier-Stokes codes by which future CFD investigators can compare and evaluate physics modeling improvements and validation and uncertainty assessments of future CFD code developments. The experimental database is presented tabulated in the Appendices describing each experiment. The database is also provided in computer-readable ASCII files located on a companion DVD.

The 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet

NASA Technical Reports Server (NTRS)

Mizukami, M.; Saunders, J. D.

1995-01-01

The supersonic diffuser of a Mach 2.68 bifurcated, rectangular, mixed-compression inlet was analyzed using a three-dimensional (3D) Navier-Stokes flow solver. A two-equation turbulence model, and a porous bleed model based on unchoked bleed hole discharge coefficients were used. Comparisons were made with experimental data, inviscid theory, and two-dimensional Navier-Stokes analyses. The main objective was to gain insight into the inlet fluid dynamics. Examination of the computational results along with the experimental data suggest that the cowl shock-sidewall boundary layer interaction near the leading edge caused a substantial separation in the wind tunnel inlet model. As a result, the inlet performance may have been compromised by increased spillage and higher bleed mass flow requirements. The internal flow contained substantial waves that were not in the original inviscid design. 3D effects were fairly minor for this inlet at on-design conditions. Navier-Stokes analysis appears to be an useful tool for gaining insight into the inlet fluid dynamics. It provides a higher fidelity simulation of the flowfield than the original inviscid design, by taking into account boundary layers, porous bleed, and their interactions with shock waves.

Base Heating Sensitivity Study for a 4-Cluster Rocket Motor Configuration in Supersonic Freestream

NASA Technical Reports Server (NTRS)

Mehta, Manish; Canabal, Francisco; Tashakkor, Scott B.; Smith, Sheldon D.

2011-01-01

In support of launch vehicle base heating and pressure prediction efforts using the Loci-CHEM Navier-Stokes computational fluid dynamics solver, 35 numerical simulations of the NASA TND-1093 wind tunnel test have been modeled and analyzed. This test article is composed of four JP-4/LOX 500 lbf rocket motors exhausting into a Mach 2 - 3.5 wind tunnel at various ambient pressure conditions. These water-cooled motors are attached to a base plate of a standard missile forebody. We explore the base heating profiles for fully coupled finite-rate chemistry simulations, one-way coupled RAMP (Reacting And Multiphase Program using Method of Characteristics)-BLIMPJ (Boundary Layer Integral Matrix Program - Jet Version) derived solutions and variable and constant specific heat ratio frozen flow simulations. Variations in turbulence models, temperature boundary conditions and thermodynamic properties of the plume have been investigated at two ambient pressure conditions: 255 lb/sq ft (simulated low altitude) and 35 lb/sq ft (simulated high altitude). It is observed that the convective base heat flux and base temperature are most sensitive to the nozzle inner wall thermal boundary layer profile which is dependent on the wall temperature, boundary layer s specific energy and chemical reactions. Recovery shock dynamics and afterburning significantly influences convective base heating. Turbulence models and external nozzle wall thermal boundary layer profiles show less sensitivity to base heating characteristics. Base heating rates are validated for the highest fidelity solutions which show an agreement within +/-10% with respect to test data.

Shock wave interactions between slender bodies. Some aspects of three-dimensional shock wave diffraction

NASA Astrophysics Data System (ADS)

Hooseria, S. J.; Skews, B. W.

2017-01-01

A complex interference flowfield consisting of multiple shocks and expansion waves is produced when high-speed slender bodies are placed in close proximity. The disturbances originating from a generator body impinge onto the adjacent receiver body, modifying the local flow conditions over the receiver. This paper aims to uncover the basic gas dynamics produced by two closely spaced slender bodies in a supersonic freestream. Experiments and numerical simulations were used to interpret the flowfield, where good agreement between the predictions and measurements was observed. The numerical data were then used to characterise the attenuation associated with shock wave diffraction, which was found to be interdependent with the bow shock contact perimeter over the receiver bodies. Shock-induced boundary layer separation was observed over the conical and hemispherical receiver bodies. These strong viscous-shock interactions result in double-reflected, as well as double-diffracted shock wave geometries in the interference region, and the diffracting waves progress over the conical and hemispherical receivers' surfaces in "lambda" type configurations. This gives evidence that viscous effects can have a substantial influence on the local bow shock structure surrounding high-speed slender bodies in close proximity.

Single-bubble and multibubble cavitation in water triggered by laser-driven focusing shock waves

NASA Astrophysics Data System (ADS)

Veysset, D.; Gutiérrez-Hernández, U.; Dresselhaus-Cooper, L.; De Colle, F.; Kooi, S.; Nelson, K. A.; Quinto-Su, P. A.; Pezeril, T.

2018-05-01

In this study a single laser pulse spatially shaped into a ring is focused into a thin water layer, creating an annular cavitation bubble and cylindrical shock waves: an outer shock that diverges away from the excitation laser ring and an inner shock that focuses towards the center. A few nanoseconds after the converging shock reaches the focus and diverges away from the center, a single bubble nucleates at the center. The inner diverging shock then reaches the surface of the annular laser-induced bubble and reflects at the boundary, initiating nucleation of a tertiary bubble cloud. In the present experiments, we have performed time-resolved imaging of shock propagation and bubble wall motion. Our experimental observations of single-bubble cavitation and collapse and appearance of ring-shaped bubble clouds are consistent with our numerical simulations that solve a one-dimensional Euler equation in cylindrical coordinates. The numerical results agree qualitatively with the experimental observations of the appearance and growth of large bubble clouds at the smallest laser excitation rings. Our technique of shock-driven bubble cavitation opens interesting perspectives for the investigation of shock-induced single-bubble or multibubble cavitation phenomena in thin liquids.

Holt film wall shear instrumentation for boundary layer transition research

NASA Technical Reports Server (NTRS)

Schneider, Steven P.

1994-01-01

Measurements of the performance of hot-film wall-shear sensors were performed to aid development of improved sensors. The effect of film size and substrate properties on the sensor performance was quantified through parametric studies carried out both electronically and in a shock tube. The results show that sensor frequency response increases with decreasing sensor size, while at the same time sensitivity decreases. Substrate effects were also studied, through parametric variation of thermal conductivity and heat capacity. Early studies used complex dual-layer substrates, while later studies were designed for both single-layer and dual-layer substrates. Sensor failures and funding limitations have precluded completion of the substrate thermal-property tests.

Results and current status of the NPARC alliance validation effort

NASA Technical Reports Server (NTRS)

Towne, Charles E.; Jones, Ralph R.

1996-01-01

The NPARC Alliance is a partnership between the NASA Lewis Research Center (LeRC) and the USAF Arnold Engineering Development Center (AEDC) dedicated to the establishment of a national CFD capability, centered on the NPARC Navier-Stokes computer program. The three main tasks of the Alliance are user support, code development, and validation. The present paper is a status report on the validation effort. It describes the validation approach being taken by the Alliance. Representative results are presented for laminar and turbulent flat plate boundary layers, a supersonic axisymmetric jet, and a glancing shock/turbulent boundary layer interaction. Cases scheduled to be run in the future are also listed. The archive of validation cases is described, including information on how to access it via the Internet.

Performance of a Blunt-lip Side Inlet with Ramp Bleed, Bypass, and a Long Constant-area Duct Ahead of the Engine : Mach Numbers 0.66 and 1.5 to 2.1

NASA Technical Reports Server (NTRS)

Allen, John L

1956-01-01

Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.

An analysis for high Reynolds number inviscid/viscid interactions in cascades

NASA Technical Reports Server (NTRS)

Barnett, Mark; Verdon, Joseph M.; Ayer, Timothy C.

1993-01-01

An efficient steady analysis for predicting strong inviscid/viscid interaction phenomena such as viscous-layer separation, shock/boundary-layer interaction, and trailing-edge/near-wake interaction in turbomachinery blade passages is needed as part of a comprehensive analytical blade design prediction system. Such an analysis is described. It uses an inviscid/viscid interaction approach, in which the flow in the outer inviscid region is assumed to be potential, and that in the inner or viscous-layer region is governed by Prandtl's equations. The inviscid solution is determined using an implicit, least-squares, finite-difference approximation, the viscous-layer solution using an inverse, finite-difference, space-marching method which is applied along the blade surfaces and wake streamlines. The inviscid and viscid solutions are coupled using a semi-inverse global iteration procedure, which permits the prediction of boundary-layer separation and other strong-interaction phenomena. Results are presented for three cascades, with a range of inlet flow conditions considered for one of them, including conditions leading to large-scale flow separations. Comparisons with Navier-Stokes solutions and experimental data are also given.

Driving mechanism of unsteady separation shock motion in hypersonic interactive flow

NASA Technical Reports Server (NTRS)

Dolling, D. S.; Narlo, J. C., II

1987-01-01

Wall pressure fluctuations were measured under the steady separation shock waves in Mach 5 turbulent interactions induced by unswept circular cylinders on a flat plate. The wall temperature was adiabatic. A conditional sampling algorithm was developed to examine the statistics of the shock wave motion. The same algorithm was used to examine data taken in earlier studies in the Princeton University Mach 3 blowdown tunnel. In these earlier studies, hemicylindrically blunted fins of different leading-edge diameters were tested in boundary layers which developed on the tunnel floor and on a flat plate. A description of the algorithm, the reasons why it was developed and the sensitivity of the results to the threshold settings, are discussed. The results from the algorithm, together with cross correlations and power spectral density estimates suggests that the shock motion is driven by the low-frequency unsteadiness of the downstream separated, vortical flow.

Effect of non-equilibrium flow chemistry on the heating distribution over the MESUR forebody during a Martian entry

NASA Technical Reports Server (NTRS)

Chen, Yih-Kang

1992-01-01

Effect of flow field properties on the heating distribution over a 140 deg blunt cone was determined for a Martian atmosphere using Euler, Navier-Stokes (NS), viscous shock layer (VSL), and reacting boundary layer (BLIMPK) equations. The effect of gas kinetics on the flow field and the surface heating distribution were investigated. Gas models with nine species and nine reactions were implemented into the codes. Effects of surface catalysis on the heating distribution were studied using a surface kinetics model having five reactions.

Control of shock-wave boundary layer interaction using steady micro-jets

NASA Astrophysics Data System (ADS)

Verma, S. B.; Manisankar, C.; Akshara, P.

2015-09-01

An experimental investigation was conducted to control the amplitude of shock unsteadiness associated with the interaction induced by a cylindrical protuberance on a flat plate in a Mach 2.18 flow. The control was applied in the form of an array of steady micro air-jets of different configurations with variation in pitch and skew angle of the jets. The effect of air-jet supply pressure on control was also studied. Each of the micro-jet configurations was placed 20 boundary layer thicknesses upstream of the leading edge of the cylinder. The overall interaction is seen to get modified for all control configurations and shows a reduction in both separation- and bow-shock strengths and in triple-point height. A significant reduction in the peak rms value is also observed in the intermittent region of separation for each case. For pitched jets placed in a zig-zag configuration, good control effectiveness is achieved at control pressures similar to the stagnation pressure of the freestream. At higher control pressures, however, their obstruction component increases and if these jets are not spaced sufficiently far apart, the effectiveness of their control begins to drop due to the beginning of spanwise jet-to-jet interaction. On the other hand, pitching or skewing the jets to reduces the obstruction component considerably which at lower control pressures shows lower effectiveness. But at higher control pressure, the effectiveness of these configurations continues to increase unlike the pitched jets.

Evaluation of Full Reynolds Stress Turbulence Models in FUN3D

NASA Technical Reports Server (NTRS)

Dudek, Julianne C.; Carlson, Jan-Renee

2017-01-01

Full seven-equation Reynolds stress turbulence models are a relatively new and promising tool for todays aerospace technology challenges. This paper uses two stress-omega full Reynolds stress models to evaluate challenging flows including shock-wave boundary layer interactions, separation and mixing layers. The Wilcox and the SSGLRR full second-moment Reynolds stress models are evaluated for four problems: a transonic two-dimensional diffuser, a supersonic axisymmetric compression corner, a compressible planar shear layer, and a subsonic axisymmetric jet. Simulation results are compared with experimental data and results using the more commonly used Spalart-Allmaras (SA) one-equation and the Menter Shear Stress Transport (SST) two-equation models.

The Formation of Boundary Clinopyroxenes and Associated Glass Veins in Type B1 CAIs

DOE Office of Scientific and Technical Information (OSTI.GOV)

Paque, J M; Beckett, J R; Ishii, H A

2008-05-18

We used focused ion beam thin section preparation and scanning transmission electron microscopy (FIB/STEM) to examine the interfacial region between spinel and host melilite for three spinel grains, two from the mantle and one from the core of an Allende type B1 inclusion, and a second pair of spinel grains from a type B1 inclusion from the Leoville carbonaceous chondrite. The compositions of boundary clinopyroxenes decorating spinel surfaces are generally consistent with those of coarser clinopyroxenes from the same region of the inclusion, suggesting little movement of spinels between mantle and core regions after the formation of boundary clinopyroxenes. Themore » host melilite displays no anomalous compositions near the interface, and anorthite or other late-stage minerals are not observed, suggesting that crystallization of residual liquid was not responsible for the formation of boundary clinopyroxenes. Allende spinels display either direct spinel-melilite contact or an intervening boundary clinopyroxene between the two phases. In the core, boundary clinopyroxene is mantled by a thin (1-2 {micro}m thick) layer of normally zoned (X{sub Ak} increasing away from the melilite-clinopyroxene contact) melilite with X{sub Ak} matching that of the host melilite at the melilite-melilite contact. In the mantle, X{sub Ak} near boundary spinels is constant. Spinels in a Leoville type B1 inclusion are more complex with boundary clinopyroxene, as observed in Allende, but also variable amounts of glass ({approx}1 {micro}m width), secondary calcite, perovskite, and an unknown Mg-, Al-, OH-rich and Ca-, Si-poor crystalline phase that may be a layered double hydrate. Glass compositions are consistent to first order with a precursor consisting mostly of Mg-carpholite or sudoite with some aluminous diopside. One possible scenario of formation for the glass veins is that open system alteration of melilite produced a porous, hydrated aggregate of Mg-carpholite or sudoite + aluminous diopside that was shock melted and quenched to a glass. The unknown crystalline phase may be a shocked remnant of the precursor phase assemblage but is more likely to have formed later by alteration of the glass. Calcite appears to be an opportunistic fracture filling that postdated all major shock events. Boundary clinopyroxenes probably share a common origin with coarser-grained pyroxenes from the same region of the inclusion. In the mantle, these crystals may represent clinopyroxene crystallized in Ti-rich liquids caused by the direct dissolution of perovskite and an associated Sc-Zr-rich phase or as a reaction product between dissolving perovskite and liquid (i.e., indirect dissolution of perovskite). In the core, any perovskite and associated Ti-enriched liquids that may have originally been present disappeared before the growth of boundary clinopyroxene.« less

Investigation of Three-Dimensional Unsteady Flow Characteristics in Transonic Diffusers

NASA Astrophysics Data System (ADS)

Proshchanka, Dzianis; Yonezawa, Koichi; Tsujimoto, Yoshinobu

Three-dimensional characteristics of unsteady flow in supercritical transonic diffuser are investigated. For various pressure ratios three-dimensional flow containing a normal shock/turbulent boundary layer interaction regions with shockwave and pseudo-shockwaves fluctuating in longitudinal and spanwise directions is observed. Experimental and numerical investigations show details of the flowfield in the vicinity of terminal shock, interaction regions and downstream turbulent unsteady flow. Spectral analysis of pressure fluctuations reveals existence of two characteristic frequencies attributed to the shockwave fluctuation in longitudinal direction for the lower frequency case and acoustic resonance in spanwise direction for the higher one. Vortices appear at each corner in transversal sections modifying the core flow. As a result, size and depth of longitudinal and vertical penetration of separation regions impelled by the terminal shock is either increased or decreased.

Enthalpy effects on hypervelocity boundary layers

NASA Astrophysics Data System (ADS)

Adam, Philippe H.

Shots with air and carbon dioxide were carried out in the T5 shock tunnel at GALCIT to study enthalpy effects on hypervelocity boundary layers. The model tested was a 1-meter long, 5-deg half-angle cone. It was instrumented with 51 chromel-constantan coaxial thermocouples and the surface heat transfer rate was computed to deduce the state of the boundary layer. Transitional boundary layers obtained confirm the stabilizing effect of enthalpy. As the reservoir enthalpy is increased, the transition Reynolds number evaluated at the reference conditions increases. This stabilizing effect is more rapid in gases with lower dissociation energy and it seems to level off when no further dissociation can be achieved. Normalizing the reservoir enthalpy with the edge enthalpy appears to collapse the data for all gases onto a single curve. A similar collapse is obtained when normalizing both the transition location and the reservoir enthalpy with the maximum temperature conditions obtained with BLIMPK, a nonequilibrium boundary layer code. The observation that reference conditions are more appropriate to normalize high enthalpy transition data was taken a step further by comparing the tunnel data with results from a reentry experiment. When the edge conditions are used, the tunnel and flight data are around an order of magnitude apart. This is commonly attributed to high disturbance levels in tunnels that cause the boundary layer to transition early. However, when the reference conditions are used instead, the tunnel and flight data come within striking distance of one another although the trends with enthalpy are reversed. This difference could be due to the cone bending and nose blunting. Experimental laminar heat transfer levels were compared to numerical results obtained with BLIMPK. Results for air indicate that the reactions are probably in nonequilibrium and that the wall is catalytic. The catalycity is seen to yield higher surface heat transfer rates than the noncatalytic and frozen chemistry models. The results for carbon dioxide, however, are inconclusive. This is, perhaps, because of inadequate modeling of the reactions. Experimentally, an anomalous yet repeatable, rise in the laminar heat transfer level can be seen at medium enthalpies in carbon dioxide boundary layers.

A Study of Normal Shock-Wave Turbulent Boundary-Layer Interactions at Mach Numbers of 1.3, 1.4 and 1.5.

DTIC Science & Technology

1982-10-01

Veigas 13used a third technique which was to hold the normal shock wave steady in a supersonic tube by the adjustment of a conical choke downstream of...the equilibrium functioii E (equation (3) as tie eqoi ihri:’ fl -it I lows of East, Sawyer and Nash 2 8 . Although the maximum values of ti wakce :, i...f.. ..l.. .. .. ..t.. fl iC!alaCC CCCCCCCCt t CCCCCifllt CCCCCCCCit C C--- ---- -- C A CCC5C tCC C NC C - f CC1S $ C C A * E~ - - C 2

Diagnostics for Hypersonic Engine Control

DTIC Science & Technology

2013-02-01

weredirected across the flow at the entrance to the isolator just downstream of the facility nozzle . The near-infrared beams were frequency tuned across...the facility nozzle were used to study the dynamics of the shock train structure during these transient combustor events. They revealed the...entropy fluctuations in supersonic boundary layers can be quite short in time – on the order of tens of microseconds. We therefore sought data

Shock Wave / Boundary Layer Interaction Experiment on Control Surface

DTIC Science & Technology

2007-06-01

attachment points to the cold structure of the capsule (see Figure 16, left). Vibrational and acoustical loads are relevant for electronic components. Noise...thermal detector subsystems. Table 1: Summary of infrared technologies considered. Thermal Detectors Quantum Detectors Bolometer Pyrometer InGaAs...holes but a decrease in sensitivity at lower temperature results. Pyrometers are suitable for high temperature measurement, but they respond only to

Aerothermal and Propulsion Ground Testing That Can Be Conducted to Increase Chances for Successful Hypervelocity Flight Experiments

DTIC Science & Technology

2005-10-01

interaction • Turbulence/ flow chemistry plus combustion interaction • Transpiration Cooling and ablation – Ram/Scramjet Technology – Ignition, mixing...turbulence models for separated regions of shock wave/turbulent boundary layer interaction – Modeling turbulence/ flow chemistry /combustion...Minutes FLOW DURATION Flow velocity Reynolds number Mach number Velocity Temperature Vehicle length NoneLengthVelocity Flow Chemistry Total temperature

Dynamic High-Pressure Behavior of Hierarchical Heterogeneous Geological Materials

DTIC Science & Technology

2016-04-01

sandwiched between two 25µm FEP copolymer layers attached to the copper driver plate . The total package thickness with thin-film epoxy on all bonding...public release. 3 OUTLINE Page # ABSTRACT 2 1. BACKGROUND 4 2. CHARACTERISTICS OF SAND INVESTIGATED 8 3. PLATE ...constituents, phases, inter-phase boundaries ; distributions in shock states; as well as the structural evolutions which can result in strain

The Influence of Ablation on Radiative Heating for Earth Entry

NASA Technical Reports Server (NTRS)

Johnston, Christopher O.; Gnoffo, Peter A.; Sutton, Kenneth

2008-01-01

Using the coupled ablation and radiation capability recently included in the LAURA flowfield solver, this paper investigates the influence of ablation on the shock-layer radiative heating for Earth entry. The extension of the HARA radiation model, which provides the radiation predictions in LAURA, to treat a gas consisting of the elements C, H, O, and N is discussed. It is shown that the absorption coefficient of air is increased with the introduction of the C and H elements. A simplified shock layer model is studied to show the impact of temperature, as well as the abundance of C and H, on the net absorption or emission from an ablation contaminated boundary layer. It is found that the ablation species reduce the radiative flux in the vacuum ultraviolet, through increased absorption, for all temperatures. However, in the infrared region of the spectrum, the ablation species increase the radiative flux, through strong emission, for temperatures above 3,000 K. Thus, depending on the temperature and abundance of ablation species, the contaminated boundary layer may either provide a net increase or decrease in the radiative flux reaching the wall. To assess the validity of the coupled ablation and radiation LAURA analysis, a previously analyzed Mars-return case (15.24 km/s), which contains significant ablation and radiation coupling, is studied. Exceptional agreement with previous viscous shock-layer results is obtained. A 40% decrease in the radiative flux is predicted for ablation rates equal to 20% of the free-stream mass flux. The Apollo 4 peak-heating case (10.24 km/s) is also studied. For ablation rates up to 3.4% of the free-stream mass flux, the radiative heating is reduced by up to 19%, while the convective heating is reduced by up to 87%. Good agreement with the Apollo 4 radiometer data is obtained by considering absorption in the radiometer cavity. For both the Mars return and the Apollo 4 cases, coupled radiation alone is found to reduce the radiative heating by 30 60% and the convective heating by less than 5%.

A closed-form analytical model for predicting 3D boundary layer displacement thickness for the validation of viscous flow solvers

NASA Astrophysics Data System (ADS)

Kumar, V. R. Sanal; Sankar, Vigneshwaran; Chandrasekaran, Nichith; Saravanan, Vignesh; Natarajan, Vishnu; Padmanabhan, Sathyan; Sukumaran, Ajith; Mani, Sivabalan; Rameshkumar, Tharikaa; Nagaraju Doddi, Hema Sai; Vysaprasad, Krithika; Sharan, Sharad; Murugesh, Pavithra; Shankar, S. Ganesh; Nejaamtheen, Mohammed Niyasdeen; Baskaran, Roshan Vignesh; Rahman Mohamed Rafic, Sulthan Ariff; Harisrinivasan, Ukeshkumar; Srinivasan, Vivek

2018-02-01

A closed-form analytical model is developed for estimating the 3D boundary-layer-displacement thickness of an internal flow system at the Sanal flow choking condition for adiabatic flows obeying the physics of compressible viscous fluids. At this unique condition the boundary-layer blockage induced fluid-throat choking and the adiabatic wall-friction persuaded flow choking occur at a single sonic-fluid-throat location. The beauty and novelty of this model is that without missing the flow physics we could predict the exact boundary-layer blockage of both 2D and 3D cases at the sonic-fluid-throat from the known values of the inlet Mach number, the adiabatic index of the gas and the inlet port diameter of the internal flow system. We found that the 3D blockage factor is 47.33 % lower than the 2D blockage factor with air as the working fluid. We concluded that the exact prediction of the boundary-layer-displacement thickness at the sonic-fluid-throat provides a means to correctly pinpoint the causes of errors of the viscous flow solvers. The methodology presented herein with state-of-the-art will play pivotal roles in future physical and biological sciences for a credible verification, calibration and validation of various viscous flow solvers for high-fidelity 2D/3D numerical simulations of real-world flows. Furthermore, our closed-form analytical model will be useful for the solid and hybrid rocket designers for the grain-port-geometry optimization of new generation single-stage-to-orbit dual-thrust-motors with the highest promising propellant loading density within the given envelope without manifestation of the Sanal flow choking leading to possible shock waves causing catastrophic failures.

Method of high speed flow field influence and restrain on laser communication

NASA Astrophysics Data System (ADS)

Meng, Li-xin; Wang, Chun-hui; Qian, Cun-zhu; Wang, Shuo; Zhang, Li-zhong

2013-08-01

For laser communication performance which carried by airplane or airship, due to high-speed platform movement, the air has two influences in platform and laser communication terminal window. The first influence is that aerodynamic effect causes the deformation of the optical window; the second one is that a shock wave and boundary layer would be generated. For subsonic within the aircraft, the boundary layer is the main influence. The presence of a boundary layer could change the air density and the temperature of the optical window, which causes the light deflection and received beam spot flicker. Ultimately, the energy hunting of the beam spot which reaches receiving side increases, so that the error rate increases. In this paper, aerodynamic theory is used in analyzing the influence of the optical window deformation due to high speed air. Aero-optics theory is used to analyze the influence of the boundary layer in laser communication link. Based on this, we focused on working on exploring in aerodynamic and aero-optical effect suppression method in the perspective of the optical window design. Based on planning experimental aircraft types and equipment installation location, we optimized the design parameters of the shape and thickness of the optical window, the shape and size of air-management kit. Finally, deformation of the optical window and air flow distribution were simulated by fluid simulation software in the different mach and different altitude fly condition. The simulation results showed that the optical window can inhibit the aerodynamic influence after optimization. In addition, the boundary layer is smoothed; the turbulence influence is reduced, which meets the requirements of the airborne laser communication.

Boundary Layer Control for Hypersonic Airbreathing Vehicles

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Nowak, Robert J.; Horvath, Thomas J.

2004-01-01

Active and passive methods for tripping hypersonic boundary layers have been examined in NASA Langley Research Center wind tunnels using a Hyper-X model. This investigation assessed several concepts for forcing transition, including passive discrete roughness elements and active mass addition (or blowing), in the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air Tunnels. Heat transfer distributions obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the Hyper-X nominal Mach 7 flight test-point of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For passive roughness, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The passive roughness study resulted in a swept ramp configuration, scaled to be roughly 0.6 of the calculated boundary layer thickness, being selected for the Mach 7 flight vehicle. For the active blowing study, the manifold pressure was systematically varied (while monitoring the mass flow) for each configuration to determine the jet penetration height, with schlieren, and transition movement, with the phosphor system, for comparison to the passive results. All the blowing concepts tested, which included various rows of sonic orifices (holes), two- and three-dimensional slots, and random porosity, provided transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model surface static pressure, which is adequate to ensure sonic jets. The present results indicate that the jet penetration height for blowing was roughly half the height required with passive roughness elements for an equivalent amount of transition movement.

Experimental investigation of starting characteristics and wave propagation from a shallow open cavity and its acoustic emission at supersonic speed

NASA Astrophysics Data System (ADS)

Pandian, S.; Desikan, S. L. N.; Niranjan, Sahoo

2018-01-01

Experiments were carried out on a shallow open cavity (L/D = 5) at a supersonic Mach number (M = 1.8) to understand its transient starting characteristics, wave propagation (inside and outside the cavity) during one vortex shedding cycle, and acoustic emission. Starting characteristics and wave propagation were visualized through time resolved schlieren images, while acoustic emissions were captured through unsteady pressure measurements. Results showed a complex shock system during the starting process which includes characteristics of the bifurcated shock system, shock train, flow separation, and shock wave boundary layer interaction. In one vortex shedding cycle, vortex convection from cavity leading edge to cavity trailing edge was observed. Flow features outside the cavity demonstrated the formation and downstream movement of a λ-shock due to the interaction of shock from the cavity leading edge and shock due to vortex and generation of waves on account of shear layer impingement at the cavity trailing edge. On the other hand, interesting wave structures and its propagation were monitored inside the cavity. In one vortex shedding cycle, two waves such as a reflected compression wave from a cavity leading edge in the previous vortex shedding cycle and a compression wave due to the reflection of Mach wave at the cavity trailing edge corner in the current vortex shedding cycle were visualized. The acoustic emission from the cavity indicated that the 2nd to 4th modes/tones are dominant, whereas the 1st mode contains broadband spectrum. In the present studies, the cavity feedback mechanism was demonstrated through a derived parameter coherence coefficient.

Impact wave deposits provide new constraints on the location of the K/T boundary impact

NASA Technical Reports Server (NTRS)

Hildebrand, A. R.; Boynton, W. V.

1988-01-01

All available evidence is consistent with an impact into oceanic crust terminating the Cretaceous Period. Although much of this evidence is incompatible with an endogenic origin, some investigators still feel that a volcanic origin is possible for the Cretaceous/Tertiary (K/T) boundary clay layers. The commonly cited evidence for a large impact stems from delicate clay layers and their components and the impact site has not yet been found. Impact sites have been suggested all over the globe. The impact is felt to have occurred near North America by: the occurrence of a 2 cm thick ejecta layer only at North American locales, the global variation of shocked quartz grain sizes peaking in North America, the global variation of spinel compositions with most refractory compositions occurring in samples from the Pacific region and possibly uniquely severe plant extinctions in the North American region. The K/T boundary interval was investigated as preserved on the banks of the Brazos River, Texas. The K/T fireball and ejecta layers with associated geochemical anomalies were found interbedded with this sequence which apparently allows a temporal resolution 4 orders of magnitude greater than typical K/T boundary sections. A literature search reveals that such coarse deposits are widely preserved at the K/T boundary. Impact wave deposits have not been found elsewhere on the globe, suggesting the impact occurred between North and South America. The coarse deposits preserved in Deep Sea Drilling Project (DSDP) holes 151-3 suggest the impact occurred nearby. Subsequent tectonism has complicated the picture.

Acoustically excited heated jets. 2: In search of a better understanding

NASA Technical Reports Server (NTRS)

Lepicovsky, J.; Ahuja, K. K.; Brown, W. H.; Salikuddin, M.; Morris, P. J.

1988-01-01

The second part of a three-part report on the effects of acoustic excitation on jet mixing includes the results of an experimental investigation directed at resolving the question of poor excitability of some of the heated jets. The theoretical predictions discussed in Part 1 are examined to find explanations for the observed discrepancies between the measured and the predicted results. Additional testing was performed by studying the self excitation of the shock containing hot jets and also by exciting the jet by sound radiated through source tubes located externally around the periphery of the jet. The effects of nozzle-exit boundary layer conditions on jet excitability was also investigated. It is concluded that high-speed, heated jet mixing rates and consequently also the jet excitability strongly depends on nozzle exit boundary layer conditions.

Measurements of density, temperature, and their fluctuations in turbulent supersonic flow using UV laser spectroscopy

NASA Technical Reports Server (NTRS)

Fletcher, Douglas G.; Mckenzie, R. L.

1992-01-01

Nonintrusive measurements of density, temperature, and their turbulent fluctuation levels were obtained in the boundary layer of an unseeded, Mach 2 wind tunnel flow. The spectroscopic technique that was used to make the measurements is based on the combination of laser-induced oxygen fluorescence and Raman scattering by oxygen and nitrogen from the same laser pulse. Results from this demonstration experiment are compared with previous measurements obtained in the same facility using conventional probes and an earlier spectroscopic technique. Densities and temperatures measured with the current technique agree with the previous surveys to within 3 percent and 2 percent, respectively. The fluctuation amplitudes for both variables agree with the measurements obtained using the earlier spectroscopic technique and show evidence of an unsteady, weak shock wave that perturbs the boundary layer.

International Conference on Numerical Methods in Fluid Dynamics, 7th, Stanford University, Stanford and Moffett Field, CA, June 23-27, 1980, Proceedings

NASA Technical Reports Server (NTRS)

Reynolds, W. C. (Editor); Maccormack, R. W.

1981-01-01

Topics discussed include polygon transformations in fluid mechanics, computation of three-dimensional horseshoe vortex flow using the Navier-Stokes equations, an improved surface velocity method for transonic finite-volume solutions, transonic flow calculations with higher order finite elements, the numerical calculation of transonic axial turbomachinery flows, and the simultaneous solutions of inviscid flow and boundary layer at transonic speeds. Also considered are analytical solutions for the reflection of unsteady shock waves and relevant numerical tests, reformulation of the method of characteristics for multidimensional flows, direct numerical simulations of turbulent shear flows, the stability and separation of freely interacting boundary layers, computational models of convective motions at fluid interfaces, viscous transonic flow over airfoils, and mixed spectral/finite difference approximations for slightly viscous flows.

Three-dimensional Features of the Outer Heliosphere Due to Coupling between the Interstellar and Heliospheric Magnetic Field. V. The Bow Wave, Heliospheric Boundary Layer, Instabilities, and Magnetic Reconnection

DOE Office of Scientific and Technical Information (OSTI.GOV)

Pogorelov, N. V.; Heerikhuisen, J.; Roytershteyn, V.

The heliosphere is formed due to interaction between the solar wind (SW) and local interstellar medium (LISM). The shape and position of the heliospheric boundary, the heliopause, in space depend on the parameters of interacting plasma flows. The interplay between the asymmetrizing effect of the interstellar magnetic field and charge exchange between ions and neutral atoms plays an important role in the SW–LISM interaction. By performing three-dimensional, MHD plasma/kinetic neutral atom simulations, we determine the width of the outer heliosheath—the LISM plasma region affected by the presence of the heliosphere—and analyze quantitatively the distributions in front of the heliopause. Itmore » is shown that charge exchange modifies the LISM plasma to such extent that the contribution of a shock transition to the total variation of plasma parameters becomes small even if the LISM velocity exceeds the fast magnetosonic speed in the unperturbed medium. By performing adaptive mesh refinement simulations, we show that a distinct boundary layer of decreased plasma density and enhanced magnetic field should be observed on the interstellar side of the heliopause. We show that this behavior is in agreement with the plasma oscillations of increasing frequency observed by the plasma wave instrument onboard Voyager 1. We also demonstrate that Voyager observations in the inner heliosheath between the heliospheric termination shock and the heliopause are consistent with dissipation of the heliospheric magnetic field. The choice of LISM parameters in this analysis is based on the simulations that fit observations of energetic neutral atoms performed by Interstellar Boundary Explorer .« less

Current Issues in Unsteady Turbomachinery Flows (Images)

NASA Technical Reports Server (NTRS)

Povinelli, Louis

2004-01-01

Among the numerous causes for unsteadiness in turbo machinery flows are turbulence and flow environment, wakes from stationary and rotating vanes, boundary layer separation, boundary layer/shear layer instabilities, presence of shock waves and deliberate unsteadiness for flow control purposes. These unsteady phenomena may lead to flow-structure interactions such as flutter and forced vibration as well as system instabilities such as stall and surge. A major issue of unsteadiness relates to the fact that a fundamental understanding of unsteady flow physics is lacking and requires continued attention. Accurate simulations and sufficient high fidelity experimental data are not available. The Glenn Research Center plan for Engine Component Flow Physics Modeling is part of the NASA 21st Century Aircraft Program. The main components of the plan include Low Pressure Turbine National Combustor Code. The goals, technical output and benefits/impacts of each element are described in the presentation. The specific areas selected for discussion in this presentation are blade wake interactions, flow control, and combustor exit turbulence and modeling.

Depletion of solar wind plasma near a planetary boundary

NASA Technical Reports Server (NTRS)

Zwan, B. J.; Wolf, R. A.

1976-01-01

A mathematical model is presented that describes the squeezing of solar wind plasma out along interplanetary magnetic field lines in the region between the bow shock and the effective planetary boundary (in the case of the earth, the magnetopause). In the absence of local magnetic merging the squeezing process should create a 'depletion layer', a region of very low plasma density just outside the magnetopause. Numerical solutions are obtained for the dimensionless magnetohydrodynamic equations describing this depletion process for the case where the solar wind magnetic field is perpendicular to the solar wind flow direction. For the case of the earth, the theory predicts that the density should be reduced by a factor exceeding 2 in a layer about 700-1300 km thick if the Alfven Mach number in the solar wind, is equal to 8. Scaling of the model calculations to Venus and Mars suggests layer thicknesses about 1/10 and 1/15 those of the earth, respectively, neglecting diffusion and ionospheric effects.

Application of holography to the determination of flow conditions within the rotating blade row of a compressor

NASA Technical Reports Server (NTRS)

Hantman, R. G.; Burr, R. J.; Alwang, W. G.; Williams, M. C.

1973-01-01

The double-pulse, double-exposure holography technique was applied to visualize the flow field within a transonic compressor rotor with a tip speed of 1800 ft/sec. The principal objective was to visualize the shock waves created in the flow field which was supersonic relative to the rotating blade row. The upstream rotor blade bow shocks and, at high speed, the outermost portion of the leading edge passage shock were successfully observed in the holograms. Techniques were devised for locating these shocks in three dimensions, and the results were compared with theoretical predictions. Density changes between the two pulses due to motion of the shocks were large and, therefore, it was not possible to resolve the fringe systems in detail for the 100% speed conditions. However, gross features of the shocks were easily observed, and the upstream shocks were well displayed. In all cases the shock angles were somewhat larger than predicted by theory, and a distinct increase in angle near the outer wall was observed, which may be attributed to endwall boundary layer effects. The location and orientation of the observed leading edge passage shocks were in good agreement with static pressure contours obtained from measurements in the outer casing over the rotor tip.

Flow Visualization by Elastic Light Scattering in the Boundary Layer of a Supersonic Flow

NASA Technical Reports Server (NTRS)

Herring, G. C.; Hillard, Mervin E., Jr.

2000-01-01

We demonstrate instantaneous flow visualization of the boundary layer region of a Mach 2.5 supersonic flow over a flat plate that is interacting with an impinging shock wave. Tests were performed in the Unitary Plan Wind Tunnel (UPWT) at NASA Langley Research Center. The technique is elastic light scattering using 10-nsec laser pulses at 532 nm. We emphasize that no seed material of any kind, including water (H2O), is purposely added to the flow. The scattered light comes from a residual impurity that normally exists in the flow medium after the air drying process. Thus, the technique described here differs from the traditional vapor-screen method, which is typically accomplished by the addition of extra H2O vapor to the airflow. The flow is visualized with a series of thin two-dimensional light sheets (oriented perpendicular to the streamwise direction) that are located at several positions downstream of the leading edge of the model. This geometry allows the direct observation of the unsteady flow structure in the spanwise dimension of the model and also allows the indirect observation of the boundary layer growth in the streamwise dimension.

X-33 Hypersonic Boundary Layer Transition

NASA Technical Reports Server (NTRS)

Berry, Scott A.; Horvath, Thomas J.; Hollis, Brian R.; Thompson, Richard A.; Hamilton, H. Harris, II

1999-01-01

Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

Computations of Axisymmetric Flows in Hypersonic Shock Tubes

NASA Technical Reports Server (NTRS)

Sharma, Surendra P.; Wilson, Gregory J.

1995-01-01

A time-accurate two-dimensional fluid code is used to compute test times in shock tubes operated at supersonic speeds. Unlike previous studies, this investigation resolves the finer temporal details of the shock-tube flow by making use of modern supercomputers and state-of-the-art computational fluid dynamic solution techniques. The code, besides solving the time-dependent fluid equations, also accounts for the finite rate chemistry in the hypersonic environment. The flowfield solutions are used to estimate relevant shock-tube parameters for laminar flow, such as test times, and to predict density and velocity profiles. Boundary-layer parameters such as bar-delta(sub u), bar-delta(sup *), and bar-tau(sub w), and test time parameters such as bar-tau and particle time of flight t(sub f), are computed and compared with those evaluated by using Mirels' correlations. This article then discusses in detail the effects of flow nonuniformities on particle time-of-flight behind the normal shock and, consequently, on the interpretation of shock-tube data. This article concludes that for accurate interpretation of shock-tube data, a detailed analysis of flowfield parameters, using a computer code such as used in this study, must be performed.

Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Hollis, Brian R.; Johnston, Christopher O.; Bose, Deepak; White, Todd R.; Mahzari, Milad

2013-01-01

The Mars Science Laboratory heatshield was designed to withstand a fully turbulent heat pulse based on test results and computational analysis on a pre-flight design trajectory. Instrumentation on the flight heatshield measured in-depth temperatures in the thermal protection system. The data indicate that boundary layer transition occurred at 5 of 7 thermocouple locations prior to peak heating. Data oscillations at 3 pressure measurement locations may also indicate transition. This paper presents the heatshield temperature and pressure data, possible explanations for the timing of boundary layer transition, and a qualitative comparison of reconstructed and computational heating on the as-flown trajectory. Boundary layer Reynolds numbers that are typically used to predict transition are compared to observed transition at various heatshield locations. A uniform smooth-wall transition Reynolds number does not explain the timing of boundary layer transition observed during flight. A roughness-based Reynolds number supports the possibility of transition due to discrete or distributed roughness elements on the heatshield. However, the distributed roughness height would have needed to be larger than the pre-flight assumption. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the leeside shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent calculations on the reconstructed trajectory, indicating that augmentation due to surface roughness probably did not occur. Turbulent heating on the downstream side of the heatshield nose exceeded smooth-wall computations, indicating that roughness may have augmented heating. The stagnation region also experienced heating that exceeded computational levels, but shock layer radiation does not fully explain the differences.

Tetrahedral-Mesh Simulation of Turbulent Flows with the Space-Time Conservative Schemes

NASA Technical Reports Server (NTRS)

Chang, Chau-Lyan; Venkatachari, Balaji; Cheng, Gary C.

2015-01-01

Direct numerical simulations of turbulent flows are predominantly carried out using structured, hexahedral meshes despite decades of development in unstructured mesh methods. Tetrahedral meshes offer ease of mesh generation around complex geometries and the potential of an orientation free grid that would provide un-biased small-scale dissipation and more accurate intermediate scale solutions. However, due to the lack of consistent multi-dimensional numerical formulations in conventional schemes for triangular and tetrahedral meshes at the cell interfaces, numerical issues exist when flow discontinuities or stagnation regions are present. The space-time conservative conservation element solution element (CESE) method - due to its Riemann-solver-free shock capturing capabilities, non-dissipative baseline schemes, and flux conservation in time as well as space - has the potential to more accurately simulate turbulent flows using unstructured tetrahedral meshes. To pave the way towards accurate simulation of shock/turbulent boundary-layer interaction, a series of wave and shock interaction benchmark problems that increase in complexity, are computed in this paper with triangular/tetrahedral meshes. Preliminary computations for the normal shock/turbulence interactions are carried out with a relatively coarse mesh, by direct numerical simulations standards, in order to assess other effects such as boundary conditions and the necessity of a buffer domain. The results indicate that qualitative agreement with previous studies can be obtained for flows where, strong shocks co-exist along with unsteady waves that display a broad range of scales, with a relatively compact computational domain and less stringent requirements for grid clustering near the shock. With the space-time conservation properties, stable solutions without any spurious wave reflections can be obtained without a need for buffer domains near the outflow/farfield boundaries. Computational results for the isotropic turbulent flow decay, at a relatively high turbulent Mach number, show a nicely behaved spectral decay rate for medium to high wave numbers. The high-order CESE schemes offer very robust solutions even with the presence of strong shocks or widespread shocklets. The explicit formulation in conjunction with a close to unity theoretical upper Courant number bound has the potential to offer an efficient numerical framework for general compressible turbulent flow simulations with unstructured meshes.

Program and charts for determining shock tube, and expansion tunnel flow quantities for real air

NASA Technical Reports Server (NTRS)

Miller, C. G., III; Wilder, S. E.

1975-01-01

A computer program in FORTRAN 4 language was written to determine shock tube, expansion tube, and expansion tunnel flow quantities for real-air test gas. This program permits, as input data, a number of possible combinations of flow quantities generally measured during a test. The versatility of the program is enhanced by the inclusion of such effects as a standing or totally reflected shock at the secondary diaphragm, thermochemical-equilibrium flow expansion and frozen flow expansion for the expansion tube and expansion tunnel, attenuation of the flow in traversing the acceleration section of the expansion tube, real air as the acceleration gas, and the effect of wall boundary layer on the acceleration section air flow. Charts which provide a rapid estimation of expansion tube performance prior to a test are included.

Stratigraphy and sedimentology of the K/T boundary deposit in Haiti

NASA Technical Reports Server (NTRS)

Carey, S.; Sigurdsson, H.; Dhondt, S.; Espindola, J. M.

1993-01-01

The K/T boundary sequence is exposed in uplifted carbonate sediments of the southwest peninsula of Haiti. It is found at 15 localities within the Beloc formation, a sequence of limestone and marls interpreted as a monoclinal nappe structure thrust to the north. This tectonic deformation has affected the K/T boundary deposit to varying degrees. In some cases the less competent K/T deposit has acted as a slip plane leading to extensive shearing of the boundary layer, as well as duplication of the section. The presence of glassy tektites, shocked quartz, and an Ir anomaly directly link the deposit to a bolide impact. Stratigraphic and sedimentological features of the tripartite sequence indicate that it was formed by deposition from ballistic fallout of coarse tektites, emplacement of particle gravity flows and fine grained fallout of widely dispersed impact ejecta.

European Science Notes Information Bulletin Reports on Current European/ Middle Eastern Science

DTIC Science & Technology

1991-06-01

particularly those that involve shock wave/boundary layer cell-centered, finite-volume, explicit, Runge-Kutta interactions , still prov;de considerble...aircraft configuration attributed to using an interactive vcual grid generation was provided by A. Bocci and A. Baxendale, the Aircraft system developed...the surface pressure the complex problem of wing/body/pylon/store distributions with and without the mass flow through the interaction . Reasonable

Effects of Convoluted Divergent Flap Contouring on the Performance of a Fixed-Geometry Nonaxisymmetric Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Hunter, Craig A.

1999-01-01

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the effects of convoluted divergent-flap contouring on the internal performance of a fixed-geometry, nonaxisymmetric, convergent-divergent exhaust nozzle. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and four convoluted configurations. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at overexpanded conditions. Convoluted configurations were found to significantly reduce, and in some cases totally alleviate separation at overexpanded conditions. This result was attributed to the ability of convoluted contouring to energize and improve the condition of the nozzle boundary layer. Separation alleviation offers potential for installed nozzle aeropropulsive (thrust-minus-drag) performance benefits by reducing drag at forward flight speeds, even though this may reduce nozzle thrust ratio as much as 6.4% at off-design conditions. At on-design conditions, nozzle thrust ratio for the convoluted configurations ranged from 1% to 2.9% below the baseline configuration; this was a result of increased skin friction and oblique shock losses inside the nozzle.

Results of a flow field survey conducted using the 0.0175 scale orbiter model 29-0 in AEDC VKF tunnel B during test OH52. [atmospheric entry simulation

NASA Technical Reports Server (NTRS)

Herrera, B. J.

1976-01-01

Static pressure data and flow field surveys of the boundary layer and shock layer on the lower surface of a 0.0175 scale model of the space shuttle orbiter were obtained in a hypersonic wind tunnel. The tests were conducted at Mach number 7.9 and Reynolds number based on the model length of 1.3 x 1 million to simulate atmospheric entry. Twenty-six stations were surveyed at 30 and 35 degree angles of attack.

A Numerical Investigation of the Burnett Equations Based on the Second Law

NASA Technical Reports Server (NTRS)

Comeaux, Keith A.; Chapman, Dean R.; MacCormack, Robert W.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

The Burnett equations have been shown to potentially violate the second law of thermodynamics. The objective of this investigation is to correlate the numerical problems experienced by the Burnett equations to the negative production of entropy. The equations have had a long history of numerical instability to small wavelength disturbances. Recently, Zhong corrected the instability problem and made solutions attainable for one dimensional shock waves and hypersonic blunt bodies. Difficulties still exist when attempting to solve hypersonic flat plate boundary layers and blunt body wake flows, however. Numerical experiments will include one-dimensional shock waves, quasi-one dimensional nozzles, and expanding Prandlt-Meyer flows and specifically examine the entropy production for these cases.

Center manifolds for a class of degenerate evolution equations and existence of small-amplitude kinetic shocks

NASA Astrophysics Data System (ADS)

Pogan, Alin; Zumbrun, Kevin

2018-06-01

We construct center manifolds for a class of degenerate evolution equations including the steady Boltzmann equation and related kinetic models, establishing in the process existence and behavior of small-amplitude kinetic shock and boundary layers. Notably, for Boltzmann's equation, we show that elements of the center manifold decay in velocity at near-Maxwellian rate, in accord with the formal Chapman-Enskog picture of near-equilibrium flow as evolution along the manifold of Maxwellian states, or Grad moment approximation via Hermite polynomials in velocity. Our analysis is from a classical dynamical systems point of view, with a number of interesting modifications to accommodate ill-posedness of the underlying evolution equation.

A Factor Affecting Transonic Leading-edge Flow Separation

NASA Technical Reports Server (NTRS)

Wood, George P; Gooderum, Paul B

1956-01-01

A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock-wave - boundary layer interaction are proposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. Some experimental evidence in support of the importance of this factor is presented.

Recalibration of the Shear Stress Transport Model to Improve Calculation of Shock Separated Flows

NASA Technical Reports Server (NTRS)

Georgiadis, Nicholas J.; Yoder, Dennis A.

2013-01-01

The Menter Shear Stress Transport (SST) k . turbulence model is one of the most widely used two-equation Reynolds-averaged Navier-Stokes turbulence models for aerodynamic analyses. The model extends Menter s baseline (BSL) model to include a limiter that prevents the calculated turbulent shear stress from exceeding a prescribed fraction of the turbulent kinetic energy via a proportionality constant, a1, set to 0.31. Compared to other turbulence models, the SST model yields superior predictions of mild adverse pressure gradient flows including those with small separations. In shock - boundary layer interaction regions, the SST model produces separations that are too large while the BSL model is on the other extreme, predicting separations that are too small. In this paper, changing a1 to a value near 0.355 is shown to significantly improve predictions of shock separated flows. Several cases are examined computationally and experimental data is also considered to justify raising the value of a1 used for shock separated flows.

Prediction of space shuttle fluctuating pressure environments, including rocket plume effects

NASA Technical Reports Server (NTRS)

Plotkin, K. J.; Robertson, J. E.

1973-01-01

Preliminary estimates of space shuttle fluctuating pressure environments have been made based on prediction techniques developed by Wyle Laboratories. Particular emphasis has been given to the transonic speed regime during launch of a parallel-burn space shuttle configuration. A baseline configuration consisting of a lightweight orbiter and monolithic SRB, together with a typical flight trajectory, have been used as models for the predictions. Critical fluctuating pressure environments are predicted at transonic Mach numbers. Comparisons between predicted environments and wind tunnel test results, in general, showed good agreement. Predicted one-third octave band spectra for the above environments were generally one of three types: (1) attached turbulent boundary layer spectra (typically high frequencies); (2) homogeneous separated flow and shock-free interference flow spectra (typically intermediate frequencies); and (3) shock-oscillation and shock-induced interference flow spectra (typically low frequencies). Predictions of plume induced separated flow environments were made. Only the SRB plumes are important, with fluctuating levels comparable to compression-corner induced separated flow shock oscillation.

DSMC Simulation of Separated Flows About Flared Bodies at Hypersonic Conditions

NASA Technical Reports Server (NTRS)

Moss, James N.

2000-01-01

This paper describes the results of a numerical study of interacting hypersonic flows at conditions that can be produced in ground-based test facilities. The computations are made with the direct simulation Monte Carlo (DSMC) method of Bird. The focus is on Mach 10 flows about flared axisymmetric configurations, both hollow cylinder flares and double cones. The flow conditions are those for which experiments have been or will be performed in the ONERA R5Ch low-density wind tunnel and the Calspan-University of Buffalo Research Center (CUBRC) Large Energy National Shock (LENS) tunnel. The range of flow conditions, model configurations, and model sizes provides a significant range of shock/shock and shock/boundary layer interactions at low Reynolds number conditions. Results presented will highlight the sensitivity of the calculations to grid resolution, contrast the differences in flow structure for hypersonic cold flows and those of more energetic but still low enthalpy flows, and compare the present results with experimental measurements for surface heating, pressure, and extent of separation.

Review Of Existing Facilities: Shock Tunnels, Part of AIAA Short Course On Aerothermodynamic Facilities and Applications

NASA Technical Reports Server (NTRS)

Bogdanoff, David W.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

This review is divided into two main sections. The first section described the various types of shock tunnel facilities - reflected shock tunnels, non-reflected shock tunnels and expansion tubes/tunnels. Driver technology is then described, followed by a discussion of the performance obtainable from various driver-driven combinations. A survey of a number of facilities is then presented. The second part of the review deals with details of the operation of the facilities. Operation of combustion drivers, electrically heated drivers and piston compression drivers is discussed in some detail. Main diaphragm break techniques are discussed, with particular attention being paid to maintaining the integrity of the diaphragm petals. Secondary diaphragm techniques are discussed. Phenomena which limit test time are discussed and a number of techniques to increase test time are presented. Contamination of the flow with material ablated from the wall is discussed along with the relative suitability of various materials for lining the tubes and nozzle. Finally, boundary layer effects in shock tunnels and expansion tubes are discussed.

Using second-sound shock waves to probe the intrinsic critical velocity of liquid helium II

NASA Technical Reports Server (NTRS)

Turner, T. N.

1983-01-01

A critical velocity truly intrinsic to liquid helium II is experimentally sought in the bulk fluid far from the apparatus walls. Termed the 'fundamental critical velocity,' it necessarily is caused by mutual interactions which operate between the two fluid components and which are activated at large relative velocities. It is argued that flow induced by second-sound shock waves provides the ideal means by which to activate and isolate the fundamental critical velocity from other extraneous fluid-wall interactions. Experimentally it is found that large-amplitude second-sound shock waves initiate a breakdown in the superfluidity of helium II, which is dramatically manifested as a limit to the maximum attainable shock strength. This breakdown is shown to be caused by a fundamental critical velocity. Secondary effects include boiling for ambient pressures near the saturated vapor pressure or the formation of helium I boundary layers at higher ambient pressures. When compared to the intrinsic critical velocity discovered in highly restricted geometries, the shock-induced critical velocity displays a similar temperature dependence and is the same order of magnitude.

MHD heat flux mitigation in hypersonic flow around a blunt body with ablating surface

NASA Astrophysics Data System (ADS)

Bityurin, V. A.; Bocharov, A. N.

2018-07-01

One of the possible applications of magnetohydrodynamic flow control is considered. Namely, the surface heat flux mitigation by means of magnetohydrodynamic (MHD) interaction in hypersonic flow around a blunt body. The 2D computational model realizes a coupled solution of chemically non-equilibrium ionized airflow in magnetic field. Heat- and mass-transfer due to the ablation of materials from the body surface is taken into account. Two cases of free-stream flow conditions are considered: moderate free-stream velocity (7500 m s‑1) case and high free-stream velocity (11 000 m s‑1) case. It is shown that the first flow case results in moderate ionization in the shock layer, while the second flow case results in high ionization. In the first case, the Hall effect is significant, and effective electrical conductivity in the shock layer is rather low. In the second case, the Hall effect reduces, and effective conductivity is high. Even if the Hall effect is strong, as in the first case, intensive MHD deceleration of the flow behind the shock is provided due to the presence of insulating boundaries, the bow shock front and non-conductive wall of the blunt body. In the second case, high effective conductivity provides a high intensity of MHD flow deceleration. In both cases, a strong effect of MHD interaction on the flow structure is observed. As a consequence, a noticeable reduction of the surface heat flux is revealed for reasonable values of magnetic induction. The new treatment of mechanism for the surface heat flux reduction is proposed, which is different from commonly used one assuming that MHD interaction increases the bow shock stand-off distance, and, consequently results in a decrease of the mean temperature drop across the shock layer. The new effect of ‘saturation of heat flux’ is discussed.

Numerical techniques for the solution of the compressible Navier-Stokes equations and implementation of turbulence models. [separated turbulent boundary layer flow problems

NASA Technical Reports Server (NTRS)

Baldwin, B. S.; Maccormack, R. W.; Deiwert, G. S.

1975-01-01

The time-splitting explicit numerical method of MacCormack is applied to separated turbulent boundary layer flow problems. Modifications of this basic method are developed to counter difficulties associated with complicated geometry and severe numerical resolution requirements of turbulence model equations. The accuracy of solutions is investigated by comparison with exact solutions for several simple cases. Procedures are developed for modifying the basic method to improve the accuracy. Numerical solutions of high-Reynolds-number separated flows over an airfoil and shock-separated flows over a flat plate are obtained. A simple mixing length model of turbulence is used for the transonic flow past an airfoil. A nonorthogonal mesh of arbitrary configuration facilitates the description of the flow field. For the simpler geometry associated with the flat plate, a rectangular mesh is used, and solutions are obtained based on a two-equation differential model of turbulence.

The Sheath Transport Observer for the Redistribution of Mass (STORM) Image

NASA Technical Reports Server (NTRS)

Kuntz, Kip; Collier, Michael; Sibeck, David G.; Porter, F. Scott; Carter, J. A.; Cravens, Thomas; Omidi, N.; Robertson, Ina; Sembay, S.; Snowden, Steven L.

2008-01-01

All of the solar wind energy that powers magnetospheric processes passes through the magnetosheath and magnetopause. Global images of the magnetosheath and magnetopause boundary layers will resolve longstanding controversy surrounding fundamental phenomena that occur at the magnetopause and provide information needed to improve operational space weather models. Recent developments showing that soft X-rays (0.15-1 keV) result from high charge state solar wind ions undergoing charge exchange recombination through collisions with exospheric neutral atoms has led to the realization that soft X-ray imaging can provide global maps of the high-density shocked solar wind within the magnetosheath and cusps, regions lying between the lower density solar wind and magnetosphere. We discuss an instrument concept called the Sheath Transport Observer for the Redistribution of Mass (STORM), an X-ray imager suitable for simultaneously imaging the dayside magnetosheath, the magnetopause boundary layers, and the cusps.

The Sheath Transport Observer for the Redistribution of Mass (STORM) Imager

NASA Technical Reports Server (NTRS)

Collier, Michael R.; Sibeck, David G.; Porter, F. Scott; Burch, J.; Carter, J. A.; Cravens, Thomas; Kuntz, Kip; Omidi, N.; Read, A.; Robertson, Ina;

Tactical missile aerodynamics

NASA Technical Reports Server (NTRS)

Hemsch, Michael J. (Editor); Nielsen, Jack N. (Editor)

1986-01-01

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

Microstructural effects on damage evolution in shocked copper polycrystals

DOE PAGES

Lieberman, Evan J.; Lebensohn, Ricardo A.; Menasche, David B.; ...

2016-07-01

Three-dimensional crystal orientation fields of a copper sample, characterized before and after shock loading using High Energy Diffraction Microscopy, are used for input and validation of direct numerical simulations using a Fast Fourier Transform (FFT)-based micromechanical model. The locations of the voids determined by X-ray tomography in the incipiently-spalled sample, predominantly found near grain boundaries, were traced back and registered to the pre-shocked microstructural image. Using FFT-based simulations with direct input from the initial microstructure, micromechanical fields at the shock peak stress were obtained. Statistical distributions of micromechanical fields restricted to grain boundaries that developed voids after the shock aremore » compared with corresponding distributions for all grain boundaries. Distributions of conventional measures of stress and strain (deviatoric and mean components) do not show correlation with the locations of voids in the post-shocked image. Neither does stress triaxiality, surface traction or grain boundary inclination angle, in a significant way. On the other hand, differences in Taylor factor and accumulated plastic work across grain boundaries do correlate with the occurrence of damage. As a result, damage was observed to take place preferentially at grain boundaries adjacent to grains having very different plastic response.« less

Shuttle Damage/Repair from the Perspective of Hypersonic Boundary Layer Transition - Experimental Results

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Merski, N. Ronald; Berger, Karen T.; Buck, Gregory M.; Liechty, Derek S.; Schneider, Steven P.

2006-01-01

An overview is provided of the experimental wind tunnel program conducted at the NASA Langley Research Center Aerothermodynamics Laboratory in support of an agency-wide effort to prepare the Shuttle Orbiter for Return-to-Flight. The effect of an isolated protuberance and an isolated rectangular cavity on hypersonic boundary layer transition onset on the windward surface of the Shuttle Orbiter has been experimentally characterized. These experimental studies were initiated to provide a protuberance and cavity effects database for developing hypersonic transition criteria to support on-orbit disposition of thermal protection system damage or repair. In addition, a synergistic experimental investigation was undertaken to assess the impact of an isolated mass-flow entrainment source (simulating pyrolysis/outgassing from a proposed tile repair material) on boundary layer transition. A brief review of the relevant literature regarding hypersonic boundary layer transition induced from cavities and localized mass addition from ablation is presented. Boundary layer transition results were obtained using 0.0075-scale Orbiter models with simulated tile damage (rectangular cavities) of varying length, width, and depth and simulated tile damage or repair (protuberances) of varying height. Cavity and mass addition effects were assessed at a fixed location (x/L = 0.3) along the model centerline in a region of near zero pressure gradient. Cavity length-to-depth ratio was systematically varied from 2.5 to 17.7 and length-to-width ratio of 1 to 8.5. Cavity depth-to-local boundary layer thickness ranged from 0.5 to 4.8. Protuberances were located at several sites along the centerline and port/starboard attachment lines along the chine and wing leading edge. Protuberance height-to-boundary layer thickness was varied from approximately 0.2 to 1.1. Global heat transfer images and heating distributions of the Orbiter windward surface using phosphor thermography were used to infer the state of the boundary layer (laminar, transitional, or turbulent). Test parametrics include angles-of-attack of 30 deg and 40 deg, sideslip angle of 0 deg, freestream Reynolds numbers from 0.02x106 to 7.3x106 per foot, edge-to-wall temperature ratio from 0.4 to 0.8, and normal shock density ratios of approximately 5.3, 6.0, and 12 in Mach 6 air, Mach 10 air, and Mach 6 CF4, respectively. Testing to simulate the effects of ablation from a proposed tile repair concept indicated that transition was not a concern. The experimental protuberance and cavity databases highlighted in this report were used to formulate boundary layer transition correlations that were an integral part of an analytical process to disposition observed Orbiter TPS damage during STS- 114.

Study of Pressure Oscillations in Supersonic Parachute

NASA Astrophysics Data System (ADS)

Dahal, Nimesh; Fukiba, Katsuyoshi; Mizuta, Kazuki; Maru, Yusuke

2018-04-01

Supersonic parachutes are a critical element of planetary mission whose simple structure, light-weight characteristics together with high ratio of aerodynamic drag makes them the most suitable aerodynamic decelerators. The use of parachute in supersonic flow produces complex shock/shock and wake/shock interaction giving rise to dynamic pressure oscillations. The study of supersonic parachute is difficult, because parachute has very flexible structure which makes obtaining experimental pressure data difficult. In this study, a supersonic wind tunnel test using two rigid bodies is done. The wind tunnel test was done at Mach number 3 by varying the distance between the front and rear objects, and the distance of a bundle point which divides suspension lines and a riser. The analysis of Schlieren movies revealed shock wave oscillation which was repetitive and had large pressure variation. The pressure variation differed in each case of change in distance between the front and rear objects, and the change in distance between riser and the rear object. The causes of pressure oscillation are: interaction of wake caused by front object with the shock wave, fundamental harmonic vibration of suspension lines, interference between shock waves, and the boundary layer of suspension lines.

Heat Transfer Due to Shock Wave Turbulent Boundary Layer Interactions on High Speed Weapon Systems

DTIC Science & Technology

1974-04-01

Thermocouple Location. ............ 20 8 Pressure Tap LoceL.lons ............ 23 9 Pitot Pressurp Rake Detail and Profile Measurement Stations... Pitot Pressure Profile Measurements - Pitot pressure measure- ments were made with a NASA Langley five-probe pitot pressure rake at the axial location...defined in Figure 6. Figure 9 is a schematic illustration providing the dimensions of the rake and the pitot pressure profile measure- ment stations. The

Modeling of structural uncertainties in Reynolds-averaged Navier-Stokes closures

NASA Astrophysics Data System (ADS)

Emory, Michael; Larsson, Johan; Iaccarino, Gianluca

2013-11-01

Estimation of the uncertainty in numerical predictions by Reynolds-averaged Navier-Stokes closures is a vital step in building confidence in such predictions. An approach to model-form uncertainty quantification that does not assume the eddy-viscosity hypothesis to be exact is proposed. The methodology for estimation of uncertainty is demonstrated for plane channel flow, for a duct with secondary flows, and for the shock/boundary-layer interaction over a transonic bump.

PROJECT SQUID. Field Survey Report. Volume 1, Part 4. Fluid Mechanics

DTIC Science & Technology

1947-06-30

title of TechnolOgyl) .. .1 Figure 4. Sketch showing position of compresion shocks in conical ditfitsers. (Aflcr Osiruaitscw) . 12 Figure 5. Maximnum...degree. However, experimen- sary to secure separation of the boundary layer. A tal investigations are being conducted on this problem blister was added at...MECHANIC v Since temperature is a transferable properky in the general nature of the mathematical background and to same sense as momentumt and vorticity

Gas flows in radial micro-nozzles with pseudo-shocks

NASA Astrophysics Data System (ADS)

Kiselev, S. P.; Kiselev, V. P.; Zaikovskii, V. N.

2018-07-01

In the present paper, results of an experimental and numerical study of supersonic gas flows in radial micro-nozzles are reported. A distinguishing feature of such flows is the fact that two factors, the nozzle divergence and the wall friction force, exert a substantial influence on the flow structure. Under the action of the wall friction force, in the micro-nozzle there forms a pseudo-shock that separates the supersonic from subsonic flow region. The position of the pseudo-shock can be evaluated from the condition of flow blockage in the nozzle exit section. A detailed qualitative and quantitative analysis of gas flows in radial micro-nozzles is given. It is shown that the gas flow in a micro-nozzle is defined by the complicated structure of the boundary layer in the micro-nozzle, this structure being dependent on the width-to-radius ratio of the nozzle and its inlet-to-outlet pressure ratio.

Source Distribution Method for Unsteady One-Dimensional Flows With Small Mass, Momentum, and Heat Addition and Small Area Variation

NASA Technical Reports Server (NTRS)

Mirels, Harold

1959-01-01

A source distribution method is presented for obtaining flow perturbations due to small unsteady area variations, mass, momentum, and heat additions in a basic uniform (or piecewise uniform) one-dimensional flow. First, the perturbations due to an elemental area variation, mass, momentum, and heat addition are found. The general solution is then represented by a spatial and temporal distribution of these elemental (source) solutions. Emphasis is placed on discussing the physical nature of the flow phenomena. The method is illustrated by several examples. These include the determination of perturbations in basic flows consisting of (1) a shock propagating through a nonuniform tube, (2) a constant-velocity piston driving a shock, (3) ideal shock-tube flows, and (4) deflagrations initiated at a closed end. The method is particularly applicable for finding the perturbations due to relatively thin wall boundary layers.

Science Objectives for a Soft X-ray Mission

NASA Astrophysics Data System (ADS)

Sibeck, D. G.; Connor, H. K.; Collier, M. R.; Collado-Vega, Y. M.; Walsh, B.

2016-12-01

When high charge state solar wind ions exchange electrons with exospheric neutrals, soft X-rays are emitted. In conjunction with flight- proven wide field-of-view soft X-ray imagers employing lobster-eye optics, recent simulations demonstrate the feasibility of imaging magnetospheric density structures such as the bow shock, magnetopause, and cusps. This presentation examines the Heliospheric scientific objectives that such imagers can address. Principal amongst these is the nature of reconnection at the dayside magnetopause: steady or transient, widespread or localized, component or antiparallel as a function of solar wind conditions. However, amongst many other objectives, soft X-ray imagers can provide crucial information concerning the structure of the bow shock as a function of solar wind Mach number and IMF orientation, the presence or absence of a depletion layer, the occurrence of Kelvin-Helmholtz or pressure-pulse driven magnetopause boundary waves, and the effects of radial IMF orientations and the foreshock upon bow shock and magnetopause location.

Research study of space plasma boundary processes

NASA Technical Reports Server (NTRS)

Greenstadt, E. W.; Taylor, W. W. L.

1984-01-01

Representation of the Earth's bow shock and magnetopause and their geometrically determined macrostructure was investigated. Computer graphic depictions of the global distributions of bow shock structures and elementary animation of the dynamics of those distributions in the changing solar wind were developed. The shock-foreshock boundary and subcritical bow shocks as observed by ISEE 1 and 2 are discussed.

Predictive Analytical Model for Isolator Shock-Train Location in a Mach 2.2 Direct-Connect Supersonic Combustion Tunnel

NASA Astrophysics Data System (ADS)

Lingren, Joe; Vanstone, Leon; Hashemi, Kelley; Gogineni, Sivaram; Donbar, Jeffrey; Akella, Maruthi; Clemens, Noel

2016-11-01

This study develops an analytical model for predicting the leading shock of a shock-train in the constant area isolator section in a Mach 2.2 direct-connect scramjet simulation tunnel. The effective geometry of the isolator is assumed to be a weakly converging duct owing to boundary-layer growth. For some given pressure rise across the isolator, quasi-1D equations relating to isentropic or normal shock flows can be used to predict the normal shock location in the isolator. The surface pressure distribution through the isolator was measured during experiments and both the actual and predicted locations can be calculated. Three methods of finding the shock-train location are examined, one based on the measured pressure rise, one using a non-physics-based control model, and one using the physics-based analytical model. It is shown that the analytical model performs better than the non-physics-based model in all cases. The analytic model is less accurate than the pressure threshold method but requires significantly less information to compute. In contrast to other methods for predicting shock-train location, this method is relatively accurate and requires as little as a single pressure measurement. This makes this method potentially useful for unstart control applications.

Diagnostics of boundary layer transition by shear stress sensitive liquid crystals

NASA Astrophysics Data System (ADS)

Shapoval, E. S.

2016-10-01

Previous research indicates that the problem of boundary layer transition visualization on metal models in wind tunnels (WT) which is a fundamental question in experimental aerodynamics is not solved yet. In TsAGI together with Khristianovich Institute of Theoretical and Applied Mechanics (ITAM) a method of shear stress sensitive liquid crystals (LC) which allows flow visualization was proposed. This method allows testing several flow conditions in one wind tunnel run and does not need covering the investigated model with any special heat-insulating coating which spoils the model geometry. This coating is easily applied on the model surface by spray or even by brush. Its' thickness is about 40 micrometers and it does not spoil the surface quality. At first the coating obtains some definite color. Under shear stress the LC coating changes color and this change is proportional to shear stress. The whole process can be visually observed and during the tests it is recorded by camera. The findings of the research showed that it is possible to visualize boundary layer transition, flow separation, shock waves and the flow image on the whole. It is possible to predict that the proposed method of shear stress sensitive liquid crystals is a promise for future research.

Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Lamar, John E.; Obara, Clifford J.; Fisher, Bruce D.; Fisher, David F.

2001-01-01

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.

Effects of variables upon pyrotechnically induced shock response spectra, part 2

NASA Technical Reports Server (NTRS)

Smith, James Lee

1988-01-01

Throughout the aerospace industry, large variations of 50 percent (6 dB) or more in shock response spectra (SRS) derived from pyrotechnic separation events continue to be reported from actual spaceflight data and from laboratory tests. As a result of these variations, NASA funded a research program for 1984 through 1986. The purpose of the 1984 through 1986 project was to analyze variations in pyrotechnically induced SRS and to determine if and to what degree manufacturing and assembly variables and tolerances, distance from the shock source, data acquisition instrumentation, and shock energy propagation affect the SRS. Sixty-four free-free boundary plate tests were performed. NASA funded an additional study for 1987 through 1988. This paper is a summary of the additional study. The purpose was to evaluate shock dissipation through various spacecraft structural joint types, to evaluate shock variation for various manufacturing and assembly variables on clamped boundary test plates, and to verify data correction techniques. Five clamped boundary plate tests investigated manufacturing and assembly variables and mass loading effects. Six free-free boundary plate tests investigated shock dissipation across spacecraft joint structures.

Shocked quartz and more: Impact signatures in K-T boundary clays and claystones

NASA Technical Reports Server (NTRS)

Bohor, Bruce F.

1988-01-01

Quartz grains displaying multiple sets of planar features are described from numerous Cretaceous-Tertiary (K-T) boundary clays and claystones at both marine and nonmarine depositional sites around the world. All these sites also show anomalously high amounts of iridium and enrichments of other siderophile elements in cosmic ratios within these boundary units. This combination of mineralogical and geochemical features are used in support of an impact hypothesis for the end-Cretaceous event. Recently, it was suggested that some combination of explosive and nonexplosive volcanism associated with the formation of the Deccan traps in India could be responsible for the mineralogy and geochemistry seen in the K-T boundary units. Besides the obvious contradition of simultaneous explosive and nonexplosive volcanism from one locality during an instant of geologic time, there remains the difficulty of spreading both iridium (and trace elements in cosmic proportions) and quartz grains around the world by volcanic (atmospheric) transport. In addition, the ability of volcanism to produce the type of shock metamorphism seen in minerals at the K-T boundary was not demonstrated. Multiple sets of shock lamellae in quartz are considered characteristic of shock metamorphism in rocks at the sites of known impact craters and are the type of deformation seen in quartz from K-T boundary clays and claystones. Single sets of poorly defined lamellae described from rare quartz grains in certain volcanic deposits are characteristic of tectonic deformation and do not correspond to the shock lamellae in quartz from K-T sediments and impact structures. So-called shock mosaicism in quartz and feldspar grains described from volcanic deposits can result from many processes other than shock metamorphism, and therefore is not considered to be an effect characteristic solely of shock. The mineralogy of shock-metamorphosed grains at the K-T boundary also argues against a volcanic origin.

The NASA Ames 16-Inch Shock Tunnel Nozzle Simulations and Experimental Comparison

NASA Technical Reports Server (NTRS)

TokarcikPolsky, S.; Papadopoulos, P.; Venkatapathy, E.; Delwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

The 16-Inch Shock Tunnel at NASA Ames Research Center is a unique test facility used for hypersonic propulsion testing. To provide information necessary to understand the hypersonic testing of the combustor model, computational simulations of the facility nozzle were performed and results are compared with available experimental data, namely static pressure along the nozzle walls and pitot pressure at the exit of the nozzle section. Both quasi-one-dimensional and axisymmetric approaches were used to study the numerous modeling issues involved. The facility nozzle flow was examined for three hypersonic test conditions, and the computational results are presented in detail. The effects of variations in reservoir conditions, boundary layer growth, and parameters of numerical modeling are explored.