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Showing papers on "Freestream published in 1996"


Journal ArticleDOI
TL;DR: In this paper, boundary-layer measurements were conducted over a flared cone in a quiet wind tunnel and the results suggest that the second mode is the dominant mode of instability, compared well with linear stability theory in the linear stability regime.
Abstract: Hypersonic boundary-layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82 x 10 6 /ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary-layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode is the dominant mode of instability. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. Furthermore, the existence of higher harmonics of the fundamental suggests that nonlinear disturbances are not associated with high freestream disturbance levels.

116 citations


Journal ArticleDOI
TL;DR: In this paper, the boundary layer statistics for the interaction between a turbulent boundary layer and a freestream with turbulence levels ranging from 10 to 20 percent were reported, showing that the mean velocity profile still exhibits a log-linear region.
Abstract: High freestream turbulence levels significantly alter the characteristics of turbulent boundary layers. Numerous studies have been conducted with freestreams having turbulence levels of 7 percent or less, but studies using turbulence levels greater than 10 percent have been essentially limited to the effects on wall shear stress and heat transfer. This paper presents measurements of the boundary layer statistics for the interaction between a turbulent boundary layer and a freestream with turbulence levels ranging from 10 to 20 percent. The boundary layer statistics reported in this paper include mean and rms velocities, velocity correlation coefficients, length scales, and power spectra. Although the freestream turbulent eddies penetrate into the boundary layer at high freestream turbulence levels, as shown through spectra and length scale measurements, the mean velocity profile still exhibits a log-linear region. Direct measurements of total shear stress (turbulent shear stress and viscous shear stress) confirm the validity of the log-law at high freestream turbulence levels. Velocity defects in the outer region of the boundary layer were significantly decreased resulting in negative wake parameters. Fluctuating rms velocities were only affected when the freestream turbulence levels exceeded the levels of the boundary layer generated rms velocities. Length scales and power spectra measurements showed large scale turbulent eddies penetrate to within y+ = 15 of the wall.

99 citations


01 Dec 1996
TL;DR: In this paper, the influence of freestream turbulence intensity and film cooling hole length-to-diameter ratio on mean velocity and turbulence intensity was studied in simulated film cooling.
Abstract: Hot-wire anemometry of simulated film cooling was used to study the influence of freestream turbulence intensity and film cooling hole length-to-diameter ratio on mean velocity and turbulence intensity. Measurements were made in the zone where the coolant and freestream flows mix. Flow from one row of film cooling holes with a streamwise injection of 35{degree} and no lateral injection and with a coolant- to-freestream flow velocity ratio of 1.0 was investigated under freestream turbulence levels of 0.5 and 12%. Coolant-to-freestream density ratio was unity. Two length-to-diameter ratios for the film cooling holes, 2.3 and 7.0, are tested. Results show that under low freestream turbulence conditions, pronounced differences exist in the flowfield between L/D=7.0 and 2.3; the differences are less prominent at high freestream turbulence intensities. Generally, short-L/D injection results in ``jetting`` of the coolant further into the freestream flow and enhanced mixing. Other changes in the flowfield attributable to a rise in freestream turbulence intensity to engine- representative conditions are documented. 15 figs, 2 tabs, refs.

93 citations


Journal ArticleDOI
TL;DR: In this article, the bow and separation shocks formed upstream of the injectant plume are examined in flowfields created by transverse injection into supersonic cross-flows. And the interaction between these features and the large-scale eddies that develop at the jet/freestream interface has been examined.
Abstract: In flowfields created by transverse injection into supersonic cross-flows, the bow and separation shocks formed upstream of the injectant plume are dominant features. In the present investigation, the interaction between these features and the large-scale eddies that develop at the jet/freestream interface has been examined.

76 citations


Journal ArticleDOI
TL;DR: In this paper, a planar laser-induced iodine fluorescence is used to map out the nonreacting mixing flowfield of an unswept ramp fuel injector using air injected at Mach 2.0 into a Mach2.9 freestream.
Abstract: Planar laser-induced iodine fluorescence is used to map out the nonreacting mixing flowfield of an unswept ramp fuel injector using air injected at Mach 2.0 into a Mach 2.9 freestream. A fully automated test setup is used to measure time-averaged pressure, temperature, velocity, and injectant mole fraction on 21 crossflow planes and 7 axial planes. The measurement uncertainties are 5-8% for temperature, 4-10% for pressure, 10-20 m/s for velocity, and 2-3% for injectant mole fraction depending on the thermodynamic conditions. The measurements allow any desired gasdynamic quantity to be determined on a three-dimensional grid that spans the entire wind-tunnel test section. The experimental data set is comparable to the completeness of results normally available only from a computational fluid dynamics simulation. Results showing detailed flow features on specific planes, as well as overall quantities, such as global conservation checks, mixing performance, and flowfield losses, are presented. Mass, momentum, and energy flux, determined at the crossflow plane locations of the data set, show about a 2% standard deviation. The results are compared to a simulation using a three-dimensional Navier-Stokes solver. Agreement is reasonable with the exception of measurements in regions very close to walls, where the intensity of scattered light is high or where optical access is limited. The ability to generate extensive data sets, such as the one presented here, demonstrates that the planar laser-induced iodine fluorescence technique can be used 1) to generate detailed test cases for the validation of computational fluid dynamics codes and 2) as an alternative to computational fluid dynamics for performing design studies and performance evaluation in complex compressible flows.

73 citations


Journal ArticleDOI
TL;DR: In this article, an experiment was conducted to determine whether the addition of swirl will improve the mixing of a supersonic jet of fuel simulant (helium or air) injected at 30 deg to the wall into a confined Mach 2 airflow.
Abstract: Hydrogen fuel injected into a scramjet combustor must mix rapidly if complete combustion is to occur within a reasonable stream wise distance. An experiment has been conducted to determine whether the addition of swirl will improve the mixing of a supersonic jet of fuel simulant (helium or air) injected at 30 deg to the wall into a confined Mach 2 airflow. The swirling jets were created by injecting the fuel simulant tangentially into a cylindrical chamber and accelerating it through a convergent-div er gent nozzle. The flow was visualized by imaging Rayleigh scattering from a laser light sheet, and the plume penetration and cross-sectional area were obtained. The plumes from the swirling and nonswirling jets had comparable penetration and area, but the swirling jets contained substantially less mass flow, suggesting better mixing efficiency. Interaction of streamwise vorticity within the plumes of the swirling jets with their images in the duct wall caused the plumes to be inclined laterally to the freestream.

58 citations


Journal ArticleDOI
TL;DR: In this paper, a new technique for measuring skin friction was employed to help document the flow on an airfoil at angles of attack from -0.5 to 11.5 deg.
Abstract: A new technique for measuring skin friction was employed to help document the flow on an airfoil at angles of attack from -0.5 to 11.5 deg. Surface pressures were also measured on both the wing and wind-tunnel walls. The experiment was conducted at a freestream Mach number of 0.2 and Reynolds numbers of 0.6, 2, and 6 x 10 6 . The objective of the study was to provide data and boundary condition information sufficient for the validation of numerical simulations. Such a simulation of the experiment was conducted using the INS2D Navier-Stokes code with the shear-stress-transport turbulence model. The computations provide a good description of both laminar and turbulent shear levels, except for turbulent flow on the top surface of the wing at the higher angles of attack.

57 citations


Proceedings ArticleDOI
01 Jan 1996
TL;DR: In this paper, boundary layer measurements were conducted over a flared cone in a quiet wind tunnel and the results indicated that the second mode disturbances were the most unstable and scaled with the boundary layer thickness.
Abstract: Hypersonic boundary layer measurements were conducted over a flared cone in a quiet wind tunnel. The flared cone was tested at a freestream unit Reynolds number of 2.82x106/ft in a Mach 6 flow. This Reynolds number provided laminar-to-transitional flow over the model in a low-disturbance environment. Point measurements with a single hot wire using a novel constant voltage anemometry system were used to measure the boundary layer disturbances. Surface temperature and schlieren measurements were also conducted to characterize the laminar-to-transitional state of the boundary layer and to identify instability modes. Results suggest that the second mode disturbances were the most unstable and scaled with the boundary layer thickness. The integrated growth rates of the second mode compared well with linear stability theory in the linear stability regime. The second mode is responsible for transition onset despite the existence of a second mode sub-harmonic. The sub-harmonic wavelength also scales with the boundary layer thickness. Furthermore, the existence of higher harmonics of the fundamental suggests that non-linear disturbances are not associated with high free stream disturbance levels.

54 citations


Proceedings ArticleDOI
15 Jan 1996
TL;DR: In this paper, Rayleigh/Mie scattering images were used to examine the convection characteristics of large-scale vortex structures developing in flowfields created by sonic transverse injection through circular and elliptical nozzles into a Mach 1.98 crossflow.
Abstract: Temporally correlated Rayleigh/Mie scattering images were used to examine the convection characteristics of the large-scale vortex structures developing in flowfields created by sonic transverse injection through circular and elliptical nozzles into a Mach 1.98 crossflow. Both compressibility and injector geometry were found to have significant influences on the convection characteristics of the large eddies. High compressibility injection cases produce eddies having dramatically larger near-field convection velocities than those in low compressibility cases. Farther downstream, the large-scale vortices tend to travel at velocities nearer the freestream veocity. Injector geometry primarily affects the near-field behavior.

52 citations


Proceedings ArticleDOI
15 Jan 1996
TL;DR: In this paper, a two-equation k-omega turbulence model has been developed and applied to a quasi-three-dimensional viscous analysis code for blade-to-blade flows in turbomachinery.
Abstract: A two-equation k-omega turbulence model has been developed and applied to a quasi-three-dimensional viscous analysis code for blade-to-blade flows in turbomachinery. the code includes the effects of rotation, radius change, and variable stream sheet thickness. The flow equations are given and the explicit runge-Kutta solution scheme is described. the k-omega model equations are also given and the upwind implicit approximate-factorization solution scheme is described. Three cases were calculated: transitional flow over a flat plate, a transonic compressor rotor, and transonic turbine vane with heat transfer. Results were compared to theory, experimental data, and to results using the Baldwin-Lomax turbulence model. The two models compared reasonably well with the data and surprisingly well with each other. Although the k-omega model behaves well numerically and simulates effects of transition, freestream turbulence, and wall roughness, it was not decisively better than the Baldwin-Lomax model for the cases considered here.

49 citations


Journal ArticleDOI
TL;DR: In this article, the mean and turbulent flow field associated with low-angled supersonic gaseous injection into a freestream was performed. And the results indicated that the turbulent flow structure of the injection plume were strongly influenced by the presence of a counter-rotating vortex pair (up'v'lpu'v', up'w'/pu'w' were in the range of 2.0-75.0% of the total shear stress level.
Abstract: An experimental study of the mean and turbulent flowfield associated with low-angled supersonic gaseous injection into a supersonic freestream was performed. Air was injected at Mach 1.8, with an effective back pressure ratio of 3.0, through an orifice at an angle of 25 deg into a Mach 2.9 air freestream (Re/m - 15 x 10 6). Cross-film anemometry and conventional mean flow probe surveys were acquired across the plume at two downstream stations (xld = 20 and 40). Schlieren photography was used for qualitative flow visualization. Turbulence measurements included contours of the turbulent kinetic energy and the full compressible Reynolds shear stresses in both the x-y and x-z planes. Mean flow data included Mach number, three-dimensional velocity components, and vorticity. The measurements indicated that the mean and turbulent flow structure of the injection plume were strongly influenced by the presence of a counter-rotating vortex pair (\ux |max « 15,000 /s). The turbulent kinetic energy was found to have two peaks colocated with the vortices. The turbulent shear stress distributions across the plume were found to be highly three dimensional and complicated by both the additional strain rates associated with the vorticity and turbulent convection. The present results also implied that the compressibility terms in the Reynolds shear stress accounted for about 67.0-75.0% of the total shear stress level, i.e., up'v'lpu'v' and up'w'/pu'w' were in the range of 2.0-3.0.

Journal ArticleDOI
TL;DR: In this paper, a tunable infrared diode laser was used in the F4 high-enthalpy hypersonic wind tunnel at ONERA to make line-of-sight measurements on nitric oxide (NO).
Abstract: A tunable infrared diode laser was used in the F4 high-enthalpy hypersonic wind tunnel at ONERA to make line-of-sight measurements on nitric oxide (NO). The NO is generated in the arc chamber and remains frozen in the hypersonic flow. We report the results of a medium enthalpy run. The Ω (14.5), Ω = 3/2 doublet was recorded simultaneously in two views : one normal to the flow and the other biased at 63-deg to the flow axis. Spectra were recorded at 2 ms intervals during a tunnel run, and 30 sets of spectra were analyzed by a least-squares fitting procedure. The mean of the velocity measurements inferred from the Doppler shift of the doublet was 2740 m/s (±5%). The velocity decayed about 10% over 60 ms. The temperature and NO column density determined by the least-squares fits were subject to much larger uncertainties. The static temperature of the freestream is estimated to be 168 K (±35%). The median inferred NO column density of 4 X 10 17 molecule/cm 2 corresponds to an NO number density of order 5 x 10 15 molecules/cm 3 in the freestream.

Journal ArticleDOI
TL;DR: In this paper, the pressure oscillations over the nose of a blunt body with a forward-facing cavity were studied both experimentally and numerically. But the results were limited to the case of the Mach 5 blowdown tunnel and the commercial finite volume code.
Abstract: Hypersonic flow over the nose of a blunt body with a forward-facing cavity is studied both experimentally (Mach 5 blowdown tunnel) and numerically (commercial finite volume code). Trends are established for the pressure oscillations within the cavity, flowfield structure, and surface heating for different cavity depths and lip radii. Resonant pressure oscillations within the cavity are an experimental flow feature but occur numerically, for the cavity geometries studied, only if freestream fluctuations are present. The oscillations are dominated by the quarter-wave frequency of the cavity. In the numerical simulations, the oscillation strength increases with cavity depth. In the experiments, oscillation strength generally increases with cavity depth; however, for a specific midrange of cavity depths, experiments show that the pressure oscillations switch randomly between two modes of behavior involving smalland large-amplitude fluctuations. Agreement between experiment and computations is good for the flowfield structure and surface heating of shallow cavity flows. Sharp lips produce both a recirculation region that cools the outer surface, and severe heating just inside the cavity. Rounding the lip eliminates the recirculation region and alleviates heating inside the cavity. Experimental results show that the strong oscillations associated with deeper cavities may produce a cooling effect.

19 Aug 1996
TL;DR: In this paper, tail buffet tests were performed on a full-scale, production model F/A-18 in the 80- by 120-foot wind tunnel at NASA Ames Research Center.
Abstract: In 1993, tail buffet tests were performed on a full-scale, production model F/A-18 in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center. Steady and unsteady pressures were recorded on both sides of the starboard vertical tail for an angle-of-attack range of 20 to 40 degrees and at a sideslip range of -16 to 16 degrees at freestream velocities up to 100 knots (Mach 0.15, Reynolds number 1.23*10super7). The aircraft was equipped with removable leading edge extension (LEX) fences that are used in Flight to reduce tail buffet loads. In 1995, tail buffet tests were performed on a 1/6-scale F-18 A/B model in the Transonic Dynamics Tunnel (TDT) at NASA Langley Research Center. Steady and unsteady pressures were recorded on both sides of both vertical tails for an angle-of-attack range of 7 to 37 degrees at freestream velocities up to 65 knots (Mach 0.10). Comparisons of steady and unsteady pressures and root bending moments are presented for these wind-tunnel models for selected test cases. Representative pressure and root bending moment power spectra are also discussed, as are selected pressure cross-spectral densities.

01 Dec 1996
TL;DR: An aerothermodynamic database has been generated through both experimental testing and computational fluid dynamics simulations for a 70 deg sphere-cone configuration based on the NASA Mars-Pathfinder entry vehicle as mentioned in this paper.
Abstract: An aerothermodynamic database has been generated through both experimental testing and computational fluid dynamics simulations for a 70 deg sphere-cone configuration based on the NASA Mars-Pathfinder entry vehicle The aerothermodynamics of several related parametric configurations were also investigated Experimental heat-transfer data were obtained at hypersonic test conditions in both a perfect-gas air wind tunnel and in a hypervelocity, high-enthalpy expansion tube in which both air and carbon dioxide were employed at test gases In these facilities, measurements were made with thin-film temperature-resistance gages on both the entry vehicle models and on the support stings of the models Computational results for freestream conditions equivalent to those of the test facilities were generated using an axisymmetric/2D laminar Navier-Stokes solver with both perfect-gas and nonequilibrium thermochemical models Forebody computational and experimental heating distributions agreed to within the experimental uncertainty for both the perfect-gas and high-enthalpy test conditions In the wake, quantitative differences between experimental and computational heating distributions for the perfect-gas conditions indicated transition of the free shear layer near the reattachment point on the sting For the high-enthalpy cases, agreement to within, or slightly greater than, the experimental uncertainty was achieved in the wake except within the recirculation region, where further grid resolution appeared to be required Comparisons between the perfect-gas and high-enthalpy results indicated that the wake remained laminar at the high-enthalpy test conditions, for which the Reynolds number was significantly lower than that of the perfect-gas conditions

Journal ArticleDOI
TL;DR: In this paper, a 50 by 25 mm combustor model using a flush-mounted, Mach 1.7 single injector introducing the fuel at an angle of 15 deg relative to the main flow was performed in the T5 free-piston shock tunnel at GALCIT.
Abstract: Transverse jet mixing and combustion experiments at a freestream velocity of 5 km/s have been conducted in the T5 free-piston shock tunnel at GALCIT. The experiments were performed in a 50 by 25 mm combustor model using a flush-mounted, Mach 1.7 single injector introducing the fuel at an angle of 15 deg relative to the main flow. The test conditions cover a range of freestream pressures from low values chosen to match previous experiments with the same duct in the HYPULSE expansion tube facility at GASL, to high pressures corresponding to the conditions closer to those in a real scramjet combustor. Tests with cold and hot hydrogen fuel were performed, the latter to better reproduce the fuel conditions of a scramjet propulsion system where the hydrogen will have to be used first as a cooling fluid. A small combustion-driven shock tunnel was used simultaneously with T5 to supply the hot fuel.

Journal ArticleDOI
TL;DR: In this article, the arrival of the driver gas in a shock tunnel is detected by choking a duct when the specific heat ratio is increased past a critical value, and the results are compared with and are in agreement with measurements made with a mass spectrometer.
Abstract: A device has been developed to detect the arrival of the driver gas in a shock tunnel. The detector is small enough to be used in conjunction with other experiments. It works by choking a duct when the specific heat ratio is increased past a critical value. Times are given for the onset of a 7.5% contamination level in flows with freestream enthalpies ranging from 3–9 MJ/kg. These results are compared with and are shown to be in agreement with measurements made with a mass spectrometer. Results displaying the rate at which the test gas is contaminated are also given.

Journal ArticleDOI
TL;DR: In this article, the power spectra were measured for the frequency range 13 Hz < f < 13 kHz and a clear overlap region between the mid and high frequency parts of the spectrum is shown.
Abstract: Silicon based pressure sensors have been used to measure turbulent wall-pressure fluctuations in a two-dimensional flat plate boundary layer, Re θ = 5072. The side lengths of the diaphragms were 100 μm (d + = 7.2 ) and 300 μm (d + = 21.6), giving a ratio of the boundary layer thickness to the diaphragm side length of the order of 240 and a resolution of eddies with wave numbers less than ten viscous units. Power spectra were measured for the frequency range 13 Hz < f < 13 kHz. Scaled in outer and inner variables a clear overlap region between the mid and high frequency parts of the spectrum is shown. In this overlap region the slope was found to be ω -1 , while in the high frequency part it was ω -5 . Correlation measurements in both the longitudinal and transversal directions were performed and compared to other investigations. Longitudinal space time correlations, including the high frequency range, indicated an advection velocity of the order of half the freestream velocity. A broad band filtering of the longitudinal correlation showed that the high frequency part of the spectrum is associated with the smaller eddies from the inner part of the boundary layer, resulting in a reduction of the correlation.

Journal ArticleDOI
TL;DR: In this article, boundary-layer transition data on swept cylinders in hypersonic flow were obtained in the Hypersonic Ludwieg-tube wind tunnel of DLR and the response of attachment-line transition to surface roughness heights due to trip wires for sweep angles of 45 and 60 deg at M∞ = 5.0 was investigated.
Abstract: We present boundary-layer transition data on swept cylinders in hypersonic flow that were obtained in the hypersonic Ludwieg-tube wind tunnel of DLR. Experiments were conducted at M∞ = 5.0 and 6.9 for three swept cylinders with and without end plates and sweep angles of 30, 45, and 60 deg, respectively. To determine the state of the boundary layers, the liquid-crystal technique was applied. In addition, the response of attachment-line transition to surface roughness heights due to trip wires for sweep angles of 45 and 60 deg at M∞ = 5.0 was investigated. Freestream Reynolds numbers based on cylinder diameter for attachment-line transition were about 0.2-0.3 x 10 6 for the end plate disturbances and 0.9-1.2 x 10 6 for no contaminations. Trends of attachment-line transition Reynolds numbers with surface roughness heights are roughly similar to subsonic ones with respect to critical roughness heights. The behavior of the attachment-line transition around the critical roughness heights, however, depended strongly on spanwise Mach numbers. With flow visualizations, off-attachment-line transition could be observed, as well as fine streak pattern in laminar flow region upstream of the transition front, confirming streamwise vortices induced by crossflow instability at hypersonic speeds.

Journal ArticleDOI
TL;DR: In this article, the effect of radiation on mixed convection from a horizontal flat plate in a saturated porous medium is investigated and the conservation equations that govern the problem are reduced to a system of nonlinear ordinary different equations.
Abstract: An analysis is presented to investigate the effect of radiation on mixed convection from a horizontal flat plate in a saturated porous medium. Both a hot surface facing upward and a cold surface facing downward are considered in the analysis. The conservation equations that govern the problem are reduced to a system of nonlinear ordinary different equations. The important parameters of this problem are the radiation parameter R, the buoyancy parameter B, and the freestream to wall temperature ratio T ∞/T w for the case of a hot surface or the wall to freestream to wall temperature T w /T ∞ for the case of a cold surface.

01 Jan 1996
TL;DR: In this article, the authors investigated the transverse injection of helium and air through circular and elliptical nozzles into a supersonic cross flow using optical diagnostics and probe-based measurement techniques.
Abstract: : Transverse injection of helium and air through circular and elliptical nozzles into a supersonic cross flow was investigated using optical diagnostics and probe-based measurement techniques. Shadowgraph visualizations documented the global characteristics of the jet/freestream interaction. Rayleigh/Mie scattering allowed ensembles of digital images from several measurement planes in each case to be collected. Pitot and concentration probes were used in the helium injection cases. The images provide mean and standard deviation statistics, spreading and penetration characteristics, large-scale mixing and convective velocity information, bow and separation shock features, and two-dimensional spatial correlation fields. Results indicate that the elliptical nozzle produces jets with greater lateral spread and suppressed transverse penetration compared to the jets from the circular nozzle. Injectant molecular weight does not strongly affect the jet's penetration, although it leads to substantial compressibility differences that dramatically influence the characteristics of the large-scale shear layer structures and the entrainment and mixing occurring between the injectant and crossflow. Probe measurements provide quantitative comparisons of total pressure losses and mixing characteristics. Results suggest better near field mixing and lower total pressure losses in the elliptical injection flowfield.

Proceedings ArticleDOI
10 Jun 1996
TL;DR: In this article, the regions of laminar and turbulent flow have been investigated in a linear cascade of a high tuming HP rotor blades, and the results are improved with the Kato-Launder modification.
Abstract: The regions of laminar and turbulent flow have been investigated in a linear cascade of a high tuming HP rotor blades. Measurements of intermittency close to the blade and end wall surfaces have shown substantial areas of laminar and transitional flow. The implications for turbulence modelling are important, and Navier-Stokes computations have been performed to investigate how well transition can be modelled in such a flow. Using the intermittency data to specify transitional areas, the mixing length model of turbulence produces excellent results, although there is some sensitivity to the assumed freestream length scale. High Reynolds k-e model results show too much turbulence and loss using the measured high inlet length scale, but the results are improved with the Kato-Launder modification. A low Reynolds number model does not seem to predict the transition effects, although more work is required with this model.Copyright © 1996 by ASME

Journal ArticleDOI
TL;DR: In this article, the authors report steady, two-dimensional, inviscid solutions for the near field and far field of a supersonic reactive flow over a variable-double-ramp geometry.
Abstract: In this article we report steady, two-dimensional, inviscid solutions for the near field and far field of a supersonic reactive flow over a variable-double-ramp geometry/The incident shock wave compresses and heats the reactants that will combust after flowing some induction length* Upon reaction, a detonation wave forms and intersects the leading wave at some distance from the ramp surface. In this article, reaction-polar diagrams are developed and the detonation branch solutions are used to investigate the wave interaction processes that may lead to a steady three-wave structure. By considering this formation process new oblique detonation wave stabilization criteria based on the freestream conditions and ramp geometry are provided.

Journal ArticleDOI
TL;DR: In this paper, the authors describe the development and structure of flow downstream of a single row of holes with compound angle orientations producing film cooling at high blowing ratios, which is important because similar arrangements are frequently employed on the first stage of rotating blades of operating gas turbine engines.
Abstract: Experimental results are presented which describe the development and structure of flow downstream of a single row of holes with compound angle orientations producing film cooling at high blowing ratios. This film cooling configuration is important because similar arrangements are frequently employed on the first stage of rotating blades of operating gas turbine engines. With this configuration, holes are spaced 6d apart in the spanwise direction, with inclination angles of 24 degrees, and angles of orientation of 50.5 degrees. Blowing ratios range from 1.5 to 4.0 and the ratio of injectant to freestream density is near 1.0. Results show that spanwise averaged adiabatic effectiveness, spanwise-averaged iso-energetic Stanton number ratios, surveys of streamwise mean velocity, and surveys of injectant distributions change by important amounts as the blowing ratio increases. This is due to injectant lift-off from the test surface just downstream of the holes.

Journal ArticleDOI
TL;DR: In this article, a joint experimental and computational study has been performed to investigate the flowfield structure created by two crossing oblique shock waves interacting with a turbulent boundary layer, which is of practical importance in the design of high-speed sidewall-compression inlets.
Abstract: A joint experimental and computational study has been performed to investigate the flowfield structure created by two crossing oblique shock waves interacting with a turbulent boundary layer. Such an interaction is of practical importance in the design of high-speed sidewall-compression inlets. The interaction is created by a test model, consisting of two sharp fins mounted at 15-deg angle of attack to a flat plate, placed in a Mach 3.85 freestream flow with a unit Reynolds number of 76 X 106/m. Two computational solutions, one using a Baldwin-Lomax algebraic turbulent eddy viscosity model and one using a modified K-£ (Rodi) turbulence model, are compared with experimental flowfield data obtained from a fast-response five-hole probe. Both the experiment and the computations show that the flowfield is dominated by a large, low-Mach-number, low-total-pressure separated region located on the interaction centerline. A comparison of the results shows significant differences between experiment and computations within this separated region. Outside the separated region, the experiment and computations are in good agreement. Additionally, the comparison shows that both turbulence models provide similar results, with neither model being clearly superior in predicting the flowfield.

Journal ArticleDOI
TL;DR: The results of an experimental investigation of high-enthalpy, hypersonic flow over sharp leading-edge compression corners are presented and discussed in this article, where the possible effects of real gas behavior are examined.
Abstract: The results of an experimental investigation of high-enthalpy, hypersonic flow over sharp leading-edge compression corners are presented and discussed. In particular, the possible effects of real gas behavior are examined. Measurements have been made of the heat transfer and pressure distributions for flat plate and compression corner flow. Some flow visualization data have also been obtained. Test flows were generated using a free-piston shock tunnel operating in the reflected mode. The reservoir enthalpy ranged from 3 to 19 MJ kg -1 , giving freestream speeds of 2.3-5.5 km s -1 . For these conditions, the flow remains laminar throughout. The flat plate data for both high- and low-enthalpy flows are in agreement with the reference enthalpy method for heat transfer and the weak interaction theory for pressure. Also, the measured flat plate boundary-layer thickness compares well with an expression strictly valid for perfect gas flows only. The high- and low-enthalpy compression corner flows have upstream influence and plateau pressure behavior similar to perfect gas flow. That is, real gas effects for the present flows appear to be negligible. This is consistent with the essentially chemically frozen viscous and inviscid flow upstream of the interaction.

Journal ArticleDOI
TL;DR: In this paper, a single-pulse coherent anti-Stokes Raman scattering (CARS) experiment employing a folded-box phase-matching geometry in a pulsed hypervelocity blunt body flow is presented.
Abstract: Broadband single pulse coherent anti-Stokes Raman scattering (CARS) experiments employing a folded-box phase-matching geometry in a pulsed hypervelocity blunt body flow are presented. Rovibrational spectra of molecular nitrogen, produced in the freestream and within the shock layer at moderately high enthalpy (8.4 MJ/kg), are examined. Difficulties peculiar to the application of a single pulse optical technique to a high enthalpy pulsed flow facility are discussed and measurements of flow temperatures are presented. Theoretically calculated values for temperatures based upon algorithms used to determine freestream and shock layer conditions agree well with experimental measurements using the CARS technique. The measurements indicate that thermal non-equilibrium conditions exist within the freestream, and that near thermal equilibrium exists at the point of measurement within the shock layer. The comparison between the experiment and theory in the shock layer is improved by using the measured freestream temperatures as input to the shock layer computations.

Journal ArticleDOI
TL;DR: In this paper, a numerical simulation of transient hypersonic flows involving shock-wave-freestream-disturbance interactions is presented using the essentially nonoscillatory (ENO) schemes.
Abstract: Numerical simulation of transient hypersonic flows involving shock-wave-freestream-disturbance interactions is presented using the essentially nonoscillatory (ENO) schemes. The ENO schemes were chosen for transient-flow simulations because they have high-order accuracy at extrema as well as in other parts of smooth solutions. First, the accuracy of the ENO schemes was tested numerically by applying them to the computations of a one-dimensional linear model equation and to an oscillating plate problem using the two-dimensional Navier-Stokes equations. Then, the third-order ENO scheme was used to compute the unsteady interaction of a freestream acoustic wave with a bow shock in hypersonic flow past a cylinder. The numerical results along the stagnation line were compared with linearized analytical solutions. The results show that the disturbance waves generated behind the bow shock are significantly amplified by the back-and-forth interactions and reflections of the acoustic waves. These results on the bow-shock-disturbance interactions will be useful in understanding the effects of the bow shock wave on the receptivity of hypersonic boundary layers to freestream disturbances.


Journal ArticleDOI
TL;DR: In this article, a 70-deg sphere-cone configuration model and the wake of the model on the sting were used to study the wake flow establishment process in a high-enthalpy impulse facility and a conventional perfect-gas air wind tunnel.
Abstract: Detailed aerodynamic heating measurements were made on a 70-deg sphere-cone configuration model and in the wake of the model on the sting. Tests were conducted in hypersonic flows in a high-enthalpy impulse facility, in which air and carbon dioxide were employed as test gases, and in a conventional perfect-gas air wind tunnel. Heating data were also obtained on three similar parametric forebody configurations. Normalized forebody Stanton number distributions were independent of Reynolds number and test gas, with the exception of smaller forebody corner heating peaks in carbon dioxide. Peak wake Stanton numbers were 5% of the forebody stagnation point values in the high-enthalpy tests and varied with Reynolds number from 7 to 15% of the stagnation point values in the perfect-gas tests. The impulse facility wake flow establishment process was studied in detail, and a criterion for determining when the wake flow becomes established was developed. Wake flow establishment was found to require on the order of 45-75 flow-path lengths as based on the model size and freestream velocity. Nomenclature C = ^Too/^ooT-* CH = Stanton number, q/[pooUoo(ho — hw)] cp = specific heat, J/kg-K h = enthalpy, J/kg k = thermal conductivity, W/m-K L = surface distance along sting from model base, m Q - heat energy, J/m2 q = heat transfer rate, W/m2