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

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Characteristics of the unsteady shock-induced laminar boundary layer on a flat plate.

The unsteady shock-induced laminar boundary layer on a flat plate is theoretically studied for shock speeds ranging from 1.12 to 9 km/sec. Boundary layer flows for real air, assumed to be in thermochemical equilibrium, are analyzed by extending Lam's work on shock-induced laminar boundary layers to equilibrium dissociated and ionized flows. A complete description of the unsteady nature of the boundary layer is presented in terms of heat transfer and several boundary layer thickness quantities as functions of a single time-position variable. Boundary layer development is considered for two points of view. Time-dependent boundary layer development and approach to steady state for any fixed position on the plate is described, as is the configuration of the boundary layer with position on the plate at any point in time.

Cook, W. J.↗
NASA NTRS1980

Coherent substructure of turbulence near the stagnation zone of a bluff body

The evolution of freestream turbulence in crossflow about a circular cylinder was studied in order to identify the existence of a coherent substructure which is the outcome of the amplification of freesteam turbulence by the stretching mechanism in diverging flow about a bluff body. Visualization of the flow events revealed the selective stretching of cross-vortex tubes and the emergence of an organized turbulent flow pattern near the cylinder stagnation zone. Significant amplification of the total turbulent energy of the streamwise fluctuating velocity was consistently monitored. Realization of selective amplification at scales larger than the neutral scale of the stagnation flow was indicated by the variation of the discrete streamwise turbulent energy. A most amplified scale, characteristic of the energy containing eddies within the coherent substructure and commensurate with the boundary-layer thickness, was detected. Penetration of the amplified turbulence into the cylinder boundary layer led to the retardation of separation and to a concurrent decrease in the drag coefficient at subcritical cylinder-diameter Reynolds numbers.

Sadeh, W. Z.↗
NASA NTRS1989

An experimental study of near wall flow parameters in the blade end-wall corner region

The near wall flow parameters in the blade end-wall corner region is investigated. The blade end-wall corner region was simulated by mounting an airfoil section (NACA 65-015 base profile) symmetric blades on both sides of the flat plate with semi-circular leading edge. The initial 7 cm from the leading edge of the flat plate was roughened by gluing No. 4 floor sanding paper to artificially increase the boundary layer thickness on the flat plate. The initial flow conditions of the boundary layer upstream of the corner region are expected to dictate the behavior of flow inside the corner region. Therefore, an experimental investigation was extended to study the combined effect of initial roughness and increased level of free stream turbulence on the development of a 2-D turbulent boundary layer in the absence of the blade. The measurement techniques employed in the present investigation included, the conventional pitot and pitot-static probes, wall taps, the Preston tube, piezoresistive transducer and the normal sensor hot-wire probe. The pitot and pitot-static probes were used to obtain mean velocity profile measurements within the boundary layer. The measurements of mean surface static pressure were obtained with the surface static tube and the conventional wall tap method. The wall shear vector measurements were made with a specially constructed Preston tube. The flush mounted piezoresistive type pressure transducer were employed to measure the wall pressure fluctuation field. The velocity fluctuation measurements, used in obtaining the wall pressure-velocity correlation data, were made with normal single sensor hot-wire probe. At different streamwise stations, in the blade end-wall corner region, the mean values of surface static pressure varied more on the end-wall surface in the corner region were mainly caused by the changes in the curvature of the streamlines. The magnitude of the wall shear stress in the blade end-wall corner region increased significantly in the close vicinity of the corner line. The maximum value of the wall shear stress and its location from the corner line, on both the surfaces forming the corner region, were observed to change along the corner. These observed changes in the maximum values of the wall shear stress and its location from the corner line could be associated with the stretching and attenuation of the horseshoe vortex. The wall shear stress vectors in the blade end-wall corner region were observed to be more skewed on the end-wall surface as compared to that on the blade surface. The differences in the wall shear stress directions obtained with the Preston tube and flow visualization method were within the range in which the Preston tube was found to be insensitive to the yaw angle.

Bhargava, Rakesh K.↗
NASA NTRS1991

Evaluation of a technique to generate artificially thickened boundary layers in supersonic and hypersonic flows

The feasibility of using a contoured honeycomb model to generate a thick boundary layer in high-speed, compressible flow was investigated. The contour of the honeycomb was tailored to selectively remove momentum in a minimum of streamwise distance to create an artificially thickened turbulent boundary layer. Three wind tunnel experiments were conducted to verify the concept. Results indicate that this technique is a viable concept, especially for high-speed inlet testing applications. In addition, the compactness of the honeycomb boundary layer simulator allows relatively easy integration into existing wind tunnel model hardware.

Porro, A. R.↗
NASA NTRS1985

Application of axisymmetric analogue for calculating heating in three-dimensional flows

A rapid, approximate method has been developed for calculating the heating rates on three-dimensional vehicles such as the Space Shuttle Orbiter and other advanced reentry configurations. The method is based on the axisymmetric analogue for three-dimensional boundary layers. It uses information obtained from a three-dimensional inviscid flowfield solution, such as HALIS, to calculate inviscid surface streamlines along which approximate heating rates are calculated independent of what happens along other streamlines. Three-dimensional effects are included through the metric coefficient that describes the divergence or convergence of streamlines. Boundary-layer edge properties are obtained from the inviscid flowfield solution by interpolating in the inviscid flowfield at a distance equal to the boundary-layer thickness away from the wall. This accounts, approximately, for the variable boundary-layer edge entropy. Using this method, heating calculations can be made along a typical streamline in a few seconds. This method has been used to accurately predict heating rates for simple shapes such as a spherically blunted cone and more complex shapes such as the Shuttle Orbiter for a variety of wind-tunnel and flight conditions. A unique feature of the method is its ability to accurately predict heating rates on the Shuttle Orbiter wing.

Hamilton, H. H.↗
NASA NTRS1975

Spinning mode sound propagation in ducts with acoustic treatment and sheared flow

The propagation of spinning mode sound was considered for a cylindrical duct with sheared steady flow. Calculations concentrated on the determination of the wall optimum acoustic impedance and the maximum possible attenuation. Both the least attenuated and higher radial modes for spinning lobe patterns were considered. A parametric study was conducted over a wide range of Mach numbers, spinning lobe numbers, sound frequency, and boundary layer thickness. A correlation equation was developed from theoretical considerations starting with the thin boundary layer approximation of Eversman. This correlation agrees well with the more exact calculations for inlets and provides a single boundary layer refraction parameter which determines the change in optimum wall impedance due to refraction effects.

Rice, E. J.↗
NASA NTRS1975

Spinning mode sound propagation in ducts with acoustic treatment and sheared flow

The propagation of spinning mode sound was considered for a cylindrical duct with sheared steady flow. The calculations concentrated on the determination of the wall optimum acoustic impedance and the maximum possible attenuation. Both the least attenuated and higher radial modes for spinning lobe patterns were considered. A parametric study was conducted over a wide range of Mach numbers, spinning lobe numbers, sound frequency, and boundary layer thickness. A correlation equation was developed from theoretical considerations starting with the thin boundary layer approximation of Eversman. This correlation agrees well with the more exact calculations for inlets and provides a single boundary layer refraction parameter which determines the change in optimum wall impedance due to refraction effects.

Rice, E. J.↗
NASA NTRS1991

A steadying effect of acoustic excitation on transitory stall

The effect of acoustic excitation on a class of separated flows with a transitional boundary layer at the point of separation is considered. Experimental results on the flow over airfoils, a two-dimensional backward-facing step, and through large angle conical diffusers are presented. In all cases, the separated flow undergoes large amplitude fluctuations, much of the energy being concentrated at unusually low frequencies. In each case, an appropriate high frequency acoustic excitation is found to be effective in reducing the fluctuations substantially. The effective excitation frequency scales on the initial boundary layer thickness and the effect is apparently achieved through acoustic tripping of the separating boundary layer.

Zaman, K. B. M. Q.↗
NASA NTRS1991

A steadying effect of acoustic excitation on transitory stall

The effect of acoustic excitation on a class of separated flows with a transitional boundary layer at the point of separation is considered. Experimental results on the flow over airfoils, a two-dimensional backward-facing step, and through large angle conical diffusers are presented. In all cases, the separated flow undergoes large amplitude fluctuations, much of the energy being concentrated at unusually low frequencies. In each case, an appropriate high frequency acoustic excitation is found to be effective in reducing the fluctuations substantially. The effective excitation frequency scales on the initial boundary layer thickness and the effect is apparently achieved through acoustic tripping of the separating boundary layer.

Zaman, K. B. M. Q.↗
NASA NTRS1975

Excitation of a Helmholtz resonator by a turbulent boundary layer

A turbulent boundary layer on the fuselage of a glider was used to excite a Helmholtz resonator. The resonator orifice was flush with the surface and smaller than the boundary layer thickness. Resonator frequencies were chosen so they would tune with different portions of the boundary layer wall pressure spectrum. The resonators were excited at both the Helmholtz frequency and a standing wave frequency. The results show a shift in the Helmholtz frequency when the boundary layer is present. This shift indicates the degree to which the turbulence interacts with the acoustic motion in the orifice and modifies the end correction. Four of the nine resonators tested were 'strongly' excited and radiated considerable acoustic sound. This phenomenon occurs when turbulent eddies about the size of the orifice are convected past the orifice with the proper velocity to impose a frequency that matches a resonant frequency of the resonator. Turbulent eddies which are either smaller or larger than the orifice will not produce strong excitation.

Panton, R. L.↗
NASA NTRS1989

The response of a laminar boundary layer in supersonic flow to small amplitude progressive waves

The effect of a small amplitude progressive wave on the laminar boundary layer on a semi-infinite flat plate, due to a uniform supersonic freestream flow, is considered. The perturbation to the flow divides into two streamwise zones. In the first, relatively close to the leading edge of the plate, on a transverse scale comparable to the boundary layer thickness, the perturbation flow is described by a form of the unsteady linearized compressible boundary layer equations. In the freestream, this component of flow is governed by the wave equation, the solution of which provides the outer velocity conditions for the boundary layer. This system is solved numerically, and also the asymptotic structure in the far downstream limit is studied. This reveals a breakdown and a subsequent second streamwise zone, where the flow disturbance is predominantly inviscid. The two zones are shown to match in a proper asymptotic sense.

Duck, Peter W.↗
NASA NTRS1976

Measurements in the near-wall region of a relaxing three-dimensional low speed turbulent air boundary layer

An experimental investigation was conducted at selected locations of the near-wall region of a three dimensional turbulent air boundary layer relaxing in a nominally zero external pressure gradient behind a transverse hump (in the form of a 30 deg swept, 5-foot chord wing-type model) faired into the side wall of a low speed wind tunnel. Wall shear stresses measured with a flush-mounted hot-film gage and a sublayer fence were in very good agreement with experimental data obtained with two Preston probes. With the upstream unit Reynolds number held constant at 325,000/ft. approximately one-fourth of the boundary layer thickness adjacent to the wall was surveyed with a single rotated hot-wire probe mounted on a specially designed minimum interference traverse mechanism. The boundary layer (approximately 3.5 in thick near the first survey station where the length Reynolds number was 5.5 million) had a maximum crossflow velocity ratio of 0.145 and a maximum crossflow angle of 21.875 deg close to the wall.

Hebbar, K. S.↗
NASA NTRS1972

Pitot-probe displacement in a supersonic turbulent boundary layer

Eight circular pitot probes ranging in size from 2 to 70 percent of the boundary-layer thickness were tested to provide experimental probe displacement results in a two-dimensional turbulent boundary layer at a nominal free-stream Mach number of 2 and unit Reynolds number of 8 million per meter. The displacement obtained in the study was larger than that reported by previous investigators in either an incompressible turbulent boundary layer or a supersonic laminar boundary layer. The large probes indicated distorted Mach number profiles, probably due to separation. When the probes were small enough to cause no appreciable distortion, the displacement was constant over most of the boundary layer. The displacement in the near-wall region decreased to negative displacement in some cases. This near-wall region was found to extend to about one probe diameter from the test surface.

Allen, J. M.↗
NASA NTRS1983

Assessment of NASA and RAE viscous-inviscid interaction methods for predicting transonic flow over nozzle afterbodies

The Langley Research Center of the National Aeronautics and Space Administration and the Royal Aircraft Establishment have undertaken a cooperative program to conduct an assessment of their patched viscous-inviscid interaction methods for predicting the transonic flow over nozzle afterbodies. The assessment was made by comparing the predictions of the two methods with experimental pressure distributions and boattail pressure drag for several convergent circular-arc nozzle configurations. Comparisons of the predictions of the two methods with the experimental data showed that both methods provided good predictions of the flow characteristics of nozzles with attached boundary layer flow. The RAE method also provided reasonable predictions of the pressure distributions and drag for the nozzles investigated that had separated boundary layers. The NASA method provided good predictions of the pressure distribution on separated flow nozzles that had relatively thin boundary layers. However, the NASA method was in poor agreement with experiment for separated nozzles with thick boundary layers due primarily to deficiencies in the method used to predict the separation location.

Putnam, L. E.↗
NASA NTRS2017

Direct Numerical Simulation of Hypersonic Turbulent Boundary Layer inside an Axisymmetric Nozzle

As a first step toward a study of acoustic disturbance field within a conventional, hypersonic wind tunnel, direct numerical simulations (DNS) of a Mach 6 turbulent boundary layer on the inner wall of a straight axisymmetric nozzle are conducted and the results are compared with those for a flat plate. The DNS results for a nozzle radius to boundary-layer thickness ratio of 5:5 show that the turbulence statistics of the nozzle-wall boundary layer are nearly unaffected by the transverse curvature of the nozzle wall. Before the acoustic waves emanating from different parts of the nozzle surface can interfere with each other and undergo reflections from adjacent portions of the nozzle surface, the rms pressure fluctuation beyond the boundary layer edge increases toward the nozzle axis, apparently due to a focusing effect inside the axisymmetric configuration. Spectral analysis of pressure fluctuations at both the wall and the freestream indicates a similar distribution of energy content for both the nozzle and the flat plate, with the peak of the premultiplied frequency spectrum at a frequency of [(omega)(delta)]/U(sub infinity) approximately 6.0 inside the free stream and at [(omega)(delta)]/U(sub infinity) approximately 2.0 along the wall. The present results provide the basis for follow-on simulations involving reverberation effects inside the nozzle.

Huang, Junji↗
NASA NTRS1978

Thin-layer approximation and algebraic model for separated turbulent flows

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.

Baldwin, B.↗
NASA NTRS1993

A feasibility study of using Langley 0.3-m transonic cryogenic tunnel sidewall boundary-layer removal system for heavy gas testing

This report presents the results of a preliminary study for using the 0.3-m Transonic Cryogenic Tunnel sidewall boundary-layer removal system with heavy gas sulfur hexafluoride as the test medium. It is shown that the drive motor speed/power of the existing system and the additional heat load on the tunnel heat exchanger are the major problems limiting the boundary-layer removal system performance. Overcoming these problems can provide the capability to remove about 1.5 percent of the test section mass flow at Mach number M = 0.8 and about 5 percent at M = 0.25. Previous studies have shown that these boundary-layer mass flow removal rates can reduce the boundary-layer thickness by a factor of two at the model station. Also the effect of upstream boundary-layer removal on the airfoil test data is not likely to be significant under high lifting conditions. Near design conditions, corrections to the test Mach number may be necessary to account for sidewall boundary-layer effects.

Murthy, A. V.↗
NASA NTRS1983

An experimental study of surface curvature effects on a supersonic turbulent boundary layer

The mean flow results of an experimental study of compressible turbulent boundary layers in an adverse pressure gradient with and without surface curvature effects are presented. The test was conducted in an axisymmetric flow facility. The upstream Reynolds number based on boundary layer momentum thickness was 5884 and the boundary layer thickness was 0.90 cm. The curvature effects were examined by studying two flows with essentially identical adverse pressure gradients. One flow was along a concave compression surface test section, while the other was along a straight-walled test section. Mean flow measurements included wall static pressure distributions, wall temperatures, pitot pressure profiles and total temperature profiles. The mean flow results indicated that the surface curvature resulted in a definite increase of turbulent mixing in the boundary layer.

Chou, J. H.↗