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At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Turbulent boundary layer in an adverse pressure gradient without effect of wall curvature

The hypersonic compressible turbulent boundary layer in an adverse pressure gradient along a cylindrical axisymmetric body was studied. The tests were conducted in a Mach 6 contoured axisymmetric nozzle. An external compression cowl was used to produce the gradual adverse pressure gradient and a maximum pressure rise of 7 times the freestream static pressure was achieved in a test region of 23 cm. Boundary layer profiles of static pressure, total pressure, and total temperature, as well as wall transient heat transfer rates were measured. Comparisons of the velocity total temperature profiles to linear and quadratic relations were made. Measured heat transfer data were in good agreement with the prediction from the flat-plate reference enthalpy method. Integral parameters were also in good agreement with results of numerical solutions for compressible turbulent boundary layer equations.

Zakkay, V.↗
NASA NTRSDate not supplied

CFD Simulation of an Inward-Turning Supersonic Inlet Unstart at Flight Mach Number 1.7

Interest in inward-turning supersonic inlets renewed during NASA’s Commercial Supersonic Technology Project, as a means of reducing disturbances caused by propulsion integration. Inward-turning designs have the potential to greatly reduce the strength of shock waves emanating outward from the nacelle, and therefore their contribution to the overall sonic boom signature. For top-mounted, or over-wing installations, disturbances to aircraft aerodynamics are reduced. The nature of the inward-turning geometry results in a boundary-layer thickening effect, opposite that of the familiar thinning of an outward-turning cone boundary-layer. The application of a small amount of boundary-layer bleed at the throat has been shown to alleviate the negative effects of the boundary-layer on subsonic diffuser performance. Boundary-layer bleed, however, comes at the expense of increased complexity, and external disturbances at bleed exits that would partially defeat the purpose of the inward-turning scheme. Vortex generators were proposed as an alternative to bleed, and an initial CFD study of their effectiveness in this application forms the basis of the present work. A source-term model in the Wind-US CFD code was used to model the effects of an array of vortex generators placed at the entrance to the subsonic diffuser. Baseline results with no vortex generators compared favorably to previously reported results with no bleed at back-pressures near the design point, but stability margin in the present case was limited by a curious non-linear phenomena similar to unstart in mixed-compression inlets. While some improvement was noted with vortex generators, their effectiveness was diminished by the limited stability margin. A time-accurate calculation was subsequently performed to better understand this process. Results of this calculation are presented and instantaneous solutions compared with steady-state, subcritical results previously reported. The stability margin limitation is attributed to the use of a constant-pressure outflow boundary condition, which differed from that used in the prior simulations.

Pre-Compressor Cooling, Water Injection, Hypersoni↗
NASA NTRSDate not supplied

Thermomechanical Relaxation of Vibrationally Excited NO in a Hypersonic Turbulent Boundary Layer

The effects of thermal nonequilibrium are studied using a DC glow discharge plasma to selectively modify the vibrational temperature of an otherwise cold flow. These experiments were performed at 𝑴=5.7and 𝑹𝒆=6×10 6 /m in the Actively Controlled Expansion tunnel at Texas A&M University on a canonical 2.75° half-angle wedge test article. A∼47 W plasma was generated across spanwise electrodes, and removable trips produced either a laminar or turbulent boundary layer. NO planar laser-induced fluorescence (PLIF) quantified the extent of the NO 𝑿 2 𝚷 (𝒗=1) vibrational perturbation due to the plasma and its subsequent interplay with the boundary layer, the first such data of which the authors are aware. NOPLIF measurements were taken at three locations along the test article, and quasi-temperature measurements are reported at each. Despite not quantifiably perturbing the boundary layer, the plasma was shown to alter the NO vibrational temperature profile therein. Independent analysis of the thermodynamics and fluid mechanics of the system suggests each set of physics could explain the results observed near and far from the wall, respectively.

Casey J. Broslawski↗
NASA NTRSDate not supplied

Thermomechanical Relaxation of Vibrationally Excited NO in a Hypersonic Turbulent Boundary Layer

The effects of thermal nonequilibrium are studied using a DC glow discharge plasma to selectively modify the vibrational temperature of an otherwise cold flow. These experiments were performed at 𝑴=5.7and 𝑹𝒆=6×10 6 /m in the Actively Controlled Expansion tunnel at Texas A&M University on a canonical 2.75° half-angle wedge test article. A∼47 W plasma was generated across spanwise electrodes, and removable trips produced either a laminar or turbulent boundary layer. NO planar laser-induced fluorescence (PLIF) quantified the extent of the NO 𝑿 2 𝚷 (𝒗=1) vibrational perturbation due to the plasma and its subsequent interplay with the boundary layer, the first such data of which the authors are aware. NOPLIF measurements were taken at three locations along the test article, and quasi-temperature measurements are reported at each. Despite not quantifiably perturbing the boundary layer, the plasma was shown to alter the NO vibrational temperature profile therein. Independent analysis of the thermodynamics and fluid mechanics of the system suggests each set of physics could explain the results observed near and far from the wall, respectively.

Casey J. Broslawski↗
NASA NTRS1950

Investigation of Operating Pressure Ratio of a Supersonic Wind Tunnel Utilizing Distributed Boundary-layer Suction in Test Section

Effect of distributed boundary-layer suction on operating pressure ratio of a supersonic wind tunnel was investigated. Investigation was made in 3.84- by 10-inch supersonic tunnel operating at Mach number 2.0 and suction was applied in neighborhood of the normal shock to two walls of a constant-area extension of test section. A reduction of 4 percent of operating pressure ratio was attributed to improved flow conditions at subsonic-diffuser inlet. The theoretical normal shock was, in practice, replaced by a multiple-branch shock configuration across which the flow parameters changed in approximate accordance with the Rankine-Hugoniot values.

Cohen, C B↗
NASA NTRS1990

Modelling the effect of vegetation on the formation of boundary layer cumulus during FIFE

A 1D model is discussed which permits prediction of the effect of vegetation cover on the area-averaged boundary layer. Special attention is given to the manner in which the relative contributions of sensible and latent heat fluxes to the surface-energy budget are determined by the vegetation. The model can be effectively applied to the heterogeneous prairie environment to predict cloud cover at different times of the year demonstrating good agreement with observed conditions.

Wetzel, Peter J.↗
NASA NTRS1989

Curvature effects on the stability of laminar boundary layers on swept wings

The stability of the laminar boundary layer on a swept wing is examined. An improved linear stability theory which includes the effects of body and streamline curvature and compressibility is utilized for the calculations. The computed N-factor is correlated with the onset of transition. For this study, only test conditions where transition is due to the growth of highly amplified crossflow instabilities on convex surfaces are examined. The calculations show that the effect of the curvature terms is to dramatically reduce local amplification rates in regions where body and streamline curvature are large. For the cases where transition occurred ahead of the pressure minimum on the upper surface of the wing, the N-factor at transition onset is near 9 when the effects of body and streamline curvature are included in the computations. When the curvature terms are neglected, the average N-factor is about 17. The calculations show that traveling crossflow waves are most amplified.

Collier, F. S., Jr.↗
NASA NTRS1978

A computational scheme usable for calculating the plume backflow region

The effects of the nozzle wall boundary layer on the plume flowfield are neglected in the majority of computational schemes which exist for the calculation of rocket engine exhaust plume flowfields. This neglect, which is unimportant in many applications, becomes unacceptable for applications where a surface which can be adversely affected by plume impingement forces, heating, or contamination is located behind the nozzle exit plane in what is called the 'plume backflow region'. The flow in this region originates in, and is highly affected by, the nozzle wall boundary layer. The inclusion of the effects of the boundary layer in the calculations is required for an appropriate determination of the flowfield properties within this region. A description is presented of the results of modifications of a method-of-characteristics computer program. The modifications were made to include the effects of the nozzle wall boundary layer on the plume flowfield. A comparison of computed and experimental data indicates that the employed computer program may be a useful tool for calculating the entire plume flowfield for liquid propellant rocket engines.

Cooper, B. P., Jr.↗
NASA NTRS1986

Effects of compressibility and free-stream turbulence on boundary layer transition in high-subsonic and transonic flows

Based on the existing boundary layer transition data, the effects of compressibility, pressure fluctuations, and free-stream turbulence have been reexamined for subsonic and transonic flow speeds. It is confirmed that the compressibility effects may be adequately expressed in terms of a simple correlation with free-stream Mach number. Pressure fluctuations, especially at low levels, do not seem to significantly affect the transition phenomenon. Effects of free-stream turbulence in high-subsonic and transonic flows are similar to the trends observed for low-speed flows and the transition process is hastened. The trends, as seen from slender cone flow data, seem to suggest power law correlations between transition Reynolds number and free-stream turbulence.

Murthy, S. V.↗
NASA NTRS1981

Numerical Boundary Condition Procedures

Topics include numerical procedures for treating inflow and outflow boundaries, steady and unsteady discontinuous surfaces, far field boundaries, and multiblock grids. In addition, the effects of numerical boundary approximations on stability, accuracy, and convergence rate of the numerical solution are discussed.

Source record↗
NASA NTRS1960

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

Boltz, Frederick W.↗
NASA NTRS1979

The Jovian boundary layer as formed by magnetic-anomaly effects

A model is presented in which a plasma boundary layer of Jupiter is formed from plasma of internal origin. It is proposed that, unlike the Earth's boundary layer, which is thought to consist principally of solar wind plasma, Jupiter's boundary layer consists principally of sulphur and oxygen from the Io plasma torus, plus a small component of hydrogen from Jupiter's ionosphere. Fresh plasma is supplied to the boundary layer once each planetary rotation period by a convection pattern that rotates with Jupiter.

Dessler, A. J.↗
NASA NTRS1991

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 5: Propagation of propeller tone noise through a fuselage boundary layer

An analysis of tone noise propagation through a boundary layer and fuselage scattering effects was derived. This analysis is a three dimensional and the complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The outer wave field is constructed analytically from an incident wave appropriate to the source and a scattered wave in the standard Hankel function form. For the incident wave, an existing function - domain propeller noise radiation theory is used. In the boundary layer region, the wave equation is solved by numerical methods. The theoretical analysis is embodied in a computer program which allows the calculation of correction factors for the fuselage scattering and boundary layer refraction effects. The effects are dependent on boundary layer profile, flight speed, and frequency. Corrections can be derived for any point on the fuselage, including those on the opposite side from the source. The theory was verified using limited cases and by comparing calculations with available measurements from JetStar tests of model prop-fans. For the JetStar model scale, the boundary layer refraction effects produce moderate fuselage pressure reinforcements aft of and near the plane of rotation and significant attenuation forward of the plane of rotation at high flight speeds. At lower flight speeds, the calculated boundary layer effects result in moderate amplification over the fuselage area of interest. Apparent amplification forward of the plane of rotation is a result of effective changes in the source directivity due to boundary layer refraction effects. Full scale effects are calculated to be moderate, providing fuselage pressure amplification of about 5 dB at the peak noise location. Evaluation using available noise measurements was made under high-speed, high-altitude flight conditions. Comparisons of calculations made of free field noise, using a current frequency-domain propeller noise prediction method, and fuselage effects using this new procedure show good agreement with fuselage measurements over a wide range of flight speeds and frequencies. Correction factors for the JetStar measurements made on the fuselage are provided in an Appendix.

Magliozzi, B.↗
NASA NTRS1997

Effect of Flow Misalignment and Multi-Hole Interaction on Boundary-Layer Bleed Hole Flow Coefficient Behavior

The effect of flow misalignment on the flow coefficient behavior of a 20 deg. boundary-layer bleed hole and the effect of the interaction between two 90 deg. bleed holes separated by two hole diameters on flow coefficient behavior has been studied experimentally. Both tests were run at freestream Mach numbers of 0.61, 1.62 and 2.49. The flow misalignment study was conducted over a range of 0 to 30 deg. The results show that neither flow misalignment nor hole interaction has much effect on the flow coefficient for the subsonic case. For the supersonic cases, flow misalignment causes significant degradation in the performance of the slant hole. For the supersonic normal hole interaction cases, depending on the hole orientation, either an increase or decrease in overall flow coefficient was observed. The largest change in flow coefficient, 6% increase at near choke conditions, occurred when the holes were oriented in line with the flow direction.

Davis, David O.↗
NASA NTRS1994

Boundary Conditions for Jet Flow Computations

Ongoing activities are focused on capturing the sound source in a supersonic jet through careful large eddy simulation (LES). One issue that is addressed is the effect of the boundary conditions, both inflow and outflow, on the predicted flow fluctuations, which represent the sound source. In this study, we examine the accuracy of several boundary conditions to determine their suitability for computations of time-dependent flows. Various boundary conditions are used to compute the flow field of a laminar axisymmetric jet excited at the inflow by a disturbance given by the corresponding eigenfunction of the linearized stability equations. We solve the full time dependent Navier-Stokes equations by a high order numerical scheme. For very small excitations, the computed growth of the modes closely corresponds to that predicted by the linear theory. We then vary the excitation level to see the effect of the boundary conditions in the nonlinear flow regime.

Hayder, M. E.↗
NASA NTRS1981

Turbulent boundary layer heat transfer experiments - A separate effects study on a convexly-curved wall

Surface heat transfer rates have been measured for several different flows on an isothermal, convexly curved surface. The freestream velocity, boundary layer thickness, acceleration parameter, and unheated starting length were varied systematically, and both turbulent and transitional boundary layers were studied. The effect of convex curvature on heat transfer rates is significant with Stanton numbers reduced 20-25% below flat wall values for the same enthalpy thickness Reynolds number. Heat transfer rates recovered slowly on a flat wall downstream of the curved wall, and after 60 cm, the Stanton numbers were still 15-20% below flat wall values. The behavior of the boundary layer suggests the existence of an asymptotic condition. Boundary layer thickness, freestream velocity, and boundary layer maturity affect the initial response to the introduction of curvature and the rate at which the asymptotic state is approached. Convex curvature appears to increase the boundary layer's sensitivity to acceleration; it also delays and retards transition. Near-laminar or early-transitional boundary layers recover from curvature rapidly, whereas late-transitional and mature boundary layers recover slowly.

Simon, T. W.↗