SEARCH · Search NASA

Results for “Boundary layer thickness”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 235 records · Page 13

All sources NASA NTRS DOE OSTI

NASA NTRS1983

Airfoil self noise - Effect of scale

Key data from a comprehensive airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional sharp trailing-edge models. The models include seven NACA 0012 airfoil sections and five 'flat plate' sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity, angle of attack to the flow, and boundary layer turbulence through natural transition and through tripping. Detailed aerodynamic measurements, of pertinence to the scaling problem of airfoil self-noise, were conducted in the near-wake of the trailing edges. Presented are mean and rms turbulent velocity profiles as well as boundary layer thicknesses and integral thickness parameters for a large range of conditions. The noise spectra of the self-noise sources were determined by the use of a cross-spectral technique. The spectra were normalized using the measured aerodynamic parameters in order to evaluate the most common scaling law now in use. An examination of the Reynolds number dependence of the overall self-noise levels has revealed a new and useful scaling result. This result appears to quantify the transition between turbulent-boundary-layer trailing edge noise and laminar-boundary-layer vortex shedding noise.

Brooks, T. F.↗

NASA NTRS1974

A review of turbulence measurements in compressible flow

A survey of turbulence measurements in compressible flows is presented. The majority of turbulence measurements at super- and hypersonic speeds were made for the zero pressure gradient, turbulent boundary layer. It was found that the nondimensional turbulent stress terms for the zero pressure gradient flow appear to agree closely with equivalent incompressible measurements in the outer part of the boundary layer. The stress terms were nondimensionalized by the wall value of shear stress and plotted versus the distance from the wall, nondimensionalized by the boundary-layer thickness. Indirect evaluation of the total shear stress distribution from mean velocity measurements for both super- and hypersonic flows (zero pressure gradient, two-dimensional flows) indicate a near universal distribution. These total shear stress curves also agree very closely with measured incompressible shear stress distributions. Recent laser anemometer measurements of the turbulent Reynolds shear stress (puv), reported by Johnson and Rose for a Mach number 2.9 flow, are in reasonable agreement with the expected total shear stress curve over the outer 60% of the boundary layer.

Sandborn, V. A.↗

NASA NTRS1974

Heat-transfer and pressure distributions for laminar separated flows downstream of rearward-facing steps with and without mass suction

Heat-transfer and pressure distributions were measured for laminar separated flows downstream of rearward-facing steps with and without mass suction. The flow conditions were such that the boundary-layer thickness was comparable to or larger than the step height. For both suction and no-suction cases, an increase in the step height resulted in a sharp decrease in the initial heat-transfer rates behind the step. Downstream, however, the heat transfer gradually recovered back to less than or near attached-flow values. Mass suction from the step base area increased the local heat-transfer rates; however, this effect was relatively weak for the laminar flows considered. Even removal of the entire approaching boundary layer raised the post-step heat-transfer rates only about 10 percent above the flatplate values. Post-step pressure distributions were found to depend on the entrainment conditions at separation. In the case of the solid-faced step, a sharp pressure drop behind the step was followed by a very short plateau and relatively fast recompression. For the slotted-step connected to a large plenum but without suction, the pressure drop at the base was much smaller and the downstream recompression more gradual than that for solid-faced step.

Brown, R. D.↗

NASA NTRS2018

Influence of a Backward-Facing Step on Swept-Wing Boundary-Layer Transition

Experimental measurements were performed on a swept flat-plate model with an airfoil leading edge and imposed chordwise pressure gradient to determine the effects of a backward-facing step on transition in a low-speed stationary crossflow-dominated boundary layer. Detailed hot-wire measurements were performed for three step heights ranging from 36 to 49% of the boundary-layer thickness at the step and corresponding to subcritical, nearly critical, and critical cases. In general, the step had a small localized effect on the growth of the stationary crossflow vortex, whereas the unsteady disturbance amplitudes increased with increasing step height. Intermittent spikes in instantaneous velocity began to appear for the two larger step heights. A physical explanation was provided for the mechanism leading to transition and the sudden movement in the transition front due to the critical steps. The large localized velocity spikes, which ultimately led to an intermittent breakdown of the boundary layer, were the result of nonlinear interactions of the different types of unsteady instabilities with each other and with the stationary crossflow vortices. Thus, the unsteady disturbances played the most important role in transition, but the stationary crossflow vortices also had a significant role via the modulation and the increased amplitude of the unsteady disturbances.

Eppink, Jenna L.↗

NASA NTRS1982

Viscous effect on airfoils for unsteady transonic flows

The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.

Lee, S. C.↗

NASA NTRS1983

Embedded cavity drag in steady and unsteady flows

The numerical solution of the laminar boundary-layer flow over an embedded cavity is studied. The purpose is to examine the relevant drag characteristics of laminar cavity flow. The solution field is obtained in terms of velocity and vorticity variables, with the stream function and pressure derivable from the directly computed variables. An analysis and comparison is made among four square cavities, ranging in size from 0.25 to 1.00 boundary-layer thicknesses deep. The dominant flow features are examined in the vicinity of the cavity by means of the stream function and iso-vorticity contours. The dominant physics in the overall drag characteristics of the flow is examined by an analysis of the pressure and wall shear stress distributions in the cavity, and upstream and downstream of the cavity. Pressure forces and frictional forces in, and in the vicinity of, the cavity are determined. Stress relaxation distances, both upstream and downstream of the cavity, are calculated and analyzed. The flow dynamics of the boundary-layer flow over an embedded cavity is summarized. Finally, the relevance of the results to the control of flow separation in such flows is discussed.

Gatski, T. B.↗

NASA NTRS1984

Embedded cavity drag in steady and unsteady flows

The numerical solution of the laminar boundary-layer flow over an embedded cavity is studied. The purpose of the study is to examine the relevant drag characteristics of laminar cavity flow. The solution field is obtained in terms of velocity and vorticity variables, with the stream function and pressure derivable from the directly computed variables. An analysis and comparison is made among four square cavities, ranging in size from 0.25 to 1.00 boundary-layer thicknesses deep. The dominant flow features are examined in the vicinity of the cavity by means of the stream function and iso-vorticity contours. The dominant physics in the overall drag characteristics of the flow is examined by an analysis of the pressure and wall shear stress distributions in the cavity, and upstream and downstream of the cavity. Pressure forces and frictional forces in, and in the vicinity of, the cavity are determined. Stress relaxation distances, both upstream and downstream of the cavity, are calculated and analyzed. The flow dynamics of the boundary-layer flow over an embedded cavity is summarized. Finally, the relevance of the present results to the control of flow separation in such flows is discussed.

Gatski, T. B.↗

NASA NTRS1987

What happens to the large eddies when net drag reduction is achieved by outer flow manipulators?

The tandem-arranged parallel plate manipulator (TAPPM) presently used to generate a flow with net drag reduction is conditionally sampled and found to exhibit significant changes of the boundary layer's large-scale motions at 20 delta(0), where delta(0) is the boundary layer thickness at the upstream plate of the TAPPM. Flow fields in the 'valleys' are equally affected. Both turbulent and irrotational flows then reestablish themselves by 50 delta(0). Flow visualization indicates that the TAPPM wake is very coherent at 20 delta(0), and has not yet spread into the wall region, while large scale motions and the ambient response flow continue to exhibit dynamic changes. This indicates that the shielding effect of the TAPPM's wake prolongs suppression of large-scale motions, thereby reducing skin friction over a sufficiently long extent of the boundary layer to overcome device drag and yield net drag reductions.

Falco, R. E.↗

NASA NTRS1995

Experimental study of the use of vortex generators to reduce fluctuating pressure loads in shock wave turbulent boundary layer interactions

Fluctuating wall pressure measurements have been made in a separated shock wave/turbulent boundary layer interaction produced by an unswept compression corner in a Mach 5 flow. Wheeler doublet vortex generators were placed 15.8 boundary layer thicknesses upstream of the corner to study their effect on the fluctuating pressure loads produced by the translating separation shock. The vortex generators produced significant three-dimensionality in an otherwise two-dimensional interaction. They reduced the upstream influence and the length of the region of shock motion by 60% and 64%, respectively, decreased the maximum wall pressure rms by 23%, and shifted the fluctuations to a higher frequency band. The maximum fraction of energy in the 100-500 Hz frequency band is decreased by 11%. These changes are due to a fuller boundary layer profile, a weaker separation shock, and increased boundary layer turbulence causing increased separation shock jitter.

Barter, John W.↗

NASA NTRS1972

Compressible boundary layer with normal pressure gradients: Quasi-similarity equations - Their properties at the wall and at sharp and blunt leading edges.

The failure of most viscous-inviscid interaction methods at strong interactions is attributed to the presence of a normal pressure gradient. A new theory is proposed for supersonic laminar boundary layers that can generate normal pressure gradients. The Navier-Stokes equations are reexamined by an order of magnitude analysis and all first and second order terms are retained. The approximation is found to be dependent not only on the boundary layer thickness but also on the ratio of the dimensionless viscosity and density. The equations are transformed into two quasi-similar, nonlinear, third order, ordinary integro-differential equations for the velocity and pressure as functions of a single transverse variable. The properties of the equations at the boundaries are discussed.

Seginer, A.↗

NASA NTRSDate not supplied

On A Stabilization of the Ingard-Myers Impedance Boundary Condition

It has been well-known that the Ingard-Myers impedance condition, while simple to apply, is subject to the hydrodynamic Kelvin-Helmholtz instability due to its use of a vortex sheet in modeling the flow at the liner boundary. Recently, in the development of a time domain boundary element method for acoustic scattering by treated surfaces, it was found that by neglecting a certain second-order spatial derivative term in the Ingard-Myers formulation, the hydrodynamic instability can be avoided. The present paper aims to provide further analysis of this modified condition, hereby referred to as Truncated Ingard-Myers Impedance Boundary Condition (TIMIBC). It will be shown, based on the dispersion relations of linear waves, that the instability intrinsic to the Ingard-Myers condition is eliminated in the proposed new formulation. Quantitative assessments on the accuracy of TIMIBC for scattering of acoustic waves by lined surfaces will be carried out, and its effectiveness will be demonstrated by numerical examples. Specifically, the accuracy is assessed by comparing solutions obtained by the Ingard-Myers condition with that by the proposed TIMIBC where theoretical reflection coefficients at a lined surface are derived for cases of plane and spherical incident waves. It is found that the TIMIBC provides a good approximation to the original Ingard-Myers condition for flows of low to mid subsonic Mach numbers. Time domain implementations of TIMIBC are also discussed and illustrated with a numerical example using a finite difference scheme. As many studies have shown the Ingard-Myers condition to be the correct limit of boundary layer thickness going to zero, the proposed TIMIBC can offer a practical solution for overcoming the intrinsic instability associated with the Ingard-Myers condition.

Time Domain BEM↗

NASA NTRSDate not supplied

On A Stabilization of the Ingard-Myers Impedance Boundary Condition

Time Domain BEM↗

NASA NTRS1989

Boundary layer simulator improvement

Boundary Layer Integral Matrix Procedure (BLIMPJ) has been identified by the propulsion community as the rigorous boundary layer program in connection with the existing JANNAF reference programs. The improvements made to BLIMPJ and described herein have potential applications in the design of the future Orbit Transfer Vehicle engines. The turbulence model is validated to include the effects of wall roughness and a way is devised to treat multiple smooth-rough surfaces. A prediction of relaminarization regions is examined as is the combined effects of wall cooling and surface roughness on relaminarization. A turbulence model to represent the effects of constant condensed phase loading is given. A procedure is described for thrust decrement calculation in thick boundary layers by coupling the T-D Kinetics Program and BLIMPJ and a way is provided for thrust loss optimization. Potential experimental studies in rocket nozzles are identified along with the required instrumentation to provide accurate measurements in support of the presented new analytical models.

Praharaj, Sarat C.↗

NASA NTRS1996

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Giel, P. W.↗

NASA NTRS1992

Hypersonic turbulent expansion-corner flow with shock impingement

Mean and fluctuating surface pressure data were obtained in a Mach 8, turbulent, cold flow past an expansion corner subjected to shock impingement. The expansion corner of 2.5 or 4.25 deg was located at 0.77 m (30.25 in.) from the leading edge of a shape-edged flat plate while an external shock, generated by either a 2- or 4-deg sharp wedge, impinged at the corner, or at one boundary layer thickness ahead or behind the corner. The mean pressure distribution was strongly influenced by the mutual interaction between the shock and the expansion. For example, the upstream influence decreased when the shock impinged downstream of the corner. Also, the unsteadiness of the interactions was characterized by an intermittent region and a local rms pressure peak near the upstream influence line. The peak rms pressure fluctuations increased with a larger overall interaction strength. Shock impingement downstream of the corner resulted in lower peaks and also in a shorter region of reduced fluctuation levels. These features may be exploited in inlet design by impinging the cowl shock downstream of an expansion corner instead of at the corner. In addition, a limited Pitot pressure survey showed a thinning of the boundary layer downstream of the corner.

Chung, Kung-Ming↗

NASA NTRS1945

Wind-Tunnel Tests of a 1/6-Scale Model of Republic XF-12 Vertical Tail with Stub Fuselage and Stub Horizontal Tail

A 1/6-scale model of the Republic XF-12 vertical tail with stub fuselage and stub horizontal tail was tested in the Langley stability tunnel to determine the aerodynamic characteristics of the model, The investigation included a study of the effects of boundarylayer thickness, rudder area, and cover-plate alinement on the aerodynamic characteristics. Tuft studies were made in the vicinity of the junction of the vertical and stub horizontal tails. The results of the investigation indicated that the flow in the vicinity of the junction of the vertical and stub horizontal tails was only slightly improved by the addition of a fillet. An increase in boundary-layer thickness produced a slight decrease in rudder effectiveness. The increase in lift of the combined rudders over that of the upper rudder alone was not proportional at low deflections and was approximately proportional at high deflections to the increase in rudder area. When the balance-chamber cover plates were bowed out, the change in rudder hinge moment with rudder angle was less negative. The variation of the lift coefficient with angle of attack and the variation, at small values of angle of attack, of rudder hinge-noment coefficient with angle of attack was approximately the same for all model configurations tested. The upper rudder used in conjunction with a tab was found to satisfy the Army specifications regarding asymmetric power on a multiengine airplane.

MacLachlan, Robert↗

NASA NTRS1985

Foundations for the numerical solution of the Euler equations

The Navier-Stokes equations represent an extremely good model of the physical phenomena encountered in most aeronautical problems. However, the computational resource needed to solve the Navier-Stokes equations are so large that even with today's supercomputers, it is necessary to make use of simpler models. A large number of external aerodynamic problems can be accurately described by a simpler model. This model consists of an outer inviscid flow plus a boundary-layer thickness correction for the vehicle shape. The outer inviscid model may be represented by the potential equation or by the Euler equation. The present paper provides the foundations for the numerical solution of the Euler equations. The governing equations are considered, taking into account conservation laws, the medium, the differential form of the conservation laws, generalized solutions, shock-fitting, and characteristics. Attention is also given to initial and boundary conditions, existence and uniqueness, and rotational phenomena.

Salas, M. D.↗

NASA NTRS2018

Design of Distributed Mail-Slot Propulsion System on a Hybrid Wingbody Aircraft

This paper deals with designing a thrust distribution strategy when a Turboelectric Distributed Propulsion (TeDP) system of 16 embedded propulsors is installed on an aerodynamically optimized hybrid wing-body configuration. This HWB previously designed to satisfy conditions of trim, longitudinally static stability and specific cargo space is employed as the baseline configuration for the current study of seeking an optimal propulsion/power system. According to the nature of the entrance flow condition for each distributed propulsion passage in hybrid wing-body aircraft, the ingested boundary layer thickness differs and results in different propulsive reaction. An optimal distribution of thrust and power output is determined by how the system utilizes the propulsive characteristics of each passage. The design space and the number of design variables are selected and described accordingly. An actuator disk model is employed to model thrust generation and shaft power from the propulsor. To carry out the optimization of the propulsion/power system on a computationally expensive CFD model, a Kriging method in conjunction with a Genetic Algorithm (GA) is applied. Throughout the design process, the propulsion performances of the sampled propulsion/power system are analyzed and compared to those of a clean flow engine. The performance metrics includes mass flow rate, fan pressure ratio besides the thrust and shaft power. Minimization of total shaft power from the distributed engine is performed at multiple thrust levels. The benefit of boundary layer ingestion propulsion system is quantified via comparison of thrust equivalent, shaft power and mass flow equivalent clean flow engines with CFD based system design.

Hybrid Wingbody↗