Search NASA⌕ Search

SEARCH · Search NASA

Results for “boundary layer analysis”

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.

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

At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

WMLES of Boundary-Layer Transition

An analysis of the potential for the accurate prediction of laminar-turbulent transition by Wall-Modeled Large-Eddy Simulation (WMLES) indicates that satisfactory prediction of transition by WMLES is possible, provided a) a sufficiently accurate wall model for the initial stage of transition is employed, b) a satisfactory transition is made between this initial-stage wall model and the wall model for the downstream fully-developed turbulence, and c) limitations on temporal and spatial resolution, and on the thickness of the wall-modeled region, imposed by the need to accurately resolve the disturbance waves and the transition region are observed. This is illustrated with WMLES computations of subharmonic transition in a flat-plate boundary layer that yield reasonable agreement with DNS results. However, the wall-modeled region thickness, and thus the cost savings, is less than is typical for WMLES of fully developed turbulence.

Wall-modeled large-eddy simulation↗
NASA NTRS1976

A complete model of turbulence

A set of constitutive equations suitable for a priori computation of turbulent shear flows has been developed. Since no properties of a given turbulent flow need be known in advance in order to obtain a solution, the equations comprise a complete model of turbulence. Perturbation analysis shows that the model predicts a composite five-layer structure for an incompressible turbulent boundary layer, viz, a defect layer, a law-of-the-wall layer, a viscous sublayer, a near-surface roughness layer, and a viscous superlayer at the boundary-layer edge. Analysis of the defect layer demonstrates the key improvement of the model over its predecessor, the Saffman-Wilcox two-equation model of turbulence. Examination of model-predicted sublayer structure yields model-parameter boundary conditions appropriate for surfaces with roughness and mass injection. Results of numerical computations of compressible and incompressible equilibrium boundary layers show that, for such flows, the model is as accurate as mixing-length theory. Applications to transitional boundary layers and to nonequilibrium relaxation of a boundary layer passing from a rough to a smooth surface indicate that the model's applicability extends far beyond that of mixing-length theory's.

Wilcox, D. C.↗
NASA NTRS1981

Efficient computation of the stability of three-dimensional compressible boundary layers

Methods for the computer analysis of the stability of three-dimensional compressible boundary layers are discussed and the user-oriented Compressible Stability Analysis (COSAL) computer code is described. The COSAL code uses a matrix finite-difference method for local eigenvalue solution when a good guess for the eigenvalue is available and is significantly more computationally efficient than the commonly used initial-value approach. The local eigenvalue search procedure also results in eigenfunctions and, at little extra work, group velocities. A globally convergent eigenvalue procedure is also developed which may be used when no guess for the eigenvalue is available. The global problem is formulated in such a way that no unstable spurious modes appear so that the method is suitable for use in a black-box stability code. Sample stability calculations are presented for the boundary layer profiles of an LFC swept wing.

Malik, M. R.↗
NASA NTRS2014

Forced Response Analysis of a Fan with Boundary Layer Inlet Distortion

Boundary layer ingesting propulsion systems have the potential to significantly reduce fuel burn for future generations of commercial aircraft, but these systems must be designed to overcome the challenge of high dynamic stresses in fan blades due to forced response. High dynamic stresses can lead to high cycle fatigue failures. High-fidelity computational analysis of the fan aeromechanics is integral to an ongoing effort to design a boundary layer ingesting inlet and fan for a wind-tunnel test. An unsteady flow solution from a Reynoldsaveraged Navier Stokes analysis of a coupled inlet-fan system is used to calculate blade unsteady loading and assess forced response of the fan to distorted inflow. Conducted prior to the mechanical design of a fan, the initial forced response analyses performed in this study provide an early look at the levels of dynamic stresses that are likely to be encountered. For the boundary layer ingesting inlet, the distortion contains strong engine order excitations that act simultaneously. The combined effect of these harmonics was considered in the calculation of the forced response stresses. Together, static and dynamic stresses can provide the information necessary to evaluate whether the blades are likely to fail due to high cycle fatigue. Based on the analyses done, the overspeed condition is likely to result in the smallest stress margin in terms of the mean and alternating stresses. Additional work is ongoing to expand the analyses to off-design conditions, on-resonance conditions, and to include more detailed modeling of the blade structure.

Vibrational stress↗
NASA NTRS1951

Analysis of the effects of boundary-layer control in the take-off and power-off landing performance characteristics of a liaison type of airplane

A performance analysis has been made to determine whether boundary-layer control by suction might reduce the minimum take-off and landing distances of a four-place or five-place airplane or a liaison type of airplane below those obtainable with conventional high-lift devices. The airplane was assumed to have a cruise duration of 5 hours at 60-percent power and to be operating from airstrips having a ground friction coefficient of 0.2 or a combined ground and braking coefficient of 0.4. The payload was fixed at 1500 pounds, the wing span was varied from 25 to 100 feet, the aspect ratio was varied from 5 to 15, and the power was varied from 300 to 1300 horsepower. Maximum lift coefficients of 5.0 and 2.8 were assumed for the airplanes with and without boundary-layer-control --equipment weight was included. The effects of the boundary-layer control on total take-off distance, total power-off landing distance, landing and take-off ground run, stalling speed, sinking speed, and gliding speed were determined.

Horton, Elmer A↗
NASA NTRS1989

The role of thermal and lubricant boundary layers in the transient thermal analysis of spur gears

An improved convection heat-transfer model has been developed for the prediction of the transient tooth surface temperature of spur gears. The dissipative quality of the lubricating fluid is shown to be limited to the capacity extent of the thermal boundary layer. This phenomenon can be of significance in the determination of the thermal limit of gears accelerating to the point where gear scoring occurs. Steady-state temperature prediction is improved considerably through the use of a variable integration time step that substantially reduces computer time. Computer-generated plots of temperature contours enable the user to animate the propagation of the thermal wave as the gears come into and out of contact, thus contributing to better understanding of this complex problem. This model has a much better capability at predicting gear-tooth temperatures than previous models.

El-Bayoumy, L. E.↗
NASA NTRS1989

The role of thermal and lubricant boundary layers in the transient thermal analysis of spur gears

An improved convection heat-transfer model has been developed for the prediction of the transient tooth surface temperature of spur gears. The dissipative quality of the lubricating fluid is shown to be limited to the capacity extent of the thermal boundary layer. This phenomenon can be of significance in the determination of the thermal limit of gears accelerating to the point where gear scoring occurs. Steady-state temperature prediction is improved considerably through the use of a variable integration time step that substantially reduces computer time. Computer-generated plots of temperature contours enable the user to animate the propagation of the thermal wave as the gears come into and out of contact, thus contributing to better understanding of this complex problem. This model has a much better capability at predicting gear-tooth temperatures than previous models.

El-Bayoumy, L. E.↗
NASA NTRS1980

Viscous flow drag reduction; Symposium, Dallas, Tex., November 7, 8, 1979, Technical Papers

The symposium focused on laminar boundary layers, boundary layer stability analysis of a natural laminar flow glove on the F-111 TACT aircraft, drag reduction of an oscillating flat plate with an interface film, electromagnetic precipitation and ducting of particles in turbulent boundary layers, large eddy breakup scheme for turbulent viscous drag reduction, blowing and suction, polymer additives, and compliant surfaces. Topics included influence of environment in laminar boundary layer control, generation rate of turbulent patches in the laminar boundary layer of a submersible, drag reduction of small amplitude rigid surface waves, and hydrodynamic drag and surface deformations generated by liquid flows over flexible surfaces.

Hough, G. R.↗
NASA NTRS1971

Some characteristics of turbulent boundary layers in rapidly accelerated flows

An analysis of time-mean-turbulent boundary layer velocity profiles measured in a rapidly accelerating flow suggests that the outer region of the velocity profiles consists of essentially inviscid, rotational flow. The extent of this inviscid outer region was observed in some cases to exceed 90 percent of what is ordinarily thought of as the turbulent boundary layer thickness. On the other hand, the inner frictional region of these velocity profiles appears to have turbulent characteristics similar to those of more conventional turbulent boundary layers. Hence, the outer edge boundary condition for this inner region is more properly the external rotational flow region than the free stream.

Brinich, P. F.↗
NASA NTRS1983

Turbulence and surface heat transfer near the stagnation point of a circular cylinder in turbulent flow

A turbulent boundary layer flow analysis of the momentum and thermal flow fields near the forward stagnation point due to a circular cylinder in turbulent cross flow is presented. Turbulence modeling length scale, anisotropic turbulence initial profiles and boundary conditions were identified as functions of the cross flow turbulence intensity and the boundary layer flow far field velocity. These parameters were used in a numerical computational procedure to calculate the mean velocity, mean temperature, and turbulence double correlation profiles within the flow field. The effects of the cross flow turbulence on the stagnation region momentum and thermal flow fields were investigated. This analysis predicted the existing measurements of the stagnation region mean velocity and surface heat transfer rate with cross flow Reynolds number and turbulence intensity less than 250,000 and 0.05, respectively.

Wang, C. R.↗
NASA NTRS1995

Analysis of in-flight boundary-layer state measurements on a subsonic transport wing in high-lift configuration

Flight experiments on NASA Langley's B737-100 (TSRV) airplane have been conducted to document flow characteristics in order to further the understanding of high-lift flow physics, and to correlate and validate computational predictions and wind-tunnel measurements. The project is a cooperative effort involving NASA, industry, and universities. In addition to focusing on in-flight measurements, the project includes extensive application of various computational techniques, and correlation of flight data with computational results and wind-tunnel measurements. Results obtained in the most recent phase of flight experiments are analyzed and presented in this paper. In-flight measurements include surface pressure distributions, measured using flush pressure taps and pressure belts on the slats, main element, and flap elements; surface shear stresses, measured using Preston tubes; off-surface velocity distributions, measured using shear-layer rakes; aeroelastic deformations of the flap elements, measured using an optical positioning system; and boundary-layer transition phenomena, measured using hot-film anemometers and an infrared imaging system. The analysis in this paper primarily focuses on changes in the boundary-layer state that occurred on the slats, main element, and fore flap as a result of changes in flap setting and/or flight condition. Following a detailed description of the experiment, the boundary-layer state phenomenon will be discussed based on data measured during these recent flight experiments.

vanDam, C. P.↗
NASA NTRS1978

Derivation and evaluation of an approximate analysis for three-dimensional viscous subsonic flow with large secondary velocities

An approximate analysis is presented for calculating three-dimensional, low Mach number, laminar viscous flows in curved passages with large secondary flows and corner boundary layers. The analysis is based on the decomposition of the overall velocity field into inviscid and viscous components with the overall velocity being determined from superposition. An incompressible vorticity transport equation is used to estimate inviscid secondary flow velocities to be used as corrections to the potential flow velocity field. A parabolized streamwise momentum equation coupled to an adiabatic energy equation and global continuity equation is used to obtain an approximate viscous correction to the pressure and longitudinal velocity fields. A collateral flow assumption is invoked to estimate the viscous correction to the transverse velocity fields. The approximate analysis is solved numerically using an implicit ADI solution for the viscous pressure and velocity fields. An iterative ADI procedure is used to solve for the inviscid secondary vorticity and velocity fields. This method was applied to computing the flow within a turbine vane passage with inlet flow conditions of M = 0.1 and M = 0.25, Re = 1000 and adiabatic walls, and for a constant radius curved rectangular duct with R/D = 12 and 14 and with inlet flow conditions of M = 0.1, Re = 1000, and adiabatic walls.

Anderson, O. L.↗
NASA NTRS1950

An analysis of supersonic aerodynamic heating with continuous fluid injection

An analysis of the steady-state aerodynamic heating problem at high-supersonic speeds is made for two-dimensional flows with laminar boundary layers. The aerodynamic heating is shown to be reduced substantially by injecting a small amount of coolant through a porous surface into the boundary layer. The analysis includes calculations of the cooling requirements and equilibrium surface temperatures for flat plates and for an equilibrium surface temperatures for flat plates and for flat porous surfaces with several rates of fluid injection at Mach numbers from 5 to 15 and altitudes from sea level to 200,000 feet. Some calculations of the skin friction are also included.

Klunker, E G↗
NASA NTRS2013

Geometric Effects on the Amplification of First Mode Instability Waves

The effects of geometric changes on the amplification of first mode instability waves in an external supersonic boundary layer were investigated using numerical techniques. Boundary layer stability was analyzed at Mach 6 conditions similar to freestream conditions obtained in quiet ground test facilities so that results obtained in this study may be applied to future test article design to measure first mode instability waves. The DAKOTA optimization software package was used to optimize an axisymmetric geometry to maximize the amplification of the waves at first mode frequencies as computed by the 2D STABL hypersonic boundary layer stability analysis tool. First, geometric parameters such as nose radius, cone half angle, vehicle length, and surface curvature were examined separately to determine the individual effects on the first mode amplification. Finally, all geometric parameters were allowed to vary to produce a shape optimized to maximize the amplification of first mode instability waves while minimizing the amplification of second mode instability waves. Since first mode waves are known to be most unstable in the form of oblique wave, the geometries were optimized using a broad range of wave frequencies as well as a wide range of oblique wave angles to determine the geometry that most amplifies the first mode waves. Since first mode waves are seen most often in flows with low Mach numbers at the edge of the boundary layer, the edge Mach number for each geometry was recorded to determine any relationship between edge Mach number and the stability of first mode waves. Results indicate that an axisymmetric cone with a sharp nose and a slight flare at the aft end under the Mach 6 freestream conditions used here will lower the Mach number at the edge of the boundary layer to less than 4, and the corresponding stability analysis showed maximum first mode N factors of 3.

Kirk, Lindsay C.↗
NASA NTRS1991

Assessment of meanflow solutions for instability analysis of transitioning flows

Meanflow solutions of 3D supersonic flow past a cone at incidence and a swept leading edge wing have been obtained by thre methods, viz., boundary-layer, parabolized Navier-Stokes, and thin shear-layer Navier-Stokes solvers. The smoothness and accuracy of the solution profiles are compared with a view to applying the meanflow solution to boundary-layer stability analysis.

Iyer, Venkit↗
NASA NTRS1990

Variable Sweep Transition Flight Experiment (VSTFE): Unified Stability System (USS). Description and Users' Manual

NASA initiated the Variable Sweep Transition Flight Experiment (VSTFE) to establish a boundary layer transition database for laminar flow wing design. For this experiment, full-span upper surface gloves were fitted to a variable sweep F-14 aircraft. The development of an improved laminar boundary layer stability analysis system called the Unified Stability System (USS) is documented and results of its use on the VSTFE flight data are shown. The USS consists of eight computer codes. The theoretical background of the system is described, as is the input, output, and usage hints. The USS is capable of analyzing boundary layer stability over a wide range of disturbance frequencies and orientations, making it possible to use different philosophies in calculating the growth of disturbances on sweptwings.

Rozendaal, Rodger A.↗
NASA NTRS2009

Transition Analysis for the HIFiRE-5 Vehicle

The Hypersonic International Flight Research and Experimentation (HIFiRE) 5 flight experiment by Air Force Research Laboratories and Australian Defense Science and Technology Organization is designed to provide in-flight boundary-layer transition data for a canonical 3D configuration at hypersonic Mach numbers. This paper outlines the progress, to date, on boundary layer stability analysis for the HIFiRE-5 flight configuration, as well as for selected test conditions from the wind tunnel experiments supporting the flight test. At flow conditions corresponding to the end of the test window, rather large values of linear amplification factor are predicted for both second mode (N>40) and crossflow (N>20) instabilities, strongly supporting the feasibility of first in-flight measurements of natural transition on a fully three-dimensional hypersonic configuration. Additional results highlight the rich mixture of instability mechanisms relevant to a large segment of the flight trajectory, as well as the effects of angle of attack and yaw angle on the predicted transition fronts for ground facility experiments at Mach 6.

Choudhari, Meelan M.↗
NASA NTRS1980

Prediction method for two-dimensional aerodynamic losses of cooled vanes using integral boundary-layer parameters

A generalized analysis to predict the two-dimensional aerodynamic losses of film-cooled vanes by using integral boundary-layer parameters is presented. Heat-transfer and trailing-edge injection effects are included in the method. An approximate solution of the generalized equations is also included to show more clearly the effect of the different boundary-layer and cooling parameters on the losses. The analytical predictions agree well with the experimental results, indicating that available boundary-layer calculations for cooled vanes are of sufficient accuracy to use in the prediction method.

Goldman, L. J.↗