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At least 523 records · Page 29
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

A mathematical formulation of the Scole control problem, part 1

A mathematical formulation of the SCOLE control problem in terms of a continuous model described by partial differential equations with delta functions on the boundary is presented along with three techniques of solution. The abstract wave equation approach leads immediately to a linear feedback law that can ensure (strong) stability. The boundary control approach yields an explicit solution, albeit in a simple case.

Balakrishnan, A. V.↗
NASA NTRS1994

Instabilities orginating from suction holes used for Laminar Flow Control (LFC)

A small-scale wind tunnel previously used for turbulent boundary layer studies has been modified for experiments in laminar flow control. The facility incorporates suction through interchangeable porous test surfaces which are used to stabilize the boundary layer and delay transition to turbulent flow. The thin porous test surfaces are supported by a baffled plenum chamber box which also acts to gather the flow through the surface into tubes which are routed to a high pressure fan. An elliptic leading edge is attached to the assembly to establish a new layer on the test plate. A slot is used to remove the test section flow below the leading edge. The test section was lengthened and fitted with a new ceiling. Substantial modifications were also made to the 3D probe traverse. Detailed studies have been made using isolated holes to explore the underlying instability mechanisms. The suction is perturbed harmonically and data are averaged on the basis of the phase of the disturbance. Conditions corresponding to strong suction and without suction have been studied. In both cases, 3D contour surfaces in the vicinity of the hole show highly three-dimensional T-S waves that fan out away from the hole with streamwise distance. With suction, the perturbations on the centerline are much stronger and decay less rapidly, while the far field is similar to the case without suction. Downstream the contour surfaces of the bow-shaped TS waves develop spanwise irregularities which eventually form into clumps. The contours remain smooth when suction is not applied. Even without suction, the harmonic point source is challenging for CFD; e.g. DNS has been used for streamwise growth. With suction, grid resources are consumed by the hole and this makes DNS even more expensive. Limited DNS results so far indicate that the vortices which emanate from suction holes appear to be stable. The spanwise clumping observed in the experiment is evidence of a secondary instability that could be associated with suction vortices. A typical porous surface for LFC consists of 0.002 inch diameter holes with 0.020 inch grid spacing L, which is too small to resolve disturbances. A 20:1 scale porous test surface has been machined for improved spatial resolution while the L/d is still representative of flight conditions. Designers of porous surfaces use Goldsmith's criterion to minimize crossstream interaction. However nothing is known about the streamwise interactions. Results using two holes, aligned but displaced in the streamwise direction, indicate that partial TS wave cancellation is possible, depending on the hole spacing and disturbance frequency. Using DNS for streamwise interaction studies will be prohibitively expensive if linear superposition cannot be used for the multiple holes.

Watmuff, Jonathan H.↗
NASA NTRS1991

On the interaction between first- and second-mode waves in a supersonic boundary layer

Linear stability theory predicts two or more types of unstable disturbances in a sufficiently high-speed boundary layer. These include the first mode, which is similar to the Tollmien-Schlichting waves found in low-speed flows, and the second mode, which does not depend strongly on the viscosity. Generally, the most unstable first mode is three-dimensional while the most unstable second mode is two-dimensional. The interaction between these two spatially unstable modes are studied by direct solution of the three-dimensional Navier-Stokes equations. It is found that the two-dimensional second mode causes a significant increase in the nonlinearity and in the three-dimensionality of the flowfield. The results suggest that this interaction may accelerate transition for flows where the second mode has a significant growth rate.

Maestrello, L.↗
NASA NTRS1998

Flight Research at Ames: Fifty-Seven Years of Development and Validation of Aeronautical Technology

This NASA special publication presents a general overview of the flight research that has been conducted at Ames Research Center over the last 57 years. Icing research, transonic model testing, aerodynamics, variable stability aircraft, boundary layer control, short takeoff and landing (STOL), vertical/ short takeoff and landing (V/STOL) and rotorcraft research are among the major topics of interest discussed. Flying qualities, stability and control, performance evaluations, gunsight tracking and guidance and control displays research are also presented. An epilogue is included which presents the significant contributions that came about as a result of research and development conducted at Ames.

Borchers, Paul F.↗
NASA NTRS2004

LASTRAC.3d: Transition Prediction in 3D Boundary Layers

Langley Stability and Transition Analysis Code (LASTRAC) is a general-purpose, physics-based transition prediction code released by NASA for laminar flow control studies and transition research. This paper describes the LASTRAC extension to general three-dimensional (3D) boundary layers such as finite swept wings, cones, or bodies at an angle of attack. The stability problem is formulated by using a body-fitted nonorthogonal curvilinear coordinate system constructed on the body surface. The nonorthogonal coordinate system offers a variety of marching paths and spanwise waveforms. In the extreme case of an infinite swept wing boundary layer, marching with a nonorthogonal coordinate produces identical solutions to those obtained with an orthogonal coordinate system using the earlier release of LASTRAC. Several methods to formulate the 3D parabolized stability equations (PSE) are discussed. A surface-marching procedure akin to that for 3D boundary layer equations may be used to solve the 3D parabolized disturbance equations. On the other hand, the local line-marching PSE method, formulated as an easy extension from its 2D counterpart and capable of handling the spanwise mean flow and disturbance variation, offers an alternative. A linear stability theory or parabolized stability equations based N-factor analysis carried out along the streamline direction with a fixed wavelength and downstream-varying spanwise direction constitutes an efficient engineering approach to study instability wave evolution in a 3D boundary layer. The surface-marching PSE method enables a consistent treatment of the disturbance evolution along both streamwise and spanwise directions but requires more stringent initial conditions. Both PSE methods and the traditional LST approach are implemented in the LASTRAC.3d code. Several test cases for tapered or finite swept wings and cones at an angle of attack are discussed.

Chang, Chau-Lyan↗
NASA NTRS1985

Flight-measured laminar boundary-layer transition phenomena including stability theory analysis

Flight experiments were conducted on a single-engine turboprop aircraft fitted with a 92-in-chord, 3-ft-span natural laminar flow glove at glove section lift coefficients from 0.15 to 1.10. The boundary-layer transition measurement methods used included sublimating chemicals and surface hot-film sensors. Transition occurred downstream of the minimum pressure point. Hot-film sensors provided a well-defined indication of laminar, laminar-separation, transitional, and turbulent boundary layers. Theoretical calculations of the boundary-layer parameters provided close agreement between the predicted laminar-separation point and the measured transition location. Tollmien-Schlichting (T-S) wave growth n-factors between 15 and 17 were calculated at the predicted point of laminar separation. These results suggest that for many practical airplane cruise conditions, laminar separation (as opposed to T-S instability) is the major cause of transition in predominantly two-dimensional flows.

Obara, C. J.↗
NASA NTRS1989

Stability and transition of three-dimensional boundary layers

The most recent efforts on the stability and transition of three-dimensional flows are reviewed. These include flows over swept wings, rotating disks, and attachment lines. The generic similarities of their stability behavior is discussed. It is shown that the breakdown process if very complex, often leading to contradictory results. Particular attention is paid to opposing observations of stationary and traveling wave distances.

Saric, William S.↗
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 NTRSDate not supplied

Investigating the Numerical Stability of Using an Impedance Boundary Condition to Model Broadband Noise Scattering With Acoustic Liners

Reducing aircraft noise is a major objective in the field of computational aeroacoustics. When designing next generation quiet aircraft, it is important to be able to accurately and efficiently predict the acoustic scattering by an aircraft body from a given noise source. Acoustic liners are an effective tool for achieving aircraft noise reduction and are characterized by a frequency-dependent impedance value. Converted into the time-domain using Fourier transforms, an impedance boundary condition can be used to simulate the acoustic wave scattering by geometric bodies treated with acoustic liners. A Broadband Impedance Model will be discussed in which the liner impedance is specified along a wide range of frequencies. The liner impedance boundary condition will be derived and coupled with a time-domain boundary integral equation to model acoustic scattering by a flat plate consisting of both un-lined and lined surfaces. The scattering solution will be obtained iteratively using both spatial and temporal basis functions and the stability will be demonstrated through eigenvalue analysis. Stability will be assessed for its dependence on time step, spatial discretization, as well as temporal basis function order. Both second- and third-order Lagrange temporal basis functions are considered.

acoustics↗
NASA NTRSDate not supplied

Investigating the Numerical Stability of Using an Impedance Boundary Condition to Model Broadband Noise Scattering With Acoustic Liners

Reducing aircraft noise is a major objective in the field of computational aeroacoustics. When designing next generation quiet aircraft, it is important to be able to accurately and efficiently predict the acoustic scattering by an aircraft body from a given noise source. Acoustic liners are an effective tool for achieving aircraft noise reduction and are characterized by a frequency-dependent impedance value. Converted into the time-domain using Fourier transforms, an impedance boundary condition can be used to simulate the acoustic wave scattering by geometric bodies treated with acoustic liners. A Broadband Impedance Model will be discussed in which the liner impedance is specified along a wide range of frequencies. The liner impedance boundary condition will be derived and coupled with a time-domain boundary integral equation to model acoustic scattering by a flat plate consisting of both un-lined and lined surfaces. The scattering solution will be obtained iteratively using both spatial and temporal basis functions and the stability will be demonstrated through eigenvalue analysis. Stability will be assessed for its dependence on time step, spatial discretization, as well as temporal basis function order. Both second- and third-order Lagrange temporal basis functions are considered.

acoustics↗
NASA NTRS1996

Transition Studies on a Swept-Wing Model

The present investigation contributes to the understanding of boundary-layer stability and transition by providing detailed measurements of carefully-produced stationary crossflow vortices. It is clear that a successful prediction of transition in swept-wing flows must include an understanding of the detailed physics involved. Receptivity and nonlinear effects must not be ignored. Linear stability theory correctly predicts the expected wavelengths and mode shapes for stationary crossflow, but fails to predict the growth rates, even for low amplitudes. As new computational and analytical methods are developed to deal with three-dimensional boundary layers, the data provided by this experiment will serve as a useful benchmark for comparison.

Saric, William S.↗
NASA NTRS1990

Advanced Mach 3.5 Axisymmetric Quiet Nozzle

To advance boundary-layer stability and transition research and to ultimately provide reliable predictions of transition for supersonic flight vehicles, a wind tunnel is required with very low stream disturbance levels comparable to free flight conditions. Experimental and theoretical research to develop a low-disturbance supersonic wind tunnel has achieved a breakthrough. A new concept for nozzle design is presented which promises a large increase in the length of the quiet test core. The Advanced Mach 3.5 Axisymmetric Quiet Nozzle is the first prototype built to prove the new design concept. Experimental results from this new nozzle on the extent of laminar wall boundary layers are compared with data from other nozzles and with theoretical predictions based on linear stability theory. The Reynolds numbers based on the measured length of the quiet test core for this new nozzle are in excellent agreement with the theoretical predictions. The effect of surface finish on the nozzle performance is also discussed.

Chen, Fang-Jenq↗
NASA NTRS1984

Effects of nozzle design parameters on the extent of quiet test flow at Mach 3.5

Tests results at the NASA Langley Research Center, involving a Mach 3.5 pilot quiet tunnel, have shown that laminar-layered nozzle walls improve boundary layer stability and reduce stream disturbance levels caused by eddy Mach wave radiation. This type of wall design is required to obtain transition Reynolds numbers on tests models as high as those previously observed in supersonic flight vehicles. The Mach 3.5 pilot nozzle wall boundary layers were tested for Tollmein-Schlichting and Goertler linear amplification, and, in an analysis of Goertler vortices in two axisymmetric Mach 5 nozzles, transition values were found to vary. These values were applied to several nozzles with similar throat heights but different expansion rates. Among the nozzles included in the study, a flat-wall radial flow nozzle and a proposed rod-wall nozzle were tested. For the highest test unit Reynolds number, it was determined that the nozzle wall surface finish should not exceed 0.3 micron. Oil flow studies have indicated that Goertler vortex disturbances were the dominant mechanism causing transition on the walls of the pilot nozzle.

Beckwith, I. E.↗
NASA NTRS1955

An Investigation of the Effects of Heat Transfer on Boundary-Layer Transition on a Parabolic Body of Revolution (NACA RM-10) at a Mach Number of 1.61

Report presents the results of an investigation conducted to determine the effects of heat transfer on boundary-layer transition on a parabolic body of revolution (NACA rm-10 without fins) at Mach number of 1.61 and over a Reynolds number range from 2.5 x 10(6) to 35 x 10(6). The maximum cooling of the model used in these tests corresponded to a temperature ratio (ratio of model-surface temperature to free-stream temperature) of 1.12, a value somewhat higher than the theoretical value required for infinite boundary-layer stability at this Mach number. The maximum heating corresponded to a temperature ratio of about 1.85. Included in the investigation was a study of the effects of surface irregularities and disturbances generated in the airstream on the ability of heat transfer to influence boundary-layer transition.

Czarnecki, K R↗
NASA NTRS1994

Supersonic laminar flow control research

The objective of the research is to understand supersonic laminar flow stability, transition and active control. Some prediction techniques will be developed or modified to analyze laminar stability. The effects of supersonic laminar flow with distributed heating and cooling on active control will be studied. The primary tasks of the research applying to the NASA/Ames POC and LFSWT's nozzle design with laminar flow control are as follows: (1) supersonic laminar boundary layer stability and transition prediction; (2) effects of heating and cooling for supersonic laminar flow control; and (3) POC and LFSWT nozzle design with heating and cooling effects combining wall contour and length changes.

Lo, C. F.↗
NASA NTRS1987

Stability of three-dimensional boundary layers

The computational modeling of the transition process characteristic of flows over swept wings is discussed. Specifically, the crossflow instability and crossflow/Tollmien-Schlichting wave interactions are analyzed through the numerical solution of the full three-dimensional Navier-Stokes equations including unsteadiness, curvature, and sweep. This approach is chosen because of the complexity of the problem and because it appears that regular stability theory is insufficient to explain the discrepancies between experiments and between theory and experiment. The leading edge region of a swept wing will be considered in a three-dimensional spatial simulation with random disturbances as the initial conditions.

Reed, Helen L.↗
NASA NTRS1991

A numerical investigation of supersonic nozzle boundary layer transition

The boundary layer stability within the high-area-ratio nozzle at NASA Lewis was tested at five chamber pressure conditions. For all nozzle conditions examined, Taylor-Goertler vortices grew more rapidly than Tollmien-Schlichting waves. The experimental heat flux was accurately predicted by a laminar boundary layer computation, thereby confirming the laminar nature of the nozzle boundary layer flow. When the chamber pressure was increased in a series of cases, the transition point occurred farther upstream. The number of vortices contained in the dominant instability increased with chamber pressure.

Pauley, Laura L.↗