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At least 271 records · Page 15

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NASA NTRS1987

Stability and transition of three-dimensional flows

The role of secondary instabilities in the transition to turbulence in three-dimensional boundary layers (especially in swept-wing flows) is characterized, reviewing the results of recent theoretical and experimental investigations. Consideration is given to cross-flow vortices, boundary-layer profile measurements, flow visualization, spanwise wavelength determinations, interactions, problems involving rotating disks and cones, and leading-edge contamination. All of the flows investigated are shown to exhibit streamwise vorticity and to depend strongly on initial conditions.

Reed, H. L.↗

NASA NTRS2011

Flow Field Characterization of an Angled Supersonic Jet Near a Bluff Body

An experiment was performed to acquire data from a hot supersonic jet in cross flow for the purpose of validating computational fluid dynamics (CFD) turbulence modeling relevant to the Orion Launch Abort System. Hot jet conditions were at the highest temperature and pressure that could be acquired in the test facility. The nozzle pressure ratio was 28.5, and the nozzle temperature ratio was 3. These conditions are different from those of the flight vehicle, but sufficiently high to model the observed turbulence features. Stereo Particle Image Velocimetry (SPIV) data and capsule pressure data are presented. Features of the flow field are presented and discussed

Wolter, John D.↗

NASA NTRS1973

Numerical calculations of flow fields

Numerical calculations were made of flow fields generated by various aerodynamic configurations. Data cover flow fields generated by a finitely thick lifting three dimensional wing with subsonic tips moving at supersonic speeds, cross flow instability associated with lifting delta wing configurations such as space shuttles, and flow fields produced by a lifting elliptic cone. Finite difference techniques were used to determine elliptic cone flow.

Anderson, D.↗

NASA NTRS1978

Wall region of a relaxing three-dimensional incompressible turbulent boundary layer

A three-dimensional turbulent boundary layer relaxing behind a transverse hump (shaped as a 30-deg swept 5-ft chord wing-type model) was studied in a low-speed wind tunnel. Data obtained with hot-wire probes showed that the apparent dimensionless velocity profiles in the viscous sublayer prevail universally; evidence for wall similarity in the relaxing flow field was confirmed in the form of a log law. An unusual region of slightly decreasing cross-flow angle was found in collateral regions, and a near-wall noncollateral flow was posited. Streamwise relaxation of the mean flow field was also investigated.

Hebbar, K. S.↗

NASA NTRS1985

Experimental and numerical investigation of supersonic flow through a square duct

Steady, developing, adiabatic supersonic flow in a square duct is investigated for an inlet Mach number of 3.91 and a unit Reynolds number of 1.8 x 10 to the 6th/m. The numerical results for laminar flow show that two secondary flow cells develop in the near vicinity of the corner which are centered about the corner bisector and distort the primary flow in this region. For turbulent flow, the experimental results indicate that two secondary flow cells also develop about the corner bisector, but are directed in an opposite sense to that observed for the laminar case. Numerical results based on the Baldwin-Lomax model show that this model is incapable of predicting turbulence-generated secondary flow cells. For a suitable choice of constants, the Gessner-Emery model is able to predict the strength of these cells, but is deficient with respect to predicting their positions in the flow and their distorting influence on the primary flow. These observations are based on comparisons made in this paper between predicted and measured total pressure contours, cross flow velocity profiles, and local wall shear stress distributions.

Davis, D. O.↗

NASA NTRS1973

Hot-wire coil probe for high-speed flows

Small-diameter-wire coil probes developed for use in a hypersonic helium tunnel are discussed. The springlike properties of the coil minimize strain-gauge effects, and allow to use a higher length-to-diameter ratio for a given flow. In addition, the coil is more rugged for sudden flow changes, and since it can be mounted straight across the support tips, there is less support interference in cross flows. In addition to measuring fluctuating quantities in a boundary layer, the probes were used with a constant temperature anemometer for measuring mean mass flow profiles, and with a constant current anemometer for measuring mean total temperature profiles.

Weinstein, L. M.↗

NASA NTRS1996

Experimental and Computational Boundary-Layer Studies in a Supersonic Two-Dimensional Nozzle

Calculations have been carried out on the adiabatic laminar boundary layer developing on the surface of a two-dimensional supersonic nozzle, consisting of contoured nozzle blocks and flat sidewalls. Two- and three-dimensional Navier-Stokes codes, as well as two-dimensional boundary-layer codes have been employed. Thee codes have been adapted to the characteristics of a specific wind tunnel nozzle, so that their numerical results could be directly compared with experimental data obtained in the same nozzle. Such comparisons have been made for the boundary-layer growth on the contoured nozzle, and for the boundary-layer growth, surface streamlines and surface shear on the sidewalls. The three-dimensional Navier-Stokes code was found to be the only one to correctly predict the mean boundary-layer flow on both the sidewalls and the contoured nozzle. Theory and experiment both indicate that the sidewall flow is highly three-dimensional, with non-uniform shear, comer vortices and a boundary layer strongly distorted by cross flows induced by lateral pressure gradients.

Brogan, Torence P.↗

NASA NTRS1998

Experimental and Computational Boundary-Layer Studies In a Supersonic Two-Dimensional Nozzle

Experiments and computations were carried out on the adiabatic laminar boundary layer developing along the surfaces of a two-dimensional supersonic nozzle, consisting of upper and lower contoured nozzle blocks and flat sidewalls. Two- and three-dimensional Navier-Stokes codes, as well as two-dimensional boundary-layer codes were employed. These codes were adapted to the characteristics of a specific wind tunnel nozzle, so that their numerical results could be directly compared with experimental data obtained in the same nozzle. Such comparisons were made for the boundary-layer growth on the nozzle contoured surfaces, and for the boundary-layer growth, surface streamlines and surface shear on the sidewalls. The three-dimensional Navier-Stokes code was found to be the only one to correctly predict the mean laminar boundary-layer flow on both the sidewalls and the contoured surfaces. Theory and experiment both indicated that the sidewall flow is highly three-dimensional, with non-uniform shear, corner vortices and a boundary layer strongly distorted by cross flows induced by lateral pressure gradients.

Demetriades, Anthony↗

NASA NTRS2006

CFD Simulation of the Space Shuttle Launch Vehicle with Booster Separation Motor and Reaction Control System Plumes

The primary objective of this paper is to demonstrate the capability of computational fluid dynamics (CFD) to simulate a very complicated flow field encountered during the space shuttle ascent. The flow field features nozzle plumes from booster separation motor (BSM) and reaction control system (RCS) jets with a supersonic incoming cross flow at speed of Mach 4. The overset Navier-Stokes code OVERFLOW, was used to simulate the flow field surrounding the entire space shuttle launch vehicle (SSLV) with high geometric fidelity. The variable gamma option was chosen due to the high temperature nature of nozzle flows and different plume species. CFD predicted Mach contours are in good agreement with the schlieren photos from wind tunnel test. Flow fields are discussed in detail and the results are used to support the debris analysis for the space shuttle Return To Flight (RTF) task.

Gea, L. M.↗

NASA NTRS2006

CFD Simulation of the Space Shuttle Launch Vehicle with Booster Separation Motor and Reaction Control Plumes

Gea, L. M.↗

NASA NTRS1974

Flutter of elastically supported orthotropic panels including the effects of flow angle

A theoretical panel flutter analysis and computer program was developed which is capable of analyzing an orthotropic panel with in-plane loads, at various angles of cross flow and various edge-support conditions. The resulting modal analysis uses linear piston-theory aerodynamics and includes both aerodynamic and structural damping. Calculations made for typical panels with no in-plane forces show that large reductions in dynamic presure for flutter are possible with only small changes in flow angle. The reduction in dynamic pressure with flow angle is greater for elastically supported panels than for simply supported panels. Aerodynamic damping has a significant stabilizing effect at all flow angles except zero, whereas structural damping has negligible effect.

Sawyer, J. W.↗

NASA NTRS1984

Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation

Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.

Pittman, J. L.↗

NASA NTRS1980

Calculation of three-dimensional unsteady transonic flows past helicopter blades

A finite difference code for predicting the high speed flow over the advancing helicopter rotor is presented. The code solves the low frequency, transonic small disturbance equation and is suitable for modeling the effects of advancing blade unsteadiness on blades of nearly arbitrary planform. The method employs a quasi-conservative mixed differencing scheme and solves the resulting difference equations by an alternating direction scheme. Computed results showed good agreement with experimental blade pressure data and illustrate some of the effects of varying the rotor planform. The flow unsteadiness is shown to be an indispensible part of a transonic solution. Close to the tip at high advance ratio, cross flow effects can significantly affect the solution.

Chattot, J. J.↗

NASA NTRS1981

Aerodynamic heating on the corrugated surface of a 10.2 deg half-angle blunted cone at Mach 6.7

A 10.2 deg half-angle blunted cone with corrugated surfaces was tested in the Langley 8-foot high-temperature structures tunnel to measure the aerodynamic heating of its surfaces. The tests were made in a turbulent boundary layer at angles of attack of 0 deg, 5 deg, and 10 deg. Heating of the windward side was in reasonable agreement with theoretical turbulent predictions for a smooth cone, while heating on the leeward side was between laminar and turbulent predictions as a result of local transitional flow or flow separation produced by high lee-side pressures. Localized heating measurements indicated a significant increase in heating at large cross-flow angles, with the maximum heating rates occurring where the flow reattaches on the upstream side of the corrugation crest and the minimum occurring on the downstream side where the flow is separated.

Weinstein, I.↗

NASA NTRS1987

Experimental studies on Goertler vortices

Goertler vortices arise in laminar boundary layers along concave walls due to an imbalance between pressure and centrifugal forces. In advanced laminar-flow control (LFC) supercritical airfoil designs, boundary-layer suction is primarily used to control Tollmien-Schlichting instability and cross-flow vortices in the concave region near the leading edge of the airfoil lower surface. The concave region itself is comprised of a number of linear segments positioned to limit the total growth of Goertler vortices. Such an approach is based on physical reasonings but rigorous theoretical justification or experimental evidence to support such an approach does not exist. An experimental project was initiated at NASA Langley to verify this concept. In the first phase of the project an experiment was conducted on an airfoil whose concave region has a continuous curvature distribution. Some results of this experiment were previously reported and significant features are summarized.

Mangalam, Siva M.↗

NASA NTRS1965

Microbial profile of laminar flow clean rooms

Microbial profile of laminar flow clean rooms - viable particle concentrations on surfaces of spacecraft and table tops and in air of occupied and empty downflow and cross flow rooms

PARTICLE PRODUCTION↗

NASA NTRS1985

Double-branched vortex generator

In order to assess the suitability of using a double branched vortex generator in parametric studies involving vortex interactions, an experimental study of the main vortex and secondary flows produced by a double branched vortex generator was conducted in a 20-by-40 cm indraft wind tunnel. Measurements of the cross flow velocities were made with a five hole pressure probe from which vorticity contours and vortex parameters were derived. The results showed that the optimum configuration consisted of chord extensions with the absence of a centerbody.

Cantwell, E. R.↗

NASA NTRS2010

PIV Measurements of the CEV Hot Abort Motor Plume for CFD Validation

NASA s next manned launch platform for missions to the moon and Mars are the Orion and Ares systems. Many critical aspects of the launch system performance are being verified using computational fluid dynamics (CFD) predictions. The Orion Launch Abort Vehicle (LAV) consists of a tower mounted tractor rocket tasked with carrying the Crew Module (CM) safely away from the launch vehicle in the event of a catastrophic failure during the vehicle s ascent. Some of the predictions involving the launch abort system flow fields produced conflicting results, which required further investigation through ground test experiments. Ground tests were performed to acquire data from a hot supersonic jet in cross-flow for the purpose of validating CFD turbulence modeling relevant to the Orion Launch Abort Vehicle (LAV). Both 2-component axial plane Particle Image Velocimetry (PIV) and 3-component cross-stream Stereo Particle Image Velocimetry (SPIV) measurements were obtained on a model of an Abort Motor (AM). Actual flight conditions could not be simulated on the ground, so the highest temperature and pressure conditions that could be safely used in the test facility (nozzle pressure ratio 28.5 and a nozzle temperature ratio of 3) were used for the validation tests. These conditions are significantly different from those of the flight vehicle, but were sufficiently high enough to begin addressing turbulence modeling issues that predicated the need for the validation tests.

Wernet, Mark↗