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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.

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At least 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Predicted and measured in-flight wing deformations of a forward-swept-wing aircraft

An electrooptical flight-deflection measurement system (FDMS) is described in terms of its use in structural testing of the composite forward-swept wing of the X-29 aircraft. The wing deflected shapes measured using the present system are compared to the shapes predicted by NASTRAN and other codes as well as data from ground-test load measurements. The electrooptical FDMS is based on a control unit, two receivers, a target driver, and 12-16 IR LED targets. The FDMS determines the in-flight deflected wing shapes at a variety of altitudes at Mach 0.9, and the results are compared to the analytically predicted wing-twist distributions. The FDMS data describe the predicted increasing streamwise twist with increasing dynamic pressure and suggest that the streamwise twist is more prevalent at the inboard measurement station than at the wing tip. This hook shape is not represented in the predicted data, and suggestions are given for improving the modeling of the X-29 wing.

Lokos, William A.↗
NASA NTRS1992

Computational study of transition front on a swept wing leading-edge model

The e(N) method is employed with Navier-Stokes calculations to study 3D supersonic boundary layers at the leading-edge region of swept wings with attention given to the conditions for linear stability. Mean-flow profiles are computed as solutions to the Navier-Stokes equations, and the e(N) method is used to identify the onset of the transition. Specific treatment is given to the identification of parameters which affect the transition such as free-stream Reynolds number, wall cooling, and boundary-layer suction levels. The boundary layer can be stabilized to below N=10 levels in all locations by utilizing distributed suction, and enhanced stabilization can also be achieved by employing wall cooling and reduced free-stream Reynolds numbers. This study presents key calculations for corresponding experimental investigations by shedding light on requirements for test conditions and measurement requirements.

Iyer, Venkit↗
NASA NTRS1992

Forebody vortex control for suppressing wing rock on a highly-swept wing configuration

Free-to-roll tests were conducted in a wind tunnel with a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. A limit cycle oscillation was observed for angles of attack between 22 and 30 deg. In general, the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. Various blowing techniques were evaluated as means of wing rock suppression. Blowing tangentially aft from leeward side nozzles near the forebody tip can damp the roll motion at low blowing rates and stop it completely at higher blowing rates. At the high rates, significant vortex asymmetries are created, causing the model to stop at a non-zero roll angle. Forward blowing and alternating right/left pulsed blowing appear to be more efficient techniques for suppressing wing rock. The oscillations can be damped almost completely at lower blowing coefficients, and, apparently, no major vortex asymmetries are induced. Good agreement is observed between this study and previous water tunnel tests on the same configuration.

Suarez, Carlos J.↗
NASA NTRS1993

Results of low power deicer tests on a swept inlet component in the NASA Lewis icing research tunnel

Tests were conducted under a USAF/NASA Low Power Deicer program on two expulsive technologies to examine system performance on hardware representative of a modern aircraft part. The BF Goodrich Electro-Expulsive Deicing System and Pneumatic Impulse Ice Protection System were installed on a swept, compound curve, engine inlet component with varying leading edge radius, and tested through a range of icing and system operating conditions in the NASA Lewis Icing Research Tunnel. A description of the experimental procedure and results, including residual ice thickness, shed ice particle size, and changes in system energy/pressure characteristics are presented.

Bond, Thomas H.↗
NASA NTRS1992

Crossflow Stability and Transition Experiments in a Swept-Wing Flow

An experimental examination of crossflow instability and transition on a 45 degree swept wing is conducted in the Arizona State University Unsteady Wind Tunnel. The stationary-vortex pattern and transition location are visualized using both sublimating-chemical and liquid-crystal coatings. Extensive hot-wire measurements are conducted at several measurement stations across a single vortex track. The mean and travelling-wave disturbances are measured simultaneously. Stationary-crossflow disturbance profiles are determined by subtracting either a reference or a span-averaged velocity profile from the mean-velocity data. Mean, stationary-crossflow, and travelling-wave velocity data are presented as local boundary-layer profiles and as contour plots across a single stationary-crossflow vortex track. Disturbance-mode profiles and growth rates are determined. The experimental data are compared to predictions from linear stability theory.

Dagenhart, John Ray↗
NASA NTRS1993

Effect of micron-sized roughness on transition in swept-wing flows

Boundary-layer transition-to-turbulence studies are conducted in the Arizona State University Unsteady Wind Tunnel on a 45-degree swept airfoil. The pressure gradient is designed so that the initial stability characteristics are purely crossflow-dominated. Flow visualization and hot-wire measurements show that the development of the crossflow vortices is influenced by roughness near the attachment-line. Comparisons of transition location are made between a painted surface, a machine-polished surface, and a hand-polished surface. Then, isolated 6 micron roughness elements are placed near the attachment line on the airfoil surface under conditions of the final polish (0.25 micron rms). These elements amplify a centered stationary crossflow vortex and its neighbors, resulting in localized early transition. The diameter, height, and location of these roughness elements are varied in a systematic manner. Spanwise hot-wire measurements are taken behind the roughness element to document the enhanced vortices. These scans are made at several different chord locations to examine vortex growth.

Radeztsky, Ronald H., Jr.↗
NASA NTRS1993

Stability and transition on swept wings

This paper describes the extension and application of the Parabolized Stability Equations (PSE) to the stability and transition of the supersonic three-dimensional laminar boundary layer on a swept wing. The problem formulation uses a general coordinate transformation for arbitrary curvilinear body-fitted computational grids. Some testing using these coordinates is briefly described to help validate the software used for the investigation. The disturbance amplitude ratios as a function of chord position for supersonic (Mach 1.5) boundary layers on untapered, untwisted wings of different sweep angles are then presented and compared with those obtained from local parallel analyses.

Stuckert, Greg↗
NASA NTRS1992

Development of a structural optimization capability for the aeroelastic tailoring of composite rotor blades with straight and swept tips

This paper describes the development of a new structural optimization capability aimed at the aeroelastic tailoring of composite rotor blades with straight and swept tips. The primary objective is to reduce vibration levels in forward flight without diminishing the aeroelastic stability margins of the blade. In the course of this research activity a number of complicated tasks have been addressed: (1) development of a new, aeroelastic stability and response analysis; (2) formulation of a new comprehensive sensitive analysis, which facilitates the generation of the appropriate approximations for the objective and the constraints; (3) physical understanding of the new model and, in particular, determination of its potential for aeroelastic tailoring, and (4) combination of the newly developed analysis capability, the sensitivity derivatives and the optimizer into a comprehensive optimization capability. The first three tasks have been completed and the fourth task is in progress.

Friedmann, P. P.↗
NASA NTRS1992

A study of hypersonic swept shock wave/turbulent boundary layer interactions using a conical Navier-Stokes code

A computational study has been performed of sharp fin-induced swept shock wave/turbulent boundary layer interactions at low hypersonic Mach numbers. The objective was to determine if results obtained using a conical Navier-Stokes code, particularly the peak heating and pressure, are adequate for engineering predictions. The advantage of the conical approach is that the problem becomes two-dimensional and requires much less computational effort than a fully three-dimensional calculation. In this code, the standard Baldwin-Lomax model is used for turbulent closure and its performance is studied in some detail. To assess the approach interactive flowfields generated by unswept sharp fins at two angles of attack at each of three hyprsonic freestream Mach numbers (5, 6, 11) have been calculated and the results compared with experimental wall pressure and heat transfer data. Although the conical Navier-Stokes/Baldwin-Lomax approach is reasonably successful at Mach numbers up to 5, the performance deteriorates as the Mach number is increased. Nevertheless, the approach could be a valuable tool in preliminary parametric design studies.

Rodi, Patrick E.↗
NASA NTRS1993

Results of Low Power Deicer tests on a swept inlet component in the NASA Lewis Icing Research Tunnel

Tests were conducted under a USAF/NASA Low Power Deicer program on two expulsive technologies to examine system performance on hardware representative of a modern aircraft part. The BF Goodrich Electro-Expulsive Deicing System and Pneumatic Impulse Ice Protection system were installed on a swept, compound curve, engine inlet component with varying leading edge radius, and tested through a range of icing and system operating conditions in the NASA Lewis Icing Research Tunnel. A description of the experimental procedure and results, including residual ice thickness, shed ice particle size, and changes in system energy/pressure characteristics are presented.

Bond, Thomas H.↗
NASA NTRS1993

Wall pressure fluctuations beneath swept shock wave/boundary layer interactions

An experimental research program providing basic knowledge and establishing a database on the fluctuating pressure loads produced on aerodynamic surfaces beneath 3D shock wave/boundary layer interactions is presented. A turbulent boundary layer on a flat plate is subjected to interactions with swept planar shock waves generated by sharp fins at angle of attack. Measurements are made for the first time in the aft areas of these interactions, showing fluctuating pressure levels as high as 160 dB.

Garg, S.↗
NASA NTRS1992

Temperature and suction effects on the instability of an infinite swept attachment line

It is known that the incompressible, infinite swept attachment line flow is unstable to streamwise disturbances that originate in the boundary layer when the cross-flow exceeds a critical magnitude. Furthermore, a small degree of suction at the surface has a significant stabilizing influence while a small degree of blowing has a considerable destabilizing influence. This paper investigates the stabilizing and destabilizing effects of, respectively, cooling or heating the plate and the competing or enhancing effects of suction or blowing. A nonorthogonal flow with respect to the attachment line is also considered by adding a component of shear to the mean flow.

Lasseigne, D. G.↗
NASA NTRS1993

An experimental study of the effects of bodyside compression on forward swept sidewall compression inlets ingesting a turbulent boundary layer

Forward swept sidewall compression inlets have been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to study the effects of bodyside compression surfaces on inlet performance in the presence of an incoming turbulent boundary layer. The measurements include mass flow capture and mean surface pressure distributions obtained during simulated combustion pressure increases downstream of the inlet. The kerosene-lampblack surface tracer technique has been used to obtain patterns of the local wall shear stress direction. Inlet performance is evaluated using starting and unstarting characteristics, mass capture, mean surface pressure distributions and permissible back pressure limits. The results indicate that inlet performance can be improved with selected bodyside compression surfaces placed between the inlet sidewalls.

Rodi, Patrick E.↗
NASA NTRS1993

Experimental and numerical study of swept ramp injection into a supersonic flowfield

Time-averaged measurements of pressure, temperature, velocity, and injectant mole fraction are presented using the planar laser-induced iodine fluorescence (PLIIF) techniques in the complex 3D compressible flowfield around a swept ramp fuel injector. Within the range of thermodynamic conditions present in the test case studied, the technique's accuracy is estimated to be 4 percent for pressure, 4 percent for temperature, 4 percent for velocity, and 3 percent for injectant mole fraction. Comparisons to numerical simulations using the SPARK 3D Navier-Stokes computer code with an algebraic turbulence model are made at the centerplane of the flowfield as well as on three crossflow planes downstream of the injector, and good agreement is found. A weak asymmetry in the incoming flowfield appears to be amplified by boundary layer separation occuring when the the ramp generated shock reflects off the tunnel walls. In the near field of the injector, laminar calculations show better agreement to the measurements than turbulent calculations.

Donohue, James M.↗
NASA NTRS1951

Stability and Control Characteristics of a Complete Airplane Model Having a Wing with Quarter-chord Line Swept Back 40 Degrees, Aspect Ratio 2.50, and Taper Ratio 0.42

An investigation has been made of a complete airplane model having a wing with the quarter-chord line swept back 40 degrees, aspect ratio 2.50, and taper ratio 0.42 to determine its low-speed stability and control characteristics. The longitudinal stability investigation included stabilizer and tail-off tests with different wing dihedral angles (Gamma = 0 degrees and Gamma = -10 degrees) over an angle-of-attack range for the cruising and landing configurations and tests. with a high horizontal-tail location (Gamma = -10 degrees) for the cruising configuration. Tests were made of the wing alone and to determine the effect of wing end plates in pitch. Lateral stability characteristics were determined for the airplane with different geometric wing dihedrals, with end plates, and with several dorsal modifications. Tests were made with ailerons and spoilers to determine control characteristics.

STABILITY, STATIC↗
NASA NTRS1951

Effect of Taper Ratio on the Low-speed Rolling Stability Derivatives of Swept and Unswept Wings of Aspect Ratio 2.61

Results of tests conducted in the 6-foot-diameter rolling-flow test section of the Langley stability tunnel to determine the effects of varying taper ratio on the rolling and static stability characteristics of a swept wing are presented; results are also given for the effects of varying taper ratio on an unswept wing and for the effects of sweep on a tapered wing. All the models were of aspect ratio 2.61 and had NACA 0012 sections normal to the quarter-chord line. Taper ratios of 1.00, 0.50, and 0.25 and sweep angles of 0 degrees and 45 degrees were investigated.

WINGS, COMPLETE - SWEEP↗
NASA NTRS1957

Sidewash in the Vicinity of Lifting Swept Wings at Supersonic Speeds

Equations and charts have been presented for the conical part of the sidewash above and below swept wings at small angles of attack. This flow-field component is necessary in making calculations of the wing-induced forces and moments on stores, missiles, and pylons. All supersonic Mach numbers and wing leading-edge sweep angles have been considered.

AERODYNAMIC LOADS, BODIES↗
NASA NTRS1948

Charts for Determining Preliminary Values of Span-load, Shear, Bending-moment, and Accumulated-torque Distributions of Swept Wings of Various Taper Ratios

Contains charts for use in determining preliminary values of the spanwise-load, shear, bending-moment, and accumulated-torque distributions of swept wings. The charts are based on strip theory and include four aerodynamic-load distributions, two section-moment distributions, and two inertia-load distributions. The taper ratios considered cover the range from 1.0 to 0 and the results are applicable to any angle of sweep.

LOADS, MANEUVERING - WINGS↗