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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 397 records · Page 22

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NASA NTRS1984

An Analysis of Wave Interactions in Swept-Wing Flows

Crossflow instabilities dominate disturbance growth in the leading-edge region of swept wings. Streamwise vortices in a boundary layer strongly influence the behavior of other disturbances. Amplification of crossflow vortices near the leading edge produces a residual spanwise nonuniformity in the mid-chord regions where Tollmien-Schlichting (T-S) waves are strongly amplified. Should the T-S wave undergo double-exponential growth because of this effect, the usual transition prediction methods would fail. The crossflow/Tollmien-Schlichting wave interaction was modeled as a secondary instability. The effects of suction are included, and different stability criteria are examined. The results are applied to laminar flow control wings characteristic of energy-efficient aircraft designs.

Reed, H. L.↗

NASA NTRS1983

Study of transonic flow over swept wings

The design of shockless airfoils that are appropriate for experimental work on a supersonic transport with an oblique wing are examined. A series of computer codes for the design and analysis of airfoils and wings in two dimensional and three dimensional transonic flow are studied. The oblique wing 3-D code was the forerunner of the later swept wing code. Techniques to incorporate the effect of an engine or fuselage in the inverse design code while using a minimum of computer resources are developed.

Source record↗

NASA NTRS1984

Lightning swept-stroke attachment patterns and flight conditions for storm hazards 1981

As part of the NASA Langley Research Center Storm Hazards Program, 111 thunderstorm penetrations were made in 1981 with an F-106B airplane in order to record direct-strike lightning data and the associated flight conditions. Ground-based weather radar measurements in conjunction with these penetrations were made by NOAA National Severe Storms Laboratory in Oklahoma and by NASA Wallops Flight Facility in Virginia. In 1981, the airplane received 10 direct lightning strikes; in addition, lightning transient data were recorded from 22 nearby flashes. Following each flight, the airplane was thoroughly inspected for evidence of lightning attachment, and the individual lightning attachment points were plotted on isometric projections of the airplane to identify swept-flash patterns. This report shows the strike attachment patterns that were found, and tabulates the flight conditions at the time of each lightning event. Finally, this paper contains a table in which the data in this report are cross-referenced with the previously published electromagnetic waveform data recorded onboard the airplane.

Fisher, B. D.↗

NASA NTRS1984

Aerodynamic design of the contoured wind-tunnel liner for the NASA supercritical, laminar-flow-control, swept-wing experiment

An overview is presented of the entire procedure developed for the aerodynamic design of the contoured wind tunnel liner for the NASA supercritical, laminar flow control (LFC), swept wing experiment. This numerical design procedure is based upon the simple idea of streamlining and incorporates several transonic and boundary layer analysis codes. The liner, presently installed in the Langley 8 Foot Transonic Pressure Tunnel, is about 54 ft long and extends from within the existing contraction cone, through the test section, and into the diffuser. LFC model testing has begun and preliminary results indicate that the liner is performing as intended. The liner design results presented in this paper, however, are examples of the calculated requirements and the hardware implementation of them.

Newman, P. A.↗

NASA NTRS1983

Applicability of the independence principle to subsonic turbulent flow over a swept rearward-facing step

Prandtl (1946) has concluded that for yawed laminar incompressible flows the streamwise flow is independent of the spanwise flow. However, Ashkenas and Riddell (1955) have reported that for turbulent flow the 'independence principle' does not apply to yawed flat plates. On the other hand, it was also found that this principle may be applicable to many turbulent flows. As the sweep angle is increased, a sweep angle is reached which defines the interval over which the 'independence principle' is valid. The results obtained in the present investigation indicate the magnitude of the critical angle for subsonic turbulent flow over a swept rearward-facing step.

Selby, G. V.↗

NASA NTRS1984

Flowfield scaling of a swept compression corner interaction A comparison of experiment and computation

Over the past decade, there has been some important progress in understanding the problems of three-dimensional (3D) shock wave/turbulent boundary layer interactions. However, the problems are by no means solved. The present investigation has the objective to determine the flowfield structure of a swept compression corner interaction and to perform a test of an equation considered by Settles and Bogdonoff (1982). The experimental data obtained in a 20 x 20 cm high Reynolds number supersonic wind tunnel are compared with the predictions of a state-of-the-art numerical solution of the Navier-Stokes equations. It is found that a 3D scaling law for Reynolds number effects, previously established for interaction 'footprints', is equally valid when applied to the present flowfield.

Settles, G. S.↗

NASA NTRS1984

Wave interactions in swept-wing flows

Crossflow instabilities dominate disturbance growth in the leading-edge region of swept wings. It is well known that streamwise vortices in a boundary layer strongly influence the behavior of other disturbances. Amplification of crossflow vortices near the leading edge produces a residual spanwise nonuniformity in the mid-chord regions where Tollmien-Schlichting (T-S) waves are strongly amplified. Should the T-S wave undergo double-exponential growth because of this effect, the usual transition prediction methods would fail. Thus, it is important to study interactions of this sort and to develop more realistic criteria for transition prediction.

Reed, H. L.↗

NASA NTRS1983

The NYU inverse swept wing code

An inverse swept wing code is described that is based on the widely used transonic flow program FLO22. The new code incorporates a free boundary algorithm permitting the pressure distribution to be prescribed over a portion of the wing surface. A special routine is included to calculate the wave drag, which can be minimized in its dependence on the pressure distribution. An alternate formulation of the boundary condition at infinity was introduced to enhance the speed and accuracy of the code. A FORTRAN listing of the code and a listing of a sample run are presented. There is also a user's manual as well as glossaries of input and output parameters.

Bauer, F.↗

NASA NTRS1985

Effects of leading-edge devices on the low-speed aerodynamic characteristics of a highly-swept arrow-wing

An investigation was conducted in the Texas A&M University 7 by 10 foot Low Speed Wind Tunnel to provide a direct comparison of the effect of several leading edge devices on the aerodynamic performance of a highly swept wing configuration. Analysis of the data indicates that for the configuration with undeflected leading edges, vortex separation first occurs on the outboard wing panel for angles of attack of approximately 2, and wing apex vorticies become apparent for alpha or = 4 deg. However, the occurrence of the leading edge vortex flow may be postponed with leading edge devices. Of the devices considered, the most promising were a simple leading edge deflection of 30 deg and a leading edge slat system. The trailing edge flap effectiveness was found to be essentially the same for the configuration employing either of these more promising leading edge devices. Analysis of the lateral directional data showed that for all of the concepts considered, deflecting leading edge downward in an attempt to postpone leading edge vortex flows, has the favorable effect of reducing the effective dihedral.

Scott, S. J.↗

NASA NTRS1984

Noise produced by the interaction of a rotor wake with a swept stator blade

An analysis is developed for the noise generated by the interaction of rotor viscous wakes and a single swept stator vane. The stator vane spans a channel with infinite parallel walls which contains a uniform subsonic mean flow. High frequency wakes, for which the noise generation is concentrated at the vane leading edge, are considered. The general wake pattern is expanded in spanwise modes and solutions for each mode are derived using the Wiener-Hopf technique applied to the equations in the nonorthogonal coordinates. Closed form expressions for the acoustic farfield are obtained. The results of the analysis are used in parametric calculations of rotor viscous wake-stator vane interactions in order to study the effectiveness of sweep as a noise reduction mechanism. For the cases studied, moderate stator sweep angles produce sizeable reductions in the level of the farfield noise. The presence of rotor wake circumferential lean actually increases the noise reduction produced by moderate stator sweep angles.

Envia, E.↗

NASA NTRS1984

Applicability of a panel method, which includes nonlinear effects, to a forward-swept-wing aircraft

The ability of a lower order panel method VSAERO, to accurately predict the lift and pitching moment of a complete forward-swept-wing/canard configuration was investigated. The program can simulate nonlinear effects including boundary-layer displacement thickness, wake roll up, and to a limited extent, separated wakes. The predictions were compared with experimental data obtained using a small-scale model in the 7- by 10- Foot Wind Tunnel at NASA Ames Research Center. For the particular configuration under investigation, wake roll up had only a small effect on the force and moment predictions. The effect of the displacement thickness modeling was to reduce the lift curve slope slightly, thus bringing the predicted lift into good agreement with the measured value. Pitching moment predictions were also improved by the boundary-layer simulation. The separation modeling was found to be sensitive to user inputs, but appears to give a reasonable representation of a separated wake. In general, the nonlinear capabilities of the code were found to improve the agreement with experimental data. The usefullness of the code would be enhanced by improving the reliability of the separated wake modeling and by the addition of a leading edge separation model.

Ross, J. C.↗

NASA NTRS1985

Transonic test of a forward swept wing configuration exhibiting Body Freedom Flutter

The aeroelastic dynamic instability designated Body Freedom Flutter (BFF) involves aircraft pitch and wing bending motions characteristic of forward swept wing (FSW) aircraft. Attention is presently given to the results of tests conducted on a 1/2-scale cable-mounted FSW wind tunnel model, with and without relaxed static stability (RSS) control conditions. BFF instability boundaries were found to occur at significantly lower air speeds than those associated with aeroelastic wing divergence on the same model. Servoaeroelastic stability analyses have been conducted which proved capable of predicting the measured onset of BFF, in both the statically stable and RSS configurations tested.

Chipman, R.↗

NASA NTRS1985

Flutter of swept fan blades

The effect of sweep on fan blade flutter is studied by applying the analytical methods developed for aeroelastic analysis of advance turboprops. Two methods are used. The first method utilizes an approximate structural model in which the blade is represented by a swept, nonuniform beam. The second method utilizes a finite element technique to conduct modal flutter analysis. For both methods the unsteady aerodynamic loads are calculated using two dimensional cascade theories which are modified to account for sweep. An advanced fan stage is analyzed with 0, 15 and 30 degrees of sweep. It is shown that sweep has a beneficial effect on predominantly torsional flutter and a detrimental effect on predominantly bending flutter. This detrimental effect is shown to be significantly destabilizing for 30 degrees of sweep.

Kielb, R. E.↗

NASA NTRS1985

Conical similarity of shock/boundary-layer interactions generated by swept and unswept fins

A parametric experimental investigation has been made of the class of three-dimensional shock wave/turbulent boundary layer interactions generated by swept and unswept leading-edge fins. The fin sweepback angles were 0-65 deg at 5, 9, and 15 deg angles of attack. Two equilibrium two-dimensional turbulent boundary layers with a freestream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All of the resulting interactions were found to possess conical symmetry of the surface flow patterns and pressures outside of an initial inception zone. Further, these interactions were found to obey a simple conical similarity rule based on inviscid shock wave strength, irrespective of fin sweepback or angle of attack. This is one of the first demonstrations of similarity among three-dimensional interactions produced by geometrically dissimilar shock generators.

Settles, G. S.↗

NASA NTRS1985

The buffeting pressure field of a high-aspect-ratio swept wing

An experimental study of the fluctuating pressure field of a high-aspect-ratio swept transport-type wing model in transonic buffeting flow is examined. A high degree of similarity between the buffet-related features of the wing flowfield and 2-D transonic flowfields is indicated. At several spanwise stations, unsteady lift produced by the pressure fluctuations is evaluated and found to increase in intensity as local flow separation develops. Results show that although the wing does vibrate, the measured lift fluctuations are driving, and not driven by wing vibrations. The influence of tunnel-flow unsteadiness on the data is considered.

Roos, F. W.↗

NASA NTRS1985

Installation effects of long-duct pylon-mounted nacelles on a twin-jet transport model with swept supercritical wing

The installation interference effects of an underwing-mounted, long duct, turbofan nacelle were evaluated in the Langley 16-Foot Transonic Tunnel with two different pylon shapes installed on a twin engine transport model having a supercritical wing swept 30 deg. Wing, pylon, and nacelle pressures and overall model force data were obtained at Mach numbers from 0.70 to 0.83 and nominal angles of attack from -2 deg to 4 deg at an average unit Reynolds number of 11.9 x 1,000,000 per meter. The results show that adding the long duct nacelles to the supercritical wing, in the near sonic flow field, changed the magnitude and direction of flow velocities over the entire span, significantly reduced cruise lift, and caused large interference drag on the nacelle afterbody.

Lee, E. E., Jr.↗

NASA NTRS1986

ALESEP. Part 2: A computer program for the analysis of leading edge separation bubbles on infinite swept wings

A program called ALESEP is presented for the analysis of the inviscid-viscous interaction which occurs due to the presence of a closed laminar-transitional separation bubble on an airfoil or infinite swept wing. The ALESEP code provides an iterative solution of the boundary layer equations expressed in an inverse formulation coupled to a Cauchy integral representation of the inviscid flow. This interaction analysis is treated as a local perturbation to a known solution obtained from a global airfoil analysis; hence, part of the required input to the ALESEP code are the reference displacement thickness and tangential velocity distributions. Special windward differencing may be used in the reversed flow regions of the separation bubble to accurately account for the flow direction in the discretization of the streamwise convection of momentum. The ALESEP code contains a forced transition model based on a streamwise intermittency function, a natural transition model based on a solution of the integral form of the turbulent kinetic energy equation, and an empirical natural transition model.

Davis, R. L.↗

NASA NTRS1986

Divergence study of a high-aspect ratio, forward-swept wing

An experimental wind-tunnel study to determine the divergence characteristics of a high-aspect ratio, forward-swept wing has been conducted in the NASA Langley Research Center (LaRC) Transonic Dynamics Tunnel (TDT). The rectangular wing used for this study had a panel aspect ratio of 9.16 (lambda = 0 deg.) and the sweep angle could be set at lambda = 0 deg., -15 deg., -30 deg., -45 deg., or -60 deg. A rectangular wing tip shape was tested at each of these sweep angles. In addition, a tip shape parallel to the freestream flow was tested for a wing sweep angle of lambda = -45 deg. The root of the wing was cantilever mounted to the wall of the wind tunnel. Divergence conditions were measured at M = 0.4 for each sweep angle and tip configuration tested. Subcritical response techniques were used to extrapolate to the divergence conditions during the wind-tunnel test. The primary objective of this test was to obtain data which could be used to verify for this configuration the divergence prediction capability of an aeroelastic analysis code. Subsonic lifting surface theory (kernel function) aerodynamics are utilized by this particular code. The analytical predictions of divergence were found to be significantly conservative at all forward sweep angles. At lambda = -45 deg., the analysis was 14 percent conservative. The effect of the two tip shapes on the divergence dynamic pressure was predicted accurately by the analysis. The divergence condition for the tip shape parallel to the flow occurred at a dynamic pressure 14 percent higher than the divergence condition with a rectangular tip shape.

Cole, S. R.↗