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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

AEROX: Computer program for transonic aircraft aerodynamics to high angles of attack. Volume 1: Aerodynamic methods and program users' guide

The AEROX program estimates lift, induced-drag and pitching moments to high angles (typ. 60 deg) for wings and for wingbody combinations with or without an aft horizontal tail. Minimum drag coefficients are not estimated, but may be input for inclusion in the total aerodynamic parameters which are output in listed and plotted formats. The theory, users' guide, test cases, and program listing are presented.

Axelson, J. A.↗
NASA NTRS1984

Evaluation of a flow direction probe and a pitot-static probe on the F-14 airplane at high angles of attack and sideslip

The measurement performance of a hemispherical flow-angularity probe and a fuselage-mounted pitot-static probe was evaluated at high flow angles as part of a test program on an F-14 airplane. These evaluations were performed using a calibrated pitot-static noseboom equipped with vanes for reference flow direction measurements, and another probe incorporating vanes but mounted on a pod under the fuselage nose. Data are presented for angles of attack up to 63, angles of sideslip from -22 deg to 22 deg, and for Mach numbers from approximately 0.3 to 1.3. During maneuvering flight, the hemispherical flow-angularity probe exhibited flow angle errors that exceeded 2 deg. Pressure measurements with the pitot-static probe resulted in very inaccurate data above a Mach number of 0.87 and exhibited large sensitivities with flow angle.

Larson, T. J.↗
NASA NTRS1987

Numerical simulation by TVD schemes of complex shock reflections from airfoils at high angle of attack

The shock-capturing capability of total variation diminishing (TVD) schemes is demonstrated for a more realistic complex shock-diffraction problem for which the experimental data are available. Second-order explicit upwind and symmetric TVD schemes are used to solve the time-dependent Euler equations of gas dynamics for the interaction of a blast wave with an airfoil at high angle-of-attack. The test cases considered are a time-dependent moving curved-shock wave and a contant moving planar-shock wave impinging at an angle-of-attack 30 deg on a NACA 0018 airfoil. Good agreement is obtained between isopycnic contours computed by the TVD schemes and those from experimental interferograms. No drastic difference in flow-field structure is found between the curved- and planar-shock wave cases, except for a difference in density level near the lower surface of the airfoil. Computation for cases with higher shock Mach numbers is also possible. Numerical experiments show that the symmetric TVD scheme is less sensitive to the boundary conditions treatment than the upwind scheme.

Moon, Young J.↗
NASA NTRS1990

Comparison of high-angle-of-attack slender-body theory and exact solutions for potential flow over an ellipsoid

The accuracy of high-alpha slender-body theory (HASBT) for bodies with elliptical cross-sections is presently demonstrated by means of a comparison with exact solutions for incompressible potential flow over a wide range of ellipsoid geometries and angles of attack and sideslip. The addition of the appropriate trigonometric coefficients to the classical slender-body theory decomposition yields the formally correct HASBT, and results in accuracies previously considered unattainable.

Hemsch, Michael J.↗
NASA NTRS1999

Factors Affecting Inlet-Engine Compatibility During Aircraft Departures at High Angle of Attack for an F/A -18A Aircraft

The F404-GE-400 engine powered F/A- 18A High Alpha Research Vehicle (HARV) was used to examine the quality of inlet airflow during departed flight maneuvers, that is, during flight outside the normal maneuvering envelope where control surfaces have little or no effectiveness. A series of six nose-left and six nose-right departures were initiated at Mach numbers between 0.3 and 0.4 at an altitude of 35 kft. The yaw rates at departure recovery were in the range of 40 to 90 degrees per second. Engine surges were encountered during three of the nose-left and one of the nose-right departures. Time-variant inlet-total-pressure distortion levels at the engine face were determined to not significantly exceed those measured at maximum angle-of-attack and - sideslip maneuvers during controlled flight. Surges as a result of inlet distortion levels were anticipated to initiate in the fan. Analysis revealed that the surges initiated in the compressor and were the result of a combination of high levels of inlet distortion and rapid changes in aircraft motion. These rapid changes in aircraft motion are indicative of a combination of engine mount and gyroscopic loads being applied to the engine structure that impact the aerodynamic stability of the compressor through changes in the rotor-to-case clearances.

Steenken, W. G.↗
NASA NTRS1991

Numerical study of the X-31 high angle-of-attack flow characteristics

The application of CFD methods to high-angle-of-attack flight regimes is studied. The X-31 configuration was chosen as the model due to its inherent designed high-angle-of-attack characteristics and capabilities, the existence of a wind tunnel database, and the ongoing flight test program. The flow characteristics of the X-31 at high angle-of-attack are presented and the CFD capability for studying the effects of canard deflection on aerodynamic control during poststall is revealed.

Yeh, David T.↗
NASA NTRS1987

Computation of asymmetric supersonic flows around cones at large incidence

The Schiff-Steger parabolized Navier-Stokes (PNS) code has been modified to allow computation of conical flowfields around cones at high incidence. The improved algorithm of Degani and Schiff has been incorporated with the PNS code. This algorithm adds the cross derivative and circumferential viscous terms to the original PNS code and modifies the algebraic eddy viscosity turbulence model to take into account regions of so called cross-flow separation. Assuming the flowfield is conical (but not necessarily symmetric) a marching stepback procedure is used: the solution is marched one step downstream using improved PNS code and the flow variables are then scaled to place the solution back to the original station. The process is repeated until no change in the flow variables is observed with further marching. The flow variables are then constant along rays of the flowfield. The experiments obtained by Bannik and Nebbeling were chosen as a test case. In these experiments a cone of 7.5 deg. half angle at Mach number 2.94 and Reynolds number 1.372 x 10(7) was tested up 34 deg. angle of attack. At high angle of attack nonconical asymmetric leeward side vortex patterns were observed. In the first set of computations, using an earlier obtained solution of the above cone for angle of attack of 22.6 deg. and at station x=0.5 as a starting solution, the angle of attack was gradually increased up to 34 deg. During this procedure the grid was carfully adjusted to capture the bow shock. A stable, converged symmetric solution was obtained. Since the numerical code converged to a symmetric solution which is not the physical one, the stability was tested by a random perturbation at each point. The possible effect of surface roughness or non perfect body shape was also investigated. It was concluded that although the assumption of conical viscous flows can be very useful for certain cases, it can not be used for the present case. Thus the second part of the investigation attempted to obtain a marching (in space) solution with the PNS method using the conical solution as initial data. Finally, the solution of the full Navier-Stokes equations was carried out.

Degani, David↗
NASA NTRS1993

Robust, nonlinear, high angle-of-attack control design for a supermaneuverable vehicle

High angle-of-attack flight control laws are developed for a supermaneuverable fighter aircraft. The methods of dynamic inversion and structured singular value synthesis are combined into an approach which addresses both the nonlinearity and robustness problems of flight at extreme operating conditions. The primary purpose of the dynamic inversion control elements is to linearize the vehicle response across the flight envelope. Structured singular value synthesis is used to design a dynamic controller which provides robust tracking to pilot commands. The resulting control system achieves desired flying qualities and guarantees a large margin of robustness to uncertainties for high angle-of-attack flight conditions. The results of linear simulation and structured singular value stability analysis are presented to demonstrate satisfaction of the design criteria. High fidelity nonlinear simulation results show that the combined dynamics inversion/structured singular value synthesis control law achieves a high level of performance in a realistic environment.

Adams, Richard J.↗
NASA NTRS1989

Numerical simulation of F-18 fuselage forebody flows at high angles of attack

Fine-grid Navier-Stokes solutions were obtained for flow over the fuselage forebody and wing leading edge extension of the F/A-18 High Alpha Research Vehicle at large incidence. The resulting flows are complex, and exhibit cross flow separation from the sides of the forebody and from the leading edge extension. A well-defined vortex pattern is observed in the leeward-side flow. Results obtained for laminar flow show good agreement with flow visualizations obtained in ground-based experiments. Further, turbulent flows computed at high Reynolds-number flight-test conditions show good agreement with surface and off-surface visualizations obtained in flight.

Schiff, Lewis B.↗
NASA NTRS1990

The effect of turbulence models on the numerical prediction of the flowfield about a prolate spheroid at high angle of attack

The F3D thin-layer Navier-Stokes code presently used to numerically investigate the three-dimensional separated flow about a prolate spheroid at high incidence analyzes the effect of different turbulence models on the flowfield solution and the characteristics of the predicted flow. The Johnson-King (1984) model is applied in order to evaluate the importance of modeling nonequilibrium effects in predicting flow about a slender body at high incidence; the computations in question are for steady-state, fully turbulent flow. Insight is gained into the effects of turbulence models on flow characteristics, and model effects on the accurate prediction of highly separated and vortical flows about a slender body are demonstrated.

Gee, Ken↗
NASA NTRS1990

Unsteady aerodynamic characteristics of a fighter model undergoing large-amplitude pitching motions at high angles of attack

The effects of harmonic or constant-rate-ramp pitching motions (giving angles of attack from 0 to 75 deg) on the aerodynamic performance of a fighter-aircraft model with highly swept leading-edge extensions are investigated experimentally in the NASA Langley 12-ft low-speed wind tunnel. The model configuration and experimental setup are described, and the results of force and moment measurements and flow visualizations are presented graphically and discussed in detail. Large force overshoots and hysteresis are observed and attributed to lags in vortical-flow development and breakup. The motion variables have a strong influence on the persistence of dynamic effects, which are found to affect pitch-rate capability more than flight-path turning performance.

Brandon, Jay M.↗
NASA NTRS1951

Wind-tunnel investigation of six shielded total-pressure tubes at high angles of attack subsonic speeds

The effect of inclination of the air stream on the measured pressures of six shielded total-pressure tubes (all designed for end-mounting on a horizontal boom) has been determined for an angle-of-attack range of approximately -30° to 65°. The tests were conducted at a Mach number of 0.26 in the Langley stability tunnel and at Mach numbers of 0.50, 0.70, 0.90, and 0.95 in the Langley high-speed 7- by 10-foot tunnel. The results of the tests showed that curved venturi entries are less sensitive to inclination of the air stream than the conical entry of the standard Kiel design. Of the three curved entries tested, the tube having the bluntest entry proved the least sensitive to inclination. The critical angle (that is, the sngle.at which the total-pressure error equals 1 percent of the impact pressure) for this tube was ±63.2° at a Mach number of 0.26. The tests also showed that the critical angle of a shielded tube could be extended to higher positive angles of attack by means of a slant profile. The critical angle of a tube having a conical entry with a 10° slant profile, for example, was found to average 9° greater than that of a similar tube having a square profile. Although the critical angle of this tube at angles of yaw was less than that at angles of attack, the critical angle in yaw was equal to that of the tube with the square profile. Tests of the effect of varying the position of the total-pressure probe in the shield showed that, in general, no advantage was to be gained by changing from the position used in the standard Kiel design. Limited tests of the effect of varying the internal diameter of the shield showed that an increase in the diameter of the throat corresponding an increase in the internal area of 50 percent resulted in a relatively small increase (2.5°) in the critical angle. The effect of Mach number on the sensitivity of the tubes to inclination was shown to be appreciable. An increase in Mach number from 0.26 to 0.90, for example, resulted in a decrease in the critical angle of most of the tubes of about 4°.

Walter R Russell↗
NASA NTRS1996

X-29A Lateral-Directional Stability and Control Derivatives Extracted From High-Angle-of-Attack Flight Data

The lateral-directional stability and control derivatives of the X-29A number 2 are extracted from flight data over an angle-of-attack range of 4 degrees to 53 degrees using a parameter identification algorithm. The algorithm uses the linearized aircraft equations of motion and a maximum likelihood estimator in the presence of state and measurement noise. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics over the aircraft at angles of attack above 15 degrees. The results supported the flight-envelope-expansion phase of the X-29A number 2 by helping to update the aerodynamic mathematical model, to improve the real-time simulator, and to revise flight control system laws. Effects of the aircraft high gain flight control system on maneuver quality and the estimated derivatives are also discussed. The derivatives are plotted as functions of angle of attack and compared with the predicted aerodynamic database. Agreement between predicted and flight values is quite good for some derivatives such as the lateral force due to sideslip, the lateral force due to rudder deflection, and the rolling moment due to roll rate. The results also show significant differences in several important derivatives such as the rolling moment due to sideslip, the yawing moment due to sideslip, the yawing moment due to aileron deflection, and the yawing moment due to rudder deflection.

Iliff, Kenneth W.↗
NASA NTRS1989

Development of non-conventional control methods for high angle of attack flight using vortex manipulation

Low-speed wind tunnel experiments have been conducted to define methods for the manipulation of a generic fighter forebody's vortices in order to generate controlled yawing moments at high alpha. Attention is given to two vortex-control schemes: one involving independently controlled forebody tip strakes, and the other individually controlled blowing ports on the forebody. The most effective methods for forebody yawing-moment control were found to be the minimization of natural asymmetry with a pair of symmetrically mounted tip strakes, and the perturbation of the vortex system away from the symmetric condition with asymmetric blowing.

Malcolm, Gerald N.↗
NASA NTRS1977

Theoretical and experimental analysis of longitudinal and lateral aerodynamic characteristics of skewed wings at subsonic speeds to high angles of attack

The effects of sweep and aspect ratio on the longitudinal and lateral-directional aerodynamic characteristics of low-aspect-ratio skewed (oblique) wings having separation-induced vortex flows along leading and side edges were investigated in the Langley high-speed 7- by 10-foot tunnel at a low-subsonic Mach number. The theoretical analysis used the vortex-lattice method for estimating attached-flow aerodynamic characteristics and the leading-edge suction analogy of Polhamus for estimating separation induced vortex-flow effects. Experimental results were compared with asymmetric, separated, vortex flow theory.

Luckring, J. M.↗
NASA NTRS1978

Low-speed aerodynamic characteristics of a 0.08-scale YF-17 airplane model at high angles of attack and sideslip

Data were obtained with and without the nose boom and with several strake configurations; also, data were obtained for various control surface deflections. Analysis of the results revealed that selected strake configurations adequately provided low Reynolds number simulation of the high Reynolds number characteristics. The addition of the boom in general tended to reduce the Reynolds number effects.

Petroff, D. N.↗
NASA NTRS1980

Exploratory investigation of the effects of vortex bursting on the high angle-of-attack lateral-directional stability characteristics of highly-swept wings

A recent low-speed wind-tunnel investigation of highly-swept wings has shown that vortex breakdown at high angles of attack can cause large destabilizing effects on static lateral-directional stability characteristics, and that the destabilizing effects of vortex breakdown can be greatly aggravated by model support strut interference effects. The present paper discusses these effects based on the results of static force tests of several highly-swept wing configurations for different wind-tunnel strut setup arrangements. Also included in the paper are photographs obtained during tuft-, smoke-, and helium-bubble flow visualization studies to indicate wing flow behavior patterns

Johnson, J. L., Jr.↗
NASA NTRS1994

Navy and the HARV: High angle of attack tactical utility issues

This presentation will highlight results from the latest Navy evaluation of the HARV (March 1994) and focus primarily on the impressions from a piloting standpoint of the tactical utility of thrust vectoring. Issue to be addressed will be mission suitability of high AOA flight, visual and motion feedback cues associated with operating at high AOA, and the adaptability of a pilot to effectively use the increased control power provided by the thrust vectoring system.

Sternberg, Charles A.↗