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At least 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Experimental investigation of the effects of aft blowing with various nozzle exit geometries on a 3.0 caliber tangent ogive at high angles of attack: Forebody pressure distributions

An experimental study of the effects of aft blowing on the asymmetric vortex flow of a slender, axisymmetric body at high angles of attack was conducted. A 3.0 caliber tangent ogive body fitted with a cylindrical afterbody was tested in a wind tunnel under subsonic, laminar flow test conditions. Asymmetric blowing from both a single nozzle and a double nozzle configuration, positioned near the body apex, was studied. Aft blowing was observed to alter the vortex asymmetry by moving the blowing-side vortex closer to the body surface while moving the non-blowing-side vortex further away from the body. The effect of increasing the blowing coefficient was to move the blowing-side vortex closer to the body surface at a more upstream location. The data also showed that blowing was more effective in altering the initial vortex asymmetry at the higher angles of attack than at the lower. The effects of changing the nozzle exit geometry were studied and it was observed that blowing from a nozzle with a low, broad exit geometry was more effective in reducing the vortex asymmetry than blowing from a high, narrow exit geometry.

Chokani, Ndaona↗
NASA NTRS1976

Asymmetric aerodynamic forces on aircraft at high angles of attack - some design guides

Aerodynamic side forces on forebodies are considered that are produced by two types of flow: asymmetric vortices on bodies of revolution and nonuniform flow separation on square bodies with rounded corners under spinning conditions. Steady side forces that can be as large as the normal force are produced by asymmetric vortices on pointed forebodies. This side force has a large variation with Reynolds number, decreases rapidly with Mach number, and can be nearly eliminated with small nose bluntness or strakes. The angle of attack where the side force first occurs depends primarily on body geometry. The theoretical techniques to predict these side forces are necessarily semi-empirical because the basic phenomenon is not well understood. The side forces produced by nonuniform flow separation under spinning conditions depend extensively on spin rate, angle of attack, and Reynolds number. The application of simple crossflow theory to predict this side force is inadequate much below angles of attack of 90 deg.

Chapman, G. T.↗
NASA NTRS1991

Update of the X-29 high-angle-of-attack program

The X-29A forward-swept wing flight research aircraft flight envelope was expanded to 66 deg angle-of-attack during 1990. Following this flight envelope expansion, a military utility evaluation was performed to investigate the tactical utility of the X-29 configurations at high-angle -of-attack, slow-speed flight conditions. An overall management view and perspective of the expansion process, the technical problems encountered, and the results obtained when compared to the predictions are presented.

Trippensee, Gary↗
NASA NTRS1990

Noise of a simulated installed model counterrotation propeller at angle-of-attack and takeoff/approach conditions

Acoustic results for two model counterrotation propellers are presented. The propellers were tested over a range of rotational speeds and propeller axis angles of attack in both the baseline configuration and the installed configuration consisting of a simulated upstream nacelle support pylon and fuselage section. Acoustic data were taken with a polar microphone probe attached to the downstream propeller housing, capable of surveying directivities at several azimuthal locations. The forward and aft rotor power coefficients and fundamental rotor-alone tone levels are found to be directly controlled by propeller axis angle of attack. The second-order rotor-alone tones are strongly influenced by the upstream pylon wake at 80 percent speed; however, rotor-alone mechanisms control the tone level at 90 percent speed, while rotor-rotor interaction tones are essentially unaffected by the presence of the simulated installation.

Woodward, Richard P.↗
NASA NTRS1986

Experimental study of effects of forebody geometry on high angle of attack static and dynamic stability and control

A series of low-speed wind tunnel tests on a generic airplane model with a cylindrical fuselage were made to investigate the effects of forebody shape and fitness ratio, and fuselage/wing proximity on static and dynamic lateral/directional stability. In addition, some preliminary testing to determine the effectiveness of deflectable forebody strakes for high angle of attack yaw control was conducted. During the stability investigation, 11 forebodies were tested including three different cross-sectional shapes with fineness ratios of 2, 3, and 4. In addition, the wing was tested at two longitudinal positions to provide a substantial variation in forebody/wing proximity. Conventional force tests were conducted to determine static stability characteristics, and single-degree-of-freedom free-to-roll tests were conducted to study the wing rock characteristics of the model with the various forebodies. Flow visualization data were obtained to aid in the analysis of the complex flow phenomena involved. The results show that the forebody cross-sectional shape and fineness ratio and forebody/wing proximity can strongly affect both static and dynamic (roll) stability at high angles of attack. These characteristics result from the impact of these factors on forebody vortex development, the behavior of the vortices in sideslip, and their interaction with the wing flow field. Preliminary results from the deflectable strake investigation indicated that forebody flow control using this concept can provide very large yaw control moments at stall and post-stall angles of attack.

Brandon, J. M.↗
NASA NTRS1988

Wind-tunnel investigation of the forebody aerodynamics of a vortex-lift fighter configuration at high angles of attack

Results of a recent low-speed wind-tunnel investigation conducted to define the forebody flow on a 16-percent scale model of the NASA High Angle-of-Attack Research Vehicle, an F-18 configuration, are presented with analysis. Measurements include force and moment data, oil-flow visualizations, and surface pressure data taken at angles of attack near and above maximum lift (36 to 52 deg) at a Reynolds number of one million (based on mean aerodynamic chord). The results presented identify the key flow-field features on the forebody including the wing-body strake.

Banks, Daniel W.↗
NASA NTRS1983

Magnus effects at high angles of attack and critical Reynolds numbers

The Magnus force and moment experienced by a yawed, spinning cylinder were studied experimentally in low speed and subsonic flows at high angles of attack and critical Reynolds numbers. Flow-field visualization aided in describing a flow model that divides the Magnus phenomenon into a subcritical region, where reverse Magnus loads are experienced, and a supercritical region where these loads are not encountered. The roles of the spin rate, angle of attack, and crossflow Reynolds number in determining the boundaries of the subcritical region and the variations of the Magnus loads were studied.

Seginer, A.↗
NASA NTRS2005

Application of Piloted Simulation to High-Angle-of-Attack Flight-Dynamics Research for Fighter Aircraft

This paper reviews the use of piloted simulation at Langley Research Center as part of the NASA High-Angle-of-Attack Technology Program (HATP), which was created to provide concepts and methods for the design of advanced fighter aircraft. A major research activity within this program is the development of the design processes required to take advantage of the benefits of advanced control concepts for high-angle-of-attack agility. Fundamental methodologies associated with the effective use of piloted simulation for this research are described, particularly those relating to the test techniques, validation of the test results, and design guideline/criteria development.

Ogburn, Marilyn E.↗
NASA NTRS1957

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Brown, Clarence A , Jr↗
NASA NTRS1990

Near-field noise of a single rotation propfan at an angle of attack

The near-field noise characteristics of a propfan operating at an angle of attack are examined utilizing the unsteady pressure field obtained from a three-dimensional Euler simulation of the propfan flowfield. The near-field noise is calculated employing three different procedures: a direct computation method in which the noise field is extracted directly from the Euler solution, and two acoustic-analogy-based frequency domain methods which utilize the computed unsteady pressure distribution on the propfan blades as the source term. The inflow angles considered are -0.4, 1.6, and 4.6 degrees. The results of the direct computation method and one of the frequency domain methods show qualitative agreement with measurements. They show that an increase in the inflow angle is accompanied by an increase in the sound pressure level at the outboard wing boom locations and a decrease in the sound pressure level at the (inboard) fuselage locations. The trends in the computed azimuthal directivities of the noise field also conform to the measured and expected results.

Nallasamy, M.↗
NASA NTRS1988

Modeling of large-amplitude high-angle-of-attack maneuvers

Results from a simulation study to characterize and model rapid, large-amplitude high-angle-of-attack maneuvers of high performance aircraft are summarized. There is concern that conventional aerodynamic math models combining static and small-amplitude demping data may be deficient in properly representing the associated aerodynamics to allow accurate prediction of these types of motion. The objectives of this investigation were: (1) to characterize these maneuvers in order to establish the aerodynamic modeling requirements, (2) to define refinements to address these requirements, and (3) to examine the impact of these refinements on the calculated motions. The results show that accurate modeling of the dynamic terms is vital due to the very high angular rates generated during these maneuvers. It was found that the addition of dynamic terms to include the lateral acceleration derivatives and the rotary aerodynamics can in some cases significantly affect the simulated flight motions.

Angle of attack↗
NASA NTRS1983

Subsonic high-angle-of-attack aerodynamic characteristics of a cone and cylinder with triangular cross sections and a cone with a square cross section

Experiments were conducted in the 12-Foot Pressure Wind Tunnel at Ames Research Center on three models with noncircular cross sections: a cone having a square cross section with rounded corners and a cone and cylinder with triangular cross sections and rounded vertices. The cones were tested with both sharp and blunt noses. Surface pressures and force and moment measurements were obtained over an angle of attack range from 30 deg to 90 deg and selected oil-flow experiments were conducted to visualize surface flow patterns. Unit Reynolds numbers ranged from 0.8x1,000,000/m to 13.0x1,000,000/m at a Mach number of 0.25, except for a few low-Reynolds-number runs at a Mach number of 0.17. Pressure data, as well as force data and oil-flow photographs, reveal that the three dimensional flow structure at angles of attack up to 75 deg is very complex and is highly dependent on nose bluntness and Reynolds number. For angles of attack from 75 deg to 90 deg the sectional aerodynamic characteristics are similar to those of a two dimensional cylinder with the same cross section.

Clarkson, M. H.↗
NASA NTRS1985

Pivotable strakes for high angle of attack control

A series of pivotable strakes have been tested in combination with a 44-deg swept wing to determine their ability to provide pitch control at high angles of attack. The tests were carried out in the Langley 7 X 10-Foot High-Speed tunnel at nominal Mach numbers of 0.3 to 0.4. A total of five strakes with various aspects ratios and shapes were tested at negative deflection angles of 5, 10, 15, and 20 deg through an angle of attack range up to approximately 50 deg. In addition, a cranked delta wing with 70-deg sweep was tested with two different pivotable apex flaps. This paper shows lift, drag and moment data for the complete configurations as well as for the separate contributions of the strakes and fuselage nose which are mounted on a second balance. Generally, strake deflection was found to have a significant effect on moment while lift/drag changed very little.

Bobbitt, P. J.↗
NASA NTRS2001

Comparison of Angle of Attack Measurements for Wind Tunnel Testing

Two optical systems capable of measuring model attitude and deformation were compared to inertial devices employed to acquire wind tunnel model angle of attack measurements during the sting mounted full span 30% geometric scale flexible configuration of the Northrop Grumman Unmanned Combat Air Vehicle (UCAV) installed in the NASA Langley Transonic Dynamics Tunnel (TDT). The overall purpose of the test at TDT was to evaluate smart materials and structures adaptive wing technology. The optical techniques that were compared to inertial devices employed to measure angle of attack for this test were: (1) an Optotrak (registered) system, an optical system consisting of two sensors, each containing a pair of orthogonally oriented linear arrays to compute spatial positions of a set of active markers; and (2) Video Model Deformation (VMD) system, providing a single view of passive targets using a constrained photogrammetric solution whose primary function was to measure wing and control surface deformations. The Optotrak system was installed for this test for the first time at TDT in order to assess the usefulness of the system for future static and dynamic deformation measurements.

Jones, Thomas, W.↗
NASA NTRS1957

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

A full-scale rocket-powered model of a cruciform canard missile configuration with a low-aspect-ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed-control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift-curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift-curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta = -0.3 deg. The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic-center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number. The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Brown, C. A., Jr.↗
NASA NTRS1996

Experimental and Computational Induced Aerodynamics from Missile Jet Reaction Controls at Angles of Attack to 75 Degrees

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine induced aerodynamic effects from jet reaction controls of an advanced air-to-air missile concept. The 75-percent scale model featured independently controlled reaction jets located near the nose and tail of the model. Aerodynamic control was provided by four fins located near the tail of the model. This investigation was conducted at Mach numbers of 0.35 and 0.60, at angles of attack up to 75 deg and at nozzle pressure ratios up to 90. Jet-reaction thrust forces were not measured by the force balance but jet-induced forces were. In addition, a multiblock three-dimensional Navier-Stokes method was used to calculate the flowfield of the missile at angles of attack up to 40 deg. Results indicate that large interference effects on pitching moment were induced from operating the nose jets with the the off. Excellent correlation between experimental and computational pressure distributions and pitching moment were obtained a a Mach number of 0.35 and at angles of attack up to 40 deg.

Capone, Francis J.↗
NASA NTRS1993

Longitudinal-control design approach for high-angle-of-attack aircraft

This paper describes a control synthesis methodology that emphasizes a variable-gain output feedback technique that is applied to the longitudinal channel of a high-angle-of-attack aircraft. The aircraft is a modified F/A-18 aircraft with thrust-vectored controls. The flight regime covers a range up to a Mach number of 0.7; an altitude range from 15,000 to 35,000 ft; and an angle-of-attack (alpha) range up to 70 deg, which is deep into the poststall region. A brief overview is given of the variable-gain mathematical formulation as well as a description of the discrete control structure used for the feedback controller. This paper also presents an approximate design procedure with relationships for the optimal weights for the selected feedback control structure. These weights are selected to meet control design guidelines for high-alpha flight controls. Those guidelines that apply to the longitudinal-control design are also summarized. A unique approach is presented for the feed-forward command generator to obtain smooth transitions between load factor and alpha commands. Finally, representative linear analysis results and nonlinear batch simulation results are provided.

Ostroff, Aaron J.↗
NASA NTRS1995

Numerical simulation of the flow about the F-18 HARV at high angle of attack

This research has been aimed at validating numerical methods for computing the flow about the complete F-18 HARV at alpha = 30 deg and alpha = 45 deg. At 30 deg angle of attack, the flow about the F-18 is dominated by the formation, and subsequent breakdown, of strong vortices over the wing leading-edge extensions (LEX). As the angle of attack is increased to alpha = 45 deg, the fuselage forebody of the F-18 contains significant laminar and transitional regions which are not present at alpha = 30 deg. Further, the flow over the LEX at alpha = 45 deg is dominated by an unsteady shedding in time, rather than strong coherent vortices. This complex physics, combined with the complex geometry of a full-aircraft configuration, provides a challenge for current computational fluid dynamics (CFD) techniques. The following sections present the numerical method and grid generation scheme that was used, a review of prior research done to numerically model the F-18 HARV, and a discussion of the current research. The current research is broken into three main topics; the effect of engine-inlet mass-flow rate on the F-18 vortex breakdown position, the results using a refined F-18 computational model to compute the flow at alpha = 30 deg and alpha = 45 deg, and research done using the simplified geometry of an ogive-cylinder configuration to investigate the physics of unsteady shear-layer shedding. The last section briefly summarizes the discussion.

Murman, Scott M.↗