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Sample records for high angle-of-attack aerodynamics

  1. High Angle of Attack Aerodynamics

    DTIC Science & Technology

    1979-01-01

    describe the behaviour of the air. The Symposium made it clear that the present state of knowledge in the area of high angle-of-attack aerodynamics...TESTING EXPERIENCE by C.W. Smith and C.A.Anderson S FOREBODY-WING VORTEX INTERACTIONS AND THEIR INFLUENCE ON DEPARTURE AND SPIN RESISTANCE by A.MSkow...Figure 6) caused by asymmetric flow conditions. ., s already mentioned, asymmetric flow occurs not only when an air- craft flies at non-zero sideslip

  2. High angle of attack hypersonic aerodynamics

    NASA Technical Reports Server (NTRS)

    Harloff, Gary J.

    1987-01-01

    A new aerodynamics force model is presented which is based on modified Newtonian theory and empirical correlations. The algebraic model was developed for complete vehicles from take off to orbital speeds and for large angles of attack. Predictions are compared to results for a wind tunnel model at a Mach number of 20, and the full scale Shuttle Orbiter for Mach numbers from 0.25 to 20 for angles of attack from 0 to 50 deg. The maximum shuttle orbiter lift/drag at Mach 10 and 20 is 1.85 at 20-deg angle-of-attack. Aerodynamic force predictions are made for a transatmospheric vehicle, which is a derivative of the Shuttle Orbiter, for Mach numbers from 4 to 27 at angles of attack from 5 to 40 deg. Predicted aerodynamic force data indicate that lift/drag ratios of 5.2 at Mach number 10 and 3.6 at Mach number 26 are obtainable. Changes in force coefficients with changes in: nose angle, sweep angle, and (volume exp 2/3)/planform area are quantified for Mach numbers of 10 and 26. Lift/drag ratios increase with decreasing nose angle and (volume exp 2/3)/planform area and increasing wing sweep angle. Lift/drag ratios are independent of these variables for angles of attack in excess of 20 deg at Mach 10 and 30 deg at Mach 26.

  3. Aerodynamic characteristics of airplanes at high angles of attack

    NASA Technical Reports Server (NTRS)

    Chambers, J. R.; Grafton, S. B.

    1977-01-01

    An introduction to, and a broad overiew of, the aerodynamic characteristics of airplanes at high angles of attack are provided. Items include: (1) some important fundamental phenomena which determine the aerodynamic characteristics of airplanes at high angles of attack; (2) static and dynamic aerodynamic characteristics near the stall; (3) aerodynamics of the spin; (4) test techniques used in stall/spin studies; (5) applications of aerodynamic data to problems in flight dynamics in the stall/spin area; and (6) the outlook for future research in the area. Although stalling and spinning are flight dynamic problems of importance to all aircraft, including general aviation aircraft, commercial transports, and military airplanes, emphasis is placed on military configurations and the principle aerodynamic factors which influence the stability and control of such vehicles at high angles of attack.

  4. Development of an engineering level prediction method for high angle of attack aerodynamics

    NASA Technical Reports Server (NTRS)

    Reisenthel, Patrick H.; Rodman, Laura C.; Nixon, David

    1993-01-01

    The present work is concerned with predicting the unsteady flow considered to be the cause of the structural failure of twin vertical tail aircraft. An engineering tool has been produced for high angle of attack aerodynamics using the simplest physical models. The main innovation behind this work is its emphasis on the modeling of two key aspects of the dominant physics associated with high angle-of-attack airflows, namely unsteady separation and vortex breakdown.

  5. Calculation of aerodynamic characteristics of airplane configurations at high angles of attack

    NASA Technical Reports Server (NTRS)

    Tseng, J. B.; Lan, C. Edward

    1988-01-01

    Calculation of longitudinal and lateral directional aerodynamic characteristics of airplanes by the VORSTAB code is examined. The numerical predictions are based on the potential flow theory with corrections of high angle of attack phenomena; namely, vortex flow and boundary layer separation effects. To account for the vortex flow effect, vortex lift, vortex action point, augmented vortex lift and vortex breakdown effect through the method of suction analogy are included. The effect of boundary layer separation is obtained by matching the nonlinear section data with the three dimensional lift characteristics iteratively. Through correlation with results for nine fighter configurations, it is concluded that reasonably accurate prediction of longitudinal and static lateral directional aerodynamics can be obtained with the VORSTAB code up to an angle of attack at which wake interference and forebody vortex effect are not important. Possible reasons for discrepancy at higher angles of attack are discussed.

  6. Prediction of Unsteady Aerodynamic Coefficients at High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Pamadi, Bandu N.; Murphy, Patrick C.; Klein, Vladislav; Brandon, Jay M.

    2001-01-01

    The nonlinear indicial response method is used to model the unsteady aerodynamic coefficients in the low speed longitudinal oscillatory wind tunnel test data of the 0.1 scale model of the F-16XL aircraft. Exponential functions are used to approximate the deficiency function in the indicial response. Using one set of oscillatory wind tunnel data and parameter identification method, the unknown parameters in the exponential functions are estimated. The genetic algorithm is used as a least square minimizing algorithm. The assumed model structures and parameter estimates are validated by comparing the predictions with other sets of available oscillatory wind tunnel test data.

  7. High angle-of-attack aerodynamic characteristics of crescent and elliptic wings

    NASA Technical Reports Server (NTRS)

    Vandam, C. P.

    1989-01-01

    Static longitudinal and lateral-directional forces and moments were measured for elliptic- and crescent-wing models at high angles-of-attack in the NASA Langley 14 by 22-Ft Subsonic Tunnel. The forces and moments were obtained for an angle-of-attack range including stall and post-stall conditions at a Reynolds number based on the average wing chord of about 1.8 million. Flow-visualization photographs using a mixture of oil and titanium-dioxide were also taken for several incidence angles. The force and moment data and the flow-visualization results indicated that the crescent wing model with its highly swept tips produced much better high angle-of-attack aerodynamic characteristics than the elliptic model. Leading-edge separation-induced vortex flow over the highly swept tips of the crescent wing is thought to produce this improved behavior at high angles-of-attack. The unique planform design could result in safer and more efficient low-speed airplanes.

  8. Aerodynamic parameters of High-Angle-of attack Research Vehicle (HARV) estimated from flight data

    NASA Technical Reports Server (NTRS)

    Klein, Vladislav; Ratvasky, Thomas R.; Cobleigh, Brent R.

    1990-01-01

    Aerodynamic parameters of the High-Angle-of-Attack Research Aircraft (HARV) were estimated from flight data at different values of the angle of attack between 10 degrees and 50 degrees. The main part of the data was obtained from small amplitude longitudinal and lateral maneuvers. A small number of large amplitude maneuvers was also used in the estimation. The measured data were first checked for their compatibility. It was found that the accuracy of air data was degraded by unexplained bias errors. Then, the data were analyzed by a stepwise regression method for obtaining a structure of aerodynamic model equations and least squares parameter estimates. Because of high data collinearity in several maneuvers, some of the longitudinal and all lateral maneuvers were reanalyzed by using two biased estimation techniques, the principal components regression and mixed estimation. The estimated parameters in the form of stability and control derivatives, and aerodynamic coefficients were plotted against the angle of attack and compared with the wind tunnel measurements. The influential parameters are, in general, estimated with acceptable accuracy and most of them are in agreement with wind tunnel results. The simulated responses of the aircraft showed good prediction capabilities of the resulting model.

  9. Mathematical modeling of the aerodynamics of high-angle-of-attack maneuvers

    NASA Technical Reports Server (NTRS)

    Schiff, L. B.; Tobak, M.; Malcolm, G. N.

    1980-01-01

    This paper is a review of the current state of aerodynamic mathematical modeling for aircraft motions at high angles of attack. The mathematical model serves to define a set of characteristic motions from whose known aerodynamic responses the aerodynamic response to an arbitrary high angle-of-attack flight maneuver can be predicted. Means are explored of obtaining stability parameter information in terms of the characteristic motions, whether by wind-tunnel experiments, computational methods, or by parameter-identification methods applied to flight-test data. A rationale is presented for selecting and verifying the aerodynamic mathematical model at the lowest necessary level of complexity. Experimental results describing the wing-rock phenomenon are shown to be accommodated within the most recent mathematical model by admitting the existence of aerodynamic hysteresis in the steady-state variation of the rolling moment with roll angle. Interpretation of the experimental results in terms of bifurcation theory reveals the general conditions under which aerodynamic hysteresis must exist.

  10. Aerodynamic control of slender bodies from low to high angles of attack through flow manipulation

    NASA Astrophysics Data System (ADS)

    Lopera, Javier

    2007-12-01

    This dissertation presents experimental investigations of several novel active flow control methodologies that have been implemented for aerodynamic control and maneuvering of slender bodies at low and high angles of attack through flow manipulation. For low angles of attack, a U.S. Army Smart Cargo projectile was examined. For high angles of attack a U.S. Air Force countermeasure concept projectile termed DEX (Destructive Expendable) was examined. Low angle of attack control was attempted using two novel separation control techniques: reconfigurable porosity and miniature deployable spoilers. Results show that significant aerodynamic forces are generated by implementing reconfigurable porosity and can be effectively used to steer and maneuver air vehicles. Porous patterns with a "saw-tooth" configuration seem to be the most effective in generating consistent control forces over a wide range of angles of attack. Miniature deployable spoilers successfully demonstrated their ability in producing both positive and negative pitch and yaw controls by modulating the spoiler height and length when used on the boattail and Aero Control Fins (ACFs) of a projectile. The effect of aftbody strake parameters such as shape, locations (axial and azimuthal), deployment height, and number of strakes implemented was examined on a short blunt-nose projectile. Large yaw control authority was attained for alpha > 40 deg. The largest yaw control authority was produced by a rectangular-shaped strake. A robust closed-loop feedback controller was successfully tested using dynamic wind tunnel experiments to control the coning motion of a projectile. The controller showed good control authority and was capable of attaining and maintaining the commanded roll angle with a tolerance of +/-10 deg. A study was conducted to gain some insights into the fluid mechanics of short blunt-nose bodies of revolution at high angles of attack. Off- and on-surface flow visualization records are collected to

  11. Prediction of Aerodynamic Characteristics of Fighter Wings at High Angles of Attack.

    DTIC Science & Technology

    1984-03-01

    method coupled with iterative routines for wake location, viscous effects and vortex flows. Applications of the techniques to a number of...AD-A145 1@7 PREDICTION OF AERODYNAMIC CHARACTERISTICS OF FIGHTER i/2 WIINGS AT HIGH ANGLES OF ATTACK(U) ANALYTICAL METHODS INC REDMOND WA B MASKEW ET...ATTACK I B. !4askew T.S. Vaidyanathan J.K. Nathman F.A. Dvorak Analytical Methods , Inc. 2047 - 152nd Avenue N.E. Redmond, Washington 98052 CONTRACT

  12. Ultra high bypass Nacelle aerodynamics inlet flow-through high angle of attack distortion test

    NASA Technical Reports Server (NTRS)

    Larkin, Michael J.; Schweiger, Paul S.

    1992-01-01

    A flow-through inlet test program was conducted to evaluate inlet test methods and determine the impact of the fan on inlet separation when operating at large angles of attack. A total of 16 model configurations of approximately 1/6 scale were tested. A comparison of these flow-through results with powered data indicates the presence of the fan increased separation operation 3 degrees to 4 degrees over the flow through inlet. Rods and screens located at the fan face station, that redistribute the flow, achieved simulation of the powered-fan results for separation angle of attack. Concepts to reduce inlet distortion and increase angle of attack capability were also evaluated. Vortex generators located on the inlet surface increased inlet angle of attack capability up to 2 degrees and reduced inlet distortion in the separated region. Finally, a method of simulating the fan/inlet aerodynamic interaction using blockage sizing method has been defined. With this method, a static blockage device used with a flow-through model will approximate the same inlet onset of separation angle of attack and distortion pattern that would be obtained with an inlet model containing a powered fan.

  13. Space shuttle: Verification of transition reentry corridor at high angles of attack and determination of transition aerodynamic characteristics and subsonic aerodynamic characteristics at low angles of attack for the Boeing H-32 booster

    NASA Technical Reports Server (NTRS)

    Houser, J.; Johnson, L. J.; Oiye, M.; Runciman, W.

    1972-01-01

    Experimental aerodynamic investigations were made in a transonic wind tunnel on a 1/150-scale model of the Boeing H-32 space shuttle booster configuration. The purpose of the test was: (1) to verify the transonic reentry corridor at high angles of attack; (2) to determine the transonic aerodynamic characteristics; and (3) to determine the subsonic aerodynamic characteristics at low angles of attack. Test variables included configuration buildup, horizontal stabilizer settings of 0 and -20 deg, elevator deflections of 0 and -30 deg, and wing spoiler settings of 60 deg.

  14. Identification of an unsteady aerodynamic model up to high angle of attack regime

    NASA Astrophysics Data System (ADS)

    Fan, Yigang

    1997-12-01

    those from references, a state-space model is developed to describe the unsteady aerodynamic characteristics up to the high angle of attack regime. A nondimensional coordinate is introduced as the state variable describing the flow separation or vortex burst. First-order differential equation is used to govern the dynamics of flow separation or vortex bursting through this state variable. To be valid for general configurations, Taylor series expansions in terms of the input variables are used in the determination of aerodynamic characteristics, resembling the current approach of the stability derivatives. However, these derivatives are longer constant. They are dependent on the state variable of flow separation or vortex burst. In this way, the changes in stability derivatives with the angle of attack are included dynamically. The performance of the model is then validated by the wind-tunnel measurements of an NACA 0015 airfoil, a 70sp° delta wing and, finally two F-18 aircraft configurations. The results obtained show that within the framework of the proposed model, it is possible to obtain good agreement with different unsteady wind tunnel data in high angle-of-attack regime.

  15. Actuator and aerodynamic modeling for high-angle-of-attack aeroservoelasticity

    NASA Technical Reports Server (NTRS)

    Brenner, Martin J.

    1993-01-01

    Accurate prediction of airframe/actuation coupling is required by the imposing demands of modern flight control systems. In particular, for agility enhancement at high angle of attack and low dynamic pressure, structural integration characteristics such as hinge moments, effective actuator stiffness, and airframe/control surface damping can have a significant effect on stability predictions. Actuator responses are customarily represented with low-order transfer functions matched to actuator test data, and control surface stiffness is often modeled as a linear spring. The inclusion of the physical properties of actuation and its installation on the airframe is therefore addressed using detailed actuator models which consider the physical, electrical, and mechanical elements of actuation. The aeroservoelastic analysis procedure is described in which the actuators are modeled as detailed high-order transfer functions and as approximate low-order transfer functions. The impacts of unsteady aerodynamic modeling on aeroservoelastic stability are also investigated by varying the order of approximation, or number of aerodynamic lag states, in the analysis. Test data from a thrust-vectoring configuration of an F/A-l8 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.

  16. Aerodynamic surface distension system for high angle of attack forebody vortex control

    NASA Technical Reports Server (NTRS)

    Zell, Peter T. (Inventor)

    1994-01-01

    A deployable system is introduced for assisting flight control under certain flight conditions, such as at high angles of attack, whereby two inflatable membranes are located on the forebody portion of an aircraft on opposite sides thereof. The members form control surfaces for effecting lateral control forces if one is inflated and longitudinal control forces if both are inflated.

  17. Wing-Alone Aerodynamic Characteristics to High Angles of Attack at Subsonic and Transonic Speeds.

    DTIC Science & Technology

    1982-11-01

    indicators of symmetry since the wings were unbanked within the limits of tolerances and flow angularity. Longitudinal, spanwise, and vertical... unbanked wings at subsonic and transonic speeds from low to high angles of attack. The wing planforms varied in aspect ratio and taper ratio with

  18. Estimation of parameters involved in high angle-of-attack aerodynamic theory using spin flight test data

    NASA Technical Reports Server (NTRS)

    Taylor, L. W., Jr.; Pamadi, B. N.

    1983-01-01

    The difficulty in applying parameter estimation techniques to spinning airplanes is due in part to the unwieldy number of possible combinations of terms in the equations of motion, when the model structure is unknown. The combination of high angle of attack and high rotation rate results in aerodynamic functions which are quite complex. For wing dominated configurations it is advantageous to use aerodynamic theory to generate the model structure. In this way, the number of unknown parameters is reduced and the model accuracy may be increased. Under conditions for which the theory is inadequate, however, model accuracy may be reduced. Strip theory, for example, is incapable of predicting autorotative rolling moments indicated by wind tunnel tests at angles of attack exceeding 40 degrees. An improved aerodynamic theory would be necessary to successfully apply the technique advanced for such regions.

  19. Reentry vehicle aerodynamics and control at very high angle of attack

    NASA Astrophysics Data System (ADS)

    Merret, Jason Michael

    In recent flight tests the X-38 reentry test vehicle spins during the deployment of the drogue parachute. An experimental aerodynamic study has been conducted at the University of Illinois using a scale model of the X-38 to explore the cause of this problem. A six-component sting balance was used to measure the forces and moments on the 4.7% wind tunnel model at angles of attack from -7° to 95°. In addition, surface pressure taps and flow visualization techniques were utilized to determine the forebody pressures and surface flowfield on the model. The effect of Reynolds number and boundary-layer state were also examined. The investigation suggests that the spinning under the drogue parachute was caused by asymmetric vortex formation. At angles of attack between 75° and 90° vortex asymmetry developed in all of the cases without separation geometrically fixed. This flow asymmetry produced large side forces and yawing moments. The Reynolds number effect and the effect of the boundary-layer state were noticeable, but did not greatly change the side force and yawing moment characteristics of the model. The micro-geometry of the model had a large effect on the side force generated by the vortex positioning. The effects of forced oscillations were also examined and it was determined that the side forces were still present during the oscillations. Control of the vortices and side forces was obtained by applying strakes to the side of the forebody of the model.

  20. Experimental study of the effects of Reynolds number on high angle of attack aerodynamic characteristics of forebodies during rotary motion

    NASA Technical Reports Server (NTRS)

    Pauley, H.; Ralston, J.; Dickes, E.

    1995-01-01

    The National Aeronautics and Space Administration and the Defense Research Agency (United Kingdom) have ongoing experimental research programs in rotary-flow aerodynamics. A cooperative effort between the two agencies is currently underway to collect an extensive database for the development of high angle of attack computational methods to predict the effects of Reynolds number on the forebody flowfield at dynamic conditions, as well as to study the use of low Reynolds number data for the evaluation of high Reynolds number characteristics. Rotary balance experiments, including force and moment and surface pressure measurements, were conducted on circular and rectangular aftbodies with hemispherical and ogive noses at the Bedford and Farnborough wind tunnel facilities in the United Kingdom. The bodies were tested at 60 and 90 deg angle of attack for a wide range of Reynolds numbers in order to observe the effects of laminar, transitional, and turbulent flow separation on the forebody characteristics when rolling about the velocity vector.

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

    NASA Technical Reports Server (NTRS)

    Brandon, Jay M.; Shah, Gautam H.

    1990-01-01

    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.

  2. Validation of aerodynamic parameters at high angles of attack for RAE high incidence research models

    NASA Technical Reports Server (NTRS)

    Ross, A. Jean; Edwards, Geraldine F.; Klein, Vladislav; Batterson, James G.

    1987-01-01

    Two series of free-flight tests have been conducted for combat aircraft configuration research models in order to investigate flight behavior near departure conditions as well as to obtain response data from which aerodynamic characteristics can be derived. The structure of the mathematical model and values for the mathematical derivatives have been obtained through an analysis of the first series, using stepwise regression. The results thus obtained are the bases of the design of active control laws. Flight test results for a novel configuration are compared with predicted responses.

  3. Turbulence Effects on the High Angle of Attack Aerodynamics of a Vertically Launched Missile

    DTIC Science & Technology

    1988-06-01

    aerodynamic characteristics of the missile may vary greatly as it transitions through these operating speeds. All surface- to - air missiles ...representative of a cruciform tail - control missile with very low aspect ratio wings or dorsal fins. The model was fabricated from 6061 and 2024 aluminum alloy...Lilsewise, in the tail section, are four tail control fins which are fixed in postion at 00 incidence relative

  4. Critical evaluation of the unsteady aerodynamics approach to dynamic stability at high angles of attack

    NASA Technical Reports Server (NTRS)

    Hui, W. H.

    1985-01-01

    Bifurcation theory is used to analyze the nonlinear dynamic stability characteristics of an aircraft subject to single-degree-of-freedom. The requisite moment of the aerodynamic forces in the equations of motion is shown to be representable in a form equivalent to the response to finite amplitude oscillations. It is shown how this information can be deduced from the case of infinitesimal-amplitude oscillations. The bifurcation theory analysis reveals that when the bifurcation parameter is increased beyond a critical value at which the aerodynamic damping vanishes, new solutions representing finite amplitude periodic motions bifurcate from the previously stable steady motion. The sign of a simple criterion, cast in terms of aerodynamic properties, determines whether the bifurcating solutions are stable or unstable. For the pitching motion of flat-plate airfoils flying at supersonic/hypersonic speed and for oscillation of flaps at transonic speed, the bifurcation is subcritical, implying either the exchanges of stability between steady and periodic motion are accompanied by hysteresis phenomena, or that potentially large aperiodic departures from steady motion may develop.

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

    NASA Technical Reports Server (NTRS)

    Axelson, J. A.

    1977-01-01

    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.

  6. Prediction of static aerodynamic characteristics for slender bodies alone and with lifting surfaces to very high angles of attack

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1977-01-01

    An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. The angles of attack extend from 0 deg to 180 deg for M = 2.9 from 0 deg to 60 deg for M = 0.6 to 2.0. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 deg to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. Vapor-screen and oil-flow pictures are shown for many body, body-wing and body-wing-tail configurations. When spearation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).

  7. A study of prediction methods for the high angle-of-attack aerodynamics of straight wings and fighter aircraft

    NASA Technical Reports Server (NTRS)

    Mcmillan, O. J.; Mendenhall, M. R.; Perkins, S. C., Jr.

    1984-01-01

    Work is described dealing with two areas which are dominated by the nonlinear effects of vortex flows. The first area concerns the stall/spin characteristics of a general aviation wing with a modified leading edge. The second area concerns the high-angle-of-attack characteristics of high performance military aircraft. For each area, the governing phenomena are described as identified with the aid of existing experimental data. Existing analytical methods are reviewed, and the most promising method for each area used to perform some preliminary calculations. Based on these results, the strengths and weaknesses of the methods are defined, and research programs recommended to improve the methods as a result of better understanding of the flow mechanisms involved.

  8. Prediction of static aerodynamic characteristics for slender bodies alone and with lifting surfaces to very high angles of attack

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1976-01-01

    An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. However, with a decrease in Mach number to 0.6, the agreement for C sub N rapidly deteriorates, although the normal-force centers remain in close agreement. Vapor-screen and oil-flow pictures are shown for many body, body-wing, and body-wing-tail configurations. When separation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).

  9. A summary of the forebody high-angle-of-attack aerodynamics research on the F-18 and the X-29A aircraft

    NASA Technical Reports Server (NTRS)

    Bjarke, Lisa J.; Delfrate, John H.; Fisher, David F.

    1992-01-01

    High-angle-of-attack aerodynamic studies have been conducted on both the F18 High Alpha Research Vehicle (HARV) and the X-29A aircraft. Data obtained include on- and off-surface flow visualization and static pressure measurements on the forebody. Comparisons of similar results are made between the two aircraft where possible. The forebody shapes of the two aircraft are different and the X-29A forebody flow is affected by the addition of nose strakes and a flight test noseboom. The forebody flow field of the F-18 HARV is fairly symmetric at zero sideslip and has distinct, well-defined vortices. The X-29A forebody vortices are more diffuse and are sometimes asymmetric at zero sideslip. These asymmetries correlate with observed zero-sideslip aircraft yawing moments.

  10. Aerodynamic static stability characteristics, fin effectiveness, and fin location of the MSFC 33-foot pressure fed booster at high angles of attack

    NASA Technical Reports Server (NTRS)

    Hamilton, T.

    1972-01-01

    Experimental aerodynamic investigations were conducted in the NASA/MSFC 14 x 14 inch trisonic wind tunnel during early February 1972 on a 0.00280 scale model of the MSFC 33-foot diameter space shuttle pressure fed booster configuration. The basic configuration tested was a 40 deg cone/cylinder/7 deg frustrum designated the MSFC 33 foot pressure fed booster. Six component aerodynamic force and moment data were recorded over a Mach number range from 0.9 to 5.0 at angles of attack from +20 to +90 at 0 deg sideslip and over a sideslip range from -10 to +10 deg at 30, 50, 60 and 80 deg angles of attack. Primary configuration variables were fin deflection angle and vertical position on the frustrum.

  11. Forebody tangential blowing for control at high angles of attack

    NASA Technical Reports Server (NTRS)

    Kroo, I.; Rock, S.; Roberts, L.

    1991-01-01

    A feasibility study to determine if the use of tangential leading edge blowing over the forebody could produce effective and practical control of the F-18 HARV aircraft at high angles of attack was conducted. A simplified model of the F-18 configuration using a vortex-lattice model was developed to obtain a better understanding of basic aerodynamic coupling effects and the influence of forebody circulation on lifting surface behavior. The effect of tangential blowing was estimated using existing wind tunnel data on normal forebody blowing and analytical studies of tangential blowing over conical forebodies. Incorporation of forebody blowing into the flight control system was investigated by adding this additional yaw control and sideforce generating actuator into the existing F-18 HARV simulation model. A control law was synthesized using LQG design methods that would schedule blowing rates as a function of vehicle sideslip, angle of attack, and roll and yaw rates.

  12. A numerical analysis applied to high angle of attack three-dimensional inlets

    NASA Technical Reports Server (NTRS)

    Hwang, D. P.

    1986-01-01

    The three-dimensional analytical methods used to analyze subsonic high angle of attack inlets are described. The methods are shown to be in good agreement with experimental results for various three-dimensional high angle of attack inlets. The methods are used to predict aerodynamic characteristics of scarf and slotted-lip inlets.

  13. Supersonic aerodynamic characteristics of some reentry concepts for angles of attack to 90 deg

    NASA Astrophysics Data System (ADS)

    Spearman, M. L.

    1985-11-01

    Past studies of reentry vehicles tested to high angles of attack (up to 90 deg) in the Mach number range from 2 to 4.8 are reviewed. Two basic planforms are considered: highly-swept deltas and circular. The delta concepts include variations in cross section (and thus volume) and in camber distribution. The effectiveness of various types of aerodynamic control devices is also included. The purpose of the paper is to examine the characteristics of the vehicles with a view toward the potential usefulness of such concepts in a flight regime that would include reentry from space into the atmosphere followed by a transition to sustained atmospheric flight.

  14. Low speed aerodynamic characteristics of an 0.075-scale F-15 airplane model at high angles of attack and sideslip

    NASA Technical Reports Server (NTRS)

    Petroff, D. N.; Scher, S. H.; Cohen, L. E.

    1974-01-01

    An 0.075 scale model representative of the F-15 airplane was tested in the Ames 12 foot pressure wind tunnel at a Mach number of 0.16 to determine static longitudinal and lateral directional characteristics at spin attitudes for Reynolds numbers from 1.48 to 16.4 million per meter (0.45 to 5.0 million per foot). Angles of attack ranged from 0 to +90 deg and from -40 deg to -80 deg while angles of sideslip were varied from -20 deg to +30 deg. Data were obtained for nacelle inlet ramp angles of 0 to 11 deg with the left and right stabilators deflected 0, -25 deg, and differentially 5 deg and -5 deg. The normal pointed nose and two alternate nose shapes were also tested along with several configurations of external stores. Analysis of the results indicate that at higher Reynolds numbers there is a slightly greater tendency to spin inverted than at lower Reynolds numbers. Use of a hemispherical nose in place of the normal pointed nose provided an over correction in simulating yawing moment effects at high Reynolds numbers.

  15. X-31 high angle of attack control system performance

    NASA Technical Reports Server (NTRS)

    Huber, Peter; Seamount, Patricia

    1994-01-01

    The design goals for the X-31 flight control system were: (1) level 1 handling qualities during post-stall maneuvering (30 to 70 degrees angle-of-attack); (2) thrust vectoring to enhance performance across the flight envelope; and (3) adequate pitch-down authority at high angle-of-attack. Additional performance goals are discussed. A description of the flight control system is presented, highlighting flight control system features in the pitch and roll axes and X-31 thrust vectoring characteristics. The high angle-of-attack envelope clearance approach will be described, including a brief explanation of analysis techniques and tools. Also, problems encountered during envelope expansion will be discussed. This presentation emphasizes control system solutions to problems encountered in envelope expansion. An essentially 'care free' envelope was cleared for the close-in-combat demonstrator phase. High angle-of-attack flying qualities maneuvers are currently being flown and evaluated. These results are compared with pilot opinions expressed during the close-in-combat program and with results obtained from the F-18 HARV for identical maneuvers. The status and preliminary results of these tests are discussed.

  16. Design of Nonlinear Autopilots for High Angle of Attack Missiles

    DTIC Science & Technology

    1996-01-01

    Copyright 1996 by Optimal Synthesis . All Rights Reserved. 1 Design of Nonlinear Autopilots for High Angle of Attack Missiles By P. K. Menon* and...M. Yousefpor† Optimal Synthesis Palo Alto, CA 94306, USA Abstract Two nonlinear autopilot design approaches for a tail-controlled high angle of...Prescribed by ANSI Std Z39-18 © Copyright 1996 by Optimal Synthesis . All Rights Reserved. 2 better agility from tactical missiles. In air-to-air

  17. Pneumatic vortical flow control at high angles of attack

    NASA Technical Reports Server (NTRS)

    Tavella, Domingo A.; Schiff, Lewis B.; Cummings, Russell M.

    1990-01-01

    The injection of thin, high-momentum jets of air into the fuselage forebody boundary layers of the F-18 aircraft is explored numerically as a means of controlling the onset of fuselage vortices and of generating yaw control forces. The study was carried out for an angle of attack of 30 deg with symmetrical and asymmetrical blowing configurations. One-sided blowing results in a strongly asymmetrical flow pattern in the fore portion of the fuselage, leading to a net lateral force.

  18. Pneumatic vortical flow control at high angles of attack

    NASA Technical Reports Server (NTRS)

    Tavella, Domingo A.; Schiff, Lewis B.; Cummings, Russell M.

    1990-01-01

    The injection of thin, high-momentum jets of air into the fuselage forebody boundary layers of the F-18 aircraft is explored numerically as a means of controlling the onset of fuselage vortices and of generating yaw control forces. The study was carried out for an angle of attack of 30 deg with symmetrical and asymmetrical blowing configurations. One-sided blowing results in a strongly asymmetrical flow pattern in the fore portion of the fuselage, leading to a net lateral force.

  19. Low-speed aerodynamic characteristics of a 1/8-scale X-29A airplane model at high angles of attack and sideslip

    NASA Technical Reports Server (NTRS)

    Whipple, R. D.; Ricket, J. L.

    1986-01-01

    A 1/8-scale model of the X-29A airplane was tested in the Ames 12-Foot Pressure Wind Tunnel at a Mach number of 0.20 and Reynolds numbers of 0.13 x 10 to the 6th power to 2.00 x 10 to the 6th power based on a fuselage forebody depth of 0.4 ft, For the test series presented herein, the angle of attack ranged from 40 deg. to 90 deg. and the angle of sideslip ranged from -10 deg. to 30 deg. for the erect attitude. Tests with the model inverted covered angles of attack from -40 deg. to -90 deg. and angles of sideslip from -30 deg. to 10 deg. Data were obtained for the basic design and for several forebody strakes. An alternate forebody design was also tested. The results provided information for selection of forebody strakes for compensation of Reynolds number effect on the 1/25-scale free-spinning model tested in the Langley Spin Tunnel.

  20. A review of some Reynolds number effects related to bodies at high angles of attack

    NASA Technical Reports Server (NTRS)

    Polhamus, E. C.

    1984-01-01

    A review of some effects of Reynolds number on selected aerodynamic characteristics of two- and three-dimensional bodies of various cross sections in relation to fuselages at high angles of attack at subsonic and transonic speeds is presented. Emphasis is placed on the Reynolds number ranges above the subcritical and angles of attack where lee side vortex flow or unsteady wake type flows predominate. Lists of references, arranged in subject categories, are presented with emphasis on those which include data over a reasonable Reynolds number range. Selected Reynolds number data representative of various aerodynamic flows around bodies are presented and analyzed and some effects of these flows on fuselage aerodynamic parameters are discussed.

  1. Reconfigurable flight control for high angle of attack fighter aircraft, with wind tunnel study

    NASA Astrophysics Data System (ADS)

    Siddiqui, Bilal Ahmed

    In this work we studied Reconfigurable Flight Control Systems to achieve acceptable performance of a fighter aircraft, even in the event of wing damage to the aircraft at low speeds and high angle of attack, which is typical of many combat maneuvers. Equations of motion for the damaged aircraft were derived, which helped in building simulators. A new methodology combining experimental and numerical aerodynamic prediction was proposed and implemented. For this a wind-tunnel study of a similar configuration was carried out to study the aerodynamics at low speeds and high angle of attack. A baseline control system for undamaged aircraft was developed, and finally a reconfigurable flight control scheme was implemented to keep the aircraft flyable even after the damage.

  2. Wake vortex measurements of bodies at high angle of attack

    NASA Technical Reports Server (NTRS)

    Owen, F. K.; Johnson, D. A.

    1978-01-01

    Three-dimensional laser velocimeter measurements have been made of the wake vortices of a slender tangent-ogive body which had nose and body fineness ratios of 3.5 and 12, respectively. Data were obtained for an angle of attack to seminose angle ratio of 2.3 at a free-stream Mach number of 0.6 and unit Reynolds number of 2 million/ft. Details of the mean flow field are presented and features of the turbulent and unsteady nature of the vortex flow field are discussed. Problems associated with obtaining meaningful vortex measurements in high-speed flows are addressed.

  3. The F-18 High Alpha Research Vehicle: A High-Angle-of-Attack Testbed Aircraft

    NASA Technical Reports Server (NTRS)

    Regenie, Victoria; Gatlin, Donald; Kempel, Robert; Matheny, Neil

    1992-01-01

    The F-18 High Alpha Research Vehicle is the first thrust-vectoring testbed aircraft used to study the aerodynamics and maneuvering available in the poststall flight regime and to provide the data for validating ground prediction techniques. The aircraft includes a flexible research flight control system and full research instrumentation. The capability to control the vehicle at angles of attack up to 70 degrees is also included. This aircraft was modified by adding a pitch and yaw thrust-vectoring system. No significant problems occurred during the envelope expansion phase of the program. This aircraft has demonstrated excellent control in the wing rock region and increased rolling performance at high angles of attack. Initial pilot reports indicate that the increased capability is desirable although some difficulty in judging the size and timing of control inputs was observed. The aircraft, preflight ground testing and envelope expansion flight tests are described.

  4. Experimental Flight Characterization of Spin Stabilized Projectiles at High Angle of Attack

    DTIC Science & Technology

    2017-08-07

    of downrange travel ) is also evident in the horizontal data. Fig. 3 Center-of-gravity motion The rolling motion is captured in Fig. 4. These...data are strongly linear with distance travelled because the launch spin rate is around 90 Hz and roll decelerates slightly due to aerodynamic damping...This vehicle rolls about 70 times over 200 m of downrange travel . For some of the high-angle-of-attack flights a significant amount of roll data

  5. Reynolds Number Effects at High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Fisher, David F.; Cobleigh, Brent R.; Banks, Daniel W.; Hall, Robert M.; Wahls, Richard A.

    1998-01-01

    Lessons learned from comparisons between ground-based tests and flight measurements for the high-angle-of-attack programs on the F-18 High Alpha Research Vehicle (HARV), the X-29 forward-swept wing aircraft, and the X-31 enhanced fighter maneuverability aircraft are presented. On all three vehicles, Reynolds number effects were evident on the forebodies at high angles of attack. The correlation between flight and wind tunnel forebody pressure distributions for the F-18 HARV were improved by using twin longitudinal grit strips on the forebody of the wind-tunnel model. Pressure distributions obtained on the X-29 wind-tunnel model at flight Reynolds numbers showed excellent correlation with the flight data up to alpha = 50 deg. Above (alpha = 50 deg. the pressure distributions for both flight and wind tunnel became asymmetric and showed poorer agreement, possibly because of the different surface finish of the model and aircraft. The detrimental effect of a very sharp nose apex was demonstrated on the X-31 aircraft. Grit strips on the forebody of the X-31 reduced the randomness but increased the magnitude of the asymmetry. Nose strakes were required to reduce the forebody yawing moment asymmetries and the grit strips on the flight test noseboom improved the aircraft handling qualities.

  6. Unsteady aerodynamics of missiles. Part 3: Determination of the longitudinal stability of wings at high angles of attack in supersonic flight

    NASA Astrophysics Data System (ADS)

    Schneider, C. P.

    1980-05-01

    A theoretical method for the determination of unsteady aerodynamic coefficients associated with the longitudinal stability of slender wings in supersonic flight is presented. It is based on the indicial functional theory of Tobak. Extension to higher incidences is effected by combining the indicial functions with steady nonlinear coefficients derived from a semiempiricial procedure. The unsteady nonlinear aerodynamic coefficients are determined for delta wings with subsonic and supersonic leading edges, respectively.

  7. Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

    NASA Technical Reports Server (NTRS)

    Spahr, J. R.

    1954-01-01

    The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance

  8. Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2

    NASA Technical Reports Server (NTRS)

    Spahr, J Richard

    1954-01-01

    The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing- tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift - interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the

  9. The transient roll moment response due to forebody tangential blowing at high angles of attack

    NASA Astrophysics Data System (ADS)

    Chow, Jonathan Kwokching

    The sustained ability for controlled flight at high angles of attack is desirable for future aircraft. For combat aircraft, enhancing maneuverability is important to increasing its survivability. For future supersonic commercial aircraft, an increase in lift at high angles of attack leads to improved performance during take-offs and landing, and a reduction in noise pollution. However, nonlinear and unsteady phenomena, such as flow separation and vortex shedding dominate the aerodynamics in the high angle of attack regime. These phenomena cause the onset of lateral loads and decrease the effectiveness of conventional control surfaces. For conventional aircraft, controlled flight at high angle of attack is difficult or unfeasible without augmented means of control and a good understanding of their impact on vehicle characteristics and dynamics. The injection of thin sheets of air tangentially to the forebody of the vehicle has been found to be an extremely promising method for augmenting the control of a flight vehicle at high angles of attack. Forebody Tangential Blowing (FTB) allows the flow structure to be altered in a rational manner and increase the controllability of the vehicle under these flight conditions. The feasibility of using FTB to control the roll-yaw motion of flight vehicles has been demonstrated. Existing knowledge of FTB's nonlinear impact on the aerodynamic moment responses is limited. Currently available dynamic models predict the general trends in the behavior but do not capture important transient effects that dominate the responses when small amounts of blowing is used. These transients can be large in comparison to the steady-state values. This thesis summarizes the experimental and theoretical results of an investigation into the transient effects of Forebody Tangential Blowing. The relationship between the aerodynamic roll moment, vortical flowfield, and blowing strength is examined to obtain a fundamental understanding of the physics of

  10. Trim angle of attack of flexible wings using non-linear aerodynamics

    NASA Astrophysics Data System (ADS)

    Cohen, David Erik

    Multidisciplinary interactions are expected to play a significant role in the design of future high-performance aircraft (Blended-Wing Body, Truss-Braced wing; High Speed Civil transport, High-Altitude Long Endurance aircraft and future military aircraft). Also, the availability of supercomputers has made it now possible to employ high-fidelity models (Computational Fluid Dynamics for fluids and detailed finite element models for structures) at the preliminary design stage. A necessary step at that stage is to calculate the wing angle-of-attack at which the wing will generate the desired lift for the specific flight maneuver. Determination of this angle, a simple affair when the wing is rigid and the flow regime linear, becomes difficult when the wing is flexible and the flow regime non-linear. To solve this inherently nonlinear problem, a Newton's method type algorithm is developed to simultaneously calculate the deflection and the angle of attack. The developed algorithm is tested for a wing, used for in-house aeroelasticity research at Boeing (previously McDonnell Douglas) Long Beach. The trim angle of attack is calculated for a range of desired lift values. In addition to the Newton's method algorithm, a non derivative method (NDM) based on fixed point iteration, typical of fixed angle of attack calculations in aeroelasticity, is employed. The NDM, which has been extended to be able to calculate trim angle of attack, is used for one of the cases. The Newton's method calculation converges in fewer iterations, but requires more CPU time than the NDM method. The NDM, however, results in a slightly different value of the trim angle of attack. It should be noted that NDM will converge in a larger number of iterations as the dynamic pressure increases. For one value of the desired lift, both viscous and inviscid results were generated. The use of the inviscid flow model while not resulting in a markedly different value for the trim angle of attack, does result in a

  11. Space shuttle: High angle of attack transition and low angle of attack launch phase aerodynamic stability and control of GD/C B-18E-2, B-18E-3 delta wing booster, and launch configuration of MSC-040A orbiter and twin pressure fed boosters

    NASA Technical Reports Server (NTRS)

    Debevoise, J. M.; Mcginnis, R. F.

    1972-01-01

    The test was a conventional stability and control test except for two aspects. One was the very high angles of attack at which the delta wing configurations were tested (up to 60 degrees) at Mach numbers of 3 and 4.96. The other was the installation of the orbiter and twin boosters in a manner that caused the support system to induce normal forces and side forces on the aft portion of the boosters at all Mach numbers; i.e., the support and the booster bodies were close together, side by side.

  12. Inlet Distortion for an F/A-18A Aircraft During Steady Aerodynamic Conditions up to 60 deg Angle of Attack

    NASA Technical Reports Server (NTRS)

    Walsh, Kevin R.; Yuhas, Andrew J.; Williams, John G.; Steenken, William G.

    1997-01-01

    The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This paper addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from -10 deg to 60 deg angle of attack and from -8 deg to 11 deg angle of sideslip at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to establish a foundation for understanding inlet characteristics at high angle of attack.

  13. Full-scale high angle-of-attack tests of an F/A-18

    NASA Technical Reports Server (NTRS)

    Meyn, Larry A.; Lanser, Wendy R.; James, Kevin D.

    1992-01-01

    This paper presents an overview of high angle-of-attack tests of a full-scale F/A-18 in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California. A production aircraft was tested over an angle-of-attack range of 18 to 50 deg and at wind speeds of up to 100 knots. These tests had three primary test objectives. Pneumatic and mechanical forebody flow control devices were tested at full-scale and shown to produce significant yawing moments for lateral control of the aircraft at high angles of attack. Mass flow requirements for the pneumatic system were found to scale with freestream density and speed rather than freestream dynamic pressure. Detailed measurements of the pressures buffeting the vertical tail were made and spatial variations in the buffeting frequency were found. The LEX fence was found to have a significant effect on the frequency distribution on the outboard surface of the vertical fin. In addition to the above measurements, an extensive set of data was acquired for the validation of computational fluid dynamics codes and for comparison with flight test and small-scale wind tunnel test results.

  14. Aerodynamic Behavior at One Revolution Angle of Attack of Two-Dimensional Wings

    NASA Astrophysics Data System (ADS)

    Han, Yong; Lee, Eun; Kim, Jeong; Shin, Yong

    2011-11-01

    In order to investigate aerodynamic behaviors at extreme angles of attack beyond the normal static stall angle and in the reversed flow, lift and drag have been measured at one revolution angles of attack by rotating the wing around the 1/4 chord with use of a dynamic balance in the low speed wind tunnel. Three different geometries of wing section; a flat plate, a symmetric airfoil, NACA0018, and a cambered airfoil, Goe222, were selected for these experiments. It was turned out that the lift coefficient maintains substantially even beyond the traditional stall AoA of the wing. Drag coefficients of these wings showed sinusoidal profiles, and polar plots of Cl versus Cd provided distinctive behaviors unseen in the calculation by the classical wing theory. Application of the cyclic aerodynamic characteristics to a vertical axis wind turbine and wake characteristics around the critical angle will be displayed. This work was supported by Cooperative R&D program between Industry, Academy, and Research Institute funded Korea Small and Medium Business Administration in 2011.

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

    NASA Technical Reports Server (NTRS)

    Petroff, D. N.; Scher, S. H.; Sutton, C. E.

    1978-01-01

    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.

  16. High Angle-of-Attack Aerodynamics

    DTIC Science & Technology

    1982-12-01

    of oil-streak lines toward a particular line. Whether this is also a sufficient con- dition is a matter of current debate. Of the many attempts to...ON THE SAAB-VIGGEN CANARD-WING AIRCRAFT AT 30-DEGREE AGA INORTHROP WATER TUNNEL) I ins viiribls flowus interactin al-5 occrs’r’ betwieens 41 ligslY...lead, with the missile mass in mind , to structures with minimal structural stiffness. Consequently, effects of aeroelasticity may play a role. Due to

  17. Use of Piloted Simulation for High-Angle-of-Attack Agility Research and Design Criteria Development

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Foster, John V.; Hoffler, Keith D.

    1991-01-01

    The use of piloted simulation at Langley Research Center as part of the NASA High-Angle-of-Attack Technology Program (HATP) to provide methods and concepts for the design of advanced fighter aircraft is discussed. A major focus is to develop the design process required to fully exploit the benefits from advanced control concepts for high-angle-of attack agility.

  18. High angle of attack flying qualities criteria for longitudinal rate command systems

    NASA Technical Reports Server (NTRS)

    Wilson, David J.; Citurs, Kevin D.; Davidson, John B.

    1994-01-01

    This study was designed to investigate flying qualities requirements of alternate pitch command systems for fighter aircraft at high angle of attack. Flying qualities design guidelines have already been developed for angle of attack command systems at 30, 45, and 60 degrees angle of attack, so this research fills a similar need for rate command systems. Flying qualities tasks that require post-stall maneuvering were tested during piloted simulations in the McDonnell Douglas Aerospace Manned Air Combat Simulation facility. A generic fighter aircraft model was used to test angle of attack rate and pitch rate command systems for longitudinal gross acquisition and tracking tasks at high angle of attack. A wide range of longitudinal dynamic variations were tested at 30, 45, and 60 degrees angle of attack. Pilot comments, Cooper-Harper ratings, and pilot induced oscillation ratings were taken from five pilots from NASA, USN, CAF, and McDonnell Douglas Aerospace. This data was used to form longitudinal design guidelines for rate command systems at high angle of attack. These criteria provide control law design guidance for fighter aircraft at high angle of attack, low speed flight conditions. Additional time history analyses were conducted using the longitudinal gross acquisition data to look at potential agility measures of merit and correlate agility usage to flying qualities boundaries. This paper presents an overview of this research.

  19. An experimental study of an airfoil with a bio-inspired leading edge device at high angles of attack

    NASA Astrophysics Data System (ADS)

    Mandadzhiev, Boris A.; Lynch, Michael K.; Chamorro, Leonardo P.; Wissa, Aimy A.

    2017-09-01

    Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. Deployable aerodynamic surfaces from the wing leading or trailing edges are often used to extend the aerodynamic envelope (e.g. slats and flaps). Birds have also evolved feathers at the leading edge (LE) of their wings, known as the alula, which enables them to perform high angles of attack maneuvers. In this study, a series of wind tunnel experiments are performed to quantify the effect of various deployment parameters of an alula-like LE device on the aerodynamic performance of a cambered airfoil (S1223) at stall and post stall conditions. The alula relative angle of attack, measured from the mean chord of the airfoil, is varied to modulate tip-vortex strength, while the alula deflection angle is varied to modulate the distance between the tip vortex and the wing surface. Integrated lift force measurements were collected at various alula-inspired device configurations. The effect of the alula-inspired device on the boundary layer velocity profile and turbulence intensity were investigated through hot-wire anemometer measurements. Results show that as alula deflection angle increases, the lift coefficient also increase especially at lower alula relative angles of attack. Moreover, at post stall wing angles of attack, the wake velocity deficit is reduced in the presence of alula device, confirming the mitigation of the wing adverse pressure gradient. The results are in strong agreement with measurements taken on bird wings showing delayed flow reversal and extended range of operational angles of attack. An engineered alula-inspired device has the potential to improve mission adaptability in small unmanned air vehicles during low Reynolds number flight.

  20. Flight test of the X-29A at high angle of attack: Flight dynamics and controls

    NASA Technical Reports Server (NTRS)

    Bauer, Jeffrey E.; Clarke, Robert; Burken, John J.

    1995-01-01

    The NASA Dryden Flight Research Center has flight tested two X-29A aircraft at low and high angles of attack. The high-angle-of-attack tests evaluate the feasibility of integrated X-29A technologies. More specific objectives focus on evaluating the high-angle-of-attack flying qualities, defining multiaxis controllability limits, and determining the maximum pitch-pointing capability. A pilot-selectable gain system allows examination of tradeoffs in airplane stability and maneuverability. Basic fighter maneuvers provide qualitative evaluation. Bank angle captures permit qualitative data analysis. This paper discusses the design goals and approach for high-angle-of-attack control laws and provides results from the envelope expansion and handling qualities testing at intermediate angles of attack. Comparisons of the flight test results to the predictions are made where appropriate. The pitch rate command structure of the longitudinal control system is shown to be a valid design for high-angle-of-attack control laws. Flight test results show that wing rock amplitude was overpredicted and aileron and rudder effectiveness were underpredicted. Flight tests show the X-29A airplane to be a good aircraft up to 40 deg angle of attack.

  1. Investigation of Flying Qualities of Military Aircraft at High Angles of Attack. Volume 2. Appendices

    DTIC Science & Technology

    1974-06-01

    IWP^P^WW»» .1 9« »iitfUumiimm^mw^mMini^ ’mi AD-A015 830 INVESTIGATION OF FLYING QUALITIES OF MILITARY AIRCRAFT AT HIGH ANGLES OF ATTACK ...l l I ITTI \\ AFFDLre-7*61 \\- 296063 e CO 00 INVESTIGATION OF FLYING QUALITIES OF tO MILITARY AIRCRAFT AT HIGH ANGLES OF ^ ATTACK ^ Volume II...high Angles of Attack , Vol. II: Appendices 1. *UTMORf.J Donald E. Johnston Jeffrey R. Hogge Gary L. Teper ». PERFORMING ORGANIZATION NAME

  2. The equivalent angle-of-attack method for estimating the nonlinear aerodynamic characteristics of missile wings and control surfaces

    NASA Technical Reports Server (NTRS)

    Hemsch, M. J.; Nielsen, J. N.

    1982-01-01

    A method has been developed for estimating the nonlinear aerodynamic characteristics of missile wing and control surfaces. The method is based on the following assumption: if a fin on a body has the same normal-force coefficient as a wing alone composed of two of the same fins joined together at their root chords, then the other force and moment coefficients of the fin and the wing alone are the same including the nonlinearities. The method can be used for deflected fins at arbitrary bank angles and at high angles of attack. In the paper, a full derivation of the method is given, its accuracy demonstrated and its use in extending missile data bases is shown.

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

    NASA Technical Reports Server (NTRS)

    Murman, Scott M.

    1994-01-01

    As part of NASA's High Alpha Technology Program, research has been aimed at developing and extending numerical methods to accurately predict the high Reynolds number flow about the NASA F-18 High Alpha Research Vehicle (HARV) at large angles of attack. The HARV aircraft is equipped with a bidirectional thrust vectoring unit which enables stable, controlled flight through 70 deg angle of attack. Currently, high-fidelity numerical solutions for the flow about the HARV have been obtained at alpha = 30 deg, and validated against flight-test data. It is planned to simulate the flow about the HARV through alpha = 60 deg, and obtain solutions of the same quality as those at the lower angles of attack. This report presents the status of work aimed at extending the HARV computations to the extreme angle of attack range.

  4. Comparison of X-31 Flight and Ground-Based Yawing Moment Asymmetries at High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Cobleigh, Brent R.; Croom, Mark A.

    2001-01-01

    Significant yawing moment asymmetries were encountered during the high-angle-of-attack envelope expansion of the two X-31 aircraft. These asymmetries caused position saturations of the thrust-vectoring vanes and trailing-edge flaps during some stability-axis rolling maneuvers at high angles of attack. The two test aircraft had different asymmetry characteristics, and ship two has asymmetries that vary as a function of Reynolds number. Several aerodynamic modifications have been made to the X-31 forebody with the goal of minimizing the asymmetry. These modifications include adding transition strips on the forebody and noseboom, using two different length strakes, and increasing nose bluntness. Ultimately, a combination of forebody strakes, nose blunting, and noseboom transition strips reduced the yawing moment asymmetry enough to fully expand the high-angle-of-attack envelope. Analysis of the X-31 flight data is reviewed and compared to wind-tunnel and water-tunnel measurements. Several lessons learned are outlined regarding high-angle-of-attack configuration design and ground testing.

  5. Development of a pneumatic high-angle-of-attack flush airdata sensing system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.

    1991-01-01

    A nonintrusive high-angle-of-attack flush airdata sensing system was installed and flight tested in the F-18 High Alpha Research Vehicle. This system consists of a matrix of 25 pressure orifices arranged in concentric circles on the nose of the vehicle to determine angles of attack and sideslip, Mach number, and pressure altitude. During the course of the flight tests, it was determined that satisfactory results could be achieved using a subset of just nine ports.

  6. Development of a pneumatic high-angle-of-attack flush airdata sensing (HI-FADS) system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.

    1991-01-01

    A nonintrusive high-angle-of-attack flush airdata sensing system was installed and flight tested in the F-18 High Alpha Research Vehicle. This system consists of a matrix of 25 pressure orifices arranged in concentric circles on the nose of the vehicle to determine angles of attack and sideslip, Mach number, and pressure altitude. During the course of the flight tests, it was determined that satisfactory results could be achieved using a subset of just nine ports.

  7. Development of a pneumatic high-angle-of-attack flush airdata sensing system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.

    1991-01-01

    A nonintrusive high-angle-of-attack flush airdata sensing system was installed and flight tested in the F-18 High Alpha Research Vehicle. This system consists of a matrix of 25 pressure orifices arranged in concentric circles on the nose of the vehicle to determine angles of attack and sideslip, Mach number, and pressure altitude. During the course of the flight tests, it was determined that satisfactory results could be achieved using a subset of just nine ports.

  8. Use of piloted simulation for high-angle-of-attack agility research and design criteria development

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Foster, John V.; Hoffler, Keith D.

    1991-01-01

    The application of piloted simulations in the development of advanced fighter aircraft is reviewed in the context of the NASA High-Angle-of-Attack Technology Program (HATP). The HATP combines wind-tunnel experiments, computational aerodynamics, piloted simulations, and flight tests on a modified F-18 testbed aircraft and utilizes the experience and facilities of several NASA centers. Consideration is given to the role of simulation in the overall research process, simulation capabilities and software requirements, simulation flexibility and fidelity, evaluation maneuvers, the role of simulator pilots in evaluations, the analysis of simulation results, flight validation of maneuvers and rating approaches, and the use of simulations to define design criteria. Extensive diagrams, graphs, and flow charts are included.

  9. Evaluation of importance of lateral acceleration derivatives in extraction of lateral-directional derivatives at high angles of attack

    NASA Technical Reports Server (NTRS)

    Nguyen, L. T.

    1974-01-01

    A theoretical investigation was conducted to determine the importance of the lateral acceleration (beta) derivatives in the extraction of lateral-directional stability derivatives for swept wing airplanes at high angles of attack. Representative values of lateral acceleration derivatives in yaw and roll (Cn beta and Cl beta) were used in a computer program to generate representative flight motions at several angles of attack and altitudes. The computer-generated motions were then subjected to a parameter identification process based on a modified Newton-Raphson method. Two identification techniques were evaluated, one which included the beta derivatives and one which neglected them. The results of the study indicate that omission of the beta derivatives from mathematical models used in the derivative-extraction techniques can produce erroneous values for the lateral-directional stability derivatives particularly at high angles of attack, where the beta derivatives are large. The largest errors occur in the dynamic derivatives, but large errors may also occur in the static derivatives for cases in which the beta derivatives have large effects on the flight motions of the airplane. In addition, the resulting identified mathematical models provide poor motion prediction as well as erroneous predictions of dynamic modal characteristics. These results strongly indicate that the effects of beta derivatives should be considered in any attempt to extract lateral-directional aerodynamic parameters at high angles of attack.

  10. Simulator study of the effectiveness of an automatic control system designed to improve the high-angle-of-attack characteristics of a fighter airplane

    NASA Technical Reports Server (NTRS)

    Gilbert, W. P.; Nguyen, L. T.; Vangunst, R. W.

    1976-01-01

    A piloted, fixed-base simulation was conducted to study the effectiveness of some automatic control system features designed to improve the stability and control characteristics of fighter airplanes at high angles of attack. These features include an angle-of-attack limiter, a normal-acceleration limiter, an aileron-rudder interconnect, and a stability-axis yaw damper. The study was based on a current lightweight fighter prototype. The aerodynamic data used in the simulation were measured on a 0.15-scale model at low Reynolds number and low subsonic Mach number. The simulation was conducted on the Langley differential maneuvering simulator, and the evaluation involved representative combat maneuvering. Results of the investigation show the fully augmented airplane to be quite stable and maneuverable throughout the operational angle-of-attack range. The angle-of-attack/normal-acceleration limiting feature of the pitch control system is found to be a necessity to avoid angle-of-attack excursions at high angles of attack. The aileron-rudder interconnect system is shown to be very effective in making the airplane departure resistant while the stability-axis yaw damper provided improved high-angle-of-attack roll performance with a minimum of sideslip excursions.

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

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Wang, Kon-Sheng Charles Wang

    1996-01-01

    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.

  12. Prediction of static aerodynamic characteristics for space-shuttle-like and other bodies at angles of attack from 0 deg to 180 deg

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1973-01-01

    An engineering-type procedure is presented for computing normal-force, axial-force, and pitching-moment coefficients for bodies at angles of attack from 0 deg to 180 deg. The procedure is ideally suited for estimating the aerodynamic characteristics of space shuttle booster-like bodies because of the wide range of angles of attack, Mach numbers, and Reynolds numbers to be encountered. The analytical formulas, plots, and references given are also applicable for shuttle orbiter, missile, and aircraft-like bodies of both circular and noncircular cross section. The method for computing normal-force and pitching-moment coefficients is based upon the original proposal of Allen that the crossflow or lift distribution over a body can be expressed as the sum of slender-body potential term and an empirical viscous crossflow term. Although experimental data from which to verify the procedure at very high angles of attack are extremely limited, the comparisons made thus far of computed with experimental results are good. In this report the procedure has been shown to be capable of predicting reasonably well the experimental variation of C sub N, C sub A, C sub m, and x sub ac/l with angle of attack for nine bodies of revolution at a free-stream Mach number of 2.86.

  13. Effect of thermochemical non-equilibrium on the aerodynamics of an osculating-cone waverider under different angles of attack

    NASA Astrophysics Data System (ADS)

    Liu, Zhen; Liu, Jun; Ding, Feng; Li, Kai; Xia, Zhixun

    2017-10-01

    In order to research the effect of thermochemical non-equilibrium on the aerodynamics of an osculating-cone waverider, thermochemical non-equilibrium flow and perfect gas model are employed to study the aerodynamics of an osculating-cone waverider under different angles of attack. The obtained results show that the slope of the oblique shock wave has little difference when considering the thermochemical non-equilibrium effect under the condition of zero angle of attack. However, under the condition of other attack angles, the slope of the oblique shock wave diminishes when considering the thermochemical non-equilibrium effect. Furthermore, the non-equilibrium effect moves the pressure center of the osculating-cone waverider forward by as much as 1.53% of the whole craft's length, which must be taken into consideration in the balance design of aircraft.

  14. Roll-yaw control at high angle of attack by forebody tangential blowing

    NASA Technical Reports Server (NTRS)

    Pedreiro, N.; Rock, S. M.; Celik, Z. Z.; Roberts, L.

    1995-01-01

    The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.

  15. Roll-Yaw control at high angle of attack by forebody tangential blowing

    NASA Technical Reports Server (NTRS)

    Pedreiro, N.; Rock, S. M.; Celik, Z. Z.; Roberts, L.

    1995-01-01

    The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.

  16. High-angle-of-attack pneumatic lag and upwash corrections for a hemispherical flow direction sensor

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Heeg, Jennifer; Larson, Terry J.; Ehernberger, L. J.; Hagen, Floyd W.; Deleo, Richard V.

    1987-01-01

    As part of the NASA F-14 high angle of attack flight test program, a nose mounted hemispherical flow direction sensor was calibrated against a fuselage mounted movable vane flow angle sensor. Significant discrepancies were found to exist in the angle of attack measurements. A two fold approach taken to resolve these discrepancies during subsonic flight is described. First, the sensing integrity of the isolated hemispherical sensor is established by wind tunnel data extending to an angle of attack of 60 deg. Second, two probable causes for the discrepancies, pneumatic lag and upwash, are examined. Methods of identifying and compensating for lag and upwash are presented. The wind tunnel data verify that the isolated hemispherical sensor is sufficiently accurate for static conditions with angles of attack up to 60 deg and angles of sideslip up to 30 deg. Analysis of flight data for two high angle of attack maneuvers establishes that pneumatic lag and upwash are highly correlated with the discrepancies between the hemispherical and vane type sensor measurements.

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

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Ashbury, Scott C.; Deere, Karen A.

    1996-01-01

    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.

  18. Magnus effects at high angles of attack and critical Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Ringel, M.

    1983-01-01

    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.

  19. Magnus effects at high angles of attack and critical Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Ringel, M.

    1983-01-01

    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.

  20. Side-force alleviation on slender, pointed forebodies at high angles of attack

    NASA Technical Reports Server (NTRS)

    Rao, D. M.

    1978-01-01

    A new device was proposed for alleviating high angle-of-attack side force on slender, pointed forebodies. A symmetrical pair of separation strips in the form of helical ridges are applied to the forebody to disrupt the primary lee-side vortices and thereby avoid the instability that produces vortex asymmetry. Preliminary wind tunnel tests at Mach 0.3 and Reynolds no. 5,250,000 on a variety of forebody configurations and on a wing-body combination at angles of attack up to 56 degrees, demonstrated the effectiveness of the device.

  1. Aerodynamic characteristics of two-dimensional wing configurations at angles of attack near -90 deg

    NASA Technical Reports Server (NTRS)

    Maisel, Martin; Laub, Georgene; Mccroskey, W. J.

    1986-01-01

    Wind tunnel tests were conducted to determine the drag of two-dimensional wing sections operating in a near-vertical flow condition. Various leading- and trialing-edge configurations, including plain flaps of 25, 30, and 35% chord were tested at angles of attack from -75 to -105 deg. Reynolds numbers examined ranged from approximately 0.6 x 10 to the 6th power to 1.4 x 10 to the 6th power. The data were obtained using a wind tunnel force and moment balance system and arrays of chordwise pressure orifices. The results showed that significant reductions in drag, beyond what would be expected by virtue of the decreased frontal area, were obtainable with geometries that delayed flow separation. Rapid changes in drag with angle of attack were noted for many configurations. The results, however, were fairly insensitive to Reynolds number variations. Drag values computed from the pressure data generally agreed with the force data within 2%.

  2. X-29 High Angle-of-Attack Flying Qualities

    DTIC Science & Technology

    1991-06-01

    flhit and closure rate. Limited military utility tesIs ith thw variable gain capability. Predicted large amplitude increased roll-rate .apability... Fiscal Year Test Organization (PTO). The program consisted of an 1991 for resumption of high AOA flight testing. The intcgrated test team of AFFTC

  3. Reentry aerodynamic characteristics of a space shuttle solid rocket booster (MSFC model 454) at high angles of attack and high Mach number in the NASA/Langley four-foot unitary plan wind tunnel (SA25F)

    NASA Technical Reports Server (NTRS)

    Johnson, J. D.; Braddock, W. F.

    1975-01-01

    A force test of a 2.112 percent scale Space Shuttle Solid Rocket Booster (SRB), MSFC Model 454, was conducted in test section no. 2 of the Unitary Plan Wind Tunnel. Sixteen (16) runs (pitch polars) were performed over an angle of attack range from 144 through 179 degrees. Test Mach numbers were 2.30, 2.70, 2.96, 3.48, 4.00 and 4.63. The first three Mach numbers had a test Reynolds number of 1.5 million per foot. The remaining three were at 2.0 million per foot. The model was tested in the following configurations: (1) SRB without external protuberances, and (2) SRB with an electrical tunnel and a SRB/ET thrust attachment structure. Schlieren photographs were taken during the testing of the first configuration. The second configuration was tested at roll angles of 45, 90, and 135 degrees.

  4. The Aerodynamics of Axisymmetric Blunt Bodies Flying at Angle of Attack

    NASA Technical Reports Server (NTRS)

    Schoenenberger, Mark; Kutty, Prasad; Queen, Eric; Karlgaard, Chris

    2014-01-01

    The Mars Science Laboratory entry capsule is used as an example to demonstrate how a blunt body of revolution must be treated as asymmetric in some respects when flying at a non-zero trim angle of attack. A brief description of the axisymmetric moment equations are provided before solving a system of equations describing the lateral-directional moment equations for a blunt body trimming at an angle of attack. Simplifying assumptions are made which allow the solution to the equations to be rearranged to relate the roll and yaw stability with sideslip angle to the frequency of oscillation of the vehicle body rates. The equations show that for a blunt body the roll and yaw rates are in phase and proportional to each other. The ratio of the rates is determined by the static stability coefficients and mass properties about those axes. A trajectory simulation is used to validate the static yaw stability parameter identification equation and a simple method of identifying the oscillation frequency from the body rates. The approach is shown to successfully extract the modeled yaw stability coefficient along a simulated Mars entry in agreement with data earlier analysis of MSL flight data.

  5. Investigation of the Aerodynamic Characteristics of a Model Wing-Propeller Combination and of the Wing and Propeller Separately at Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Kuhn, Richard E; Draper, John W

    1956-01-01

    This report presents the results of an investigation conducted in the Langley 300 mph 7- by 10-foot wind tunnel for the purpose of determining the aerodynamic characteristics of a model wing-propeller combination, and of the wing and propeller separately at angles of attack up to 90 degrees. The tests covered thrust coefficients corresponding to free-stream velocities from zero forward speed to the normal range of cruising speeds. The results indicate that increasing the thrust coefficient increases the angle of attack for maximum lift and greatly diminishes the usual reduction in lift above the angle of attack for maximum lift.

  6. Measurement of Aerodynamic Forces for Various Mean Angles of Attack on an Airfoil Oscillating in Pitch and on Two Finite-span Wings Oscillating in Bending with Emphasis on Damping in the Stall

    NASA Technical Reports Server (NTRS)

    Rainey, A Gerald

    1957-01-01

    The oscillating air forces on a two-dimensional wing oscillating in pitch about the midchord have been measured at various mean angles of attack and at Mach numbers of 0.35 and 0.7. The magnitudes of normal-force and pitching-moment coefficients were much higher at high angles of attack than at low angles of attack for some conditions. Large regions of negative damping in pitch were found, and it was shown that the effect of increasing the Mach number 0.35 to 0.7 was to decrease the initial angle of attack at which negative damping occurred. Measurements of the aerodynamic damping of a 10-percent-thick and of a 3-percent-thick finite-span wing oscillating in the first bending mode indicate no regions of negative damping for this type of motion over the range of variables covered. The damping measured at high angles of attack was generally larger than that at low angles of attack. (author)

  7. Evaluation of Gritting Strategies for High Angle of Attack Using Wind Tunnel and Flight Test Data for the F/A-18

    NASA Technical Reports Server (NTRS)

    Hall, Robert M.; Erickson, Gary E.; Fox, Charles H., Jr.; Banks, Daniel W.; Fisher, David F.

    1998-01-01

    A subsonic study of high-angle-of-attack gritting strategies was undertaken with a 0.06-scale model of the F/A-18, which was assumed to be typical of airplanes with smooth-sided forebodies. This study was conducted in the Langley 7- by 10-Foot High-Speed Tunnel and was intended to more accurately simulate flight boundary layer characteristics on the model in the wind tunnel than would be possible by using classical, low-angle-of-attack gritting on the fuselage. Six-component force and moment data were taken with an internally mounted strain-gauge balance, while pressure data were acquired by using electronically scanned pressure transducers. Data were taken at zero sideslip over an angle-of-attack range from 0 deg to 40 deg and, at selected angles of attack, over sideslip angles from -10 deg to 10 deg. Free-stream Mach number was fixed at 0.30, which resulted in a Reynolds number, based on mean aerodynamic chord, of 1.4 x 10(exp 6). Pressure data measured over the forebody and leading-edge extensions are compared to similar pressure data taken by a related NASA flight research program by using a specially instrumented F/A-18, the High-Alpha Research Vehicle (HARV). Preliminary guidelines for high-angle-of-attack gritting strategies are given.

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

    NASA Technical Reports Server (NTRS)

    Adams, Richard J.

    1993-01-01

    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.

  9. Pitch control margin at high angle of attack - Quantitative requirements (flight test correlation with simulation predictions)

    NASA Technical Reports Server (NTRS)

    Lackey, J.; Hadfield, C.

    1992-01-01

    Recent mishaps and incidents on Class IV aircraft have shown a need for establishing quantitative longitudinal high angle of attack (AOA) pitch control margin design guidelines for future aircraft. NASA Langley Research Center has conducted a series of simulation tests to define these design guidelines. Flight test results have confirmed the simulation studies in that pilot rating of high AOA nose-down recoveries were based on the short-term response interval in the forms of pitch acceleration and rate.

  10. Pitch control margin at high angle of attack - Quantitative requirements (flight test correlation with simulation predictions)

    NASA Technical Reports Server (NTRS)

    Lackey, J.; Hadfield, C.

    1992-01-01

    Recent mishaps and incidents on Class IV aircraft have shown a need for establishing quantitative longitudinal high angle of attack (AOA) pitch control margin design guidelines for future aircraft. NASA Langley Research Center has conducted a series of simulation tests to define these design guidelines. Flight test results have confirmed the simulation studies in that pilot rating of high AOA nose-down recoveries were based on the short-term response interval in the forms of pitch acceleration and rate.

  11. Calculation of the flow on a blunted cone at high angle of attack

    NASA Technical Reports Server (NTRS)

    Lubard, S. C.; Rakich, J. V.

    1975-01-01

    A new technique for calculating the entire flow-field on spherically blunted circular cones at high angles of attack and high Reynolds numbers is described. The calculations are based on a single-layer system of three-dimensional parabolic equations which are approximations to the full steady Navier-Stokes equations. Initial conditions at the sphere-cone tangency plane are provided by using an inviscid time-dependent solution added to a viscous nonsimilar boundary layer solution. Calculated results are compared with experimental heat transfer and pressure data for a 15 deg half-angle cone with a 1-in. spherical nose at 15 deg angle of attack. The free-stream Mach number is 10.6, and the free-stream Reynolds number is 1,200,000 per foot. Excellent agreement between the calculated and experimental data for both pressure and heat transfer is obtained.

  12. Computational Investigation of Incompressible Airfoil Flows at High Angles of Attack

    DTIC Science & Technology

    1988-12-01

    Incompressible Airfoil Flows at High Angles of Attack by John Mark Mathre Lieutenant, United States Navy B.S., United States Naval Academy, 1978 Submitted...Similarly, in the y-direction the Navier-Stokes equation is ODv v 3v I P Z) v 32v - + U- + v- =- - + V(- + -). (2.24) Zt Zx zy p Dy x 2 Y2 11 III. STEADY

  13. Differential Game Based Guidance Law for High Angle of Attack Missiles

    DTIC Science & Technology

    1996-01-01

    Copyright 1996 by Optimal Synthesis . All Rights Reserved. Differential Game Based Guidance Law for High Angle of Attack Missiles By P. K. Menon and...G.B. Chatterji* Optimal Synthesis 450 San Antonio Road, Suite 53 Palo Alto, CA 94303 Abstract A nonlinear differential game theoretic intercept...Scientists, Optimal Synthesis Inc. Report Documentation Page Form ApprovedOMB No. 0704-0188 Public reporting burden for the collection of information is

  14. A study of roll attractor and wing rock of delta wings at high angles of attack

    NASA Technical Reports Server (NTRS)

    Niranjana, T.; Rao, D. M.; Pamadi, Bandu N.

    1993-01-01

    Wing rock is a high angle of attack dynamic phenomenon of limited cycle motion predominantly in roll. The wing rock is one of the limitations to combat effectiveness of the fighter aircraft. Roll Attractor is the steady state or equilibrium trim angle (phi(sub trim)) attained by the free-to-roll model, held at some angle of attack, and released form rest at a given initial roll (bank) angle (phi(sub O)). Multiple roll attractors are attained at different trim angles depending on initial roll angle. The test facility (Vigyan's low speed wind tunnel) and experimental work is presented here along with mathematical modelling of roll attractor phenomenon and analysis and comparison of predictions with experimental data.

  15. High-angle-of-attack stability and control improvements for the EA-6B Prowler

    NASA Technical Reports Server (NTRS)

    Jordan, Frank L.; Hahne, David E.; Masiello, Matthew F.; Gato, William

    1987-01-01

    The factors involved in high-angle-of-attack directional divergence phenomena for the EA-6B ECM aircraft have been investigated in NASA-Langley wind tunnel facilities in order to evaluate airframe modifications which would eliminate or delay such divergence to angles-of-attack farther removed from the operational flight envelope of the aircraft. The results obtained indicate that an adverse sidewash at the aft fuselage and vertical tail location is responsible for the directional stability loss, and that the sidewash is due to a vortex system generated by the fuselage-wing juncture. Modifications encompassing a wing inboard leading edge droop, a wing glove strake, and a vertical fin extension, have significantly alleviated the stability problem.

  16. Low-Speed Aerodynamic Characteristics of a Model of a Hypersonic Research Airplane at Angles of Attack up to 90 deg for a Range of Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Bowman, James S., Jr.; Grantham, William D.

    1960-01-01

    Static force tests have been made at low subsonic speeds for a model of a hypersonic research airplane in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic forces and moments up to an angle of attack of 90 deg for a range of Reynolds numbers. The Reynolds numbers, based on the mean aerodynamic chord, ranged from 740,000 to 1,900,000, which correspond to dynamic pressures from 15 to 100 lb/sq ft (Mach numbers from 0.10 to 0.27). The model was tested in the clean configuration with various horizontal-tail settings, horizontal tail off, lower rudder off, fuselage alone, and with various size strakes and slats on the nose of the model. Representative results of the present investigation are presented in plotted form, and a tabulation of all the data obtained is presented in a table. Appreciable effects on side force, yawing moment, and pitching moment are indicated by changes in Reynolds number for angles of attack of 40 to 90 deg.

  17. Aerodynamics of Tactical Weapons to Mach Number 8 and Angle-of-Attack of 180 deg

    DTIC Science & Technology

    1981-05-14

    to use Aerodynamic Prediction Code in 1971. The code was developed so as to handle fairly general wing -body- tail configurations and hence have...need exists for estimating the aerodynamic characteristics of a wide variety of tactical missile and projectile configurations , especially in the...planar or cruciform . Horizontalall-movable control deflections in the plus position are considered. Canard/ wing and

  18. Side forces on forebodies at high angles of attack and Mach numbers from 0.1 to 0.7: two tangent ogives, paraboloid and cone

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Chapman, G. T.; Taleghani, J.; Cohen, L.

    1977-01-01

    An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics of four forebodies at high angles of attack. The forebodies tested were a tangent ogive with fineness ratio of 5, a paraboloid with fineness ratio of 3.5, a 20 deg cone, and a tangent ogive with an elliptic cross section. The investigation included the effects of nose bluntness and boundary-layer trips. The tangent-ogive forebody was also tested in the presence of a short afterbody and with the afterbody attached. Static longitudinal and lateral/directional stability data were obtained. The investigation was conducted to investigate the existence of large side forces and yawing moments at high angles of attack and zero sideslip. It was found that all of the forebodies experience steady side forces that start at angles of attack of from 20 deg to 35 deg and exist to as high as 80 deg, depending on forebody shape. The side is as large as 1.6 times the normal force and is generally repeatable with increasing and decreasing angle of attack and, also, from test to test. The side force is very sensitive to the nature of the boundary layer, as indicated by large changes with boundary trips. The maximum side force caries considerably with Reynolds number and tends to decrease with increasing Mach number. The direction of the side force is sensitive to the body geometry near the nose. The angle of attack of onset of side force is not strongly influenced by Reynolds number or Mach number but varies with forebody shape. Maximum normal force often occurs at angles of attack near 60 deg. The effect of the elliptic cross section is to reduce the angle of onset by about 10 deg compared to that of an equivalent circular forebody with the same fineness ratio. The short afterbody reduces the angle of onset by about 5 deg.

  19. Experimental and theoretical study of aerodynamic characteristics of some lifting bodies at angles of attack from -10 degrees to 53 degrees at Mach numbers from 2.30 to 4.62

    NASA Technical Reports Server (NTRS)

    Spearman, M. Leroy; Torres, Abel O.

    1994-01-01

    Lifting bodies are of interest for possible use as space transportation vehicles because they have the volume required for significant payloads and the aerodynamic capability to negotiate the transition from high angles of attack to lower angles of attack (for cruise flight) and thus safely reenter the atmosphere and perform conventional horizontal landings. Results are presented for an experimental and theoretical study of the aerodynamic characteristics at supersonic speeds for a series of lifting bodies with 75 deg delta planforms, rounded noses, and various upper and lower surface cambers. The camber shapes varied in thickness and in maximum thickness location, and hence in body volume. The experimental results were obtained in the Langley Unitary Plan Wind Tunnel for both the longitudinal and the lateral aerodynamic characteristics. Selected experimental results are compared with calculated results obtained through the use of the Hypersonic Arbitrary-Body Aerodynamic Computer Program.

  20. Evaluation of High-Angle-of-Attack Handling Qualities for the X-31A Using Standard Evaluation Maneuvers

    NASA Technical Reports Server (NTRS)

    Stoliker, Patrick C.; Bosworth, John T.

    1996-01-01

    The X-31A aircraft gross-acquisition and fine-tracking handling qualities have been evaluated using standard evaluation maneuvers developed by Wright Laboratory, Wright-Patterson Air Force Base. The emphasis of the testing is in the angle-of-attack range between 30 deg and 70 deg. Longitudinal gross-acquisition handling qualities results show borderline Level 1/Level 2 performance. Lateral gross-acquisition testing results in Level 1/Level 2 ratings below 45 deg angle of attack, degrading into Level 3 as angle of attack increases. The fine-tracking performance in both longitudinal and lateral axes also receives Level 1 ratings near 30 deg angle of attack, with the ratings tending towards Level 3 at angles of attack greater than 50 deg. These ratings do not match the expectations from the extensive close-in combat testing where the X-31A aircraft demonstrated fair to good handling qualities maneuvering for high angles of attack. This paper presents the results of the high-angle-of-attack handling qualities flight testing of the X-31A aircraft. Discussion of the preparation for the maneuvers, the pilot ratings, and selected pilot comments are included. Evaluation of the results is made in conjunction with existing Neal-Smith, bandwidth, Smith-Geddes, and military specifications.

  1. Evaluation of High-Angle-of-Attack Handling Qualities for the X-31A Using Standard Evaluation Maneuvers

    NASA Technical Reports Server (NTRS)

    Stoliker, Patrick C.; Bosworth, John T.

    1997-01-01

    The X-31A aircraft gross-acquisition and fine-tracking handling qualities have been evaluated using standard evaluation maneuvers developed by Wright Laboratory, Wright Patterson Air Force Base. The emphasis of the testing is in the angle-of-attack range between 30 deg. and 70 deg. Longitudinal gross-acquisition handling qualities results show borderline Level l/Level 2 performance. Lateral gross-acquisition testing results in Level l/Level 2 ratings below 45 deg. angle of attack, degrading into Level 3 as angle of attack increases. The fine tracking performance in both longitudinal and lateral axes also receives Level 1 ratings near 30 deg. angle of attack, with the ratings tending towards Level 3 at angles of attack greater than 50 deg. These ratings do not match the expectations from the extensive close-in combat testing where the X-31A aircraft demonstrated fair to good handling qualities maneuvering for high angles of attack. This paper presents the results of the high-angle-of-attack handling qualities flight testing of the X-31A aircraft. Discussion of the preparation for the maneuvers, the pilot ratings, and selected pilot comments are included. Evaluation of the results is made in conjunction with existing Neal Smith, bandwidth, Smith-Geddes, and military specifications.

  2. Global Stability and Control Analysis of Aircraft at High Angles-of-Attack.

    DTIC Science & Technology

    1979-08-31

    6601ISOSLT 70 / .. ,.-j4tY CLASSIrIgCATION OF THIS PAGlVt’Whg Doea Knta.e) 20. (cont.) standing of the dynamic instabilities at high angles-of-attack. A basic...and this agrees with flight test results. The other groups have Cnr > 0 for high a, and so are less realistic. In addition to Cnr, the effects on...Silver Spring, MD 20910 Dover, NJ 07801 J. Wingate , Code R44 1 N. Coleman, DRDAR-SCFCC 1 ""Naval Air Test Center NASA Langley Research Center

  3. Quasi-periodic dynamics of a high angle of attack aircraft

    NASA Astrophysics Data System (ADS)

    Rohith, G.; Sinha, Nandan K.

    2017-01-01

    High angle of attack maneuvers closer to stall is a commonly accessed flight regime especially in case of fighter aircrafts. Stall and post-stall dynamics are dominated by nonlinearities which make the analysis difficult. Presence of external factors such as wind makes the system even more complex. Rich nonlinearities point to the possibility of existence of chaotic solutions. Past studies in this area confirm the development of such solutions. These studies are mainly concentrated on very high angle of attack regimes, which may not be practically easily accessible. This paper examines the possibility of existence of chaotic solutions in the lower, more accessible areas in the post stall domain. The analysis is composed of the study of effect of external wind as an agent to drive the system towards the possibility of a chaotic solution. Investigations reveal presence of quasi-periodic solutions, which are characterized by two incommensurate frequencies. This solution appears in the time simulation by varying the control parameter viz., wind. The solutions correspond to the values in the lower region of the angle of attack versus elevator bifurcation curve in the post-stall region. A steady wind is considered for the analysis and explores the possibility of chaotic motion by increasing the wind in a step wise manner. It is found that wind adds extra energy to the system which in turn drives the system in to chaos. The analysis is done with the help of phase portrait, Poincare map and amplitude spectrum and a quasi-periodic route to chaos via torus doubling is also presented.

  4. 2-D and 3-D oscillating wing aerodynamics for a range of angles of attack including stall

    NASA Technical Reports Server (NTRS)

    Piziali, R. A.

    1994-01-01

    A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted. Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment. The test has generated a very complete, detailed, and accurate body of data. These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions. This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall. This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil. Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results.

  5. Prediction of forces and moments on finned bodies at high angle of attack in transonic flow

    SciTech Connect

    Oberkampf, W. L.

    1981-04-01

    This report describes a theoretical method for the prediction of fin forces and moments on bodies at high angle of attack in subsonic and transonic flow. The body is assumed to be a circular cylinder with cruciform fins (or wings) of arbitrary planform. The body can have an arbitrary roll (or bank) angle, and each fin can have individual control deflection. The method combines a body vortex flow model and lifting surface theory to predict the normal force distribution over each fin surface. Extensive comparisons are made between theory and experiment for various planform fins. A description of the use of the computer program that implements the method is given.

  6. Fluid-structure interaction of a rolling restrained body of revolution at high angles of attack

    NASA Astrophysics Data System (ADS)

    Degani, D.; Ishay, M.; Gottlieb, O.

    2017-03-01

    The current work investigates numerically rolling instabilities of a free-to-roll slender rigid-body of revolution placed in a wind tunnel at a high angle of attack. The resistance to the roll moment is represented by a linear torsion spring and equivalent linear damping representing friction in the bearings of a simulated wind tunnel model. The body is subjected to a three-dimensional, compressible, laminar flow. The full Navier-Stokes equations are solved using the second-order implicit finite difference Beam-Warming scheme, adapted to a curvilinear coordinate system, whereas the coupled structural second order equation of motion for roll is solved by a fourth-order Runge-Kutta method. The body consists of a 3.5-diameter tangent ogive forebody with a 7.0-diameter long cylindrical afterbody extending aft of the nose-body junction to x/D = 10.5. We describe in detail the investigation of three angles of attack 20°, 40°, and 65°, at a Reynolds number of 30 000 (based on body diameter) and a Mach number of 0.2. Three distinct configurations are investigated as follows: a fixed body, a free-to-roll body with a weak torsion spring, and a free-to-roll body with a strong torsion spring. For each angle of attack the free-to-roll configuration portrays a distinct and different behavior pattern, including bi-stable limit-cycle oscillations. The bifurcation structure incorporates both large and small amplitude periodic roll oscillations where the latter lose their periodicity with increasing stiffness of the restraining spring culminating with distinct quasiperiodic oscillations. We note that removal of an applied upstream disturbance for a restrained body does not change the magnitude or complexity of the oscillations or of the flow patterns along the body. Depending on structure characteristics and flow conditions even a small rolling moment coefficient at the relatively low angle of attack of 20° may lead to large amplitude resonant roll oscillations.

  7. Preliminary results from a subsonic high angle-of-attack flush airdata sensing (HI-FADS) system: Design, calibration, and flight test evaluation

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Larson, Terry J.

    1990-01-01

    A nonintrusive high angle-of-attack flush airdata sensing (HI-FADS) system was installed and flight-tested on the F-18 high alpha research flight vehicle. The system is a matrix of 25 pressure orifices in concentric circles on the nose of the vehicle. The orifices determine angles of attack and sideslip, Mach number, and pressure altitude. Pressure was transmitted from the orifices to an electronically scanned pressure module by lines of pneumatic tubing. The HI-FADS system was calibrated and demonstrated using dutch roll flight maneuvers covering large Mach, angle-of-attack, and sideslip ranges. Reference airdata for system calibration were generated by a minimum variance estimation technique blending measurements from two wingtip airdata booms with inertial velocities, aircraft angular rates and attitudes, precision radar tracking, and meteorological analyses. The pressure orifice calibration was based on identifying empirical adjustments to modified Newtonian flow on a hemisphere. Calibration results are presented. Flight test results used all 25 orifices or used a subset of 9 orifices. Under moderate maneuvering conditions, the HI-FADS system gave excellent results over the entire subsonic Mach number range up to 55 deg angle of attack. The internal pneumatic frequency response of the system is accurate to beyond 10 Hz. Aerodynamic lags in the aircraft flow field caused some performance degradation during heavy maneuvering.

  8. In-flight leading-edge extension vortex flow-field survey measurements on a F-18 aircraft at high angle of attack

    NASA Technical Reports Server (NTRS)

    Richwine, David M.; Fisher, David F.

    1992-01-01

    Flow-field measurements on the leading-edge extension (LEX) of the F-18 High Alpha Research Vehicle (HARV) were obtained using a rotating rake with 16 hemispherical-tipped five-hole probes. Detailed pressure, velocity, and flow direction data were obtained through the LEX vortex core. Data were gathered during 1-g quasi-stabilized flight conditions at angles of attack alpha from 10 degrees to 52 degrees and at Reynolds numbers based on mean aerodynamic cord up to 16 x 10(exp 6). Normalized dynamic pressures and crossflow velocities clearly showed the primary vortex above the LEX and formation of a secondary vortex at higher angles of attack. The vortex was characterized by a ring of high dynamic pressure surrounding a region of low dynamic pressure at the vortex core center. The vortex core, subcore diameter, and vertical location of the core above the LEX increased with angle of attack. Minimum values for static pressure were obtained in the vortex subcore and decreased nearly linearly with increasing angle of attack until vortex breakdown. Rake-measured static pressures were consistent with previously documented surface pressures and showed good agreement with flow visualization flight test results. Comparison of the LEX vortex flight test data to computational solutions at alpha approximately equals 19 degrees and 30 degrees showed fair correlation.

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

    NASA Technical Reports Server (NTRS)

    Ostroff, Aaron J.; Proffitt, Melissa S.

    1993-01-01

    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.

  10. Large-Vortex Capture by a Wing at Very High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Wu, J. M.; Wu, J. Z.; Denny, G. A.; Lu, X. Y.

    1996-01-01

    In generating the lift on a wing, the static stall is a severe barrier. As the angle of attack, alpha, increases to the stall angle, alpha(sub stall) the flow separation point on the upper surface of the wing moves to the leading edge, so that on a two-dimensional airfoil or a large-aspect-ratio wing, the lift abruptly drops to a very low level. Therefore, the first generation of aeronautical flow type, i.e., the attached steady flow, has been limited to alpha less than alpha(sub stall). Owing to the obvious importance in applications, therefore, a great effort has been made in the past two decades to enlarge the range of usable angles of attack by various flow controls for a large-aspect-ratio wing. Basically, relevant works fall into two categories. The first category is usually refereed to as separation control, which concentrates on partially separated flow at alpha less than alpha(sub stall). Since the first experimental study of Collins and Zelenevitz, there has been ample literature showing that a partially separated flow can be turned to almost fully attached by flow controls, so that the lift is recovered and the stall is delayed (for a recent work see Seifert et al.). It has been well established that, in this category, unsteady controls are much more effective than steady ones and can be realized at a very low power-input level (Wu et al.; Seifert et al.). The second and more ambitious category of relevant efforts is the post-stall lift enhancement. Its possibility roots at the existence of a second lift peak at a very high angle of attack. In fact, As alpha further increases from alpha(sub stall), the completely separated flow develops and gradually becomes a bluff-body flow. This flow gives a normal force to the airfoil with a lift component, which reaches a peak at a maximum utilizable angle of attack, alpha(sub m) approx.= 40 deg. This second peak is of the same level as the first lift peak at alpha(sub stall). Meanwhile, the drag is also quickly

  11. Simulating Effects of High Angle of Attack on Turbofan Engine Performance

    NASA Technical Reports Server (NTRS)

    Liu, Yuan; Claus, Russell W.; Litt, Jonathan S.; Guo, Ten-Huei

    2013-01-01

    A method of investigating the effects of high angle of attack (AOA) flight on turbofan engine performance is presented. The methodology involves combining a suite of diverse simulation tools. Three-dimensional, steady-state computational fluid dynamics (CFD) software is used to model the change in performance of a commercial aircraft-type inlet and fan geometry due to various levels of AOA. Parallel compressor theory is then applied to assimilate the CFD data with a zero-dimensional, nonlinear, dynamic turbofan engine model. The combined model shows that high AOA operation degrades fan performance and, thus, negatively impacts compressor stability margins and engine thrust. In addition, the engine response to high AOA conditions is shown to be highly dependent upon the type of control system employed.

  12. Flight validation of ground-based assessment for control power requirements at high angles of attack

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Ross, Holly M.; Foster, John V.; Pahle, Joseph W.; Sternberg, Charles A.; Traven, Ricardo; Lackey, James B.; Abbott, Troy D.

    1994-01-01

    A review is presented in viewgraph format of an ongoing NASA/U.S. Navy study to determine control power requirements at high angles of attack for the next generation high-performance aircraft. This paper focuses on recent flight test activities using the NASA High Alpha Research Vehicle (HARV), which are intended to validate results of previous ground-based simulation studies. The purpose of this study is discussed, and the overall program structure, approach, and objectives are described. Results from two areas of investigation are presented: (1) nose-down control power requirements and (2) lateral-directional control power requirements. Selected results which illustrate issues and challenges that are being addressed in the study are discussed including test methodology, comparisons between simulation and flight, and general lessons learned.

  13. MISDRAG- AERODYNAMIC DRAG ON WING-BODY COMBINATIONS AT SMALL ANGLES OF ATTACK IN SUPERSONIC FLOW

    NASA Technical Reports Server (NTRS)

    Sawyer, W. C.

    1994-01-01

    A computer program has been written to obtain the wave and friction drag of configurations with bodies of revolution and fins. These inviscid flow fields are superimposed and the wave drag of the configuration is obtained by integration of the surface pressures. The friction drag is obtained from the viscous flow field of the body and a flat-plate friction analysis of the fins. The numerical solution of these flow fields, superposition, and integration to obtain total drag have been programmed for high-speed digital computation. A large portion of the input required by the program is involved with the description of the configuration geometry and the specific surface positions where pressures are to be evaluated. In addition to drag forces, an output is available whereby the pressure distributions on the body and fins can be obtained.

  14. An EnKF-based Flow State Estimator for Airfoils at High Angles of Attack

    NASA Astrophysics Data System (ADS)

    de Castro da Silva, Andre Fernando; Colonius, Tim

    2016-11-01

    Robust flow estimation from available measurements remains a major obstacle to successful flow control applications. Although several estimation methodologies have been developed in the past decades, the high dimensionality of fluid systems renders many of them computationally intractable. In this work, we employ the Ensemble Kalman Filter (EnKF) and the two-dimensional incompressible Navier-Stokes equations to estimate the state of the flow past a NACA 0009 airfoil at high angles of attack and moderate Reynolds number. The pressure distribution on the airfoil and the velocity field in the wake, both randomized by synthetic noise, are sampled as measurement data. In order to evaluate the relative importance of each sensor location to the estimate correction, their influence fields (also known as representers) are analyzed. The performance of the estimator is then assessed for different choices of ensemble size, noise levels, and number/location of sensors. Graduate Student.

  15. Flight-Determined Subsonic Longitudinal Stability and Control Derivatives of the F-18 High Angle of Attack Research Vehicle (HARV) with Thrust Vectoring

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Wang, Kon-Sheng Charles

    1997-01-01

    The subsonic longitudinal stability and control derivatives of the F-18 High Angle of Attack Research Vehicle (HARV) are extracted from dynamic flight data using a maximum likelihood parameter identification technique. The technique uses the linearized aircraft equations of motion in their continuous/discrete form and accounts for state and measurement noise as well as thrust-vectoring effects. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics over the aircraft, particularly at high angles of attack. Thrust vectoring was implemented using electrohydraulically-actuated nozzle postexit vanes and a specialized research flight control system. During maneuvers, a control system feature provided independent aerodynamic control surface inputs and independent thrust-vectoring vane inputs, thereby eliminating correlations between the aircraft states and controls. Substantial variations in control excitation and dynamic response were exhibited for maneuvers conducted at different angles of attack. Opposing vane interactions caused most thrust-vectoring inputs to experience some exhaust plume interference and thus reduced effectiveness. The estimated stability and control derivatives are plotted, and a discussion relates them to predicted values and maneuver quality.

  16. Optimization of lateral-directional dynamics for an aircraft operating at high angle of attack

    NASA Technical Reports Server (NTRS)

    Snell, S. A.; Garrard, William L., Jr.; Enns, Dale F.

    1991-01-01

    In this paper, the control laws for the lateral-directional dynamics of a supermaneuverable aircraft is analyzed with a view to reducing the levels of lateral acceleration and sideslip, which are encountered during aggressive rolling maneuvers at high angles of attack. The analysis uses a linearized model of the lateral-directional dynamics and thus H-free-flow techniques can be applied. It is shown that trade-offs exist between simultaneously minimizing lateral acceleration measured at the pilot's station, ny(p), minimizing sideslip and minimizing tracking errors between the roll-rate about the velocity vector and its command. The paper concludes that a significant reduction in ny(p) is only attainable by compromising the roll-rate performance.

  17. Separation control over an airfoil at high angles of attack by sound emanating from the surface

    NASA Technical Reports Server (NTRS)

    Huang, L. S.; Maestrello, L.; Bryant, T. D.

    1987-01-01

    Active control by sound emanating from a narrow gap in the vicinity of the leading edge of a symmetrical airfoil is used to study the influence of sound on the pressure distribution and the wake at high angles of attack. The results from experiments conducted at a Reynolds number based on the chord of 35,000 show that, with injection of sound at twice the shedding frequency of the shear layer, the region of separation becomes drastically reduced. The shear layer is found to be very sensitive to sound excitation in the vicinity of the separation point. The excitation sufficiently alters the global circulation to cause an increase in lift and reduction in drag. Furthermore, experimental results describing stall and post-stall conditions compare well with the limited data available and indicate that stall is delayed by sound injection into the separated region.

  18. PIV-based estimation of unsteady loads on a flat plate at high angle of attack using momentum equation approaches

    NASA Astrophysics Data System (ADS)

    Guissart, A.; Bernal, L. P.; Dimitriadis, G.; Terrapon, V. E.

    2017-05-01

    This work presents, compares and discusses results obtained with two indirect methods for the calculation of aerodynamic forces and pitching moment from 2D Particle Image Velocimetry (PIV) measurements. Both methodologies are based on the formulations of the momentum balance: the integral Navier-Stokes equations and the "flux equation" proposed by Noca et al. (J Fluids Struct 13(5):551-578, 1999), which has been extended to the computation of moments. The indirect methods are applied to spatio-temporal data for different separated flows around a plate with a 16:1 chord-to-thickness ratio. Experimental data are obtained in a water channel for both a plate undergoing a large amplitude imposed pitching motion and a static plate at high angle of attack. In addition to PIV data, direct measurements of aerodynamic loads are carried out to assess the quality of the indirect calculations. It is found that indirect methods are able to compute the mean and the temporal evolution of the loads for two-dimensional flows with a reasonable accuracy. Nonetheless, both methodologies are noise sensitive, and the parameters impacting the computation should thus be chosen carefully. It is also shown that results can be improved through the use of dynamic mode decomposition (DMD) as a pre-processing step.

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

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Foster, John V.; Hoffler, Keith D.

    2005-01-01

    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.

  20. Aerodynamic characteristics of seven symmetrical airfoil sections through 180-degree angle of attack for use in aerodynamic analysis of vertical axis wind turbines

    SciTech Connect

    Sheldahl, R E; Klimas, P C

    1981-03-01

    When work began on the Darrieus vertical axis wind turbine (VAWT) program at Sandia National Laboratories, it was recognized that there was a paucity of symmetrical airfoil data needed to describe the aerodynamics of turbine blades. Curved-bladed Darrieus turbines operate at local Reynolds numbers (Re) and angles of attack (..cap alpha..) seldom encountered in aeronautical applications. This report describes (1) a wind tunnel test series conducted at moderate values of Re in which 0 less than or equal to ..cap alpha.. less than or equal to 180/sup 0/ force and moment data were obtained for four symmetrical blade-candidate airfoil sections (NACA-0009, -0012, -0012H, and -0015), and (2) how an airfoil property synthesizer code can be used to extend the measured properties to arbitrary values of Re (10/sup 4/ less than or equal to Re less than or equal to 10/sup 7/) and to certain other section profiles (NACA-0018, -0021, -0025).

  1. Low speed rotary aerodynamics of F-18 configuration for 0 deg to 90 deg angle of attack: Test results and analysis

    NASA Technical Reports Server (NTRS)

    Hultberg, R.

    1984-01-01

    Aerodynamic characteristics obtained in a rotational flow environment, utilizing a rotary balance located in the Langley Spin Tunnel, are discussed and presented in tabular form for a 1/10 scale F-18 airplane model. The rotational aerodynamic characteristics were established for the basic airplane, as well as the influence of control deflections and the contribution of airplane components, i.e., body, wing, leading edge extension, horizontal and vertical tails, on these characteristics up to 90 deg angle of attack. Spin equilibrium conditions predicted using the measured data are also presented and compared with spin model and full scale flight results.

  2. LDV Surveys Over a Fighter Model at Moderate to High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Sellers, William L., III; Meyers, James F.; Hepner, Timothy E.

    2004-01-01

    The vortex flowfield over an advanced twin-tailed fighter configuration was measured in a low-speed wind tunnel at two angles of attack. The primary test data consisted of 3-component velocity surveys obtained using a Laser Doppler Velocimeter. Laser light sheet and surface flow visualization were also obtained to provide insight into the flowfield structure. Time-averaged velocities and the root mean square of the velocity fluctuations were obtained at two cross-sections above the model. At 15 degrees angle of attack, the vortices generated by the wing leading edge extension (LEX) were unburst over the model and passed outboard of the vertical tail. At 25 degrees angle of attack, the vortices burst in the vicinity of the wing-LEX intersection and impact directly on the vertical tails. The RMS levels of the velocity fluctuations reach values of approximately 30% in the region of the vertical tails.

  3. High angle-of-attack characteristics of three-surface fighter aircraft. [canard-wing-horizontal tail configuration for greater stability and control

    NASA Technical Reports Server (NTRS)

    Croom, M. A.; Grafton, S. B.; Nguyen, L. T.

    1982-01-01

    As part of a research program aimed at providing information on the high angle-of-attack characteristics of three-surface fighter concepts incorporating a close-coupled canard, an investigation is being conducted on two specific configurations based on the F-18 and F-15 designs. The study configurations are being subjected to a wide range of tests including wind-tunnel tests, dynamic model tests, and piloted simulation. This paper summarizes the results obtained to date in this study. High-alpha results in the areas of static stability, damping, and control characteristics are reviewed and some of the more significant aerodynamic phenomena are identified.

  4. Eigenstructure assignment for a thrust-vectored high angle-of-attack aircraft

    NASA Technical Reports Server (NTRS)

    Sobel, Kenneth M.; Lallman, Frederick J.

    1988-01-01

    Eigenstructure assignment is utilized to design flight control laws for a thrust-vectored aircraft at several different angles of attack. An interesting characteristic of the aircraft model is that the control distribution matrix is rank-deficient. Also, the effectiveness of the control inputs varies with the angle of attack. A pseudocontrol strategy is used to reduce the control space to two dimensions. After the eigenstructure assignment design is complete, the controller is mapped back to the original five-dimensional control space. The designs are shown to exhibit acceptable multivariable stability margins at the aircraft inputs.

  5. An Inlet Distortion Assessment During Aircraft Departures at High Angle of Attack for an F/A-18A Aircraft

    NASA Technical Reports Server (NTRS)

    Steenken, William G.; Williams, John G.; Yuhas, Andrew J.; Walsh, Kevin R.

    1997-01-01

    The F404-GE-400-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. 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 entry yaw rates were approximately 40 to 90 deg/sec. 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 did not significantly exceed those at maximum angle-of-attack and sideslip maneuvers during controlled flight. Surges caused by inlet distortion levels resulted from a combination of high levels of inlet distortion and rapid changes in aircraft position. These rapid changes indicate a combination of engine support 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. This document presents the slides from an oral presentation.

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

    NASA Technical Reports Server (NTRS)

    Brandon, J. M.; Nguyen, L. T.

    1986-01-01

    A series of low speed wind tunnel tests on a generic fighter model with a cylindrical fuselage were made to investigate the effects of forebody shape on static and dynamic lateral/directional stability. Five forebodies, including a chine nose of unconventional cross-sectional shape, were tested. Conventional force tests were conducted to determine static stability characteristics and single degree-of-freedom free-to-roll tests were used to study the wing rock susceptibility of the model with the various forebodies. Flow visualization data were obtained to aid in analysis of the complex flow phenomena involved. The results show that forebody cross-sectional shape can strongly effect both static and dynamic (roll) stability at high angles of attack. Large variations in stability were obtained for the various forebody geometries. These characteristics result from the impact of cross-sectional shape on forebody vortex development, the behavior of the vortices at sideslip conditions, and their interaction with the wing and empennage flow fields.

  7. A Tail Buffet Loads Prediction Method for Aircraft at High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Pototzky, Anthony S.; Moses, Robert W.

    2005-01-01

    Aircraft designers commit significant resources to the design of aircraft in meeting performance goals. Despite fulfilling traditional design requirements, many fighter aircraft have encountered buffet loads when demonstrating their high angle-of-attack maneuver capabilities. As a result, during test or initial production phases of fighter development programs, many new designs are impacted, usually in a detrimental way, by resulting in reassessing designs or limiting full mission capability. These troublesome experiences usually stem from overlooking or completely ignoring the effects of buffet during the design phase of aircraft. Perhaps additional requirements are necessary that addresses effects of buffet in achieving best aircraft performance in fulfilling mission goals. This paper describes a reliable, fairly simple, but quite general buffet loads analysis method to use in the initial design phases of fighter-aircraft development. The method is very similar to the random gust load analysis that is now commonly available in a commercial code, which this analysis capability is based, with some key modifications. The paper describes the theory and the implementation of the methodology. The method is demonstrated on a JSF prototype example problem. The demonstration also serves as a validation of the method, since, in the paper, the analysis is shown to nearly match the flight data. In addition, the paper demonstrates how the analysis method can be used to assess candidate design concepts in determining a satisfactory final aircraft configuration.

  8. Technical Evaluation Report, Part A - Vortex Flow and High Angle of Attack

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2003-01-01

    A symposium entitled Vortex Flow and High Angle of Attack was held in Loen, Norway, from May 7 through May 11, 2001. The Applied Vehicle Technology (AVT) panel, under the auspices of the Research and Technology Organization (RTO), sponsored this symposium. Forty-eight papers, organized into nine sessions, addressed computational and experimental studies of vortex flows pertinent to both aircraft and maritime applications. The studies also ranged from fundamental fluids investigations to flight test results, and significant results were contributed from a broad range of countries. The principal emphasis of this symposium was on "the understanding and prediction of separation-induced vortex flows and their effects on military vehicle performance, stability, control, and structural design loads." It was further observed by the program committee that "separation- induced vortex flows are an important part of the design and off-design performance of conventional fighter aircraft and new conventional or unconventional manned or unmanned advanced vehicle designs (UAVs, manned aircraft, missiles, space planes, ground-based vehicles, and ships)." The nine sessions addressed the following topics: vortical flows on wings and bodies, experimental techniques for vortical flows, numerical simulations of vortical flows, vortex stability and breakdown, vortex flows in maritime applications, vortex interactions and control, vortex dynamics, flight testing, and vehicle design. The purpose of this paper is to provide brief reviews of these papers along with some synthesizing perspectives toward future vortex flow research opportunities. The paper includes the symposium program. (15 refs.)

  9. Extraction of Lateral-Directional Stability and Control Derivatives for the Basic F-18 Aircraft at High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Wang, Kon-Sheng Charles

    1997-01-01

    The results of parameter identification to determine the lateral-directional stability and control derivatives of an F-18 research aircraft in its basic hardware and software configuration are presented. The derivatives are estimated from dynamic flight data using a specialized identification program developed at NASA Dryden Flight Research Center. The formulation uses the linearized aircraft equations of motions in their continuous/discrete form and a maximum likelihood estimator that accounts for both state and measurement noise. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics, such as separated and vortical flows, over the aircraft. The derivatives are plotted as functions of angle of attack between 3 deg and 47 deg and compared with wind-tunnel predictions. The quality of the derivative estimates obtained by parameter identification is somewhat degraded because the maneuvers were flown with the aircraft's control augmentation system engaged, which introduced relatively high correlations between the control variables and response variables as a result of control motions from the feedback control system.

  10. Low-speed wind tunnel performance of high-speed counterrotation propellers at angle-of-attack

    NASA Technical Reports Server (NTRS)

    Hughes, Christopher E.; Gazzaniga, John A.

    1989-01-01

    The low-speed aerodynamic performance characteristics of two advanced counterrotation pusher-propeller configurations with cruise design Mach numbers of 0.72 were investigated in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. The tests were conducted at Mach number 0.20, which is representative of the aircraft take-off/landing flight regime. The investigation determined the effect of nonuniform inflow on the propeller performance characteristics for several blade angle settings and a range of rotational speeds. The inflow was varied by yawing the propeller model to angle-of-attack by as much as plus or minus 16 degrees and by installing on the counterrotation propeller test rig near the propeller rotors a model simulator of an aircraft engine support pylon and fuselage. The results of the investigation indicated that the low-speed performance of the counterrotation propeller configurations near the take-off target operating points were reasonable and were fairly insensitive to changes in model angle-of-attack without the aircraft pylon/fuselage simulators installed on the propeller test rig. When the aircraft pylon/fuselage simulators were installed, small changes in propeller performance were seen at zero angle-of-attack, but fairly large changes in total power coefficient and very large changes of aft-to-forward-rotor torque ratio were produced when the propeller model was taken to angle-of-attack. The propeller net efficiency, though, was fairly insensitive to any changes in the propeller flowfield conditions near the take-off target operating points.

  11. Experimental investigation into wing span and angle-of-attack effects on sub-scale race car wing/wheel interaction aerodynamics

    NASA Astrophysics Data System (ADS)

    Diasinos, S.; Gatto, A.

    2008-09-01

    This paper details a quantitative 3D investigation using LDA into the interaction aerodynamics on a sub-scale open wheel race car inverted front wing and wheel. Of primary importance to this study was the influence of changing wing angle of attack and span on the resulting near-field and far-field flow characteristics. Results obtained showed that both variables do have a significant influence on the resultant flow-field, particularly on wing vortex and wheel wake development and propagation.

  12. Supersonic wing-body inteference at high angles of attack with emphasis on low aspect ratios

    NASA Technical Reports Server (NTRS)

    Nielsen, J. N.

    1986-01-01

    A systematic NASA data base for cruciform wing-body combinations has been analyzed to extract the values of the wing and body interference factors. The fin planforms vary in aspect ratio from 0.5 to 2.0, the Mach number from 2.5 to 4.5, and the angle of attack from 0 deg to 40 deg. Sufficient data are available to permit interpolation with respect to fin aspect ratio, taper ratio, Mach number, and angle of attack. The data base described in this paper can be used in an engineering prediction method to determine the normal force of planar wing-body combinations and for scaling the effects of body radius-fin semispan ratio.

  13. Analysis of flight data on boundary layer transition at high angles of attack

    NASA Technical Reports Server (NTRS)

    Haigh, W. W.; Lake, B. M.; Ko, D. R. S.

    1972-01-01

    Boundary layer transition data were obtained on the flight of two cones which reentered at velocities of about 7.0 km/sec. One cone reentered at a nominal zero degree angle of attack and the other, due to an anomaly above the earth atmosphere, reentered at local angles of attack up to 7.0 km/sec. The transition data were obtained from on-board acoustic and electrostatic sensors. A description of the design, calibration, and method used to detect transition from the sensors is included. The flow field calculation used to obtain the local flow properties on the cones is described. Finally, the transition data found from both cone flights is correlated.

  14. Space shuttle: Aerodynamic characteristics of cone-cylinder-flare-fin configurations at Mach numbers of 1.96, 2.74, and 4.96 and angles of attack from 50 to 90 degrees

    NASA Technical Reports Server (NTRS)

    Bradley, D.; Ellis, R. R.

    1972-01-01

    A 0.00227-scale parametric model of an LMSC/MSFC water recoverable booster was tested in the MSFC 14 x 14-inch trisonic wind tunnel. The purpose of the test was to obtain high angle of attack force and static stability data which could be used by MSFC in preliminary design and aerodynamic trade studies. These data were obtained using six-component internal strain gauge balances. One hundred forty-four different geometrical combinations were possible as all model parts were interchangeable (three nose cones, three cylinder lengths, four flare sections and three sets of fins, plus a no-fin case in combination with the other components). However, due to tunnel occupancy limitations, only the most representative combinations were tested. All configurations investigated were tested at Mach 1.96, 2.74 and 4.96 with data obtained at angles of attack from 50 degrees to 90 degrees and at angles of sideslip from -10 degrees to +10 degrees (at an angle of attack of 60 degrees).

  15. Lateral control at high angles of attack using pneumatic blowing through a chined forebody

    NASA Technical Reports Server (NTRS)

    Arena, A. S., Jr.; Nelson, R. C.; Schiff, L. B.

    1993-01-01

    Directional control through the use of pneumatic blowing was investigated on a generic subscale model with a chined forebody with blowing through a chine slot in a direction normal to the forebody surface. Comparisons are made with a vertical tail on and off, and with control through rudder deflection. Force and moment data were obtained for various blowing coefficients over a 0-75 deg alpha range, and flow visualization was also conducted in order to see qualitative effects on the flowfield. Blowing through a chined forebody generates yaw moments at large alpha where control surfaces lose their effectiveness; these moments are much larger than obtained by jet thrust alone, since the forebody flowfield is modified through the interaction of the jet with the chine vortices. Directional control increased with angle of attack for a given blowing coefficient until a maximum was reached. Further increases in angle of attack results in a rapid loss of effectiveness. For angles of attack above 60 deg, yaw moments are generated by simple jet thrust effect. The effectiveness of the pneumatic system depended on tail configuration.

  16. Lateral control at high angles of attack using pneumatic blowing through a chined forebody

    NASA Technical Reports Server (NTRS)

    Arena, A. S., Jr.; Nelson, R. C.; Schiff, L. B.

    1993-01-01

    Directional control through the use of pneumatic blowing was investigated on a generic subscale model with a chined forebody with blowing through a chine slot in a direction normal to the forebody surface. Comparisons are made with a vertical tail on and off, and with control through rudder deflection. Force and moment data were obtained for various blowing coefficients over a 0-75 deg alpha range, and flow visualization was also conducted in order to see qualitative effects on the flowfield. Blowing through a chined forebody generates yaw moments at large alpha where control surfaces lose their effectiveness; these moments are much larger than obtained by jet thrust alone, since the forebody flowfield is modified through the interaction of the jet with the chine vortices. Directional control increased with angle of attack for a given blowing coefficient until a maximum was reached. Further increases in angle of attack results in a rapid loss of effectiveness. For angles of attack above 60 deg, yaw moments are generated by simple jet thrust effect. The effectiveness of the pneumatic system depended on tail configuration.

  17. Aerodynamic Characteristics of a Circular Cylinder at Mach Number 6.86 and Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Penland, Jim A

    1957-01-01

    Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.88, a Reynolds number of 129,000, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and a simple modification of the Newtonian flow theory. An evaluation of the crossflow theory is made through comparison of present results with available crossflow Mach number drag coefficients.

  18. Side forces on a tangent ogive forebody with a fineness ratio of 3.5 at high angles of attack and Mach numbers from 0.1 to 0.7

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Chapman, G. T.; Cohen, L.; Taleghani, J.

    1977-01-01

    An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics, at high angles of attack, of a tangent ogive forebody with a fineness ratio of 3.5. The investigation included the effects of nose bluntness, nose strakes, nose booms, a simulated canopy, and boundary-layer trips. The forebody was also tested with a short afterbody attached. Static longitudinal and lateral-directional stability data were obtained at Reynolds numbers ranging from 0.3 mil. to 3.8 mil. (based on base diameter) at a Mach number of 0.25, and at a Reynolds number of 0.8 mil. at Mach numbers ranging from 0.1 to 0.7. Angle of attack was varied from 0 to 88 deg at zero sideslip, and the sideslip angle was varied from -10 to 30 deg at angles of attack of 40, 55, and 70 deg.

  19. Wind-tunnel research comparing lateral control devices, particularly at high angles of attack VI : skewed ailerons on rectangular wings

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Harris, Thomas A

    1934-01-01

    This report covers the sixth of a series of investigations in which various lateral control devices are compared with particular reference to their effectiveness at high angles of attack. The present report deals with flap-type ailerons hinged about axes having an angle with respect to the leading and trailing edges of the wing. Tests were made on four different skewed ailerons, including two different angles of skew and two sizes of ailerons. At the high angles of attack, all the skewed ailerons tested were slightly inferior with respect to rolling and yawing moments to straight ailerons having the same span and average chord. Computations indicate that the skewed ailerons are also inferior with respect to hinge moments.

  20. Effect of nose perturbation on asymmetric vortices over a blunt-nose body at high angle of attack

    NASA Astrophysics Data System (ADS)

    Sha, Y. X.; Wang, Y. K.; Qi, Z. Y.

    2017-04-01

    To improve the maneuverability, the modern missiles of blunt-nose slender body configuration are required to flight at high angle of attack where many research show complex asymmetric vortices flow will develop over the slender body. The effect of artificial perturbation located on the nose of a blunt slender body at high angle of attack(α=50°) on asymmetric vortices has been investigated with low speed wind tunnel test. The Reynolds number(ReD) of the experiment was 1.48×105. The experimental results are shown to be the following: Asymmetric flow over blunt-nose body was extremely sensitive to the machining tolerances of the nose and can be governed by the artificial perturbation; with the circumferential angle of artificial perturbation varied a period, asymmetric vortices present a behavior of single-period.

  1. Convergence behavior that controls adaptive wind tunnel walls near the test section in the high angle of attack range

    NASA Technical Reports Server (NTRS)

    Ziemann, J.

    1982-01-01

    The NACA 0012 profile at Mach 0.5 was investigated in a wind tunnel with adaptive walls. It is found that adaptation of the flexible walls is possible in the high angle of attack range on both sides of maximum lift. Oil film photographs of the flow at the profile surface show three dimensional effects in the region of the corners between the profile and the sidewall. It is concluded that pure two dimensional separated flow is not possible.

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

    NASA Technical Reports Server (NTRS)

    Hemsch, Michael J.

    1990-01-01

    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.

  3. Flutter Clearance of the F-18 High-angle-of-attack Research Vehicle with Experimental Wingtip Instrumentation Pods

    NASA Technical Reports Server (NTRS)

    Freudinger, Lawrence C.

    1989-01-01

    An F-18 aircraft was modified with wingtip instrumentation pods for use in NASA's high-angle-of-attack research program. Ground vibration and flight flutter testing were performed to clear an acceptable flight envelope for the aircraft. Flight test utilized atmospheric turbulence for structural excitation; the aircraft displayed no adverse aeroelastic trends within the envelope tested. The data presented in this report include mode shapes from the ground vibration and estimates of frequency and damping as a function of Mach number.

  4. Operating characteristics of an inlet model tested with a 0.5m powered fan at high angles of attack

    NASA Technical Reports Server (NTRS)

    Koncsek, J. L.; Shaw, R. J.

    1977-01-01

    An inlet model designed for high angle of attack capability, coupled to a .508 m tip diameter turbofan simulator, was tested in the NASA-Lewis Research Center's 9-by 15-ft low speed wind tunnel. The test variables were: tunnel velocity, 0 to 75 m/s; inlet angle of attack, 0 to 120 deg; and fan face corrected airflow per unit area, 75 to 200 kg/s sqm. The inlet flow separation boundaries, the fan face total pressure recovery and distortion characteristics, and the fan blade vibratory stresses were determined. The recovery, distortion, and stress levels showed no abrupt changes at the onset of separation, but became gradually more unfavorable as the size and intensity of the separation increased as induced by increasingly severe operating conditions. Performance characteristics for a large scale model of the inlet were estimated from these test results.

  5. Status of the Validation of High-Angle-Of-Attack Nose-Down Pitch Control Margin Design Guidelines

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Foster, John V.; Pahle, Joseph W.; Wilson, R. Joe; Lackey, James B.

    1992-01-01

    This paper presents a summary of results obtained to date in an ongoing cooperative research program between NASA and the U.S. Navy to develop design criteria for high-angle-of-attack nose- down pitch control for combat aircraft. A fundamental design consideration for aircraft incorporating relaxed static stability in pitch is the level of stability which achieves a proper balance between high- speed performance considerations and low-speed requirements for maneuvering at high angles of attack. A comprehensive data base of piloted simulation results was generated for parametric variations of critical parameters affecting nose-down control capability. The results showed a strong correlation of pilot rating to the short-term pitch response for nose-down commands applied at high- angle-of-attack conditions. Using these data, candidate design guidelines and flight demonstration requirements were defined. Full- scale flight testing to validate the research methodology and proposed guidelines is in progress, some preliminary results of which are reviewed.

  6. Effect of aerodynamic and angle-of-attack uncertainties on the May 1979 entry flight control system of the Space Shuttle from Mach 8 to 1.5

    NASA Technical Reports Server (NTRS)

    Stone, H. W.; Powell, R. W.

    1985-01-01

    A six degree of freedom simulation analysis was performed for the space shuttle orbiter during entry from Mach 8 to Mach 1.5 with realistic off nominal conditions by using the flight control systems defined by the shuttle contractor. The off nominal conditions included aerodynamic uncertainties in extrapolating from wind tunnel derived characteristics to full scale flight characteristics, uncertainties in the estimates of the reaction control system interaction with the orbiter aerodynamics, an error in deriving the angle of attack from onboard instrumentation, the failure of two of the four reaction control system thrusters on each side, and a lateral center of gravity offset coupled with vehicle and flow asymmetries. With combinations of these off nominal conditions, the flight control system performed satisfactorily. At low hypersonic speeds, a few cases exhibited unacceptable performances when errors in deriving the angle of attack from the onboard instrumentation were modeled. The orbiter was unable to maintain lateral trim for some cases between Mach 5 and Mach 2 and exhibited limit cycle tendencies or residual roll oscillations between Mach 3 and Mach 1. Piloting techniques and changes in some gains and switching times in the flight control system are suggested to help alleviate these problems.

  7. Hybrid structured/unstructured grid computations for the F/A-18 at high angle of attack

    NASA Technical Reports Server (NTRS)

    Biedron, Robert T.; Whitaker, David L.

    1994-01-01

    At high angles of attack, vortical flows play a crucial role in the maintenance of lift for fighter aircraft. However, under certain conditions, the vortical flow can have an adverse effect on the aircraft. On the F/A-18 the LEX vortex can impinge on the tail; at high angles of attack the unsteady flow from vortex bursting can cause structural fatigue on the vertical tails. At high angles of attack, the flow field can be quite complex, and a computational analysis challenging. A full configuration analysis with viscous structured grids can be computationally expensive and the task of generating the requisite grids can be quite difficult. To mitigate these difficulties and provide a medium-fidelity analysis tool, a hybrid structured/unstructured approach was adopted for this work. In this analysis, the formation and roll-up of the LEX vortex is computed with a structured Navier-Stokes solver, and the resulting vortex propagated downstream with an unstructured Euler solver.

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

    NASA Technical Reports Server (NTRS)

    Brandon, J. M.; Murri, D. G.; Nguyen, L. T.

    1986-01-01

    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.

  9. Qualitative evaluation of a flush air data system at transonic speeds and high angles of attack

    NASA Technical Reports Server (NTRS)

    Larson, Terry J.; Whitmore, Stephen A.; Ehernberger, L. J.; Johnson, J. Blair; Siemers, Paul M., III

    1987-01-01

    Flight tests were performed on an F-14 aircraft to evaluate the use of flush pressure orifices on the nose section for obtaining air data at transonic speeds over a large range of flow angles. This program was part of a flight test and wind tunnel program to assess the accuracies of such systems for general use on aircraft. It also provided data to validate algorithms developed for the shuttle entry air data system designed at NASA Langley. Data were obtained for Mach numbers between 0.60 and 1.60, for angles of attack up to 26.0 deg, and for sideslip angles up to 11.0 deg. With careful calibration, a flush air data system with all flush orifices can provide accurate air data information over a large range of flow angles. Several orificies on the nose cap were found to be suitable for determination of stagnation pressure. Other orifices on the nose section aft of the nose cap were shown to be suitable for determination of static pressure. Pairs of orifices on the nose cap provided the most sensitive measurements for determining angles of attack and sideslip, although orifices located farther aft on the nose section could also be used.

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

    NASA Technical Reports Server (NTRS)

    Murman, Scott M.

    1994-01-01

    This report summarizes research done over the past two years as part of NASA Grant NCC 2-729. 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 two main topics: the effect of engine-inlet mass-flow rate on the F-18 vortex breakdown position, and the results using a refined F-18 computational model to compute the flow at alpha = 30 deg and alpha = 45 deg.

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

    NASA Technical Reports Server (NTRS)

    Murman, Scott M.

    1995-01-01

    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.

  12. Effect of aerodynamic and angle-of-attack uncertainties on the blended entry flight control system of the Space Shuttle from Mach 10 to 2.5

    NASA Technical Reports Server (NTRS)

    Stone, H. W.; Powell, R. W.

    1984-01-01

    A six-degree-of-freedom simulation analysis has been performed for the Space Shuttle Orbiter during entry from Mach 10 to 2.5 with realistic off-nominal conditions using the entry flight control system specified in May 1978. The off-nominal conditions included the following: (1) aerodynamic uncertainties, (2) an error in deriving the angle of attack from onboard instrumentation, (3) the failure of two of the four reaction control-system thrusters on each side, and (4) a lateral center-of-gravity offset. With combinations of the above off-nominal conditions, the control system performed satisfactorily with a few exceptions. The cases that did not exhibit satisfactory performance displayed the following main weaknesses. Marginal performance was exhibited at hypersonic speeds with a sensed angle-of-attack error of 4 deg. At supersonic speeds the system tended to be oscillatory, and the system diverged for several cases because of the inability to hold lateral trim. Several system modifications were suggested to help solve these problems and to maximize safety on the first flight: alter the elevon-trim and speed-brake schedules, delay switching to rudder trim until the rudder effectiveness is adequate, and reduce the overall rudder loop gain. These and other modifications were incorporated in a flight-control-system redesign in May 1979.

  13. High angle of attack position sensing for the Southampton University magnetic suspension and balance system

    NASA Technical Reports Server (NTRS)

    Parker, David H.

    1987-01-01

    An all digital five channel position detection system is to be installed in the Southampton University Magnetic Suspension and Balance System (SUMSBS). The system is intended to monitor a much larger range of model pitch attitudes than has been possible hitherto, up to a maximum of a 90 degree angle of attack. It is based on the use of self-scanning photodiode arrays and illuminating laser light beams, together with purpose built processing electronics. The principles behind the design of the system are discussed, together with the results of testing one channel of the system which was used to control the axial position of a magnetically suspended model in SUMSBS. The removal of optically coupled heave position information from the axial position sensing channel is described.

  14. Vortex pattern development on the upper surface of a swept wing at high angle of attack

    NASA Technical Reports Server (NTRS)

    Mirande, J.; Schmitt, V.; Werle, H.

    1979-01-01

    An experimental study, based on a swept wing, was undertaken in the water tunnel and the wind tunnel at low speeds, with a view to improving the understanding of the intervening phenomena and to make easier their modelling. The vortex flow effects on the wing are first illustrated from global effort measurements and static pressure distributions. The domain of existence of this type of flow is deduced as a function of both sweep angle and angle of attack. By a phenomenological study in the water tunnel, an attempt is made to describe the physical pattern of the vortex flow, from its formation near the apex to its breakdown at the trailing edge. Lastly, by means of a clinometric probe, the flow field over the wing is determined.

  15. Multiaxis Thrust-Vectoring Characteristics of a Model Representative of the F-18 High-Alpha Research Vehicle at Angles of Attack From 0 deg to 70 deg

    NASA Technical Reports Server (NTRS)

    Asbury, Scott C.; Capone, Francis J.

    1995-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the multiaxis thrust-vectoring characteristics of the F-18 High-Alpha Research Vehicle (HARV). A wingtip supported, partially metric, 0.10-scale jet-effects model of an F-18 prototype aircraft was modified with hardware to simulate the thrust-vectoring control system of the HARV. Testing was conducted at free-stream Mach numbers ranging from 0.30 to 0.70, at angles of attack from O' to 70', and at nozzle pressure ratios from 1.0 to approximately 5.0. Results indicate that the thrust-vectoring control system of the HARV can successfully generate multiaxis thrust-vectoring forces and moments. During vectoring, resultant thrust vector angles were always less than the corresponding geometric vane deflection angle and were accompanied by large thrust losses. Significant external flow effects that were dependent on Mach number and angle of attack were noted during vectoring operation. Comparisons of the aerodynamic and propulsive control capabilities of the HARV configuration indicate that substantial gains in controllability are provided by the multiaxis thrust-vectoring control system.

  16. Survey of needs and capabilities for wind tunnel testing of dynamic stability of aircraft at high angles of attack

    NASA Technical Reports Server (NTRS)

    Orlik-Ruckemann, K. J.

    1973-01-01

    A survey was conducted relative to future requirements for dynamic stability information for such aerospace vehicles as the space shuttle and advanced high performance military aircraft. High-angle-of-attack and high-Reynolds number conditions were emphasized. A review was made of the wind-tunnel capabilities in North America for measuring dynamic stability derivatives, revealing an almost total lack of capabilities that could satisfy these requirements. Recommendations are made regarding equipment that should be constructed to remedy this situation. A description is given of some of the more advanced existing capabilities, which can be used to at least partly satisfy immediate demands.

  17. High angle of attack control law development for a free-flight wind tunnel model using direct eigenstructure assignment

    NASA Technical Reports Server (NTRS)

    Wendel, Thomas R.; Boland, Joseph R.; Hahne, David E.

    1991-01-01

    Flight-control laws are developed for a wind-tunnel aircraft model flying at a high angle of attack by using a synthesis technique called direct eigenstructure assignment. The method employs flight guidelines and control-power constraints to develop the control laws, and gain schedules and nonlinear feedback compensation provide a framework for considering the nonlinear nature of the attack angle. Linear and nonlinear evaluations show that the control laws are effective, a conclusion that is further confirmed by a scale model used for free-flight testing.

  18. High angle of attack control law development for a free-flight wind tunnel model using direct eigenstructure assignment

    NASA Technical Reports Server (NTRS)

    Wendel, Thomas R.; Boland, Joseph R.; Hahne, David E.

    1991-01-01

    Flight-control laws are developed for a wind-tunnel aircraft model flying at a high angle of attack by using a synthesis technique called direct eigenstructure assignment. The method employs flight guidelines and control-power constraints to develop the control laws, and gain schedules and nonlinear feedback compensation provide a framework for considering the nonlinear nature of the attack angle. Linear and nonlinear evaluations show that the control laws are effective, a conclusion that is further confirmed by a scale model used for free-flight testing.

  19. Viscous shock-layer predictions of three-dimensional nonequilibrium flows past the space shuttle at high angle of attack

    NASA Technical Reports Server (NTRS)

    Kim, M. D.; Swaminathan, S.; Lewis, C. H.

    1983-01-01

    Computational solutions were obtained for chemically reacting flowfields over the entire windward surface of the Space Shuttle at high angle of attack. The recently developed computational method for the Space Shuttle is capable of treating three dimensional viscous nonequilibrium air flow as well as equilibrium air and perfect gas flows. A general nonorthogonal computational grid system is used to treat the nonaxisymmetric geometry. Boundary conditions take into account noncatalytic wall, equilibrium with noncatalytic wall condition are compared to the fully catalytic wall solutions, the equilibrium air solutions, the perfect gas solutions, and also the Shuttle flight heating and pressure data. The comparisons show good agreement and correlations in most cases.

  20. Development of microprocessor-based laser velocimeter and its application to measurement of jet exhausts and flows over missiles at high angles of attack

    NASA Astrophysics Data System (ADS)

    Harwell, K. E.; Farmer, W. M.; Hornkohl, J. O.; Stallings, E.

    1981-03-01

    During the past three years, personnel have developed a unique three-component laser velocimeter for the in situ measurement of particle and/or gas velocities in flow fields produced behind bodies at high angles of attack and in jet exhaust plumes. This report describes the development of the laser velocimeter and its subsequent application of the measurement of the velocity distribution and vortex structure in free jets and in flows over missiles at high angles of attack.

  1. In-flight flow visualization characteristics of the NASA F-18 high alpha research vehicle at high angles of attack

    NASA Technical Reports Server (NTRS)

    Fisher, David F.; Delfrate, John H.; Richwine, David M.

    1991-01-01

    Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.

  2. Experimental aerodynamic characteristics for slender bodies with thin wings at angles of attack from 0 deg to 58 deg and Mach numbers from 0.6 to 2.0

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.; Howell, M. H.

    1976-01-01

    An experimental investigation was conducted in the Ames 6-by-6-Foot Wind Tunnel to measure the static aerodynamic characteristics for bodies of circular and elliptic cross section with various thin flat-plate wings. Eighteen configuration combinations were tested at Mach numbers of 0.6, 0.9, 1.2, 1.5, and 2.0 at angles of attack from 0 deg to 58 deg. The data demonstrate that taper ratio and aspect ratio had only small effect on the aerodynamic characteristics, especially at the higher angles of attack. Undesirable side forces and yawing moments, which developed at angles of attack greater than about 25 deg, were generally no greater than those for the bodies tested alone. As for the bodies alone, the side forces and yawing moments increased as the nose fineness ratio increased and/or as the subsonic Mach number decreased.

  3. Correlation of forebody pressures and aircraft yawing moments on the X-29A aircraft at high angles of attack

    NASA Technical Reports Server (NTRS)

    Fisher, David F.; Richwine, David M.; Landers, Stephen

    1992-01-01

    In-flight pressure distributions at four fuselage stations on the forebody of the X-29A aircraft have been reported at angles of attack from 15 to 66 deg and at Mach numbers from 0.22 to 0.60. At angles of attack of 20 deg and higher, vortices shed from the nose strake caused suction peaks in the pressure distributions that generally increased in magnitude with angle of attack. Above 30 deg-angle of attack, the forebody pressure distributions became asymmetrical at the most forward station, while they remained nearly symmetrical until 50 to 55 deg-angle of attack for the aft stations. Between 59 to 66 deg-angle of attack, the asymmetry of the pressure distributions changed direction. Yawing moments for the forebody alone were obtained by integrating the forebody pressure distributions. At 45 deg-angle of attack, the aircraft yaws to the right and at 50 deg and higher, the aircraft yaws to the left. The forebody yawing moments correlated well with the aircraft left yawing moment at an angle of attack of 50 deg or higher. At a 45 deg-angle of attack, the forebody yawing moments did not correlate well with the aircraft yawing moment, but it is suggested that this was due to asymmetric pressures on the cockpit region of the fuselage which was not instrumented. The forebody was also shown to provide a positive component of directional stability of the aircraft at angles of attack of 25 deg or higher. A Mach number effect was noted at angles of attack of 30 deg or higher at the station where the nose strake was present. At this station, the suction peaks in the pressure distributions at the highest Mach number were reduced and much more symmetrical as compared to the lower Mach number pressure distributions.

  4. Mathematical description of nonstationary aerodynamic characteristics of a passenger aircraft model in longitudinal motion at large angles of attack

    NASA Astrophysics Data System (ADS)

    Petoshin, V. I.; Chasovnikov, E. A.

    2011-05-01

    Aerodynamic loads in problems of flight dynamics of passenger aircraft in stalled flow regimes are described using a mathematical model that includes an ordinary linear first-order differential equation. A procedure for determining the parameters of the mathematical model is proposed which is based on approximating experimental frequency characteristics with the frequency characteristics of the linearized mathematical model. The mathematical model was verified by tests of a modern passenger aircraft model in a wind tunnel.

  5. Experimental investigations on airfoils with different geometries in the domain of high angles of attack-flow separation

    NASA Technical Reports Server (NTRS)

    Keil, J.

    1985-01-01

    Wind tunnel tests were conducted on airfoil models in order to study the flow separation phenomena occurring for high angles of attack. Pressure distribution on wings of different geometries were measured. Results show that for three-dimensional airfoils layout and span lift play a role. Separation effects on airfoils with moderate extension are three-dimensional. The flow domains separated from the air foil must be treated three-dimensionally. The rolling-up of separated vortex layers increases with angle in intensity and induction effect and shows strong nonlinearities. Boundary layer material moves perpendicularly to the flow direction due to the pressure gradients at the airfoil; this has a stabilizing effect. The separation starts earlier with increasing pointed profiles.

  6. Development of a pneumatic high-angle-of-attack Flush Airdata Sensing (HI-FADS) system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Leondes, Cornelius T.

    1992-01-01

    The HI-FADS system design is an evolution of the FADS systems (e.g., Larson et al., 1980, 1987), which emphasizes the entire airdata system development. This paper describes the HI-FADS measurement system, with particular consideration given to the basic measurement hardware and the development of the HI-FADS aerodynamic model and the basic nonlinear regression algorithm. Algorithm initialization techniques are developed, and potential algorithm divergence problems are discussed. Data derived from HI-FADS flight tests are used to demonstrate the system accuracies and to illustrate the developed concepts and methods.

  7. Development of a pneumatic high-angle-of-attack Flush Airdata Sensing (HI-FADS) system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Leondes, Cornelius T.

    1992-01-01

    The HI-FADS system design is an evolution of the FADS systems (e.g., Larson et al., 1980, 1987), which emphasizes the entire airdata system development. This paper describes the HI-FADS measurement system, with particular consideration given to the basic measurement hardware and the development of the HI-FADS aerodynamic model and the basic nonlinear regression algorithm. Algorithm initialization techniques are developed, and potential algorithm divergence problems are discussed. Data derived from HI-FADS flight tests are used to demonstrate the system accuracies and to illustrate the developed concepts and methods.

  8. Aerodynamic Characteristics of a Circular Cylinder at Mach Number of 6.86 and Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Penland, Jim A

    1954-01-01

    Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86, a Reynolds number of 129,000 based on diameter, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and with a simple modification of the Newtonian flow theory. The comparison of experimental results shows that either theory gives adequate general aerodynamic characteristics but that the modified Newtonian theory gives a more accurate prediction of the pressure distribution. The calculated crossflow drag coefficients plotted as a function of crossflow Mach number were found to be in reasonable agreement with similar results obtained from other investigations at lower supersonic Mach numbers. Comparison of the results of this investigation with data obtained at a lower Mach number indicates that the drag coefficient of a cylinder normal to the flow is relatively constant for Mach numbers above about 4.

  9. Simulation model of the F/A-18 high angle-of-attack research vehicle utilized for the design of advanced control laws

    NASA Technical Reports Server (NTRS)

    Strickland, Mark E.; Bundick, W. Thomas; Messina, Michael D.; Hoffler, Keith D.; Carzoo, Susan W.; Yeager, Jessie C.; Beissner, Fred L., Jr.

    1996-01-01

    The 'f18harv' six degree-of-freedom nonlinear batch simulation used to support research in advanced control laws and flight dynamics issues as part of NASA's High Alpha Technology Program is described in this report. This simulation models an F/A-18 airplane modified to incorporate a multi-axis thrust-vectoring system for augmented pitch and yaw control power and actuated forebody strakes for enhanced aerodynamic yaw control power. The modified configuration is known as the High Alpha Research Vehicle (HARV). The 'f18harv' simulation was an outgrowth of the 'f18bas' simulation which modeled the basic F/A-18 with a preliminary version of a thrust-vectoring system designed for the HARV. The preliminary version consisted of two thrust-vectoring vanes per engine nozzle compared with the three vanes per engine actually employed on the F/A-18 HARV. The modeled flight envelope is extensive in that the aerodynamic database covers an angle-of-attack range of -10 degrees to +90 degrees, sideslip range of -20 degrees to +20 degrees, a Mach Number range between 0.0 and 2.0, and an altitude range between 0 and 60,000 feet.

  10. Estimation of Static Longitudinal Stability of Aircraft Configurations at High Mach Numbers and at Angles of Attack Between 0 deg and +/-180 deg

    NASA Technical Reports Server (NTRS)

    Dugan, Duane W.

    1959-01-01

    The possibility of obtaining useful estimates of the static longitudinal stability of aircraft flying at high supersonic Mach numbers at angles of attack between 0 and +/-180 deg is explored. Existing theories, empirical formulas, and graphical procedures are employed to estimate the normal-force and pitching-moment characteristics of an example airplane configuration consisting of an ogive-cylinder body, trapezoidal wing, and cruciform trapezoidal tail. Existing wind-tunnel data for this configuration at a Mach number of 6.86 provide an evaluation of the estimates up to an angle of attack of 35 deg. Evaluation at higher angles of attack is afforded by data obtained from wind-tunnel tests made with the same configuration at angles of attack between 30 and 150 deg at five Mach numbers between 2.5 and 3.55. Over the ranges of Mach numbers and angles of attack investigated, predictions of normal force and center-of-pressure locations for the configuration considered agree well with those obtained experimentally, particularly at the higher Mach numbers.

  11. Low-speed wind-tunnel study of the high-angle-of-attack stability and control characteristics of a cranked-arrow-wing fighter configuration

    NASA Technical Reports Server (NTRS)

    Grafton, S. B.

    1984-01-01

    The low-speed, high-angle-of-attack stability and control characteristics of a fighter configuration incorporating a cranked arrow wing were investigated in the Langley 30- by 60-foot tunnel as part of a NASA/General Dynamics cooperative research program to investigate the application of advanced wing designs to combat aircraft. Tests were conducted on a baseline configuration and on several modified configurations. The results show that the baseline configuration exhibited a high level of maximum lift but displayed undesirable longitudinal and lateral-directional stability characteristics at high angles of attack. Various wing modifications were made which improved the longitudinal and lateral-directional stability characteristics of the configuration at high angles of attack. However, most of the modifications were detrimental to maximum lift.

  12. Calculation of inviscid flow over shuttle-like vehicles at high angles of attack and comparisons with experimental data

    NASA Technical Reports Server (NTRS)

    Weilmuenster, K. J.; Hamilton, H. H., II

    1983-01-01

    A computer code HALIS, designed to compute the three dimensional flow about shuttle like configurations at angles of attack greater than 25 deg, is described. Results from HALIS are compared where possible with an existing flow field code; such comparisons show excellent agreement. Also, HALIS results are compared with experimental pressure distributions on shuttle models over a wide range of angle of attack. These comparisons are excellent. It is demonstrated that the HALIS code can incorporate equilibrium air chemistry in flow field computations.

  13. Wind-tunnel static and free-flight investigation of high-angle-of-attack stability and control characteristics of a model of the EA-6B airplane

    NASA Technical Reports Server (NTRS)

    Jordan, Frank L., Jr.; Hahne, David E.

    1992-01-01

    An investigation was conducted in the Langley 30- by 60-Foot Tunnel and the Langley 12-Foot Low-Speed Tunnel to identify factors contributing to a directional divergence at high angles of attack for the EA-6B airplane. The study consisted of static wind-tunnel tests, smoke and tuft flow-visualization tests, and free-flight tests of a 1/8.5-scale model of the airplane. The results of the investigation indicate that the directional divergence of the airplane is brought about by a loss of directional stability and effective dihedral at high angles of attack. Several modifications were tested that significantly alleviate the stability problem. The results of the free-flight study show that the modified configuration exhibits good dynamic stability characteristics and could be flown at angles of attack significantly higher than those of the unmodified configuration.

  14. Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Kim, Hyun D.; Frate, Franco C.

    2001-01-01

    A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

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

    NASA Technical Reports Server (NTRS)

    Sternberg, Charles A.; Traven, Ricardo; Lackey, James B.

    1994-01-01

    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.

  16. Flight-Determined, Subsonic, Lateral-Directional Stability and Control Derivatives of the Thrust-Vectoring F-18 High Angle of Attack Research Vehicle (HARV), and Comparisons to the Basic F-18 and Predicted Derivatives

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Wang, Kon-Sheng Charles

    1999-01-01

    The subsonic, lateral-directional, stability and control derivatives of the thrust-vectoring F-1 8 High Angle of Attack Research Vehicle (HARV) are extracted from flight data using a maximum likelihood parameter identification technique. State noise is accounted for in the identification formulation and is used to model the uncommanded forcing functions caused by unsteady aerodynamics. Preprogrammed maneuvers provided independent control surface inputs, eliminating problems of identifiability related to correlations between the aircraft controls and states. The HARV derivatives are plotted as functions of angles of attack between 10deg and 70deg and compared to flight estimates from the basic F-18 aircraft and to predictions from ground and wind tunnel tests. Unlike maneuvers of the basic F-18 aircraft, the HARV maneuvers were very precise and repeatable, resulting in tightly clustered estimates with small uncertainty levels. Significant differences were found between flight and prediction; however, some of these differences may be attributed to differences in the range of sideslip or input amplitude over which a given derivative was evaluated, and to differences between the HARV external configuration and that of the basic F-18 aircraft, upon which most of the prediction was based. Some HARV derivative fairings have been adjusted using basic F-18 derivatives (with low uncertainties) to help account for differences in variable ranges and the lack of HARV maneuvers at certain angles of attack.

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

    NASA Technical Reports Server (NTRS)

    Schiff, Lewis B.; Cummings, Russell M.; Sorenson, Reese L.; Rizk, Yehia M.

    1989-01-01

    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.

  18. The geometry of high angle of attack maneuvers and the implications for Gy-induced neck injuries.

    PubMed

    Newman, David G; Ostler, David

    2011-08-01

    Modern super agile fighter aircraft have significantly expanded maneuverability envelopes, often involving very high angles of attack (AOA) in the post-stall region. One such maneuver is the high AOA velocity vector roll. The geometry of this flight maneuver is such that during the roll there is a significant lateral C load imposed on the unrestrained head-neck complex of the pilot. A mathematical analysis of the geometric relationship determining the magnitude of +/- Gy acceleration during high AOA maneuvering was conducted. This preliminary mathematical model is able to predict the Gy load imposed on the head-neck complex of the pilot for a given set of flight maneuver parameters. The analysis predicts that at an AOA of 700 and with a roll rate of 100 degrees x s(-1), the lateral G developed will be approximately 3.5 Gy. Increasing the roll rate increases the lateral G component: at 200 degrees x s(-1) the Gy, load is more than 6 Gy. There are serious potential implications of super agile maneuvers on the neck of the pilot. The G environment experienced by the pilot of super agile aircraft is increasingly multiaxial, involving +/- Gx, +/- Gy, and +/- Gz. The level of lateral G developed during these dynamic flight maneuvers should not be underestimated, as such G loads can potentially lead to neck injuries. While aircraft become ever more capable, a full understanding of the biodynamic effects on the pilot while exploiting the agility of the aircraft still needs to be developed.

  19. Investigation on Stability in Roll of Square Section Missile at High Angle of Attack

    NASA Astrophysics Data System (ADS)

    Tao, Yang; Fan, Zhaolin; Wu, Jifei; Wu, Wenhua

    An experimental investigation of the stability in roll of a square section missile at high incidence was conducted in FL-23 wind tunnel. Dynamic motions were obtained on a square section missile that is free to rotate about its longitudinal axis. Different dynamic rolling motions were observed depending on the incidence of the model sting. These dynamic regimes include damped oscillations, quasi-limit-cycle wing-rock motion, and constant rolling. A coupling numerical method was established by solving the fluid dynamics equations and the rigid-body dynamics equations synchronously in order to predict the onset and the development of uncommented motions and then explore the unsteady movement characteristics of the aircraft. The study indicates that the aircraft loss stability at high incidence is caused by the asymmetric vertex on the level fin tip liftoff and attach alternately. The computation results are in line with the experiment results.

  20. Component build-up method for engineering analysis of missiles at low-to-high angles of attack

    NASA Technical Reports Server (NTRS)

    Hemsch, Michael J.

    1992-01-01

    Methods are presented for estimating the component build-up terms, with the exception of zero-lift drag, for missile airframes in steady flow and at arbitrary angles of attack and bank. The underlying and unifying bases of all these efforts are slender-body theory and its nonlinear extensions through the equivalent angle-of-attack concept. Emphasis is placed on the forces and moments which act on each of the fins, so that control cross-coupling effects as well as longitudinal and lateral-directional effects can be determined.

  1. Correlation of forebody pressures and aircraft yawing moments on the X-29A aircraft at high angles of attack

    NASA Technical Reports Server (NTRS)

    Fisher, David F.; Richwine, David M.; Landers, Stephen

    1992-01-01

    In-flight pressure distributions are presented at angles of attack from 15 deg to 66 deg and at Mach numbers from 0.22 to 0.60 at four fuselage stations on the forebody of the X-29A aircraft. Forebody yawing moments are obtained from the integrated pressure distributions and the results are correlated with the overall aircraft yawing moments. Yawing moments created by the forebody were not significant until an angle of attack of 50 deg or above and correlated well with the aircraft left yawing moment.

  2. Effects of fuselage forebody geometry on low-speed lateral-directional characteristics of twin-tail fighter model at high angles of attack

    NASA Technical Reports Server (NTRS)

    Carr, P. C.; Gilbert, W. P.

    1979-01-01

    Low-speed, static wind-tunnel tests were conducted to explore the effects of fighter fuselage forebody geometry on lateral-directional characteristics at high angles of attack and to provide data for general design procedures. Effects of eight different forebody configurations and several add-on devices (e.g., nose strakes, boundary-layer trip wires, and nose booms) were investigated. Tests showed that forebody design features such as fineness ratio, cross-sectional shape, and add-on devices can have a significant influence on both lateral-directional and longitudinal aerodynamic stability. Several of the forebodies produced both lateral-directional symmetry and strong favorable changes in lateral-directional stability. However, the same results also indicated that such forebody designs can produce significant reductions in longitudinal stability near maximum lift and can significantly change the influence of other configuration variables. The addition of devices to highly tailored forebody designs also can significantly degrade the stability improvements provided by the clean forebody.

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

    NASA Technical Reports Server (NTRS)

    Steenken, W. G.; Williams, J. G.; Yuhas, A. J.; Walsh, K. R.

    1999-01-01

    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.

  4. Transonic Navier-Stokes wing solutions using a zonal approach. Part 2: High angle-of-attack simulation

    NASA Technical Reports Server (NTRS)

    Chaderjian, Neal M.

    1986-01-01

    A computer code is under development whereby the thin-layer Reynolds-averaged Navier-Stokes equations are to be applied to realistic fighter aircraft configurations. This transonic Navier-Stokes code (TNS) utilizes a zonal approach in order to treat complex geometries and satisfy in-core computer memory constraints. The zonal approach was applied to isolated wing geometries in order to facilitate code development. The TNS finite difference algorithm, zonal methodology, and code validation with experimental data is addressed. Also addressed are some numerical issues such as code robustness, efficiency, and accuracy at high angles of attack. Special free-stream-preserving metrics proved an effective way to treat H-mesh singularities over a large range of severe flow conditions, including strong leading edge flow gradients, massive shock induced separation, and stall. Furthermore, lift and drag coefficients were computed for a wing up through CLmax. Numerical oil flow patterns and particle trajectories are presented both for subcritical and transonic flow. These flow simulations are rich with complex separated flow physics and demonstrate the efficiency and robustness of the zonal approach.

  5. Transonic Navier-Stokes wing solutions using a zonal approach. Part 2: High angle-of-attack simulation

    NASA Technical Reports Server (NTRS)

    Chaderjian, N. M.

    1986-01-01

    A computer code is under development whereby the thin-layer Reynolds-averaged Navier-Stokes equations are to be applied to realistic fighter-aircraft configurations. This transonic Navier-Stokes code (TNS) utilizes a zonal approach in order to treat complex geometries and satisfy in-core computer memory constraints. The zonal approach has been applied to isolated wing geometries in order to facilitate code development. Part 1 of this paper addresses the TNS finite-difference algorithm, zonal methodology, and code validation with experimental data. Part 2 of this paper addresses some numerical issues such as code robustness, efficiency, and accuracy at high angles of attack. Special free-stream-preserving metrics proved an effective way to treat H-mesh singularities over a large range of severe flow conditions, including strong leading-edge flow gradients, massive shock-induced separation, and stall. Furthermore, lift and drag coefficients have been computed for a wing up through CLmax. Numerical oil flow patterns and particle trajectories are presented both for subcritical and transonic flow. These flow simulations are rich with complex separated flow physics and demonstrate the efficiency and robustness of the zonal approach.

  6. Flow-visualization study of the X-29A aircraft at high angles of attack using a 1/48-scale model

    NASA Technical Reports Server (NTRS)

    Cotton, Stacey J.; Bjarke, Lisa J.

    1994-01-01

    A water-tunnel study on a 1/48-scale model of the X-29A aircraft was performed at the NASA Dryden Flow Visualization Facility. The water-tunnel test enhanced the results of the X-29A flight tests by providing flow-visualization data for comparison and insights into the aerodynamic characteristics of the aircraft. The model was placed in the water tunnel at angles of attack of 20 to 55 deg. and with angles of sideslip from 0 to 5 deg. In general, flow-visualization techniques provided useful information on vortex formation, separation, and breakdown and their role in yaw asymmetries and tail buffeting. Asymmetric forebody vortices were observed at angles of attack greater than 30 deg. with 0 deg. sideslip and greater than 20 deg. with 5 deg. sideslip. While the asymmetric flows observed in the water tunnel did not agree fully with the flight data, they did show some of the same trends. In addition, the flow visualization indicated that the interaction of forebody vortices and the wing wake at angles of attack between 20 and 35 deg. may cause vertical-tail buffeting observed in flight.

  7. Wind tunnel research comparing lateral control devices, particularly at high angles of attack X : various control devices on a wing with a fixed auxiliary airfoil

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Noyes, Richard W

    1933-01-01

    Results are given of a series of systemic tests comparing lateral control devices with particular reference to their effectiveness at high angles of attack. These tests were made with two sizes of ordinary ailerons and different sizes of spoilers on a Clark Y wing model having a narrow auxiliary airfoil fixed ahead and above the leading edge, the chords of the main and auxiliary airfoils being parallel. In addition, the auxiliary airfoil itself was given angular deflection. The purpose was to provide rolling moments for lateral control. The tests were made in a 7 by 10 foot wind tunnel. They included both force and rotation tests to show the effect of the devices on the lift and drag characteristics of the wing and on the lateral stability characteristics, as well as lateral control. They showed that none of the aileron arrangements tried would give rolling control of an assumed satisfactory value at all angles of attack up to the stall. However, they would give satisfactory values, but at the expense of abnormally high deflections and very heavy hinge moments. The most effective combination of ailerons and spoilers gave satisfactory values of rolling moment at angles of attack below the stall, and the values did not fall off as rapidly above the stall as with ailerons alone. With an arrangement of this type having the proper relative proportions and linkage, it should be possible to obtain reasonably satisfactory yawing moments and control forces. Deflecting one-half of the auxiliary airfoil downward for the purpose of control gave strong favorable yawing moments at all angles of attack, but gave very small rolling moments at the low angles of attack.

  8. Pressures on a Slender, Axisymmetric Body at High Angle of Attack in a Very Low Turbulence Level Air Stream.

    DTIC Science & Technology

    1979-09-01

    previously associated with high incidence aerodynamics can be overcome, by ensuring that tests are always carried out in this ’regular’ state. This may...on high incidence aerodynamics contains so many anomalies. If tests have been performed at arcitrary roll angles then the results are likely to be...the model and coat it in soap solution. It has also been found that at test speeds corresponding to the Reynolds numbers used in the pressure tests more

  9. Space shuttle: Stability and control effectiveness at high and low angles of attack and effects of variations in engine shround, fin, and drag petal configurations for the Boeing 0.008899-scale pressure-fed ballistic recoverable booster, model 979-160

    NASA Technical Reports Server (NTRS)

    Hanson, R. L.; Obrien, R. G.; Oiye, M. Y.; Vanderleest, S.

    1972-01-01

    Experimental aerodynamic investigations were carried out in the Boeing transonic and supersonic wind tunnels on a 0.008899-scale model of a proposed pressure-fed ballistic recoverable booster (BRB) configuration. The purpose of the test program was to determine the stability and control effectiveness of the basic configuration at high and low angles of attack, and to conduct parametric studies of various engine shroud, fin, and drag petal configurations. Six-component force data and base pressure data were obtained over a Mach number range of 0.35 to 4.0 at angles of attack of -5 to 25 and 55 to 85 at zero degrees sideslip and over a sideslip range of -10 to +10 at angles of attack ranging from -10 to 72.5. Two-component force data were also obtained with a fin balance on selected runs.

  10. Turbofan blade stresses induced by the flow distortion of a VTOL inlet at high angles of attack

    NASA Technical Reports Server (NTRS)

    Williams, R. C.; Diedrich, J. H.; Shaw, R. J.

    1983-01-01

    A 51-cm-diameter turbofan with a tilt-nacelle VTOL inlet was tested in the Lewis Research Center's 9- by 15-Ft Low Speed Wind Tunnel at velocities up to 72 m/s and angles of attack up to 120 deg. Fan-blade vibratory stress levels were investigated over a full aircraft operating range. These stresses were due to inlet air flow distortion resulting from (1) internal flow separation in the inlet, and (2) ingestion of the exterior nacelle wake. Stress levels are presented, along with an estimated safe operating envelope, based on infinite blade fatigue life.

  11. Wind tunnel research comparing lateral control devices, particularly at high angles of attack XI : various floating tip ailerons on both rectangular and tapered wings

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Harris, Thomas A

    1933-01-01

    Discussed here are a series of systematic tests being conducted to compare different lateral control devices with particular reference to their effectiveness at high angles of attack. The present tests were made with six different forms of floating tip ailerons of symmetrical section. The tests showed the effect of the various ailerons on the general performance characteristics of the wing, and on the lateral controllability and stability characteristics. In addition, the hinge moments were measured for the most interesting cases. The results are compared with those for a rectangular wing with ordinary ailerons and also with those for a rectangular wing having full-chord floating tip ailerons. Practically all the floating tip ailerons gave satisfactory rolling moments at all angles of attack and at the same time gave no adverse yawing moments of appreciable magnitude. The general performance characteristics with the floating tip ailerons, however, were relatively poor, especially the rate of climb. None of the floating tip ailerons entirely eliminated the auto rotational moments at angles of attack above the stall, but all of them gave lower moments than a plain wing. Some of the floating ailerons fluttered if given sufficiently large deflection, but this could have been eliminated by moving the hinge axis of the ailerons forward. Considering all points including hinge moments, the floating tip ailerons on the wing with 5:1 taper are probably the best of those which were tested.

  12. Aerodynamic Characteristics of Missile Configurations with Wings of Low Aspect Ratio for Various Combinations of Forebodies, Afterbodies, and Nose Shapes for Combined Angles of Attack and Sideslip at a Mach Number of 2.01

    NASA Technical Reports Server (NTRS)

    Robinson, Ross B

    1957-01-01

    An investigation has been made in the Langley 4-by-4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a series of missile configurations having low-aspect-ratio wings at a Mach number of 2.01. The effects of wing plan form and size, length-diameter ratio, forebody and afterbody length, boattailed and flared afterbodies, and component force and moment data are presented for combined angles of attack and sideslip to about 28 degrees. No analysis of the data was made in this report.

  13. A method for estimating static aerodynamic characteristics for slender bodies of circular and noncircular cross section alone and with lifting surfaces at angles of attack from 0 deg to 90 deg

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1973-01-01

    An engineering-type method is presented for estimating normal-force, axial-force, and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. Static aerodynamic characteristics computed by the method are shown to agree closely with experimental results for slender bodies of circular and elliptic cross section and for winged-circular and winged-elliptic cones. However, the present experimental results used for comparison with the method are limited to angles of attack only up to about 20 deg and Mach numbers from 2 to 4.

  14. Comparison of X-31 flight, wind-tunnel, and water-tunnel yawing moment asymmetries at high angles of attack

    NASA Technical Reports Server (NTRS)

    Cobleigh, Brent R.; Croom, Mark A.; Tamrat, B. F.

    1994-01-01

    The X-31 aircraft are being used in the enhanced fighter maneuverability (EFM) research program, which is jointly funded by the (U.S.) Advanced Research Projects Agency (ARPA) and Germany's Federal Ministry of Defense (FMOD). The flight test portion of the program, which involves two aircraft, is being conducted by an International Test Organization (ITO) comprising the National Aeronautics and Space Administration (NASA), the U.S. Navy, the U.S. Air Force, Rockwell International, and Deutsche Aerospace (DASA). The goals of the flight program are to demonstrate EFM technologies, investigate close-in-combat exchange ratios, develop design requirements, build a database for application to future fighter aircraft, and develop and validate low-cost prototype concepts. For longitudinal control the X-31 uses canards, symmetrical movement of the trailing-edge flaps, and pitch deflection of the thrust vectoring system. The trim, inertial coupling, and engine gyroscopic coupling compensation tasks are performed primarily by the trailing-edge flaps. For lateral-directional control the aircraft uses differential deflection of the trailing-edge flaps for roll coordination and a conventional rudder combined with the thrust vectoring system to provide yaw control. The rudder is only effective up to about 40 deg angle of attack (alpha), after which the thrust vectoring becomes the primary yaw control effector. Both the leading-edge flaps and the inlet lip are scheduled with the angle of attack to provide best performance.

  15. Investigation of the effect of vehicle, angle-of-attack, and trim elevon position on lateral-directional aerodynamic parameters of the Shuttle Orbiter

    NASA Technical Reports Server (NTRS)

    Schiess, J. R.; Suit, W. T.; Scallion, W. I.

    1984-01-01

    This paper discusses the lateral-directional stability and control coefficients extracted from onboard measurements made during reentry of flights two to eight of the Space Shuttle Orbiters. Maximum likelihood estimation is applied to data derived from accelerometer and rate gyro measurements and trajectory, meteorological and control surface data to estimate these lateral-directional parameters. Comparison of estimated coefficients across the seven flights and two shuttle vehicles is made. The effect of the angle-of-attack, and trim elevon deflection on estimated values of the stability and control derivatives is also studied.

  16. A modification to linearized theory for prediction of pressure loadings on lifting surfaces at high supersonic Mach numbers and large angles of attack

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.

    1979-01-01

    A new linearized-theory pressure-coefficient formulation was studied. The new formulation is intended to provide more accurate estimates of detailed pressure loadings for improved stability analysis and for analysis of critical structural design conditions. The approach is based on the use of oblique-shock and Prandtl-Meyer expansion relationships for accurate representation of the variation of pressures with surface slopes in two-dimensional flow and linearized-theory perturbation velocities for evaluation of local three-dimensional aerodynamic interference effects. The applicability and limitations of the modification to linearized theory are illustrated through comparisons with experimental pressure distributions for delta wings covering a Mach number range from 1.45 to 4.60 and angles of attack from 0 to 25 degrees.

  17. Computation of hypersonic laminar viscous flow past spinning sharp and blunt cones at high angle of attack

    NASA Technical Reports Server (NTRS)

    Agarwal, R.; Rakich, J. V.

    1978-01-01

    Computational results, obtained with a parabolic Navier-Stokes marching code, are presented for hypersonic viscous flow past spinning sharp and blunt cones at angle of attack. The code takes into account the asymmetries in the flow field resulting from spinning motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results. In addition, a new criterion for defining crossflow separation behind spinning bodies is introduced which generalizes the Moore-Rott-Sears criterion for two-dimensional unsteady separation. A condition which characterizes the onset of separation in the flow field is defined.

  18. Heating analysis of bent-nose biconics at high angles of attack using the parabolized Navier-Stokes equations

    NASA Technical Reports Server (NTRS)

    Stephenson, B. L.; Hassan, H. A.

    1983-01-01

    A method based on the Parabolized Navier-Stokes equations is used to calculate the flow field and heat transfer of lifting entry vehicles. The method is based on the Bean and Warming implicit algorithm and uses a new procedure for preventing departure solutions. Calculations are carried out for blunt on-axis and bent biconics, assuming a perfect gas and laminar flow, and compared with available heat transfer, surface pressure and shock shape measurements for a range of Mach numbers and angles of attack. In all calculations presented here, the starting solution is obtained from available inviscid and boundary layer codes. Good agreement with experiment is indicated. Thus, the method provides an accurate and rather inexpensive procedure for calculating three-dimensional flows at supersonic Mach numbers.

  19. Numerical simulation of the viscous flow around a simplified F/A-18 at high angles of attack

    NASA Technical Reports Server (NTRS)

    Rizk, Yehia M.; Schiff, Lewis B.; Gee, Ken

    1990-01-01

    A numerical method developed for solving the viscous flow around three-dimensional complex configurations is presently used to simulate the flow around a simplified F/A-18 configuration encompassing forebody, wing, leading-edge extension, faired-over inlet, and deflected wing leading-edge flaps, at Mach 0.243 and 30.3 deg angle of attack. The computational results show the details of the flowfield structure, including primary, secondary, and tertiary separation lines, the development of forebody and leading-edge extension vortex, and the burst of this vortex. A grid-refinement study is conducted to assess the effect of grid characteristics on solution accuracy. Substantial agreement is obtained between these results and flight data.

  20. Computation of hypersonic laminar viscous flow past spinning sharp and blunt cones at high angle of attack

    NASA Technical Reports Server (NTRS)

    Agarwal, R.; Rakich, J. V.

    1978-01-01

    Computational results, obtained with a parabolic Navier-Stokes marching code, are presented for hypersonic viscous flow past spinning sharp and blunt cones at angle of attack. The code takes into account the asymmetries in the flow field resulting from spinning motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results. In addition, a new criterion for defining crossflow separation behind spinning bodies is introduced which generalizes the Moore-Rott-Sears criterion for two-dimensional unsteady separation. A condition which characterizes the onset of separation in the flow field is defined.

  1. Prediction of Wind Tunnel Effects on the Installed F/A-18A Inlet Flow Field at High Angles-of-attack

    NASA Technical Reports Server (NTRS)

    Smith, Crawford F.

    1995-01-01

    NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18A, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Wind tunnel tests were a major component of the Lewis program. Since the available wind tunnel was small (9 x 15 ft) as compared to the scale of the model of the F/A-18A (19.78 percent), there were questions about the capability to obtain useful inlet performance data. The blockage effects were expected to be very large. This report represents the results of an analysis to determine how the wind tunnel walls effect inlet performance at several angles of attack. The predictions for the external particle traces along the fuselage indicate the influence of the wind tunnel side wall under the model is greater at 30 deg angle of attack than at 50 deg angle of attack on the under Leading Edge Extension (LEX) vortex trajectory. The side wall above the model appears to have negligible influence on the under LEX vortex. This may be due to the LEX acting as 'shield' to the upper wall effects. As expected, the wind tunnel has a significant influence on the external forces. The lift and drag coefficients increase significantly for the wind tunnel model as compared to free stream conditions. The wind tunnel had a small effect on the inlet recovery and on inlet total pressure distortion patterns. The predicted recoveries for the wind tunnel model are within one percentage point of the model recoveries in

  2. Test data report: Low speed wind tunnel tests of a full scale, fixed geometry inlet, with engine, at high angles of attack

    NASA Technical Reports Server (NTRS)

    Shain, W. M.

    1976-01-01

    A full scale inlet test was to be done in the NASA-ARC 40' X 80' WT to demonstrate satisfactory inlet performance at high angles of attack. The inlet was designed to match a Hamilton-Standard 55 inch, variable pitch fan, driven by a Lycoming T55-L-11A gas generator. The test was installed in the wind tunnel on two separate occasions, but mechanical failures in the fan drive gear box early in each period terminated testing. A detailed description is included of the Model, installation, instrumentation and data reduction procedures.

  3. Wind-tunnel research comparing lateral control devices, particularly at high angles of attack IV : floating tip ailerons on rectangular wings

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Harris, Thomas A

    1933-01-01

    This report is the fourth of a series on systematic tests conducted which compares lateral control devices with particular reference to their effectiveness at high angles of attack. The present report covers tests with floating tip ailerons on rectangular Clark y wings. Ailerons of two profiles were tested - symmetrical and Clark y, both with adjustable trailing-edge flaps. Each form was tested at three hinge-axis locations, both with and without vertical end plates between the ailerons and the wing proper. The results from these tests are compared with the results from tests on a wing of the same over-all size equipped with average-sized ordinary ailerons.

  4. Ionospheric plasma flow over large high-voltage space platforms. I - Ion-plasma-time scale interactions of a plate at zero angle of attack. II - The formation and structure of plasma wake

    NASA Technical Reports Server (NTRS)

    Wang, J.; Hastings, D. E.

    1992-01-01

    The paper presents the theory and particle simulation results for the ionospheric plasma flow over a large high-voltage space platform at a zero angle of attack and at a large angle of attack. Emphasis is placed on the structures in the large, high-voltage regime and the transient plasma response on the ion-plasma time scale. Special consideration is given to the transient formation of the space-charge wake and its steady-state structure.

  5. Flow analysis for the nacelle of an advanced ducted propeller at high angle-of-attack and at cruise with boundary layer control

    NASA Technical Reports Server (NTRS)

    Hwang, D. P.; Boldman, D. R.; Hughes, C. E.

    1994-01-01

    An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

  6. An investigation of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Gittner, Nathan M.

    1992-01-01

    An experimental investigation 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 investigated. 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 investigated 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.

  7. The effects of pressure sensor acoustics on airdata derived from a High-angle-of-attack Flush Airdata Sensing (HI-FADS) system

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.

    1991-01-01

    The accuracy of a nonintrusive high angle-of-attack flush airdata sensing (HI-FADS) system was verified for quasi-steady flight conditions up to 55 deg angle of attack during the F-18 High Alpha Research Vehicle (HARV) Program. The system is a matrix of nine pressure ports arranged in annular rings on the aircraft nose. The complete airdata set is estimated using nonlinear regression. Satisfactory frequency response was verified to the system Nyquist frequency (12.5 Hz). The effects of acoustical distortions within the individual pressure sensors of the nonintrusive pressure matrix on overall system performance are addressed. To quantify these effects, a frequency-response model describing the dynamics of acoustical distortion is developed and simple design criteria are derived. The model adjusts measured HI-FADS pressure data for the acoustical distortion and quantifies the effects of internal sensor geometries on system performance. Analysis results indicate that sensor frequency response characteristics very greatly with altitude, thus it is difficult to select satisfactory sensor geometry for all altitudes. The solution used presample filtering to eliminate resonance effects, and short pneumatic tubing sections to reduce lag effects. Without presample signal conditioning the system designer must use the pneumatic transmission line to attenuate the resonances and accept the resulting altitude variability.

  8. Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

    1993-01-01

    A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.

  9. Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

    1993-01-01

    A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angle-of-attack and sideslip regions studied.

  10. Flight and wind-tunnel calibrations of a flush airdata sensor at high angles of attack and sideslip and at supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.

    1993-01-01

    A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angle-of-attack and sideslip regions studied.

  11. The effects of angle-of-attack indication on aircraft control in the event of an airspeed indicator malfunction

    NASA Astrophysics Data System (ADS)

    Boesser, Claas Tido

    Analysis of accident data by the Federal Aviation Administration, the National Transportation Safety Board, and other sources show that loss of control is the leading cause of aircraft accidents. Further evaluation of the data indicates that the majority of loss of control accidents are caused by the aircraft stalling. In response to these data, the Federal Aviation Administration and the General Aviation Joint Steering Committee emphasize the importance of stall and angle-of-attack awareness during flight. The high-profile crash of Air France Flight 447, in which pilots failed to recover from a self-induced stall, reinforced concerns over the need for improved stall and angle-of-attack awareness and reinvigorated interest in the debate over the effectiveness of angle-of-attack information displays. Further support for aerodynamic information in the form of an angle-of-attack indicator comes from core cognitive engineering principles. These principles argue for the provision of information about system functioning and dynamics as a means to ensure a human is always in position to recover a system when technology is unable. The purpose of this research was to empirically evaluate the importance of providing pilots with feedback about fundamental aircraft aerodynamics, especially during non-standard situations and unexpected disturbances. An experiment was conducted using a flight simulator to test the effects of in-cockpit angle-of-attack indication on aircraft control following an airspeed indicator malfunction on final approach. Participants flew a final approach with a target airspeed range of 60 to 65 knots. Once participants slowed the aircraft for final approach, the airspeed indicator needle would be stuck at an indication of 70 knots. One group of participants flew the final approach with an angle-of-attack indicator while the other group lacked such an instrument. Examination of aircraft performance data along the final approach showed that, when confronted

  12. Influence of the Angle of Attack on the Aerothermodynamics of the Mars Science Laboratory

    NASA Technical Reports Server (NTRS)

    Dyakonov, Artem A.; Edquist, Karl T.; Schoenenberger, Mark

    2006-01-01

    An investigation of the effects of the incidence angle on the aerothermodynamic environments of the Mars Science Laboratory has been conducted. Flight conditions of peak heating, peak deceleration and chute deploy are selected and the effects of the angle of attack on the aerodynamics and aerothermodynamics are analyzed. The investigation found that static aerodynamics are well behaved within the considered range of incidence angles. Leeside laminar and turbulent computed heating rates decrease with incidence, despite the increase in the leeside running length. Stagnation point was found to stay on the conical flank at all angles of attack, and this is linked to the rapid flow expansion around the shoulder. Hypersonic lift to drag ratio is limited by the heating rates in the region of the windside shoulder. The effects of the high angle of incidence on the dynamic aero at low Mach remains to be determined. Influence of the angle of attack on the smooth-wall transition parameter indicates, that higher angle of attack flight may result in delayed turbulence onset, however, a coupled analysis, involving flight trajectory simulation is necessary.

  13. An investigation to determine the pressure distribution on the 0.0137 scale solid rocket booster forebody (MSFC model 467) at angles of attack at or near 90 deg and high Reynolds numbers in the MSFC High Reynolds Number Wind Tunnel (SA29F)

    NASA Technical Reports Server (NTRS)

    Ramsey, P. E.

    1976-01-01

    An aerodynamic investigation was conducted in the MSFC High Reynolds Number Wind Tunnel to determine the pressure distribution over the foresection of the current 146 inch diameter shuttle SRB. The test model consisted of a 0.0137 scale version of the SRB nose cone and a forward portion of the cylindrical body which was approximately 2.7 calibers in length. The pressure distributions are plotted as a function of longitudinal station ratioed to body diameter and circumferential location for each angle of attack and Mach number. A Reynolds number variation study was made for Mach numbers of 0.4 and 0.6 at an angle of attack of 270 deg and roll angle of 180 deg.

  14. Aerodynamic Characteristics of a 10 deg Sharp Cone at Hypersonic Speeds and High Angles of Attack

    DTIC Science & Technology

    1975-06-01

    atzthrough to the fSo fection. a r me ’: : SB Base area , in \\; B S Horizontal component of model center BACKGROUND of gravity displacement measured...experimental work includes regions included the cone surface and basenear-wake data for a variety of vehicle con- areas , the cone shoulder boundary layer and...t t kisolated areas amounting to 5 t.: ;.0 of roll These units were based on a Harrison design2 5 angle each. It was further fou•. that the

  15. Investigation of High-Angle-of-Attack Maneuver-Limiting Factors. Part 3. Appendices - Aerodynamic Models

    DTIC Science & Technology

    1980-12-01

    damper (gradient .•3.5 lb/in) * Breakout 1- .5 to 2 lb ’.3 20 - ~~44 ’.I rto 0 0 r() *-d LO to. 4LO 10 -j CLC E N im0 U1- - L 210 Z. DATA SOURCES This...I;F 0.0U0 -M.090 .,U! 1 i J’ totr 03 F~igure 17- (a)- LCm1 (a, hI &3Rm1+c 1 iri - .. ’I .*Is CH -2.10 -ion Fvigure 18. Cmq (~ 35 NLCM.2CEE PER ARM...DEG) - Figur’e .39. CLF~ (m) 1-de 14 iri .......... 1:0 ia in :. o I,’A 1. I . Figure 40. OD•s•c(Op3) *1 Figure 41 . CDBASIC(CY) ,) 6.5 j ,.

  16. An experimental investigation of thrust vectoring two-dimensional convergent-divergent nozzles installed in a twin-engine fighter model at high angles of attack

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Mason, Mary L.; Leavitt, Laurence D.

    1990-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine thrust vectoring capability of subscale 2-D convergent-divergent exhaust nozzles installed on a twin engine general research fighter model. Pitch thrust vectoring was accomplished by downward rotation of nozzle upper and lower flaps. The effects of nozzle sidewall cutback were studied for both unvectored and pitch vectored nozzles. A single cutback sidewall was employed for yaw thrust vectoring. This investigation was conducted at Mach numbers ranging from 0 to 1.20 and at angles of attack from -2 to 35 deg. High pressure air was used to simulate jet exhaust and provide values of nozzle pressure ratio up to 9.

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

    NASA Technical Reports Server (NTRS)

    Larson, T. J.

    1984-01-01

    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.

  18. Large-scale wind tunnel tests of a sting-supported V/STOL fighter model at high angles of attack

    NASA Technical Reports Server (NTRS)

    Stoll, F.; Minter, E. A.

    1981-01-01

    A new sting model support has been developed for the NASA/Ames 40- by 80-Foot Wind Tunnel. This addition to the facility permits testing of relatively large models to large angles of attack or angles of yaw depending on model orientation. An initial test on the sting is described. This test used a 0.4-scale powered V/STOL model designed for testing at angles of attack to 90 deg and greater. A method for correcting wake blockage was developed and applied to the force and moment data. Samples of this data and results of surface-pressure measurements are presented.

  19. Wind-tunnel research comparing lateral control devices, particularly at high angles of attack III : ordinary ailerons rigged up 10 degrees when neutral

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Wenzinger, Carl J

    1933-01-01

    This report presents the results of wind-tunnel tests made on three model wings having different sizes of ordinary ailerons rigged up 10 degrees when neutral, the same models having previously been tested with the ailerons rigged even with the wings in the usual manner. One of the wings had ailerons of medium size, 25 per cent of the wing chord by 40 per cent of the semispan, one had long, narrow ailerons, and one had short, wide ones. These tests are part of a general investigation on lateral control devices, with particular reference to the control at high angles of attack, in which all the devices are being subjected to the same series of tests in the 7 by 10 foot wind tunnel of the National Advisory Committee for Aeronautics. Force tests of the usual type, free-autorotation tests, and forced-rotation tests were made showing the effect of the ailerons on the general performance of the wing, on the lateral controllability, and on the lateral stability.

  20. Low-speed stability and control wind-tunnel investigations of effects of spanwise blowing on fighter flight characteristics at high angles of attack. [Langely 12-ft low-speed tunnel and 30- by 60-ft tunnel

    NASA Technical Reports Server (NTRS)

    Satran, D. R.; Gilbert, W. P.; Anglin, E. L.

    1985-01-01

    The effects of spanwise blowing on two configurations representative of current fighter airplanes were investigated. The two configurations differed only in wing planform, with one incorporating a trapezoidal wing and the other a 60 delta wing. Emphasis was on determining the lateral-directional characteristics, particularly in the stall/departure angle-of-attack range; however, the effects of spanwise blowing on the longitudinal aerodynamics were also determined. The-tunnel tests included measurement of static force and forced-oscillation aerodynamic data, visualization of the airflow changes created by the spanwise blowing, and free-flight model tests. The effects of blowing rate, chordwise location of the blowing ports, asymmetric blowing, and blowing on the conventional aerodynamic control characteristics were investigated. In the angle-of-attack regions in which the spanwise blowing substantially improved the wing upper-surface flow field (i.e., provided reattachment of the flow aft of the leading-edge vortex), improvements in both static and dynamic lateral-directional stability were observed. Blowing effects on stability could be proverse or adverse depending on blowing rate, blowing port loaction, and wing planform. Free-flight model tests of the trapezoidal wing confirmed the beneficial effects of spanwise blowing measured in the static and dynamic force tests.

  1. Analysis of the longitudinal handling qualities and pilot-induced-oscillation tendencies of the High-Angle-of-Attack Research Vehicle (HARV)

    NASA Technical Reports Server (NTRS)

    Hess, Ronald A.

    1994-01-01

    The NASA High-Angle-of Attack Research Vehicle (HARV), a modified F-18 aircraft, experienced handling qualities problems in recent flight tests at NASA Dryden Research Center. Foremost in these problems was the tendency of the pilot-aircraft system to exhibit a potentially dangerous phenomenon known as a pilot-induced oscillation (PIO). When they occur, PIO's can severely restrict performance, sharply dimish mission capabilities, and can even result in aircraft loss. A pilot/vehicle analysis was undertaken with the goal of reducing these PIO tendencies and improving the overall vehicle handling qualities with as few changes as possible to the existing feedback/feedforward flight control laws. Utilizing a pair of analytical pilot models developed by the author, a pilot/vehicle analysis of the existing longitudinal flight control system was undertaken. The analysis included prediction of overall handling qualities levels and PIO susceptability. The analysis indicated that improvement in the flight control system was warranted and led to the formulation of a simple control stick command shaping filter. Analysis of the pilot/vehicle system with the shaping filter indicated significant improvements in handling qualities and PIO tendencies could be achieved. A non-real time simulation of the modified control system was undertaken with a realistic, nonlinear model of the current HARV. Special emphasis was placed upon those details of the command filter implementation which could effect safety of flight. The modified system is currently awaiting evaluation in the real-time, pilot-in-the-loop, Dual-Maneuvering-Simulator (DMS) facility at Langley.

  2. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 1: Low-wing model A. [fluid flow and vortices data for general aviation aircraft to determine aerodynamic characteristics for various designs

    NASA Technical Reports Server (NTRS)

    Hultberg, R. S.; Mulcay, W.

    1980-01-01

    Aerodynamic characteristics obtained in a rotational flow environment utilizing a rotary balance are presented in plotted form for a 1/5 scale, single engine, low-wing, general aviation airplane model. The configuration tested included the basic airplane, various control deflections, tail designs, fuselage shapes, and wing leading edges. Data are presented without analysis for an angle of attack range of 8 to 90 deg and clockwise and counterclockwise rotations covering a range from 0 to 0.85.

  3. A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack

    NASA Technical Reports Server (NTRS)

    Petrilli, Justin; Paul, Ryan; Gopalarathnam, Ashok; Frink, Neal T.

    2013-01-01

    This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.

  4. Wind-tunnel investigation of effects of wing-leading-edge modifications on the high angle-of-attack characteristics of a T-tail low-wing general-aviation aircraft

    NASA Technical Reports Server (NTRS)

    White, E. R.

    1982-01-01

    Exploratory tests have been conducted in the NASA-Langley Research Center's 12-Foot Low-Speed wind Tunnel to evaluate the application of wing-leading-edge devices on the stall-departure and spin resistance characteristics of a 1/6-scale model of a T-tail general-aviation aircraft. The model was force tested with an internal strain-gauge balance to obtain aerodynamic data on the complete configuration and with a separate wing balance to obtain aerodynamic data on the outer portion of the wing. The addition of the outboard leading-edge droop eliminated the abrupt stall of the windtip and maintained or increased the resultant-force coefficient up to about alpha = 32 degrees. This change in slope of the resultant-force coefficient curve with angle of attack has been shown to be important for eliminating autorotation and for providing spin resistance.

  5. Vibration Amplitude of a Flexible Filament Changes Non-Monotonically with Angle of Attack

    NASA Astrophysics Data System (ADS)

    Akaydin, H. Dogus; Voesenek, Cees J.; Lentink, David; Lentink Lab Team

    2013-11-01

    Certain animals exploit the interaction of their flexible, foil-like appendages with vortices to propel themselves effectively in surrounding fluids. The interaction is reciprocal because the fluid forces deform the appendage while the appendage alters the flow. As the flow separates from a foil-like structure, it rolls-up into a vortex with a low-pressure core, which deforms the structure until the vortex is shed into the wake. A control over this interaction can provide a robust aerodynamic performance. To identify the salient parameters that control fluid-structure interactions on a deformable structure, we varied the thickness, length and angle of attack of a rubber filament in a quasi-two-dimensional flow generated using a soap-film tunnel. We resolved fluid-structure interaction by simultaneously measuring deformation of the filament and motion of particles in the fluid using high-speed cameras. We show that increasing length or decreasing diameter of the filament increases its vibration amplitude monotonically. In stark contrast, the angle of attack of the filament may alter the amplitude of vibrations in a non-monotonic way: Within a certain range of angle of attack, the filament motion and vortex shedding lock-in, i.e. become synchronized, and result in a violent flapping behavior. This response can therefore be ceased not only by decreasing but also by increasing the angle of attack at the leading edge. Such an insight can help us engineer more effective bio-inspired robots, energy harvesters, and flow control devices with vibration control.

  6. Test data report, low speed wind tunnel tests of a full scale lift/cruise-fan inlet, with engine, at high angles of attack

    NASA Technical Reports Server (NTRS)

    Shain, W. M.

    1978-01-01

    A low speed wind tunnel test of a fixed lip inlet with engine, was performed. The inlet was close coupled to a Hamilton Standard 1.4 meter, variable pitch fan driven by a lycoming T55-L-11A engine. Tests were conducted with various combinations of inlet angle of attack freestream velocities, and fan airflows. Data were recorded to define the inlet airflow separation boundaries, performance characteristics, and fan blade stresses. The test model, installation, instrumentation, test, data reduction and final data are described.

  7. Angle-of-Attack-Modulated Terminal Point Control for Neptune Aerocapture

    NASA Technical Reports Server (NTRS)

    Queen, Eric M.

    2004-01-01

    An aerocapture guidance algorithm based on a calculus of variations approach is developed, using angle of attack as the primary control variable. Bank angle is used as a secondary control to alleviate angle of attack extremes and to control inclination. The guidance equations are derived in detail. The controller has very small onboard computational requirements and is robust to atmospheric and aerodynamic dispersions. The algorithm is applied to aerocapture at Neptune. Three versions of the controller are considered with varying angle of attack authority. The three versions of the controller are evaluated using Monte Carlo simulations with expected dispersions.

  8. Angle of Attack Modulation for Mars Entry Terminal State Optimization

    NASA Technical Reports Server (NTRS)

    Lafleur, Jarret M.; Cerimele, Christopher J.

    2009-01-01

    From the perspective of atmospheric entry, descent, and landing (EDL), one of the most foreboding destinations in the solar system is Mars due in part to its exceedingly thin atmosphere. To benchmark best possible scenarios for evaluation of potential Mars EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities and vehicle masses, focusing on the identification of potential benefits of enabling angle of attack modulation. The terminal state is envisioned as one appropriate for the initiation of terminal descent via parachute or other means. A particle swarm optimizer varies entry flight path angle, ten bank profile points, and ten angle of attack profile points to find maximum-final-altitude trajectories for a 10 30 m ellipsled at 180 different combinations of values for entry mass, entry velocity, terminal Mach number, and minimum allowable altitude. Parametric plots of maximum achievable altitude are shown, as are examples of optimized trajectories. It is shown that appreciable terminal state altitude gains (2.5-4.0 km) over pure bank angle control may be possible if angle of attack modulation is enabled for Mars entry vehicles. Gains of this magnitude could prove to be enabling for missions requiring high-altitude landing sites. Conclusions are also drawn regarding trends in the bank and angle of attack profiles that produce the optimal trajectories in this study, and directions for future work are identified.

  9. Computational fluid dynamics study of wind turbine blade profiles at low Reynolds numbers for various angles of attack

    NASA Astrophysics Data System (ADS)

    Sayed, Mohamed A.; Kandil, Hamdy A.; Morgan, El-Sayed I.

    2012-06-01

    Airfoil data are rarely available for Angles Of Attack (AOA) over the entire range of ±180°. This is unfortunate for the wind turbine designers, because wind turbine airfoils do operate over this entire range. In this paper, an attempt is made to study the lift and drag forces on a wind turbine blade at various sections and the effect of angle of attack on these forces. Aerodynamic simulations of the steady flow past two-dimensional wind-turbine blade-profiles, developed by the National Renewable Energy Laboratory (NREL) at low Reynolds number, will be performed. The aerodynamic simulation will be performed using Computational Fluid Dynamics (CFD) techniques. The governing equations used in the simulations are the Reynolds-Average-Navier-Stokes (RANS) equations. The simulations at different wind speeds will be performed on the S809 and the S826 blade profiles. The S826 blade profile is considered in this study because it is the most suitable blade profile for the wind conditions in Egypt in the site of Gulf El-Zayt on the red sea. Lift and drag forces along with the angle of attack are the important parameters in a wind turbine system. These parameters determine the efficiency of the wind turbine. The lift and drag forces are computed over the entire range of AOA of ±180° at low Reynolds numbers. The results of the analysis showed that the AOA between 3° and 8° have high Lift/Drag ratio regardless of the wind speed and the blade profile. The numerical results are compared with wind tunnel measurements at the available limited range of the angle of attack. In addition, the numerical results are compared with the results obtained from the equations developed by Viterna and Janetzke for deep stall. The comparisons showed that the used CFD code can accurately predict the aerodynamic loads on the wind-turbine blades.

  10. Aerodynamic Force Characteristics of a Series of Lifting Cone and Cone-Cylinder Configurations at a Mach Number of 6.83 and Angles of Attack up to 130 Deg

    NASA Technical Reports Server (NTRS)

    Penland, Jim A.

    1961-01-01

    Force tests of a series of right circular cones having semivertex angles ranging from 5 deg to 45 deg and a series of right circular cone-cylinder configurations having semivertex angles ranging from 5 deg to 20 deg and an afterbody fineness ratio of 6 have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.83, a Reynolds number of 0.24 x 10.6 per inch, and angles of attack up to 130 deg. An analysis of the results made use of the Newtonian and modified Newtonian theories and the exact theory. A comparison of the experimental data of both cone and cone-cylinder configurations with theoretical calculations shows that the Newtonian concept gives excellent predictions of trends of the force characteristics and the locations with respect to angle of attack of the points of maximum lift, maximum drag, and maximum lift-drag ratio. Both the Newtonian a.nd exact theories give excellent predictions of the sign and value of the initial lift-curve slope. The maximum lift coefficient for conical bodies is nearly constant at a value of 0.5 based on planform area for semivertex angles up to 30 deg. The maximum lift-drag ratio for conical bodies can be expected to be not greater than about 3.5, and this value might be expected only for slender cones having semivertex angles of less than 5 deg. The increments of angle of attack and lift coefficient between the maximum lift-drag ratio and the maximum lift coefficient for conical bodies decrease rapidly with increasing semivertex angles as predicted by the modified Newtonian theory.

  11. Effect of Mean Angle of Attack Modulation on Dynamic Stall

    NASA Astrophysics Data System (ADS)

    Heintz, Kyle; Corke, Thomas

    2016-11-01

    Wind tunnel experiments at M = 0 . 2 were conducted on a cambered airfoil instrumented with surface pressure transducers that was oscillated with two independent frequencies. The primary input, f1, corresponds to a range of reduced frequencies, while the slower, secondary input, f2, drives the modulation of the mean angle of attack, thus varying the stall-penetration angle, αpen. Various combinations transitioned different regimes of dynamic stall from "light" to "deep". Results suggest that when αpen is falling between consecutive cycles, the aerodynamic loads do not fully recover to the values seen when αpen is rising, even though the airfoil recedes to αpen < 0 during each oscillation. The experimental data is presented in terms of load coefficients, aerodynamic damping, and their phase relationships to pitch angle. APS Fellow.

  12. Engineering flowfield method with angle-of-attack applications

    NASA Technical Reports Server (NTRS)

    Zoby, E. V.; Simmonds, A. L.

    1984-01-01

    An approximate inviscid flowfield method has been extended to include heat-transfer predictions using a technique to account for variable-entropy edge conditions. The engineering code computes the flowfield over hyperboloids, ellipsoids, paraboloids, and sphere cones at 0 deg angle of attack (AOA). For angle-of-attack applications, an approximation to sphere-cone streamline-spreading effects on the heat transfer along the windward and leeward rays and an empirical circumferential heating technique have been incorporated also in the method. The present engineering calculations yield good comparisons with existing pressure and heating data over sphere cones even at high incidence values with the restriction that the sonic-line location remain on the spherical cap.

  13. Engineering flowfield method with angle-of-attack applications

    NASA Technical Reports Server (NTRS)

    Zoby, E. V.; Simmonds, A. L.

    1984-01-01

    An approximate inviscid flowfield method has been extended to include heat-transfer predictions using a technique to account for variable-entropy edge conditions. The engineering code computes the flowfield over hyperboloids, ellipsoids, paraboloids, and sphere cones at 0 deg angle of attack (AOA). For angle-of-attack applications, an approximation to sphere-cone streamline-spreading effects on the heat transfer along the windward and leeward rays and an empirical circumferential heating technique have been incorporated also in the method. The present engineering calculations yield good comparisons with existing pressure and heating data over sphere cones even at high incidence values with the restriction that the sonic-line location remain on the spherical cap.

  14. 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

    NASA Technical Reports Server (NTRS)

    Chokani, Ndaona; Gittner, N. M.

    1992-01-01

    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.

  15. Calculation of High Angle of Attack Aerodynamics of Fighter Configurations. Volume 2. User Manual for VORSTAB-II

    DTIC Science & Technology

    1991-04-01

    surfaces. Grouo 3 LAT = -1 if the rolling moment coefficient for a given aileron angle is to be computed. = 0 for symmetrical loading only = 1 if both...LEV=O in Group 79 and without a fuselage.) = 2, if the calculation is for one design lift coefficient based on the attached-flow theory, = 3, if it...0, otherwise CLDS = design lift coefficient if ALPCON = 2 or 3. (Used only if LEV=O in Group 79.) = 0, otherwise -16- If ALPCON = 1., set zLPA = 0

  16. Development of a nonlinear vortex method. [steady and unsteady aerodynamic loads of highly sweptback wings

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.

    1981-01-01

    Progress is reported in the development of reliable nonlinear vortex methods for predicting the steady and unsteady aerodynamic loads of highly sweptback wings at large angles of attack. Abstracts of the papers, talks, and theses produced through this research are included. The modified nonlinear discrete vortex method and the nonlinear hybrid vortex method are highlighted.

  17. Computation of transonic flow past projectiles at angle of attack

    NASA Technical Reports Server (NTRS)

    Reklis, R. P.; Sturek, W. B.; Bailey, F. R.

    1978-01-01

    Aerodynamic properties of artillery shell such as normal force and pitching moment reach peak values in a narrow transonic Mach number range. In order to compute these quantities, numerical techniques have been developed to obtain solutions to the three-dimensional transonic small disturbance equation about slender bodies at angle of attack. The computation is based on a plane relaxation technique involving Fourier transforms to partially decouple the three-dimensional difference equations. Particular care is taken to assure accurate solutions near corners found in shell designs. Computed surface pressures are compared to experimental measurements for circular arc and cone cylinder bodies which have been selected as test cases. Computed pitching moments are compared to range measurements for a typical projectile shape.

  18. Interference effects of aircraft components on the local blade angle of attack of a wing-mounted propeller

    NASA Technical Reports Server (NTRS)

    Mendoza, J. P.

    1979-01-01

    The aerodynamic interference effects on a propeller operating in the presence of different wing-body-nacelle combinations was studied. The unsteady blade angle of attack variation with azimuth angle by varying the pitch and yaw of the nacelle was minimized. Results indicate for the particular configuration of interest the minimum blade angle of attack variation occurred with the nacelle pitched downward 4.5 deg and yawed inward 3.0 deg.

  19. Experimental static aerodynamic forces and moments at high subsonic speeds on a missile model during simulated launching from the midsemispan location of a 45 degree sweptback wing-fuselage-pylon combination

    NASA Technical Reports Server (NTRS)

    Alford, William J; King, Thomas, Jr

    1957-01-01

    An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the static aerodynamic forces and moments on a missile model during simulated launching from the midsemispan location of a 45 degree sweptback wing-fuselage-pylon combination. The results indicated significant variations in all the aerodynamic components with changes in chordwise location of the missile. Increasing the angle of attack caused increases in the induced effects on the missile model because of the wing-fuselage-pylon combination. Increasing the Mach number had little effect on the variations of the missile aerodynamic characteristics with angle of attack except that nonlinearities were incurred at smaller angles of attack for the higher Mach numbers. The effects of finite wing thickness on the missile characteristics, at zero angle of attack, increase with increasing Mach number. The effects of the pylon on the missile characteristics were to causeincreases in the rolling-moment variation with angle of attack and a negative displacement of the pitching-moment curves at zero angle of attack. The effects of skewing the missile in the lateral direction relative to and sideslipping the missile with the wing-fuselage-pylon combination were to cause additional increments in side force at zero angle of attack. For the missile yawing moments the effects of changes in skew or sideslip angles were qualitatively as would be expected from consideration of the isolated missile characteristics, although there existed differences in theyawing-moment magnitudes.

  20. The radiation of sound from a propeller at angle of attack

    NASA Technical Reports Server (NTRS)

    Mani, Ramani

    1990-01-01

    The mechanism by which the noise generated at the blade passing frequency by a propeller is altered when the propeller axis is at an angle of attack to the freestream is examined. The measured noise field is distinctly non axially symmetric under such conditions with far field sound pressure levels both diminished and increased relative to the axially symmetric values produced with the propeller at zero angle of attack. Attempts have been made to explain this non axially symmetric sound field based on the unsteady (once per rev) loading experienced by the propeller blades when the propeller axis is at non zero angle of attack. A calculation based on this notion appears to greatly underestimate the measured azimuthal asymmetry of noise for high tip speed, highly loaded propellers. A new mechanism is proposed; namely, that at angle of attack, there is a non axially symmetric modulation of the radiative efficiency of the steady loading and thickness noise which is the primary cause of the non axially symmetric sound field at angle of attack for high tip speed, heavily loaded propellers with a large number of blades. A calculation of this effect to first order in the crossflow Mach number (component of freestream Mach number normal to the propeller axis) is carried out and shows much better agreement with measured noise data on the angle of attack effect.

  1. The PELskin project-part V: towards the control of the flow around aerofoils at high angle of attack using a self-activated deployable flap.

    PubMed

    Rosti, Marco E; Kamps, Laura; Bruecker, Christoph; Omidyeganeh, Mohammad; Pinelli, Alfredo

    2017-01-01

    During the flight of birds, it is often possible to notice that some of the primaries and covert feathers on the upper side of the wing pop-up under critical flight conditions, such as the landing approach or when stalking their prey (see Fig. 1) . It is often conjectured that the feathers pop up plays an aerodynamic role by limiting the spread of flow separation . A combined experimental and numerical study was conducted to shed some light on the physical mechanism determining the feathers self actuation and their effective role in controlling the flow field in nominally stalled conditions. In particular, we have considered a NACA0020 aerofoil, equipped with a flexible flap at low chord Reynolds numbers. A parametric study has been conducted on the effects of the length, natural frequency, and position of the flap. A configuration with a single flap hinged on the suction side at 70 % of the chord size c (from the leading edge), with a length of [Formula: see text] matching the shedding frequency of vortices at stall condition has been found to be optimum in delivering maximum aerodynamic efficiency and lift gains. Flow evolution both during a ramp-up motion (incidence angle from [Formula: see text] to [Formula: see text] with a reduced frequency of [Formula: see text], [Formula: see text] being the free stream velocity magnitude), and at a static stalled condition ([Formula: see text]) were analysed with and without the flap. A significant increase of the mean lift after a ramp-up manoeuvre is observed in presence of the flap. Stall dynamics (i.e., lift overshoot and oscillations) are altered and the simulations reveal a periodic re-generation cycle composed of a leading edge vortex that lift the flap during his passage, and an ejection generated by the relaxing of the flap in its equilibrium position. The flap movement in turns avoid the interaction between leading and trailing edge vortices when lift up and push the trailing edge vortex downstream when

  2. Experimental study of compressible boundary layer on a cone at angles of attack

    NASA Astrophysics Data System (ADS)

    Maslov, A.; Bountin, D.; Shiplyuk, A.; Sidorenko, A.; Shen, Q.; Bi, Z.

    2009-06-01

    Experimental study was conducted for boundary-layers on a sharp 5° half-angle cone of 400 mm length at angles of attack. The model was tested in the T-326 hypersonic wind tunnel (ITAM) at freestream Mach number M = 5.95. Mean and fluctuation wall characteristics of the boundary layer are measured at 0°, 2°, 3° and 4° angles of attack for different stagnation pressures. Pulsation measurements are carried out by means of ALTP sensor. Pressure and temperature distributions along the model are obtained, and transition beginning and end locations have been found. Boundary layer stabilization with the increase of angle of attack and the decrease of stagnation pressure is observed. High frequency pulsations inherent to hypersonic boundary layer (second mode) have been detected.

  3. Hypersonic slender-wedge analysis with gradual change in angle of attack

    NASA Technical Reports Server (NTRS)

    Gupta, R. N.; Joshi, S. P.; Rodkiewicz, C. M.

    1983-01-01

    The behavior of a narrow cross-section wedge wing moving at a high Mach number and subjected to an angle of attack changing exponentially with time is investigated. This type of wedge wing is commonly employed as a lifting surface in hypersonic vehicles. The time history of wall shear, heat transfer, displacement thickness, and viscous induced pressure are determined. Results show that for the same change in angle of attack, the flow attains the final steady state much faster when the change is exponential than when the change is made impulsively. In addition, the unsteady character of the flow is primarily confined to the initial stages of the change in the angle of attack.

  4. Supersonic flow around circular cones at angles of attack

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1951-01-01

    The properties of conical flow without axial symmetry are analyzed. The flow around cones of circular cross section at small angles of attack is determined by correctly considering the effect of the entropy gradients in the flow.

  5. Fourth High Alpha Conference, volume 1

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The goal of the Fourth High Alpha Conference was to focus on the flight validation of high angle-of-attack technologies and provide an in-depth review of the latest high angle-of-attack activities. Areas that were covered include: high angle-of-attack aerodynamics, propulsion and inlet dynamics, thrust vectoring, control laws and handling qualities, tactical utility, and forebody controls.

  6. Fourth High Alpha Conference, volume 2

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The goal of the Fourth High Alpha Conference, held at the NASA Dryden Flight Research Center on July 12-14, 1994, was to focus on the flight validation of high angle of attack technologies and provide an in-depth review of the latest high angle of attack activities. Areas that were covered include high angle of attack aerodynamics, propulsion and inlet dynamics, thrust vectoring, control laws and handling qualities, and tactical utility.

  7. Angle-of-attack validation of a new zonal CFD method for airfoil simulations

    NASA Technical Reports Server (NTRS)

    Yoo, Sungyul; Summa, J. Michael; Strash, Daniel J.

    1990-01-01

    The angle-of-attack validation of a new concept suggested by Summa (1990) for coupling potential and viscous flow methods has been investigated for two-dimensional airfoil simulations. The fully coupled potential/Navier-Stokes code, ZAP2D (Zonal Aerodynamics Program 2D), has been used to compute the flow field around an NACA 0012 airfoil for a range of angles of attack up to stall at a Mach number of 0.3 and a Reynolds number of 3 million. ZAP2D calculation for various domain sizes from 25 to 0.12 chord lengths are compared with the ARC2D large domain solution as well as with experimental data.

  8. Angle-of-attack validation of a new zonal CFD method for airfoil simulations

    NASA Technical Reports Server (NTRS)

    Yoo, Sungyul; Summa, J. Michael; Strash, Daniel J.

    1990-01-01

    The angle-of-attack validation of a new concept suggested by Summa (1990) for coupling potential and viscous flow methods has been investigated for two-dimensional airfoil simulations. The fully coupled potential/Navier-Stokes code, ZAP2D (Zonal Aerodynamics Program 2D), has been used to compute the flow field around an NACA 0012 airfoil for a range of angles of attack up to stall at a Mach number of 0.3 and a Reynolds number of 3 million. ZAP2D calculation for various domain sizes from 25 to 0.12 chord lengths are compared with the ARC2D large domain solution as well as with experimental data.

  9. Fourth High Alpha Conference, volume 3

    NASA Technical Reports Server (NTRS)

    1994-01-01

    Thie goal of this conference was to focus on the flight validation of high-angle-of-attack technologies and provide an in-depth review of the latest high-angle-of-attack activities. Areas covered include: (1) high-angle-of-attack aerodynamics; (2) propulsion and inlet dynamics; (3) thrust vectoring; (4) control laws and handling qualities; (5) tactical utility; and (6) forebody controls.

  10. Tables for Supersonic Flow Around Right Circular Cones at Small Angle of Attack

    NASA Technical Reports Server (NTRS)

    Sims, Joseph L.

    1964-01-01

    The solution of supersonic flow fields by the method of characteristics requires that starting conditions be known. Ferri, in reference 1, developed a method-of-characteristics solution for axially symmetric bodies of revolution at small angles of attack. With computing machinery that is now available, this has become a feasible method for computing the aerodynamic characteristics of bodies near zero angle of attack. For sharp-nosed bodies of revolution, the required starting line may be obtained by computing the flow field about a cone at a small angle of attack. This calculation is readily performed using Stone's theory in reference 2. Some solutions of this theory are available in reference 3. However, the manner in which these results are presented, namely in a wind-fixed coordinate system, makes their use somewhat cumbersome. Additionally, as pointed out in reference 4, the flow component perpendicular to the meridian planes was computed incorrectly. The results contained herein have been computed in the same basic manner as those of reference 3 with the correct velocity normal to the meridian planes. Also, all results have been transferred into the body-fixed coordinate system. Therefore, the values tabulated herein may be used, in conjunction with the respective zero-angle-of-attack results of reference 5, as starting conditions for the method-of-characteristics solution of the flow field about axially symmetric bodies of revolution at small angles of attack. As in the zero-angle-of-attack case (ref. 5) the present results have been computed using the ideal gas value of 1.4 for the ratio of the specific heats of air. Solutions are given for cone angles from 2.5 deg to 30 deg in increments of 2.5 deg. For each cone angle, results were computed for a constant series of free-stream Mach numbers from 1.5 to 20. In addition, a solution was computed which yielded the minimum free-stream Mach number for a completely supersonic conical flow field. For cone angles of

  11. DSMC simulation for effects of angles of attack on rarefied hypersonic cavity flows

    NASA Astrophysics Data System (ADS)

    Jin, Xuhon; Huang, Fei; Shi, Jiatong; Cheng, Xiaoli

    2016-11-01

    The present work investigates rarefied hypersonic flows over a flat plate with two-dimensional and three-dimensional cavities by employing the direct simulation Monte Carlo (DSMC) method, focusing on the effect of angles of attack (AOAs) on flow structure inside the cavity and aerodynamic surface quantities. It was found that only one primary recirculation structure was formed inside the cavity at the angle of attack (AOA) of 0°, while a second vortex system was produced just beneath the primary one with the angle of attack increased to 30°. As AOAs grow, the freestream flow is able to penetrate deeper into the cavity and attach itself to the cavity base, making the "dead-water" region shrink. Meantime, with the increment in the AOA, both heat transfer and pressure coefficients show a similar and gradual quantitative behavior, and along centerlines of the two side surfaces of the cavity, both pressure and heat transfer coefficients become growing, indicating that the increase in the AOA does enhance momentum and energy transfer to both the two aforementioned surfaces. However, the heat flux over the cavity floor does not keep increasing with the growth of the AOA, while the pressure does, indicating that augmenting AOAs does not enhance momentum to the cavity floor, but does make compressibility stronger and stronger near the cavity base.

  12. Flight Test Techniques for Quantifying Pitch Rate and Angle of Attack Rate Dependencies

    NASA Technical Reports Server (NTRS)

    Grauer, Jared A.; Morelli, Eugene A.; Murri, Daniel G.

    2017-01-01

    Three different types of maneuvers were designed to separately quantify pitch rate and angle of attack rate contributions to the nondimensional aerodynamic pitching moment coefficient. These maneuvers combined pilot inputs and automatic multisine excitations, and were own with the subscale T-2 and Bat-4 airplanes using the NASA AirSTAR flight test facility. Stability and control derivatives, in particular C(sub mq) and C(sub m alpha(.)) were accurately estimated from the flight test data. These maneuvers can be performed with many types of aircraft, and the results can be used to increase simulation prediction fidelity and facilitate more accurate comparisons with wind tunnel experiments or numerical investigations.

  13. Subsonic longitudinal and lateral-directional characteristics of a forward-swept-wing fighter configuration at angles of attack up to 47 deg

    NASA Technical Reports Server (NTRS)

    Mann, Michael J.; Huffman, Jarrett K.; Fox, Charles H., Jr.

    1987-01-01

    Subsonic lateral-direction and longitudinal characteristics of a forward-swept-wing fighter configuration were examined in wind-tunnel tests at Mach numbers of 0.2 and 0.5 for angles of attack from -7 to 47 deg. and over a sidelslip range of +/- 15 deg. The effects of a canard, strakes, vertical tail, and leading- and trailing-edge flaps are examined. The canard and strakes both reduce asymmetric moments and side forces at zero sideslip for angles of attack up to about 30 deg. The canard has a small influence on lateral-directional stability; however, strakes produce a substantial reduction in lateral stability for angles of attack greater than about 20 deg. The vertical tail improves directional stability for angles of attack up to 30 deg. Deflection of the leading-edge flap to 20 deg. at high angles of attack on the strake and canard configurations degrades lateral and directional stability. Deflection of the trailing-edge flap to 20 deg. on the canard configuration generally increases lateral and directional stability at high angles of attack. Leading- and trailing-edge flaps on the wing-body and canard configurations are effective for increased lift only for angles of attack up to about 40 deg. The leading-edge flap remains effective on the strake configuration over the entire angle-of-attack range tested.

  14. Characteristics of a wind-actuated aerodynamic braking device for high-speed trains

    NASA Astrophysics Data System (ADS)

    Takami, H.; Maekawa, H.

    2017-04-01

    To shorten the stopping distance of the high speed trains in case of emergency, we developed a small-sized aerodynamic braking unit without use of the friction between a rail and a wheel. The developed device could actuate a pair of two drag panels with a travelling wind. However, after the drag panel fully opened, vibrational movements of the drag panel characterized by its slight flutter were repeated. In this study, to stabilize the opened panel, matters pertaining to the angle of attack with respect to the drag panel and pertaining to the arrangement of the two panels were examined by a wind tunnel experiment using a scale model. As a result, to stabilize the opened panel and to keep the good performance of the braking device, it is found out that an angle of attack of 75 to 80 degrees is suitable provided that the interval of the two panels is narrow enough.

  15. The effect of asymmetric attack on trim angle of attack

    NASA Technical Reports Server (NTRS)

    Kruse, R. L.

    1983-01-01

    Ballistic range tests were conducted to determine the effect of an asymmetrically ablated heat shield on the trim angle of attack of an entry vehicle. The tests, which were in support of Project Galileo, were conducted in atmospheric air at Mach numbers from 0.7 to 2.0. For the results for the configuration that was tested, the deduced trim angle varied between 13 deg and 21 deg.

  16. Investigation of the Influence of Rotary Aerodynamics on the Study of High Angle of Attack Dynamics of the F-15B Using Bifurcation Analysis

    DTIC Science & Technology

    1991-12-01

    very nonlinear and require complex analysis tools to study the behavior. Mehra arid Carroll (26) performed fairly extensive analysis of the F-4 Phantom ...INCREMENTS + CMDSPD + DCM C C CMM1 = ATAB03 = BASIC PITCHING MOMENT COEFFICIENT - CM C CMMQ = ATAB05 = PITCH DAMPING DERIVATIVE - CMQ C QB = (QEOBB*MAC)/(2...DELTA CM DUE TO SPEEDBRAKE C SET TO 0 DUE THE REASONS GIVEN ABOVE IN CXDSPD C DCM = BTABO2 = DELTA CM DUE TO 2-PLACE CANOPY (F15B) (=0.0) C C C C C C

  17. Technical evaluation report on the fluid dynamics panel Symposium on High Angle of attack aerodynamics. [slender wings, bodies of revolution, and body-wing configurations

    NASA Technical Reports Server (NTRS)

    Polhamus, E. C.

    1979-01-01

    An overview is presented of 32 formal papers and 7 open session papers. Topics covered include: (1) studies of configurations of practical interest; (2) mathematical modelling and supporting investigations of slender wings, bodies of revolution, and body-wing configurations; (3) design methods; and (4) air intakes.

  18. Technical evaluation report on the fluid dynamics panel Symposium on High Angle of attack aerodynamics. [slender wings, bodies of revolution, and body-wing configurations

    NASA Technical Reports Server (NTRS)

    Polhamus, E. C.

    1979-01-01

    An overview is presented of 32 formal papers and 7 open session papers. Topics covered include: (1) studies of configurations of practical interest; (2) mathematical modelling and supporting investigations of slender wings, bodies of revolution, and body-wing configurations; (3) design methods; and (4) air intakes.

  19. A Preliminary Method for Calculating the Aerodynamic Characteristics of Cruciform Missiles to High Angles of Attack Including Effects of Roll Angle and Control Deflections

    DTIC Science & Technology

    1977-11-01

    lorward -- M CRM2 fradCRM4 CN3" CBM3(, CRM3 Hinge line CBMi , CR Mi CNi i Mi I NOTE: Normal forces are measured perpendicular to the fin planform and are...coefficient; axial force/q SR CBMi bending-moment coefficient of the i’th fin about fin root CD drag coefficient; lift/q SR CL lift coefficient; lift

  20. High speed aerodynamics of upper surface blowing aircraft configurations

    NASA Technical Reports Server (NTRS)

    Birckelbaw, Larry D.

    1992-01-01

    An experimental investigation of the high speed aerodynamics of Upper Surface Blowing (USB) aircraft configurations has been conducted to accurately define the magnitude and causes of the powered configuration cruise drag. A highly instrumented wind tunnel model of a realistic USB configuration was used which permitted parametric variations in the number and spanwise location of the nacelles and was powered with two turbofan engine simulators. The tests conducted in the Ames 14 Foot Transonic Wind Tunnel examined 10 different configurations at Mach numbers from 0.5 to 0.775, fan nozzle pressure ratios from 1.1 to 2.1 and angles of attack from -4 to 6 degrees. Measured force data is presented which indicates the cruise drag penalty associated with each configuration and surface pressure contour plots are used to illustrate the underlying flowfield physics. It was found that all of the tested configurations suffered from a severe drag penalty which increased with freestream Mach number, power setting and angle of attack and was associated with the presence of strong shocks and regions of separated flow in the wing/nacelle junction regions.

  1. Prediction of unsteady blade surface pressures on an advanced propeller at an angle of attack

    NASA Technical Reports Server (NTRS)

    Nallasamy, M.; Groeneweg, J. F.

    1989-01-01

    The paper considers the numerical solution of the unsteady, three-dimensional, Euler equations to obtain the blade surface pressures of an advanced propeller at an angle of attack. The specific configuration considered is the SR7L propeller at cruise conditions with a 4.6 deg inflow angle corresponding to the +2 deg nacelle tilt of the Propeller Test Assessment (PTA) flight test condition. The results indicate nearly sinusoidal response of the blade loading, with angle of attack. For the first time, detailed variations of the chordwise loading as a function of azimuthal angle are presented. It is observed that the blade is lightly loaded for part of the revolution and shocks appear from hub to about 80 percent radial station for the highly loaded portion of the revolution.

  2. Prediction of Unsteady Blade Surface Pressures on an Advanced Propeller at an Angle of Attack

    NASA Technical Reports Server (NTRS)

    Nallasamy, M.; Groeneweg, J. F.

    1989-01-01

    The numerical solution of the unsteady, three-dimensional, Euler equations is considered in order to obtain the blade surface pressures of an advanced propeller at an angle of attack. The specific configuration considered is the SR7L propeller at cruise conditions with a 4.6 deg inflow angle corresponding to the plus 2 deg nacelle tilt of the Propeller Test Assessment (PTA) flight test condition. The results indicate nearly sinusoidal response of the blade loading, with angle of attack. For the first time, detailed variations of the chordwise loading as a function of azimuthal angle are presented. It is observed that the blade is lightly loaded for part of the revolution and shocks appear from hub to about 80 percent radial station for the highly loaded portion of the revolution.

  3. Fourier functional analysis for unsteady aerodynamic modeling

    NASA Technical Reports Server (NTRS)

    Lan, C. Edward; Chin, Suei

    1991-01-01

    A method based on Fourier analysis is developed to analyze the force and moment data obtained in large amplitude forced oscillation tests at high angles of attack. The aerodynamic models for normal force, lift, drag, and pitching moment coefficients are built up from a set of aerodynamic responses to harmonic motions at different frequencies. Based on the aerodynamic models of harmonic data, the indicial responses are formed. The final expressions for the models involve time integrals of the indicial type advocated by Tobak and Schiff. Results from linear two- and three-dimensional unsteady aerodynamic theories as well as test data for a 70-degree delta wing are used to verify the models. It is shown that the present modeling method is accurate in producing the aerodynamic responses to harmonic motions and the ramp type motions. The model also produces correct trend for a 70-degree delta wing in harmonic motion with different mean angles-of-attack. However, the current model cannot be used to extrapolate data to higher angles-of-attack than that of the harmonic motions which form the aerodynamic model. For linear ramp motions, a special method is used to calculate the corresponding frequency and phase angle at a given time. The calculated results from modeling show a higher lift peak for linear ramp motion than for harmonic ramp motion. The current model also shows reasonably good results for the lift responses at different angles of attack.

  4. Progress in high-lift aerodynamic calculations

    NASA Technical Reports Server (NTRS)

    Rogers, Stuart E.

    1993-01-01

    The current work presents progress in the effort to numerically simulate the flow over high-lift aerodynamic components, namely, multi-element airfoils and wings in either a take-off or a landing configuration. The computational approach utilizes an incompressible flow solver and an overlaid chimera grid approach. A detailed grid resolution study is presented for flow over a three-element airfoil. Two turbulence models, a one-equation Baldwin-Barth model and a two equation k-omega model are compared. Excellent agreement with experiment is obtained for the lift coefficient at all angles of attack, including the prediction of maximum lift when using the two-equation model. Results for two other flap riggings are shown. Three-dimensional results are presented for a wing with a square wing-tip as a validation case. Grid generation and topology is discussed for computing the flow over a T-39 Sabreliner wing with flap deployed and the initial calculations for this geometry are presented.

  5. Applications of low lift to drag ratio aerobrakes using angle of attack variation for control

    NASA Technical Reports Server (NTRS)

    Mulqueen, J. A.

    1991-01-01

    Several applications of low lift to drag ratio aerobrakes are investigated which use angle of attack variation for control. The applications are: return from geosynchronous or lunar orbit to low Earth orbit; and planetary aerocapture at Earth and Mars. A number of aerobrake design considerations are reviewed. It was found that the flow impingement behind the aerobrake and the aerodynamic heating loads are the primary factors that control the sizing of an aerobrake. The heating loads and other loads, such as maximum acceleration, are determined by the vehicle ballistic coefficient, the atmosphere entry conditions, and the trajectory design. Several formulations for defining an optimum trajectory are reviewed, and the various performance indices that can be used are evaluated. The 'nearly grazing' optimal trajectory was found to provide the best compromise between the often conflicting goals of minimizing the vehicle propulsive requirements and minimizing vehicle loads. The relationship between vehicle and trajectory design is investigated further using the results of numerical simulations of trajectories for each aerobrake application. The data show the sensitivity of the trajectories to several vehicle parameters and atmospheric density variations. The results of the trajectory analysis show that low lift to drag ratio aerobrakes, which use angle of attack variation for control, can potentially be used for a wide range of aerobrake applications.

  6. Numerical simulation of the effects of variation of angle of attack and sweep angle on vortex breakdown over delta wings

    NASA Technical Reports Server (NTRS)

    Ekaterinaris, J. A.; Schiff, Lewis B.

    1990-01-01

    In the present investigation of the vortical flowfield structure over delta wings at high angles of attack, three-dimensional Navier-Stokes numerical simulations were conducted to predict the complex leeward flowfield characteristics; these encompass leading-edge separation, secondary separation, and vortex breakdown. Attention is given to the effect on solution accuracy of circumferential grid-resolution variations in the vicinity of the wing leading edge, and well as to the effect of turbulence modeling on the solutions. When a critical angle-of-attack was reached, bubble-type vortex breakdown was found. With further angle-of-attack increase, a change from bubble-type to spiral-type vortex breakdown was predicted by the numerical solution.

  7. Calculations of transonic boattail flow at small angle of attack

    NASA Technical Reports Server (NTRS)

    Nakayama, A.; Chow, W. L.

    1979-01-01

    A transonic flow past a boattailed afterbody under a small angle of attack was examined. It is known that the viscous effect offers significant modifications of the pressure distribution on the afterbody. Thus, the formulation for the inviscid flow was based on the consideration of a flow past a nonaxisymmetric body. The full three dimensional potential equation was solved through numerical relaxation, and quasi-axisymmetric boundary layer calculations were performed to estimate the displacement effect. It was observed again that the viscous effects were not negligible. The trend of the final results agreed well with the experimental data.

  8. High-Alfa Aerodynamics with Separated Flow Modeled as a Single Nascent Vortex

    NASA Astrophysics Data System (ADS)

    Antony, Samuel B.; Mukherjee, Rinku

    2017-04-01

    A numerical iterative vortex lattice method is developed to study flow past wing(s) at high angles of attack where the separated flow is modelled using NY nascent vortex filaments. The wing itself is modelled using NX × NY bound vortex rings, where NX and NY are the number of sections along the chord and span of the wing respectively. The strength and position of the nascent vortex along the chord corresponding to the local effective angle of attack are evaluated from the residuals in viscous and potential flow, i.e. (Cl)visc - (Cl)pot and (Cm)visc - (Cm)pot. Hence, the 2D airfoil viscous Cl - α and Cm - α is required as input (from experiment, numerical analysis or CFD). Aerodynamic characteristics and section distribution along span are predicted for 3D wings at a high angle of attack. Effect of initial conditions and existence of multiple solutions in the post-stall region is studied. Results are validated with experiment.

  9. Comparison of Angle of Attack Measurements for Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Jones, Thomas, W.; Hoppe, John C.

    2001-01-01

    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.

  10. High supersonic aerodynamic characteristics of five irregular planform wings with systematically varying wing fillet geometry tested in the NASA/LaRC 4-foot UPWT (LEG 2) (LA45A/B)

    NASA Technical Reports Server (NTRS)

    1976-01-01

    An experimental and analytical aerodynamic program to develop predesign guides for irregular planform wings is reported. The benefits are linearization of subsonic lift curve slope to high angles of attack and avoidance of subsonic pitch instabilities at high lift by proper tailoring of the planform fillet wing combination while providing the desired hypersonic trim angle and stability. The two prime areas of concern are to optimize shuttle orbiter landing and entry characteristics. Basic longitudinal aerodynamic characteristics at high supersonic speeds are developed.

  11. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 2: Influence of horizontal tail location for Model D

    NASA Technical Reports Server (NTRS)

    Barnhart, B.

    1982-01-01

    The influence of horizontal tail location on the rotational flow aerodynamics is discussed for a 1/6-scale general aviation airplane model. The model was tested using various horizontal tail positions, with both a high and a low-wing location and for each of two body lengths. Data were measured, using a rotary balance, over an angle-of-attack range of 8 to 90 deg, and for clockwise and counter-clockwise rotations covering an Omega b/2V range of 0 to 0.9.

  12. Triservice Program for Extending Missile Aerodynamic Data Base and Prediction Program Using Rational Modeling

    DTIC Science & Technology

    1983-08-01

    Aerodynamic Characteristics of Cruciform Missiles to High Angles of Attack Including Effects of Roll Angle and Control ...Deflections. NEAR TR 152, Nov., 1977. 2. Smith, C.A., and Nielsen, J.N.: Prediction of Aerodynamic Characteristics of Cruciform Missiles to High Angles... characteristics of body- tail and canard ( wing )- body- tail missiles . Under the same contract, the data base will be incorporated into

  13. Effect of underwing aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane

    NASA Technical Reports Server (NTRS)

    Abeyounis, W. K.; Patterson, J. C., Jr.

    1985-01-01

    As part of a propulsion/airframe integration program, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the longitudinal aerodynamic effects of installing flow through engine nacelles in the aft underwing position of a high wing transonic transfer airplane. Mixed flow nacelles with circular and D-shaped inlets were tested at free stream Mach numbers from 0.70 to 0.85 and angles of attack from -2.5 deg to 4.0 deg. The aerodynamic effects of installing antishock bodies on the wing and nacelle upper surfaces as a means of attaching and supporting nacelles in an extreme aft position were investigated.

  14. Review of Research On Angle-of-Attack Indicator Effectiveness

    NASA Technical Reports Server (NTRS)

    Le Vie, Lisa R.

    2014-01-01

    The National Aeronautics and Space Administration (NASA) conducted a literature review to determine the potential benefits of a display of angle-of-attack (AoA) on the flight deck of commercial transport that may aid a pilot in energy state awareness, upset recovery, and/or diagnosis of air data system failure. This literature review encompassed an exhaustive list of references available and includes studies on the benefits of displaying AoA information during all phases of flight. It also contains information and descriptions about various AoA indicators such as dial, vertical and horizontal types as well as AoA displays on the primary flight display and the head up display. Any training given on the use of an AoA indicator during the research studies or experiments is also included for review

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

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.

    1990-01-01

    Two modern high-speed advanced counterrotation propellers, F7/A7 and F7/A3 were tested in the NASA Lewis Research Center's 9- by 15-Foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach. Both rotors were of similar diameter on the F7/A7 propeller, while the aft diameter of the F7/A3 propeller was 85 percent of the forward propeller to reduce tip vortex-aft rotor interaction. The two propellers were designed for similar performance. The propellers were tested in both the baseline configuration and installed configuration consisting of a simulated upstream nacelle support pylon and fuselage section. Acoustic measurements were made with a polar microphone probe which recorded sideline directivities at various azimuthal locations. Aerodynamic measurements were also made to establish propeller operating conditions. The propellers were run at initial blade setting angles adjusted to achieve equal forward/aft torque ratios at angle of attack with the pylon and fuselage simulation in place. Data are presented for propeller operation at 80 and 90 percent of design speed (the forward rotor design tip speed was 238 m/sec (780 ft/sec). Both propellers were tested at the maximum rotor-rotor spacing of 14.99 cm (5.90 in.) based on the pitch change axis separation.

  16. Prediction of noise field of a propfan at angle of attack

    NASA Technical Reports Server (NTRS)

    Envia, Edmane

    1991-01-01

    A method for predicting the noise field of a propfan operating at an angle of attack to the oncoming flow is presented. The method takes advantage of the high-blade-count of the advanced propeller designs to provide an accurate and efficient formula for predicting their noise field. The formula, which is written in terms of the Airy function and its derivative, provides a very attractive alternative to the use of numerical integration. A preliminary comparison shows rather favorable agreement between the predictions from the present method and the experimental data.

  17. Estimation of aerodynamic parameters from flight data of a high incidence research model

    NASA Technical Reports Server (NTRS)

    Klein, V.; Mayo, M. H.

    1986-01-01

    A procedure for the determination of an aerodynamic model structure and aerodynamic parameters is applied to flight data from a high-incidence research model (HIRM) within an angle of attack range of 18 to 40 degrees. The HIRM is a three-surface unpowered model with a swept wing, an all-moving canard and stabilator, and a vertical tail with rudder. The motion of the HIRM was excited first by its release from the helicopter and then by the activation of control surfaces. This paper briefly describes the model, flight and wind tunnel data available, equations of motion and techniques for data analysis. The results presented contain an example of a measured data compatibility and the variation of some important stability derivatives with the angle of attack and canard setting. The derivatives were obtained from various maneuvers and subsets of joined data from several maneuvers by using a stepwise-regression technique. These derivatives agreed, in general, with the results of wind-tunnel measurements. The resulting lateral aerodynamic model equations could predict the motion of the HIRM reasonably well.

  18. Numerical analysis on the effect of angle of attack on evaluating radio-frequency blackout in atmospheric reentry

    NASA Astrophysics Data System (ADS)

    Jung, Minseok; Kihara, Hisashi; Abe, Ken-ichi; Takahashi, Yusuke

    2016-06-01

    A three-dimensional numerical simulation model that considers the effect of the angle of attack was developed to evaluate plasma flows around reentry vehicles. In this simulation model, thermochemical nonequilibrium of flowfields is considered by using a four-temperature model for high-accuracy simulations. Numerical simulations were performed for the orbital reentry experiment of the Japan Aerospace Exploration Agency, and the results were compared with experimental data to validate the simulation model. A comparison of measured and predicted results showed good agreement. Moreover, to evaluate the effect of the angle of attack, we performed numerical simulations around the Atmospheric Reentry Demonstrator of the European Space Agency by using an axisymmetric model and a three-dimensional model. Although there were no differences in the flowfields in the shock layer between the results of the axisymmetric and the three-dimensional models, the formation of the electron number density, which is an important parameter in evaluating radio-frequency blackout, was greatly changed in the wake region when a non-zero angle of attack was considered. Additionally, the number of altitudes at which radio-frequency blackout was predicted in the numerical simulations declined when using the three-dimensional model for considering the angle of attack.

  19. Missile Aerodynamics

    DTIC Science & Technology

    1979-02-01

    4 HIGH-ANGLE-OF-ATTACK MISSILE AERODYNAMICS by A.B.Wardlaw, Jr ECOULEMENTS DE CU LOT par i.Delery et M.Sirieix 6 THE CONTROL OF GUIDED WEAPONS by...in Figure 3, or the bomblet shown in Figure 4 . It must be conceded that the weapon aero- dynamicist does not usually have to concern himself with the...conveniently divided into the four categories shown in Figure 5: I Unguided and unpropelled. 2 Unguided and propelled. 3 Guided and unpropelled. 4 Guided

  20. Noise of the SR-3 propeller model at 2 deg and 4 deg angle of attack

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Jeracki, R. J.

    1981-01-01

    The noise effect of operating supersonic tip speed propellers at angle of attack with respect to the incoming flow was determined. Increases in the maximum blade passage noise were observed for the propeller operating at angle of attack. The noise increase was not symmetrical with one wall of the wind tunnel having significantly more noise increase than the other wall. This was apparently the result of the rotational direction of the propeller. The lack of symmetry of the noise at angle of attack to the use of oppositely rotating propellers on opposite sides of an airplane fuselage as a way of minimizing the noise due to operation at angle of attack.

  1. Post-Stall Aerodynamic Modeling and Gain-Scheduled Control Design

    NASA Technical Reports Server (NTRS)

    Wu, Fen; Gopalarathnam, Ashok; Kim, Sungwan

    2005-01-01

    A multidisciplinary research e.ort that combines aerodynamic modeling and gain-scheduled control design for aircraft flight at post-stall conditions is described. The aerodynamic modeling uses a decambering approach for rapid prediction of post-stall aerodynamic characteristics of multiple-wing con.gurations using known section data. The approach is successful in bringing to light multiple solutions at post-stall angles of attack right during the iteration process. The predictions agree fairly well with experimental results from wind tunnel tests. The control research was focused on actuator saturation and .ight transition between low and high angles of attack regions for near- and post-stall aircraft using advanced LPV control techniques. The new control approaches maintain adequate control capability to handle high angle of attack aircraft control with stability and performance guarantee.

  2. Micro air vehicle motion tracking and aerodynamic modeling

    NASA Astrophysics Data System (ADS)

    Uhlig, Daniel V.

    exhibited quasi-steady effects caused by small variations in the angle of attack. The quasi-steady effects, or small unsteady effects, caused variations in the aerodynamic characteristics (particularly incrementing the lift curve), and the magnitude of the influence depended on the angle-of-attack rate. In addition to nominal gliding flight, MAVs in general are capable of flying over a wide flight envelope including agile maneuvers such as perching, hovering, deep stall and maneuvering in confined spaces. From the captured motion trajectories, the aerodynamic characteristics during the numerous unsteady flights were gathered without the complexity required for unsteady wind tunnel tests. Experimental results for the MAVs show large flight envelopes that included high angles of attack (on the order of 90 deg) and high angular rates, and the aerodynamic coefficients had dynamic stall hysteresis loops and large values. From the large number of unsteady high angle-of-attack flights, an aerodynamic modeling method was developed and refined for unsteady MAV flight at high angles of attack. The method was based on a separation parameter that depended on the time history of the angle of attack and angle-of-attack rate. The separation parameter accounted for the time lag inherit in the longitudinal characteristics during dynamic maneuvers. The method was applied to three MAVs and showed general agreement with unsteady experimental results and with nominal gliding flight results. The flight tests with the MAVs indicate that modern motion tracking systems are capable of capturing the flight trajectories, and the captured trajectories can be used to determine the aerodynamic characteristics. From the captured trajectories, low Reynolds number MAV flight is explored in both nominal gliding flight and unsteady high angle-of-attack flight. Building on the experimental results, a modeling method for the longitudinal characteristics is developed that is applicable to the full flight

  3. Missile aerodynamics; Proceedings of the Conference, Monterey, CA, Oct. 31-Nov. 2, 1988

    SciTech Connect

    Mendenhall, M.R.; Nixon, D.; Dillenius, M.F.E.

    1989-01-01

    The present conference discusses the development status of predictive capabilities for missile aerodynamic characteristics, the application of experimental techniques to missile-release problems, prospective high-performance missile designs, the use of lateral jet controls for missile guidance, and the integration of stores on modern tactical aircraft. Also discussed are semiempirical aerodynamic methods for preliminary design, high angle-of-attack behavior for an advanced missile, and the dynamic derivatives of missiles and fighter-type configurations at high angles-of-attack.

  4. Tests of Aerodynamically Heated Multiweb Wing Structures in a Free Jet at Mach Number 2: Five Aluminum-Alloy Models of 20-Inch Chord with 0.064-Inch-Thick Skin, 0.025-Inch-Thick Webs, and Various Chordwise Stiffening at 2 deg Angle of Attack

    NASA Technical Reports Server (NTRS)

    Trussell, Donald H.; Thomson, Robert G.

    1960-01-01

    An experimental study was made on five 2024-T3 aluminum-alloy multiweb wing structures (MW-2-(4), MW-4-(3), mw-16, MW-17, and MW-18), at a Mach number of 2 and an angle of attack of 2 deg under simulated supersonic flight conditions. These models, of 20-inch chord and semi-span and 5-percent-thick circular-arc airfoil section, were identical except for the type and amount of chordwise stiffening. One model with no chordwise ribs between root and tip bulkhead fluttered and failed dynamically partway through its test. Another model with no chordwise ribs (and a thinner tip bulkhead) experienced a static bending type of failure while undergoing flutter. The three remaining models with one, two, or three chordwise ribs survived their tests. The test results indicate that the chordwise shear rigidity imparted to the models by the addition of even one chordwise rib precludes flutter and subsequent failure under the imposed test conditions. This paper presents temperature and strain data obtained from the tests and discusses the behavior of the models.

  5. Effects of angle of attack and vertical fin on transonic flutter characteristics of an arrow-wing configuration

    NASA Technical Reports Server (NTRS)

    Doggett, R. V., Jr.; Ricketts, R. A.

    1980-01-01

    Experimental transonic flutter results are presented for a simplified 1/50 size, aspect ratio 1.77, wind tunnel model of an arrow wing design. Flutter results are presented for two configurations; namely, one with and one without a ventral fin mounted at the 0.694 semispan station. Results are presented for both configurations trimmed to zero lift and in a lifting condition at angles of attack up to 4 deg. The results show that the flutter characteristics of both configurations are similar to those usually observed. Increasing angle of attack reduces the flutter dynamic pressure by a small amount (about 13 percent maximum) for both configurations. The addition of the fin to the basic wing increases the flutter dynamic pressure. Calculated results for both configurations in the nonlifting condition obtained by using subsonic doublet lattice unsteady aerodynamic theory correlate reasonably well with the experimental results. Calculated results for the basic wing obtained by using subsonic kernal function unsteady aerodynamic theory did not agree as well with the experimental data.

  6. Investigation of load prediction on the Mexico rotor using the technique of determination of the angle of attack

    NASA Astrophysics Data System (ADS)

    Yang, Hua; Shen, Wenzhong; Sørensen, Jens Nørkær; Zhu, Weijun

    2012-05-01

    Blade element moment (BEM) is a widely used technique for prediction of wind turbine aerodynamics performance, the reliability of airfoil data is an important factor to improve the prediction accuracy of aerodynamic loads and power using a BEM code. The method of determination of angle of attack on rotor blades developed by SHEN, et al is successfully used to extract airfoil data from experimental characteristics on the MEXICO (Model experiments in controlled conditions) rotor. Detailed surface pressure and particle image velocimetry (PIV) flow fields at different rotor azimuth positions are examined to determine the sectional airfoil data. The present technique uses simultaneously both PIV data and blade pressure data that include the actual flow conditions (for example, tunnel effects), therefore it is more advantageous than other techniques which only use the blade loading (pressure data). The extracted airfoil data are put into a BEM code, and the calculated axial and tangential forces are compared to both computations using BEM with Glauert's and SHEN's tip loss correction models and experimental data. The comparisons show that the present method of determination of angle of attack is correct, and the re-calculated forces have good agreements with the experiment.

  7. Rotary balance data for a single-engine trainer design for an angle-of-attack range of 8 deg to 90 deg. [conducted in langely spin tunnel

    NASA Technical Reports Server (NTRS)

    Pantason, P.; Dickens, W.

    1979-01-01

    Aerodynamic characteristics obtained in a rotational flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/6 scale, single engine trainer airplane model. The configurations tested included the basic airplane, various wing leading edge devices, elevator, aileron and rudder control settings as well as airplane components. Data are presented without analysis for an angle of attack range of 8 to 90 degrees and clockwise and counter-clockwise rotations.

  8. Identification of aerodynamic models for maneuvering aircraft

    NASA Technical Reports Server (NTRS)

    Lan, C. Edward; Hu, C. C.

    1992-01-01

    A Fourier analysis method was developed to analyze harmonic forced-oscillation data at high angles of attack as functions of the angle of attack and its time rate of change. The resulting aerodynamic responses at different frequencies are used to build up the aerodynamic models involving time integrals of the indicial type. An efficient numerical method was also developed to evaluate these time integrals for arbitrary motions based on a concept of equivalent harmonic motion. The method was verified by first using results from two-dimensional and three-dimensional linear theories. The developed models for C sub L, C sub D, and C sub M based on high-alpha data for a 70 deg delta wing in harmonic motions showed accurate results in reproducing hysteresis. The aerodynamic models are further verified by comparing with test data using ramp-type motions.

  9. Flow Control of the Stingray UAV at Low Angles of Attack

    NASA Astrophysics Data System (ADS)

    Farnsworth, John; Vaccaro, John; Amitay, Michael

    2007-11-01

    The effectiveness of active flow control, via synthetic jets and steady blowing jets, on the aerodynamic performance of the Stingray UAV was investigated experimentally in a wind tunnel. Global flow measurements were conducted using a six component sting balance, static pressure, and Particle Image Velocimetry (PIV) measurements. Using active control for trimming the Stingray UAV in pitch and roll at low angles of attack has similar effects to those with conventional control effectors. The synthetic jets were able to alter the local streamlines through the formation of a quasi-steady interaction region on the suction surface of the vehicle's wing. Phase locked data were acquired to provide insight into the growth, propagation, and decay of the synthetic jet impulse and its interaction with the cross-flow. The changes induced on the moments and forces can be proportionally controlled by either changing the momentum coefficient or by driving the synthetic jets with a pulse modulation waveform. This can lead the way for future development of closed-loop control models.

  10. Effect of location of aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane

    NASA Technical Reports Server (NTRS)

    Abeyounis, William K.; Patterson, James C., Jr.

    1990-01-01

    As part of a propulsion/airframe integration program at Langley Research Center, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of locating flow-through mixed flow engine nacelles in several aft underwing positions on the longitudinal aerodynamics of a high wing transport airplane. D-shaped inlet nacelles were used in the test. Some configurations with antishock bodies and with nacelle toe-in were also tested. Data were obtained for a free stream Mach number range of 0.70 to 0.85 and a model angle-of-attack range from -2.5 to 4.0 degrees.

  11. X-29 at High Angle of Attack with Smoke Generators

    NASA Image and Video Library

    1991-09-10

    This photo shows the X-29 during a 1991 research flight. Smoke generators in the nose of the aircraft were used to help researchers see the behavior of the air flowing over the aircraft. The smoke here is demonstrating forebody vortex flow. This mission was flown September 10, 1991, by NASA research pilot Rogers Smith.

  12. Wind Tunnel Corrections for High Angle of Attack Models,

    DTIC Science & Technology

    1981-02-01

    FROM A WALL PRESSURE SIGNATURE 1 In essence, the above analysis is an engineering solution of a set of algebraic equations involving five nonlinear...uncontrolled boundary layer on the tunnel floor, the litt’.’correlation between the four models is excellent over the lineal C|_ - a range. There was

  13. Measurements of the unsteady vortex flow over a wing-body at angle of attack

    NASA Technical Reports Server (NTRS)

    Debry, Benoit; Komerath, Narayanan M.; Liou, Shiuh-Guang; Caplin, J.; Lenakos, Jason

    1992-01-01

    Measurements of the unsteady vortex flow over a wing-body at high angles of attack were carried out on a generic test model of a pointed body of revolution with double-delta wings. Vortex patterns and trajectories were quantified from digitized laser sheet video images. The velocity-field measurements showed the jetlike flow in the unburst vortex, unsteady secondary structures below the primary core, and then the reversed flow in the burst vortex. Results of hot-film anemometry revealed the presence of peak frequencies in the velocity spectra over the wing and near the trailing edge, which varied linearly with freestream speed and increased as the measurement point moved upstream. Good Strouhal correlation was found with previous results obtained for a smaller generic wing-body model.

  14. Measurements of the unsteady vortex flow over a wing-body at angle of attack

    NASA Technical Reports Server (NTRS)

    Debry, Benoit; Komerath, Narayanan M.; Liou, Shiuh-Guang; Caplin, J.; Lenakos, Jason

    1992-01-01

    Measurements of the unsteady vortex flow over a wing-body at high angles of attack were carried out on a generic test model of a pointed body of revolution with double-delta wings. Vortex patterns and trajectories were quantified from digitized laser sheet video images. The velocity-field measurements showed the jetlike flow in the unburst vortex, unsteady secondary structures below the primary core, and then the reversed flow in the burst vortex. Results of hot-film anemometry revealed the presence of peak frequencies in the velocity spectra over the wing and near the trailing edge, which varied linearly with freestream speed and increased as the measurement point moved upstream. Good Strouhal correlation was found with previous results obtained for a smaller generic wing-body model.

  15. Assessment of aerodynamic performance of V/STOL and STOVL fighter aircraft

    NASA Technical Reports Server (NTRS)

    Nelms, W. P.

    1984-01-01

    The aerodynamic performance of V/STOL and STOVL fighter/attack aircraft was assessed. Aerodynamic and propulsion/airframe integration activities are described and small and large scale research programs are considered. Uncertainties affecting aerodynamic performance that are associated with special configuration features resulting from the V/STOL requirement are addressed. Example uncertainties relate to minimum drag, wave drag, high angle of attack characteristics, and power induced effects.

  16. Lateral aerodynamic parameters extracted from flight data for the F-8C airplane in maneuvering flight

    NASA Technical Reports Server (NTRS)

    Suit, W. T.

    1977-01-01

    Flight test data are used to extract the lateral aerodynamic parameters of the F-8C airplane at moderate to high angles of attack. The data were obtained during perturbations of the airplane from steady turns with trim normal accelerations from 1.5g to 3.0g. The angle-of-attack variation from trim was negligible. The aerodynamic coefficients extracted from flight data were compared with several other sets of coefficients, and the extracted coefficients resulted in characteristics for the Dutch roll mode (at the highest angles of attack) similar to those of a set of coefficients that have been the basis of several simulations of the F-8C.

  17. Wind-tunnel calibration and requirements for in-flight use of fixed hemispherical head angle-of-attack and angle-of-sideslip sensors

    NASA Technical Reports Server (NTRS)

    Montoya, E. J.

    1973-01-01

    Wind-tunnel tests were conducted with three different fixed pressure-measuring hemispherical head sensor configurations which were strut-mounted on a nose boom. The tests were performed at free-stream Mach numbers from 0.2 to 3.6. The boom-angle-of-attack range was -6 to 15 deg, and the angle-of-sideslip range was -6 to 6 deg. The test Reynolds numbers were from 3.28 million to 65.6 million per meter. The results were used to obtain angle-of-attack and angle-of-sideslip calibration curves for the configurations. Signal outputs from the hemispherical head sensor had to be specially processed to obtain accurate real-time angle-of-attack and angle-of-sideslip measurements for pilot displays or aircraft systems. Use of the fixed sensors in flight showed them to be rugged and reliable and suitable for use in a high temperature environment.

  18. Aerodynamic effects of flexibility in flapping wings

    PubMed Central

    Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P.

    2010-01-01

    Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re ≈ 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic

  19. Aerodynamic effects of flexibility in flapping wings.

    PubMed

    Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P

    2010-03-06

    Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re approximately 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small

  20. Rotary balance data and analysis for the X-29A airplane for an angle-of-attack range of 0 deg to 90 deg

    NASA Technical Reports Server (NTRS)

    Ralston, J. N.

    1984-01-01

    The rotational aerodynamic characteristics are discussed for a 1/8 scale model of the X-29A airplane. The effects of rotation on the aerodynamics of the basic model were determined, as well as the influence of airplane components, various control deflections, and several forebody modifications. These data were measured using a rotary balance, over an angle of attack range of 0 to 90 deg, for clockwise and counter clockwise rotations covering an omega b/2V range of 0 to 0.4.

  1. Flight Dynamics of an Aeroshell Using an Attached Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Cruz, Juan R.; Schoenenberger, Mark; Axdahl, Erik; Wilhite, Alan

    2009-01-01

    An aeroelastic analysis of the behavior of an entry vehicle utilizing an attached inflatable aerodynamic decelerator during supersonic flight is presented. The analysis consists of a planar, four degree of freedom simulation. The aeroshell and the IAD are assumed to be separate, rigid bodies connected with a spring-damper at an interface point constraining the relative motion of the two bodies. Aerodynamic forces and moments are modeled using modified Newtonian aerodynamics. The analysis includes the contribution of static aerodynamic forces and moments as well as pitch damping. Two cases are considered in the analysis: constant velocity flight and planar free flight. For the constant velocity and free flight cases with neutral pitch damping, configurations with highly-stiff interfaces exhibit statically stable but dynamically unstable aeroshell angle of attack. Moderately stiff interfaces exhibit static and dynamic stability of aeroshell angle of attack due to damping induced by the pitch angle rate lag between the aeroshell and IAD. For the free-flight case, low values of both the interface stiffness and damping cause divergence of the aeroshell angle of attack due to the offset of the IAD drag force with respect to the aeroshell center of mass. The presence of dynamic aerodynamic moments was found to influence the stability characteristics of the vehicle. The effect of gravity on the aeroshell angle of attack stability characteristics was determined to be negligible for the cases investigated.

  2. Theoretical Evaluation of the Pressures, Forces, and Moments at Hypersonic Speeds Acting on Arbitrary Bodies of Revolution Undergoing Separate and Combined Angle-of-attack and Pitching Motions

    NASA Technical Reports Server (NTRS)

    Margolis, Kenneth

    1961-01-01

    Equations based on Newtonian impact theory have been derived and a computational procedure developed with the aid of several design-type charts which enable the determination of the aerodynamic forces and moments acting on arbitrary bodies of revolution undergoing either separate or combined angle-of-attack and pitching motions. Bodies with axially increasing and decreasing cross-sectional area distributions are considered; nose shapes may be sharp, blunt, or flat faced. The analysis considers variations in angle of attack from -90 degrees to 90 degrees and allows for both positive and negative pitching rates of arbitrary magnitude. The results are also directly applicable to bodies in either separate or combined sideslip and yawing maneuvers.

  3. Flow around a slotted circular cylinder at various angles of attack

    NASA Astrophysics Data System (ADS)

    Gao, Dong-Lai; Chen, Wen-Li; Li, Hui; Hu, Hui

    2017-10-01

    We experimentally investigated the flow characteristics around a circular cylinder with a slot at different angles of attack. The experimental campaign was performed in a wind tunnel at the Reynolds number of Re = 2.67 × 104. The cylindrical test model was manufactured with a slot at the slot width S = 0.075 D ( D is the diameter of the cylinder). The angle of attack α was varied from 0° to 90°. In addition to measuring the pressure distributions around the cylinder surface, a digital particle image velocimetry (PIV) system was employed to quantify the wake flow characteristics behind the baseline cylinder (i.e., baseline case of the cylinder without slot) and slotted cylinder at various angles of attack. Measurement results suggested that at low angles of attack, the passive jet flow generated by the slot would work as an effective control scheme to modify the wake flow characteristics and contribute to reducing the drag and suppressing the fluctuating lift. The flip-flop phenomenon was also identified and discussed with the slot at 0° angle of attack. As the angle of attack α became 45°, the effects of the slot were found to be minimal. When the angle of attack α of the slot approached 90°, the self-organized boundary layer suction and blowing were realized. As a result, the flow separations on both sides of the test model were found to be notably delayed, the wake width behind the slotted cylinder was decreased and the vortex formation length was greatly shrunk, in comparison with the baseline case. Instantaneous pressure measurement results revealed that the pressure difference between the two slot ends and the periodically fluctuating pressure distributions would cause the alternative boundary layer suction and blowing at α = 90°.

  4. Aerodynamics of the Cyclogyro

    NASA Astrophysics Data System (ADS)

    Iosilevski, Gil; Levy, Yuval; Weihs, Daniel

    2001-11-01

    The Cyclogiro is the name given by NASA researchers in the '30s to an aerodynamic configuration of several large aspect ratio rectangular airfoils with horizontal span, placed on the circumference of a vertical circle of radius of the order of the airfoil chord, and rotating around the circle center at high speed, with periodically changing angle of attack. This configuration produces aerodynamic forces that can be applied to lift and thrust, depending on the phase angle between the instantaneous position and angle of attack. The original approach was to install such rotors instead of an aircraft wing, and thus combine the lift & thrust producing functions. As a result of the state of knowledge of unsteady aerodynamics at the time disparities between predictions and measured forces remained unexplained. This, combined with low efficiency resulted in the concept being abandoned. In the present study the concept is revisited, as a possible propulsor/lift generator for a hover-capable micro-UAV. Preliminary analysis showed that scaling down to rotor airfoil sizes of 10-15 cm span and 2 cm chord will reduce the centrifugal forces to manageable proportions while the aerodynamic forces would be comparable to those obtained by conventional rotors. A series of experiments was performed, showing disparities of up to 30theory. Visualization showed that this difference resulted mainly from interactions between single foil wakes with the following foils, and a numerical study confirmed the magnitude of the effects, in good agreement with the experiments.

  5. Effects of wing deformation on aerodynamic forces in hovering hoverflies.

    PubMed

    Du, Gang; Sun, Mao

    2010-07-01

    We studied the effects of wing deformation on the aerodynamic forces of wings of hovering hoverflies by solving the Navier-Stokes equations on a dynamically deforming grid, employing the recently measured wing deformation data of hoverflies in free-flight. Three hoverflies were considered. By taking out the camber deformation and the spanwise twist deformation one by one and by comparing the results of the deformable wing with those of the rigid flat-plate wing (the angle of attack of the rigid flat-plate wing was equal to the local angle of attack at the radius of the second moment of wing area of the deformable wing), effects of camber deformation and spanwise twist were identified. The main results are as follows. For the hovering hoverflies considered, the time courses of the lift, drag and aerodynamic power coefficients of the deformable wing are very similar to their counterparts of the rigid flat-plate wing, although lift of the deformable wing is about 10% larger, and its aerodynamic power required about 5% less than that of the rigid flat-plate wing. The difference in lift is mainly caused by the camber deformation, and the difference in power is mainly caused by the spanwise twist. The main reason that the deformation does not have a very large effect on the aerodynamic force is that, during hovering, the wing operates at a very high angle of attack (about 50 deg) and the flow is separated, and separated flow is not very sensitive to wing deformation. Thus, as a first approximation, the deformable wing in hover flight could be modeled by a rigid flat-plate wing with its angle of attack being equal to the local angle of attack at the radius of second moment of wing area of the deformable wing.

  6. Wind turbine trailing edge aerodynamic brakes

    SciTech Connect

    Migliore, P G; Miller, L S; Quandt, G A

    1995-04-01

    Five trailing-edge devices were investigated to determine their potential as wind-turbine aerodynamic brakes, and for power modulation and load alleviation. Several promising configurations were identified. A new device, called the spoiler-flap, appears to be the best alternative. It is a simple device that is effective at all angles of attack. It is not structurally intrusive, and it has the potential for small actuating loads. It is shown that simultaneous achievement of a low lift/drag ratio and high drag is the determinant of device effectiveness, and that these attributes must persist up to an angle of attack of 45{degree}. It is also argued that aerodynamic brakes must be designed for a wind speed of at least 45 m/s (100 mph).

  7. Tactical missile aerodynamics - General topics. Progress in Astronautics and Aeronautics. Vol. 141

    SciTech Connect

    Hemsch, M.J. )

    1992-01-01

    The present volume discusses the development history of tactical missile airframes, aerodynamic considerations for autopilot design, a systematic method for tactical missile design, the character and reduction of missile observability by radar, the visualization of high angle-of-attack flow phenomena, and the behavior of low aspect ratio wings at high angles of attack. Also discussed are airbreathing missile inlets, 'waverider' missile configurations, bodies with noncircular cross-sections and bank-to-turn missiles, asymmetric flow separation and vortex shedding on bodies-of-revolution, unsteady missile flows, swept shock-wave/boundary-layer interactions, pylon carriage and separation of stores, and internal stores carriage and separation.

  8. Noise of the SR-6 propeller model at 2 deg and 4 deg angles of attack

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Stefko, G. L.

    1983-01-01

    The noise generated by supersonic-tip-speed propellers creates a cabin noise problem for future airplanes powered by these propellers. Noise of a number of propeller models were measured in the NASA Lewis 8- by 6-Foot Wind Tunnel with flow parallel to the propeller axis. In flight, as a result of the induced upwash from the airplane wing, the propeller is at an angle of attack with respect to the incoming flow. Therefore, the 10-blade SR-6 propeller was operated at angle of attack to determine its noise behavior. Higher blade passage tones were observed for the propeller operating at angle of attack in a 0.6 axial Mach number flow. The noise increase was not symmetrical, with one wall of the wind tunnel showing a larger noise increase than the other wall. No noise increase was observed at angle of attack in a 0.8 axial Mach number flow. For this propeller the dominance of thickness noise, which does not increase with angle of attack, explains the lack of noise increase at the higher 0.8 Mach number.

  9. In-blade angle of attack measurement and comparison with models

    NASA Astrophysics Data System (ADS)

    Gallant, T. E.; Johnson, D. A.

    2016-09-01

    The torque generated by a wind turbine blade is dependent on several parameters, one of which is the angle of attack. Several models for predicting the angle of attack in yawed conditions have been proposed in the literature, but there is a lack of experimental data to use for direct validation. To address this problem, experiments were conducted at the University of Waterloo Wind Generation Research Facility using a 3.4 m diameter test turbine. A five-hole pressure probe was installed in a modular 3D printed blade and was used to measure the angle of attack, a, as a function of several parameters. Measurements were conducted at radial positions of r/R = 0.55 and 0.72 at tip speed ratios of λ = 5.0, 3.6, and 3.1. The yaw offset of the turbine was varied from -15° to +15°. Experimental results were compared directly to angle of attack values calculated using a model proposed by Morote in 2015. Modeled values were found to be in close agreement with the experimental results. The angle of attack was shown to vary cyclically in the yawed case while remaining mostly constant when aligned with the flow, as expected. The quality of results indicates the potential of the developed instrument for wind turbine measurements.

  10. Analysis of flight data from a High-Incidence Research Model by system identification methods

    NASA Technical Reports Server (NTRS)

    Batterson, James G.; Klein, Vladislav

    1989-01-01

    Data partitioning and modified stepwise regression were applied to recorded flight data from a Royal Aerospace Establishment high incidence research model. An aerodynamic model structure and corresponding stability and control derivatives were determined for angles of attack between 18 and 30 deg. Several nonlinearities in angles of attack and sideslip as well as a unique roll-dominated set of lateral modes were found. All flight estimated values were compared to available wind tunnel measurements.

  11. Influence of Thickness and Angle of Attack on the Dynamics of Rectangular Cylinder Wakes

    NASA Astrophysics Data System (ADS)

    Mohebi, Meraj

    Stereoscopic Particle Image Velocimetry measurements were taken in the turbulent wake of two-dimensional rectangular cylinders. The influence of post-stall angles of attack and Reynolds number on the flow behind a thin at plate, and for the normal case, the effect of thickness to chord (t=d) ratio over a family of rectangular cylinders were investigated. At all cases, quasi-periodic vortex shedding is observed, the normal direction Reynolds stress becomes very large just downstream of the mean recirculation zone, and the spanwise motions were uncorrelated to the main vortex shedding process. The data were processed to obtain the mean velocities, Reynolds stresses, and forces on the body. All terms in the turbulent kinetic energy equations were measured with the exception of dissipation which was found by difference. The pressure-related terms were estimated from the numerical solution of the Poisson equation for the instantaneous velocity field. Proper Orthogonal Decomposition modes are related via mean-field theory to construct generalized phase-averaging and low-order models capturing coherent cycle-to-cycle variations. The advection, production and pressure diffusion were all significant and mostly coherent. It is shown that high, average, and low amplitude vortex shedding cycles are different in terms of vortex street dimensions, vortex topology, circulation, and decay rate. It is also shown that these flows experience irregular significant decreases in the shedding amplitude associated with shedding of disorganized vortices in a large wake. Reynolds number was found to have imperceptible effects on the wake of a normal thin plate. A reduction in the angle of attack caused the wake to decrease in size and increase in shedding frequency but the global characteristics vary non-linearly. An increase in thickness from thin plate (t=d=0.05), caused the wake to shrink, low cycles to diminish, and local turbulence increase to a peak at t=d=1.0, identified as a

  12. Investigation on the movement of vortex burst position with dynamically changing angle of attack for a schematic deltawing in a watertunnel with correlation to similar studies in windtunnel

    NASA Astrophysics Data System (ADS)

    Wolffelt, Karl W.

    1987-06-01

    The requirements for modern military aircraft to maintain good handling qualities at very high angles of attack is one of many reasons why an increased knowledge is necessary regarding the aerodynamic behavior of vortex flows at nonstationary conditions. Linearized theory as it has been utilized in flight mechanics simulation using damping derivatives derived from forced oscillation technique, for example, may no longer be valid at such conditions. With this background some investigations have been made by SAAB-SCANIA with the aim to study the hysteresis effects for nonstationary vortex flows. A schematic delta-wing model which could also be equipped with a similar canard wing has been tested in a water tunnel. The model was supported in the tunnel by a simple mechanism by which it could be forced to move in one of four different modes, pitching or plunging with either ramp or harmonic motion. The flow over the model was visualized with air bubbles and sequences were recorded on videotape. The sequences were analyzed and the movements of the leading edge vortex burst have been studied with the main interest focused on the hysteresis effects.

  13. Effects of Angle of Attack and Velocity on Trailing Edge Noise

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2004-01-01

    Trailing edge (TE) noise measurements for a NACA 63-215 airfoil model are presented, providing benchmark experimental data for a cambered airfoil. The effects of flow Mach number and angle of attack of the airfoil model with different TE bluntnesses are shown. Far-field noise spectra and directivity are obtained using a directional microphone array. Standard and diagonal removal beamforming techniques are evaluated employing tailored weighting functions for quantitatively accounting for the distributed line character of TE noise. Diagonal removal processing is used for the primary database as it successfully removes noise contaminates. Some TE noise predictions are reported to help interpret the data, with respect to flow speed, angle of attack, and TE bluntness on spectral shape and peak levels. Important findings include the validation of a TE noise directivity function for different airfoil angles of attack and the demonstration of the importance of the directivity function s convective amplification terms.

  14. Effects of Angle of Attack and Velocity on Trailing Edge Noise

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2006-01-01

    Trailing edge (TE) noise measurements for a NACA 63-215 airfoil model are presented, providing benchmark experimental data for a cambered airfoil. The effects of flow Mach number and angle of attack of the airfoil model with different TE bluntnesses are shown. Far-field noise spectra and directivity are obtained using a directional microphone array. Standard and diagonal removal beamforming techniques are evaluated employing tailored weighting functions for quantitatively accounting for the distributed line character of TE noise. Diagonal removal processing is used for the primary database as it successfully removes noise contaminates. Some TE noise predictions are reported to help interpret the data, with respect to flow speed, angle of attack, and TE bluntness on spectral shape and peak levels. Important findings include the validation of a TE noise directivity function for different airfoil angles of attack and the demonstration of the importance of the directivity function s convective amplification terms.

  15. A two camera video imaging system with application to parafoil angle of attack measurements

    NASA Technical Reports Server (NTRS)

    Meyn, Larry A.; Bennett, Mark S.

    1991-01-01

    This paper describes the development of a two-camera, video imaging system for the determination of three-dimensional spatial coordinates from stereo images. This system successfully measured angle of attack at several span-wise locations for large-scale parafoils tested in the NASA Ames 80- by 120-Foot Wind Tunnel. Measurement uncertainty for angle of attack was less than 0.6 deg. The stereo ranging system was the primary source for angle of attack measurements since inclinometers sewn into the fabric ribs of the parafoils had unknown angle offsets acquired during installation. This paper includes discussions of the basic theory and operation of the stereo ranging system, system measurement uncertainty, experimental set-up, calibration results, and test results. Planned improvements and enhancements to the system are also discussed.

  16. Improved Correction System for Vibration Sensitive Inertial Angle of Attack Measurement Devices

    NASA Technical Reports Server (NTRS)

    Crawford, Bradley L.; Finley, Tom D.

    2000-01-01

    Inertial angle of attack (AoA) devices currently in use at NASA Langley Research Center (LaRC) are subject to inaccuracies due to centrifugal accelerations caused by model dynamics, also known as sting whip. Recent literature suggests that these errors can be as high as 0.25 deg. With the current AoA accuracy target at LaRC being 0.01 deg., there is a dire need for improvement. With other errors in the inertial system (temperature, rectification, resolution, etc.) having been reduced to acceptable levels, a system is currently being developed at LaRC to measure and correct for the sting-whip-induced errors. By using miniaturized piezoelectric accelerometers and magnetohydrodynamic rate sensors, not only can the total centrifugal acceleration be measured, but yaw and pitch dynamics in the tunnel can also be characterized. These corrections can be used to determine a tunnel's past performance and can also indicate where efforts need to be concentrated to reduce these dynamics. Included in this paper are data on individual sensors, laboratory testing techniques, package evaluation, and wind tunnel test results on a High Speed Research (HSR) model in the Langley 16-Foot Transonic Wind Tunnel.

  17. Impingement of Water Droplets on NACA 65A004 Airfoil at 8 deg Angle of Attack

    NASA Technical Reports Server (NTRS)

    Brun, R. J.; Gallagher, H. M.; Vogt, D. E.

    1954-01-01

    The trajectories of droplets in the air flowing past an NACA 65AO04 airfoil at an angle of attack of 8 deg were determined.. The amount of water in droplet form impinging on the airfoil, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface were calculated from the trajectories and presented to cover a large range of flight and atmospheric conditions. These impingement characteristics are compared briefly with those previously reported for the same airfoil at an angle of attack of 4 deg.

  18. Low-speed aerodynamic characteristics of a 42 deg swept high-wing model having a double-slotted flap system and a supercritical airfoil

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Goodson, K. W.

    1974-01-01

    A low-speed investigation was conducted over an angle-of-attack range from about -4 deg to 20 deg in the Langley V/STOL tunnel to determine the effects of a double-slotted flap, high-lift system on the aerodynamic characteristics of a 42 deg swept high-wing model having a supercritical airfoil. The wing had an aspect ratio of 6.78 and a taper ratio of 0.36; the double-slotted flap consisted of a 35-percent-chord flap with a 15-percent-chord vane. The model was tested with a 15-percent-chord leading-edge slat.

  19. Aerodynamic characteristics of a high-wing transport configuration with a over-the-wing nacelle-pylon arrangement

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.; Abeyounis, W. K.

    1985-01-01

    An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on the aerodynamic characteristics of a high-wing transport configuration of installing an over-the-wing nacelle-pylon arrangement. The tests are conducted at Mach numbers from 0.70 to 0.82 and at angles of attack from -2 deg to 4 deg. The configurational variables under study include symmetrical and contoured nacelles and pylons, pylon size, and wing leading-edge extensions. The symmetrical nacelles and pylons reduce the lift coefficient, increase the drag coefficient, and cause a nose-up pitching-moment coefficient. The contoured nacelles significantly reduce the interference drag, though it is still excessive. Increasing the pylon size reduces the drag, whereas adding wing leading-edge extension does not affect the aerodynamic characteristics significantly.

  20. Stability of Supersonic Boundary Layers on a Cone at an Angle of Attack

    NASA Technical Reports Server (NTRS)

    Balakumar, Ponnampalam

    2009-01-01

    The stability and receptivity of three-dimensional supersonic boundary layers over a 7deg sharp tipped straight cone at an angle of attack of 4.2deg is numerically investigated at a free stream Mach number of 3.5 and at two high Reynolds numbers, 0.25 and 0.50x10(exp 6)/inch. The generation and evolution of stationary crossflow vortices are also investigated by performing simulations with three-dimensional roughness elements located on the surface of the cone. The flow fields with and without the roughness elements are obtained by solving the full Navier-Stokes equations in cylindrical coordinates using the fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and using the third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. Stability computations reveal that the azimuthal wavenumbers are in the range of m approx. 25-50 for the most amplified traveling disturbances and in the range of m approx. 40-70 for the stationary disturbances. The N-Factor computations predicted that transition would occur further forward in the middle of the cone compared to the transition fronts near the windward and the leeward planes. The simulations revealed that the crossflow vortices originating from the nose region propagate towards the leeward plane. No perturbations were observed in the lower part of the cone.

  1. Unsteady Euler analysis of the flow field of a propfan at an angle of attack

    NASA Technical Reports Server (NTRS)

    Nallasamy, M.; Groeneweg, J. F.

    1990-01-01

    The effects of angle of attack of a propfan on the blade loading and details of the flow field by solving the unsteady three-dimensional Euler equations are examined. The configuration considered is the SR7L propeller at cruise condition and the inflow angles considered are 4.6 degrees, 1.6 degrees and -0.4 degrees. The results indicate that the blade response is nearly sinusoidal at low inflow angles (1.6 degrees and -0.4 degrees) and significant deviations from sinusoidal behavior occur at an inflow angle of 4.6 degrees due to the presence of strong shocks on both suction and pressure surfaces of the blade. The detailed flow in the blade passages shows that a shock formed on the suction surface during the highly loaded portion of the revolution extends across the passage to the pressure surface. An increase in inflow angle results in an increase in blade loading on the down-going side and a decrease in loading on the up-going side.

  2. Rotary balance data for a single-engine agricultural airplane configuration for an angle-of-attack range of 8 deg to 90 deg

    NASA Technical Reports Server (NTRS)

    Mulcay, W. J.; Chu, J.

    1980-01-01

    Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/10 scale single engine agricultural airplane model. The configurations tested include the basic airplane, various wing leading edge and wing tip devices, elevator, aileron, and rudder control settings, and other modifications. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg, and clockwise and counter-clockwise rotations covering a spin coefficient range from 0 to .9.

  3. Introductory remarks. [fluid mechanics research for the National Transonic Facility: theoretical aerodynamics

    NASA Technical Reports Server (NTRS)

    Gessow, A.

    1977-01-01

    Suggested fluid mechanics research to be conducted in the National Transonic Facility include: wind tunnel calibration; flat plate skin friction, flow visualization and measurement techniques; leading edge separation; high angle of attack separation; shock-boundary layer interaction; submarine shapes; low speed studies of cylinder normal to flow; and wall interference effects. These theoretical aerodynamic investigations will provide empirical inputs or validation data for computational aerodynamics, and increase the usefulness of existing wind tunnels.

  4. Effects of wing leading-edge deflection on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Weston, R. P.

    1978-01-01

    Wing leading-edge deflection effects on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration were determined. Static force tests were conducted in a V/STOL tunnel at a Reynolds number of about 2.5 x 1 million for an angle-of-attack range from -10 deg to 17 deg and an angle-of-sideslip range from -5 deg to 5 deg. Limited flow visualization studies were also conducted in order to provide a qualitative assessment of leading-edge upwash characteristics.

  5. Low-speed longitudinal and lateral-directional aerodynamic characteristics of the X-31 configuration

    NASA Technical Reports Server (NTRS)

    Banks, Daniel W.; Gatlin, Gregory M.; Paulson, John W., Jr.

    1992-01-01

    An experimental investigation of a 19 pct. scale model of the X-31 configuration was completed in the Langley 14 x 22 Foot Subsonic Tunnel. This study was performed to determine the static low speed aerodynamic characteristics of the basic configuration over a large range of angle of attack and sideslip and to study the effects of strakes, leading-edge extensions (wing-body strakes), nose booms, speed-brake deployment, and inlet configurations. The ultimate purpose was to optimize the configuration for high angle of attack and maneuvering-flight conditions. The model was tested at angles of attack from -5 to 67 deg and at sideslip angles from -16 to 16 deg for speeds up to 190 knots (dynamic pressure of 120 psf).

  6. Heat transfer characteristics of staggered wing-shaped tubes bundle at different angles of attack

    NASA Astrophysics Data System (ADS)

    Sayed Ahmed, Sayed Ahmed E.; Ibrahiem, Emad Z.; Mesalhy, Osama M.; Abdelatief, Mohamed A.

    2014-08-01

    An experimental and numerical study has been conducted to clarify heat transfer characteristics and effectiveness of a cross-flow heat exchanger employing staggered wing-shaped tubes at different angels of attack. The water-side Rew and the air-side Rea were at 5 × 102 and at from 1.8 × 103 to 9.7 × 103, respectively. The tubes arrangements were employed with various angles of attack θ1,2,3 from 0° to 330° at the considered Rea range. Correlation of Nu, St, as well as the heat transfer per unit pumping power (ɛ) in terms of Rea and design parameters for the studied bundle were presented. The temperature fields around the staggered wing-shaped tubes bundle were predicted by using commercial CFD FLUENT 6.3.26 software package. Results indicated that the heat transfer increased with the angle of attack in the range from 0° to 45°, while the opposite was true for angles of attack from 135° to 180°. The best thermal performance and hence the efficiency η of studied bundle occurred at the lowest Rea and/or zero angle of attack. Comparisons between the experimental and numerical results of the present study and those, previously, obtained for similar available studies showed good agreements.

  7. The N.A.C.A. Recording Tachometer and Angle of Attack Recorder

    NASA Technical Reports Server (NTRS)

    Reid, H J E

    1923-01-01

    This note contains photos and descriptions of airplane flight apparatus for use in conjunction with a recording galvanometer. In measuring the angle of attack a variable resistance is used, being controlled by a vane in the airstream. Thus it is only necessary to measure the change of resistance.

  8. High lift aerodynamics

    NASA Technical Reports Server (NTRS)

    Sullivan, John; Schneider, Steve; Campbell, Bryan; Bucci, Greg; Boone, Rod; Torgerson, Shad; Erausquin, Rick; Knauer, Chad

    1994-01-01

    The current program is aimed at providing a physical picture of the flow physics and quantitative turbulence data of the interaction of a high Reynolds number wake with a flap element. The impact of high lift on aircraft performance is studied for a 150 passenger transport aircraft with the goal of designing optimum high lift systems with minimum complexity.

  9. Three-dimensional compressible laminar boundary layers on sharp and blunt circular cones at angle of attack

    NASA Technical Reports Server (NTRS)

    Popinski, Z.; Davis, R. T.

    1973-01-01

    A method for solving the three-dimensional compressible laminar boundary layer equations for the case of a circular cone and a sphere-cone body at an angle of attack is presented. The governing equations are modified by a similarity type transformation and then transformed into a Crocco-type form. The resulting set of equations is solved simultaneously by an iterative method using an implicit finite difference scheme by means of an efficient algorithm for equations of tridiagonal form. The effects of streamline swallowing on a sharp cone are included by introducing the true inviscid edge conditions at the distance from the wall equal to the boundary layer thickness. The validity of the approach was established by comparison of the computational results with similar results by other methods and with experimental data. It was concluded that at sufficiently high Mach number and moderate to large angles of attack, the streamline swallowing effects on a sharp cone result in higher values of skin friction and heat transfer as compared with the classical results for constant entropy.

  10. Comparison of theoretically predicted lateral-directional aerodynamic characteristics with full-scale wind tunnel data on the ATLIT airplane

    NASA Technical Reports Server (NTRS)

    Griswold, M.; Roskam, J.

    1980-01-01

    An analytical method is presented for predicting lateral-directional aerodynamic characteristics of light twin engine propeller-driven airplanes. This method is applied to the Advanced Technology Light Twin Engine airplane. The calculated characteristics are correlated against full-scale wind tunnel data. The method predicts the sideslip derivatives fairly well, although angle of attack variations are not well predicted. Spoiler performance was predicted somewhat high but was still reasonable. The rudder derivatives were not well predicted, in particular the effect of angle of attack. The predicted dynamic derivatives could not be correlated due to lack of experimental data.

  11. Wind-tunnel investigation of the flow correction for a model-mounted angle of attack sensor at angles of attack from -10 deg to 110 deg. [Langley 12-foot low speed wind tunnel test

    NASA Technical Reports Server (NTRS)

    Moul, T. M.

    1979-01-01

    A preliminary wind tunnel investigation was undertaken to determine the flow correction for a vane angle of attack sensor over an angle of attack range from -10 deg to 110 deg. The sensor was mounted ahead of the wing on a 1/5 scale model of a general aviation airplane. It was shown that the flow correction was substantial, reaching about 15 deg at an angle of attack of 90 deg. The flow correction was found to increase as the sensor was moved closer to the wing or closer to the fuselage. The experimentally determined slope of the flow correction versus the measured angle of attack below the stall angle of attack agreed closely with the slope of flight data from a similar full scale airplane.

  12. Wind tunnel investigation of the aerodynamic characteristics of symmetrically deflected ailerons of the F-8C airplane. [conducted in the Langley 8-foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Gera, J.

    1977-01-01

    A .042-scale model of the F-8C airplane was investigated in a transonic wind tunnel at high subsonic Mach numbers and a range of angles of attack between-3 and 20 degrees. The effect of symmetrically deflected ailerons on the longitudinal aerodynamic characteristics was measured. Some data were also obtained on the lateral control effectiveness of asymmetrically deflected horizontal tail surfaces.

  13. 75 FR 80735 - Special Conditions: Gulfstream Model GVI Airplane; High Incidence Protection

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-12-23

    ... which an aerodynamic stall would occur. (c) Alpha-Limit--The maximum angle of attack at which the... weight; S = Aerodynamic reference wing area; and q = Dynamic pressure. (e) V SR must be determined with... other drag devices when used as air brakes, and thrust reversers. High incidence maneuver demonstrations...

  14. A New Method for Calculating Wing Along Aerodynamics to Angle of Attack 180 deg

    DTIC Science & Technology

    1994-03-01

    FY94, funding for documentation was provided by the Air Weaponry Technology Program managed at the Naval Air Warfare Center, China Lake , California, by...WARMINSTER NAVAL AIR WARFARE CENTER WARMINSTER PA 18974-5C00 WEAPONS DIVI.-ION CHINA LAKE CA 93555-6001 ATTN HEAD WEAPONS DEPT 1 HEAD SCIENCE DEPT 1 ATTN...AUBURN UNIVERSITY AL 36849-5338 CANOGA PARK CA 91304-7928 ATTN ROBERT ENGLAR 1 ATTN M DILLENIUS GEORGIA TECH RESEARCH INSTITUTE NIELSEN ENGINEERING AND

  15. Satellite Aerodynamics and Density Determination from Satellite Dynamic Response

    NASA Technical Reports Server (NTRS)

    Karr, G. R.

    1972-01-01

    The aerodynamic drag and lift properties of a satellite are first expressed as a function of two parameters associated with gas-surface interaction at the satellite surface. The dynamic response of the satellite as it passes through the atmosphere is then expressed as a function of the two gas-surface interaction parameters, the atmospheric density, the satellite velocity, and the satellite orientation to the high speed flow. By proper correlation of the observed dynamic response with the changing angle of attack of the satellite, it is found that the two unknown gas-surface interaction parameters can be determined. Once the gas-surface interaction parameters are known, the aerodynamic properties of the satellite at all angles of attack are also determined.

  16. GASP- General Aviation Synthesis Program. Volume 3: Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hague, D.

    1978-01-01

    Aerodynamics calculations are treated in routines which concern moments as they vary with flight conditions and attitude. The subroutines discussed: (1) compute component equivalent flat plate and wetted areas and profile drag; (2) print and plot low and high speed drag polars; (3) determine life coefficient or angle of attack; (4) determine drag coefficient; (5) determine maximum lift coefficient and drag increment for various flap types and flap settings; and (6) determine required lift coefficient and drag coefficient in cruise flight.

  17. Aerodynamics of the EXPERT Reentry Capsule Along the Descent Trajectory

    NASA Astrophysics Data System (ADS)

    Vashchenkov, P.; Kashkovsky, A.; Ivanov, M.

    2009-01-01

    Results of numerical simulations of high-altitude aero thermodynamics of the EXPERT reentry capsule along its descent trajectory are presented. Aerodynamic characteristics for different angles of attack and rolling of the capsule at altitude of 150 down to 20 km are studied. An engineering local bridging method is used in computations. The uncertainty of the engineering method in the transitional regime is determined by comparisons with results obtained by DSMC simulations.

  18. Investigation of nose bluntness and angle of attack effects on slender bodies in viscous hypersonic flows

    NASA Technical Reports Server (NTRS)

    Sehgal, A. K.; Tiwari, S. N.; Singh, D. J.

    1991-01-01

    Hypersonic flows over cones and straight biconic configurations are calculated for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. Effect of angle of attack and nose bluntness on these slender cones in air is studied extensively. The numerical procedures are based on the solution of complete Navier-Stokes equations at the nose section and parabolized Navier-Stokes equations further downstream. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The complete flow field is thoroughly analyzed with respect to velocity, temperature, pressure, and entropy profiles. The post shock flow field is studied in detail from the contour plots of Mach number, density, pressure, and temperature. The effect of nose bluntness for slender cones persists as far as 200 nose radii downstream.

  19. Laser velocimeter survey about a NACA 0012 wing at low angles of attack

    NASA Technical Reports Server (NTRS)

    Hoad, D. R.; Meyers, J. F.; Young, W. H., Jr.; Hepner, T. E.

    1978-01-01

    An investigation was conducted in the Langley V/STOL tunnel with a laser velocimeter to obtain measurements of airflow velocities about a wing at low angles of attack. The applicability of the laser velocimeter technique for this purpose in the V/STOL tunnel was demonstrated in this investigation with measurement precision bias calculated at -1.33 percent to 0.91 percent and a random uncertainty calculated at + or - 0.47 percent. Free stream measurements were obtained with this device and compared with velocity calculations from pitot static probe data taken near the laser velocimeter measurement location. The two measurements were in agreement to within 1 percent. Velocity measurement results about the centerline at 0.6 degrees angle of attack were typically those expected. At 4.75 degrees, the velocity measurements indicated that a short laminar separation bubble existed near the leading edge with an oscillating shear layer.

  20. Time-marching transonic flutter solutions including angle-of-attack effects

    NASA Technical Reports Server (NTRS)

    Edwards, J. W.; Bennett, R. M.; Whitlow, W., Jr.; Seidel, D. A.

    1982-01-01

    Transonic aeroelastic solutions based upon the transonic small perturbation potential equation were studied. Time-marching transient solutions of plunging and pitching airfoils were analyzed using a complex exponential modal identification technique, and seven alternative integration techniques for the structural equations were evaluated. The HYTRAN2 code was used to determine transonic flutter boundaries versus Mach number and angle-of-attack for NACA 64A010 and MBB A-3 airfoils. In the code, a monotone differencing method, which eliminates leading edge expansion shocks, is used to solve the potential equation. When the effect of static pitching moment upon the angle-of-attack is included, the MBB A-3 airfoil can have multiple flutter speeds at a given Mach number.

  1. Doppler radiation sensing of Shuttle angle of attack and TAS during entry

    NASA Astrophysics Data System (ADS)

    Foale, C. M.

    Space Shuttle true airspeed, angle of attack, and sideslip angle are currently derived from inertial guidance information. A new method is proposed which offers a potential improvement in Shuttle safety during entry. Angle of attack, sideslip angle and true airspeed could be measured directly at heights from 120 km down to 20 km by Doppler sensing three independent true airspeeds along the Shuttle body axes. Two types of Doppler measurement sensors, employing either passive detection of atmospheric radiation or coherent detection of scattered laser light are discussed. The proposed technique is essentially solid-state and robust, and is well suited for use in future small hypersonic vehicles that require flight control in the Upper Atmosphere of the earth or in probes destined for the other planets.

  2. Doppler radiation sensing of Shuttle angle of attack and TAS during entry

    NASA Technical Reports Server (NTRS)

    Foale, C. M.

    1984-01-01

    Space Shuttle true airspeed, angle of attack, and sideslip angle are currently derived from inertial guidance information. A new method is proposed which offers a potential improvement in Shuttle safety during entry. Angle of attack, sideslip angle and true airspeed could be measured directly at heights from 120 km down to 20 km by Doppler sensing three independent true airspeeds along the Shuttle body axes. Two types of Doppler measurement sensors, employing either passive detection of atmospheric radiation or coherent detection of scattered laser light are discussed. The proposed technique is essentially solid-state and robust, and is well suited for use in future small hypersonic vehicles that require flight control in the Upper Atmosphere of the earth or in probes destined for the other planets.

  3. Effects of wing-leading-edge modifications on a full-scale, low-wing general aviation airplane: Wind-tunnel investigation of high-angle-of-attack aerodynamic characteristics. [conducted in Langley 30- by 60-foot tunnel

    NASA Technical Reports Server (NTRS)

    Newsom, W. A., Jr.; Satran, D. R.; Johnson, J. L., Jr.

    1982-01-01

    Wing-leading-edge modifications included leading-edge droop and slat configurations having full-span, partial-span, or segmented arrangements. Other devices included wing-chord extensions, fences, and leading-edge stall strips. Good correlation was apparent between the results of wind-tunnel data and the results of flight tests, on the basis of autorotational stability criterion, for a wide range of wing-leading-edge modifications.

  4. Formation of asymmetric separated flow past slender bodies of revolution at large angles of attack

    NASA Technical Reports Server (NTRS)

    Goman, M. G.; Khrabrov, A. N.

    1986-01-01

    The paper examines the problem of determining stationary positions of pairs of vortices of unequal intensity in the flow behind a cylinder modeling the axisymmetric separated flow past a slender body at large angles of attack. The possible asymmetric stationary positions of two vortices are calculated, and their stability with respect to small perturbations is determined. Bifurcations of the flow field with changes in vortex intensity are analyzed.

  5. Response of Earth Penetrator Structures in Angle-of-Attack Impacts

    DTIC Science & Technology

    1977-06-01

    decelerates by impacting the energy-absorbing aluminum honeycomb. Experimental Setup To simulate the load on a penetrator that impacts at an angle of attack...absorber ( aluminum honeycomb or Styrofoam) located in a 5-foot-long (1.52-m-long) safety shroud. The shroud ensures containment of the model. penetrator...uniform cross section (except at the loaded hemispherical end). The solid aluminum model is geometrically identical to the solid steel model but is made

  6. Effects of nose bluntness and angle of attack on slender bodies in hypersonic flows

    NASA Technical Reports Server (NTRS)

    Tiwari, S. N.; Sehgal, A. K.; Singh, D. J.

    1992-01-01

    The effects of angle of attack and nose bluntness on the flow field and wall quantities are investigated for hypersonic flows of air over slender bodies. The bodies considered are slender cones and straight biconic configurations. The numerical procedures used are based on the solution of complete Navier-Stokes equations in the nose region and parabolized Navier-Stokes equations in the downstream region. Results are obtained for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The postshock flow field is studied in detail from the contour plots of Mach number, density, and temperature. Flow separation is observed on the leeward plane for an on-axis, 12.84 deg/7 deg (fore-cone and aft-cone angles) biconic geometry at 12 deg angle of attack. Also, the windward and leeward heating rates for the fore-cone section decrease by a factor of four and five, respectively, when the nose bluntness is increased by an order of magnitude. The effect of nose bluntness for slender cone persists as far as 200 nose radii downstream.

  7. Cruise noise of counterrotation propeller at angle of attack in wind tunnel

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.

    1986-01-01

    The noise of a counterrotation propeller at angle of attack was measured in the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel at cruise conditions. Noise increases of as much as 4 dB were measured at positive angles of attack on the tunnel side wall, which represented an airplane fuselage. These noise increases could be minimized or eliminated by operating the counterrotation propeller with the front propeller turning up-inboard. This would require oppositely rotating propellers on opposite sides of the airplane. Noise analyses at different bandwidths enabled the separate front- and rear-propeller tones, as well as the total noise, at each harmonic to be determined. A simplified noise model was explored to show how the observed circumferential noise patterns of the separate propeller tones might have occurred. The total noise pattern, which represented the sum of the front- and rear-propeller tones at a particular harmonic, showed trends that would be hard to interpret without the separate-tone results. Therefore it is important that counterrotation angle-of-attack noise data be taken in such a manner that the front- and rear-propeller tones can be separated.

  8. Evaluation of electrolytic tilt sensors for measuring model angle of attack in wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Wong, Douglas T.

    1992-01-01

    The results of a laboratory evaluation of electrolytic tilt sensors as potential candidates for measuring model attitude or angle of attack in wind tunnel tests are presented. The performance of eight electrolytic tilt sensors was compared with that of typical servo accelerometers used for angle-of-attack measurements. The areas evaluated included linearity, hysteresis, repeatability, temperature characteristics, roll-on-pitch interaction, sensitivity to lead-wire resistance, step response time, and rectification. Among the sensors being evaluated, the Spectron model RG-37 electrolytic tilt sensors have the highest overall accuracy in terms of linearity, hysteresis, repeatability, temperature sensitivity, and roll sensitivity. A comparison of the sensors with the servo accelerometers revealed that the accuracy of the RG-37 sensors was on the average about one order of magnitude worse. Even though a comparison indicates that the cost of each tilt sensor is about one-third the cost of each servo accelerometer, the sensors are considered unsuitable for angle-of-attack measurements. However, the potential exists for other applications such as wind tunnel wall-attitude measurements where the errors resulting from roll interaction, vibration, and response time are less and sensor temperature can be controlled.

  9. A Method for Integrating Thrust-Vectoring and Actuated Forebody Strakes with Conventional Aerodynamic Controls on a High-Performance Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Lallman, Frederick J.; Davidson, John B.; Murphy, Patrick C.

    1998-01-01

    A method, called pseudo controls, of integrating several airplane controls to achieve cooperative operation is presented. The method eliminates conflicting control motions, minimizes the number of feedback control gains, and reduces the complication of feedback gain schedules. The method is applied to the lateral/directional controls of a modified high-performance airplane. The airplane has a conventional set of aerodynamic controls, an experimental set of thrust-vectoring controls, and an experimental set of actuated forebody strakes. The experimental controls give the airplane additional control power for enhanced stability and maneuvering capabilities while flying over an expanded envelope, especially at high angles of attack. The flight controls are scheduled to generate independent body-axis control moments. These control moments are coordinated to produce stability-axis angular accelerations. Inertial coupling moments are compensated. Thrust-vectoring controls are engaged according to their effectiveness relative to that of the aerodynamic controls. Vane-relief logic removes steady and slowly varying commands from the thrust-vectoring controls to alleviate heating of the thrust turning devices. The actuated forebody strakes are engaged at high angles of attack. This report presents the forward-loop elements of a flight control system that positions the flight controls according to the desired stability-axis accelerations. This report does not include the generation of the required angular acceleration commands by means of pilot controls or the feedback of sensed airplane motions.

  10. Experimental investigation of the aerodynamic characteristics for a winged-cone concept

    NASA Technical Reports Server (NTRS)

    Phillips, W. Pelham; Brauckmann, Gregory J.; Micol, John R.; Woods, William C.

    1987-01-01

    Experimental longitudinal and lateral-directional aerodynamics were obtained for a generic aerodynamics were obtaiend for a generic winged-cone configuration having possible application as a transatmospheric vehicle concept. Data were obtained at Mach numbers from 0.6 to 20.0; Reynolds numbers, based on model length, between 2.5 and 5.3 million; and angles of attack from -4 to 20 deg. Results indicate a longitudinal center-of-pressure travel of about 23 percent of the fuselage length for the test Mach number range, with longitudinal instabilities noted at high-supersonic to hypersonic Mach numbers. These instabilities are coupled with directional instability at similar Mach numbers. Predictions with analytic codes, namely, the USAF DATCOM and the tangent-cone option of the Hypersonic Arbitrary Body Program, provided fair agreement with the experimental aerodynamic characteristics at low angles-of-attack.

  11. Effect of nozzle lateral spacing, engine interfairing shape, and angle of attack on the performance of a twin-jet afterbody model with cone plug nozzles

    NASA Technical Reports Server (NTRS)

    Berrier, B. L.

    1973-01-01

    Twin-jet afterbody models were investigated by using two balances to measure separately the thrust minus total axial force and the afterbody drag at Mach numbers from 0 to 1.3. Angle of attack was varied from minus 2 deg to 8.5 deg. Translating shroud cone plug nozzles were tested at dry-power and maximum-afterburning-power settings with a high-pressure air system used to provide jet total-pressure ratios up to 9.0. Two nozzle lateral spacings were studied by using afterbodies with several interfairing shapes. The close- and wide-spaced afterbodies had identical cross-sectional area distributions when similar interfairings were installed on each. The results show that the highest overall performance was obtained with the close-spaced afterbody and basic interfairings. Increasing angle of attack decreased performance for all configurations and conditions investigated.

  12. The aerodynamic forces and pressure distribution of a revolving pigeon wing

    PubMed Central

    Usherwood, James R.

    2012-01-01

    The aerodynamic forces acting on a revolving dried pigeon wing and a flat card replica were measured with a propeller rig, effectively simulating a wing in continual downstroke. Two methods were adopted: direct measurement of the reaction vertical force and torque via a forceplate, and a map of the pressures along and across the wing measured with differential pressure sensors. Wings were tested at Reynolds numbers up to 108,000, typical for slow-flying pigeons, and considerably above previous similar measurements applied to insect and hummingbird wing and wing models. The pigeon wing out-performed the flat card replica, reaching lift coefficients of 1.64 compared with 1.44. Both real and model wings achieved much higher maximum lift coefficients, and at much higher geometric angles of attack (43°), than would be expected from wings tested in a windtunnel simulating translating flight. It therefore appears that some high-lift mechanisms, possibly analogous to those of slow-flying insects, may be available for birds flapping with wings at high angles of attack. The net magnitude and orientation of aerodynamic forces acting on a revolving pigeon wing can be determined from the differential pressure maps with a moderate degree of precision. With increasing angle of attack, variability in the pressure signals suddenly increases at an angle of attack between 33° and 38°, close to the angle of highest vertical force coefficient or lift coefficient; stall appears to be delayed compared with measurements from wings in windtunnels. PMID:22736891

  13. The aerodynamic forces and pressure distribution of a revolving pigeon wing

    NASA Astrophysics Data System (ADS)

    Usherwood, James R.

    2009-05-01

    The aerodynamic forces acting on a revolving dried pigeon wing and a flat card replica were measured with a propeller rig, effectively simulating a wing in continual downstroke. Two methods were adopted: direct measurement of the reaction vertical force and torque via a forceplate, and a map of the pressures along and across the wing measured with differential pressure sensors. Wings were tested at Reynolds numbers up to 108,000, typical for slow-flying pigeons, and considerably above previous similar measurements applied to insect and hummingbird wing and wing models. The pigeon wing out-performed the flat card replica, reaching lift coefficients of 1.64 compared with 1.44. Both real and model wings achieved much higher maximum lift coefficients, and at much higher geometric angles of attack (43°), than would be expected from wings tested in a windtunnel simulating translating flight. It therefore appears that some high-lift mechanisms, possibly analogous to those of slow-flying insects, may be available for birds flapping with wings at high angles of attack. The net magnitude and orientation of aerodynamic forces acting on a revolving pigeon wing can be determined from the differential pressure maps with a moderate degree of precision. With increasing angle of attack, variability in the pressure signals suddenly increases at an angle of attack between 33° and 38°, close to the angle of highest vertical force coefficient or lift coefficient; stall appears to be delayed compared with measurements from wings in windtunnels.

  14. The aerodynamic forces and pressure distribution of a revolving pigeon wing

    NASA Astrophysics Data System (ADS)

    Usherwood, James R.

    The aerodynamic forces acting on a revolving dried pigeon wing and a flat card replica were measured with a propeller rig, effectively simulating a wing in continual downstroke. Two methods were adopted: direct measurement of the reaction vertical force and torque via a forceplate, and a map of the pressures along and across the wing measured with differential pressure sensors. Wings were tested at Reynolds numbers up to 108,000, typical for slow-flying pigeons, and considerably above previous similar measurements applied to insect and hummingbird wing and wing models. The pigeon wing out-performed the flat card replica, reaching lift coefficients of 1.64 compared with 1.44. Both real and model wings achieved much higher maximum lift coefficients, and at much higher geometric angles of attack (43°), than would be expected from wings tested in a windtunnel simulating translating flight. It therefore appears that some high-lift mechanisms, possibly analogous to those of slow-flying insects, may be available for birds flapping with wings at high angles of attack. The net magnitude and orientation of aerodynamic forces acting on a revolving pigeon wing can be determined from the differential pressure maps with a moderate degree of precision. With increasing angle of attack, variability in the pressure signals suddenly increases at an angle of attack between 33° and 38°, close to the angle of highest vertical force coefficient or lift coefficient; stall appears to be delayed compared with measurements from wings in windtunnels.

  15. The aerodynamic forces and pressure distribution of a revolving pigeon wing.

    PubMed

    Usherwood, James R

    2009-05-01

    The aerodynamic forces acting on a revolving dried pigeon wing and a flat card replica were measured with a propeller rig, effectively simulating a wing in continual downstroke. Two methods were adopted: direct measurement of the reaction vertical force and torque via a forceplate, and a map of the pressures along and across the wing measured with differential pressure sensors. Wings were tested at Reynolds numbers up to 108,000, typical for slow-flying pigeons, and considerably above previous similar measurements applied to insect and hummingbird wing and wing models. The pigeon wing out-performed the flat card replica, reaching lift coefficients of 1.64 compared with 1.44. Both real and model wings achieved much higher maximum lift coefficients, and at much higher geometric angles of attack (43°), than would be expected from wings tested in a windtunnel simulating translating flight. It therefore appears that some high-lift mechanisms, possibly analogous to those of slow-flying insects, may be available for birds flapping with wings at high angles of attack. The net magnitude and orientation of aerodynamic forces acting on a revolving pigeon wing can be determined from the differential pressure maps with a moderate degree of precision. With increasing angle of attack, variability in the pressure signals suddenly increases at an angle of attack between 33° and 38°, close to the angle of highest vertical force coefficient or lift coefficient; stall appears to be delayed compared with measurements from wings in windtunnels.

  16. The Effect of Rate of Change of Angle of Attack on the Maximum Lift Coefficient of a Pursuit Airplane

    DTIC Science & Technology

    1951-10-01

    The effect of rate of change of angle of attack on the maximum lift coefficient of a pursuit airplane equipped with a low-drag-type wing has been...lift coefficients were found to increase linearly with increasing rate of change of angle of attack per chord length of travel up to the maximum rate...indicated that the Mach and Reynolds numbers effects were of sufficient importance to produce more than a twofold variation in the increment of due to a given rate of change of angle of attack.

  17. The turbulence structure of the wake of a thin flat plate at post-stall angles of attack

    NASA Astrophysics Data System (ADS)

    Mohebi, Meraj; Wood, David H.; Martinuzzi, Robert J.

    2017-06-01

    The influence of post-stall angles of attack, α, on the turbulent flow characteristics behind a thin high aspect ratio flat plate was investigated experimentally. Time-resolved stereo particle image velocimetry was used in an open-section wind tunnel at a Reynolds number of 6600. The mean field was determined along with the wake topology, force coefficients, vortex shedding frequency, and the terms in the transport equation for the turbulent kinetic energy k. Coherent and incoherent contributions to the Reynolds stress and k-transport terms were estimated. Over the measured range of 20° ≤ α ≤ 90°, quasi-periodic vortex shedding is observed and it is shown that most of the fluctuation energy contribution in the wake arises from coherent fluctuations associated with vortex shedding. As the angle of attack is reduced from 90°, the length of the recirculation region and the drag decrease, while the shedding frequency increases monotonically. In contrast, mean lift and k are maximized at α ≈ 40°, suggesting a relationship between the bound vortex circulation and the levels of k. Structural differences in the mean strain field, wake topology, relative contributions to the k-production terms, and significant differences in the incoherent field suggest changes in the wake dynamics for α > 40° and 20° ≤ α ≤ 40°. For α > 40°, coherent contributions to the fluctuation field result in a large region close to the plate exhibiting small levels of negative mean production and generally low levels of advection, despite very high levels of production just downstream of the recirculation region.

  18. HYSHOT-2 Aerodynamics

    NASA Astrophysics Data System (ADS)

    Cain, T.; Owen, R.; Walton, C.

    2005-02-01

    The scramjet flight test Hyshot-2, flew on the 30 July 2002. The programme, led by the University of Queensland, had the primary objective of obtaining supersonic combustion data in flight for comparison with measurements made in shock tunnels. QinetiQ was one of the sponsors, and also provided aerodynamic data and trajectory predictions for the ballistic re-entry of the spinning sounding rocket. The unconventional missile geometry created by the nose-mounted asymmetric-scramjet in conjunction with the high angle of attack during re-entry makes the problem interesting. This paper presents the wind tunnel measurements and aerodynamic calculations used as input for the trajectory prediction. Indirect comparison is made with data obtained in the Hyshot-2 flight using a 6 degree-of-freedom trajectory simulation.

  19. Analytical aerodynamic model of a high alpha research vehicle wind-tunnel model

    NASA Technical Reports Server (NTRS)

    Cao, Jichang; Garrett, Frederick, Jr.; Hoffman, Eric; Stalford, Harold

    1990-01-01

    A 6 DOF analytical aerodynamic model of a high alpha research vehicle is derived. The derivation is based on wind-tunnel model data valid in the altitude-Mach flight envelope centered at 15,000 ft altitude and 0.6 Mach number with Mach range between 0.3 and 0.9. The analytical models of the aerodynamics coefficients are nonlinear functions of alpha with all control variable and other states fixed. Interpolation is required between the parameterized nonlinear functions. The lift and pitching moment coefficients have unsteady flow parts due to the time range of change of angle-of-attack (alpha dot). The analytical models are plotted and compared with their corresponding wind-tunnel data. Piloted simulated maneuvers of the wind-tunnel model are used to evaluate the analytical model. The maneuvers considered are pitch-ups, 360 degree loaded and unloaded rolls, turn reversals, split S's, and level turns. The evaluation finds that (1) the analytical model is a good representation at Mach 0.6, (2) the longitudinal part is good for the Mach range 0.3 to 0.9, and (3) the lateral part is good for Mach numbers between 0.6 and 0.9. The computer simulations show that the storage requirement of the analytical model is about one tenth that of the wind-tunnel model and it runs twice as fast.

  20. Hypersonic Laminar Viscous Flow Past Spinning Cones at Angle of Attack

    NASA Technical Reports Server (NTRS)

    Agarwal, Ramesh; Rakich, John V.

    1982-01-01

    Computational results are presented for hypersonic viscous flow past spinning sharp and blunt cones of angle of attack, obtained with a parabolic Navier-Stokes marching code. The code takes into account the asymmetries in the flowfield resulting from spinning motion and computes the asymmetric shock shape, cross-flow and streamwise shear, heat transfer, cross-flow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results.

  1. Miniature On-Board Angle of Attack Measurement System for Hypersonic Facilities

    NASA Technical Reports Server (NTRS)

    Crawford, Bradley L.; Rhode, Matthew N.

    2006-01-01

    The most prevalent method of establishing model angle of attack (AoA) in hypersonic wind tunnel facilities is using an encoder in the model support system then calculating sting/balance deflections based on balance output. This method has been shown to be less accurate than on-board methods in subsonic and transonic facilities and preliminary indications, as compared to optical methods, show large discrepancies in a hypersonic facility as well. With improvements in Micro-Electro- Mechanical Systems (MEMS) accelerometer technology more accurate onboard AoA measurement systems are now available for the small models usually found in hypersonic research facilities.

  2. Tables for Supersonic Flow of Helium Around Right Circular Cones at Zero Angle of Attack

    NASA Technical Reports Server (NTRS)

    Sims, J. L.

    1973-01-01

    The results of the calculation of supersonic flow of helium about right circular cones at zero angle of attack are presented in tabular form. The calculations were performed using the Taylor-Maccoll theory. Numerical integrations were performed using a Runge-Kutta method for second-order differential equations. Results were obtained for cone angles from 2.5 to 30 degrees in regular increments of 2.5 degrees. In all calculations the desired free-stream Mach number was obtained to five or more significant figures.

  3. Determination of angles of attack and sideslip from radar data and a roll-stabilized platform

    NASA Technical Reports Server (NTRS)

    Preisser, J. S.

    1972-01-01

    Equations for angles of attack and sideslip relative to both a rolling and nonrolling body axis system are derived for a flight vehicle for which radar and gyroscopic attitude data are available. The method is limited to application where a flat, nonrotating earth may be assumed. The gyro measures attitude relative to an inertial reference in an Euler angle sequence. In particular, a pitch, yaw, and roll sequence is used as an example in the derivation. Sample calculations based on flight data are presented to illustrate the method. Results obtained with the present gyro method are compared with another technique that uses onboard camera data.

  4. Hypersonic Laminar Viscous Flow Past Spinning Cones at Angle of Attack

    NASA Technical Reports Server (NTRS)

    Agarwal, Ramesh; Rakich, John V.

    1982-01-01

    Computational results are presented for hypersonic viscous flow past spinning sharp and blunt cones of angle of attack, obtained with a parabolic Navier-Stokes marching code. The code takes into account the asymmetries in the flowfield resulting from spinning motion and computes the asymmetric shock shape, cross-flow and streamwise shear, heat transfer, cross-flow separation, and vortex structure. The Magnus force and moments are also computed. Comparisons are made with other theoretical analyses based on boundary-layer and boundary-region equations, and an anomaly is discovered in the displacement thickness contribution to the Magnus force when compared with boundary-layer results.

  5. Effect of periodic changes of angle of attack on behavior of airfoils

    NASA Technical Reports Server (NTRS)

    Katzmayr, R

    1922-01-01

    This report presents the results of a series of experiments, which gave some quantitative results on the effect of periodic changes in the direction of the relative air flow against airfoils. The first series of experiments concerned how the angle of attack of the wing model was changed by causing the latter to oscillate about an axis parallel to the span and at right angles to the air flow. The second series embraced all the experiments in which the direction of the air flow itself was periodically changed.

  6. Numerical solutions of turbulent models for flow over a flat plate with angle of attack

    SciTech Connect

    Truncellito, N.T.; Yeh, H.; Lior, N.

    1985-03-01

    Numerical solutions of the two-dimensional boundary layer equations were developed as applied to flow over a flat plate at various angles of attack. Three methods of approach were examined. An integral solution was constructed for laminar and turbulent flow, as well as finite difference solutions for zeroth- and first-order turbulence models. The models also account for buoyancy effects. A three part mixing length model was employed in the zeroth-order model, and an additional turbulent kinetic energy equation was utilized for the first-order model. The computational method utilized Patankar-Spalding coordinates and differs from other methods in that no matching procedure is required for the inner and outer flow regions. The Falkner-Skan velocity profile is applied as an edge boundary condition while variable wall temperature conditions can be imposed. The effects of freestream velocity and angle of attack on skin friction and heat transfer were established, and the velocity and temperature fields were determined. Results of the zeroth-order solution are in excellent agreement with the Colburn equation and several other data sources. These solutions provide correlations in terms of Nusselt number and skin friction coefficient versus local Reynolds number which can be used for estimating heat transfer and wind loadings on a flat plate. Results generated are especially useful in predicting the performance of solar system designs.

  7. Rarefield-Flow Shuttle Aerodynamics Flight Model

    NASA Technical Reports Server (NTRS)

    Blanchard, Robert C.; Larman, Kevin T.; Moats, Christina D.

    1994-01-01

    A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.

  8. Identification of aerodynamic models for maneuvering aircraft

    NASA Technical Reports Server (NTRS)

    Chin, Suei; Lan, C. Edward

    1990-01-01

    Due to the requirement of increased performance and maneuverability, the flight envelope of a modern fighter is frequently extended to the high angle-of-attack regime. Vehicles maneuvering in this regime are subjected to nonlinear aerodynamic loads. The nonlinearities are due mainly to three-dimensional separated flow and concentrated vortex flow that occur at large angles of attack. Accurate prediction of these nonlinear airloads is of great importance in the analysis of a vehicle's flight motion and in the design of its flight control system. A satisfactory evaluation of the performance envelope of the aircraft may require a large number of coupled computations, one for each change in initial conditions. To avoid the disadvantage of solving the coupled flow-field equations and aircraft's motion equations, an alternate approach is to use a mathematical modeling to describe the steady and unsteady aerodynamics for the aircraft equations of motion. Aerodynamic forces and moments acting on a rapidly maneuvering aircraft are, in general, nonlinear functions of motion variables, their time rate of change, and the history of maneuvering. A numerical method was developed to analyze the nonlinear and time-dependent aerodynamic response to establish the generalized indicial function in terms of motion variables and their time rates of change.

  9. Low Speed Analysis of Mission Adaptive Flaps on a High Speed Civil Transport Configuration

    NASA Technical Reports Server (NTRS)

    Lessard, Victor R.

    1999-01-01

    Thin-layer Navier-Stokes analyses were done on a high speed civil transport configuration with mission adaptive leading-edge flaps. The flow conditions simulated were Mach = 0.22 and Reynolds number of 4.27 million for angles-of-attack ranging from 0 to 18 degrees. Two turbulence closure models were used. Analyses were done exclusively with the Baldwin-Lomax turbulence model at low angle-of-attack conditions. At high angles-of-attack where considerable flow separation and vortices occurred the Spalart-Allmaras turbulence model was also considered. The effects of flow transition were studied. Predicted aerodynamic forces, moment, and pressure are compared to experimental data obtained in the 14- by 22-Foot Subsonic Tunnel at NASA Langley. The forces and moments correlated well with experimental data in terms of trends. Drag and pitching moment were consistently underpredicted. Predicted surface pressures compared well with experiment at low angles-of-attack. Above 10 angle-of-attack the pressure comparisons were not as favorable. The two turbulent models affected the pressures on the flap considerably and neither produced correct results at the high angles-of-attack.

  10. Deep-Stall Aerodynamic Characteristics of T-Tail Aircraft

    NASA Technical Reports Server (NTRS)

    Taylor, Robert T.; Ray, Edward J.

    1965-01-01

    A wind-tunnel research program has been under-taken by the NASA to study the aerodynamic characteristics of T-tail aircraft at high angles of attack. The program was designed to show the effects on longitudinal stability and control of several configuration variables. The results to date do not allow the formulation of general design rules, but the effects of several configuration variables have been noted to have a prime influence on the post-stall characteristics. An increase in tail size, changes in the location of fuselage-mounted engine nacelles, and reduced fuselage-forebody lift were all found to have a beneficial effect on static longitudinal stability at high angles of attack.

  11. Effects of vortex flaps on the low-speed aerodynamic characteristics of an arrow wing

    NASA Technical Reports Server (NTRS)

    Yip, L. P.; Murri, D. G.

    1981-01-01

    Tests were conducted in the Langley 12-foot low-speed wind-tunnel to determine the longitudinal and lateral-directional aerodynamic effects of plain and tabbed vortex flaps on a flat-plate, highly swept arrow-wing model. Flow-visualization studies were made using a helium-bubble technique. Static forces and moments were measured over an angle-of-attack range from 0 deg to 50deg for sideslip angles of 0 deg and + or - 4 deg.

  12. Wheel squeal noise: A simplified model to simulate the effect of rolling speed and angle of attack

    NASA Astrophysics Data System (ADS)

    Liu, Xiaogang; Meehan, Paul A.

    2015-03-01

    The sound pressure level of wheel squeal has been shown to increase with angle of attack and rolling speed in both field and laboratory tests. However, the exact causes behind the manner of increase are still unknown. To investigate this, a simplified analytical vibration model in the time domain is integrated with nonlinear rolling contact theory developed for wheel squeal. This model is used to simulate the vibration velocity of a test rig wheel at different rolling speeds and angles of attack. The simulated vibration velocities correlate well in the trend with the recorded sound pressure levels of wheel squeal in laboratory tests. Lateral creepage and force at various angles of attack and rolling speeds in the rolling contact are simulated. It is found that due to the interaction of wheel vibration, lateral force and creepage, the vibration velocity amplitude of the wheel increases with angle of attack and rolling speed. The generation mechanism of wheel squeal is explained from the view of energy input per cycle of vibration. Furthermore, the reasons why the sound pressure levels of wheel squeal increase with rolling speed and angle of attack are investigated, and these phenomena are explained theoretically based on energy input and the nonlinear creep behaviour.

  13. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  14. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  15. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  16. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  17. Aerodynamic characteristics of a propeller-powered high-lift semispan wing

    NASA Technical Reports Server (NTRS)

    Gentry, Garl L., Jr.; Takallu, M. A.; Applin, Zachary T.

    1994-01-01

    A small-scale semispan high-lift wing-flap system equipped under the wing with a turboprop engine assembly was tested in the LaRC 14- by 22-Foot Subsonic Tunnel. Experimental data were obtained for various propeller rotational speeds, nacelle locations, and nacelle inclinations. To isolate the effects of the high lift system, data were obtained with and without the flaps and leading-edge device. The effects of the propeller slipstream on the overall longitudinal aerodynamic characteristics of the wing-propeller assembly were examined. Test results indicated that the lift coefficient of the wing could be increased by the propeller slipstream when the rotational speed was increased and high-lift devices were deployed. Decreasing the nacelle inclination (increased pitch down) enhanced the lift performance of the system much more than varying the vertical or horizontal location of the nacelle. Furthermore, decreasing the nacelle inclination led to higher lift curve slope values, which indicated that the powered wing could sustain higher angles of attack near maximum lift performance. Any lift augmentation was accompanied by a drag penalty due to the increased wing lift.

  18. Surface Pressure Distribution at Hypersonic Speeds for Blunt Delta Wings at Angle of Attack

    NASA Technical Reports Server (NTRS)

    Creager, Marcus O.

    1959-01-01

    Surface pressures were measured over a blunt 60 deg delta wing with extended trailing edge at a Mach number of 5.7, a free-stream Reynolds number of 20,000 per inch, and angles of attack from -10 to +10 deg. Aft of four leading-edge thicknesses the pressure distributions evidenced no appreciable three-dimensional effects and were predicted qualitatively by a method described herein for calculation of pressure distribution in two-dimensional flow. Results of tests performed elsewhere on blunt triangular wings were found to substantiate the near two-dimensionality of the flow and were used to extend the range of applicability of the method of surface pressure predictions to Mach numbers of 11.5 in air and 13.3 in helium.

  19. Study of optical techniques for the Ames unitary wind tunnels. Part 3: Angle of attack

    NASA Technical Reports Server (NTRS)

    Lee, George

    1992-01-01

    A review of optical sensors that are capable of accurate angle of attack measurements in wind tunnels was conducted. These include sensors being used or being developed at NASA Ames and Langley Research Centers, Boeing Airplane Company, McDonald Aircraft Company, Arnold Engineering Development Center, National Aerospace Laboratory of the Netherlands, National Research Council of Canada, and the Royal Aircraft Establishment of England. Some commercial sensors that may be applicable to accurate angle measurements were also reviewed. It was found that the optical sensor systems were based on interferometers, polarized light detector, linear or area photodiode cameras, position sensing photodetectors, and laser scanners. Several of the optical sensors can meet the requirements of the Ames Unitary Plan Wind Tunnel. Two of these, the Boeing interferometer and the Complere lateral effect photodiode sensors are being developed for the Ames Unitary Plan Wind Tunnel.

  20. Predicted changes in advanced turboprop noise with shaft angle of attack

    NASA Technical Reports Server (NTRS)

    Padula, S. L.; Block, P. J. W.

    1984-01-01

    Advanced turboprop blade designs and new propeller installation schemes motivated an effort to include unsteady loading effects in existing propeller noise prediction computer programs. The present work validates the prediction capability while studing the effects of shaft inclination on the radiated sound field. Classical methods of propeller performance analysis supply the time-dependent blade loading needed to calculate noise. Polar plots of the sound pressure level (SPL) of the first four harmonics and overall SPL are indicative of the change in directivity pattern as a function of propeller angle of attack. Noise predictions are compared with newly available wind tunnel data and the accuracy and applicability of the prediction method are discussed. It is concluded that unsteady blade loading caused by inclining the propeller with respect to the flow changes the directionality and the intensity of the radiated noise. These changes are well modeled by the present quasi-steady prediction method.

  1. Evaluation of electrolytic tilt sensors for wind tunnel model angle-of-attack (AOA) measurements

    NASA Technical Reports Server (NTRS)

    Wong, Douglas T.

    1991-01-01

    The results of a laboratory evaluation of three types of electrolytic tilt sensors as potential candidates for model attitude or angle of attack (AOA) measurements in wind tunnel tests are presented. Their performance was also compared with that from typical servo accelerometers used for AOA measurements. Model RG-37 electrolytic tilt sensors were found to have the highest overall accuracy among the three types. Compared with the servo accelerometer, their accuracies are about one order of magnitude worse and each of them cost about two-thirds less. Therefore, the sensors are unsuitable for AOA measurements although they are less expensive. However, the potential for other applications exists where the errors resulting from roll interaction, vibration, and response time are less, and sensor temperature can be controlled.

  2. An evaluation of several methods of determining the local angle of attack on wind turbine blades

    NASA Astrophysics Data System (ADS)

    Guntur, S.; Sørensen, N. N.

    2014-12-01

    Several methods of determining the angles of attack (AOAs) on wind turbine blades are discussed in this paper. A brief survey of the methods that have been used in the past are presented, and the advantages of each method are discussed relative to their application in the BEM theory. Data from existing as well as new full rotor CFD computations of the MEXICO rotor are used in this analysis. A more accurate estimation of the AOA is possible from 3D full rotor CFD computations, but when working with experimental data, pressure measurements and sectional forces are often the only data available. The aim of this work is to analyse the reliability of some of the simpler methods of estimating the 3D effective AOA compared some of the more rigorous CFD based methods.

  3. Lock-in of elastically mounted airfoils at a 90° angle of attack

    NASA Astrophysics Data System (ADS)

    Ehrmann, R. S.; Loftin, K. M.; Johnson, S.; White, E. B.

    2014-01-01

    Reducing vortex-induced vibration (VIV) of elastically mounted cylinders has applications to petroleum, nuclear, and civil engineering. One simple method is streamlining the cylinder into an airfoil shape. However, if flow direction changes, an elastic airfoil could experience similar oscillations with even more drag. To better understand a general airfoil's response, three elastically mounted airfoil shapes are tested at a 90° angle of attack in a 3 ft by 4 ft wind tunnel. The shapes are a NACA 0018, a sharp leading- and trailing-edge (sharp-sharp) model, and a round leading- and trailing-edge (round-round) model. Mass-damping ranges from 0.96 to 1.44. For comparison to canonical VIV research, a cylinder is also tested. Since lock-in occurs near Rec=125×103, the models are also tested with a trip strip. The NACA 0018 and sharp-sharp configuration show nearly identical responses. The cylinder and round-round airfoil have responses five to eight times larger. Thus, the existence of a single sharp edge is sufficient to greatly reduce VIV at 90° angle of attack. Whereas the cylinder and round-round maximum response amplitudes are similar, cylinder lock-in occurs over a velocity range three times larger than the round-round. The tripped cylinder and round-round models' response is attenuated by 70% compared to their respective clean configurations. Hysteresis is only observed in the circular cylinder and round-round models. Hotwire data indicates the clean cylinder has a unique vortex pattern compared to the other configurations.

  4. Liftoff and Transition Aerodynamics of the Ares I (A106) Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Capone, Francis J.; Paulson, John W., Jr.; Erickson, Gary E.

    2011-01-01

    An investigation has been conducted in the NASA Langley Research Center 14- by 22- Foot Subsonic Wind Tunnel to obtain the liftoff and transition aerodynamics of the Ares I (A106) Crew Launch Vehicle. Data were obtained in free-air at angles of attack from 10 to 90 at various roll angles and at roll angles of 0 to 360 at various angles of attack. In addition, tower effects were assessed by testing with and without a mobile launcher/tower at all wind azimuth angles and at various model heights to simulate the rise of the vehicle as it clears the tower on launch. The free-air data will be used for low speed high angle of attack flight simulation and as a bridge to the low angle of attack ascent database (0.5 < Mach < 5.0) being developed with data from the Langley Unitary Plan Wind Tunnel and Boeing Polysonic Wind Tunnel. The Ares I Database Development Team will add incremental tower effects data to the free-air data to develop the database for tower clearance.

  5. Approximate heating analysis for the windward-symmetry plane of Shuttle-like bodies at large angle of attack

    NASA Technical Reports Server (NTRS)

    Zoby, E. V.

    1981-01-01

    An engineering method has been developed for computing the windward-symmetry plane convective heat-transfer rates on Shuttle-like vehicles at large angles of attack. The engineering code includes an approximate inviscid flowfield technique, laminar and turbulent heating-rate expressions, an approximation to account for the variable-entropy effects on the surface heating and the concept of an equivalent axisymmetric body to model the windward-ray flowfields of Shuttle-like vehicles at angles of attack from 25 to 45 degrees. The engineering method is validated by comparing computed heating results with corresponding experimental data measured on Shuttle and advanced transportation models over a wide range of flow conditions and angles of attack from 25 to 40 degrees and also with results of existing prediction techniques. The comparisons are in good agreement.

  6. Mach 6 experimental and theoretical stability and performance of a cruciform missile at angles of attack up to 65 degrees

    NASA Technical Reports Server (NTRS)

    Hartman, Edward R.; Johnston, Patrick J.

    1987-01-01

    An experimental and theoretical investigation of the longitudinal and lateral-directional stability and control of an axisymmetric cruciform-finned missile has been conducted at Mach 6. The angle-of-attack range extended from 20 to 65 deg to encompass maximum lift. Longitudinal stability, performance, and trim could be accurately predicted with the fins at a fin roll angle of 0 deg but not when the fins were at a fin roll angle of 45 deg. At this roll angle, windward fin choking occurred at angles of attack above 50 deg and reduced the effectiveness of the fins and caused pitch-up.

  7. A computational/experimental study of the flow around a body of revolution at angle of attack

    NASA Technical Reports Server (NTRS)

    Zilliac, Gregory G.

    1986-01-01

    The incompressible Navier-Stokes equations are numerically solved for steady flow around an ogive-cylinder (fineness ration 4.5) at angle of attack. The three-dimensional vortical flow is investigated with emphasis on the tip and the near wake region. The implicit, finite-difference computation is performed on the CRAY X-MP computer using the method of pseudo-compressibility. Comparisons of computational results with results of a companion towing tank experiment are presented for two symmetric leeside flow cases of moderate angles of attack. The topology of the flow is discussed and conclusions are drawn concerning the growth and stability of the primary vortices.

  8. AIAA Applied Aerodynamics Conference, 10th, Palo Alto, CA, June 22-24, 1992, Technical Papers. Pts. 1 AND 2

    SciTech Connect

    Not Available

    1992-01-01

    Consideration is given to vortex physics and aerodynamics; supersonic/hypersonic aerodynamics; STOL/VSTOL/rotors; missile and reentry vehicle aerodynamics; CFD as applied to aircraft; unsteady aerodynamics; supersonic/hypersonic aerodynamics; low-speed/high-lift aerodynamics; airfoil/wing aerodynamics; measurement techniques; CFD-solvers/unstructured grid; airfoil/drag prediction; high angle-of-attack aerodynamics; and CFD grid methods. Particular attention is given to transonic-numerical investigation into high-angle-of-attack leading-edge vortex flow, prediction of rotor unsteady airloads using vortex filament theory, rapid synthesis for evaluating the missile maneuverability parameters, transonic calculations of wing/bodies with deflected control surfaces; the static and dynamic flow field development about a porous suction surface wing; the aircraft spoiler effects under wind shear; multipoint inverse design of an infinite cascade of airfoils, turbulence modeling for impinging jet flows; numerical investigation of tail buffet on the F-18 aircraft; the surface grid generation in a parameter space; and the flip flop nozzle extended to supersonic flows.

  9. Low-speed aerodynamic characteristics of a highly swept, untwisted uncambered arrow wing

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Kjelgaard, S. O.; Gentry, G. L., Jr.

    1983-01-01

    An investigation was conducted in the Langley 4- by 7-Meter Tunnel to provide a detailed study of wing pressure distributions and forces and moments acting on a highly swept arrow-wing model at low Mach numbers (0.25). A limited investigation of the effect of spoilers at several locations was also conducted. Analysis of the pressure data shows that for the configuration with undeflected leading edges, vortex separation occurs on the outboard wing panel for angles of attack on the order of only 3 deg, whereas conventional leading-edge separation occurs at a nondimensional semispan station of 0.654 for the same incidence angle. The pressure data further show that vortex separation exists at wing stations more inboard for angles of attack on the order of 7 deg and that these vortices move inboard and forward with increasing angle of attack. The force and moment data show the expected nonlinear increments in lift and pitching moment and the increased drag associated with the vortex separation. The pressure data and corresponding force and moment data confirm that deflecting the entire wing leading edge uniformly to 30 deg is effective in forestalling the onset of flow separation to angles of attack greater than 8.6 deg; however, the inboard portion of the leading edge is overdeflected. The investigation further identifies the contribution of the trailing-edge flap deflection to the leading-edge upwash fields.

  10. Rarefied-flow Shuttle aerodynamics model

    NASA Technical Reports Server (NTRS)

    Blanchard, Robert C.; Larman, Kevin T.; Moats, Christina D.

    1993-01-01

    A rarefied-flow shuttle aerodynamic model spanning the hypersonic continuum to the free molecule-flow regime was formulated. The model development has evolved from the High Resolution Accelerometer Package (HiRAP) experiment conducted on the Orbiter since 1983. The complete model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as functions of angle-of-attack, body flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle-of-attack are presented, along with flight derived rarefied-flow transition bridging formulae. Comparisons are made with data from the Operational Aerodynamic Design Data Book (OADDB), applicable wind-tunnel data, and recent flight data from STS-35 and STS-40. The flight-derived model aerodynamic force coefficient ratio is in good agreement with the wind-tunnel data and predicts the flight measured force coefficient ratios on STS-35 and STS-40. The model is not, however, in good agreement with the OADDB. But, the current OADDB does not predict the flight data force coefficient ratios of either STS-35 or STS-40 as accurately as the flight-derived model. Also, the OADDB differs with the wind-tunnel force coefficient ratio data.

  11. Effects of vortex breakdown on longitudinal and lateral-directional aerodynamics of slender wings by the suction analogy

    NASA Technical Reports Server (NTRS)

    Lan, C. E.; Hsu, C.-H.

    1982-01-01

    A semi-empirical method based on the suction analogy is developed to predict longitudinal aerodynamics and lateral-directional characteristics of slender wings at high angles of attack, including effects of vortex breakdown. The latter is based on a correlation parameter derived from the predicted leading-edge suction distribution in the attached flow. Empirical formulas, derived from a least-square analysis of data, for the vortex-breakdown angle of attack at the trailing edge, the progression rate of breakdown points and the vortex lift recovery factor in the breakdown region are given. Comparison of predicted results with data in longitudinal aerodynamics and lateral-directional characteristics for wings exhibiting strong vortex flow shows that the present method is reasonably accurate. Explanation for peculiar lateral-directional characteristics is given.

  12. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 1: Influence of airplane components for model D. [Langley spin tunnel tests

    NASA Technical Reports Server (NTRS)

    Ralston, J.

    1983-01-01

    The influence of airplane components, as well as wing location and tail length, on the rotational flow aerodynamics is discussed for a 1/6 scale general aviation airplane model. The airplane was tested in a built-up fashion (i.e., body, body-wing, body-wing-vertical, etc.) in the presence of two wing locations and two body lengths. Data were measured, using a rotary balance, over an angle-of-attack range of 8 deg to 90 deg, and for clockwise and counter-clockwise rotations covering an omega b/2V range of 0 to 0.9.

  13. Control of the Periodic Turbulent Flow over a Semicircular Airfoil with the Use of the Slot Suction of the Air from a Circular Vortex Cell at Small Angles of Attack

    NASA Astrophysics Data System (ADS)

    Isaev, S. I.; Baranov, P. A.; Sudakov, A. G.; Usachev, A. E.

    2016-11-01

    It is shown that, in the case where, into the back wall of a semicircular airfoil with an angle of attack of 5°, a vortex cell of diameter 0.2 in fractions of the airfoil chord is built in and the mean-mass rate of slot suction of the air from this cell is larger than 0.15 of the incident-flow velocity, the pattern of the turbulent flow over the airfoil is transformed, and, at an optimum suction rate of 0.75, the lift coefficient of the airfoil reaches a maximum value of the order of 1.7 at an aerodynamic efficiency of 10.

  14. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 2: Low-wing model B

    NASA Technical Reports Server (NTRS)

    Bihrle, W., Jr.; Hultberg, R. S.

    1979-01-01

    Aerodynamic characteristics obtained in a rotational flow environment utilizing a rotary balance located in the spin tunnel are presented in plotted form for a 1/6.5 scale, single engine, low wing, general aviation airplane model. The configurations tested included the basic airplane, various wing leading-edge devices, tail designs, and rudder control settings as well as airplane components. Data are presented without analysis for an angle-of-attack range of 8 deg to 90 deg and clockwise and counter-clockwise rotations covering an (omega)(b)/2V range from 0 to 0.85.

  15. Impingement of Water Droplets on an NACA 65(sub 1) -212 Airfoil at an Angle of Attack of 4 Deg

    NASA Technical Reports Server (NTRS)

    Brun, Rinaldo J.; Serafini, John S.; Moshos, George J.

    1952-01-01

    The trajectories of droplets in the air flowing past an NACA 651-212 airfoil at an angle of attack of 40 were determined. The collection efficiency, the area of droplet impingement, and the rate of droplet impingement were calculated from the trajectories and are presented herein.

  16. An NACA Vane-Type Angle-of-Attack Indicator for use at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitchell, Jesse L.; Peck, Robert F.

    1949-01-01

    A vane-type angle-of-attack indicator suitable for measurements at both subsonic and supersonic speeds has been developed by the National Advisory Committee for Aeronautics. A brief history is given of the development, and a wind-tunnel calibration of the indicator is presented, together with a discussion of the corrections to be applied to the indicated readings.

  17. Effect of Varying the Angle of Attack of the Scales on a Biomimetic Shark Skin Model on Embedded Vortex Formation

    NASA Astrophysics Data System (ADS)

    Wheelus, Jennifer; Lang, Amy; Bradshaw, Michael; Jones, Emily; Afroz, Farhana; Motta, Philip; Habegger, Maria

    2012-11-01

    The skin of fast-swimming sharks is proposed to have mechanisms to reduce drag and delay flow separation. The skin of fast-swimming and agile sharks is covered with small teeth-like denticles on the order of 0.2 mm. The shortfin mako is one of the fastest and most agile ocean predators creating the need to minimize its pressure drag by controlling flow separation. Biological studies of the shortfin mako skin have shown the passive bristling angle of their denticles to exceed 50 degrees in areas on the flank corresponding to the locations likely to experience separation first. It has been shown that for an angle of attack of 90 degrees, vortices form within these cavities and impose a partial slip condition at the surface of the cavity. This experiment focuses on smaller angles of attack for denticle bristling, closer to the range thought to be achieved on real shark skin. A 3-D bristled shark skin model with varying angle of attack, embedded below a boundary layer, was used to study the formation of cavity vortices through fluorescent dye visualization and Digital Particle Image Velocimetry (DPIV). The effect of varying angle of attack on vortex formation will be discussed.

  18. Tonal noise of a controlled-diffusion airfoil at low angle of attack and Reynolds number.

    PubMed

    Padois, Thomas; Laffay, Paul; Idier, Alexandre; Moreau, Stéphane

    2016-07-01

    The acoustic signature of a controlled-diffusion airfoil immersed in a flow is experimentally characterized. Acoustic measurements have been carried out in an anechoic open-jet-wind-tunnel for low Reynolds numbers (from 5 × 10(4) to 4.3 × 10(5)) and several angles of attack. As with the NACA0012, the acoustic spectrum is dominated by discrete tones. These tonal behaviors are divided into three different regimes. The first one is characterized by a dominant primary tone which is steady over time, surrounded by secondary peaks. The second consists of two unsteady primary tones associated with secondary peaks and the third consists of a hump dominated by several small peaks. A wavelet study allows one to identify an amplitude modulation of the acoustic signal mainly for the unsteady tonal regime. This amplitude modulation is equal to the frequency interval between two successive tones. Finally, a bispectral analysis explains the presence of tones at higher frequencies.

  19. Stability of Hypersonic Boundary Layers on a Cone at an Angle of Attack

    NASA Technical Reports Server (NTRS)

    Balakumar, P.; Owens, Lewis R.

    2010-01-01

    The stability and receptivity of a three-dimensional hypersonic boundary layer over a 7deg half-angle straight cone at an angle of attack of 6deg is numerically investigated at a freestream Mach number of 6.0 and a Reynolds number of 10.4x10(exp 6)/m. The generation and evolution of stationary crossflow vortices are also investigated by performing simulations with three-dimensional roughness elements located on the surface of the cone. The flow fields with and without the roughness elements are obtained by solving the full Navier- Stokes equations in cylindrical coordinates using a fifth-order accurate weighted essentially non-oscillatory (WENO) scheme for spatial discretization and a third-order total-variation-diminishing (TVD) Runge-Kutta scheme for temporal integration. Stability computations produced azimuthal wavenumbers in the range of m approx. 20-50 for the most amplified traveling disturbances and in the range of m approx.30-70 for the stationary disturbances. The frequency of the unstable second-mode ranges from 400 kHz to 900 kHz along the windward ray. The N-Factor computations predicted transition would occur more forward on the sides of the cone as compared to the transition fronts near the windward and the leeward rays. The simulations also show the crossflow vortices originating from the nose region propagate towards the leeward ray. No perturbations were observed toward the windward half of the cone.

  20. Crossflow transition at Mach 6 on a cone at low angles of attack

    NASA Astrophysics Data System (ADS)

    Henderson, Ryan O.

    Experiments on a sharp 7-degree cone at low angles of attack were conducted at Mach 6 to understand the stationary and traveling modes of crossflow disturbances, the interaction between them, and the development of other instabilities that can lead to transition. Using the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT), pressure and temperature measurements were collected to better describe crossflow characteristics. Noisy and quiet flow conditions were compared to understand crossflow development. Temperature Sensitive Paint (TSP) was used to measure the global surface temperatures on the model. Schmidt-Boelter (SB) gauges were used to convert the surface temperatures to heat transfer. The global heat transfer then allowed the stationary crossflow to be visualized and quantified in terms of heat flux. Integrating heat fluxes azimuthally, the amplitudes of the stationary crossflow vortices were compared. against the amplitudes of the traveling waves. PCB 132A31 and Kulite XCQ-062-15A transducers were used to measure pressure fluctuations over a broad range of frequencies. The traveling crossflow instability, the second-mode instability, and possibly the secondary-instability of the stationary crossflow mode were found at certain tunnel conditions. A grouping of Kulites was used to determine traveling wave speed and direction. Roughness elements were added to the model to excite discrete stationary vortices. The roughness elements provided a method to alter the strength of the stationary vortices. This technique allowed traveling-mode amplitudes to be compared to varying stationary-mode amplitudes.

  1. Effect of angle of attack on the flow past a harbor seal vibrissa shaped cylinder

    NASA Astrophysics Data System (ADS)

    Kim, Hyo Ju; Yoon, Hyun Sik

    2016-11-01

    The present study considered the geometric disturbance inspired by a harbor seal vibrissa of which undulated surface structures are known as a detecting device to capture the water movement induced by prey fish. In addition, this vibrissa plays an important role to suppress vortex-induced vibration, which has been reported by the previous researches. The present study aims at finding the effect of the angle of attack (AOA) on flow characteristics around the harbor seal vibrissa shaped cylinder, since the flow direction facing the harbor seal vibrissa with the elliptic shape can be changed during the harbor seal's movements and surrounding conditions. Therefore, we considered a wide range of AOA varying from 0 to 90 degree. We carried out large eddy simulation (LES) to investigate the flow around inclined vibrissa shaped cylinder for the Reynolds number (Re) of 500 based hydraulic diameter of a harbor seal vibrissa shape. For comparison, the flow over the elliptic cylinder was also simulated according to AOA at the same Re. The vortical structures of both vibrissa shaped and elliptic cylinders have been compared to identify the fundamental mechanism making the difference flow quantities. This subject is supported by Korea Ministry of Environment (MOE) as "the Chemical Accident Prevention Technology Development Project.", National Research Foundation of Korea (NRF) Grant through GCRCSOP (No.20110030013) and (NRF-2015R1D1A3A01020867).

  2. Effects of wing leading-edge radius and Reynolds number on longitudinal aerodynamic characteristics of highly swept wing-body configurations at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.

    1976-01-01

    An investigation was conducted in the Langley low turbulence pressure tunnel to determine the effects of wing leading edge radius and Reynolds number on the longitudinal aerodynamic characteristics of a series of highly swept wing-body configurations. The tests were conducted at Mach numbers below 0.30, angles of attack up to 16 deg, and Reynolds numbers per meter from 6.57 million to 43.27 million. The wings under study in this investigation had leading edge sweep angles of 61.7 deg, 64.61 deg, and 67.01 deg in combination with trailing edge sweep angles of 0 deg and 40.6 deg. The leading edge radii of each wing planform could be varied from sharp to nearly round.

  3. Effect of milling machine roughness and wing dihedral on the supersonic aerodynamic characteristics of a highly swept wing

    NASA Technical Reports Server (NTRS)

    Darden, Christine M.

    1989-01-01

    An experimental investigation was conducted to assess the effect of surface finish on the longitudinal and lateral aerodynamic characteristics of a highly-swept wing at supersonic speeds. A study of the effects of wing dihedral was also made. Included in the tests were four wing models: three models having 22.5 degrees of outboard dihedral, identical except for surface finish, and a zero-dihedral, smooth model of the same planform for reference. Of the three dihedral models, two were taken directly from the milling machine without smoothing: one having a maximum scallop height of 0.002 inches and the other a maximum scallop height of 0.005 inches. The third dihedral model was handfinished to a smooth surface. Tests were conducted in Test Section 1 of the Unitary Plan Wind Tunnel at NASA-Langley over a range of Mach numbers from 1.8 to 2.8, a range of angle of attack from -5 to 8 degrees, and at a Reynolds numbers per foot of 2 x 10(6). Selected data were also taken at a Reynolds number per foot of 6 x 10(6). Drag coefficient increases, with corresponding lift-drag ratio decreases were the primary aerodynamic effects attributed to increased surface roughness due to milling machine grooves. These drag and lift-drag ratio increments due to roughness increased as Reynolds number increased.

  4. Development of an Apparatus for Wind Tunnel Dynamic Experiments at High-alpha

    NASA Technical Reports Server (NTRS)

    Pedreiro, Nelson

    1997-01-01

    A unique experimental apparatus that allows a wind tunnel model two degrees of freedom has been designed and built. The apparatus was developed to investigate the use of new methods to augment aircraft control in the high angle of attack regime. The model support system provides a platform in which the roll-yaw coupling at high angles of attack can be studied in a controlled environment. Active cancellation of external effects is used to provide a system in which the dynamics are dominated by the aerodynamic loads acting on the wind tunnel model.

  5. Installed F/A-18 inlet flow calculations at 30 degrees angle-of-attack: A comparative study

    NASA Technical Reports Server (NTRS)

    Smith, C. Frederic; Podleski, Steve D.

    1994-01-01

    NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high-performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Full Navier-Stokes (FNS) calculations on the installed F/A-18 inlet at 30 degrees angle of attack, 0 degrees yaw, and a freestream Mach number of 0.2 have been obtained in this study using an algebraic turbulence model with two grids (original and revised). Results obtained with the original grid were used to determine where further grid refinements and additional geometry were needed. In order to account properly for the external effects, the forebody, leading edge extension (LEX), ramp, and wing were included with inlet geometry. In the original grid, the diverter, LEX slot, and leading edge flap were not included due to insufficient geometry definition, but were included in a revised grid. In addition, a thin-layer Navier-Stokes (TLNS) code is used with the revised grid and the numerical results are compared to those obtained with the FNS code. The TLNS code was used to evaluate the effects on the solution using a code with more recent CFD developments such as upwinding with TVD schemes versus central differencing with artificial dissipation. The calculations are compared to a limited amount of available experimental data. The predicted forebody/fuselage surface static pressures compared well with data of all solutions. The predicted trajectory of the vortex generated under the LEX was

  6. Determination of slender body aerodynamics using discrete vortex methods

    NASA Astrophysics Data System (ADS)

    Gebert, G. A.

    1994-03-01

    Current aerodynamic interest has turned to the study of supermaneuverable fighters and weapon performance when launched in extreme flight conditions. The evaluation of design missile performance requires multiple runs of six degree-of-freedom (6-DOF) simulations, analyzing the missile behavior for a variety of launch and flight conditions. Before wind-tunnel tests, it is necessary to produce the aerodynamic loading of candidate missiles for 6-DOF analyses. Since semi-empirical formulas fail in regions of nonlinear aerodynamics, and solutions to the full Navier-Stokes equations are too costly and time consuming, an alternative method of discrete vortex analysis is re-examined. The present theory examines the three-dimensional nature of the shed vorticity and generalizes previous discrete vortex analyses. Consequently, the results demonstrate relative user independence in determining all slender-body loading at angles of attack from 0 to 70 deg. The rapid calculations of the discrete vortex method makes it a prime candidate for the determinations of high angle-of-attack aerodynamic databases.

  7. Linearized Poststall Aerodynamic and Control Law Models of the X-31A Aircraft and Comparison with Flight Data

    NASA Technical Reports Server (NTRS)

    Stoliker, Patrick C.; Bosworth, John T.; Georgie, Jennifer

    1997-01-01

    The X-31A aircraft has a unique configuration that uses thrust-vector vanes and aerodynamic control effectors to provide an operating envelope to a maximum 70 deg angle of attack, an inherently nonlinear portion of the flight envelope. This report presents linearized versions of the X-31A longitudinal and lateral-directional control systems, with aerodynamic models sufficient to evaluate characteristics in the poststall envelope at 30 deg, 45 deg, and 60 deg angle of attack. The models are presented with detail sufficient to allow the reader to reproduce the linear results or perform independent control studies. Comparisons between the responses of the linear models and flight data are presented in the time and frequency domains to demonstrate the strengths and weaknesses of the ability to predict high-angle-of-attack flight dynamics using linear models. The X-31A six-degree-of-freedom simulation contains a program that calculates linear perturbation models throughout the X-31A flight envelope. The models include aerodynamics and flight control system dynamics that are used for stability, controllability, and handling qualities analysis. The models presented in this report demonstrate the ability to provide reasonable linear representations in the poststall flight regime.

  8. Computational Aerodynamic Analysis of a Micro-CT Based Bio-Realistic Fruit Fly Wing

    PubMed Central

    Brandt, Joshua; Doig, Graham; Tsafnat, Naomi

    2015-01-01

    The aerodynamic features of a bio-realistic 3D fruit fly wing in steady state (snapshot) flight conditions were analyzed numerically. The wing geometry was created from high resolution micro-computed tomography (micro-CT) of the fruit fly Drosophila virilis. Computational fluid dynamics (CFD) analyses of the wing were conducted at ultra-low Reynolds numbers ranging from 71 to 200, and at angles of attack ranging from -10° to +30°. It was found that in the 3D bio-realistc model, the corrugations of the wing created localized circulation regions in the flow field, most notably at higher angles of attack near the wing tip. Analyses of a simplified flat wing geometry showed higher lift to drag performance values for any given angle of attack at these Reynolds numbers, though very similar performance is noted at -10°. Results have indicated that the simplified flat wing can successfully be used to approximate high-level properties such as aerodynamic coefficients and overall performance trends as well as large flow-field structures. However, local pressure peaks and near-wing flow features induced by the corrugations are unable to be replicated by the simple wing. We therefore recommend that accurate 3D bio-realistic geometries be used when modelling insect wings where such information is useful. PMID:25954946

  9. Computational Aerodynamic Analysis of a Micro-CT Based Bio-Realistic Fruit Fly Wing.

    PubMed

    Brandt, Joshua; Doig, Graham; Tsafnat, Naomi

    2015-01-01

    The aerodynamic features of a bio-realistic 3D fruit fly wing in steady state (snapshot) flight conditions were analyzed numerically. The wing geometry was created from high resolution micro-computed tomography (micro-CT) of the fruit fly Drosophila virilis. Computational fluid dynamics (CFD) analyses of the wing were conducted at ultra-low Reynolds numbers ranging from 71 to 200, and at angles of attack ranging from -10° to +30°. It was found that in the 3D bio-realistic model, the corrugations of the wing created localized circulation regions in the flow field, most notably at higher angles of attack near the wing tip. Analyses of a simplified flat wing geometry showed higher lift to drag performance values for any given angle of attack at these Reynolds numbers, though very similar performance is noted at -10°. Results have indicated that the simplified flat wing can successfully be used to approximate high-level properties such as aerodynamic coefficients and overall performance trends as well as large flow-field structures. However, local pressure peaks and near-wing flow features induced by the corrugations are unable to be replicated by the simple wing. We therefore recommend that accurate 3D bio-realistic geometries be used when modelling insect wings where such information is useful.

  10. Effect of twist and camber on the low-speed aerodynamic characteristics of a powered close-coupled wing-canard configuration

    NASA Technical Reports Server (NTRS)

    Paulson, J. W., Jr.; Thomas, J. L.

    1978-01-01

    A series of wind-tunnel tests were conducted in a V/STOL tunnel to determine the low-speed longitudinal aerodynamic characteristics of a powered close-coupled wing/canard fighter configuration. The data was obtained for a high angle-of-attack maneuvering configuration and a takeoff and landing configuration. The data presented in tabulated form are intended for reference purposes.

  11. Aerodynamics of high frequency flapping wings

    NASA Astrophysics Data System (ADS)

    Hu, Zheng; Roll, Jesse; Cheng, Bo; Deng, Xinyan

    2010-11-01

    We investigated the aerodynamic performance of high frequency flapping wings using a 2.5 gram robotic insect mechanism developed in our lab. The mechanism flaps up to 65Hz with a pair of man-made wing mounted with 10cm wingtip-to-wingtip span. The mean aerodynamic lift force was measured by a lever platform, and the flow velocity and vorticity were measured using a stereo DPIV system in the frontal, parasagittal, and horizontal planes. Both near field (leading edge vortex) and far field flow (induced flow) were measured with instantaneous and phase-averaged results. Systematic experiments were performed on the man-made wings, cicada and hawk moth wings due to their similar size, frequency and Reynolds number. For insect wings, we used both dry and freshly-cut wings. The aerodynamic force increase with flapping frequency and the man-made wing generates more than 4 grams of lift at 35Hz with 3 volt input. Here we present the experimental results and the major differences in their aerodynamic performances.

  12. Modeling of turbulent separated flows for aerodynamic applications

    NASA Technical Reports Server (NTRS)

    Marvin, J. G.

    1983-01-01

    Steady, high speed, compressible separated flows modeled through numerical simulations resulting from solutions of the mass-averaged Navier-Stokes equations are reviewed. Emphasis is placed on benchmark flows that represent simplified (but realistic) aerodynamic phenomena. These include impinging shock waves, compression corners, glancing shock waves, trailing edge regions, and supersonic high angle of attack flows. A critical assessment of modeling capabilities is provided by comparing the numerical simulations with experiment. The importance of combining experiment, numerical algorithm, grid, and turbulence model to effectively develop this potentially powerful simulation technique is stressed.

  13. PIV-based study of the gliding osprey aerodynamics in a wind tunnel

    NASA Astrophysics Data System (ADS)

    Gurka, Roi; Liberzon, Alex; Kopp, Gregory; Kirchhefer, Adam; Weihs, Daniel

    2009-11-01

    The hunting flight of an osprey consists of periods where the bird glides while foraging for prey. High quality measurements of aerodynamics in this flight mode are needed in order to estimate the daily energy expenditure of the bird accurately. An experimental study of an osprey model in a wind tunnel (BLWTL, UWO) was performed in order to characterize the aerodynamic forces using particle image velocimetry (PIV). The model was a stuffed osprey with mechanical joints allowing control of the the wing (angle of attack, tilt) and tail orientation. Two-dimensional velocity realizations in the streamwise-normal plane were obtained simultaneously in the two fields of view: above the wing and in the wake of the wing. Mean and turbulent flow characteristics are presented as function of angle of attack based on measurements taken at 4 different angles of attack at three different locations over the wingspan. The main outcome is the accurate estimate of the drag from the measurements of momentum thickness in the turbulent boundary layer of the osprey wing. Moreover, the gradient of the momentum thickness method was applied to identify the separation point in the boundary layer. This estimate has been compared to the total drag calculated from measurements in the wake of the wing and with a theoretical prediction.

  14. 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Baize, Daniel G. (Editor)

    1999-01-01

    The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.

  15. The development of a capability for aerodynamic testing of large-scale wing sections in a simulated natural rain environment

    NASA Technical Reports Server (NTRS)

    Bezos, Gaudy M.; Cambell, Bryan A.; Melson, W. Edward

    1989-01-01

    A research technique to obtain large-scale aerodynamic data in a simulated natural rain environment has been developed. A 10-ft chord NACA 64-210 wing section wing section equipped with leading-edge and trailing-edge high-lift devices was tested as part of a program to determine the effect of highly-concentrated, short-duration rainfall on airplane performance. Preliminary dry aerodynamic data are presented for the high-lift configuration at a velocity of 100 knots and an angle of attack of 18 deg. Also, data are presented on rainfield uniformity and rainfall concentration intensity levels obtained during the calibration of the rain simulation system.

  16. Toward Improved Predictions of Slender Airframe Aerodynamics Using the F-16XL Aircraft

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Rizzi, Arthur; Davis, M. Bruce

    2016-01-01

    A coordinated project has been underway to improve computational fluid dynamics predictions of slender airframe aerodynamics. The work is focused on two flow conditions and leverages a unique flight data set obtained with an F-16XL aircraft. These conditions, a low-speed high angle-of-attack case and a transonic low angle-of-attack case, were selected from a prior prediction campaign wherein the computational fluid dynamics failed to provide acceptable results. In this paper, the background, objectives, and approach to the current project are presented. The work embodies predictions from multiple numerical formulations that are contributed from multiple organizations, and the context of this campaign to other multicode, multi-organizational efforts is included. The relevance of this body of work toward future supersonic commercial transport concepts is also briefly addressed.

  17. Low-speed aerodynamic test of an axisymmetric supersonic inlet with variable cowl slot

    NASA Technical Reports Server (NTRS)

    Powell, A. G.; Welge, H. R.; Trefny, C. J.

    1985-01-01

    The experimental low-speed aerodynamic characteristics of an axisymmetric mixed-compression supersonic inlet with variable cowl slot are described. The model consisted of the NASA P-inlet centerbody and redesigned cowl with variable cowl slot powered by the JT8D single-stage fan simulator and driven by an air turbine. The model was tested in the NASA Lewis Research Center 9- by 15-foot low-speed tunnel at Mach numbers of 0, 0.1, and 0.2 over a range of flows, cowl slot openings, centerbody positions, and angles of attack. The variable cowl slot was effective in minimizing lip separation at high velocity ratios, showed good steady-state and dynamic distortion characteristics, and had good angle-of-attack tolerance.

  18. Effects of wing deformation on aerodynamic performance of a revolving insect wing

    NASA Astrophysics Data System (ADS)

    Noda, Ryusuke; Nakata, Toshiyuki; Liu, Hao

    2014-12-01

    Flexible wings of insects and bio-inspired micro air vehicles generally deform remarkably during flapping flight owing to aerodynamic and inertial forces, which is of highly nonlinear fluid-structure interaction (FSI) problems. To elucidate the novel mechanisms associated with flexible wing aerodynamics in the low Reynolds number regime, we have built up a FSI model of a hawkmoth wing undergoing revolving and made an investigation on the effects of flexible wing deformation on aerodynamic performance of the revolving wing model. To take into account the characteristics of flapping wing kinematics we designed a kinematic model for the revolving wing in two-fold: acceleration and steady rotation, which are based on hovering wing kinematics of hawkmoth, Manduca sexta. Our results show that both aerodynamic and inertial forces demonstrate a pronounced increase during acceleration phase, which results in a significant wing deformation. While the aerodynamic force turns to reduce after the wing acceleration terminates due to the burst and detachment of leading-edge vortices (LEVs), the dynamic wing deformation seem to delay the burst of LEVs and hence to augment the aerodynamic force during and even after the acceleration. During the phase of steady rotation, the flexible wing model generates more vertical force at higher angles of attack (40°-60°) but less horizontal force than those of a rigid wing model. This is because the wing twist in spanwise owing to aerodynamic forces results in a reduction in the effective angle of attack at wing tip, which leads to enhancing the aerodynamics performance by increasing the vertical force while reducing the horizontal force. Moreover, our results point out the importance of the fluid-structure interaction in evaluating flexible wing aerodynamics: the wing deformation does play a significant role in enhancing the aerodynamic performances but works differently during acceleration and steady rotation, which is mainly induced by

  19. Assessing Uncertainties in Boundary Layer Transition Predictions for HIFiRE-1 at Non-zero Angles of Attack

    NASA Technical Reports Server (NTRS)

    Marek, Lindsay C.

    2011-01-01

    Boundary layer stability was analyzed for the HIFiRE-1 flight vehicle geometry for ground tests conducted at the CUBRC LENS I hypersonic shock test facility and the Langley Research Center (LaRC) 20- inch Mach 6 Tunnel. Boundary layer stability results were compared to transition onset location obtained from discrete heat transfer measurements from thin film gauges during the CUBRC test and spatially continuous heat transfer measurements from thermal phosphor paint data during the LaRC test. The focus of this analysis was on conditions at non-zero angles of attack as stability analysis has already been performed at zero degrees angle of attack. Also, the transition onset data obtained during flight testing was at nonzero angles of attack, so this analysis could be expanded in the future to include the results of the flight test data. Stability analysis was performed using the 2D parabolized stability software suite STABL (Stability and Transition Analysis for Hypersonic Boundary Layers) developed at the University of Minnesota and the mean flow solutions were computed using the DPLR finite volume Navier-Stokes computational fluid dynamics (CFD) solver. A center line slice of the 3D mean flow solution was used for the stability analysis to incorporate the angle of attack effects while still taking advantage of the 2D STABL software suite. The N-factors at transition onset and the value of Re(sub theta)/M(sub e), commonly used to predict boundary layer transition onset, were compared for all conditions analyzed. Ground test data was analyzed at Mach 7.2 and Mach 6.0 and angles of attack of 1deg, 3deg and 5deg. At these conditions, the flow was found to be second mode dominant for the HIFiRE-1 slender cone geometry. On the leeward side of the vehicle, a strong trend of transition onset location with angle of attack was observed as the boundary layer on the leeward side of the vehicle developed inflection points at streamwise positions on the vehicle that correlated to

  20. Lee surface flow phenomena over space shuttle at large angles of attack at M sub infinity equal 6

    NASA Technical Reports Server (NTRS)

    Zakkay, V.; Miyazawa, M.; Wang, C. R.

    1974-01-01

    Surface pressure and heat transfer, flow separation, flow field, and oil flow patterns on the leeward side of a space shuttle orbiter model are investigated at a free stream Mach number of 6. The free stream Reynolds numbers are between 1.64 times 10 to the 7th power and 1.31 times 10 to the 8th power per meter, and the angle of attack is varied between 0 deg and 40 deg for the present experiments. The stagnation temperatures for the tests are approximately 500 K and the wall temperature is maintained at 290 K. Existing numerical methods of three-dimensional inviscid supersonic flow theory and compressible boundary layer theory are used to predict the present experimental measurements. Results of the present tests indicate two distinct types of flow separation and surface peak heating depending on the angle of attack.

  1. Computation of the inviscid supersonic flow about cones at large angles of attack by a floating discontinuity approach

    NASA Technical Reports Server (NTRS)

    Daywitt, J.; Kutler, P.; Anderson, D.

    1977-01-01

    The technique of floating shock fitting is adapted to the computation of the inviscid flowfield about circular cones in a supersonic free stream at angles of attack that exceed the cone half-angle. The resulting equations are applicable over the complete range of free-stream Mach numbers, angles of attack and cone half-angles for which the bow shock is attached. A finite difference algorithm is used to obtain the solution by an unsteady relaxation approach. The bow shock, embedded cross-flow shock, and vortical singularity in the leeward symmetry plane are treated as floating discontinuities in a fixed computational mesh. Where possible, the flowfield is partitioned into windward, shoulder, and leeward regions with each region computed separately to achieve maximum computational efficiency. An alternative shock fitting technique which treats the bow shock as a computational boundary is developed and compared with the floating-fitting approach. Several surface boundary condition schemes are also analyzed.

  2. Unsteady Thick Airfoil Aerodynamics: Experiments, Computation, and Theory

    NASA Technical Reports Server (NTRS)

    Strangfeld, C.; Rumsey, C. L.; Mueller-Vahl, H.; Greenblatt, D.; Nayeri, C. N.; Paschereit, C. O.

    2015-01-01

    An experimental, computational and theoretical investigation was carried out to study the aerodynamic loads acting on a relatively thick NACA 0018 airfoil when subjected to pitching and surging, individually and synchronously. Both pre-stall and post-stall angles of attack were considered. Experiments were carried out in a dedicated unsteady wind tunnel, with large surge amplitudes, and airfoil loads were estimated by means of unsteady surface mounted pressure measurements. Theoretical predictions were based on Theodorsen's and Isaacs' results as well as on the relatively recent generalizations of van der Wall. Both two- and three-dimensional computations were performed on structured grids employing unsteady Reynolds-averaged Navier-Stokes (URANS). For pure surging at pre-stall angles of attack, the correspondence between experiments and theory was satisfactory; this served as a validation of Isaacs theory. Discrepancies were traced to dynamic trailing-edge separation, even at low angles of attack. Excellent correspondence was found between experiments and theory for airfoil pitching as well as combined pitching and surging; the latter appears to be the first clear validation of van der Wall's theoretical results. Although qualitatively similar to experiment at low angles of attack, two-dimensional URANS computations yielded notable errors in the unsteady load effects of pitching, surging and their synchronous combination. The main reason is believed to be that the URANS equations do not resolve wake vorticity (explicitly modeled in the theory) or the resulting rolled-up un- steady flow structures because high values of eddy viscosity tend to \\smear" the wake. At post-stall angles, three-dimensional computations illustrated the importance of modeling the tunnel side walls.

  3. A Synthesis of Hybrid RANS/LES CFD Results for F-16XL Aircraft Aerodynamics

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Park, Michael A.; Hitzel, Stephan M.; Jirasek, Adam; Lofthouse, Andrew J.; Morton, Scott A.; McDaniel, David R.; Rizzi, Arthur M.

    2015-01-01

    A synthesis is presented of recent numerical predictions for the F-16XL aircraft flow fields and aerodynamics. The computational results were all performed with hybrid RANS/LES formulations, with an emphasis on unsteady flows and subsequent aerodynamics, and results from five computational methods are included. The work was focused on one particular low-speed, high angle-of-attack flight test condition, and comparisons against flight-test data are included. This work represents the third coordinated effort using the F-16XL aircraft, and a unique flight-test data set, to advance our knowledge of slender airframe aerodynamics as well as our capability for predicting these aerodynamics with advanced CFD formulations. The prior efforts were identified as Cranked Arrow Wing Aerodynamics Project International, with the acronyms CAWAPI and CAWAPI-2. All information in this paper is in the public domain.

  4. Method of Estimating the Incompressible-flow Pressure Distribution of Compressor Blade Sections at Design Angle of Attack

    NASA Technical Reports Server (NTRS)

    Erwin, John R; Yacobi, Laura A

    1953-01-01

    A method was devised for estimating the incompressible-flow pressure distribution over compressor blade sections at design angle of attack. The theoretical incremental velocities due to camber and thickness of the section as an isolated airfoil are assumed proportional to the average passage velocity and are modified by empirically determined interference factors. Comparisons were made between estimated and test pressure distributions of NACA 65-series sections for typical conditions. Good agreement was obtained.

  5. The method of characteristics for the determination of supersonic flow over bodies of revolution at small angles of attack

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1951-01-01

    The method of characteristics has been applied for the determination of the supersonic-flow properties around bodies of revolution at a small angle of attack. The system developed considers the effect of the variation of entropy due to the curved shock and determines a flow that exactly satisfies the boundary conditions in the limits of the simplifications assumed. Two practical methods for numerical calculations are given. (author)

  6. Design of high performance multivariable control systems for supermaneuverable aircraft at high angle of attack

    NASA Technical Reports Server (NTRS)

    Valavani, Lena

    1995-01-01

    The main motivation for the work under the present grant was to use nonlinear feedback linearization methods to further enhance performance capabilities of the aircraft, and robustify its response throughout its operating envelope. The idea was to use these methods in lieu of standard Taylor series linearization, in order to obtain a well behaved linearized plant, in its entire operational regime. Thus, feedback linearization was going to constitute an 'inner loop', which would then define a 'design plant model' to be compensated for robustness and guaranteed performance in an 'outer loop' application of modern linear control methods. The motivation for this was twofold; first, earlier work had shown that by appropriately conditioning the plant through conventional, simple feedback in an 'inner loop', the resulting overall compensated plant design enjoyed considerable enhancement of performance robustness in the presence of parametric uncertainty. Second, the nonlinear techniques did not have any proven robustness properties in the presence of unstructured uncertainty; a definition of robustness (and performance) is very difficult to achieve outside the frequency domain; to date, none is available for the purposes of control system design. Thus, by proper design of the outer loop, such properties could still be 'injected' in the overall system.

  7. Two-Dimensional High-Lift Aerodynamic Optimization Using Neural Networks

    NASA Technical Reports Server (NTRS)

    Greenman, Roxana M.

    1998-01-01

    The high-lift performance of a multi-element airfoil was optimized by using neural-net predictions that were trained using a computational data set. The numerical data was generated using a two-dimensional, incompressible, Navier-Stokes algorithm with the Spalart-Allmaras turbulence model. Because it is difficult to predict maximum lift for high-lift systems, an empirically-based maximum lift criteria was used in this study to determine both the maximum lift and the angle at which it occurs. The 'pressure difference rule,' which states that the maximum lift condition corresponds to a certain pressure difference between the peak suction pressure and the pressure at the trailing edge of the element, was applied and verified with experimental observations for this configuration. Multiple input, single output networks were trained using the NASA Ames variation of the Levenberg-Marquardt algorithm for each of the aerodynamic coefficients (lift, drag and moment). The artificial neural networks were integrated with a gradient-based optimizer. Using independent numerical simulations and experimental data for this high-lift configuration, it was shown that this design process successfully optimized flap deflection, gap, overlap, and angle of attack to maximize lift. Once the neural nets were trained and integrated with the optimizer, minimal additional computer resources were required to perform optimization runs with different initial conditions and parameters. Applying the neural networks within the high-lift rigging optimization process reduced the amount of computational time and resources by 44% compared with traditional gradient-based optimization procedures for multiple optimization runs.

  8. Nonlinear prediction of the aerodynamic loads on lifting surfaces

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.; Mook, D. T.; Nayfeh, A. H.

    1974-01-01

    A numerical procedure is used to predict the nonlinear aerodynamic characteristics of lifting surfaces of low aspect ratio at high angles of attack for low subsonic Mach numbers. The procedure utilizes a vortex-lattice method and accounts for separation at sharp tips and leading edges. The shapes of the wakes emanating from the edges are predicted, and hence the nonlinear characteristics are calculated. Parallelogram and delta wings are presented as numerical examples. The numerical results are in good agreement with the experimental data.

  9. 1.5 μm lidar anemometer for true air speed, angle of sideslip, and angle of attack measurements on-board Piaggio P180 aircraft

    NASA Astrophysics Data System (ADS)

    Augere, B.; Besson, B.; Fleury, D.; Goular, D.; Planchat, C.; Valla, M.

    2016-05-01

    Lidar (light detection and ranging) is a well-established measurement method for the prediction of atmospheric motions through velocity measurements. Recent advances in 1.5 μm Lidars show that the technology is mature, offers great ease of use, and is reliable and compact. A 1.5 μm airborne Lidar appears to be a good candidate for airborne in-flight measurement systems. It allows measurements remotely, outside aircraft aerodynamic disturbance, and absolute air speed (no need for calibration) with great precision in all aircraft flight domains. In the framework of the EU AIM2 project, the ONERA task has consisted of developing and testing a 1.5 μm anemometer sensor for in-flight airspeed measurements. The objective of this work is to demonstrate that the 1.5 μm Lidar sensor can increase the quality of the data acquisition procedure for aircraft flight test certification. This article presents the 1.5 μm anemometer sensor dedicated to in-flight airspeed measurements and describes the flight tests performed successfully on-board the Piaggio P180 aircraft. Lidar air data have been graphically compared to the air data provided by the aircraft flight test instrumentation (FTI) in the reference frame of the Lidar sensor head. Very good agreement of true air speed (TAS) by a fraction of ms-1, angle of sideslip (AOS), and angle of attack (AOA) by a fraction of degree were observed.

  10. 1999 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hahne, David E. (Editor)

    1999-01-01

    NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 1/Part 1 publication covers configuration aerodynamics.

  11. High speed civil transport aerodynamic optimization

    NASA Technical Reports Server (NTRS)

    Ryan, James S.

    1994-01-01

    This is a report of work in support of the Computational Aerosciences (CAS) element of the Federal HPCC program. Specifically, CFD and aerodynamic optimization are being performed on parallel computers. The long-range goal of this work is to facilitate teraflops-rate multidisciplinary optimization of aerospace vehicles. This year's work is targeted for application to the High Speed Civil Transport (HSCT), one of four CAS grand challenges identified in the HPCC FY 1995 Blue Book. This vehicle is to be a passenger aircraft, with the promise of cutting overseas flight time by more than half. To meet fuel economy, operational costs, environmental impact, noise production, and range requirements, improved design tools are required, and these tools must eventually integrate optimization, external aerodynamics, propulsion, structures, heat transfer, controls, and perhaps other disciplines. The fundamental goal of this project is to contribute to improved design tools for U.S. industry, and thus to the nation's economic competitiveness.

  12. Advanced Aerodynamic Design of Passive Porosity Control Effectors

    NASA Technical Reports Server (NTRS)

    Hunter, Craig A.; Viken, Sally A.; Wood, Richard M.; Bauer, Steven X. S.

    2001-01-01

    This paper describes aerodynamic design work aimed at developing a passive porosity control effector system for a generic tailless fighter aircraft. As part of this work, a computational design tool was developed and used to layout passive porosity effector systems for longitudinal and lateral-directional control at a low-speed, high angle of attack condition. Aerodynamic analysis was conducted using the NASA Langley computational fluid dynamics code USM3D, in conjunction with a newly formulated surface boundary condition for passive porosity. Results indicate that passive porosity effectors can provide maneuver control increments that equal and exceed those of conventional aerodynamic effectors for low-speed, high-alpha flight, with control levels that are a linear function of porous area. This work demonstrates the tremendous potential of passive porosity to yield simple control effector systems that have no external moving parts and will preserve an aircraft's fixed outer mold line.

  13. HSCT high lift system aerodynamic requirements

    NASA Technical Reports Server (NTRS)

    Paulson, John A.

    1992-01-01

    The viewgraphs and discussion of high lift system aerodynamic requirements are provided. Low speed aerodynamics has been identified as critical to the successful development of a High Speed Civil Transport (HSCT). The airplane must takeoff and land at a sufficient number of existing or projected airports to be economically viable. At the same time, community noise must be acceptable. Improvements in cruise drag, engine fuel consumption, and structural weight tend to decrease the wing size and thrust required of engines. Decreasing wing size increases the requirements for effective and efficient low speed characteristics. Current design concepts have already been compromised away from better cruise wings for low speed performance. Flap systems have been added to achieve better lift-to-drag ratios for climb and approach and for lower pitch attitudes for liftoff and touchdown. Research to achieve improvements in low speed aerodynamics needs to be focused on areas most likely to have the largest effect on the wing and engine sizing process. It would be desirable to provide enough lift to avoid sizing the airplane for field performance and to still meet the noise requirements. The airworthiness standards developed in 1971 will be the basis for performance requirements for an airplane that will not be critical to the airplane wing and engine size. The lift and drag levels that were required to meet the performance requirements of tentative airworthiness standards established in 1971 and that were important to community noise are identified. Research to improve the low speed aerodynamic characteristics of the HSCT needs to be focused in the areas of performance deficiency and where noise can be reduced. Otherwise, the wing planform, engine cycle, or other parameters for a superior cruising airplane would have to be changed.

  14. Transonic-supersonic high Reynolds number stability and control characteristics of a 0.015-scale (remotely controlled elevon) model 44-0 of the space shuttle orbiter tested in the VSD high speed wind tunnel (LA67)

    NASA Technical Reports Server (NTRS)

    1976-01-01

    A detailed aerodynamic data base which can be used to substantiate the aerodynamic design data book for the current shuttle orbiter configuration was generated. Special attention was directed to definition of non-linear aerodynamic characteristics by taking data at small increments in the angle of attack, angle of sideslip, Mach number, and elevon position. Six-component aerodynamic force and moment and elevon position data were recorded over an angle-of-attack range from -2 deg to as high as 32 deg at angles of sideslip of 0 deg, 1 deg, and +2 deg. The test Mach numbers were 0.60, 0.80, 0.90, 1.2, 1.5, 2.0, 3.0, and 4.6. The effects of Reynolds number were investigated and covered a range from 5.0 to 16.0 million per foot.

  15. Flow-field surveys on the windward side of the NASA 040A space shuttle orbiter at 31 deg angle of attack and Mach 20 in helium

    NASA Technical Reports Server (NTRS)

    Ashby, G. C., Jr.; Helms, V. T., III

    1977-01-01

    Pitot pressure and flow angle distributions in the windward flow field of the NASA 040A space shuttle orbiter configuration and surface pressures were measured, at a Mach number of 20 and an angle of attack of 31 deg. The free stream Reynolds number, based on model length, was 5.39 x 10 to the 6th power. Results show that cores of high pitot pressure, which are related to the body-shock-wing-shock intersections, occur on the windward plane of symmetry in the vicinity of the wing-body junction and near midspan on the wing. Theoretical estimates of the flow field pitot pressures show that conical flow values for the windward plane of symmetry surface are representative of the average level over the entire lower surface.

  16. Preliminary Wind-Tunnel Investigation of the Performance of Republic F-105 Wing-Root Inlet Configurations at Various Angles of Attack and a Mach Number of 2.01. [Coord. No. AF-163

    NASA Technical Reports Server (NTRS)

    Kouyoumjian, W. L.

    1957-01-01

    A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.

  17. Aerodynamic investigations of a disc-wing

    NASA Astrophysics Data System (ADS)

    Dumitrache, Alexandru; Frunzulica, Florin; Grigorescu, Sorin

    2017-01-01

    The purpose of this paper is to evaluate the aerodynamic characteristics of a wing-disc, for a civil application in the fire-fighting system. The aerodynamic analysis is performed using a CFD code, named ANSYS Fluent, in the flow speed range up to 25 m/s, at lower and higher angle of attack. The simulation is three-dimensional, using URANS completed by a SST turbulence model. The results are used to examine the flow around the disc with increasing angle of attack and the structure of the wake.

  18. 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Baize, Daniel G. (Editor)

    1999-01-01

    The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in area of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodyamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.

  19. Asymmetric Uncertainty Expression for High Gradient Aerodynamics

    NASA Technical Reports Server (NTRS)

    Pinier, Jeremy T

    2012-01-01

    When the physics of the flow around an aircraft changes very abruptly either in time or space (e.g., flow separation/reattachment, boundary layer transition, unsteadiness, shocks, etc), the measurements that are performed in a simulated environment like a wind tunnel test or a computational simulation will most likely incorrectly predict the exact location of where (or when) the change in physics happens. There are many reasons for this, includ- ing the error introduced by simulating a real system at a smaller scale and at non-ideal conditions, or the error due to turbulence models in a computational simulation. The un- certainty analysis principles that have been developed and are being implemented today do not fully account for uncertainty in the knowledge of the location of abrupt physics changes or sharp gradients, leading to a potentially underestimated uncertainty in those areas. To address this problem, a new asymmetric aerodynamic uncertainty expression containing an extra term to account for a phase-uncertainty, the magnitude of which is emphasized in the high-gradient aerodynamic regions is proposed in this paper. Additionally, based on previous work, a method for dispersing aerodynamic data within asymmetric uncer- tainty bounds in a more realistic way has been developed for use within Monte Carlo-type analyses.

  20. Aerodynamics of Tactical Weapons to Mach Number 3 and Angle of Attack 15 deg. Part I. Theory and Application.

    DTIC Science & Technology

    1977-02-01

    or = M or = 3 and O or = alpha or = 15 deg, respectively. Body and wing geometries can be quite general in that pointed or blunt nose bodies and sharp or blunt leading edge wings can be assumed. Computed results for several configurations compare well with experimental and other analytical results. The computer program is cost effective as it costs less than ten dollars per Mach number to compute the lift, drag, pitching moment, magnus moment, roll damping moment, and pitch damping moment of a typical wing-body shape on the CDC 6700 computer. Thus, a reasonably accurate

  1. Aerodynamics of Tactical Weapons to Mach Number 8 and Angle of Attack 180 deg. Part 1. Theory and Application.

    DTIC Science & Technology

    1980-10-01

    distance from the comer, and k, = Pa is the asymptote for the pressure distribution. For 8 > 0 on Section 2-3 and s’ very long, Pa equals the cone...pressure at the freestream Mach number. For 8 < 0, the assump- tion is made that Pa = p., or the freestream pressure. This last assumption can be relatively...distribution is then given by P = Pa + (P2 - Pa )e - 7 (4) (P2- Pa ) (5) where P2 is obtained from the Prandtl-Meyer relationship for an isentropic expansion

  2. Preliminary aerodynamic design considerations for advanced laminar flow aircraft configurations

    NASA Technical Reports Server (NTRS)

    Johnson, Joseph L., Jr.; Yip, Long P.; Jordan, Frank L., Jr.

    1986-01-01

    Modern composite manufacturing methods have provided the opportunity for smooth surfaces that can sustain large regions of natural laminar flow (NLF) boundary layer behavior and have stimulated interest in developing advanced NLF airfoils and improved aircraft designs. Some of the preliminary results obtained in exploratory research investigations on advanced aircraft configurations at the NASA Langley Research Center are discussed. Results of the initial studies have shown that the aerodynamic effects of configuration variables such as canard/wing arrangements, airfoils, and pusher-type and tractor-type propeller installations can be particularly significant at high angles of attack. Flow field interactions between aircraft components were shown to produce undesirable aerodynamic effects on a wing behind a heavily loaded canard, and the use of properly designed wing leading-edge modifications, such as a leading-edge droop, offset the undesirable aerodynamic effects by delaying wing stall and providing increased stall/spin resistance with minimum degradation of laminar flow behavior.

  3. Aerodynamic Characteristics of Water Rocket and Stabilization of Flight Trajectory

    NASA Astrophysics Data System (ADS)

    Watanabe, Rikio; Tomita, Nobuyuki; Takemae, Toshiaki

    The aerodynamic characteristics of water rockets are analyzed experimentally by wind tunnel testing. Aerodynamic devices such as vortex generators and dimples are tested and their effectiveness to the flight performance of water rocket is discussed. Attaching vortex generators suppresses the unsteady body fluttering. Dimpling the nose reduces the drag coefficient in high angles of attack. Robust design approach is applied to water rocket design for flight stability and optimum water rocket configuration is determined. Semi-sphere nose is found to be effective for flight stability and it is desirable for the safety of landing point. Stiffed fin attachment is required for fins to work properly as aerodynamic device and it enhances the flight stability of water rockets.

  4. Kinematic control of aerodynamic forces on an inclined flapping wing with asymmetric strokes.

    PubMed

    Park, Hyungmin; Choi, Haecheon

    2012-03-01

    In the present study, we conduct an experiment using a one-paired dynamically scaled model of an insect wing, to investigate how asymmetric strokes with different wing kinematic parameters are used to control the aerodynamics of a dragonfly-like inclined flapping wing in still fluid. The kinematic parameters considered are the angles of attack during the mid-downstroke (α(md)) and mid-upstroke (α(mu)), and the duration (Δτ) and time of initiation (τ(p)) of the pitching rotation. The present dragonfly-like inclined flapping wing has the aerodynamic mechanism of unsteady force generation similar to those of other insect wings in a horizontal stroke plane, but the detailed effect of the wing kinematics on the force control is different due to the asymmetric use of the angle of attack during the up- and downstrokes. For example, high α(md) and low α(mu) produces larger vertical force with less aerodynamic power, and low α(md) and high α(mu) is recommended for horizontal force (thrust) production. The pitching rotation also affects the aerodynamics of a flapping wing, but its dynamic rotational effect is much weaker than the effect from the kinematic change in the angle of attack caused by the pitching rotation. Thus, the influences of the duration and timing of pitching rotation for the present inclined flapping wing are found to be very different from those for a horizontal flapping wing. That is, for the inclined flapping motion, the advanced and delayed rotations produce smaller vertical forces than the symmetric one and the effect of pitching duration is very small. On the other hand, for a specific range of pitching rotation timing, delayed rotation requires less aerodynamic power than the symmetric rotation. As for the horizontal force, delayed rotation with low α(md) and high α(mu) is recommended for long-duration flight owing to its high efficiency, and advanced rotation should be employed for hovering flight for nearly zero horizontal force. The

  5. 1998 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    McMillin, S. Naomi (Editor)

    1999-01-01

    NASA's High-Speed Research Program sponsored the 1998 Aerodynamic Performance Technical Review on February 9-13, in Los Angeles, California. The review was designed to bring together NASA and industry HighSpeed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of. Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, and Flight Controls. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working HSCT aerodynamics. In particular, single and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program.

  6. 1998 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    McMillin, S. Naomi (Editor)

    1999-01-01

    NASA's High-Speed Research Program sponsored the 1998 Aerodynamic Performance Technical Review on February 9-13, in Los Angeles, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working HSCT aerodynamics. In particular, single and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program.

  7. 1999 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hahne, David E. (Editor)

    1999-01-01

    NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 1/Part 2 publication covers the design optimization and testing sessions.

  8. Aerodynamic Focusing Of High-Density Aerosols

    SciTech Connect

    Ruiz, D. E.; Fisch, Nathaniel

    2014-02-24

    High-density micron-sized particle aerosols might form the basis for a number of applications in which a material target with a particular shape might be quickly ionized to form a cylindrical or sheet shaped plasma. A simple experimental device was built in order to study the properties of high-density aerosol focusing for 1 m silica spheres. Preliminary results recover previous findings on aerodynamic focusing at low densities. At higher densities, it is demonstrated that the focusing properties change in a way which is consistent with a density dependent Stokes number.

  9. Aerodynamic Characteristics of a Feathered Dinosaur Measured Using Physical Models. Effects of Form on Static Stability and Control Effectiveness

    PubMed Central

    Evangelista, Dennis; Cardona, Griselda; Guenther-Gleason, Eric; Huynh, Tony; Kwong, Austin; Marks, Dylan; Ray, Neil; Tisbe, Adrian; Tse, Kyle; Koehl, Mimi

    2014-01-01

    We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While M. gui lived after Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies are tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver. PMID:24454820

  10. A study of aerodynamic heating distributions on a tip-fin controller installed on a Space Shuttle Orbiter model

    NASA Technical Reports Server (NTRS)

    Wittliff, C. E.

    1982-01-01

    The aerodynamic heating of a tip-fin controller mounted on a Space Shuttle Orbiter model was studied experimentally in the Calspan Advanced Technology Center 96 inch Hypersonic Shock Tunnel. A 0.0175 scale model was tested at Mach numbers from 10 to 17.5 at angles of attack typical of a shuttle entry. The study was conducted in two phases. In phase 1 testing a thermographic phosphor technique was used to qualitatively determine the areas of high heat-transfer rates. Based on the results of this phase, the model was instrumented with 40 thin-film resistance thermometers to obtain quantitative measurements of the aerodynamic heating. The results of the phase 2 testing indicate that the highest heating rates, which occur on the leading edge of the tip-fin controller, are very sensitive to angle of attack for alpha or = 30 deg. The shock wave from the leading edge of the orbiter wing impinges on the leading edge of the tip-fin controller resulting in peak values of h/h(Ref) in the range from 1.5 to 2.0. Away from the leading edge, the heat-transfer rates never exceed h/h(Ref) = 0.25 when the control surface, is not deflected. With the control surface deflected 20 deg, the heat-transfer rates had a maximum value of h/h(Ref) = 0.3. The heating rates are quite nonuniform over the outboard surface and are sensitive to angle of attack.

  11. Crossflow instability and transition on a circular cone at angle of attack in a Mach-6 quiet tunnel

    NASA Astrophysics Data System (ADS)

    Ward, Christopher A. C.

    To investigate the effect of roughness on the stationary crossflow vortices, roughness elements 2 inches from the nosetip were created in a Torlon insert section of a 7-deg half-angle cone at 6-deg angle of attack in the Boeing/AFOSR Mach-6 Quiet Tunnel. Roughness elements with depths and diameters on the order of 500 microns were found to have a significant effect on the generation of the stationary vortices and the location of crossflow-induced transition under quiet flow. Crossflow-induced transition was measured under fully quiet flow. The controlled roughness elements also appeared to dominate the generation of the vortices, overwhelming the effect that the random roughness of the cone had on the stationary vortices. It was surprising that the stationary vortices did not break down until close to the lee ray, disagreeing with linear stability computations. The roughness elements had the biggest effect on the stationary vortices at approximately 150-deg to 180-deg from the windward ray, depending on conditions. The roughness elements had a minimal impact on boundary-layer transition under noisy flow. The travelling crossflow waves were measured with Kulite pressure transducers under both noisy and quiet flow. The wave properties agreed well with computations by Texas A&M and experiments by TU Braunschweig under similar conditions. The amplitude of the waves was reduced by approximately 20 times when the noise levels in the BAM6QT were reduced from noisy to quiet. An interaction between the stationary and travelling crossflow waves was observed. When large stationary waves were induced with either controlled or uncontrolled roughness and the stationary waves passed near or over a fast pressure transducer, this fast pressure transducer measured a damped or distorted travelling crossflow wave. The nature of this stationary-travelling wave interaction is poorly understood but it may be a significant factor in crossflow-induced transition. A high

  12. Effect of tail size reductions on longitudinal aerodynamic characteristics of a three surface F-15 model with nonaxisymmetric nozzles

    NASA Technical Reports Server (NTRS)

    Frassinelli, Mark C.; Carson, George T., Jr.

    1990-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of horizontal and vertical tail size reductions on the longitudinal aerodynamic characteristics of a modified F-15 model with canards and 2-D convergent-divergent nozzles. Quantifying the drag decrease at low angles of attack produced by tail size reductions was the primary focus. The model was tested at Mach numbers of 0.40, 0.90, and 1.20 over an angle of attack of -2 degree to 10 degree. The nozzle exhaust flow was simulated using high pressure air at nozzle pressure ratios varying from 1.0 (jet off) to 7.5. Data were obtained on the baseline configuration with and without tails as well as with reduced horizontal and/or vertical tail sizes that were 75, 50, and 25 percent of the baseline tail areas.

  13. Transonic flow past a symmetrical airfoil at high angle of attack

    NASA Technical Reports Server (NTRS)

    Johnson, D. A.; Owen, F. K.; Bachalo, W. D.

    1979-01-01

    The results of an experimental investigation of shock-induced stall and leading-edge stall on a 64A010 airfoil section are presented. Advanced nonintrusive techniques - laser velocimetry, holographic interferometry, and buried-wire anemometry - were used in characterizing the inviscid and viscous flow regions. The measurements include Mach contours of the inviscid flow regions, and mean velocity, flow direction and Reynolds shear stress profiles in the separated regions. The experimental observations of this study are relevant to efforts to improve surface pressure prediction methods for airfoils at or near stall.

  14. Global Stability and Control Analysis of Aircraft at High Angles-of-Attack.

    DTIC Science & Technology

    1977-06-30

    500 deg/sec to 500 deg/sec and com - puting the eigenvalues of the linearized system around different equilibrium points for 6e = 2°. It is noticed...SYSTEMS, INC. 0" -114- 2 Ä»" • + + + + + _+ + + + + + + + + M I o o o a i i 8 8 in 8 8 XNXX * x x» x x x< x...history com - parisons show significant improvement in dynamic response and performance using the BACTM ARI gain, as opposed to nulling the rudder

  15. Basic Studies of the Unsteady Flow Past High Angle of Attack Airfoils

    DTIC Science & Technology

    1989-05-15

    30 0 ) NACA 0012 airfoil started impul- quantity defired in terms of local veloc- sively from rest. ity gradients. In contrast, the currently...ambiguity in the formed, for example, by focusing the laser direction of the velocity vector. In beam first with a long focal length spher- addition...Velocimetry, reserving the Adrian 1 2 , commonly known as "velocity bias term Laser Speckle Velocimetry for the technique", Consists of recording the flow

  16. Thrust Vectoring for Advanced Fighter Aircraft - High Angle of Attack Intake Investigations -

    DTIC Science & Technology

    2001-06-01

    radius Tt total temperature Numerical flow calculations ( CFD ) were to be performed to WAT normalized engine mass flow support the analysis of the... CFD ) investigations will be but could also result in damages of the engine and/or aircraft detailed. Results and comparisons between flows at small...dominated measured, by the shielding of the fuselage and the canard. 4.4 Data analysis During all the testing the intake lip position has been held fixed at

  17. Flowfield Measurements in the Vortex Wake of a Missile at High Angle of Attack in Turbulence

    DTIC Science & Technology

    1988-12-01

    4520 THYME = TIMER 4530 CALL OUTPUT(RELAY.ACT.01,STEPPER) 4540 CHKTIME = TIMER 4550 IF CHKTIME < ( THYME + DELAY 1) GOTO 4540 4560 CALL OPEN.CHANNEL...CALL OUTPUT(RELAY.ACT.0 ],STEPPER) 4640 CHKTIME = TIMER 4650 IF CHKTIME < ( THYME + DELAY2) GOTO 4640 4660 REM EACH PORT SAMPLE 10 TIMES 4670 FOR 11=1

  18. Fundamentals and Methods of High Angle-of-Attack Flying Qualities Research

    DTIC Science & Technology

    1988-01-01

    the a-, versus oxt criterion orginally developed by Jenny of McDonnell Aircraft Company (McAir) (1971) (see Reference (115) and (2)). The more recent...equations of motion. Other forms of this criteron are also found in the literature such as the a-, versus "b der ,rture criterion ( Jenny ). The C, R criterion...Dynamics." Journal of the Aeronautical Sciences, Vol. 21, No. 11, November 1954, pp. 749-762. 122. Johnston, D.E. and Hogge , J.R., "The Effects of Non

  19. Investigation of the High Angle of Attack Dynamics of the F-15B Using Bifurcation Analysis

    DTIC Science & Technology

    1990-12-01

    the formation of a family of limit cycles at the Hopf point. 3 Limit cycles are found by putting y1 and Yz into polar coordinates yI- pcos Y2 - psinO...C4L1,CQN1) BY M{AIMING THAT PARTICULARIC COEFICIETS SIGN (POSITIVE OR NEGATIVE) DPEDENT ON THE SIGN C OF THE SIDESLIP ANGLE ( BErA ). IF BETA IS NEGTIVE...EPA02L= -1.OdO ENDIF IF ((BETA .GE. -5.0) .AND. (BETA .LE. 5.0)) TH2 EPA02L=-l.0d0+(0 .OEOdO*( ( BErA +5.OdO)**2 .OdO) )- 1 (0.0040d0*(( BErA +5.odo)**3

  20. Plasma Control of Separated Flows on Delta Wings at High Angles of Attack

    DTIC Science & Technology

    2009-03-18

    breakdown or vortex burst, leads to rapid decrease of the lift to drag ratio. If the vortex burst on one wing side occurs earlier than on the other, then...discharge actuators which may be advantageous compared to other modern control techniques (synthetic jets, MEMS and etc.). Dielectric barrier...field. The experiments will be carried out using modern measurement and visualization techniques . They will demonstrate robustness of plasma

  1. Global Stability and Control Analysis of Aircraft at High Angles-of-Attack.

    DTIC Science & Technology

    1978-06-30

    7. 1 4 22 Ficken , 1951, Davidenko, 1953) to solve (2.1.4), in combination with the Newton method and Adams integration formulas. This particular...Atmospheric Flight, New York: Wiley. Ficken , F. A. (1951), "The Continuation Method for Functional Equations," Comm. Pure Appl. Math., Vol. 4. Gilbert

  2. A Lateral-Directional Controller for High-Angle-of-Attack Flight.

    DTIC Science & Technology

    1983-03-01

    0.004852; (ADi ’nput aC)volts ) v2out ’ 204.7; ( volts .a) 0/A out ) P :S92; 1/0 OTcmaadrgster) FVT 1; (locuf dtaotC porPCI( $a oc of !195 data Porn C...I - 178 - diptr is * pic ( Set interrupt jup) cbpto :* d( W); intptrl is ar( 2’ a~JKr~m ) ) ta i adiIm), ( befhm

  3. Effects of nose bluntness, roughness, and surface perturbations on the asymmetric flow past slender bodies at large angles of attack

    NASA Technical Reports Server (NTRS)

    Moskovitz, Cary A.; Dejarnette, F. R.; Hall, Robert M.

    1989-01-01

    The effects of such geometric perturbations as variations of model-tip sharpness and roughness, as well as discrete surface perturbations, on the asymmetric flow past slender bodies is experimentally investigated for the cases of a cone/cylinder model having a 10-deg semiapex angle and a 3.0-caliber tangent ogive model. Both models have base diameters of 3.5 inches, and were tested in laminar flow conditions at angles-of-attack in the 30-60 deg range. Single, discrete roughness elements were represented by beads; bead effectiveness was judged on the basis of the extent to which they affected the flowfield in various conditions.

  4. Design of an all-attitude flight control system to execute commanded bank angles and angles of attack

    NASA Technical Reports Server (NTRS)

    Burgin, G. H.; Eggleston, D. M.

    1976-01-01

    A flight control system for use in air-to-air combat simulation was designed. The input to the flight control system are commanded bank angle and angle of attack, the output are commands to the control surface actuators such that the commanded values will be achieved in near minimum time and sideslip is controlled to remain small. For the longitudinal direction, a conventional linear control system with gains scheduled as a function of dynamic pressure is employed. For the lateral direction, a novel control system, consisting of a linear portion for small bank angle errors and a bang-bang control system for large errors and error rates is employed.

  5. A method for predicting shock shapes and pressure distributions on two dimensional airfoils at large angles of attack

    NASA Technical Reports Server (NTRS)

    Kaattari, G. E.

    1973-01-01

    A method is presented for determining shock envelopes and pressure distributions for two-dimensional airfoils at angles of attack sufficiently large to cause shock detachment and subsonic flow over the windward surface of the airfoil. Correlation functions obtained from exact solutions are used to relate the shock standoff distance at the stagnation and sonic points of the body through a suitable choice for the shock shape. The necessary correlation functions were obtained from perfect gas solutions but may be extended to any gas flow for which the normal shock-density ratio can be specified.

  6. Lee-surface heating and flow phenomena on space shuttle orbiters at large angles of attack and hypersonic speeds

    NASA Technical Reports Server (NTRS)

    Hefner, J. N.

    1972-01-01

    The lee-surface flow phenomena on a delta-wing orbiter and a straight-wing orbiter have been investigated at angles of attack between 0 deg and 50 deg at a Mach number of 6. Limited studies of the delta-wing orbiter were conducted at a Mach number of 19. Heat-transfer data, pressure distributions, and oil-flow studies were employed to experimentally examine the nature of the surface flow and the severity of the lee-surface heating. The effects of Reynolds number on the flow field and heating were investigated. Problem areas are defined and areas for further study are recommended.

  7. Nacelle Aerodynamic and Inertial Loads (NAIL) project

    NASA Technical Reports Server (NTRS)

    1982-01-01

    A flight test survey of pressures measured on wing, pylon, and nacelle surfaces and of the operating loads on Boeing 747/Pratt & Whitney JT9D-7A nacelles was made to provide information on airflow patterns surrounding the propulsion system installations and to clarify processes responsible for inservice deterioration of fuel economy. Airloads at takeoff rotation were found to be larger than at any other normal service condition because of the combined effects of high angle of attack and high engine airflow. Inertial loads were smaller than previous estimates indicated. A procedure is given for estimating inlet airloads at low speeds and high angles of attack for any underwing high bypass ratio turbofan installation approximately resembling the one tested. Flight procedure modifications are suggested that may result in better fuel economy retention in service. Pressures were recorded on the core cowls and pylons of both engine installations and on adjacent wing surfaces for use in development of computer codes for analysis of installed propulsion system aerodynamic drag interference effects.

  8. Prediction of Aerodynamic Coefficients using Neural Networks for Sparse Data

    NASA Technical Reports Server (NTRS)

    Rajkumar, T.; Bardina, Jorge; Clancy, Daniel (Technical Monitor)

    2002-01-01

    Basic aerodynamic coefficients are modeled as functions of angles of attack and sideslip with vehicle lateral symmetry and compressibility effects. Most of the aerodynamic parameters can be well-fitted using polynomial functions. In this paper a fast, reliable way of predicting aerodynamic coefficients is produced using a neural network. The training data for the neural network is derived from wind tunnel test and numerical simulations. The coefficients of lift, drag, pitching moment are expressed as a function of alpha (angle of attack) and Mach number. The results produced from preliminary neural network analysis are very good.

  9. Effects of wing leading-edge deflection on low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Weston, R. P.

    1979-01-01

    Static force tests were conducted in the Langley V/STOL tunnel at a Reynolds number (based on the mean aerodynamic chord) of about 2.0 x 10 to the 6th power for an angle-of-attack range from about - 10 deg to 17 deg and angles of sideslip of 0 and + or - 5 deg. Limited flow visualization studies were also conducted in order to provide a qualitative assessment of leading-edge upwash characteristics.

  10. Low-speed aerodynamic characteristics of a transport configuration having a 42 deg swept supercritical airfoil wing and three tail height positions

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Sleeman, W. C., Jr.

    1974-01-01

    A low speed investigation was conducted in the Langley V/STOL tunnel to define the static stability characteristics of an advanced high subsonic speed transport aircraft model in the cruise configuration (no high lift system). The wing of the model had 42 deg sweep of the quarter chord line, an aspect ratio of 6.78, and supercritical airfoil sections. Three different horizontal tail configurations (high, mid, and low) were investigated on the complete model and for the model with the wing removed in order to assess effects of the wing flow field on the tail contributions to both longitudinal and lateral stability characteristics. All the model configurations investigated were tested over an angle of attack range from approximately -5 to 23 deg. Some model configurations were also tested over an angle of attack range from about 11 to 38 deg in order to explore the aerodynamic characteristics in the deep stall region.

  11. CFD Simulations in Support of Shuttle Orbiter Contingency Abort Aerodynamic Database Enhancement

    NASA Technical Reports Server (NTRS)

    Papadopoulos, Periklis E.; Prabhu, Dinesh; Wright, Michael; Davies, Carol; McDaniel, Ryan; Venkatapathy, E.; Wercinski, Paul; Gomez, R. J.

    2001-01-01

    Modern Computational Fluid Dynamics (CFD) techniques were used to compute aerodynamic forces and moments of the Space Shuttle Orbiter in specific portions of contingency abort trajectory space. The trajectory space covers a Mach number range of 3.5-15, an angle-of-attack range of 20deg-60deg, an altitude range of 100-190 kft, and several different settings of the control surfaces (elevons, body flap, and speed brake). Presented here are details of the methodology and comparisons of computed aerodynamic coefficients against the values in the current Orbiter Operational Aerodynamic Data Book (OADB). While approximately 40 cases have been computed, only a sampling of the results is provided here. The computed results, in general, are in good agreement with the OADB data (i.e., within the uncertainty bands) for almost all the cases. However, in a limited number of high angle-of-attack cases (at Mach 15), there are significant differences between the computed results, especially the vehicle pitching moment, and the OADB data. A preliminary analysis of the data from the CFD simulations at Mach 15 shows that these differences can be attributed to real-gas/Mach number effects. The aerodynamic coefficients and detailed surface pressure distributions of the present simulations are being used by the Shuttle Program in the evaluation of the capabilities of the Orbiter in contingency abort scenarios.

  12. Locally linearized longitudinal and lateral-directional aerodynamic stability and control derivaties for the X-29A aircraft

    NASA Technical Reports Server (NTRS)

    Budd, G. D.

    1984-01-01

    The locally linearized longitudinal and lateral-directional aerodynamic stability and control derivatives for the X-29A aircraft were calculated for altitudes ranging from sea level to 50,000 ft, Mach numbers from 0.2 to 1.5, and angles of attack from -5 deg to 25 deg. Several other parameters were also calculated, including aerodynamic force and moment coefficients, control face position, normal acceleration, static margin, and reference angle of attack.

  13. Experimental investigation of a new device to control the asymmetric flowfield on forebodies at large angles of attack

    NASA Technical Reports Server (NTRS)

    Moskovitz, Cary A.; Hall, Robert M.; Dejarnette, F. R.

    1990-01-01

    An exploratory experimental investigation of a new device to control the asymmetric flowfield on forebodies at large angles of attack has been conducted. The device is a rotatable forebody tip, which varies in cross section from circular at its base to elliptic at its tip. The device itself extends over a small portion of the aircraft or missile forebody. The device provides two important improvements. First, it replaced the normally random behavior of the nose side force as a function of nose tip orientation with a predictable and generally sinusoidal distribution and, second, the device showed promise for use as part of a vehicle control system, to be deflected in a prescribed manner to provide additional directional control for the vehicle. The device was tested on a cone/cylinder model having a 10 deg semiapex angle and on a 3.0 caliber tangent ogive model, each with a base diameter of 3.5 in, for angles of attack from 30 to 60 deg. Data were taken from 3 circumferential rows of pressure taps on each model at a Reynolds number of 84,000 based on cylinder diameter and by a helium-bubble flow visualization technique at a Reynolds number of 24,000.

  14. Effect of a rotating propeller on the separation angle of attack and distortion in ducted propeller inlets

    NASA Technical Reports Server (NTRS)

    Boldman, D. R.; Iek, C.; Hwang, D. P.; Larkin, M.; Schweiger, P.

    1993-01-01

    The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propeller removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

  15. Longitudinal aerodynamic and propulsion characteristics of a propulsive-wing V/STOL model at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Salters, L. B., Jr.; Schmeer, J. W.

    1973-01-01

    The aerodynamic and propulsion characteristics of a 1/6-scale propulsive-wing V/STOL air-powered model was investigated over the Mach number range from 0.40 to 0.96 and at angles of attack from -5 deg to 15 deg for several fan rotational speeds. Three fanduct-exit configurations were tested, including two exit areas. The model with 25-percent-thick wing had a drag-rise Mach number of 0.85, which is typical of aircraft with thinner, conventional, unswept wings.

  16. Stability Analysis of a mortar cover ejected at various Mach numbers and angles of attack

    NASA Astrophysics Data System (ADS)

    Schwab, Jane; Carnasciali, Maria-Isabel; Andrejczyk, Joe; Kandis, Mike

    2011-11-01

    This study utilized CFD software to predict the aerodynamic coefficient of a wedge-shaped mortar cover which is ejected from a spacecraft upon deployment of its Parachute Recovery System (PRS). Concern over recontact or collision between the mortar cover and spacecraft served as the impetus for this study in which drag and moment coefficients were determined at Mach numbers from 0.3 to 1.6 at 30-degree increments. These CFD predictions were then used as inputs to a two-dimensional, multi-body, three-DoF trajectory model to calculate the relative motion of the mortar cover and spacecraft. Based upon those simulations, the study concluded a minimal/zero risk of collision with either the spacecraft or PRS. Sponsored by Pioneer Aerospace.

  17. X-34 Vehicle Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Brauckmann, Gregory J.

    1998-01-01

    The X-34, being designed and built by the Orbital Sciences Corporation, is an unmanned sub-orbital vehicle designed to be used as a flying test bed to demonstrate key vehicle and operational technologies applicable to future reusable launch vehicles. The X-34 will be air-launched from an L-1011 carrier aircraft at approximately Mach 0.7 and 38,000 feet altitude, where an onboard engine will accelerate the vehicle to speeds above Mach 7 and altitudes to 250,000 feet. An unpowered entry will follow, including an autonomous landing. The X-34 will demonstrate the ability to fly through inclement weather, land horizontally at a designated site, and have a rapid turn-around capability. A series of wind tunnel tests on scaled models was conducted in four facilities at the NASA Langley Research Center to determine the aerodynamic characteristics of the X-34. Analysis of these test results revealed that longitudinal trim could be achieved throughout the design trajectory. The maximum elevon deflection required to trim was only half of that available, leaving a margin for gust alleviation and aerodynamic coefficient uncertainty. Directional control can be achieved aerodynamically except at combined high Mach numbers and high angles of attack, where reaction control jets must be used. The X-34 landing speed, between 184 and 206 knots, is within the capabilities of the gear and tires, and the vehicle has sufficient rudder authority to control the required 30-knot crosswind.

  18. Transonic aerodynamic and aeroelastic characteristics of a variable sweep wing

    NASA Technical Reports Server (NTRS)

    Goorjian, P. M.; Guruswamy, G. P.; Ide, H.; Miller, G.

    1985-01-01

    The flow over the B-1 wing is studied computationally, including the aeroelastic response of the wing. Computed results are compared with results from wind tunnel and flight tests for both low-sweep and high-sweep cases, at 25.0 deg. and 67.5 deg., respectively, for selected transonic Mach numbers. The aerodynamic and aeroelastic computations are made by using the transonic unsteady code ATRAN3S. Steady aerodynamic computations compare well with wind tunnel results for the 25.0 deg. sweep case and also for small angles of attack at the 67.5 deg. sweep case. The aeroelastic response results show that the wing is stable at the low sweep angle for the calculation at the Mach number at which there is a shock wave. In the higher sweep case, for the higher angle of attack at which oscillations were observed in the flight and wind tunnel tests, the calculations do not show any shock waves. Their absence lends support to the hypothesis that the observed oscillations are due to the presence of leading edge separation vortices and are not due to shock wave motion as was previously proposed.

  19. Improvements to the missile aerodynamic prediction code DEMON3

    NASA Technical Reports Server (NTRS)

    Dillenius, Marnix F. E.; Johnson, David L.; Lesieutre, Daniel J.

    1992-01-01

    The computer program DEMON3 was developed for the aerodynamic analysis of nonconventional supersonic configurations comprising a body with noncircular cross section and up to two wing or fin sections. Within a wing or fin section, the lifting surfaces may be cruciform, triform, planar, or low profile layouts; the planforms of the lifting surfaces allow for breaks in sweep. The body and fin sections are modeled by triplet and constant u-velocity panels, respectively, accounting for mutual body-fin interference. Fin thickness effects are included for the use of supersonic planar source panels. One of the unique features of DEMON3 is the modeling of high angle of attack vortical effects associated with the lifting surfaces and the body. In addition, shock expansion and Newtonian pressure calculation methods can be optionally engaged. These two dimensional nonlinear methods are augmented by aerodynamic interference determined from the linear panel methods. Depending on the geometric details of the body, the DEMON3 program can be used to analyze nonconventional configurations at angles of attack up to 25 degrees for Mach numbers from 1.1 to 6. Calculative results and comparisons with experimental data demonstrate the capabilities of DEMON3. Limitations and deficiencies are listed.

  20. Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

    2011-01-01

    Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

  1. Hypersonic and Supersonic Static Aerodynamics of Mars Science Laboratory Entry Vehicle

    NASA Technical Reports Server (NTRS)

    Dyakonov, Artem A.; Schoenenberger, Mark; Vannorman, John W.

    2012-01-01

    This paper describes the analysis of continuum static aerodynamics of Mars Science Laboratory (MSL) entry vehicle (EV). The method is derived from earlier work for Mars Exploration Rover (MER) and Mars Path Finder (MPF) and the appropriate additions are made in the areas where physics are different from what the prior entry systems would encounter. These additions include the considerations for the high angle of attack of MSL EV, ablation of the heatshield during entry, turbulent boundary layer, and other aspects relevant to the flight performance of MSL. Details of the work, the supporting data and conclusions of the investigation are presented.

  2. Identification of integro-differential systems for application to unsteady aerodynamics and aeroelasticity

    NASA Technical Reports Server (NTRS)

    Gupta, N. K.; Iliff, K. W.

    1985-01-01

    Integrodifferential equations for unsteady aerodynamic and aeroelastic phenomena are identified by means of several approaches. When the product of the frequency of motion and maximum time delay is much smaller than unity, the integral term can be approximated by a constant; when greater than unity, however, approximation of the integral is not possible. Approximations of integrodifferential models are needed to obtain identifiability. While the least-squares method may be used for model determination, the maximum likelihood technique is needed for accurate parameter estimation. High angle of attack and post stall/spin regions appear to have characteristics that can be satisfied by indicial models.

  3. Calculation of the flow field including boundary layer effects for supersonic mixed compression inlets at angles of attack

    NASA Technical Reports Server (NTRS)

    Vadyak, J.; Hoffman, J. D.

    1982-01-01

    The flow field in supersonic mixed compression aircraft inlets at angle of attack is calculated. A zonal modeling technique is employed to obtain the solution which divides the flow field into different computational regions. The computational regions consist of a supersonic core flow, boundary layer flows adjacent to both the forebody/centerbody and cowl contours, and flow in the shock wave boundary layer interaction regions. The zonal modeling analysis is described and some computational results are presented. The governing equations for the supersonic core flow form a hyperbolic system of partial differential equations. The equations for the characteristic surfaces and the compatibility equations applicable along these surfaces are derived. The characteristic surfaces are the stream surfaces, which are surfaces composed of streamlines, and the wave surfaces, which are surfaces tangent to a Mach conoid. The compatibility equations are expressed as directional derivatives along streamlines and bicharacteristics, which are the lines of tangency between a wave surface and a Mach conoid.

  4. Computation of supersonic laminar viscous flow past a pointed cone at angle of attack in spinning and coning motion

    NASA Technical Reports Server (NTRS)

    Agarwal, R.; Rakich, J. V.

    1978-01-01

    Computational results obtained with a parabolic Navier-Stokes marching code are presented for supersonic viscous flow past a pointed cone at angle of attack undergoing a combined spinning and coning motion. The code takes into account the asymmetries in the flow field resulting from the motion and computes the asymmetric shock shape, crossflow and streamwise shear, heat transfer, crossflow separation and vortex structure. The side force and moment are also computed. Reasonably good agreement is obtained with the side force measurements of Schiff and Tobak. Comparison is also made with the only available numerical inviscid analysis. It is found that the asymmetric pressure loads due to coning motion are much larger than all other viscous forces due to spin and coning, making viscous forces negligible in the combined motion.

  5. Experimental investigation of boundary layer transition on rotating cones in axial flow in 0 and 35 degrees angle of attack

    NASA Astrophysics Data System (ADS)

    Kargar, Ali; Mansour, Kamyar

    2015-11-01

    In this paper, experimental results using hot wire anemometer and smoke visualization are presented. The results obtained from the hot wire anemometer for critical Reynolds number and transitional Reynolds number are compared with previous results. Excellent agreement is found for the transitional Reynolds number. The results for the transitional Reynolds number are also compared to previous linear stability results. The results from the smoke visualization clearly show the crossflow vortices which arise in the transition process from a laminar to a turbulent flow. A non-zero angle of attack is also considered. we compare our results by linear stability theory which was done by. We just emphasis. Also we compare visualization and hot wire anemometer results graphically, our goal in this paper is to check reliability of using hot wire anemometer and smoke visualization in stability problem and check reliability of linear stability theory for this two cases and compare our results with some trusty experimental works.

  6. Vortex flow structures and interactions for the optimum thrust efficiency of a heaving airfoil at different mean angles of attack

    SciTech Connect

    Martín-Alcántara, A.; Fernandez-Feria, R.

    2015-07-15

    The thrust efficiency of a two-dimensional heaving airfoil is studied computationally for a low Reynolds number using a vortex force decomposition. The auxiliary potentials that separate the total vortex force into lift and drag (or thrust) are obtained analytically by using an elliptic airfoil. With these auxiliary potentials, the added-mass components of the lift and drag (or thrust) coefficients are also obtained analytically for any heaving motion of the airfoil and for any value of the mean angle of attack α. The contributions of the leading- and trailing-edge vortices to the thrust during their down- and up-stroke evolutions are computed quantitatively with this formulation for different dimensionless frequencies and heave amplitudes (St{sub c} and St{sub a}) and for several values of α. Very different types of flows, periodic, quasi-periodic, and chaotic described as St{sub c}, St{sub a}, and α, are varied. The optimum values of these parameters for maximum thrust efficiency are obtained and explained in terms of the interactions between the vortices and the forces exerted by them on the airfoil. As in previous numerical and experimental studies on flapping flight at low Reynolds numbers, the optimum thrust efficiency is reached for intermediate frequencies (St{sub c} slightly smaller than one) and a heave amplitude corresponding to an advance ratio close to unity. The optimal mean angle of attack found is zero. The corresponding flow is periodic, but it becomes chaotic and with smaller average thrust efficiency as |α| becomes slightly different from zero.

  7. 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

    NASA Technical Reports Server (NTRS)

    Brown, Clarence A , Jr

    1957-01-01

    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.

  8. 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

    NASA Technical Reports Server (NTRS)

    Brown, C. A., Jr.

    1957-01-01

    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.

  9. Prediction of Aerodynamic Coefficient using Genetic Algorithm Optimized Neural Network for Sparse Data

    NASA Technical Reports Server (NTRS)

    Rajkumar, T.; Bardina, Jorge; Clancy, Daniel (Technical Monitor)

    2002-01-01

    Wind tunnels use scale models to characterize aerodynamic coefficients, Wind tunnel testing can be slow and costly due to high personnel overhead and intensive power utilization. Although manual curve fitting can be done, it is highly efficient to use a neural network to define the complex relationship between variables. Numerical simulation of complex vehicles on the wide range of conditions required for flight simulation requires static and dynamic data. Static data at low Mach numbers and angles of attack may be obtained with simpler Euler codes. Static data of stalled vehicles where zones of flow separation are usually present at higher angles of attack require Navier-Stokes simulations which are costly due to the large processing time required to attain convergence. Preliminary dynamic data may be obtained with simpler methods based on correlations and vortex methods; however, accurate prediction of the dynamic coefficients requires complex and costly numerical simulations. A reliable and fast method of predicting complex aerodynamic coefficients for flight simulation I'S presented using a neural network. The training data for the neural network are derived from numerical simulations and wind-tunnel experiments. The aerodynamic coefficients are modeled as functions of the flow characteristics and the control surfaces of the vehicle. The basic coefficients of lift, drag and pitching moment are expressed as functions of angles of attack and Mach number. The modeled and training aerodynamic coefficients show good agreement. This method shows excellent potential for rapid development of aerodynamic models for flight simulation. Genetic Algorithms (GA) are used to optimize a previously built Artificial Neural Network (ANN) that reliably predicts aerodynamic coefficients. Results indicate that the GA provided an efficient method of optimizing the ANN model to predict aerodynamic coefficients. The reliability of the ANN using the GA includes prediction of aerodynamic

  10. The aerodynamics of insect flight.

    PubMed

    Sane, Sanjay P

    2003-12-01

    The flight of insects has fascinated physicists and biologists for more than a century. Yet, until recently, researchers were unable to rigorously quantify the complex wing motions of flapping insects or measure the forces and flows around their wings. However, recent developments in high-speed videography and tools for computational and mechanical modeling have allowed researchers to make rapid progress in advancing our understanding of insect flight. These mechanical and computational fluid dynamic models, combined with modern flow visualization techniques, have revealed that the fluid dynamic phenomena underlying flapping flight are different from those of non-flapping, 2-D wings on which most previous models were based. In particular, even at high angles of attack, a prominent leading edge vortex remains stably attached on the insect wing and does not shed into an unsteady wake, as would be expected from non-flapping 2-D wings. Its presence greatly enhances the forces generated by the wing, thus enabling insects to hover or maneuver. In addition, flight forces are further enhanced by other mechanisms acting during changes in angle of attack, especially at stroke reversal, the mutual interaction of the two wings at dorsal stroke reversal or wing-wake interactions following stroke reversal. This progress has enabled the development of simple analytical and empirical models that allow us to calculate the instantaneous forces on flapping insect wings more accurately than was previously possible. It also promises to foster new and exciting multi-disciplinary collaborations between physicists who seek to explain the phenomenology, biologists who seek to understand its relevance to insect physiology and evolution, and engineers who are inspired to build micro-robotic insects using these principles. This review covers the basic physical principles underlying flapping flight in insects, results of recent experiments concerning the aerodynamics of insect flight, as well

  11. Experimental Aerodynamic Characteristics of the Pegasus Air-Launched Booster and Comparisons with Predicted and Flight Results

    NASA Technical Reports Server (NTRS)

    Rhode, M. N.; Engelund, Walter C.; Mendenhall, Michael R.

    1995-01-01

    Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for the Pegasus and Pegasus XL configurations over a Mach number range from 1.6 to 6 and angles of attack from -4 to +24 degrees. Angle of sideslip was varied from -6 to +6 degrees, and control surfaces were deflected to obtain elevon, aileron, and rudder effectiveness. Experimental data for the Pegasus configuration are compared with engineering code predictions performed by Nielsen Engineering & Research, Inc. (NEAR) in the aerodynamic design of the Pegasus vehicle, and with results from the Aerodynamic Preliminary Analysis System (APAS) code. Comparisons of experimental results are also made with longitudinal flight data from Flight #2 of the Pegasus vehicle. Results show that the longitudinal aerodynamic characteristics of the Pegasus and Pegasus XL configurations are similar, having the same lift-curve slope and drag levels across the Mach number range. Both configurations are longitudinally stable, with stability decreasing towards neutral levels as Mach number increases. Directional stability is negative at moderate to high angles of attack due to separated flow over the vertical tail. Dihedral effect is positive for both configurations, but is reduced 30-50 percent for the Pegasus XL configuration because of the horizontal tail anhedral. Predicted longitudinal characteristics and both longitudinal and lateral-directional control effectiveness are generally in good agreement with experiment. Due to the complex leeside flowfield, lateral-directional characteristics are not as well predicted by the engineering codes. Experiment and flight data are in good agreement across the Mach number range.

  12. Control research in the NASA high-alpha technology program

    NASA Technical Reports Server (NTRS)

    Gilbert, William P.; Nguyen, Luat T.; Gera, Joseph

    1990-01-01

    NASA is conducting a focused technology program, known as the High-Angle-of-Attack Technology Program, to accelerate the development of flight-validated technology applicable to the design of fighters with superior stall and post-stall characteristics and agility. A carefully integrated effort is underway combining wind tunnel testing, analytical predictions, piloted simulation, and full-scale flight research. A modified F-18 aircraft has been extensively instrumented for use as the NASA High-Angle-of-Attack Research Vehicle used for flight verification of new methods and concepts. This program stresses the importance of providing improved aircraft control capabilities both by powered control (such as thrust-vectoring) and by innovative aerodynamic control concepts. The program is accomplishing extensive coordinated ground and flight testing to assess and improve available experimental and analytical methods and to develop new concepts for enhanced aerodynamics and for effective control, guidance, and cockpit displays essential for effective pilot utilization of the increased agility provided.

  13. Wing kinematics measurement and aerodynamics of a dragonfly in turning flight.

    PubMed

    Li, Chengyu; Dong, Haibo

    2017-02-03

    This study integrates high-speed photogrammetry, 3D surface reconstruction, and computational fluid dynamics to explore a dragonfly (Erythemis Simplicicollis) in free flight. Asymmetric wing kinematics and the associated aerodynamic characteristics of a turning dragonfly are analyzed in detail. Quantitative measurements of wing kinematics show that compared to the outer wings, the inner wings sweep more slowly with a higher angle of attack during the downstroke, whereas they flap faster with a lower angle of attack during the upstroke. The inner-outer asymmetries of wing deviations result in an oval wingtip trajectory for the inner wings and a figure-eight wingtip trajectory for the outer wings. Unsteady aerodynamics calculations indicate significantly asymmetrical force production between the inner and outer wings, especially for the forewings. Specifically, the magnitude of the drag force on the inner forewing is approximately 2.8 times greater than that on the outer forewing during the downstroke. In the upstroke, the outer forewing generates approximately 1.9 times greater peak thrust than the inner forewing. To keep the body aloft, the forewings contribute approximately 64% of the total lift, whereas the hindwings provide 36%. The effect of forewing-hindwing interaction on the aerodynamic performance is also examined. It is found that the hindwings can benefit from this interaction by decreasing power consumption by 13% without sacrificing force generation.

  14. Unstructured Grid Viscous Flow Simulation Over High-Speed Research Technology Concept Airplane at High-Lift Conditions

    NASA Technical Reports Server (NTRS)

    Ghaffari, Farhad

    1999-01-01

    Numerical viscous solutions based on an unstructured grid methodology are presented for a candidate high-speed civil transport configuration, designated as the Technology Concept Airplane (TCA), within the High-Speed Research (HSR) program. The numerical results are obtained on a representative TCA high-lift configuration that consisted of the fuselage and the wing, with deflected full-span leading-edge and trailing-edge flaps. Typical on-and off-surface flow structures, computed at high-lift conditions appropriate for the takeoff and landing, indicated features that are generally plausible. Reasonable surface pressure correlations between the numerical results and the experimental data are obtained at free-stream Mach number M(sub infinity) = 0.25 and Reynolds number based on bar-c R(sub c) = 8 x 10(exp 6) for moderate angles of attack of 9.7 deg. and 13.5 deg. However, above and below this angle-of-attack range, the correlation between computed and measured pressure distributions starts to deteriorate over the examined angle-of-attack range. The predicted longitudinal aerodynamic characteristics are shown to correlate very well with existing experimental data across the examined angle-of-attack range. An excellent agreement is also obtained between the predicted lift-to-drag ratio and the experimental data over the examined range of flow conditions.

  15. Unsteady Aerodynamic Modeling in Roll for the NASA Generic Transport Model

    NASA Technical Reports Server (NTRS)

    Murphy, Patrick C.; Klein, Vladislav; Frink, Neal T.

    2012-01-01

    Reducing the impact of loss-of-control conditions on commercial transport aircraft is a primary goal of the NASA Aviation Safety Program. One aspect in developing the supporting technologies is to improve the aerodynamic models that represent these adverse conditions. Aerodynamic models appropriate for loss of control conditions require a more general mathematical representation to predict nonlinear unsteady behaviors. In this paper, a more general mathematical model is proposed for the subscale NASA Generic Transport Model (GTM) that covers both low and high angles of attack. Particular attention is devoted to the stall region where full-scale transports have demonstrated a tendency for roll instability. The complete aerodynamic model was estimated from dynamic wind-tunnel data. Advanced computational methods are used to improve understanding and visualize the flow physics within the region where roll instability is a factor.

  16. High speed propeller acoustics and aerodynamics - A boundary element approach

    NASA Technical Reports Server (NTRS)

    Farassat, F.; Myers, M. K.; Dunn, M. H.

    1989-01-01

    The Boundary Element Method (BEM) is applied in this paper to the problems of acoustics and aerodynamics of high speed propellers. The underlying theory is described based on the linearized Ffowcs Williams-Hawkings equation. The surface pressure on the blade is assumed unknown in the aerodynamic problem. It is obtained by solving a singular integral equation. The acoustic problem is then solved by moving the field point inside the fluid medium and evaluating some surface and line integrals. Thus the BEM provides a powerful technique in calculation of high speed propeller aerodynamics and acoustics.

  17. High speed propeller acoustics and aerodynamics - A boundary element approach

    NASA Technical Reports Server (NTRS)

    Farassat, F.; Myers, M. K.; Dunn, M. H.

    1989-01-01

    The Boundary Element Method (BEM) is applied in this paper to the problems of acoustics and aerodynamics of high speed propellers. The underlying theory is described based on the linearized Ffowcs Williams-Hawkings equation. The surface pressure on the blade is assumed unknown in the aerodynamic problem. It is obtained by solving a singular integral equation. The acoustic problem is then solved by moving the field point inside the fluid medium and evaluating some surface and line integrals. Thus the BEM provides a powerful technique in calculation of high speed propeller aerodynamics and acoustics.

  18. Aerodynamics and vortical structures in hovering fruitflies

    NASA Astrophysics Data System (ADS)

    Meng, Xue Guang; Sun, Mao

    2015-03-01

    We measure the wing kinematics and morphological parameters of seven freely hovering fruitflies and numerically compute the flows of the flapping wings. The computed mean lift approximately equals to the measured weight and the mean horizontal force is approximately zero, validating the computational model. Because of the very small relative velocity of the wing, the mean lift coefficient required to support the weight is rather large, around 1.8, and the Reynolds number of the wing is low, around 100. How such a large lift is produced at such a low Reynolds number is explained by combining the wing motion data, the computed vortical structures, and the theory of vorticity dynamics. It has been shown that two unsteady mechanisms are responsible for the high lift. One is referred as to "fast pitching-up rotation": at the start of an up- or downstroke when the wing has very small speed, it fast pitches down to a small angle of attack, and then, when its speed is higher, it fast pitches up to the angle it normally uses. When the wing pitches up while moving forward, large vorticity is produced and sheds at the trailing edge, and vorticity of opposite sign is produced near the leading edge and on the upper surface, resulting in a large time rate of change of the first moment of vorticity (or fluid impulse), hence a large aerodynamic force. The other is the well known "delayed stall" mechanism: in the mid-portion of the up- or downstroke the wing moves at large angle of attack (about 45 deg) and the leading-edge-vortex (LEV) moves with the wing; thus, the vortex ring, formed by the LEV, the tip vortices, and the starting vortex, expands in size continuously, producing a large time rate of change of fluid impulse or a large aerodynamic force.

  19. The Effect of Trailing Vortices on the Production of Lift on an Airfoil Undergoing a Constant Rate of Change of Angle of Attack.

    DTIC Science & Technology

    1983-12-01

    The purpose of this study was to investigate the effect a trailing vortex wake has on an airfoil undergoing a constant rate of change of angle of...When applied to the constant rate - of - change of angle-of-attack problem, the results showed that a trailing vortex wake has a measurable and

  20. Stability and control characteristics of a three-surface advanced fighter configuration at angles of attack up to 45 deg. [conducted in the Langley 16-foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.; Leavitt, L. D.

    1981-01-01

    The tests were conducted at Mach numbers from 0.40 to 0.90, at angles of attack up to 45 deg for the lower Mach numbers, and at angles of sideslip up to 15 deg. The model variations under study included adding a canard surface and deflecting horizontal tails, ailerons, and rudders.