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Sample records for high angle-of-attack aerodynamics

  1. 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.

  2. High-angle-of-attack aerodynamics - Lessons learned

    NASA Technical Reports Server (NTRS)

    Chambers, J. R.

    1986-01-01

    Recently, the military and civil technical communities have undertaken numerous studies of the high angle-of-attack aerodynamic characteristics of advanced airplane and missile configurations. The method of approach and the design methodology employed have necessarily been experimental and exploratory in nature, due to the complex nature of separated flows. However, despite the relatively poor definition of many of the key aerodynamic phenomena involved for high-alpha conditions, some generic guidelines for design consideration have been identified. The present paper summarizes some of the more important lessons learned in the area of high angle-of-attack aerodynamics with examples of a number of key concepts and with particular emphasis on high-alpha stability and control characteristics of high performance aircraft. Topics covered in the discussion include the impact of design evolution, forebody flows, control of separated flows, configuration effects, aerodynamic controls, wind-tunnel flight correlation, and recent NASA research activities.

  3. Aerodynamic characteristics of cruciform missiles at high angles of attack

    NASA Technical Reports Server (NTRS)

    Lesieutre, Daniel J.; Mendenhall, Michael R.; Nazario, Susana M.; Hemsch, Michael J.

    1987-01-01

    An aerodynamic prediction method for missile aerodynamic performance and preliminary design has been developed to utilize a newly available systematic fin data base and an improved equivalent angle of attack methodology. The method predicts total aerodynamic loads and individual fin forces and moments for body-tail (wing-body) and canard-body-tail configurations with cruciform fin arrangements. The data base and the prediction method are valid for angles of attack up to 45 deg, arbitrary roll angles, fin deflection angles between -40 deg and 40 deg, Mach numbers between 0.6 and 4.5, and fin aspect ratios between 0.25 and 4.0. The equivalent angle of attack concept is employed to include the effects of vorticity and geometric scaling.

  4. Aerodynamic characteristics of general aviation at high angle of attack with the propeller slipstream

    NASA Technical Reports Server (NTRS)

    Matsuo, N.; Hirano, S.

    1986-01-01

    The aerodynamic characteristics of the FA-300 business aircraft at high angle of attack with the propeller stream are described. The FA-300 offers two types, FA-300-700 for 340 HP, and -710 for 450 Hp of the engine. The effects of the propeller slipstream on the high angle of the attack are discussed.

  5. 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.

  6. An Overview of the RTO Symposium on Vortex Flow and High Angle of Attack Aerodynamics

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2002-01-01

    In May of 2001 the Research and Technology Organization (RTO) sponsored a symposium on Vortex Flow and High Angle of Attack aerodynamics. Forty-six 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. Selected highlights are included in this paper to provide a perspective toward the scope of the full symposium.

  7. Shuttle Orbiter Contingency Abort Aerodynamics: Real-Gas Effects and High Angles of Attack

    NASA Technical Reports Server (NTRS)

    Prabhu, Dinesh K.; Papadopoulos, Periklis E.; Davies, Carol B.; Wright, Michael J.; McDaniel, Ryan D.; Venkatapathy, Ethiraj; Wercinski, Paul F.

    2005-01-01

    An important element of the Space Shuttle Orbiter safety improvement plan is the improved understanding of its aerodynamic performance so as to minimize the "black zones" in the contingency abort trajectories [1]. These zones are regions in the launch trajectory where it is predicted that, due to vehicle limitations, the Orbiter will be unable to return to the launch site in a two or three engine-out scenario. Reduction of these zones requires accurate knowledge of the aerodynamic forces and moments to better assess the structural capability of the vehicle. An interesting aspect of the contingency abort trajectories is that the Orbiter would need to achieve angles of attack as high as 60deg. Such steep attitudes are much higher than those for a nominal flight trajectory. The Orbiter is currently flight certified only up to an angle of attack of 44deg at high Mach numbers and has never flown at angles of attack larger than this limit. Contingency abort trajectories are generated using the data in the Space Shuttle Operational Aerodynamic Data Book (OADB) [2]. The OADB, a detailed document of the aerodynamic environment of the current Orbiter, is primarily based on wind-tunnel measurements (over a wide Mach number and angle-of-attack range) extrapolated to flight conditions using available theories and correlations, and updated with flight data where available. For nominal flight conditions, i.e., angles of attack of less than 45deg, the fidelity of the OADB is excellent due to the availability of flight data. However, at the off-nominal conditions, such as would be encountered on contingency abort trajectories, the fidelity of the OADB is less certain. The primary aims of a recent collaborative effort (completed in the year 2001) between NASA and Boeing were to determine: 1) accurate distributions of pressure and shear loads on the Orbiter at select points in the contingency abort trajectory space; and 2) integrated aerodynamic forces and moments for the entire

  8. 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.

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

    NASA Technical Reports Server (NTRS)

    Chapman, G. T.; Keener, E. R.; Malcolm, G. N.

    1976-01-01

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

  10. 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.

  11. Aerodynamic Loads on Tails at High Angles of Attack and Sideslip

    NASA Technical Reports Server (NTRS)

    Polhamus, E. C.; Spahr, J. R.

    1957-01-01

    Results are presented for the loads and moments acting on the individual tail surfaces of a body-tail combination over a wide range of angles of attack and sideslip. The effects of forebody length and panel-panel interference on the characteristics are included. It is shown that large nonlinear variations in these loads and moments, which occur at some combinations of angle of attack and sideslip, cannot be predicted by low-angle theory. A relatively simple, but general, theoretical method for calculating these load and moment characteristics is described, and the results from this method are found to be in good agreement with experiment provided the initial positions of the forebody vortices are known. It is shown that a simple application of slender-body theory can be used to predict the side loads due to sideslip that are contributed by a vertical tail on a wide variety of wing-body-tail combinations at low angles of attack. For several configurations, changes are indicated which reduced the vertical-tail loads per unit yawing moment of each complete configuration at large angles of attack. Some results are presented on the effect of high angle of attack on the induced-flow field and tail loads due to a wing at supersonic speed.

  12. 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.

  13. 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 in this paper 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 in this paper 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-18 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.

  14. Determination of aerodynamic parameters of a fighter airplane from flight data at high angles of attack

    NASA Technical Reports Server (NTRS)

    Klein, V.; Batterson, J. G.; Abbasy, I.

    1983-01-01

    A procedure for the estimation of airplane model structure and parameters is applied to data from a modern fighter airplane operating within an angle of attack range of 5 to 60 deg. The paper briefly describes the airplane, flight and wind tunnel data available, and the estimation method. The results presented contain basic longitudinal characteristics of the airplane and the estimates of aerodynamic parameters in the yawing-moment equations. These estimates are obtained from small and large amplitude maneuvers. Because the latter set of data was not suitable for airplane identification, some of the large amplitude maneuvers were joined together and then partitioned into subsets according to the values of angle of attack. Each subset was then analyzed as a separate data set. Most of the estimated parameters and functions are in good agreement with the wind tunnel measurements. The estimated lateral parameters in the model equations also demonstrate good prediction capabilities.

  15. Implications of scale effect for the prediction of high angle of attack aerodynamics

    NASA Astrophysics Data System (ADS)

    Munro, C. D.; Krus, P.; Jouannet, C.

    2005-04-01

    Scale effect is of significance in all experimental aerodynamics testing. At high angles of attack such issues as Reynolds and Mach number scaling are further complicated by the presence of complex, unsteady separated flow structures. The understanding and quantification of such flow effects remains a substantial challenge to the combat aircraft designer. The challenge, especially during conceptual and preliminary design stages, is to obtain sufficiently accurate information to make relevant design decisions, addressing potential weaknesses before proceeding to detail design. The focus in the present review is on the application of the water tunnel in such cases. The combination of qualitative and quantitative data obtainable from the water tunnel provides a useful complement to other tools during the early design stages. The effect of Reynolds number in particular is however critical; and appreciation of its effects are critical to the effective use of such a tool. It is shown however that for certain configurations such effects can be small in comparison to other experimental effects.

  16. Aerodynamic characteristics of two general aviation canard configurations at high angles of attack

    NASA Technical Reports Server (NTRS)

    Chambers, J. R.; Yip, L. P.

    1984-01-01

    The results of wind tunnel tests of two propeller driven canard general aviation aircraft models at high angles of attack are reported. Both tractor and pusher prop configurations were examined. Angles of attack of -6 to 40 deg were used with the pusher model at Re of 1,600,000, and from -30 to 90 deg and Re of 550,000 for the tractor model. The tests showed that the canard would stall long before the wing and produce a nose-down tendency, thus effectively keeping the aircraft out of the stall regime. However, a sequence of pilot actions or design factors such as the airfoils, relative geometry of the canard and wing, the propeller location and the center of gravity location could introduce a wide variance in stall characteristics from one aircraft to another.

  17. Partitioning of flight data for aerodynamic modeling of aircraft at high angles of attack

    NASA Technical Reports Server (NTRS)

    Batterson, James G.; Klein, Vladislav

    1987-01-01

    It is sometimes necessary to determine aerodynamic model structure and estimate associated stability and control derivatives for airplanes from flight data that cover a large range of angle of attack or sideslip. One method of dealing with that problem is through data partitioning. The main purpose of this paper is to provide an explanation of a data partitioning procedure and its application and to discuss both the power and limitations of that procedure for the analysis of large maneuvers of aircraft. The partitioning methodology is shown to provide estimates for coefficients of those regressors that are well excited in the aircraft motion. In particular, primary lateral stability and damping derivatives are identified throughout the maneuver ranges.

  18. High angle-of-attack aerodynamics of a strake-canard-wing V/STOL fighter configuration

    NASA Technical Reports Server (NTRS)

    Durston, D. A.; Schreiner, J. A.

    1983-01-01

    High angle-of-attack aerodynamic data are analyzed for a strake-canard-wing V/STOL fighter configuration. The configuration represents a twin-engine supersonic V/STOL fighter aircraft which uses four longitudinal thrust-augmenting ejectors to provide vertical lift. The data were obtained in tests of a 9.39 percent scale model of the configuration in the NASA Ames 12-Foot Pressure Wind Tunnel, at a Mach number of 0.2. Trimmed aerodynamic characteristics, longitudinal control power, longitudinal and lateral/directional stability, and effects of alternate strake and canard configurations are analyzed. The configuration could not be trimmed (power-off) above 12 deg angle of attack because of the limited pitch control power and the high degree of longitudinal instability (28 percent) at this Mach number. Aerodynamic center location was found to be controllable by varying strake size and canard location without significantly affecting lift and drag. These configuration variations had relatively little effect on the lateral/directional stability up to 10 deg angle of attack.

  19. Method for estimating the aerodynamic coefficients of wind turbine blades at high angles of attack

    NASA Astrophysics Data System (ADS)

    Beans, E. W.; Jakubowski, G. S.

    1983-12-01

    The method is based on the hypothesis that at high angles of attack the force on an airfoil is produced by the deflection of the fluid across the lower surface. It is also hypothesized that all airfoils behave the same regardless of shape and that the effects of circulation and skin friction are small. It is pointed out that the expression for the force N normal to the airfoil due to momentum exchange can be written in terms of the component parallel to the flow (drag) and the component perpendicular to the flow (lift). A comparison of estimated values with measured values and generally accepted data indicates that the method given here estimates coefficients which are low. It is thought that the difference may derive from the persistence of circulation at high angles of attack. Low estimates are not seen as a serious limitation to the designer of wind turbines. Owing to the fifth power diameter relation, the effect of a low estimate of performance on the inner portion of the blade is minimized.

  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. Aerodynamic damping of blade flap motions at high angles of attack

    SciTech Connect

    Hansen, A.C.

    1995-09-01

    The YawDyn computer code is used to calculate the aerodynamic damping for a typical teetering rotor configuration. The code has been modified to calculate the net work done by aerodynamic forces in one complete rotor revolution. All cases were run for a teetering rotor similar to the ESI-80 with a specified teeter angle motion in order to isolate the aerodynamic effects from the inertial and gravitational loads. Effects of nonlinear static stall, dynamic stall, dynamic inflow, and delayed static stall due to rotation stability of flap motions in high winds. Contributions of the various steady and unsteady aerodynamic effects are presented for two airfoils: the LS(1) and the NREL Thin Airfoil Family (S805A, S806, S807). Teeter stability is compared for a blade with 10{degree} of linear twist and a blade with optimum aerodynamic twist.

  2. Theoretical model of aerodynamic forces at high speeds and angle of attack in a stratified flowfield of UF6

    SciTech Connect

    Harloff, G.J.

    1985-09-01

    A theoretical aerodynamic model of lift and drag forces on a flat plate at angle of attack and at hypersonic speeds is presented. Real gas effects and friction drag are accounted for. Theoretical results are presented as a function of the viscous interaction parameter. The performance for two geometries is presented. 3 refs., 8 figs., 4 tabs.

  3. 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.

  4. 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.

  5. 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.

  6. X-31 at High Angle of Attack

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The X-31 aircraft, on a research mission from NASA's Dryden Flight Research Center, Edwards, California, is flying nearly perpendicular to the flight path while performing the Herbst maneuver. Effectively using the entire airframe as a speed brake and using the aircraft's unique thrust vectoring system to maintain control, the pilot rapidly rolls the aircraft to reverse the direction of flight, completing the maneuver with acceleration back to high speed in the opposite direction. This type of turning capability could reduce the turning time of a fighter aircraft by 30 percent. The Herbst maneuver was first conducted in an X-31 on April 29, 1993, in the No. 2 X-31 aircraft by German test pilot Karl-Heinz Lang. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal

  7. X-31 at High Angle of Attack

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The X-31 aircraft, on a research mission from NASA's Dryden Flight Research Center, Edwards, California, is flying nearly perpendicular to the flight path while performing the Herbst maneuver. Effectively using the entire airframe as a speed brake and using the aircrafts unique thrust vectoring system to maintain control, the pilot rapidly rolls the aircraft to reverse the direction of flight, completing the maneuver with acceleration back to high speed in the opposite direction. This type of turning capability could reduce the turning time of a fighter aircraft by 30 percent. The Herbst maneuver was first conducted in an X-31 on April 29, 1993, in the No. 2 X-31 aircraft by German test pilot Karl-Heinz Lang. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal

  8. X-31 at High Angle of Attack

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The X-31 aircraft on a research mission from NASA's Dryden Flight Research Center, Edwards, California, is flying nearly perpendicular to the flight path while performing the Herbst maneuver. Effectively using the entire airframe as a speed brake and using the aircrafts unique thrust vectoring system to maintain control, the pilot rapidly rolls the aircraft to reverse the direction of flight, completing the maneuver with acceleration back to high speed in the opposite direction. This type of turning capability could reduce the turning time of a fighter aircraft by 30 percent. The Herbst maneuver was first conducted in an X-31 on April 29, 1993, in the No. 2 X-31 aircraft by German test pilot Karl-Heinz Lang. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal

  9. 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.

  10. 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).

  11. 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.

  12. 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).

  13. Aerodynamics of slender finned bodies at large angles of attack

    NASA Technical Reports Server (NTRS)

    Agnone, A. M.; Zakkay, V.; Tory, E.; Stallings, R.

    1977-01-01

    In certain missions finned missiles perform slewing maneuvers. Here, large angles of attack are attained. Experimental data needed to understand the aerodynamics of such vehicles are presented. The purpose of this investigation was to study the interaction of the body flow field with that produced by the fins and the resulting effects on the aerodynamic forces and moments. The experiments were conducted at a nominal Mach number of 2.7 and angles of attack from 0 to 50 deg, with two different models. The tests were performed in a range of Reynolds number from 1.5 x 10 to the 6th to 4 x 10 to the 7th per foot (to cover both the laminar and fully turbulent regimes.) Several fin roll angles were investigated. Static pressures on both body and fin surfaces are reported.

  14. X-29 high angle of attack

    NASA Technical Reports Server (NTRS)

    Ishmael, Stephen D.; Smith, Rogers E.; Purifoy, Dana D.; Womer, Rodney K.

    1990-01-01

    Flight test program highlights are discussed for the X-29 high angle-of-attack (AOA) aircraft. The AOA envelope extended from 10 to 66 deg; the X-29 exhibited precise pitch control, allowing AOA to be maintained within 1 deg during stabilized points as well as permitting rapid recoveries from all AOAs. Attention is given to controllability degradation above 40-deg AOA due to asymmetric yawing moments. The use of this aircraft as a fundamental research tool which complements analytical methods is powerfully justified by the obviation of scaling effects.

  15. 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.

  16. 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.

  17. X-31 Demonstrating High Angle of Attack - Herbst Maneuver

    NASA Technical Reports Server (NTRS)

    1994-01-01

    The X-31 aircraft on a research mission from NASA's Dryden Flight Research Facility, Edwards, California, is flying nearly perpendicular to the flight path while performing the Herbst maneuver. Effectively using the entire airframe as a speed brake and using the aircraft's unique thrust vectoring system to maintain control, the pilot rapidly rolls the aircraft to reverse the direction of flight, completing the maneuver with acceleration back to high speed in the opposite direction. This type of turning capability could reduce the turning time of a fighter aircraft by 30 percent. The Herbst maneuver was first conducted in an X-31 on April 29, 1993, in the No. 2 aircraft by German test pilot Karl-Heinz Lang. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal stall

  18. Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2

    NASA Technical Reports Server (NTRS)

    Keener, E. R.; Taleghani, J.

    1975-01-01

    Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.

  19. High-angle-of-attack yawing moment asymmetry of the X-31 aircraft from flight test

    NASA Technical Reports Server (NTRS)

    Cobleigh, Brent R.

    1994-01-01

    Significant yawing moment asymmetries were encountered during the high-angle-of-attack envelope expansion of the two X-31 aircraft. These asymmetries led to position saturations of the thrust vector vanes and trailing-edge flaps during some of the dynamic stability axis rolling maneuvers at high angles of attack. This slowed the high-angle-of-attack envelope expansion and resulted in maneuver restrictions. Several aerodynamic modifications were made to the X-31 forebody with the goal of minimizing the asymmetry. A method for determining the yawing moment asymmetry from flight data was developed and an analysis of the various configuration changes completed. The baseline aircraft were found to have significant asymmetries above 45 deg angle of attack with the largest asymmetry typically occurring around 60 deg angle of attack. Applying symmetrical boundary layer transition strips along the forebody sides increased the magnitude of the asymmetry and widened the angle-of-attack range over which the largest asymmetry acted. Installing longitudinal forebody strakes and rounding the sharp nose of the aircraft caused the yawing moment asymmetry magnitude to be reduced. The transition strips and strakes made the asymmetry characteristic of the aircraft more repeatable than the clean forebody configuration. Although no geometric differences between the aircraft were known, ship 2 consistently had larger yawing moment asymmetries than ship 1.

  20. 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.

  1. Buffeting of NACA 0012 airfoil at high angle of attack

    NASA Astrophysics Data System (ADS)

    Zhou, Tong; Dowell, Earl

    2014-11-01

    Buffeting is a fluid instability caused by flow separation or shock wave oscillations in the flow around a bluff body. Typically there is a dominant frequency of these flow oscillations called Strouhal or buffeting frequency. In prior work several researchers at Duke University have noted the analogy between the classic Von Karman Vortex Street behind a bluff body and the flow oscillations that occur for flow around a NACA 0012 airfoil at sufficiently large angle of attack. Lock-in is found for certain combinations of airfoil oscillation (pitching motion) frequencies and amplitudes when the frequency of the airfoil motion is sufficiently close to the buffeting frequency. The goal of this paper is to explore the flow around a static and an oscillating airfoil at high angle of attack by developing a method for computing buffet response. Simulation results are compared with experimental data. Conditions for the onset of buffeting and lock-in of a NACA 0012 airfoil at high angle of attack are determined. Effects of several parameters on lift coefficient and flow response frequency are studied including Reynolds number, angle of attack and blockage ratio of the airfoil size to the wind tunnel dimensions. Also more detailed flow field characteristics are determined. For a static airfoil, a universal Strouhal number scaling has been found for angles of attack from 30° to 90°, where the flow around airfoil is fully separated. For an oscillating airfoil, conditions for lock-in are discussed. Differences between the lock-in case and the unlocked case are also studied. The second affiliation: Duke University.

  2. 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.

  3. 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.

  4. 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.

  5. A high angle of attack inviscid shuttle orbiter computation

    NASA Technical Reports Server (NTRS)

    Kleb, William L.; Weilmuenster, K. James

    1992-01-01

    As a preliminary step toward predicting the leeside thermal environment for winged reentry vehicles at flight conditions, a computational solution for the flow about the Shuttle Orbiter at wind tunnel conditions was made using a point-implicit, finite volume scheme known as the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). The surface pressures resulting from the computational solution are compared with wind tunnel data. The results indicate that the dominant inviscid flow features are being accurately predicted on the leeside of the Shuttle Orbiter at a moderately high angle of attack.

  6. Linear parameter-varying control of an F-16 aircraft at high angle of attack

    NASA Astrophysics Data System (ADS)

    Lu, Bei

    To improve the aircraft capability at high angle of attack and expand the flight envelope, advanced linear parameter-varying (LPV) control methodologies are studied in this thesis with particular applications of actuator saturation control and switching control. A standard two-step LPV antiwindup control scheme and a systematic switching LPV control approach are derived, and the advantages of LPV control techniques are demonstrated through nonlinear simulations of an F-16 longitudinal autopilot control system. The aerodynamic surface saturation is one of the major issues of flight control in the high angle of attack region. The incorporated unconventional actuators such as thrust vectoring can provide additional control power, but may have a potentially significant pay-off. The proposed LPV antiwindup control scheme is advantageous from the implementation standpoint because it can be thought of as an augmented control algorithm to the existing control system. Moreover, the synthesis condition for an antiwindup compensator is formulated as a linear matrix inequality (LMI) optimization problem and can be solved efficiently. By treating the input saturation as a sector bounded nonlinearity with a tight sector bound, the synthesized antiwindup compensator can stabilize the open-loop exponentially unstable systems. The LPV antiwindup control scheme is applied to the nonlinear F-16 longitudinal model, and compared with the thrust vectoring control approach. The simulation results show that the LPV antiwindup compensator improves the flight quality, and offers advantages over thrust vectoring in a high angle of attack region. For a thrust vectoring augmented aircraft, the actuator sets may be different at low and high angles of attack. Also due to different control objectives, a single controller may not exist over a wide angle of attack region. The proposed switching LPV control approach based on multiple parameter-dependent Lyapunov functions provides a flexible design

  7. 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.

  8. 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.

  9. Full-scale wind-tunnel investigation of the effects of wing leading-edge modifications on the high angle-of-attack aerodynamic characteristics of a low-wing general aviation airplane

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    The paper presents the results of a recent investigation to determine the effects of wing leading-edge modifications on the high angle-of-attack aerodynamic characteristics of a low-wing general aviation airplane in the Langley Full-Scale Wind Tunnel. The investigation was conducted to provide aerodynamic information for correlation and analysis of flight-test results obtained for the configuration. The wind-tunnel investigation consisted of force and moment measurements, wing pressure measurements, flow surveys, and flow visualization studies utilizing a tuft grid, smoke and nonintrusive mini-tufts which were illuminated by ultra-violet light. In addition to the tunnel scale system which measured overall forces and moments, the model was equipped with an auxiliary strain-gage balance within the left wing panel to measure lift and drag forces on the outer wing panel independent of the tunnel scale system. The leading-edge modifications studied included partial- and full-span leading-edge droop arrangements as well as leading-edge slats.

  10. Exploratory studies of actuated forebody strakes for yaw control at high angles of attack

    NASA Technical Reports Server (NTRS)

    Murri, Daniel G.; Rao, Dhanvada M.

    1987-01-01

    Wind-tunnel studies have been conducted to evaluate the potential of actuated forebody strakes to provide increased levels of yaw control on fighter aircraft at extremely high angles of attack where conventional aerodynamic controls are ineffective. The studies involved low-speed wind-tunnel tests of actuated forebody strake concepts applied to a generic fighter model and included circumferential pressure and flow visualization surveys on an isolated forebody. Results showed that the actuated forebody strake concept can provide high levels of yaw control over wide ranges of angle-of-attack and sideslip. However, when lifting surfaces were placed in close proximity to the forebody/strake combination, significant interaction effects occurred which reduced the yaw control effectiveness of the strakes and induced coupled rolling and pitching moments.

  11. 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

  12. 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

  13. Experimental investigation of the high angle of attack characteristics of a high performance general aviation aircraft

    NASA Technical Reports Server (NTRS)

    Meyer, H. F.; Yip, L. P.; Perkins, J. N.; Vess, R. J.

    1989-01-01

    Scale model wind tunnel tests and full scale flight tests have been conducted for the Questair Venture general aviation aircraft configuration in order to investigate its high angle-of-attack aerodynamics. Attention is given to the formulation of a wing leading-edge modification capable of enhancing stall departure characteristics. This modification, which involved both outboard wing leading-edge droop and two chordwise leading-edge slots, is found to produce almost no wing-rock tendency when tested on a wind tunnel free-to-roll apparatus; in the full-scale aircraft, the modification yielded gentle, controllable stall characteristics with little cruise and climb performance penalty.

  14. 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.

  15. A preliminary look at techniques used to obtain airdata from flight at high angles of attack

    NASA Technical Reports Server (NTRS)

    Moes, Timothy R.; Whitmore, Stephen A.

    1990-01-01

    Flight research at high angles of attack has posed new problems for airdata measurements. New sensors and techniques for measuring the standard airdata quantities of static pressure, dynamic pressure, angle of attack, and angle of sideslip were subsequently developed. The ongoing airdata research supporting NASA's F-18 high alpha research program is updated. Included are the techniques used and the preliminary results. The F-18 aircraft was flown with three research airdata systems: a standard airdata probe on the right wingtip, a self-aligning airdata probe on the left wingtip, and a flush airdata system on the nose cone. The primary research goal was to obtain steady-state calibrations for each airdata system up to an angle of attack of 50 deg. This goal was accomplished and preliminary accuracies of the three airdata systems were assessed and are presented. An effort to improve the fidelity of the airdata measurements during dynamic maneuvering is also discussed. This involved enhancement of the aerodynamic data with data obtained from linear accelerometers, rate gyros, and attitude gyros. Preliminary results of this technique are presented.

  16. 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.

  17. 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.

  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. Aerodynamic Characteristics at High and Low Subsonic Mach Numbers of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 Airfoil Sections at Angles of Attack from -2 Degrees to 30 Degrees

    NASA Technical Reports Server (NTRS)

    Critzos, Chris C.

    1954-01-01

    An investigation has been made in the Langley low-turbulence pressure tunnel of the aerodynamic characteristics of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 airfoil sections. Data were obtained at Mach numbers from 0.3 to that for tunnel choke, at angles of attack from -2deg to 30deg, and with the surface. of each airfoil smooth-and with roughness applied at the leading edge.The Reynolds numbers of the tests ranged from 0.8 x 10(exp 6) to 4.4 x 10(exp 6). The results are presented as variations of lift, drag, and quarter-chord pitching-moment coefficients with Mach number.

  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. Multiaxis aircraft control power from thrust vectoring at high angles of attack

    NASA Technical Reports Server (NTRS)

    Capone, F. J.; Mason, M. L.

    1986-01-01

    Extensive research programs conducted at the Langley Research Center have shown that thrust vectoring can be provided by multifunction (nonaxisymmetric) nozzles. Most of this research has been conducted on pitch vectoring at both static and forward flight conditions. Recent efforts have been aimed at evaluating yaw vectoring concepts at static (wind off) conditions. This paper summarizes results for three different twin-engine fighter configurations tested over a Mach number range of 0.15 to 2.47 at angles of attack up to 35 deg. The objective of these investigations was to determine the multiaxis control power characteristics provided by thrust vectoring. All three configurations employed two-dimensional convergent-divergent nozzles which provided pitch vectoring by differential deflection of the upper and lower nozzle divergent flaps. Three different means of yaw vectoring were tested: (1) a translating nozzle sidewall; (2) yaw flaps located in the nozzle sidewalls; and (3) canted nozzles. These investigations were conducted in the Langley 16-Foot Transonic Tunnel and the Lewis 10x10-Foot Supersonic Tunnel. Longitudinal and direction control power from thrust vectoring was greater than that provided by aerodynamic control effectors at low speed or at high angles of attack.

  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. Delayed detached eddy simulations of fighter aircraft at high angle of attack

    NASA Astrophysics Data System (ADS)

    Xu, Guoliang; Jiang, Xiong; Liu, Gang

    2016-05-01

    The massively separated flows over a realistic aircraft configuration at 40°, 50°, and 60° angles of attack are studied using the delayed detached eddy simulation (DDES). The calculations are carried out at experimental conditions corresponding to a mean aerodynamic chord-based Reynolds number of 8.93× 105 and Mach number of 0.088. The influence of the grid size is investigated using two grids, 20.0× 106 cells and 31.0× 106 cells. At the selected conditions, the lift, drag, and pitching moment from DDES predictions agree with the experimental data better than that from the Reynolds-averaged Navier-Stokes. The effect of angle of attack on the flow structure over the general aircraft is also studied, and it is found that the dominated frequency associated with the vortex shedding process decreases with increasing angle of attack.

  4. Delayed detached eddy simulations of fighter aircraft at high angle of attack

    NASA Astrophysics Data System (ADS)

    Xu, Guoliang; Jiang, Xiong; Liu, Gang

    2016-08-01

    The massively separated flows over a realistic aircraft configuration at 40°, 50°, and 60° angles of attack are studied using the delayed detached eddy simulation (DDES). The calculations are carried out at experimental conditions corresponding to a mean aerodynamic chord-based Reynolds number of 8.93× 105 and Mach number of 0.088. The influence of the grid size is investigated using two grids, 20.0× 106 cells and 31.0× 106 cells. At the selected conditions, the lift, drag, and pitching moment from DDES predictions agree with the experimental data better than that from the Reynolds-averaged Navier-Stokes. The effect of angle of attack on the flow structure over the general aircraft is also studied, and it is found that the dominated frequency associated with the vortex shedding process decreases with increasing angle of attack.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. Parameter identification for X-31A at high angles of attack

    NASA Technical Reports Server (NTRS)

    Weiss, S.; Rohlf, D.; Plaetschke, Ermin

    1994-01-01

    The U.S./German experimental aircraft X-31A was designed and constructed to demonstrate enhanced fighter maneuverability. Post-stall maneuvering is enabled by applying new technologies such as high angle of attack aerodynamics and flight control system integrated thrust vectoring. Two demonstrator aircraft have been built by the main contractors, Rockwell International and Deutsche Aerospace (formerly MBB). Flight testing started in October 1990 and before the end of 1992 both aircraft had accomplished a significant number of flights covering the entire AoA regime from about -5 to 70 deg. Throughout the envelope expansion, DLR Institute of Flight Mechanics conducted parameter identification (PID) to determine the aerodynamic parameters of the aircraft from flight test data and to compare the results to the predictions from the aerodynamic dataset (ADS). The application of system identification to high AoA / post-stall flight data raises some major problems, which are discussed in this paper. Results from both longitudinal and lateral-directional motion will be presented.

  10. 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.

  11. 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.

  12. 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.

  13. Optimum subsonic, high-angle-of-attack nacelles

    NASA Technical Reports Server (NTRS)

    Luidens, R. W.; Stockman, N. O.; Diedrich, J. H.

    1979-01-01

    The optimum design of nacelles that operate over a wide range of aerodynamic conditions and their inlets is described. For low speed operation the optimum internal surface velocity distributions and skin friction distributions are described for three categories of inlets: those with BLC, and those with blow in door slots and retractable slats. At cruise speed the effect of factors that reduce the nacelle external surface area and the local skin friction is illustrated. These factors are cruise Mach number, inlet throat size, fan-face Mach number, and nacelle contour. The interrelation of these cruise speed factors with the design requirements for good low speed performance is discussed.

  14. Investigation of Model Wake Blockage Effects at High Angles of Attack in Low-Speed Wind Tunnel

    NASA Astrophysics Data System (ADS)

    Shyu, Lih-Shyng; Chuang, Shu-Hao

    To improve the fidelity of measured aerodynamic characteristics at high angle of attack for modern jet fighters, this paper examines the model wake blockage effect. The wake blockage effect in a 2.2×3.1 m low-speed wind tunnel is investigated by analyzing drag and wall pressure measurements. Circular flat plates of different sizes are used to simulate a test model at high angles of attack. The present analysis results in simple formulas for corrections of model wake blockage effect. To verify the present correction formula, the NASA TP-1803 model is force-tested in the tunnel. The corrected test data agree well with the NASA TP-1803 data.

  15. 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.

  16. Exploratory investigation of deflectable forebody strakes for high angle of attack yaw control

    NASA Technical Reports Server (NTRS)

    Rao, D. M.; Murri, D. G.

    1986-01-01

    A deflectable strake concept was investigated on a conical forebody to evaluate its yaw control potential at high angles of attack. In exploratory low-speed tunnel tests using a generic delta wing fighter configuration, antisymmetrically deflected strakes provided useful levels of yaw power at angles of attack when the conventional rudder became totally degraded. Symmetrical strakes prevented side force development at high angles of attack, and provided pitch control through symmetrical deflection. The strake performance was sensitive to its circumferential position on the forebody due to varying interaction of strake vortices with the wing and vertical tail. The low Reynolds number results of this study provided a favorable initial validation of the concept, subject to verification in regard to scale effects.

  17. 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.

  18. 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.

  19. A Semigraphical Method of Applying Impact Theory to an Arbitrary Body to Obtain the Hypersonic Aerodynamic Characteristics at Angle of Attack and Sideslip

    NASA Technical Reports Server (NTRS)

    Jackson, Charlie M., Jr.

    1961-01-01

    A simple semigraphical method of applying impact theory to obtain the aerodynamic characteristics of an arbitrary body at combined angle of attack and sideslip is presented. The necessary equations are derived, a general procedure for application is outlined, and the effects of graphical errors and areas of application are discussed. One of the features of the present method is the requirement of only one graphical construction for any combination of angle of attack and sideslip. As an example application the present method is applied to a blunted elliptical cone in order to obtain the longitudinal aerodynamic characteristics at an angle of attack of 40 degrees and an angle of sideslip of 0 degrees.

  20. 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.

  1. 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.

  2. An aerodynamic design study of a series of lifting bodies at angles of attack from 10 to 53 degrees at Mach numbers from 2.30 to 4.62

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.; Torres, Abel O.

    1992-01-01

    The aerodynamic characteristics in the transition from high to low angles of attack at supersonic speeds have been experimentally and theoretically studied for a series of lifting bodies with various upper and lower surface camber designs. The configurations under consideration have a 75-degree swept delta planform with a rounded nose. Data obtained indicate that changes in the camber design cause some distinct changes in the aerodynamic characteristics that shoud be taken into account in the selection of a lifting body shape. The flat bottom designs with upper surface camber are found to provide greater drag for retardation at high angles of attack but are considerably out of trim longitudinally. The flat top designs with lower surface camber provide less drag at high angles of attack but can be more easily trimmed. Calculated results are found to be in good agreement with the experimental data.

  3. 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.

  4. Preliminary results from an airdata enhancement algorithm with application to high-angle-of-attack flight

    NASA Technical Reports Server (NTRS)

    Moes, Timothy R.; Whitmore, Stephen A.

    1991-01-01

    A technique was developed to improve the fidelity of airdata measurements during dynamic maneuvering. This technique is particularly useful for airdata measured during flight at high angular rates and high angles of attack. To support this research, flight tests using the F-18 high alpha research vehicle (HARV) were conducted at NASA Ames Research Center, Dryden Flight Research Facility. A Kalman filter was used to combine information from research airdata, linear accelerometers, angular rate gyros, and attitude gyros to determine better estimates of airdata quantities such as angle of attack, angle of sideslip, airspeed, and altitude. The state and observation equations used by the Kalman filter are briefly developed and it is shown how the state and measurement covariance matrices were determined from flight data. Flight data are used to show the results of the technique and these results are compared to an independent measurement source. This technique is applicable to both postflight and real-time processing of data.

  5. 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.

  6. 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.

  7. 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.

  8. Numerical simulation of vortical flow over an elliptical-body missile at high angles of attack

    NASA Technical Reports Server (NTRS)

    Newsome, R. N.; Adams, M. S.

    1986-01-01

    Numerical solutions to the Reynolds-averaged Navier-Stokes equations are given for the flow about an elliptical body missile (3:1 ellipse) at a Mach number of 2.5 and a unit Reynolds number of 6.56 x 10 to the 6th/m. At high angles of attack, the flow is dominated by large-scale free vortices which occur in the lee-side flow field due to crossflow boundary-layer separation. Emphasis is focused on the accurate prediction of the lee-side vortical flow. Solutions are presented for both symmetric and asymmetric (body rolled 45 deg) configurations at 10 deg and 20 deg angle of attack. The computed results are compared with experimental surface pressure coefficients and vapor-screen photographs. Excellent agreement is obtained in all cases.

  9. A system for testing airdata probes at high angles of attack using a ground vehicle

    NASA Technical Reports Server (NTRS)

    Geenen, Robert J.; Moulton, Bryan J.; Haering, Edward A., Jr.

    1991-01-01

    A system to calibrate airdata probes at angles of attack between 0 and 90 deg was developed and tested at the NASA Ames Dryden Flight Research Facility. This system used a test fixture mounted to the roof of a ground vehicle and included an onboard instrumentation and data acquisition system for measuring pressures and flow angles. The data could be easily transferred to the facility mainframe computer for further analysis. The system was designed to provide convenient and inexpensive airdata probe calibrations for projects which require airdata at high angles of attack, such as the F-18 High Alpha Research Program. This type of probe was tested to 90 deg angle of attack in a wind tunnel and using the ground vehicle system. The results of both tests are in close agreement. An airdata probe with a swiveling pilot-static tube was also calibrated with the ground vehicle system. This paper presents the results of these tests and gives a detailed description of the test system.

  10. 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.

  11. 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.

  12. Application of dynamical systems theory to the high angle of attack dynamics of the F-14

    NASA Technical Reports Server (NTRS)

    Jahnke, Craig C.; Culick, Fred E. C.

    1990-01-01

    Dynamical systems theory has been used to study the nonlinear dynamics of the F-14. An eight degree of freedom model that does not include the control system present in operational F-14s has been analyzed. The aerodynamic model, supplied by NASA, includes nonlinearities as functions of the angles of attack and sideslip, the rotation rate, and the elevator deflection. A continuation method has been used to calculate the steady states of the F-14 as continuous functions of the control surface deflections. Bifurcations of these steady states have been used to predict the onset of wing rock, spiral divergence, and jump phenomena which cause the aircraft to enter a spin. A simple feedback control system was designed to eliminate the wing rock and spiral divergence instabilities. The predictions were verified with numerical simulations.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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

  20. Development of a preliminary high-angle-of-attack nose-down pitch control requirement for high-performance aircraft

    NASA Technical Reports Server (NTRS)

    Nguyen, Luat T.; Foster, John V.

    1990-01-01

    The requirements for high-angle-of-attack nose-down pitch control for advanced high-performance aircraft are discussed. Background information on fundamental factors that influence and, to a large extent, determine the high angle-of-attack nose-down control requirement is briefly reviewed. Guidelines currently proposed by other sources which attempt to define these requirements are discussed. A requirement based on NASA analysis of the characteristics of existing relaxed static stability (RSS) aircraft is presented. This analysis could provide the basis for a preliminary design guide.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Flight determination of the aerodynamic stability and control characteristics of the NASA SGS 1-36 sailplane in the conventional and deep stall angles-of-attack of between -5 and 75 degrees

    NASA Technical Reports Server (NTRS)

    Mahdavi, F. A.; Sandlin, D. R.

    1984-01-01

    The flight test procedure and the preliminary analysis of the results obtained from twenty manned flights of the SGS 1-36 in the high angles of attack Deep Stall region are discussed. A comparison of the flight determined stability and control derivatives, those of the wind tunnel, and the estimated aerodynamic data is also presented. Furthermore, deep stall dynamics response of the SGS 1-36 is discussed briefly to explain some of the unexpected flight observations.

  6. In-flight flow visualization results from the X-29A aircraft at high angles of attack

    NASA Technical Reports Server (NTRS)

    Delfrate, John H.; Saltzman, John A.

    1992-01-01

    Flow visualization techniques were used on the X-29A aircraft at high angles of attack to study the vortical flow off the forebody and the surface flow on the wing and tail. The forebody vortex system was studied because asymmetries in the vortex system were suspected of inducing uncommanded yawing moments at zero sideslip. Smoke enabled visualization of the vortex system and correlation of its orientation with flight yawing moment data. Good agreement was found between vortex system asymmetries and the occurrence of yawing moments. Surface flow on the forward-swept wing of the X-29A was studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread outboard encompassing the full wing by 30 deg angle of attack. In general, the progression of the separated flow correlated well with subscale model lift data. Surface flow on the vertical tail was also studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread upward. The area of separated flow on the vertical tail at angles of attack greater than 20 deg correlated well with the marked decrease in aircraft directional stability.

  7. 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, Amandine; Bernal, Luis; Dimitriadis, Gregorios; Terrapon, Vincent

    2015-11-01

    The direct measurement of loads with force balance can become challenging when the forces are small or when the body is moving. An alternative is the use of Particle Image Velocimetry (PIV) velocity fields to indirectly obtain the aerodynamic coefficients. This can be done by the use of control volume approaches which lead to the integration of velocities, and other fields deriving from them, on a contour surrounding the studied body and its supporting surface. This work exposes and discusses results obtained with two different methods: the direct use of the integral formulation of the Navier-Stokes equations and the so-called Noca's method. The latter is a reformulation of the integral Navier-Stokes equations in order to get rid of the pressure. Results obtained using the two methods are compared and the influence of different parameters is discussed. The methods are applied to PIV data obtained from water channel testing for the flow around a 16:1 plate. Two cases are considered: a static plate at high angle of attack and a large amplitude imposed pitching motion. Two-dimensional PIV velocity fields are used to compute the aerodynamic forces. Direct measurements of dynamic loads are also carried out in order to assess the quality of the indirectly calculated coefficients.

  8. 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.

    1992-01-01

    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, is reviewed. 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.

  9. 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.

  10. Development of high-angle-of-attack nose-down pitch control margin design guidelines for combat aircraft

    NASA Technical Reports Server (NTRS)

    Ogburn, Marilyn E.; Foster, John V.

    1993-01-01

    A broad research program to identify maneuvering requirements for advanced fighters and the corresponding design criteria to aid in making critical design tradeoffs is being conducted under the NASA High-Angle-of-Attack Technology Program (HATP). As part of this activity, NASA and the U.S. Navy are conducting cooperative research to develop high-angle-of-attack control margin requirements. This paper will summarize the status of this program. Following some background information, the simulation study conducted to develop a set of preliminary guidelines for nose-down pitch control is reviewed, and the results of some very limited flight tests are described.

  11. 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).

  12. 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.

  13. 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.

  14. 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.)

  15. 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.

  16. Vortex force generation of an impulsively started wing at high angle of attack

    NASA Astrophysics Data System (ADS)

    Fu, Xiang; Wang, Fuxin; Liu, Hong; Qin, Suyang; Xiang, Yang

    2015-11-01

    A wing at high angle of attack (AoA) impulsively started from rest is a fundamental motion employed by insects during flight. Previous studies have almost solely focused on the lift enhancement by the leading-edge vortex (LEV). However, the influences of the starting vortex and secondary vortex on both the lift and drag generation have been less studied. In this paper, the vorticity fields for three AoAs of 45°, 58.5° and 72° are obtained numerically. The roles of the LEV, starting vortex and secondary vortex in generating the lift and drag are quantitatively studied using the vorticity moment theory. It is revealed that the LEV provides positive lift whereas the starting vortex and secondary vortex provide negative lift during the whole motion. The negative lift produced by the starting vortex or secondary vortex is not trivial and cannot be ignored. Regarding the drag, the LEV reduces the total drag whereas the starting vortex, the secondary vortex increases the total drag. As the AoA increases, the drag resulting from the starting vortex increases quickly and comprises almost all the total drag for the AoA of 72°. The relations between the motion of the vortical structures and the forces are also investigated. Financial support from the State Key Development Program of Basic Research of China (2014CB744802) is gratefully acknowledged.

  17. 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.

  18. 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.

  19. 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.

  20. 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 mode 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.

  1. 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.

  2. Forebody vortex management for yaw control at high angles of attack

    NASA Technical Reports Server (NTRS)

    Rao, D. M.; Moskovitz, C.; Murri, D. G.

    1986-01-01

    The yaw-control potential of deployable forebody strakes at angles of attack above the range of conventional rudder effectiveness has been investigated. The conformally-stored strakes when deployed force asymmetrical vortex shedding from the forebody, thereby generating a controlled yawing moment. The concept was explored through low-speed wind tunnel tests on a conical forebody in isolation and in a generic fighter configuration. Force and moment measurements on the complete model were supplemented with circumferential pressure and flow-visualization surveys on an isolated forebody, in order to gain insight into the vortex flow mechanisms resulting from forced asymmetrical separations and to quantify the obtainable yaw power at angles of attack to 80 deg. This preliminary, low-Reynolds-number study showed asymmetrically-deployed forebody strakes to have considerable yaw control potential, whose sensitivity to scale effects needs further investigation.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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. J.; Lackey, James B.

    1993-01-01

    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.

  9. 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.

  10. Large-scale wind-tunnel investigation of a close-coupled canard-delta-wing fighter model through high angles of attack

    NASA Technical Reports Server (NTRS)

    Stoll, F.; Koenig, D. G.

    1983-01-01

    Data obtained through very high angles of attack from a large-scale, subsonic wind-tunnel test of a close-coupled canard-delta-wing fighter model are analyzed. The canard delays wing leading-edge vortex breakdown, even for angles of attack at which the canard is completely stalled. A vortex-lattice method was applied which gave good predictions of lift and pitching moment up to an angle of attack of about 20 deg, where vortex-breakdown effects on performance become significant. Pitch-control inputs generally retain full effectiveness up to the angle of attack of maximum lift, beyond which, effectiveness drops off rapidly. A high-angle-of-attack prediction method gives good estimates of lift and drag for the completely stalled aircraft. Roll asymmetry observed at zero sideslip is apparently caused by an asymmetry in the model support structure.

  11. Vortex Effects for Canard-wing Configurations at High Angles of Attack in Subsonic Flow

    NASA Technical Reports Server (NTRS)

    Desilva, B. M. E.; Medan, R. T.

    1978-01-01

    A fully three-dimensional subsonic panel method that can handle arbitrary shed vortex wakes is used to compute the nonlinear forces and moments on simple canard-wing configurations. The lifting surfaces and wakes are represented by doublet panels. The Mangler-Smith theory is used to provide an initial estimate for the vortex sheet shed from the leading edge. The trailing-edge wake and the leading-edge wake downstream of the trailing edge are assumed to be straight and leave the trailing edge at an angle of alpha/2. Results indicate good agreement with experimental data up to 40 degs angle of attack.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Exploratory low-speed wind-tunnel study of concepts designed to improve aircraft stability and control at high angles of attack. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Hahne, D. E.

    1985-01-01

    A wind tunnel investigation of concepts to improve the high angle-of-attack stability and control characteristics of a high performance aircraft was conducted. The effect of vertical tail geometry on stability and the effectiveness of several conventional and unusual control concepts was determined. These results were obtained over a large angle-of-attack range. Vertical tail location, cant angle and leading edge sweep could influence both longitudinal and lateral-directional stability. The control concepts tested were found to be effective and to provide control into the post stall angle-of-attack region.

  18. 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.

  19. 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.

  20. 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.

  1. Impulsive Start of a Symmetric Airfoil at High Angle of Attack

    NASA Technical Reports Server (NTRS)

    Katz, Joseph; Yon, Steven; Rogers, Stuart E.

    1996-01-01

    The fluid dynamic phenomena following the impulsive start of a NACA 0015 airfoil were studied by using a time accurate solution of the incompressible laminar Navier-Stokes equations. Angle of attack was set at 10 deg to simulate steady-state poststall conditions at a Reynolds number of 1.2 x 10(exp 4). The calculation revealed that large initial lift values can be obtained, immediately following the impulsive start, when a trapped vortex develops above the airfoil. Before the buildup of this trapped vortex and immediately after the airfoil was set into motion, the fluid is attached to the airfoil's surface and flows around the trailing edge, demonstrating the delay in the buildup of the classical Kutta condition. The transient of this effect is quite short and is followed by an attached How event that leads to the trapped vortex that has a longer duration. The just described initial phenomenon eventually transits into a fully developed separated flow pattern identifiable by an alternating, periodic vortex shedding.

  2. Reynolds number effects on supersonic asymmetrical flows over a cone at high angle of attack

    NASA Technical Reports Server (NTRS)

    Thomas, J. L.

    1991-01-01

    The supersonic viscous flow over a 5-degree half-angle cone at an angle of attack of four times the cone half-angle is studied computationally using both the conical and the three-dimensional Navier-Stokes equations. The numerical solutions were obtained with an implicit, upwind-biased algorithm. Asymmetrical flowfields of the absolute-instability type are found using the conical-flow equations which agree with published results. However, the absolute instabilities of the originally symmetric flow found with the conical equations do not occur in the three-dimensional simulations, although spurious asymmetric three-dimensional flows for symmetric bodies arise if the grid resolution is insufficient in the nose region. The asymmetric flows computed with the three-dimensional equations are convective instabilities and are possible if the local Reynolds number exceeds a critical value and a fixed geometric asymmetry is imposed. A continuous range of asymmetries can be developed, depending on the size of the disturbance and the Reynolds number. As the Reynolds number is increased, the asymmetries demonstrate a bistable behavior at levels of side force consistent with those predicted using the conical equations. Below a certain critical Reynolds number, any flow asymmetries arising from a geometrical asymmetry are damped with increasing distance downstream from the geometrical asymmetry.

  3. 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.

  4. Numerical simulation by TVD schemes of complex shock reflections from airfoils at high angle of attack. [Total Variation Diminishing

    NASA Technical Reports Server (NTRS)

    Moon, Young J.; Yee, H. C.

    1987-01-01

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

  5. 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.

  6. Wind-tunnel investigation to determine the low speed yawing stability derivatives of a twin jet fighter model at high angles of attack

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Newsom, W. A., Jr.

    1974-01-01

    An investigation was conducted to determine the low-speed yawing stability derivatives of a twin-jet fighter airplane model at high angles of attack. Tests were performed in a low-speed tunnel utilizing variable-curvature walls to simulate pure yawing motion. The results of the study showed that at angles of attack below the stall the yawing derivatives were essentially independent of the yawing velocity and sideslip angle. However, at angles of attack above the stall some nonlinear variations were present and the derivatives were strongly dependent upon sideslip angle. The results also showed that the rolling moment due to yawing was primarily due to the wing-fuselage combination, and that at angles of attack below the stall both the vertical and horizontal tails produced significant contributions to the damping in yaw. Additionally, the tests showed that the use of the forced-oscillation data to represent the yawing stability derivatives is questionable, at high angles of attack, due to large effects arising from the acceleration in sideslip derivatives.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Preliminary results from a subsonic high-angle-of-attack flush airdata sensing (HI-FADS) system - Design, calibration, algorithm development, 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 vehicle. This paper discusses the airdata algorithm development and composite results expressed as airdata parameter estimates and describes the HI-FADS system hardware, calibration techniques, and algorithm development. An independent empirical verification was performed over a large portion of the subsonic flight envelope. Test points were obtained for Mach numbers from 0.15 to 0.94 and angles of attack from -8.0 to 55.0 deg. Angles of sideslip ranged from -15.0 to 15.0 deg, and test altitudes ranged from 18,000 to 40,000 ft. 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.

  13. Airfoil Lift with Changing Angle of Attack

    NASA Technical Reports Server (NTRS)

    Reid, Elliott G

    1927-01-01

    Tests have been made in the atmospheric wind tunnel of the National Advisory Committee for Aeronautics to determine the effects of pitching oscillations upon the lift of an airfoil. It has been found that the lift of an airfoil, while pitching, is usually less than that which would exist at the same angle of attack in the stationary condition, although exceptions may occur when the lift is small or if the angle of attack is being rapidly reduced. It is also shown that the behavior of a pitching airfoil may be qualitatively explained on the basis of accepted aerodynamic theory.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Experimental Investigations of the Initial Growth of Flow Asymmetries over a Slender Body of Revolution at High Angles of Attack

    NASA Astrophysics Data System (ADS)

    Zhu, Yiding; Yuan, Huijing; Lee, Cunbiao

    2015-11-01

    This paper describes an experimental investigation of the initial growth of flow asymmetries over a slender body of revolution at high angles of attack with natural and disturbed noses. Time-resolved particle image velocimetry (PIV) was used to investigate the flow field around the body. The experimental results show that initially different amplitudes of unsteady disturbances near the tip are established owing to the tip imperfections. These unsteady disturbances experience a super-exponential growth near the tip and continue to grow exponentially due to linear instabilities. Attachment of a piece to the tip brings a larger initial difference and extends the super-exponential growth region. Thus, the disturbance amplitudes and their differences are larger for the disturbed case than for the natural case before reaching the neutral point of linear instability. The amplified disturbances lead to different instability vortex strengths in the separated shear layers, which feed continuously into the two primary concentrated vortices. As a result, the primary vortex strengths differ which result in the initial vortex asymmetry. The experiment results demonstrate that the initial flow asymmetry arises from an asymmetric development of the boundary layer instability.

  19. Experimental investigations of the initial growth of flow asymmetries over a slender body of revolution at high angles of attack

    NASA Astrophysics Data System (ADS)

    Zhu, Yiding; Yuan, Huijing; Lee, Cunbiao

    2015-08-01

    This paper describes an experimental investigation of the initial growth of flow asymmetries over a slender body of revolution at high angles of attack with natural and disturbed noses. Time-resolved particle image velocimetry was used to investigate the flow field around the body. The experimental results show that initially different amplitudes of unsteady disturbances near the tip are established owing to the tip imperfections. These unsteady disturbances experience a super-exponential growth near the tip and continue to grow exponentially due to linear instabilities. Attachment of a piece to the tip brings a larger initial difference and extends the super-exponential growth region. Thus, the disturbance amplitudes and their differences are larger for the disturbed case than for the natural case before reaching the neutral point of linear instability. The amplified disturbances lead to different instability vortex strengths in the separated shear layers, which feed continuously into the two primary concentrated vortices. As a result, the primary vortex strengths differ, which result in the initial vortex asymmetry. The experiment results demonstrate that the initial flow asymmetry arises from an asymmetric development of the boundary layer instability.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 2: High-wing model C

    NASA Technical Reports Server (NTRS)

    Hultberg, R. S.; Chu, J.

    1980-01-01

    Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin g tunnel are presented in plotted form for a 1/6 scale, single engine, high wing, general aviation model. The configurations tested included the basic airplane and control deflections, wing leading edge devices, tail designs, and 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 a spin coefficient range from 0 to 0.9.

  5. 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.

  6. 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.

  7. Wall-modeled large eddy simulation of high-lift devices from low to post-stall angle of attacks

    NASA Astrophysics Data System (ADS)

    Bodart, Julien; Larsson, Johan; Moin, Parviz

    2013-11-01

    The flow around a McDonnell-Douglas 30P/30N multi-element airfoil at the flight Reynolds number of 9 million (based on chord) is computed using LES with an equilibrium wall-model with special treatment for transitional flows. Several different angles of attack are considered, up to and including stall, challenging the wall-model in several flow regimes. The maximum lift coefficient, which is generally difficult to predict with RANS approaches, is accurately predicted, as compared to experiments performed in the NASA LPT wind-tunnel. NASA grant: NNX11AI60A.

  8. Wind-tunnel research comparing lateral control devices, particularly at high angles of attack V : spoilers and ailerons on rectangular wings

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Shortal, Joseph A

    1933-01-01

    This report covers the fifth of a series of systematic investigations in which lateral control devices are compared with particular reference to their effectiveness at high angles of attack. The present report deals with tests of spoilers and ordinary ailerons on rectangular Clark y wing models. In an effort to obtain satisfactory control throughout the entire angle-of-attack range that can be maintained in flight, various spoilers were tested in combination with two sizes of previously tested ordinary ailerons - one of average proportions and the other short and wide. In addition, one large spoiler was tested alone. It was found that when ailerons and spoilers are used together the full effect of both is not obtained if the spoilers are located directly in front of the ailerons. With the proper combination of spoiler and aileron, however, it is possible to obtain satisfactory rolling control up to high angles of attack (15 degrees to 20 degrees), together with favorable yawing moments and small control forces. A moderate amount of rolling control with favorable yawing moments and small control forces was obtained with the large spoiler alone.

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

    NASA Technical Reports Server (NTRS)

    Johnson, J. L., Jr.; Grafton, S. B.; Yip, L. P.

    1980-01-01

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

  10. Navier-Stokes computations of the NREL airfoil using a {kappa}-{omega} turbulent model at high angles of attack

    SciTech Connect

    Yang, S.L.; Chang, Y.L.; Arici, O.

    1995-11-01

    This paper presents a two-dimensional numerical simulation of the turbulent flow fields for the NREL (National Renewable Energy Laboratory) S809 airfoil. The flow is modeled as steady, viscous, turbulent, and incompressible. The pseudo-compressible formulation is used for the time-averaged Navier-Stokes equations so that a time marching scheme developed for the compressible flow can be applied directly. The turbulent flow is simulated using Wilcox`s modified {kappa}-{omega} model to account for the low Reynolds number effects near a solid wall and the model`s sensitivity to the freestream conditions. The governing equations are solved by an implicit approximate-factorization scheme. To correctly model the convection terms in the mean-flow and turbulence model equations, the symmetric TVD (Total Variational Diminishing) scheme is incorporated. The methodology developed is then applied to analyze the NREL S809 airfoil at various angles of attack ({alpha}) from 1 to 45 degrees. The accuracy of the numerical results is compared with the available Delft wind tunnel test data. For comparison, two Eppler code results at low angles of attack are also included. Depending on the value of {alpha}, preliminary results show excellent to fairly good agreement with the experimental data. Directions for future work are also discussed.

  11. 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.

  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. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Wind-Tunnel Investigation at a Mach Number of 2.01 of the Aerodynamic Characteristics in Combined Angles of Attack and Sideslip of Several Hypersonic Missile Configurations with Various Canard Controls

    NASA Technical Reports Server (NTRS)

    Robinson, R. B.

    1958-01-01

    An investigation of the aerodynamic characteristics of several hypersonic missile configurations with various canard controls for an angle-of-attack range from 0 deg to about 28 deg at sideslip angles of about 0 deg and 4 deg at a Mach number of 2.01 has been made in the Langley 4- by 4-foot supersonic pressure tunnel. The configurations tested we re a body alone which had a ratio of length to diameter of 10, the b ody with a 10 deg flare, the body with cruciform fins of 5 deg or 15 deg apex angle, and a flare-stabilized rocket model with a modified Von Karman nose. Various canard surfaces for pitch control only were te sted on the body with the 10 deg flare and on the body with both sets of fins. The results indicated that the addition of a flared afterbody or cruciform fins produced configurations which were longitudinally and directionally stable. The body with 5 deg fins should be capable of producing higher normal accelerations than the flared body. A l l of the canard surfaces were effective longitudinal controls which produced net positive increments of normal force and pitching moments which progressively decreased with increasing angle of attack.

  18. BiGlobal linear stability analysis on low-Re flow past an airfoil at high angle of attack

    NASA Astrophysics Data System (ADS)

    Zhang, Wei; Samtaney, Ravi

    2016-04-01

    We perform BiGlobal linear stability analysis on flow past a NACA0012 airfoil at 16° angle of attack and Reynolds number ranging from 400 to 1000. The steady-state two-dimensional base flows are computed using a well-tested finite difference code in combination with the selective frequency damping method. The base flow is characterized by two asymmetric recirculation bubbles downstream of the airfoil whose streamwise extent and the maximum reverse flow velocity increase with the Reynolds number. The stability analysis of the flow past the airfoil is carried out under very small spanwise wavenumber β = 10-4 to approximate the two-dimensional perturbation, and medium and large spanwise wavenumbers (β = 1-8) to account for the three-dimensional perturbation. Numerical results reveal that under small spanwise wavenumber, there are at most two oscillatory unstable modes corresponding to the near wake and far wake instabilities; the growth rate and frequency of the perturbation agree well with the two-dimensional direct numerical simulation results under all Reynolds numbers. For a larger spanwise wavenumber β = 1, there is only one oscillatory unstable mode associated with the wake instability at Re = 400 and 600, while at Re = 800 and 1000 there are two oscillatory unstable modes for the near wake and far wake instabilities, and one stationary unstable mode for the monotonically growing perturbation within the recirculation bubble via the centrifugal instability mechanism. All the unstable modes are weakened or even suppressed as the spanwise wavenumber further increases, among which the stationary mode persists until β = 4.

  19. 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.

  20. 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.

  1. 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.

  2. Similarity for high angle-of-attack subsonic/transonic slender-body aerodynamics

    NASA Technical Reports Server (NTRS)

    Hemsch, M. J.

    1988-01-01

    The work of Sychev (1960) is extended to sharp-edged wings and smooth bodies. A simple fit to the correlations based on the Polhamus (1966) suction analogy is shown for several affine families of bodies with elliptical cross sections. The results suggest that simple models for nonlinear subsonic/transonic flows over not-so-slender wings and bodies can be obtained from a combination of well-known linear-theory models for the attached flow and slender-body concepts for the vortical flow.

  3. 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

  4. 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.

  5. 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.

  6. Investigation of the asymmetric aerodynamic characteristics of cylindrical bodies of revolution with variations in nose geometry and rotational orientation at angles of attack to 58 degrees and Mach numbers to 2

    NASA Technical Reports Server (NTRS)

    Kruse, R. L.; Keener, E. R.; Chapman, G. T.; Claser, G.

    1979-01-01

    Wind-tunnel tests were conducted to investigate the side forces and yawing moments that can occur at high angles of attack and zero sideslip for cylindrical bodies of revolution. Two bodies having several tangent ogive forebodies with fineness ratios of 0.5, 1.5, 2.5, and 3.5 were tested. The forebodies with fineness ratios of 2.5 and 3.5 had several bluntnesses. The cylindrical afterbodies had fineness ratios of 7 and 13. The model components - tip, forebody, and afterbody - were tested in various rotational positions about their axes of symmetry. Most of the tests were conducted at a Mach number of 0.25, a Reynolds number of 0.32 x 10 to the 6th power, and with the afterbody that had a fineness ratio of 7 and with selected forebodies. The effect of Mach number was determined with the afterbody that had a fineness ratio of 13 and with selected forebodies at mach numbers from 0.25 to 2 at Reynolds number = 0.32 X 10 to the 6th power. Maximum angle of attack was 58 deg.

  7. 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.

  8. Ultrasonic anemometer angle of attack errors under turbulent conditions

    NASA Astrophysics Data System (ADS)

    Nakai, T.

    2009-12-01

    Measurements of eddy fluxes are premised on the assumption that wind speeds are measured accurately by an ultrasonic anemometer. Recently, ultrasonic anemometers have been shown to suffer errors depending on the angle of attack, which is the angle between the wind vector and the horizontal. The correction of these errors resulted in general increases in eddy fluxes. However, since the check of the angle of attack dependent error was carried out in the wind tunnel experiment, which would be under the condition of nearly laminar flow, the applicability of this correction to the field data under turbulent conditions has been questioned. In this study, angle of attack dependencies of wind speeds measured by Gill Windmaster ultrasonic anemometers were assessed by field experiment over meadow, considered to be turbulent conditions. By using five identical anemometers, two pairs of systems were prepared: two anemometers for references and one between them for tilt. The dependencies of (co)sine responses of anemometers on angles of attack of 0 to -90 degrees in 10-degree steps and 45 degrees were checked, and clarified that the angle of attack dependent errors occur also under turbulent conditions, with results similar to the wind tunnel experiments. Sine responses of vertical wind speeds depended not only on vertical angle of attack but also on horizontal wind direction, which had not been considered in previous studies. For more robust correction, alternative calibration functions were obtained empirically so as to reasonably explain our field experimental results. Applying this new correction, eddy fluxes increased substantially even over meadow, which is somewhat aerodynamically smooth compared with forests or agricultural fields.

  9. Three-dimensional flows about simple components at angle of attack

    NASA Technical Reports Server (NTRS)

    Peake, D. J.; Tobak, M.

    1982-01-01

    The structures of three dimensional separated flow about some chosen aerodynamic components at angle of attack are synthesized, holding strictly to the notion that streamlines in the external flow (viscous plus inviscid) and skin friction lines on the body surface may be considered as trajectories having properties consistent with those of continuous vector fields. Singular points in the fields are of limited number and are classified as simple nodes and saddles. Analogous flow structures at high angles of attack about blunt and pointed bodies, straight and swept wings, etc., are discussed, highlighting the formation of spiral nodes (foci) in the pattern of the skin friction lines. How local and global three dimensional separation lines originate and form is addressed, and the characteristics of both symmetric and asymmetric leeward wakes are described.

  10. 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.

  11. Experimental aerodynamic characteristics for slender bodies with thin wings and tail 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.; Nelson, E. R.

    1976-01-01

    An experimental investigation was conducted by wind tunnel to measure the static aerodynamic characteristics for bodies of circular and elliptic cross section with various thin flat plate wings and a thin tail consisting of horizontal and vertical parts. The wings had aspect ratios of 4 and taper ratios of about 0, 0.25, and 0.5. Two additional wings, which had taper ratios near 0.25 and aspect ratios of about 3 and 5, were also tested in combination with the bodies and tail. All wings had about the same planform area. The exposed area of the horizontal portion of the tail was about 33 to 36 percent of the exposed area of the wings. The exposed area of the vertical tail fin was about 22 to 24 percent of the exposed area of the wings. The elliptic body, with an a/b = 2 cross section, had the same length and axial distribution of cross sectional area as the circular body. The circular body had a cylindrical aftersection of fineness ratio 7, and it was tested with the wings and tail in combination with tangent ogive noses that had fineness ratios of 2.5, 3.0, 3.5, and 5.0. In addition, an ogive nose with a rounded tip and an ogive nose with two different nose strake arrangements were used. Nineteen configuration combinations were tested at Mach numbers of 0.6, 0.9, 1.5, and 2.0 at angles of attack from 0 to 58 deg. The Reynolds numbers, based on body base diameter, were about 4.3 X 100,000.

  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: 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Aerothermal tests of a 12.5 percent cone at Mach 6.7 for various Reynolds numbers, angles of attack and nose shapes. [conducted in Langley 8-foot high temperature tunnel

    NASA Technical Reports Server (NTRS)

    Nowak, R. J.; Albertson, C. W.; Hunt, L. R.

    1984-01-01

    The effects of free-stream unit Reynolds number, angle of attack, and nose shape on the aerothermal environment of a 3-ft basediameter, 12.5 deg half-angle cone were investigated in the Langley 8-foot high temperature tunnel at Mach 6.7. The average total temperature was 3300 R, the freestream unit Reynolds number ranged from 400,000 to 1,400,000 per foot, and the angle of attack ranged from 0 deg to 10 deg. Three nose configurations were tested on the cone: a 3-in-radius tip, a 1-in-radius tip on an ogive frustum, and a sharp tip on an ogive frustum. Surface-pressure and cold-wall heating-rate distributions were obtained for laminar, transitional temperature in the shock layer were obtained. The location of the start of transition moved forward both on windward and leeward sides with increasing free-stream Reynolds numbers, increasing angle of attack, and decreasing nose bluntness.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. An optical angle of attack sensor

    NASA Astrophysics Data System (ADS)

    McDevitt, T. Kevin; Owen, F. Kevin

    A major source of transonic and supersonic wind-tunnel test data uncertainty is due to angle of attack (alpha) measurement errors caused by unknown sting and balance deflections under load. A novel laser-based instrument has been developed to enable continuous time-dependent alpha measurements to be made without signal dropout. Detectors capable of 0.01-deg resolution over an 18-deg range and 0.03-deg resolution over a 44-deg range with time-dependent outputs of 60 Hz have been developed. This capability is sufficient to provide accurate real-time alpha information for correlation with model balance measurements during transport and fighter model testing. Proof-of-concept experiments, along with the results of recent measurements conducted at the NASA Ames 9 x 7-ft supersonic wind tunnel, are presented. Experiments were also conducted to determine the reliable range, sensitivity, and long-term stability of the instrument.

  3. 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.

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

    NASA Technical Reports Server (NTRS)

    Gee, Ken; Cummings, Russell M.; Schiff, Lewis B.

    1990-01-01

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

  5. Yaw Control At High Angles Of Attack

    NASA Technical Reports Server (NTRS)

    Murri, Daniel G.; Rao, Dhanvada M.

    1988-01-01

    Hinged, conformal forebody strakes provide control when rudders become ineffective. Device consists of symmetric pair of longitudinally hinged strakes designed to fold completely into forebody contour. Strakes rotate individually out into external flow. Asymmetric flow produced by deployed strake generates sideward force causing aircraft to yaw.

  6. Aerodynamic Characteristics of Airfoils at High Speeds

    NASA Technical Reports Server (NTRS)

    Briggs, L J; Hull, G F; Dryden, H L

    1925-01-01

    This report deals with an experimental investigation of the aerodynamical characteristics of airfoils at high speeds. Lift, drag, and center of pressure measurements were made on six airfoils of the type used by the air service in propeller design, at speeds ranging from 550 to 1,000 feet per second. The results show a definite limit to the speed at which airfoils may efficiently be used to produce lift, the lift coefficient decreasing and the drag coefficient increasing as the speed approaches the speed of sound. The change in lift coefficient is large for thick airfoil sections (camber ratio 0.14 to 0.20) and for high angles of attack. The change is not marked for thin sections (camber ratio 0.10) at low angles of attack, for the speed range employed. At high speeds the center of pressure moves back toward the trailing edge of the airfoil as the speed increases. The results indicate that the use of tip speeds approaching the speed of sound for propellers of customary design involves a serious loss in efficiency.

  7. Sensor for measuring instantaneous angle of attack of helicopter blades

    NASA Technical Reports Server (NTRS)

    Barna, P. S.

    1980-01-01

    Systematic investigations were performed on a variety of probes to determine their potential for possible application as sensors attached to helicopter blades to measure both the instantaneous angle of attack as well as the dynamic head during actual flight operations. After some preliminary considerations a sensor of essentially spherical shape, about 30 mm in diameter, was designed. The sensor was provided with three pressure ports, and it housed two pressure transducers required for sensing the prevailing pressures acting outside on the surface. The sensors were subsequently tested in the laboratory under a variety of flow conditions to determine their aerodynamic characteristics. Two series of tests were performed: in the first series the sensor was fixed in space while exposed to steady uniform flow, while in the second series the sensor was made to oscillate, thus simulating the cyclic pitch change of the helicopter blades. While the cyclic pitch frequencies were of about the same magnitude as encountered in flight, the flow velocities during tests fell well below those experienced in a rotating blade. The tests showed that the sensors performed satisfactorily under low subsonic flow conditions with frequencies not exceeding five Hz.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Impact of high-alpha aerodynamics on dynamic stability parameters of aircraft and missiles

    NASA Technical Reports Server (NTRS)

    Malcolm, G. N.

    1981-01-01

    The aerodynamic phenomena associated with high angles of attack and their effects on the dynamic stability characteristics of airplane and missile configurations are examined. Information on dynamic effects is limited. Steady flow phenomena and their effects on the forces and moments are reviewed. The effects of asymmetric vortices and of vortex bursting on the dynamic response of flight vehicles are reviewed with respect to their influence on: (1) nonlinearity of aerodynamic coefficients with attitude, rates, and accelerations; (2) cross coupling between longitudinal and lateral directional models of motion; (3) time dependence and hysteresis effects; (4) configuration dependencey; and (5) mathematical modeling of the aerodynamics.

  15. Determination of an angle of attack sensor correction for a general aviation airplane at large angles of attack as determined from wind tunnel and flight tests

    NASA Technical Reports Server (NTRS)

    Moul, T. M.; Taylor, L. W., Jr.

    1980-01-01

    A comprehensive investigation into the flow correction for an angle of attack sensor mounted ahead of the wing tip of a general aviation research airplane has been conducted at the Langley Research Center. This correction has been determined in wind tunnels using a full-scale model up to angles of attack of 45 deg and a 1/5-scale model up to 80 deg angle of attack. The flow correction has also been obtained in flight by using a standard technique at low angles of attack and in spinning flight at larger angles of attack, by using both a simple approximate technique and a parameter estimation technique. The results show the correction is significant, reaching 10 deg at a measured angle of attack of about 90 deg. The flow correction was sensitive to the angle of sideslip at measured angles of attack greater than 60 deg and was not influenced by wing leading-edge modifications or aileron deflections.

  16. Aerodynamic characteristics of a propeller powered high lift semispan wing

    NASA Technical Reports Server (NTRS)

    Takallu, M. A.; Gentry, G. L., Jr.

    1992-01-01

    An experimental investigation was conducted on the engine/airframe integration aerodynamics for potential high-lift aircraft configurations. The model consisted of a semispan wing with a double-isolated flap system and a Krueger leading edge device. The advanced propeller and the powered nacelle were tested and aerodynamic characteristics of the combined system are presented. It was found that the lift coefficient of the powered wing could be increased by the propeller slipstream when the rotational speed was increased and high-lift devices were deployed. Moving the nacelle/propeller closer to the wing in the vertical direction indicated higher lift augmentation than a shift in the longitudinal direction. A pitch-down nacelle inclination enhanced the lift performance of the system much better than vertical and horizontal variation of the nacelle locations and showed that the powered wing can sustain higher angles of attack near maximum lift performance.

  17. High-attitude low-speed static aerodynamic characteristics of an F-4D fighter airplane model with leading edge slats

    NASA Technical Reports Server (NTRS)

    Monfort, J. C.; Whitcomb, W. M.

    1975-01-01

    An investigation was conducted to determine the effects of two-position leading edge slats on the low speed aerodynamic characteristics of a swept wing twin-jet supersonic fighter airplane model at high angle of attack and various Reynolds numbers. The investigation was performed at a Mach number of 0.20 over a range of angle of attack from 19 deg to 90 deg and angles of slideslip from -10 deg to 30 deg and Reynolds numbers from 1.97 to 13.12 million per meter.

  18. 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.

  19. 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.

  20. 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.

  1. Potential flow past axisymmetric bodies at angle of attack

    NASA Technical Reports Server (NTRS)

    Kuhlman, J. M.; Shu, J.-Y.

    1984-01-01

    The Karamcheti (1966) suggestion concerning the use of higher order singularity techniques has been developed for the calculation of incompressible flow past an axisymmetric body at angle of attack. Attention is given to the results of a convergence study using this axial singularity method, where solution accuracy has been investigated for ellipsoids of slenderness ratio in the 1-10 range for both axial and inclined flow. Effects of singularity type, element number and size distribution, and singularity line inset distance, are noted, and a paneling scheme is developed which yields accurate results for the class of axisymmetric bodies having continuous body slopes with discontinuous curvature jumps.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Low-Speed Tests of Semispan-Wing Models at Angles of Attack from 0 to 180 degrees

    NASA Technical Reports Server (NTRS)

    Koenig, David G.

    1959-01-01

    Semispan-wing models were tested at angles of attack from 0 to 180 deg at low subsonic speeds. Eight plan forms were considered, both swept and unswept with aspect ratios ranging from 2 to 6. Except for a delta-wing model of aspect ratio 2. all models had a taper ratio of 0.5 and an NACA 64AO10 airfoil section. The delta-wing model had an NACA 0005 (modified) airfoil section. With two exceptions, the models were tested both with and without a full-span trailing-edge flap deflected 25 deg. The Reynolds numbers based on the mean aerodynamic chord were between 1.5 and 2.2 million. Lift, drag, and pitching-moment coefficients are presented as functions of angle of attack. Approximate corrections for the effects of blockage were applied to the data.

  8. Tactical missile aerodynamics

    NASA Technical Reports Server (NTRS)

    Hemsch, Michael J. (Editor); Nielsen, Jack N. (Editor)

    1986-01-01

    The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

  9. Flight Investigation of a Roll-stabilized Missile Configuration at Varying Angles of Attack at Mach Numbers Between 0.8 and 1.79

    NASA Technical Reports Server (NTRS)

    Zarovsky, Jacob; Gardiner, Robert A

    1957-01-01

    A missile research model was flown at supersonic speed to determine the quality of automatic roll stabilization at varying angles of attack. Aerodynamic rolling and pitching derivatives were determined from the flight record. It was concluded that the combination of the gyro-actuated automatic pilot with wing-tip ailerons provided adequate roll stabilization under conditions encountered in flight.

  10. 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.

  11. Rotary balance data for a single engine general aviation design having a high aspect-ratio canard for an angle-of-attack range of 30 deg to 90 deg

    NASA Technical Reports Server (NTRS)

    Mulcay, W. J.; Rose, R.

    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. The configurations tested included the basic airplane, various control deflections, two canard locations, and wing leading edge modifications, as well as airplane components.

  12. 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

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Effect of angle of attack on rotor trailing-edge noise

    NASA Astrophysics Data System (ADS)

    Chou, S.-T.; George, A. R.

    1984-12-01

    Previous analyses of boundary layer trailing edge noise for large rotors have used zero blade angle of attack as input data. Attention is presently given to the important effects of blade angle of attack changes on rotor trailing edge noise in the case of a UH-1 helicopter. The primary effect is in the low to mid-frequency range, where noise level increases with angle of attack.

  18. 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.

  19. F-15 rotary balance data for an angle-of-attack range of 8 deg to 90 deg

    NASA Technical Reports Server (NTRS)

    Barnhart, B.

    1982-01-01

    Aerodynamic characteristics obtained in a rotational flow environment, utilizing a rotary balance are presented in plotted form for a 1/12 scale F-15 airplane model. The configurations tested included the buildup of airplane components and the basic airplane with various control deflections. Data are presented for all configurations without analysis for an angle of attack range of 8 to 90 deg, and clockwise and counterclockwise rotations covering an omega b/2V range from 0 to 0.4. Selected configurations are presented over an extended omega b/2V range from 0 to 0.9.

  20. Trailing Edge Noise-- Effect of Angle of Attack

    NASA Astrophysics Data System (ADS)

    Goody, Michael

    2003-03-01

    The broadband sound radiated by turbulent pressures that are convected past a sharp edge (i.e. trailing edge) can be estimated in three basic steps: (1) estimate the mean flow field and boundary layer development, (2) use this information to estimate the turbulent pressure fluctuations at the edge, and (3) use a Greens function to transform the convected pressure into the scattered pressure that is radiated as broadband sound. Moderately changing the angle of attack of a hydrofoil modifies the first two steps of this estimation process. Several methods, of varying level of complexity (and cost), for handling these modifications are presented for a NACA 0017 hydrofoil in water. The mean flow and boundary layer development are estimated (a) by assuming that the airfoil behaves as a flat plate and using standard boundary layer empirical formulas, (b) by using a panel method for the mean flow field and a simple boundary layer code to estimate the boundary layer development, and (c) by performing a RANS simulation of the flow field and boundary layer development. The turbulent pressure fluctuations are estimated using (a) an empirical model and (b) a statistical model. The performance of each method is discussed.

  1. 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.

  2. 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.

  3. 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.

  4. Inviscid Flow Computations of the Shuttle Orbiter for Mach 10 and 15 and Angle of Attack 40 to 60 Degrees

    NASA Technical Reports Server (NTRS)

    Prabhu, Ramadas K.; Sutton, Kenneth (Technical Monitor)

    2001-01-01

    This report documents the results of a computational study done to compute the inviscid longitudinal aerodynamic characteristics of the Space Shuttle Orbiter for Mach numbers 10 and 15 at angles of attack of 40, 50, 55, and 60 degrees. These computations were done to provide limited aerodynamic data in support of the Orbiter contingency abort task. The Orbiter had all the control surfaces in the undeflected position. The unstructured grid software FELISA was used for these computations with the equilibrium air option. Normal and axial force coefficients and pitching moment coefficients were computed. The hinge moment coefficients of the body flap and the inboard and outboard elevons were also computed. These results were compared with Orbiter Air Data Book (OADB) data and those computed using GASP. The comparison with the GASP results showed very good agreement in Cm and Ca at all the points. The computed axial force coefficients were smaller than those computed by GASP. There were noticeable differences between the present results and those in the OADB at angles of attack greater than 50 degrees.

  5. 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

  6. 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.

  7. 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.

  8. 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.

  9. Design of a two-element airfoil in a range of angles of attack

    NASA Astrophysics Data System (ADS)

    Abzalilov, D. F.

    2008-11-01

    A numerical-analytical solution of an inverse boundary-value problem of aerohydrodynamics is obtained for a two-element airfoil in the full formulation, based on the velocity distribution defined on the sought airfoil contours in a range of angles of attack. It is demonstrated that flow separation does not occur in the entire range considered for a specified non-separated velocity distribution on the upper surfaces at the maximum angle of attack and on the lower surface at the minimum angle of attack. An example of constructing a sectional airfoil is given; verification of the results obtained is performed with the use of the Fluent software package.

  10. Hypersonic crossflow instability and transition on a circular cone at angle of attack

    NASA Astrophysics Data System (ADS)

    Ward, Christopher A. C.

    To investigate the effect of Reynolds number, tunnel noise and surface roughness on the crossflow instability, a 7° half-angle cone was tested at 6° angle of attack with temperature-sensitive paint. All tests were performed in the Boeing/AFOSR Mach-6 Quiet Tunnel. The temperature-sensitive paint was successfully used to qualitatively visualize boundary-layer transition and stationary crossflow vortices. Tunnel noise was found to have a significant effect on the crossflow instability, promoting transition when the tunnel noise was high. Transition was never realized under fully quiet flow. Increasing Reynolds number also proved to have an effect on the crossflow instability. Under quiet flow, an increase in Reynolds number created larger amplitude stationary vortices, while under noisy flow the increase in Reynolds number caused transition to move upstream. The effect of surface roughness on the crossflow instability was also studied. This surface roughness effect was studied in two ways: by varying the frustum roughness and by varying the nosetip roughness. The frustum roughness was varied by altering the finish of the temperature-sensitive paint. The nosetip roughness was varied by polishing the stainless steel nosetip. For both cases, it was found that the magnitude and the azimuthal variations of the forward-facing step at the nosetip-frustum junction was most likely dominating the formation of the streamwise vorticity. Thus, before the effect of distributed surface roughness can be studied, the forward-facing step needs to be rendered negligible. Preliminary tests were also performed with the 7° half-angle cone placed at 0° angle of attack in order to validate a procedure for obtaining quantitative heat transfer from temperature-sensitive paint. However, this effort is not yet successful.