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Sample records for 60-deg delta wing

  1. An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

    NASA Technical Reports Server (NTRS)

    Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

    1983-01-01

    An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

  2. Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing

    NASA Technical Reports Server (NTRS)

    Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

    1983-01-01

    Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.

  3. Subsonic balance and pressure investigation of a 60 deg delta wing with leading edge devices

    NASA Technical Reports Server (NTRS)

    Tingas, S. A.; Rao, D. M.

    1982-01-01

    Low supersonic wave drag makes the thin highly swept delta wing the logical choice for use on aircraft designed for supersonic cruise. However, the high-lift maneuver capability of the aircraft is limited by severe induced-drag penalties attributed to loss of potential flow leading-edge suction. This drag increase may be alleviated through leading-edge flow control to recover lost aerodynamic thrust through either retention of attached leading-edge flow to higher angles of attack or exploitation of the increased suction potential of separation-induced vortex flow. A low-speed wind-tunnel investigation was undertaken to examine the high-lift devices such as fences, chordwise slots, pylon vortex generators, leading-edge vortex flaps, and sharp leading-edge extensions. The devices were tested individually and in combinations in an attempt to improve high-alpha drag performance with a minimum of low-alpha drag penalty. This report presents an analysis of the force, moment, and static pressure data obtained in angles of attack up to 23 deg, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter, respectively. The results indicate that all the devices produced drag and longitudinal/lateral stability improvements at high lift with, in most cases, minor drag penalties at low angles of attack.

  4. Subsonic balance and pressure investigation of a 60-deg delta wing with leading-edge devices (data report)

    NASA Technical Reports Server (NTRS)

    Rao, D. M.; Tingas, S. A.

    1981-01-01

    The drag reduction potential of leading edge devices on a 60 degree delta wing at high lift was examined. Geometric variations of fences, chordwise slots, pylon type vortex generators, leading edge vortex flaps, and sharp leading edge extensions were tested individually and in specific combinations to improve high-alpha drag performance with a minimum of low-alpha drag penalty. The force, moment, and surface static pressure data for angles of attack up to 23 degrees, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter are documented.

  5. Low-Speed Investigation of the Effects of Frequency and Amplitude of Oscillation in Sideslip on the Lateral Stability Derivatives of a 60 deg Delta Wing, a 45 deg Sweptback Wing and an Unswept Wing

    NASA Technical Reports Server (NTRS)

    Lichtenstein, Jacob H.; Williams, James L.

    1961-01-01

    A low-speed investigation has been conducted in the Langley stability tunnel to study the effects of frequency and amplitude of sideslipping motion on the lateral stability derivatives of a 60 deg. delta wing, a 45 deg. sweptback wing, and an unswept wing. The investigation was made for values of the reduced-frequency parameter of 0.066 and 0.218 and for a range of amplitudes from +/- 2 to +/- 6 deg. The results of the investigation indicated that increasing the frequency of the oscillation generally produced an appreciable change in magnitude of the lateral oscillatory stability derivatives in the higher angle-of-attack range. This effect was greatest for the 60 deg. delta wing and smallest for the unswept wing and generally resulted in a more linear variation of these derivatives with angle of attack. For the relatively high frequency at which the amplitude was varied, there appeared to be little effect on the measured derivatives as a result of the change in amplitude of the oscillation.

  6. Heat-Transfer Measurements on the Apexes of Two 60 deg Sweptback Delta Wings (Panel Semiapex Angle of 30 deg) Having 0 deg and 45 deg Dihedral at a Mach Number of 4.95

    NASA Technical Reports Server (NTRS)

    Gunn, Charles R.

    1961-01-01

    An experimental investigation was conducted to evaluate the heat-transfer rates at the apex of two 60 degree sweptback delta wings (panel semi-apex angle of 30 degrees) having cylindrical leading edges and 0 degrees and 45 degree positive dihedral. The models tested might correspond to the first several feet of a hypersonic reentry vehicle. The tests were conducted at a Mach number of 4.95 and a stagnation temperature of 400 F. nominal test-section unit Reynolds numbers varied from 2 x 10(exp 6) to 12 x 10(exp 6) per foot. The results of the investigation indicated that the laminar heat-transfer distributions (ratio of local to stagnation-line heating rate) about the models normal to the leading edges were in close agreement with two-dimensional blunt-body theory. The three-dimensional stagnation point heat-transfer rate on the 0 degree dihedral model was in excellent agreement with theory and the stagnation-line heat transfer on the straight portion of the leading edge of both models approached a constant level 12 percent above the theoretical stagnation-line level on an isolated swept infinite cylinder. When the heating rates on the 45 degree dihedral model (planform sweep of 69.3 degree) were compared with those on the 0 degree dihedral model (planform sweep of 60 degrees) at equal angles of attack and equal lifts greater than zero, the stagnation-line heating rates on the 45 degrees dihedral model were, in general, considerably lower as a result of the difference in effective sweeps of the leading edges. On the wing panels inboard from the stagnation lines, the differences in heating were very small. The stagnation-line heat-transfer variation with angle of attack, the shift in stagnation-line location, and the reduction in stagnation-line heat transfer resulting from the increase in effective sweep when positive dihedral is incorporated into a constant-panel 0 degree dihedral wing, all agreed with the results of a theoretical study made of highly swept delta

  7. A low speed wind tunnel investigation of Reynolds number effects on a 60-deg swept wing configuration with leading and trailing edge flaps

    NASA Technical Reports Server (NTRS)

    Rao, Dhanvada M.; Hoffler, Keith D.

    1988-01-01

    A low-speed wind tunnel test was performed to investigate Reynolds number effects on the aerodynamic characteristics of a supersonic cruise wing concept model with a 60-deg swept wing incorporating leading-edge and trailing-edge flap deflections. The Reynolds number ranged from 0.3 to 1.6 x 10 to the 6th, and corresponding Mach numbers from .05 to 0.3. The objective was to define a threshold Reynolds number above which the flap aerodynamics basically remained unchanged, and also to generate a data base useful for validating theoretical predictions for the Reynolds number effects on flap performance. This report documents the test procedures used and the basic data acquired in the investigation.

  8. Low-speed wind tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 60 deg

    NASA Technical Reports Server (NTRS)

    Moul, Thomas M.; Fears, Scott P.; Ross, Holly M.; Foster, John V.

    1995-01-01

    A wind tunnel investigation was conducted in the Langley 12-Foot Low-Speed Wind Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 60 deg, and all the trailing-edge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved pitching-moment characteristics and lateral stability and had three sets of trailing-edge flaps that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Top bodies of three widths and twin vertical tails of various sizes and locations were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced radar cross section and good flight dynamic characteristics.

  9. Wing loading on a 60 degree delta wing with vortex flaps

    NASA Technical Reports Server (NTRS)

    Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

    1983-01-01

    Wind tunnel tests were conducted on a 60 deg delta wing with three vortex flap designs to determine pressure distributions over the wing and flap. The results showed that an optimum vortex flap design depends on proper definition of the vortex flap deflection angle. They also revealed that flap thickness plays an important role in the behavior of the vortex flow over the flap and wing and can have a substantial effect on wing and flap pressure loading. Design codes which fail to account for thickness may result in a much less than optimum flap and deprive the designer of an important tool in designing an effective flap with optimum loading.

  10. Lift force of delta wings

    SciTech Connect

    Lee, M.; Ho, Chihming )

    1990-09-01

    On a delta wing, the separation vortices can be stationary due to the balance of the vorticity surface flux and the axial convection along the swept leading edge. These stationary vortices keep the wing from losing lift. A highly swept delta wing reaches the maximum lift at an angle of attack of about 40, which is more than twice as high as that of a two-dimensional airfoil. In this paper, the experimental results of lift forces for delta wings are reviewed from the perspective of fundamental vorticity balance. The effects of different operational and geometrical parameters on the performance of delta wings are surveyed.

  11. Supersonic aerodynamics of delta wings

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.

    1988-01-01

    Through the empirical correlation of experimental data and theoretical analysis, a set of graphs has been developed which summarize the inviscid aerodynamics of delta wings at supersonic speeds. The various graphs which detail the aerodynamic performance of delta wings at both zero-lift and lifting conditions were then employed to define a preliminary wing design approach in which both the low-lift and high-lift design criteria were combined to define a feasible design space.

  12. Self-induced wing rock of slender delta wings

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

    As part of a research program aimed at exploring basic mechanisms that cause wing rock in combat aircraft, an investigation was conducted to study the aerodynamic factors which cause the low-speed wing rock exhibited by slender delta wings. A flat-plate delta wing with 80 deg leading-edge sweep was subjected to conventional static-force tests and dynamic wind-tunnel experiments which included forced-oscillation, rotary, and free-to-roll tests. In addition, visualization of the flow phenomena involved was obtained by observing tuft patterns and using a helium-bubble technique. This paper summarizes the results of this study. Fundamental information is presented on the aerodynamic mechanisms that cause the wing rock and the problem of mathematically modeling the aerodynamics and motions is discussed.

  13. Similitude relations for buffet and wing rock on delta wings

    NASA Astrophysics Data System (ADS)

    Mabey, D. G.

    1997-08-01

    Vortex flow phenomena at high angles of incidence are of great interest to the designers of advanced combat aircraft. The steady phenomena (such as steady lift and pitching moments) are understood fairly well, whereas the unsteady phenomena are still uncertain. This paper addresses two important unsteady phenomena on delta wings. With regard to the frequency parameter of the quasi-periodic excitation caused by vortex bursting, a new correlation is established covering a range of sweep back from 60 to 75°. With regard to the much lower frequency parameter of limit-cycle rigid-body wing-rock, a new experiment shows conclusively that although the motion is non-linear, the frequency parameter can be predicted by quasi-steady theory. As a consequence, for a given sweep angle, the frequency parameter is inversely proportional to the square root of the inertia in roll. This is an important observation when attempting to extrapolate from model tests in wind tunnels to predict the wing-rock characteristics of aircraft.

  14. Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

    NASA Technical Reports Server (NTRS)

    Buchholz, Mark D.; Tso, Jin

    1993-01-01

    Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

  15. Basic studies on delta wing flow modifications by means of apex fences

    NASA Technical Reports Server (NTRS)

    Hoffler, K. D.; Rao, D. M.; Frassinelli, M. C.

    1986-01-01

    The effectiveness of apex fences on a 60-deg delta wing at low speeds was experimentally investigated. Resembling highly swept spoilers in appearance, the fences are designed to fold out of the wing apex region upper surface near the leading edges, where they generate a powerful vortex pair. The intense suction of the fence vortices augments lift in the apex region, the resulting positive pitching moment being utilized to trim trailing edge flaps for lift augmentation during approach and landing at relatively low angles of attack. The fences reduce the apex lift at high angles of attack, leading to a desirable nose-down moment. The above projected functions of the apex fence device were validated and quantified through low speed tunnel tests, comprising upper surface pressure surveys on a semispan model and balance measurements on a geometrically similar fully span wing/body configuration. Fence parameters such as area, shape, hinge position and deflection angle were investigated. Typical results are presented indicating the apex fence potential in controlling the longitudinal characteristics of a tail-less delta.

  16. Theoretical studies on flapped delta wings

    NASA Technical Reports Server (NTRS)

    Oh, S.; Tavella, D.; Roberts, L.

    1988-01-01

    The effects of leading edge flaps on the aerodynamic characteristics of a low aspect-ratio delta wing are studied theoretically. As an extension of the classical crossflow plane analysis and in order to include separated shear layers, an analogy between three dimensional steady conical and two dimensional unsteady self-similar flows is explored. This analogy provides a simple steady-unsteady relationship. The criteria for the validity of the steady-unsteady analogy are also examined. Two different theoretical techniques are used to represent the separated shear layers based on the steady-unsteady analogy, neglecting the trailing edge effect. In the first approach, each vortex system is represented by a pair of concentrated vortices connected to the separation points by straight feeding sheets. In the second approach, the vortex cloud method is adopted for simulating the flow field in the crossflow plane. The separated shear layers are replaced with a cloud of discrete vortices and the boundary element method is employed to represent the wing trace by a vorticity distribution. A simple merging scheme is used to model the core region of the vortical flow as a single vortex by imposing a restriction on the shear layer rotation angle. The results are compared with experiments and with results from 3-D panel calculations.

  17. Birds' tails do act like delta wings but delta-wing theory does not always predict the forces they generate.

    PubMed Central

    Evans, Matthew R

    2003-01-01

    Delta-wing theory, which predicts the aerodynamics of aircraft like the Concorde, is the conventional explanation for the way in which a bird's tail operates in flight. Recently, doubt has been cast on the validity of applying a theory devised for supersonic aircraft to the small tails of slow-flying birds. By testing delta-wing models and birds' tails behind bodies with wings, I empirically show that the tails of birds produce lift in a very similar way to conventional delta-wing models. Both Perspex and birds' tail models produce lift similar to that predicted by delta-wing theory when narrowly spread and at low angles of attack. However, when widely spread and at high angles of attack, both tails and Perspex models produce much less lift than predicted, owing to vortex breakdown after which the assumptions of delta-wing theory are violated. These results indicate that birds' tails can be regarded as delta wings but that the theory predicting the forces produced by delta wings can only be applied within acceptable limits (i.e. tails spread less than 60 degrees and at angles of attack of less than 20 degrees). PMID:12965029

  18. Subsonic investigations of vortex interaction control for enhanced high-alpha aerodynamics of a chine forebody/Delta wing configuration

    NASA Technical Reports Server (NTRS)

    Rao, Dhanvada M.; Bhat, M. K.

    1992-01-01

    A proposed concept to alleviate high alpha asymmetry and lateral/directional instability by decoupling of forebody and wing vortices was studied on a generic chine forebody/ 60 deg. delta configuration in the NASA Langley 7 by 10 foot High Speed Tunnel. The decoupling technique involved inboard leading edge flaps of varying span and deflection angle. Six component force/moment characteristics, surface pressure distributions and vapor-screen flow visualizations were acquired, on the basic wing-body configuration and with both single and twin vertical tails at M sub infinity = 0.1 and 0.4, and in the range alpha = 0 to 50 deg and beta = -10 to +10 degs. Results are presented which highlight the potential of vortex decoupling via leading edge flaps for enhanced high alpha lateral/directional characteristics.

  19. Navier-Stokes and Euler solutions for lee-side flows over supersonic delta wings. A correlation with experiment

    NASA Technical Reports Server (NTRS)

    Mcmillin, S. Naomi; Thomas, James L.; Murman, Earll M.

    1990-01-01

    An Euler flow solver and a thin layer Navier-Stokes flow solver were used to numerically simulate the supersonic leeside flow fields over delta wings which were observed experimentally. Three delta wings with 75, 67.5, and 60 deg leading edge sweeps were computed over an angle-of-attack range of 4 to 20 deg at a Mach number 2.8. The Euler code and Navier-Stokes code predict equally well the primary flow structure where the flow is expected to be separated or attached at the leading edge based on the Stanbrook-Squire boundary. The Navier-Stokes code is capable of predicting both the primary and the secondary flow features for the parameter range investigated. For those flow conditions where the Euler code did not predict the correct type of primary flow structure, the Navier-Stokes code illustrated that the flow structure is sensitive to boundary layer model. In general, the laminar Navier-Stokes solutions agreed better with the experimental data, especially for the lower sweep delta wings. The computational results and a detailed re-examination of the experimental data resulted in a refinement of the flow classifications. This refinement in the flow classification results in the separation bubble with the shock flow type as the intermediate flow pattern between separated and attached flows.

  20. An aerodynamic model for one and two degree of freedom wing rock of slender delta wings

    NASA Technical Reports Server (NTRS)

    Hong, John

    1993-01-01

    The unsteady aerodynamic effects due to the separated flow around slender delta wings in motion were analyzed. By combining the unsteady flow field solution with the rigid body Euler equations of motion, self-induced wing rock motion is simulated. The aerodynamic model successfully captures the qualitative characteristics of wing rock observed in experiments. For the one degree of freedom in roll case, the model is used to look into the mechanisms of wing rock and to investigate the effects of various parameters, like angle of attack, yaw angle, displacement of the separation point, and wing inertia. To investigate the roll and yaw coupling for the delta wing, an additional degree of freedom is added. However, no limit cycle was observed in the two degree of freedom case. Nonetheless, the model can be used to apply various control laws to actively control wing rock using, for example, the displacement of the leading edge vortex separation point by inboard span wise blowing.

  1. Flutter of pairs of aerodynamically interfering delta wings.

    NASA Technical Reports Server (NTRS)

    Chipman, R. R.; Rauch, F. J.; Hess, R. W.

    1973-01-01

    To examine the effect on flutter of the aerodynamic interference between pairs of closely spaced delta wings, several structurally uncoupled 1/80th-scale models were studied by experiment and analysis. Flutter test boundaries obtained in NASA Langley's 26-in. transonic blowdown wind tunnel were compared with subsonic analytical results generated using the doublet lattice method. Trends for several combinations of vertical and longitudinal wing separation were determined, showing flutter speed significantly affected in the closely spaced configurations. A new flutter mechanism coupling one wing's first bending mode with the other wing's first torsion mode was predicted and observed.

  2. The effect of asymmetric vortex wake characteristics on a slender delta wing undergoing wing rock motion

    NASA Technical Reports Server (NTRS)

    Arena, A. S., Jr.; Nelson, R. C.

    1989-01-01

    An experimental investigation into the fluid mechanisms responsible for wing rock on a slender delta wing with 80 deg leading edge sweep has been conducted. Time history and flow visualization data are presented for a wide angle-of-attack range. The use of an air bearing spindle has allowed the motion of the wing to be free from bearing friction or mechanical hysteresis. A bistable static condition has been found in vortex breakdown at an angle of attack of 40 deg which causes an overshoot of the steady state rocking amplitude. Flow visualization experiments also reveal a difference in static and dynamic breakdown locations on the wing. A hysteresis loop in dynamic breakdown location similar to that seen on pitching delta wings was observed as the wing was undergoing the limit cycle oscillation.

  3. Lift augmentation on a delta wing via leading edge fences and the Gurney flap. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Buchholz, Mark D.

    1992-01-01

    Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

  4. Effect of leading edge roundness on a delta wing in wing-rock motion

    NASA Technical Reports Server (NTRS)

    Ng, T. Terry; Malcolm, Gerald N.

    1990-01-01

    The effect of wing leading-edge roundness on wing rock was investigated using flow visualization in a water tunnel. Eighty degree delta wing models were tested on free-to-roll and forced oscillation rigs. The onset of wing rock was delayed by increasing the roundness of the leading edges. The wing rock amplitude and frequency results suggested that damping was increased at lower angles of attack but reduced at higher angles of attack. Vortex lift-off and vortex breakdown, especially during dynamic situations, were strongly affected by the leading edge roundness. Different forms of wing rock motion could be sustained by combinations of vortex breakdown and vortex lift-off. Behaviors of the wing and vortex motions were explained by the influence of leading edge roundness on the separation location, vortex trajectory, and vortex breakdown.

  5. The DELTA MONSTER: An RPV designed to investigate the aerodynamics of a delta wing platform

    NASA Technical Reports Server (NTRS)

    Connolly, Kristen; Flynn, Mike; Gallagher, Randy; Greek, Chris; Kozlowski, Marc; Mcdonald, Brian; Mckenna, Matt; Sellar, Rich; Shearon, Andy

    1989-01-01

    The mission requirements for the performance of aerodynamic tests on a delta wind planform posed some problems, these include aerodynamic interference; structural support; data acquisition and transmission instrumentation; aircraft stability and control; and propulsion implementation. To eliminate the problems of wall interference, free stream turbulence, and the difficulty of achieving dynamic similarity between the test and actual flight aircraft that are associated with aerodynamic testing in wind tunnels, the concept of the remotely piloted vehicle which can perform a basic aerodynamic study on a delta wing was the main objective for the Green Mission - the Delta Monster. The basic aerodynamic studies were performed on a delta wing with a sweep angle greater than 45 degrees. These tests were performed at various angles of attack and Reynolds numbers. The delta wing was instrumented to determine the primary leading edge vortex formation and location, using pressure measurements and/or flow visualization. A data acquisition system was provided to collect all necessary data.

  6. 65 Degree Delta Wing in the LTPT-Test 360

    NASA Technical Reports Server (NTRS)

    1991-01-01

    This 65-degree delta wing model test is part of an experimental and computational effort aimed at isolating the effects of Reynolds number, Mach number, and leading-edge radius on partial leading-edge vortex separation. Surface pressure measurments were taken in the Low-Turbulence Pressure Tunnel (LTPT) for a Reynolds number range from 1.5 to 12 Million over an angle of attack range from -2 to 30 degrees at Mach numbers of 0.2 and 0.4. This model is a 3/4-scale version of another 65-degree delta wing tested to full-scale Reynolds numbers in the National Transonic Facility.

  7. Static measurements of slender delta wing rolling moment hysteresis

    NASA Technical Reports Server (NTRS)

    Katz, Joseph; Levin, Daniel

    1991-01-01

    Slender delta wing planforms are susceptible to self-induced roll oscillations due to aerodynamic hysteresis during the limit cycle roll oscillation. Test results are presented which clearly establish that the static rolling moment hysteresis has a damping character; hysteresis tends to be greater when, due to either wing roll or side slip, the vortex burst moves back and forth over the wing trailing edge. These data are an indirect indication of the damping role of the vortex burst during limit cycle roll oscillations.

  8. Pitot-pressure distributions of the flow field of a delta-wing orbiter

    NASA Technical Reports Server (NTRS)

    Cleary, J. W.

    1972-01-01

    Pitot pressure distributions of the flow field of a 0.0075-scale model of a typical delta wing shuttle orbiter are presented. Results are given for the windward and leeward sides on centerline in the angle-of-attack plane from wind tunnel tests conducted in air. Distributions are shown for three axial stations X/L = .35, .60, and .98 and for angles of attack from 0 to 60 deg. The tests were made at a Mach number of 7.4 and for Reynolds numbers based on body length from 1,500,000 to 9,000,000. The windward distributions at the two survey stations forward of the body boat tail demonstrate the compressive aspects of the flow from the shock wave to the body. Conversely, the distributions at the aft station display an expansion of the flow that is attributed to body boat tail. On the lee side, results are given at low angles of attack that illustrate the complicating aspects of the canopy on the flow field, while results are given to show the effects of flow separation at high angles of attack.

  9. Lee side flow for slender delta wings of finite thickness

    NASA Technical Reports Server (NTRS)

    Szodruch, J. G.

    1980-01-01

    An experimental and theoretical investigation carried out to determine the lee side flow field over delta wings at supersonic speeds is presented. A theoretical method to described the flow field is described, where boundary conditions as a result of the experimental study are needed. The computed flow field with shock induced separation is satisfactory.

  10. Flutter analysis of highly swept delta wings by conventional methods

    NASA Technical Reports Server (NTRS)

    Gibbons, M. D.; Soistmann, D. L.; Bennett, R. M.

    1988-01-01

    The flutter boundaries of six thin highly-swept delta-platform wings have been calculated. Comparisons are made between experimental data and results using several aerodynamic methods. The aerodynamic methods used include a subsonic and supersonic kernel function, second order piston theory, and a transonic small disturbance code. The dynamic equations of motion are solved using analytically calculated mode shapes and frequencies.

  11. A vortex entrainment model applied to slender delta wings

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.

    1974-01-01

    A mathematical model of the vortex flow over a slender sharp-edged delta wing is proposed, and is shown to provide good agreement with the experiment. Although the technique requires experimental data in the form of the vortex core locations, it does account for the previously ignored mass entrainment of the vortex core.

  12. Leading edge vortex dynamics on a pitching delta wing

    NASA Technical Reports Server (NTRS)

    Lemay, S. P.; Batill, S. M.; Nelson, R. C.

    1988-01-01

    A study of the dynamic behavior of the leading edge vortices on a delta wing undergoing oscillatory pitching motion is presented. A sharp edge, flat plate, delta wing having a sweep angle of 70 deg was used in this investigation. The wing was sinusoidally pitched about its 1/2 chord position at reduced frequencies ranging from k = 2(pi)fc/u = 0.05 to 0.30 at chord Reynolds numbers between 90,000 and 350,000, for angle of attack ranges of 29 to 39 deg and 0 to 45 deg. During these dynamic motions, visualization of the leading edge vortices was obtained by marking the vortices with TiCl4 introduced through ports located near the model apex. The location of vortex breakdown was recorded using high speed motion picture photography. The motion picture records were analyzed to determine vortex trajectory and breakdown position as a function of angle of attack. When the wing was sinusoidally pitched, a hysteresis was observed in the location of breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing, and the phase lag between model motion and breakdown location were also determined. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

  13. Forced unsteady separated flows on a 45 degree delta wing

    NASA Astrophysics Data System (ADS)

    Huyer, Stephen Albert

    A great deal of current research activities has focused on possible exploitation of forced unsteady separated flows to provide enhanced lift and maneuvering characteristics. The formal and intentional utilization of these flows is currently being manifested in the form of the Advanced Tactical Fighter. The wing planform geometry of the ATF and other fighter aircraft is a delta wing. Under steady conditions, leading edge vortices are formed on each side of a delta wing. These vortices are mostly responsible for the attainment of lift to high angles of attack. Unsteady motion histories will likely alter the characteristics of this vortex as well as its development history. This will then present new difficulties in terms of lift enhancement and control. In order to successfully predict and optimally exploit the flight regimes offered by the ATF, greater understanding of the underlying physical mechanisms responsible for these unsteady flow fields must be obtained. The vortex dominated flow fields produced by an oscillating 45 degree delta wing were examined across a wide range of unsteady motion histories. Still and high speed video photography were employed to document the flow development processes and cortex kinematics. Force balance data recorded the unsteady aerodynamic loading produced. These methods allowed for a thorough qualitative and quantitative examination of the flow fields elicited by a pitching delta wing. The wide range of motion histories employed were found to have a tremendous impact in terms of flow development lift enhancement, drag reduction, and overall aerodynamic performance. Integrations of the data permitted speculation regarding the underlying physical mechanisms responsible for the observed phenomena. Experimental evidence allowed for hypotheses regarding the physical mechanisms of vorticity production, accumulation, convection, and diffusion.

  14. Aerodynamic performance of a wing with a deflected tip-mounted reverse half-delta wing

    NASA Astrophysics Data System (ADS)

    Lee, T.; Su, Y. Y.

    2012-11-01

    The impact of a tip-mounted 65°-sweep reverse half-delta wing (RHDW), set at different deflections, on the aerodynamic performance of a rectangular NACA 0012 wing was investigated experimentally at Re = 2.45 × 105. This study is a continuation of the work of Lee and Su (Exp Fluids 52(6):1593-1609, 2012) on the passive control of wing tip vortex by the use of a reverse half-delta wing. The present results show that for RHDW deflection with -5° ≤ δ ≤ +15°, the lift was found to increase nonlinearly with increasing δ compared to the baseline wing. The lift increment was accompanied by an increased total drag. For negative RHDW deflection with δ < -5°, the RHDW-induced lift decrement was, however, accompanied by an improved drag. The deflected RHDW also significantly modified and weakened the tip vortex, leading to a persistently lowered lift-induced drag, regardless of its deflection, compared to the baseline wing. Physical mechanisms responsible for the observed RHDW-induced phenomenon were also discussed.

  15. Delta Wing Vortex Breakdown Suppression by Vortex Core Oscillation

    NASA Astrophysics Data System (ADS)

    Cain, Charles

    2000-11-01

    The flow over a delta wing is characterized by two counter-rotating vortices that can undergo a sudden radial expansion at high angles of attack known as vortex breakdown. Downstream of this breakdown is a region of organized unsteady flow that can cause tail buffeting and structural fatigue, especially on twin-tailed aircraft. The recent self-induction theory of vortex breakdown points to the "pile-up" of vorticity due to the linear addition of vorticity in the spiraling shear layer that surrounds the vortex core as a principal cause of vortex breakdown (Kurosaka 1998). Based on that theory, this research attempts to relieve vorticity pile-up by altering the straight-line path of the vortex core and preventing the linear addition of vorticity. This is accomplished by applying a combination of periodic blowing and suction with low mass and momentum flux. The blowing and suction are directed normal to the low-pressure surface and supplied from ports under the vortex core which are near the forward tip of the delta wing. This oscillating input causes the vortex core to transition into a spiral formation downstream of the input ports. Initial results indicate that this change in the vortex core path may prevent vortex breakdown over the surface of the delta wing.

  16. Conical Euler solution for a highly-swept delta wing undergoing wing-rock motion

    NASA Technical Reports Server (NTRS)

    Lee, Elizabeth M.; Batina, John T.

    1990-01-01

    Modifications to an unsteady conical Euler code for the free-to-roll analysis of highly-swept delta wings are described. The modifications involve the addition of the rolling rigid-body equation of motion for its simultaneous time-integration with the governing flow equations. The flow solver utilized in the Euler code includes a multistage Runge-Kutta time-stepping scheme which uses a finite-volume spatial discretization on an unstructured mesh made up of triangles. Steady and unsteady results are presented for a 75 deg swept delta wing at a freestream Mach number of 1.2 and an angle of attack of 30 deg. The unsteady results consist of forced harmonic and free-to-roll calculations. The free-to-roll case exhibits a wing rock response produced by unsteady aerodynamics consistent with the aerodynamics of the forced harmonic results. Similarities are shown with a wing-rock time history from a low-speed wind tunnel test.

  17. Delta wings with shock-free cross flow

    NASA Technical Reports Server (NTRS)

    Sritharan, S. S.

    1984-01-01

    In order to have a high level of maneuverability, supersonic delta wings should have a cross flow that is free of embedded shock waves. The conical cross flow sonic surface differs from that of plane transonic flow in many aspects. Well-known properties such as the monotone law are not true for conical cross flow sonic surfaces. By using a local analysis of the cross flow sonic line, relevant conditions for smooth cross flow are obtained. A technique to artificially construct a smooth sonic surface and an efficient numerical method to calculate the flow field are used to obtain cones with smooth cross flow.

  18. Delta wing flutter based on doublet lattice method in NASTRAN

    NASA Technical Reports Server (NTRS)

    Jew, H.

    1975-01-01

    The subsonic doublet-lattice method (DLM) aeroelastic analysis in NASTRAN was successfully applied to produce subsonic flutter boundary data in parameter space for a large delta wing configuration. Computed flow velocity and flutter frequency values as functions of air density ratio, flow Mach number, and reduced frequency are tabulated. The relevance and the meaning of the calculated results are discussed. Several input-deck problems encountered and overcome are cited with the hope that they may be helpful to NASTRAN Rigid Format 45 users.

  19. Study of lee-side flows over conically cambered delta wings at supersonic speeds, part 1

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Watson, Carolyn B.

    1987-01-01

    An experimental investigation was performed in which surface pressure data, flow visualization data, and force and moment data were obtained on four conical delta wing models which differed in leading-edge camber only. Wing leading-edge camber was achieved through a deflection of the outboard 30% of the local wind semispan of a reference 75 degrees swept flat delta wing. The four wing models have leading-edge deflection angles delta sub F of 0, 5, 10, and 15 degrees measured streamwise. Data for the wings with delta sub F = 10 and 15 degrees showed that hinge-line separation dominated the lee-side wing loading and prohibited the develpment of leading-edge separation on the deflected portion of wing leading edge. However, data for the wing with delta sub F = 5 degrees, a vortex was positioned on the deflected leading edge with reattachment at the hinge line. Flow visualization results were presented which detail the influence of Mach number, angle of attack, and camber on the lee-side flow characteristics of conically cambered delta wings. Analysis of photgraphic data identified the existence of 12 distinctive lee-side flow types. In general, the aerodynamic force and moment data correlated well with the pressure and flow visualization data.

  20. Unsteadiness of Flow Structure on Low Swept Delta Wing

    NASA Astrophysics Data System (ADS)

    Zharfa, Mohammadreza; Ozturk, Ilhan; Yavuz, Mehmet Metin

    2013-11-01

    The flow structure of low-sweep 35 degree delta wing has been analyzed experimentally using flow visualization and flow measurement techniques. Laser illuminated smoke visualization, Laser Doppler Anemometry (LDA), and surface pressure measurements are performed to understand the steady and unsteady behavior of the flow regimes. Reynolds number varying from 10,000 to 100,000 and attack angles varying from 3 to 10 are tested. For the corresponding Reynolds Numbers and attack angles, prestall and poststall regimes are identified. The amplitude and frequency of the pressure and velocity fluctuations at different locations are compared with the regions of reattachment, vortex breakdown, and stall. Using statistics and spectral analysis, the unsteadiness of flow is studied in detail. Both the lift performance of the wing and the regions possibly exposed to surface buffeting are tried to be interpreted. Using the results of the study, an effective active flow control strategy to delay stall and to reduce surface buffeting is going to be determined next. The project was supported by the Turkish Scientific and Technological Research Council of Turkey (3501 - 111M732).

  1. Vortex structures for flow over a delta wing with sinusoidal leading edge

    NASA Astrophysics Data System (ADS)

    Chen, Huang; Wang, Jin-Jun

    2014-06-01

    The idea of using sinusoidal leading edge as a kind of passive flow control method was inspired by observing the flipper movement of the humpback whale. It was believed that the protuberances along the whale's pectoral fin could delay stall, thus would enhance the maneuverability of the whale. It has also been shown that when equipped with sinusoidal leading edges, the stall of a delta wing could be delayed. In this paper, stereoscopic particle image velocimetry was adopted to study the vortex structures for the flow over a 52° swept delta wing with sinusoidal leading edges. A direct comparison with the flow over a baseline delta wing was made to illustrate the different vortex structures of these two kinds of models. Results have shown that the flow over the baseline delta wing was dominated by dual leading-edge vortices (LEVs), a structure that only existed for flow over nonslender delta wing at certain Reynolds number. On the other hand, the flow over the one with sinusoidal leading edge showed a very different pattern. It has been found in this paper that there were several pairs of LEVs existed on the leeward side of the wing, which might explain the stall-delaying effect of the delta wing with sinusoidal leading edges.

  2. How do birds' tails work? Delta-wing theory fails to predict tail shape during flight.

    PubMed Central

    Evans, Matthew R; Rosén, Mikael; Park, Kirsty J; Hedenström, Anders

    2002-01-01

    Birds appear to use their tails during flight, but until recently the aerodynamic role that tails fulfil was largely unknown. In recent years delta-wing theory, devised to predict the aerodynamics of high-performance aircraft, has been applied to the tails of birds and has been successful in providing a model for the aerodynamics of a bird's tail. This theory now provides the conventional explanation for how birds' tails work. A delta-wing theory (slender-wing theory) has been used, as part of a variable-geometry model to predict how tail and wing shape should vary during flight at different airspeeds. We tested these predictions using barn swallows flying in a wind tunnel. We show that the predictions are not quantitatively well supported. This suggests that a new theory or a modified version of delta-wing theory is needed to adequately explain the way in which morphology varies during flight. PMID:12028763

  3. Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Covell, Peter F.; Wood, Richard M.; Miller, David S.

    1987-01-01

    An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

  4. Euler and Potential Experiment/CFD Correlations for a Transport and Two Delta-Wing Configurations

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Cliff, S. E.; Melton, J. E.; Langhi, R. G.; Goodsell, A. M.; Robertson, D. D.; Moyer, S. A.

    1990-01-01

    A selection of successes and failures of Computational Fluid Dynamics (CFD) is discussed. Experiment/CFD correlations involving full potential and Euler computations of the aerodynamic characteristics of four commercial transport wings and two low aspect ratio, delta wing configurations are shown. The examples consist of experiment/CFD comparisons for aerodynamic forces, moments, and pressures. Navier-Stokes equations are not considered.

  5. A Discrete-Vortex Method for Studying the Wing Rock of Delta Wings

    NASA Technical Reports Server (NTRS)

    Gainer, Thomas G.

    2002-01-01

    A discrete-vortex method is developed to investigate the wing rock problem associated with highly swept wings. The method uses two logarithmic vortices placed above the wing to represent the vortex flow field and uses boundary conditions based on conical flow, vortex rate of change of momentum, and other considerations to position the vortices and determine their strengths. A relationship based on the time analogy and conical-flow assumptions is used to determine the hysteretic positions of the vortices during roll oscillations. Static and dynamic vortex positions and wing rock amplitudes and frequencies calculated by using the method are generally in good agreement with available experimental data. The results verify that wing rock is caused by hysteretic deflections of the vortices and indicate that the stabilizing moments that limit wing rock amplitudes are the result of the one primary vortex moving outboard of the wing where it has little influence on the wing.

  6. Computation of vortex-dominated flow for a delta wing undergoing pitching oscillation

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Chuang, H. Andrew

    1990-01-01

    The conservative, unsteady Euler equations for the flow relative to a moving frame of reference are used to solve for the three-dimensional steady and unsteady flows around a sharp-edged delta wing. The resulting equations are solved by using an implicit, approximately factored, finite-volume scheme. Implicit second-order and explicit second- and fourth-order dissipations are added to the scheme. The boundary conditions are explicitly satisfied. The grid is generated by locally using a modified Joukowski transformation in crossflow planes at the grid-chord stations. The computational applications cover a steady flow around a delta wing, whose results serve as the initial conditions for the unsteady flow around a pitching delta wing at a large mean angle of attack. The steady results are compared with the experimental data, and the unsteady results are compared with results of a flux-difference splitting scheme.

  7. Leading edge vortex-flap experiments on a 74 deg delta wing

    NASA Technical Reports Server (NTRS)

    Rao, D. M.

    1979-01-01

    Exploratory wind tunnel tests are reported on a 74 deg. delta wing model. The potential of a vortex flap concept in reducing the subsonic lift dependent drag of highly swept, slender wings is examined. The suction effect of coiled vortices generated through controlled separation over leading edge flap surfaces to produce a thrust component is discussed. A series of vortex-flap configurations were investigated to explore the effect of some primary geometric variables.

  8. The lift of sharp-leading-edged delta wings with blowing

    NASA Technical Reports Server (NTRS)

    Tavella, D. A.

    1985-01-01

    An analysis of the lift augmentation due to a thin jet of air issuing from a slot along the leading edge of a delta wing is presented. The problem is treated with an extension of the method of Brown and Michael, representing the separated flow on the lee side of the wing by a pair of concentrated vortices and corresponding feeding sheets. It is assumed that the jet is not affected by Coanda forces. The analysis produces qualitative agreement with experiments.

  9. An experimental investigation of delta wing vortex flow with and without external jet blowing

    NASA Technical Reports Server (NTRS)

    Iwanski, Kenneth P.; Ng, T. Terry; Nelson, Robert C.

    1989-01-01

    A visual and quantitative study of the vortex flow field over a 70-deg delta wing with an external jet blowing parallel to and at the leading edge was conducted. In the experiment, the vortex core was visually marked with TiCl4, and LDA was used to measure the velocity parallel and normal to the wing surface. It is found that jet blowing moved vortex breakdown farther downstream from its natural position and influenced the breakdown characteristics.

  10. Computation of transonic viscous flow past the NTF 65-degree Delta Wing

    NASA Technical Reports Server (NTRS)

    Wu, Chivey

    1993-01-01

    This project is a continuation of the work performed in the summers of 1991 and 1992, during which a 9-block structured grid for the computational domain around each of the four NTF 65-degree Delta Wing models with the sting mount were created. The objective of the project is to validate and supplement the test data on the wing models by computing the viscous flow field about the models.

  11. Deflection and Stress Analysis of Thin Solid Wings of Arbitrary Plan Form with Particular Reference to Delta Wings

    NASA Technical Reports Server (NTRS)

    Stein, Manuel; Anderson, Edward; Hedgepeth, John M

    1953-01-01

    The structural analysis of arbitrary solid cantilever wings by small-deflection thin-plate theory is reduced to the solution of linear ordinary differential equations by the assumption that the chordwise deflections at any spanwise station may be expressed in the form of a power series in which the coefficients are functions of the spanwise coordinate. If the series is limited to the first two and three terms (that is, if linear and parabolic chordwise deflections, respectively, are assumed), the differential equations for the coefficients are solved exactly for uniformly loaded solid delta wings of constant thickness ratio. For cases for which exact solutions to the differential equations cannot be obtained, a numerical procedure is derived. Experimental deflection and stress data for constant-thickness delta-plate specimens of 45 degree and 60 degree sweep are presented and are found to compare favorably with the present theory.

  12. Numerical simulation of incidence and sweep effects on delta wing vortex breakdown

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

    The structure of the vortical flowfield over delta wings at high angles of attack was investigated. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics, including leading-edge separation, secondary separation, and vortex breakdown. Flows over a 75- and a 63-deg sweep delta wing with sharp leading edges were investigated and compared with available experimental data. The effect of variation of circumferential grid resolution grid resolution in the vicinity of the wing leading edge on the accuracy of the solutions was addressed. Furthermore, the effect of turbulence modeling on the solutions was investigated. The effects of variation of angle of attack on the computed vortical flow structure for the 75-deg sweep delta wing were examined. At moderate angles of attack no vortex breakdown was observed. When a critical angle of attack was reached, bubble-type vortex breakdown was found. With further increase in angle of attack, a change from bubble-type breakdown to spiral-type vortex breakdown was predicted by the numerical solution. The effects of variation of sweep angle and freestream Mach number were addressed with the solutions on a 63-deg sweep delta wing.

  13. Evaluation of the Levy Method as Applied to Vibrations of a 45 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Kruszewski, Edwin T.; Waner, Paul G., Jr.

    1959-01-01

    The Levy method which deals with an idealized structure was used to obtain the natural modes and frequencies of a large-scale built-up 45 deg. delta wing. The results from this approach, both with and without the effects of transverse shear, were compared with the results obtained experimentally and also with those calculated by the Stein-Sanders method. From these comparisons it was concluded that the method as proposed by Levy gives excellent results for thin-skin delta wings, provided that corrections are made for the effect of transverse shear.

  14. Space shuttle: Static stability and control investigation of NR/GD delta wing booster (B-20) and delta wing orbiter (134-D), volume 3

    NASA Technical Reports Server (NTRS)

    Allen, E. C., Jr.; Eder, F. W.

    1972-01-01

    Experimental aerodynamic investigations have been made on a .0035 scale model North American Rockwell/General Dynamics version of the space shuttle in the NASA/MSFC 14 x 14 Inch Trisonic Wind Tunnel. Static stability and control data were obtained on the delta wing booster alone (B-20) and with the delta wing orbiter (134D) mounted in various positions on the booster. Six component aerodynamic force and moment data were recorded over an angle of attack range from -10 to 24 deg at 0 and 6 deg sideslip angles and from -10 to +10 deg sideslip at 0 deg angle of attack. Mach number ranged from 0.6 to 4.96.

  15. Pressure measurements on a thick cambered and twisted 58 deg delta wing at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Lamar, John E.

    1987-01-01

    A pressure experiment at high subsonic speeds was conducted by a cambered and twisted thick delta wing at the design condition (Mach number 0.80), as well as at nearby Mach numbers (0.75 and 0.83) and over an angle-of-attack range. Effects of twin vertical tails on the wing pressure measurements were also assessed. Comparisons of detailed theoretical and experimental surface pressures and sectional characteristics for the wing alone are presented. The theoretical codes employed are FLO-57, FLO-28, PAN AIR, and the Vortex Lattice Method-Suction Analogy.

  16. Leading edge vortex dynamics on a pitching delta wing. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Lemay, Scott P.

    1988-01-01

    The leading edge flow structure was investigated on a 70 deg flat plate delta wing which was pitched about its 1/2 chord position, to increase understanding of the high angle of attack aerodynamics on an unsteady delta wing. The wing was sinusoidally pitched at reduced frequencies ranging from k being identical with 2pi fc/u = 0.05 to 0.30 at chord Reynolds numbers between 90,000 and 350,000, for angle of attack ranges of alpha = 29 to 39 deg and alpha = 0 to 45 deg. The wing was also impulsively pitched at an approximate rate of 0.7 rad/s. During these dynamic motions, visualization of the leading edge vorticies was obtained by entraining titanium tetrachloride into the flow at the model apex. The location of vortex breakdown was recorded using 16mm high speed motion picture photography. When the wing was sinusoidally pitched, a hysteresis was observed in the location of breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing, and the phase lag between model motion and breakdown location were also determined. When the wing was impulsively pitched, several convective times were required for the vortex flow to reach a steady state. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

  17. Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Hernandez, Gloria; Wood, Richard M.; Covell, Peter F.

    1994-01-01

    An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

  18. Conical Euler analysis and active roll suppression for unsteady vortical flows about rolling delta wings

    NASA Technical Reports Server (NTRS)

    Lee-Rausch, Elizabeth M.; Batina, John T.

    1993-01-01

    A conical Euler code was developed to study unsteady vortex-dominated flows about rolling, highly swept delta wings undergoing either forced motions or free-to-roll motions that include active roll suppression. The flow solver of the code involves a multistage, Runge-Kutta time-stepping scheme that uses a cell-centered, finite-volume, spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free to-roll case by simultaneously integrating in time the rigid-body equation of motion with the governing flow equations. Results are presented for a delta wing with a 75 deg swept, sharp leading edge at a free-stream Mach number of 1.2 and at 10 deg, 20 deg, and 30 deg angle of attack alpha. At the lower angles of attack (10 and 20 deg), forced-harmonic analyses indicate that the rolling-moment coefficients provide a positive damping, which is verified by free-to-roll calculations. In contrast, at the higher angle of attack (30 deg), a forced-harmonic analysis indicates that the rolling-moment coefficient provides negative damping at the small roll amplitudes. A free-to-roll calculation for this case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation, which is characteristic of highly swept delta wings. This limit cycle oscillation may be actively suppressed through the use of a rate-feedback control law and antisymmetrically deflected leading-edge flaps. Descriptions of the conical Euler flow solver and the free-to roll analysis are included in this report. Results are presented that demonstrate how the systematic analysis of the forced response of the delta wing can be used to predict the stable, neutrally stable, and unstable free response of the delta wing. These results also give insight into the flow physics associated with unsteady vortical flows about delta wings undergoing forced

  19. Flow Structure over Moderate Swept Delta Wing: Effects of Reynolds Number and Attack Angle

    NASA Astrophysics Data System (ADS)

    Ozturk, Ilhan; Zharfa, Mohammadreza; Yavuz, Mehmet Metin

    2013-11-01

    Recent investigations have revealed the appearance of a distinctive type of leading edge vortex, dual vortex structure, over simple delta wing planforms having moderate sweep angles. Flow over a moderate swept 45-degree wing has been investigated using laser illuminated smoke visualization, Laser Doppler Anemometry (LDA), and surface pressure measurements. The effects of Reynolds number and attack angles on dual vortex structure, vortex breakdown, and poststall regime are reported. The footprint of flow regimes on the surface of the planform is captured by the pressure measurements, and the lift performance of the wing is tried to be extracted. The relation between surface pressure fluctuations and near surface velocity fluctuations is investigated. The reattachment region of the separated shear layer on the surface, vortex breakdown, and stall regime are studied with considering the aforementioned relation, which will enlighten some of the aspects of the buffeting on the wing planform.

  20. Control of Flow Structure on Low Swept Delta Wing with Steady Leading Edge Blowing

    NASA Astrophysics Data System (ADS)

    Ozturk, Ilhan; Zharfa, Mohammadreza; Yavuz, Mehmet Metin

    2014-11-01

    Interest in unmanned combat air vehicles (UCAVs) and micro air vehicles (MAVs) has stimulated investigation of the flow structure, as well as its control, on delta wings having low and moderate values of sweep angle. In the present study, the flow structure is characterized on a delta wing of low sweep 35-degree angle, which is subjected to steady leading edge blowing. The techniques of laser illuminated smoke visualization, laser Doppler anemometry (LDA), and surface pressure measurements are employed to investigate the steady and unsteady nature of the flow structure on delta wing, in relation to the dimensionless magnitude of the blowing coefficient. Using statistics and spectral analysis, unsteadiness of the flow structure is studied in detail. Different injection locations are utilized to apply different blowing patterns in order to identify the most efficient control, which provides the upmost change in the flow structure with the minimum energy input. The study aims to find the optimum flow control strategy to delay or to prevent the stall and possibly to reduce the buffeting on the wing surface. Since the blowing set-up is computer controlled, the unsteady blowing patterns compared to the present steady blowing patterns will be studied next. This project was supported by the Scientific and Technological Research Council of Turkey (Project Number: 3501 111M732).

  1. Natural Rolling Responses of a Delta Wing in Transonic and Subsonic Flows

    NASA Technical Reports Server (NTRS)

    Menzies, Margaret A.; Kandil, Osama A.

    1996-01-01

    The unsteady, three-dimensional, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the natural rolling response of slender delta wings of zero thickness at moderate to high angles of attack, to transonic and subsonic flows. The governing equations of fluid flow and dynamics of the present multi-disciplinary problem are solved using the time-accurate solution of the NS equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The main focus is to analyze the effect of Mach number and angle of attack on the leading edge vortices and their breakdown, the resultant rolling motion, and overall aerodynamic response of the wing. Three cases demonstrate the natural response of a 65 deg swept, cropped delta wing in a transonic flow with breakdown of the leading edge vortices and an 80 deg swept delta wing in a subsonic flow undergoing either damped or self-excited limit-cycle rolling oscillations as a function of angle of attack. Comparisons with an experimental investigation completes this study, validating the analysis and illustrating the complex details afforded by computational investigations.

  2. Simulation of tail buffet using delta wing-vertical tail configuration

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Kandil, Hamdy A.; Massey, Steven J.

    1993-01-01

    Computational simulation of the vertical tail buffet problem is accomplished using a delta wing-vertical tail configuration. Flow conditions are selected such that the wing primary-vortex cores experience vortex breakdown and the resulting flow interacts with the vertical tail. This multidisciplinary problem is solved successively using three sets of equations for the fluid flow, aeroelastic deflections and grid displacements. For the fluid dynamics part, the unsteady, compressible, full Navier-Stokes equations are solved accurately in time using an implicit, upwind, flux-difference splitting, finite-volume scheme. For the aeroelastic part, the aeroelastic equation for bending vibrations is solved accurately in time using the Galerkin method and the four-stage Runge-Kutta scheme. The grid for the fluid dynamics computations is updated every few time steps using a third set of interpolation equations. The computational application includes a delta wing of aspect ratio 1 and a rectangular vertical tail of aspect ratio 2, which is placed at 0.5 root-chord length downstream of the wing trailing edge. The wing angle of attack is 35 deg and the flow Mach number and Reynolds number are 0.4 and 10,000, respectively.

  3. Conical Euler simulation and active suppression of delta wing rocking motion

    NASA Technical Reports Server (NTRS)

    Lee, Elizabeth M.; Batina, John T.

    1990-01-01

    A conical Euler code was developed to study unsteady vortex-dominated flows about rolling highly-swept delta wings, undergoing either forced or free-to-roll motions including active roll suppression. The flow solver of the code involves a multistage Runge-Kutta time-stepping scheme which uses a finite volume spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free-to-roll case, by including the rigid-body equation of motion for its simultaneous time integration with the governing flow equations. Results are presented for a 75 deg swept sharp leading edge delta wing at a freestream Mach number of 1.2 and at alpha equal to 10 and 30 deg angle of attack. A forced harmonic analysis indicates that the rolling moment coefficient provides: (1) a positive damping at the lower angle of attack equal to 10 deg, which is verified in a free-to-roll calculation; (2) a negative damping at the higher angle of attack equal to 30 deg at the small roll amplitudes. A free-to-roll calculation for the latter case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation. The wing rocking motion may be actively suppressed, however, through the use of a rate-feedback control law and antisymmetrically deflected leading edge flaps. The descriptions of the conical Euler flow solver and the free-to-roll analysis are presented. Results are also presented which give insight into the flow physics associated with unsteady vortical flows about forced and free-to-roll delta wings, including the active roll suppression of this wing-rock phenomenon.

  4. Reynolds Number, Compressibility, and Leading-Edge Bluntness Effects on Delta-Wing Aerodynamics

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2004-01-01

    An overview of Reynolds number, compressibility, and leading edge bluntness effects is presented for a 65 degree delta wing. The results of this study address both attached and vortex-flow aerodynamics and are based upon a unique data set obtained in the NASA-Langley National Transonic Facility (NTF) for i) Reynolds numbers ranging from conventional wind-tunnel to flight values, ii) Mach numbers ranging from subsonic to transonic speeds, and iii) leading-edge bluntness values that span practical slender wing applications. The data were obtained so as to isolate the subject effects and they present many challenges for Computational Fluid Dynamics (CFD) studies.

  5. Some aspects of hybrid-zeppelins. [optimization of delta wings for airships

    NASA Technical Reports Server (NTRS)

    Mackrodt, P. A.

    1975-01-01

    To increase an airship's maneuverability and payload capacity as well as to save bouyant gas it is proposed to outfit it with a slender delta-wing, which carries about one half of the total take-off weight of the vehicle. An optimization calculation based on the data of LZ 129 (the last airship, which saw passenger-service) leads to a Hybrid-Zeppelin with a wing of aspect-ratio 1.5 and 105 m span. The vehicle carries a payload of 40% of it's total take-off weight and consumes 0.8 t fuel per ton payload over a distance of 10000 km.

  6. Experimental study of supersonic viscous leeside flow over a slender delta wing

    NASA Technical Reports Server (NTRS)

    Szodruch, J.

    1980-01-01

    An investigation was conducted to study in detail the vortical flow over the leeward side of a 70 deg swept delta wing having subsonic and supersonic leading edges. Two types of flow were encountered and studied, namely leading edge separation and separation with a shock. Especially for the latter type, Reynolds number plays an important role and unexpected strong streamwise vortices were observed. An optical method is described to obtain a first aproximation of shear stress values in the streamwise direction across the wing span.

  7. Unsteady surface pressure measurements on a slender delta wing undergoing limit cycle wing rock

    NASA Technical Reports Server (NTRS)

    Arena, Andrew S., Jr.; Nelson, Robert C.

    1991-01-01

    An experimental investigation of slender wing limit cycle motion known as wing rock was investigated using two unique experimental systems. Dynamic roll moment measurements and visualization data on the leading edge vortices were obtained using a free to roll apparatus that incorporates an airbearing spindle. In addition, both static and unsteady surface pressure data was measured on the top and bottom surfaces of the model. To obtain the unsteady surface pressure data a new computer controller drive system was developed to accurately reproduce the free to roll time history motions. The data from these experiments include, roll angle time histories, vortex trajectory data on the position of the vortices relative to the model's surface, and surface pressure measurements as a function of roll angle when the model is stationary or undergoing a wing rock motion. The roll time history data was numerically differentiated to determine the dynamic roll moment coefficient. An analysis of these data revealed that the primary mechanism for the limit cycle behavior was a time lag in the position of the vortices normal to the wing surface.

  8. Effects of leading-edge flap oscillation on unsteady delta wing flow and rock control

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1991-01-01

    The isolated and interdisciplinary problems of unsteady fluid dynamics and rigid-body dynamics and control of delta wings with and without leading-edge flap oscillation are considered. For the fluid dynamics problem, the unsteady, compressible, thin-layer Navier-Stokes (NS) equations, which are written relative to a moving frame of reference, are solved along with the unsteady, linearized, Navier-displacement (ND) equations. The NS equations are solved for the flowfield using an implicit finite-volume scheme. The ND equations are solved for the grid deformation, if the leading-edge flaps oscillate, using an ADI scheme. For the dynamics and control problem, the Euler equation of rigid-body rolling motion for a wing and its flaps are solved interactively with the fluid dynamics equations for the wing-rock motion and subsequently for its control. A four-stage Runge-Kutta scheme is used to explicitly integrate the dynamics equation.

  9. A direct and inverse boundary layer method for subsonic flow over delta wings

    NASA Technical Reports Server (NTRS)

    Woodson, S. H.; Dejarnette, F. R.

    1986-01-01

    A new inverse boundary layer method is developed and applied to incompressible flows with laminar separation and reattachment. Test cases for two dimensional flows are computed and the results are compared with those of other inverse methods. One advantage of the present method is that the calculation of the inviscid velocities may be determined at each marching step without having to iterate. The inverse method was incorporated with the direct method to calculate the incompressible, conical flow over a slender delta wing at incidence. The location of the secondary separation line on the leeward surface of the wing is determined and compared with experiment for a unit aspect ratio wing at 20.5 deg incidence. The viscous flow in the separated region was calculated using prescribed skin friction coefficients.

  10. Unsteady, Transonic Flow Around Delta Wings Undergoing Coupled and Natural Modes Response: A Multidisciplinary Problem

    NASA Technical Reports Server (NTRS)

    Menzies, Margaret Anne

    1996-01-01

    The unsteady, three-dimensional Navier-Stokes equations coupled with the Euler equations of rigid-body dynamics are sequentially solved to simulate and analyze the aerodynamic response of a high angle of attack delta wing undergoing oscillatory motion. The governing equations of fluid flow and dynamics of the multidisciplinary problem are solved using a time-accurate solution of the laminar, unsteady, compressible, full Navier- Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The primary model under consideration consists of a 65 deg swept, sharp-edged, cropped delta wing of zero thickness at 20 deg angle of attack. In a freestream of Mach 0.85 and Reynolds number of 3.23 x 10(exp 6), the flow over the upper surface of the wing develops a complex shock system which interacts with the leading-edge primary vortices producing vortex breakdown. The effect of the oscillatory motion of the wing on the vortex breakdown and overall aerodynamic response is detailed to provide insight to the complicated physics associated with unsteady flows and the phenomenon of wing rock. Forced sinusoidal single and coupled mode rolling and pitching motion is presented for the wing in a transonic freestream. The Reynolds number, frequency of oscillation, and the phase angle are varied. Comparison between the single and coupled mode forced rolling and pitching oscillation cases illustrate the effects of coupling the motion. This investigation shows that even when coupled, forced rolling oscillation at a reduced frequency of 2(pi) eliminates the vortex breakdown which results in an increase in lift. The coupling effect for in phase forced oscillations show that the lift coefficient of the pitching-alone case and the rolling-moment coefficient of the rolling-alone case dominate the resulting response. However, with a phase lead in the pitching motion, the coupled motion results in a non

  11. Hypersonic incipient separation on delta wing with trailing-edge flap

    NASA Technical Reports Server (NTRS)

    Rao, D. M.

    1975-01-01

    The paper reviews the experimental data on the incipient separation characteristics of planar delta wings of 75 degree sharp leading edges, with full-span trailing edge flap deflected into the windward flow. The local Reynolds number range for these investigations covered laminar, transitional and turbulent conditions. It is shown that, while turbulent boundary layer data correlates with two dimensional results, in the laminar and transitional cases, there is a nearly parallel shift to higher flap angles for incipient separation.

  12. A Study of Grid Resolution, Transition and Turbulence Model Using the Transonic Simple Straked Delta Wing

    NASA Technical Reports Server (NTRS)

    Bartels, Robert E.

    2001-01-01

    Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.

  13. Visualization of leading edge vortices on a series of flat plate delta wings

    NASA Technical Reports Server (NTRS)

    Payne, Francis M.; Ng, T. Terry; Nelson, Robert C.

    1991-01-01

    A summary of flow visualization data obtained as part of NASA Grant NAG2-258 is presented. During the course of this study, many still and high speed motion pictures were taken of the leading edge vortices on a series of flat plate delta wings at varying angles of attack. The purpose is to present a systematic collection of photographs showing the state of vortices as a function of the angle of attack for the four models tested.

  14. A streamwise upwind algorithm applied to vortical flow over a delta wing

    NASA Technical Reports Server (NTRS)

    Goorjian, Peter M.; Obayashi, Shigeru

    1989-01-01

    Improvements were made to a streamwise upwind algorithm so that it can be used for calculating flows with vortices. A calculation is shown of flow over a delta wing at an angle of attack. The laminar, thin layer, Navier-Stokes equations are used for the calculation. The results are compared with another upwind method, a central differencing method, and experimental data. The present method shows improvements in accuracy and convergence properties.

  15. Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

    NASA Technical Reports Server (NTRS)

    Clark, L. E.; Richie, C. B.

    1977-01-01

    The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

  16. 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.

  17. Review of delta wing space shuttle vehicle dynamics

    NASA Technical Reports Server (NTRS)

    Reding, J. P.; Ericsson, L. E.

    1971-01-01

    The unsteady aerodynamics of the proposed delta planform, high cross range, shuttle orbiters, are investigated. It is found that these vehicles are subject to five unsteady-flow phenomena that could compromise the flight dynamics. The phenomena are as follows: (1) leeside shock-induced separation, (2) sudden leading-edge stall, (3) vortex burst, (4)bow shock-flap shock interaction, and (5) forebody vorticity. Trajectory shaping is seen as the most powerful means of avoiding deterimental effects of the stall phenomena; however, stall must be fixed or controlled when traversing the stall region. Other phenomana may be controlled by carefully programmed control deflections and some configuration modifications. Ways to alter the occurrence of the various flow conditions are explored.

  18. An Attached Flow Design of a Noninterferring Leading Edge Extension to a Thick Delta Wing

    NASA Technical Reports Server (NTRS)

    Lamar, John E.; Ghaffari, Farhad

    1985-01-01

    An analytical procedure for the determination of the shape of a Leading-Edge Extension (LEE) which satisfies design criteria, including especially noninterference at the wing design point, has been developed for thick delta wings. The LEE device best satisfying all criteria is designed to be mounted on a wing along a dividing stream surface associated with an attached flow design lift coefficient (C(sub L,d)) of greater than zero. This device is intended to improve the aerodynamic performance of transonic aircraft at C(sub L) greater than C(sub L,d) system emanating from the LEE leading edge. In order to quantify this process a twisted and cambered thick delta wing was chosen for the initial application of this design procedure. Appropriate computer codes representing potential and vortex flows were employed to determine the dividing stream surface at C(sub L,d) and an optimized LEE planform shape at C(sub L) greater than C(sub L,d), respectively. To aid in the LEE selection, the aerodynamic effectiveness of 36 planforms was investigated at C(sub L) greater than C(sub L,d). This study showed that reducing the span of the candidate LEEs has the most detrimental effect on overall aerodynamic efficiency, regardless of the shape or area. Furthermore, for a fixed area, constant-chord LEE candidates were relatively more efficient than those with sweep less than the wing. At C(sub L,d), the presence of the LEE planform best satisfying the design criteria was found to have no effect on the wing alone aerodynamic performance.

  19. Experimental Investigation on a Pitching Motion Delta Wing in Unsteady Free Stream

    NASA Astrophysics Data System (ADS)

    Shi, Zhiwei; Ming, Xiao

    As combat aircraft becomes more and more maneuverable, the need to understand the unsteady behavior of aircraft in dynamic flow fields becomes more important. Usually researchers pay more attention to the effects on the changes of AOA, but ignore the effects of velocity variations. It is known that the velocity of aircraft changes greatly when the aircraft undergoes a high angle of attack maneuver, like "cobra" maneuver. To completely simulate and study the effect of rapid changes in both free stream velocity and angle of attack, a pitching motion setup is developed in the unsteady wind tunnel of NUAA. By measuring unsteady loads, unsteady pressure distribution and flow visualization, the unsteady aerodynamic behavior of a pitching isolated delta wing and the pitching delta wing coupled with unsteady free stream are investigated. It is found that the oscillating free stream velocity affects the hysteresis characteristics of the pitching delta wing further. The pressure distribution and flow visualization measurements show that the changes in the structure of the leading-edge vortices are the main reason. These studies conclude that a good understanding of the unsteady aerodynamics is vitally important in the design of super-maneuverable aircraft.

  20. A Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion

    NASA Technical Reports Server (NTRS)

    Turner, Thomas R.

    1960-01-01

    An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.

  1. Unsteady aerodynamic analysis of space shuttle vehicles. Part 2: Steady and unsteady aerodynamics of sharp-edged delta wings

    NASA Technical Reports Server (NTRS)

    Ericsson, L. E.; Reding, J. P.

    1973-01-01

    An analysis of the steady and unsteady aerodynamics of sharp-edged slender wings has been performed. The results show that slender wing theory can be modified to give the potential flow static and dynamic characteristics in incompressible flow. A semiempirical approximation is developed for the vortex-induced loads, and it is shown that the analytic approximation for sharp-edged slender wings gives good prediction of experimentally determined steady and unsteady aerodynamics at M = 0 and M = 1. The predictions are good not only for delta wings but also for so-called arrow and diamond wings. The results indicate that the effects of delta planform lifting surfaces can be included in a simple manner when determining elastic launch vehicle dynamic characteristics. For Part 1 see (N73-32763).

  2. Analytical observations on the aerodynamics of a delta wing with leading edge flaps

    NASA Technical Reports Server (NTRS)

    Oh, S.; Tavella, D.

    1986-01-01

    The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

  3. An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

    NASA Technical Reports Server (NTRS)

    Ghaffari, F.; Chaturvedi, S. K.

    1984-01-01

    An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

  4. Force and moment measurements on a 74 deg delta wing with an apex flap

    NASA Technical Reports Server (NTRS)

    Buter, T. A.; Rao, D. M.

    1984-01-01

    Results are presented of a subsonic experimental investigation of an apex flap concept on a 74 deg swept delta wing with trailing-edge flaps. The apex flap comprised approximately 6 percent of the wing area forward of a transverse hinge, allowing for upward and downward deflection angles from +40 deg to -20 deg. Upward deflection forces leading-edge vortex formation on the apex flap, resulting in an increased lift component on the apex area. The associated nose-up moment balances the nose-down moment due to trailing-edge flaps, resulting in sizeable increase in the trimmed lift coefficient particularly at low angles of attack. Nose-down apex deflection may be used to augment the pitch control for rapid recovery from high-alpha maneuvers. This report presents the balance data without analysis.

  5. A Computational and Experimental Investigation of a Delta Wing with Vertical Tails

    NASA Technical Reports Server (NTRS)

    Krist. Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

    2004-01-01

    The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 and a Reynolds number of 500; 000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

  6. A computational and experimental investigation of a delta wing with vertical tails

    NASA Technical Reports Server (NTRS)

    Krist, Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

    1993-01-01

    The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 deg and a Reynolds number of 500,000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

  7. Three-dimensional simulation of slender delta wing rock and divergence

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1992-01-01

    Computational simulation of three-dimensional flows around a delta wing undergoing rock and roll-divergence motions is presented. The problem is a multidisciplinary one where fluid-dynamics equations and rigid-body-dynamics equations are sequentially solved. For the fluid-dynamics part, the unsteady Euler equations, which are written relative to a moving frame of reference, are solved using an implicit, approximately-factored, central-difference, finite-volume scheme. For the rigid-body dynamics part, the Euler equation of rigid-body rolling motion is solved using a four-stage Runge-Kutta scheme. Since the applications do not include deforming wings or relative-rigid-body motions, the computational-fluid-dynamics grid, which is fixed in the moving frame of reference, does not need to be updated once it is generated.

  8. Adaptive computations of flow around a delta wing with vortex breakdown

    NASA Technical Reports Server (NTRS)

    Modiano, David L.; Murman, Earll M.

    1993-01-01

    An adaptive unstructured mesh solution method for the three-dimensional Euler equations was used to simulate the flow around a sharp edged delta wing. Emphasis was on the breakdown of the leading edge vortex at high angle of attack. Large values of entropy, which indicate vortical regions of the flow, specified the region in which adaptation was performed. The aerodynamic normal force coefficients show excellent agreement with wind tunnel data measured by Jarrah, and demonstrate the importance of adaptation in obtaining an accurate solution. The pitching moment coefficient and the location of vortex breakdown are compared with experimental data measured by Hummel and Srinivasan, showing good agreement in cases in which vortex breakdown is located over the wing.

  9. Surface pressure distributions on a delta wing undergoing large amplitude pitching oscillations. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Thompson, Scott A.

    1989-01-01

    Wind tunnel experiments were performed on a 70 deg sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Twelve pressure taps were placed from 35 to 90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as a function of Reynolds number and pitch rate. The pressure coefficient was seen to vary at approximately the same frequency as the pitching frequency. The relative pressure variation at each chord location was comparable for each case. The average pressure distribution through each periodic motion was near the static distribution for the average angle of attack. Upon comparing the upstroke and downstroke pressures for a specific angle of attack, the downstroke pressures were slightly larger. Vortex breakdown was seen to have the most significant effect at the 40 to 45 percent chord location, where a decrease in pressure was apparent.

  10. An evaluation of the relative merits of wing-canard, wing-tail, and tailless arrangements for advanced fighter applications

    NASA Technical Reports Server (NTRS)

    Nicholas, W. U.; Naville, G. L.; Hoffschwelle, J. E.; Huffman, J. K.; Covell, P. F.

    1984-01-01

    Two sets of wind tunnel tests were performed to examine the relative merits of wing-canard, wing-tail and tailless configurations for advanced fighters. Both sessions focused on variable camber using automated, prescheduled leading and trailing edge flap positioning. The trials considered a modified F-16 tail and canard configuration at subsonic, transonic and supersonic speeds, a 60 deg delta wing sweep, a 44 deg leading edge trapezoidal wing at subsonic and supersonic speeds, vortex flow effects, and flow interactions in the canard-wing-tail-tailless variations. The results showed that large negative stabilities would need to be tolerated in wing-canard arrangements to make them competitive with wing-tail arrangements. Subsonic polar shapes for canard and tailless designs were more sensitive to static design margins than were wing-tail arrangements. Canards provided better stability at supersonic speeds. The static margin limits were a critical factor in control surface selection. Finally, a tailless delta wing configuration exhibited the lowest projected gross take-off weight and drag values.

  11. Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2002-01-01

    A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  12. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  13. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.

  14. Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.

  15. Reynolds Number and Leading-Edge Bluntness Effects on a 65 Deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2002-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  16. Measurement of Leading Edge Vortices from a Delta Wing Using a Three Component Laser Velocimeter

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Hepner, Timothy E.

    1988-01-01

    A demonstration of the capabilities of a three component laser velocimeter to provide a detailed experimental database of a complex flow field i s presented. The orthogonal three component laser velocimeter was used to measure the leading edge vortex flow field above a 75 degrees delta wing at angles-of-attack of 20.5 degrees and 40.0 degrees. The resulting mean velocity and turbulence intensity measurements are presented. The laser velocimeter is described in detail including a description of the data processing algorithm. A full error analysis was conducted and the results presented.

  17. Compressibility and Leading-Edge Bluntness Effects for a 65 Deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2004-01-01

    A 65 deg. delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the compressibility and bluntness effects primarily at a Reynolds number of 6 million from this data set. Emphasis is placed upon on the onset and progression of leading-edge vortex separation, and compressibility is shown to promote this separation. Comparisons with recent publications show that compressibility and Reynolds number have opposite effects on blunt leading edge vortex separation

  18. Numerical simulation of the vortical flow over a round-edged double-delta wing

    NASA Technical Reports Server (NTRS)

    Hsu, C.-H.; Liu, C. H.

    1988-01-01

    Simulations of three-dimensional vortical flows over a thin double-delta wing with an aspect ratio of 2.05 have been performed. Steady-state solutions to the unsteady incompressible Navier-Stokes equations are obtained using a new implicit upwind-relaxation finite-difference scheme. The method is second-order accurate spatially and naturally dissipative. Numerical results indicate that key features of both vortical interaction and vortex breakdown are successfully simulated. The computed lift coefficients and lateral trajectories of the vortical cores are in good agreement with the experimental data.

  19. Visualization and flow surveys of the leading edge vortex structure on delta wing planforms

    NASA Technical Reports Server (NTRS)

    Payne, F. M.; Ng, T. T.; Nelson, R. C.; Schiff, L. B.

    1986-01-01

    In the present experimental investigation of thin delta wing vortex breakdown, for the cases of sweep angles of 70, 75, 80, and 85 deg, and smoke flow visualization/laser light sheet technique is used to obtain cross sectional views of the leading edge vortices as they break down. A combination of lateral and longitudinal cross sectional views furnishes data on the three-dimensional character of the vortex before, during, and after breakdown. Velocity measurements conducted with a laser Doppler anemometer on the 70 deg sweep delta, at 30 deg angle-of-attack, indicate that when breakdown occurs the core flow is transformed from a jet-like to a wake-like flow.

  20. Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.

    2000-01-01

    Steady and unsteady measured pressures for a Clipped Delta Wing (CDW) undergoing pitching oscillations and trailing-edge control surface oscillations have been presented . From the several hundred compiled data points, 22 static cases, 12 pitching-oscillation cases, and 12 control-surface-oscillation cases have been proposed for Computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The planform for this wing was derived by simplifying the planform of a proposed design for a supersonic transport which is described as the Boeing 2707-300. The strake was deleted, the resulting planform was approximated by a trapezoid with an unswept trailing edge, and the twist and camber were removed. In order to facilitate pressure instrumentation, the thickness was increased to 6 percent from the typical 2.5 to 3 percent for the supersonic transport. The airfoil is thus a symmetrical circular arc section with t/c = 0.06. A wing of similar planform but with a thinner airfoil of t/c = 0.03 was used in the flutter investigations, and the buffet and stall flutter investigation . Flutter results are also reported both for the 3 per cent thick simplified wing and for a more complex SST model. One of the consequences of the increased thickness of the clipped delta wing is that transonic effects are enhanced for Mach numbers near one. They are significantly stronger than would be the case for the thinner wing. Also, with the combination of high leading edge sweep of 50.5, and the sharp leading edge, a leading edge vortex forms on the wing at relatively low angles of attack, on the order of three degrees. The Appendix discusses some of the vortex flow effects. In addition, a shock develops over the aft portion of the wing at transonic speeds such that at some angles of attack, there is both a leading edge vortex and a shock wave on the wing. Such cases are a computational challenge. Some previous applications of this data set have been

  1. A note on the drag due to lift of delta wings at Mach numbers up to 2.0

    NASA Technical Reports Server (NTRS)

    Osborne, Robert S; Kelly, Thomas C

    1953-01-01

    In order to indicate the effects of Reynolds number and other variables on the drag due to lift of delta wings for Mach numbers up to 2.0, the results of several investigations of wing-body combinations employing delta wings with aspect ratios from 2 to 4 have been assembled for comparison. Effects of Reynolds number, leading-edge radius, and thickness ratio could be correlated with Reynolds number based on the leading-edge radius as a parameter. The results indicated that leading-edge Reynolds number effects were large at low speeds, but decreased with increases in Mach number. The effects of aspect ratio, wing modifications, and trim requirements are discussed.

  2. Static and unsteady pressure measurements on a 50 degree clipped delta wing at M = 0.9

    NASA Technical Reports Server (NTRS)

    Hess, R. W.; Wynne, E. C.; Cazier, F. W.

    1982-01-01

    Static and unsteady pressures were measured on a 50.45 degree clipped delta wing in the Langley Transonic Dynamics Tunnel with Freon as the test medium. Data taken at M = 0.9 is presented for static and oscillatory deflections of the trailing edge control surface and for the wing in pitch. Comparisons of the static measured data are made with results computed using the Bailey-Ballhaus small disturbance code.

  3. Twin Tail/Delta Wing Configuration Buffet Due to Unsteady Vortex Breakdown Flow

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Sheta, Essam F.; Massey, Steven J.

    1996-01-01

    The buffet response of the twin-tail configuration of the F/A-18 aircraft; a multidisciplinary problem, is investigated using three sets of equations on a multi-block grid structure. The first set is the unsteady, compressible, full Navier-Stokes equations. The second set is the coupled aeroelastic equations for bending and torsional twin-tail responses. The third set is the grid-displacement equations which are used to update the grid coordinates due to the tail deflections. The computational model consists of a 76 deg-swept back, sharp edged delta wing of aspect ratio of one and a swept-back F/A-18 twin-tails. The configuration is pitched at 32 deg angle of attack and the freestream Mach number and Reynolds number are 0.2 and 0.75 x 10(exp 6) respectively. The problem is solved for the initial flow conditions with the twin tail kept rigid. Next, the aeroelastic equations of the tails are turned on along with the grid-displacement equations to solve for the uncoupled bending and torsional tails response due to the unsteady loads produced by the vortex breakdown flow of the vortex cores of the delta wing. Two lateral locations of the twin tail are investigated. These locations are called the midspan and inboard locations.

  4. Navier-Stokes computations of lee-side flows over delta wings

    NASA Technical Reports Server (NTRS)

    Thomas, J. L.; Newsome, R. W.

    1986-01-01

    Solutions to the Navier-Stokes equations for the flow over delta wings are computed with emphasis on the separated vortical flows developing on the lee side at high angles of attack. A recently developed implicit algorithm is used which employs upwind differencing for the pressure and convection terms and central differencing for the shear stress and heat transfer terms. Solutions to both the three-dimensional equations and the approximate conical flow equations are compared parametrically with an extensive experimental data base at supersonic speeds. The computations indicate that the conical flow approximation provides results in close agreement with the three-dimensional equations, even to angles of attack as high as 20 degrees. Good agreement with experimentally measured pressures and vapor screen photographs is obtained for the conditions investigated. The method predicts the classical pattern of vortical flow over a delta wing and transition to other flow patterns as the leading edge sweep angle and leading edge normal Mach number are varied.

  5. Unsteady supersonic flow around delta wings with symmetric and asymmetric flaps oscillation

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1991-01-01

    A parametric study is presented to investigate the effect of reduced frequency of the leading-edge flaps on the locally-conical, unsteady, supersonic flow around a delta wing. This study covers symmetric and antisymmetric forced oscillation of the leading-edge flaps. The effects of the freestream Mach number and angle of attack are also presented. The problem is solved using time-accurate integration of the unsteady, compressible, thin-layer Navier-Stokes equations and the unsteady, linearized, Navier-displacement equations. The delta wing is of aspect ratio of 1.5 and its leading-edge flaps are hinged at 65 percent of the local-half span length. The reduced frequency is varied between 2 pi and pi/2. Two supersonic flow conditions have been investigated; the first is for a freestream Mach number of 2.4 and an angle of attack of 19 deg and the second is for a freestream Mach number of 1.5 and an angle attack of 15 deg.

  6. A study of the vortex flow over 76/40-deg double-delta wing

    NASA Technical Reports Server (NTRS)

    Verhaagen, N. G.; Jenkins, L. N.; Kern, S. B.; Washburn, A. E.

    1995-01-01

    A low-speed wind-tunnel study of the flow about a 76/40-deg double-delta wing is described for angles of attack ranging from -10 to 25 deg and Reynolds numbers ranging from 0.5 to 1.5 Million. The study was conducted to provide data for the purpose of understanding the vortical flow behavior and for validating Computational Fluid Dynamics methods. Flow visualization tests have provided insight into the effect of the angle of attack and Reynolds number of the vortex-dominated flow both on and off of the surface of the double-delta wing. Upper surface pressure recordings from pressure orifices and Pressure Sensitive Paint have provided data on the pressures induced by the vortices. Flowfield surveys were carried out at an angle of attack of 10 deg by using a thin 5-hole probe. Numerical solutions of the compressible thin-layer Navier-Stokes equations were conducted and compared to the experimental data.

  7. Aeroelastic prediction of the limit cycle oscillations of a cropped delta wing

    NASA Astrophysics Data System (ADS)

    Attar, P. J.; Gordnier, R. E.

    2006-01-01

    The flutter and limit cycle oscillation (LCO) behavior of a cropped delta wing are investigated using a newly developed computational aeroelastic solver. This computational model includes a well-validated Euler finite difference solver coupled to a high-fidelity finite element structural solver. The nonlinear structural model includes geometric nonlinearities which are modelled using a co-rotational formulation. The LCOs of the cropped delta wing are computed and the results are compared to previous computations and to experiment. Over the range of dynamic pressures for which experimental results are reported, the LCO magnitudes computed using the current model are comparable to those from a previous computation which used a lower-order von Karman structural model. However, for larger dynamic pressures, the current computational model and the model which used the von Karman theory start to differ significantly, with the current model predicting larger deflections for a given dynamic pressure. This results in a LCO curve which is in better qualitative agreement with experiment. Flow features which were present in the previous computational model such as a leading edge vortex and a shock wave are enhanced in the current model due to the prediction of larger deflections and rotations at the higher dynamic pressures.

  8. Controlled vortical flow on delta wings through unsteady leading edge blowing

    NASA Technical Reports Server (NTRS)

    Lee, K. T.; Roberts, Leonard

    1990-01-01

    The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions.

  9. Detailed flow-field measurements over a 75 deg swept delta wing

    NASA Technical Reports Server (NTRS)

    Kjelgaard, Scott O.; Sellers, William L., III

    1990-01-01

    Results from an experimental investigation documenting the flowfield over a 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0, and 1.5 million based on the root chord were conducted with both a Pitot pressure probe and a 5-hole pressure probe; and 3-component laser velocimeter surveys were conducted at a Reynolds number of 1.0 million. The Pitot pressure surveys were obtained at 5 chordwise stations, the 5-hole probe surveys were obtained at 3 chordwise stations and the laser velocimeter surveys were obtained at one station. The results confirm the classical roll up of the flow into a pair of primary vortices over the delta wing. The velocity measurements indicate that Reynolds number has little effect on the global structure of the flowfield for the Reynolds number range investigated. Measurements of the non-dimensional axial velocity in the core of the vortex indicate a jet like flow with values greater than twice freestream. Comparisons between velocity measurements from the 5-hole pressure probe and the laser velocimeter indicate that the pressure probe does a reasonable job of measuring the flowfield quantities where the velocity gradients in the flowfield are low.

  10. Forced Rolling Oscillation of a 65 deg-Delta Wing in Transonic Vortex-Breakdown Flow

    NASA Technical Reports Server (NTRS)

    Menzies, Margaret A.; Kandil, Osama A.; Kandil, Hamdy A.

    1996-01-01

    Unsteady, transonic, vortex dominated flow over a 65 deg. sharp-edged, cropped-delta wing of zero thickness undergoing forced rolling oscillations is investigated computationally. The wing angle of attack is 20 deg. and the free stream Mach number and Reynolds number are 0.85 and 3.23 x 10(exp 6), respectively. The initial condition of the flow is characterized by a transverse terminating shock which induces vortex breakdown of the leading edge vortex cores. The computational investigation uses the time accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux difference splitting, finite-volume scheme. While the maximum roll amplitude is kept constant at 4.0 deg., both Reynolds number and roll frequency are varied covering three cases of forced sinusoidal rolling. First, the Reynolds number is held at 3.23 x 10(exp 6) and the wing is forced to oscillate in roll around the axis of geometric symmetry at a reduced frequency of 2(pi). Second, the Reynolds number is reduced to 0.5 x 10(exp 6) to observe the effects of added viscosity on the vortex breakdown. Third, with the Reynolds number held at 0.5 x 10(exp 6), the roll frequency is reduced to 1(pi) to complete the study.

  11. Analysis of the Flow About Delta Wings with Leading Edge Separation at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Nenni, J. P.; Tung, C.

    1973-01-01

    A research program was conducted to develop an improved theoretical flow model for the flow about sharp edge delta wings with leading-edge separation at supersonic speeds. The flow model incorporates a representation of the secondary separation region which occurs just inboard of the leading edge on such wings and is based on a slender-wing theory whereby the full three-dimensional problem is reduced to a quasi two-dimensional problem in the cross-flow plane. The secondary separation region was modeled by a surface distribution of singularities or a linearized type of cavity representation. The primary vortex and separation were modeled by a concentrated vortex and cut in the cross-flow potential which represents its feeding sheet. The cross-flow solutions for the cavity model were obtained, but these solutions have physical significance only in a very restricted range of angle of attack. The reasons for the failure of the flow model are discussed. The analysis is presented so that other interested researchers may critically review the work.

  12. Effects of Horizontal-Control Planform and Wing-Leading-Edge Modification on Low-Speed Longitudinal Aerodynamic Characteristics of a Canard Airplane Configuration

    NASA Technical Reports Server (NTRS)

    Spencer, Bernard, Jr.

    1981-01-01

    An investigation at low subsonic speeds has been conducted in the Langley 300-MPH 7- by 10-foot tunnel. The basic wing had a trapezoidal planform, an aspect ratio of 3.0., a taper ratio of 0.143, and an unswept 80-percent-chord line. Modifications to the basic wing included deflectable full-span and partial-span leading-edge chord-extensions. A trapezoidal horizontal control similar in planform to the basic wing and a 60 deg sweptback delta horizontal control were tested in conjunction with the wing. The total planform area of each horizontal control was 16 percent of the total basic-wing area. Modifications to these horizontal controls included addition of a full-span chord-extension to the trapezoidal planform and a fence to the delta planform.

  13. Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Gonzalez, Hugo A.

    2004-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76o/40o double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M =0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

  14. Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Gonzalez, Hugo A.

    2005-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76 deg/40 deg double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 30 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M = 0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

  15. Open-Type Separation on Delta Wings for Leading-Edge Bluntness

    NASA Astrophysics Data System (ADS)

    Chiba, Kazuhisa; Obayashi, Shigeru; Nakahashi, Kazuhiro

    A numerical simulation was carried out corresponding to recent experiments using delta wings with sharp and blunt leading-edges, which indicates the second primary vortex, at NASA Langley Research Center. However, the experimental data did not reveal the detailed physical phenomena regarding the second primary vortex, because the experiment used only on-the-body-surface data. In the present study, the physical phenomena were revealed using Reynolds-averaged Navier-Stokes computations with three one-equation turbulence models on an unstructured hybrid mesh. The adaptive mesh refinement method in the vicinity of the vortex center was also applied to have more mesh resolution. Consequently, the result quantitatively revealed that appropriate modeling regarding turbulent kinematic viscosity was significant. Moreover, the three-dimensional visualization of the computational fluid dynamics results suggested that the second primary vortex was a developing shear layer merging into an open-type separation generated late by the primary vortex.

  16. Calculation of supersonic viscous flow over delta wings with sharp subsonic leading edges

    NASA Technical Reports Server (NTRS)

    Vigneron, Y. C.; Rakich, J. V.; Tannehill, J. C.

    1978-01-01

    Two complementary procedures were developed to calculate the viscous supersonic flow over conical shapes at large angles of attack, with application to cones and delta wings. In the first approach the flow is assumed to be conical and the governing equations are solved at a given Reynolds number with a time-marching explicit finite-difference algorithm. In the second method the parabolized Navier-Stokes equations are solved with a space-marching implicit noniterative finite-difference algorithm. This latter approach is not restricted to conical shapes and provides a large improvement in computational efficiency over published methods. Results from the two procedures agree very well with each other and with available experimental data.

  17. Thickening oscillation of a delta wing using Navier-Stokes and Navier-displacement equations

    NASA Technical Reports Server (NTRS)

    Chuang, Hsin-Kung A.; Kandil, Osama A.

    1989-01-01

    The problem of unsteady, supersonic, locally conical, vortical flow around a delta wing undergoing thickening oscillation is solved using the unsteady, thin-layer Navier-Stokes equations and the unsteady linearized Navier-displacement equations. The unsteady, thin-layer Navier-Stokes equations are solved using the implicit approximate-factorization finite-volume scheme to compute the conservative components of the flow vector field. With the conservative components known at any time step, the linearized, Navier-displacement equations are solved using the alternating, direction-implicit scheme to obtain the grid points displacements due to known displacement boundary conditions. A grid-displacements limiter, in the form of a low mesh Reynolds number, is used to limit grid-folding in regions of highly reversed flow.

  18. Transonic Navier-Stokes calculations about a 65 deg delta wing

    NASA Technical Reports Server (NTRS)

    Londenberg, W. Kelly

    1994-01-01

    A computational study has been conducted in which the CFL3D Navier-Stokes solver coupled with an algebraic and a one-equation nonequilibrium turbulence model has been used to predict the flow over a 65 degree delta wing at transonic conditions for Reynolds numbers ranging from 6 x 10(exp 6) to 120 x 10(exp 6) based on mean aerodynamic chord. Solutions obtained indicated that the computational method when used with the one-equation turbulence model predicts results that compare well with experiment for attached flow conditions. Comparisons with experimental pressure at separated conditions show that the computational method, even though primary flow-field features are predicted well, does not predict secondary flow features.

  19. Numerical study of the effect of tangential leading edge blowing on delta wing vortical flow

    NASA Technical Reports Server (NTRS)

    Yeh, David T.; Tavella, Domingo A.; Roberts, Leonard; Fujii, Kozo

    1989-01-01

    A numerical simulation of tangential blowing along the leading edge of a delta wing is analyzed as a means of controlling the position and strength of the leading-edge vortices. The computation is done by numerical solutions of the three-dimensional thin-layer Navier-Stokes equations. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of tangential leading-edge blowing at low to moderate angles of attack tends to reduce the pressure peaks associated with leading-edge vortices and to increase the suction peak around the leading edge, such that the integrated value of the surface pressure remains about the same.

  20. The lateral-directional characteristics of a 74-degree Delta wing employing gothic planform vortex flaps

    NASA Technical Reports Server (NTRS)

    Grantz, A. C.

    1984-01-01

    The low speed lateral/directional characteristics of a generic 74 degree delta wing body configuration employing the latest generation, gothic planform vortex flaps was determined. Longitudinal effects are also presented. The data are compared with theoretical estimates from VORSTAB, an extension of the Quasi vortex lattice Method of Lan which empirically accounts for vortex breakdown effects in the calculation of longitudinal and lateral/directional aerodynamic characteristics. It is indicated that leading edge deflections of 30 and 40 degrees reduce the magnitude of the wing effective dihedral relative to the baseline for a specified angle of attack or lift coefficient. For angles of attack greater than 15 degrees, these flap deflections reduce the configuration directional stability despite improved vertical tail effectiveness. It is shown that asymmetric leading edge deflections are inferior to conventional ailerons in generating rolling moments. VORSTAB calculations provide coarse lateral/directional estimates at low to moderate angles of attack. The theory does not account for vortex flow induced, vertical tail effects.

  1. Unsteady surface pressure distributions on a delta wing undergoing large amplitude pitching motions

    NASA Technical Reports Server (NTRS)

    Thompson, S. A.; Batill, S. M.; Nelson, R. C.

    1990-01-01

    Wind tunnel experiments were performed on a 70-deg-sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Pressure taps were placed from 35-90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as functions of Reynolds number and pitch rate. The surface pressure distribution was seen to vary at the same frequency as the pitching frequency, though distortion due to the vortex breakdown was observed. Comparing the upstroke (angle of attack increasing) and downstroke (angle of attack decreasing) pressures for a specific angle of attack, a time lag in the pressure distribution was observed. The downstroke pressures were slightly larger at the forward chord locations. Vortex breakdown was seen to have the most significant effect at the 40-45-percent chord location, where an increase in local pressure was apparent, as well as a distortion of the periodic pressure fluctuation.

  2. Nonequilibrium turbulence modeling effects on transonic vortical flows about delta wings

    NASA Technical Reports Server (NTRS)

    Kaynak, Unver; Tu, Eugene; Dindar, Mustafa; Barlas, Remzi

    1991-01-01

    The Johnson-King turbulence model that is a viable method for calculating two dimensional transonic separated flows was extended into three dimensions. The implementation was done for Navier Stokes flow solvers written in general curvilinear coordinates. The present approach used in turbulence modeling is based on streamwise integration of an ordinary differential equation (o.d.e.) that governs the maximum Reynolds shear stress behavior. Streamwise integration of the o.d.e. approach was found to offer great mathematical simplicity and economy for three dimensional Navier Stokes methods. Thus, the new method is quick, simple, and very cheap. The new method was first checked against the data of a well known transonic axisymmetric bump experiment, and a good agreement was obtained. Later, the new method was used to compute the flow around a low aspect ratio wing in a transonic wind tunnel. Finally it was employed to study the nonequilibrium turbulence effects on the transonic vortical flows about a 65 deg sweep round leading edge delta wing.

  3. Reynolds number effects on the vortical-flow structure generated by a double-delta wing

    NASA Astrophysics Data System (ADS)

    Hebbar, S. K.; Platzer, M. F.; Fritzelas, A. E.

    An experimental investigation of the high-incidence vortical flowfield over a 76/40° double-delta wing model with sharp leading edges was conducted in the Naval Postgraduate School water tunnel facility at three nominal flow Reynolds numbers of 15000, 45000, and 75000 (based on centerline chord). Extensive flow visualization studies were performed with the dye-injection technique, followed by laser Doppler velocity measurements. The primary objective of this investigation was the determination of the influence of Reynolds number on vortex interactions/trajectories, and breakdown. It was found that there is a significant influence of Reynolds number. Specifically, with the increase of flow Reynolds number the strake and wing vortex trajectories tend to move outboards and closer to the model surface, and the vortex breakdown location moves forwards toward the apex of the model. The intertwining or coiling-up feature of the vortex interaction phenomenon becomes less dominant and disappears altogether at high Reynolds numbers. These trends in the vortex interaction and bursting data are found to be in good agreement with previous wind tunnel data.

  4. A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

    NASA Technical Reports Server (NTRS)

    Rao, Dhanvada M.; Bhat, M. K.

    1992-01-01

    A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift.

  5. Flow Structure on a Delta Wing of Moderate Sweep Angle During and After Pitch-Up Maneuver

    NASA Astrophysics Data System (ADS)

    Celik, Alper; Ozturk, Ilhan; Tunc, Habib Can; Yavuz, Mehmet Metin

    2012-11-01

    The flow structure over a moderate sweep angle delta wing is investigated during and after the pitch-up maneuver and compared to the corresponding stationary wing results. The effects of pitch-up rate and Reynolds number on flow structures and their transformations are also studied. Dye visualization is used for qualitative studies and particle image velocimetry is used for quantitative analysis. At early stages of the maneuver the transformation of flow is initiated by the intertwinement of the dual vortex structure. Increasing the angle of attack results in disappearance of the vortex located closest to the leading-edge of the wing which results in a single, large scale, leading-edge vortex that undergoes a distinctive form of breakdown. It is found that the motion of the wing creates a significant time lag on the development of the flow pattern, when compared to stationary wing. In the relaxation period, the vorticity concentrations become more diffuse and elongated as they move towards the plane of symmetry, and away from the surface of the wing. All these features and transformations occur irrespective of values of pitch rates and Reynold's number. On the other hand, it is seen that the lag of flow pattern is a function of pitch rate and Reynold's number.

  6. Exhaust Plume Effects on Sonic Boom for a Delta Wing and a Swept Wing-Body Model

    NASA Technical Reports Server (NTRS)

    Castner, Raymond; Lake, Troy

    2012-01-01

    Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic boom. Research has been performed on sonic boom reduction and has included the contribution of the exhaust nozzle plume. Plume effect on sonic boom has progressed from the study of isolated nozzles to a study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline analysis of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field pressure profile. Reductions occurred in the peak-to-peak magnitude of the pressure profile for a swept wing-body vehicle. The exhaust plumes also had a favorable effect as the nozzles were moved outward along the wing-span.

  7. Piezo-stack vortex generators for boundary layer control of a delta wing micro-aerial vehicle

    NASA Astrophysics Data System (ADS)

    Mystkowski, Arkadiusz

    2013-11-01

    This paper presents an idea for the control of flow separation over solid surfaces by piezo-stack vortex generators. The vortex generators are small vibrating plates attached to the delta wing surface. A model of the micro-aerial vehicle (MAV) controlled by vortex piezo-generators is presented. The vortex generators are applied to produce the appropriate aerodynamical forces and moments controlling the flight of the aircraft. The efficiency of the vortex generators is proved by the wind tunnel test results. The oscillatory added lift and drag coefficients versus angle of attack are presented. The optimal vortex generator amplitude and frequency are investigated. Boundary layer control (BLC) for delta wing micro-aircraft increases the manoeuvrability and performance of the MAV.

  8. Myoblast cytonemes mediate Wg signaling from the wing imaginal disc and Delta-Notch signaling to the air sac primordium

    PubMed Central

    Huang, Hai; Kornberg, Thomas B

    2015-01-01

    The flight muscles, dorsal air sacs, wing blades, and thoracic cuticle of the Drosophila adult function in concert, and their progenitor cells develop together in the wing imaginal disc. The wing disc orchestrates dorsal air sac development by producing decapentaplegic and fibroblast growth factor that travel via specific cytonemes in order to signal to the air sac primordium (ASP). Here, we report that cytonemes also link flight muscle progenitors (myoblasts) to disc cells and to the ASP, enabling myoblasts to relay signaling between the disc and the ASP. Frizzled (Fz)-containing myoblast cytonemes take up Wingless (Wg) from the disc, and Delta (Dl)-containing myoblast cytonemes contribute to Notch activation in the ASP. Wg signaling negatively regulates Dl expression in the myoblasts. These results reveal an essential role for cytonemes in Wg and Notch signaling and for a signal relay system in the myoblasts. DOI: http://dx.doi.org/10.7554/eLife.06114.001 PMID:25951303

  9. A Note on the Drag Due to Lift of Delta Wings at Mach Numbers up to 2.0

    NASA Technical Reports Server (NTRS)

    Osborne, Robert S.; Kelly, Thomas C.

    1960-01-01

    In order to indicate the effects of Reynolds number and other variables on the drag due to lift of delta wings for Mach numbers up to 2.0, the results of several investigations of wing-body combinations having plane delta wings with aspect ratios from 2 to 4 have been assembled for comparison and brief analysis. The effects of Reynolds number, leading-edge radius, and thickness ratio could generally be correlated with Reynolds number based on the leading-edge radius as a parameter. The effects of leading-edge Reynolds number on drag due to lift were large at Mach numbers less than 0.25. However, with increases in Mach number, the effects decreased and were almost negligible at a Mach number of 2.0. and trimming were large, as would be expected. The effects of aspect ratio and trimming were large, as would be expected. It was indicated at least for subsonic and transonic speeds that improvement in the drag due to lift might be obtained from wing modifications designed to inhibit flow separation.

  10. Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Rogers, Lawrence W.

    1992-01-01

    A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

  11. Effect of configuration modification on the hypersonic aerodynamic characteristics of a blended delta wing-body entry vehicle

    NASA Technical Reports Server (NTRS)

    Arrington, J. P.; Ashby, G. C., Jr.

    1972-01-01

    The longitudinal, lateral, and directional aerodynamic characteristics of a delta-wing configuration were obtained experimentally at Mach 20 in helium with Reynolds numbers, based on model length, of 1.5 million and 2.9 million and at a Mach number of 6 in air with a Reynolds number, based on model length, of 4.8 million. The angles of attack varied from 0 deg to 55 deg for two sideslip angles. The effects of the addition of dorsal fins, the removal of wing tip fins, an increase in elevon span, and changes in elevon hinge-line sweep angle are discussed. The unmodified vehicle had a maximum lift-drag ratio of 2.1 at Mach 19 and of 2.4 at Mach 6 with about the same lateral and directional stability level at both Mach numbers. As the Mach number increased from 6 to 20, the longitudinal center of pressure moved forward and more positive elevon deflection was therefore required to maintain a given trim angle. The removal of wing tip fins increased the maximum lift-drag ratio and had a negligible effect on longitudinal stability, but caused directional instability that was not corrected by the dorsal fins examined. The shape of the wing and elevon hinge-line sweep had a large influence on the induced yawing moment due to roll control.

  12. Steady and unsteady transonic pressure measurements on a clipped delta wing for pitching and control-surface oscillations

    NASA Technical Reports Server (NTRS)

    Hess, Robert W.; Cazier, F. W., Jr.; Wynne, Eleanor C.

    1986-01-01

    Steady and unsteady pressures were measured on a clipped delta wing with a 6-percent circular-arc airfoil section and a leading-edge sweep angle of 50.40 deg. The model was oscillated in pitch and had an oscillating trailing-edge control surface. Measurements were concentrated over a Mach number range from 0.88 to 0.94; less extensive measurements were made at Mach numbers of 0.40, 0.96, and 1.12. The Reynolds number based on mean chord was approximately 10 x 10 to the 6th power. The interaction of wing or control-surface deflection with the formation of shock waves and with a leading-edge vortex generated complex pressure distributions that were sensitive to frequency and to small changes in Mach number at transonic speeds.

  13. Effects of Coupled Rolling and Pitching Oscillations on Transonic Shock-Induced Vortex-Breakdown Flow of a Delta Wing

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Menzies, Margaret A.

    1996-01-01

    Unsteady, transonic vortex-breakdown flow over a 65 deg. sharp edged, cropped-delta wing of zero thickness undergoing forced coupled pitching and rolling oscillations is investigated computationally. The initial condition of the flow is characterized by a transverse terminating shock which induces of the leading edge vortex cores to breakdown. The computational investigation uses the time-accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme. The main focus is to analyze the effects of coupled motion on the wing response and vortex-breakdown flow by varying oscillation frequency and phase angle while keeping the maximum pitch and roll amplitude equal.

  14. 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.

  15. Heavy metals in the near-surface aerosol over the Atlantic Ocean from 60 deg South to 54 deg North

    NASA Astrophysics Data System (ADS)

    Voelkening, Joachim; Heumann, Klaus G.

    1990-11-01

    Results are presented on determinations (by isotope dilution mass spectrometry) of the particulate heavy metal concentrations of Cr, Fe, Ni, Cu, Zn, Cd, Tl, and Pb in the atmospheric samples collected during the cruise of the Polarstern research ship over the Atlantic Ocean from 60 deg S to 54 deg N. The lowest abundance of all heavy metals was exhibited in the remote areas of the South Atlantic, whereas the highest abundances were detected around the English Channel. The data clearly identify the anthropogenic influence, e.g., over the South Atlantic near the La Plata area, the area around the English Channel, and over the North Sea.

  16. Static internal performance of a two-dimensional convergent nozzle with thrust-vectoring capability up to 60 deg

    NASA Technical Reports Server (NTRS)

    Leavitt, L. D.

    1985-01-01

    An investigation was conducted at wind-off conditions in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance characteristics of a two-dimensional convergent nozzle with a thrust-vectoring capability up to 60 deg. Vectoring was accomplished by a downward rotation of a hinged upper convergent flap and a corresponding rotation of a center-pivoted lower convergent flap. The effects of geometric thrust-vector angle and upper-rotating-flap geometry on internal nozzle performance characteristics were investigated. Nozzle pressure ratio was varied from 1.0 (jet off) to approximately 5.0.

  17. Static internal performance of single-expansion-ramp nozzles with thrust-vectoring capability up to 60 deg

    NASA Technical Reports Server (NTRS)

    Berrier, B. L.; Leavitt, L. D.

    1984-01-01

    An investigation has been conducted at static conditions (wind off) in the static-test facility of the Langley 16-Foot Transonic Tunnel. The effects of geometric thrust-vector angle, sidewall containment, ramp curvature, lower-flap lip angle, and ramp length on the internal performance of nonaxisymmetric single-expansion-ramp nozzles were investigated. Geometric thrust-vector angle was varied from -20 deg. to 60 deg., and nozzle pressure ratio was varied from 1.0 (jet off) to approximately 10.0.

  18. Experimental Determination of Effects of Frequency and Amplitude on the Lateral Stability Derivatives for a Delta, a Swept, and Unswept Wing Oscillating in Yaw

    NASA Technical Reports Server (NTRS)

    Fisher, Lewis R

    1958-01-01

    Three wing models were oscillated in yaw about their vertical axes to determine the effects of systematic variations of frequency and amplitude of oscillation on the in-phase and out-of-phase combination lateral stability derivatives resulting from this motion. The tests were made at low speeds for a 60 degree delta wing, a 45 degree swept wing, and an unswept wing; the swept and unswept wings had aspect ratios of 4. The results indicate that large changes in the magnitude of the stability derivatives due to the variation of frequency occur at high angles of attack, particularly for the delta wing. The greatest variations of the derivatives with frequency take place for the lowest frequencies of oscillation; at the higher frequencies, the effects of frequency are smaller and the derivatives become more linear with angle of attack. Effects of amplitude of oscillation on the stability derivatives for delta wings were evident for certain high angles of attack and for the lowest frequencies of oscillation. As the frequency became high, the amplitude effects tended to disappear.

  19. Correlation of Predicted and Flight Derived Stability and Control Derivatives with Particular Application to Tailless Delta Wing Configurations

    NASA Technical Reports Server (NTRS)

    Weil, J.

    1981-01-01

    Flight derived longitudinal and lateral-directional stability and control derivatives were compared to wind-tunnel derived values. As a result of these comparisons, boundaries representing the uncertainties that could be expected from wind-tunnel predictions were established. These boundaries provide a useful guide for control system sensitivity studies prior to flight. The primary application for this data was the space shuttle, and as a result the configurations included in the study were those most applicable to the space shuttle. The configurations included conventional delta wing aircraft as well as the X-15 and lifting body vehicles.

  20. Computational Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Walker, Charlotte E.

    1999-01-01

    Computational test cases have been selected from the data set for a clipped delta wing with a six-percent-thick circular-arc airfoil section that was tested in the NASA Langley Transonic Dynamics Tunnel. The test cases include parametric variation of static angle of attack, pitching oscillation frequency, trailing-edge control surface oscillation frequency, and Mach numbers from subsonic to low supersonic values. Tables and plots of the measured pressures are presented for each case. This report provides an early release of test cases that have been proposed for a document that supplements the cases presented in AGARD Report 702.

  1. Large-Amplitude, High-Rate Roll Oscillations of a 65 deg Delta Wing at High Incidence

    NASA Technical Reports Server (NTRS)

    Chaderjian, Neal M.; Schiff, Lewis B.

    2000-01-01

    The IAR/WL 65 deg delta wing experimental results provide both detail pressure measurements and a wide range of flow conditions covering from simple attached flow, through fully developed vortex and vortex burst flow, up to fully-stalled flow at very high incidence. Thus, the Computational Unsteady Aerodynamics researchers can use it at different level of validating the corresponding code. In this section a range of CFD results are provided for the 65 deg delta wing at selected flow conditions. The time-dependent, three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate the unsteady vertical flow. Two sting angles and two large- amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental pressures, forces, moments and roll angle time history. In addition, surface and off-surface flow particle streaks are also presented.

  2. Coupled Rolling and Pitching Oscillation Effects on Transonic Shock-Induced Vortex-Breakdown Flow of a Delta Wing

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Menzies, Margaret A.

    1996-01-01

    Unsteady, transonic vortex dominated flow over a 65 deg. sharp edged, cropped-delta wing of zero thickness undergoing forced coupled pitching and rolling oscillations is investigated computationally. The wing mean angle of attack is 20 deg. and the free stream Mach number and Reynolds number are 0.85 and 3.23 x 10(exp 6), respectively. The initial condition of the flow is characterized by a transverse terminating shock and vortex breakdown of the leading edge vortex cores. The computational investigation uses the time-accurate solution of the laminar, unsteady, compressible, full Navier-Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite volume scheme. The main focus is to analyze the effects of coupled motion on the wing response and vortex breakdown flow by varying oscillation frequency and phase angle while the maximum pitch and roll amplitude is kept constant at 4.0 deg. Four cases demonstrate the following: simultaneous motion at a frequency of 1(pi), motion with a 90 deg. phase lead in pitch, motion with a rolling frequency of twice the pitching frequency, and simultaneous motion at a frequency of 2(pi). Comparisons with single mode motion at these frequencies complete this study and illustrate the effects of coupling the oscillations.

  3. Breaking down the delta wing vortex: The role of vorticity in the breakdown process. Ph.D. Thesis Final Report

    NASA Technical Reports Server (NTRS)

    Nelson, Robert C.; Visser, Kenneth D.

    1990-01-01

    Experimental x-wire measurements of the flowfield above a 70 and 75 deg flat plate delta wing were performed at a Reynolds number of 250,000. Grids were taken normal to the wing at various chordwise locations for angles of attack of 20 and 30 deg. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core and on chordwise location was also examined. The effects of nondimensionalization in comparison with other experimental data is made. The results indicate that the circulation distribution scales with the local semispan and grows in a nearly linear fashion in the chordwise direction. The spanwise distribution of axial vorticity is severely altered through the breakdown. The axial vorticity components with a negative sense, such as that found in the secondary vortex, seem to remain unaffected by changes in wind sweep or angle of attack, in direct contrast to the positive components. In addition, the inclusion of the local wing geometry into a previously derived correlation parameter allows the circulation of growing leading edge vortex flows to be reduced into a single curve.

  4. Experimental Investigation of the Flow about a 65 deg Delta Wing in the NASA Langley National Transonic Facility. Chapter 4

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2009-01-01

    An experimental investigation for the flow about a 65 deg. delta wing has been conducted in the NASA Langley National Transonic Facility (NTF). The tests were conducted at Reynolds numbers, based on the mean aerodynamic chord, ranging from 6 million to 120 million and at Mach numbers ranging from 0.4 to 0.9. The model incorporated four different leading-edge bluntness values. The data include detailed static surfacepressure distributions as well as normal-force and pitching-moment coefficients. The test program was designed to quantify the effects of Mach number, Reynolds number, and leading-edge bluntness on the onset and progression of leading-edge vortex separation.

  5. Investigation of flow structure on a stationary and pitching delta wing of moderate sweep angle using stereoscopic particle image velocimetry

    NASA Astrophysics Data System (ADS)

    Goruney, Tunc

    Near-surface flow patterns along a basic delta wing of moderate sweep angle, representative of key features of Unmanned Combat Air Vehicles (UCAVs) and Micro Air Vehicles (MAVs), are visualized by a technique of high-image-density digital particle image velocimetry (DPIV), which provides quantitative representations of the whole-field flow patterns. Due to the highly three-dimensional nature of the flow patterns, they are also visualized by stereoscopic particle image velocimetry (SPIV). Qualitative dye visualization is employed to complement the DPIV technique. The flow structure is represented by patterns of dye, velocity vectors, streamwise, transverse and out-of-plane velocity components, streamline topology and vorticity. The surface topology, i.e., surface streamlines, and patterns of surface velocity and vorticity oriented normal to the surface of the wing, are investigated by making use of topological rules and critical point theory. For the case of DPIV measurements, the focus is on the time evolution of the surface topology during relaxation of the flow after termination of a pitching maneuver, for a wide range of pitch rates. It is demonstrated that there exists a critical universal state, which marks an abrupt transformation between two distinctly different states of the near-surface pattern of critical points. Moreover, an approach that predicts the occurrence of three-dimensional separation from the surface of the wing, for a wide range of pitch rate, is introduced. For the case of SPIV measurements, the relationship between the three-dimensional flow structure above the surface of the wing and the near-surface topology along the wing has been established, at successive instants following termination of the maneuver. Features of the leading-edge vortex and its breakdown location were quantitatively determined at the termination of the pitching maneuver. For the relaxed state of the flow structure, there is a reference elevation above the wing surface

  6. Turbulent Vortex-Flow Simulation Over a 65 deg Sharp and Blunt Leading-Edge Delta Wing at Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Ghaffari, Farhad

    2005-01-01

    Turbulent thin-layer, Reynolds-Averaged Navier-Stokes solutions, based on a multi-block structured grid, are presented for a 65 deg delta wing having either a sharp leading edge (SLE) or blunt leading edge (BLE) geometry. The primary objective of the study is to assess the prediction capability of the method for simulating the leading-edge flow separation and the ensuing vortex flow characteristics. Computational results are obtained for two angles of attack of approximately 13 and 20 deg, at free-stream Mach number of 0.40 and Reynolds number of 6 million based on the wing mean aerodynamic chord. The effects of two turbulence models of Baldwin-Lomax with Degani-Schiff (BL/DS) and the Spalart-Allmaras (SA) on the numerical results are also discussed. The computations also explore the effects of two numerical flux-splitting schemes, i.e., flux difference splitting (fds) and flux vector splitting (fvs), on the solution development and convergence characteristics. The resulting trends in solution sensitivity to grid resolution for the selected leading-edge geometries, angles of attack, turbulence models and flux splitting schemes are also presented. The validity of the numerical results is evaluated against a unique set of experimental wind-tunnel data that was obtained in the National Transonic Facility at the NASA Langley Research Center.

  7. Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms

    NASA Technical Reports Server (NTRS)

    Holdaway, George H.; Mellenthin, Jack A.

    1960-01-01

    The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory

  8. Effects of wing position and fuselage size on the low-speed static rolling stability characteristics of a delta-wing model

    NASA Technical Reports Server (NTRS)

    Goodman, Alex; Thomas, David T , Jr

    1955-01-01

    An investigation was made to determine the effects of wing position and fuselage size on the low-speed static and rolling stability characteristics of airplane models having a triangular wing and vertical tail surfaces. (author)

  9. Vibration characteristics of Z-ring-stiffened 60 deg conical shell models of a planetary entry spacecraft

    NASA Technical Reports Server (NTRS)

    Naumann, E. C.; Mixon, J. S.

    1971-01-01

    An experimental investigation of the vibration characteristics of a 60 deg conical shell model of a planetary entry vehicle is described and the results presented. Model configurations include the shell with or without one or two Z-ring stiffeners and with or without a simulated payload. Tests were conducted with the model clamped at the small diameter and with the model suspended at the simulated payload. Additionally, calculated results obtained from application of several analytical procedures reported in the literature are presented together with comparisons between experimental and calculated frequencies and meridional mode shapes. Generally, very good frequency agreement between experimental and calculated results was obtained for all model configurations. For small values of circumferential mode number, however, the frequency agreement decreased as the number of ring stiffeners increased. Overall agreement between experimental and calculated mode shapes was generally good. The calculated modes usually showed much larger curvatures in the vicinity of the rings than were observed in the experimentally measured mode shapes. Dual resonances associated with modal preference were noted for the shell without Z-ring stiffeners, whereas the addition of stiffeners produced resonances for which the model responded in two or more modes over different sections of the shell length.

  10. Effect of configuration modifications on the low-subsonic aerodynamic characteristics of a space shuttle orbiter concept with a blended delta wing-body

    NASA Technical Reports Server (NTRS)

    Freeman, D. C., Jr.

    1972-01-01

    An investigation of several configuration modifications to improve the subsonic stability and performance of a blended delta wing-body space shuttle-orbiter concept has been conducted in the Langley low-turbulence pressure tunnel. These modifications included variations in vertical-tail location and orientation, wing planform shape, and afterbody shape. The model was tested at a Reynolds number, based on body length, of 17 x one million, at a Mack number of 0.25, and at angles of attack from about -4 deg to 22 deg.

  11. Investigation of a delta-wing fighter model flow field at transonic speeds

    NASA Technical Reports Server (NTRS)

    Bare, E. Ann; Reubush, David E.; Haddad, Raymond; Hathaway, Ross W.; Compton, Mike

    1987-01-01

    The paper reports a flow-field investigation on a 7.52-percent scale model of an advanced fighter aircraft design conducted in the NASA-Langley 16-ft Transonic Tunnel. The effects of free-stream Mach number, angle-of-attack, angle of sideslip, and various vortex control devices on the local flow values were studied. The model was tested at Mach numbers of 0.6, 0.9, and 1.2 and the angles of sideslip of 0 and +/- 5 deg; the model angle-of-attack was varied from -4 to 30 deg. Results are presented in terms of contour plots of local total pressure recovery. The dominant influence on the over-wing flow field was found to be the wing leading-edge vortex which first appears in the survey region at an angle-of-attack of 8 deg and increases in strength and influence with increasing angle-of-attack, finally dominating the entire survey region at very high angles-of-attack.

  12. Experimental study of delta wing leading-edge devices for drag reduction at high lift

    NASA Technical Reports Server (NTRS)

    Johnson, T. D., Jr.; Rao, D. M.

    1982-01-01

    The drag reduction devices selected for evaluation were the fence, slot, pylon-type vortex generator, and sharp leading-edge extension. These devices were tested on a 60 degree flatplate delta (with blunt leading edges) in the Langley Research Center 7- by 10-foot high-speed tunnel at low speed and to angles of attack of 28 degrees. Balance and static pressure measurements were taken. The results indicate that all the devices had significant drag reduction capability and improved longitudinal stability while a slight loss of lift and increased cruise drag occurred.

  13. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 2; Small-Radius Leading Edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  14. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  15. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  16. Stability and control characteristics of a monoplane missile with large delta wings and various tail controls at Mach 1.90 to 2.86

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.

    1974-01-01

    Wind tunnel tests were conducted to determine the stability and control characteristics of a monoplane missile with large delta wings and various tail configurations. The tests were conducted at angles of attack of approximately 24 degrees and angles of sideslip of about 22 degrees. Tail control configurations were the cruciform, conventional airplane type, and an inverted airplane configuration. Control effectiveness was measured for pitch, yaw, and roll.

  17. Application of the Method of Stein and Sanders to the Calculation of Vibration Characteristics of a 45 deg Delta-Wing Specimen

    NASA Technical Reports Server (NTRS)

    Hedgepeth, John M.; Warner, Paul G., Jr.

    1959-01-01

    Generalized influence coefficients are calculated by the method of NACA TN 3640 for a large-scale, built-up, 450 delta-wing specimen. These are used together with appropriate generalized masses to obtain the natural modes and frequencies in symmetric and antisymmetric free-free vibration. The resulting frequencies are compared with those obtained experimentally and are found to be consistently high. Possible sources of the disparities are discussed.

  18. Heat transfer investigation of two Langley Research Center delta wing configurations at a Mach number of 10.5, volume 1

    NASA Technical Reports Server (NTRS)

    Eaves, R. H.; Buchanan, T. D.; Warmbrod, J. D.; Johnson, C. B.

    1972-01-01

    Heat transfer tests for two delta wing configurations were conducted in the hypervelocity wind tunnel. The 24-inch long models were tested at a Mach number of approximately 10.5 and at angles of attack of 20, 40, and 60 degrees over a length Reynolds number range from 5 million to 23 million on 4 May to 4 June 1971. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint.

  19. Aerodynamic characteristics of a 55 deg clipped-delta-wing orbiter model at Mach numbers from 1.60 to 4.63

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.; Grow, J.

    1973-01-01

    Wind tunnel tests to determine the supersonic aerodynamic characteristics of a delta wing space shuttle orbiter model were conducted. The model was tested at Mach numbers from 1.60 to 4.63, at nominal angles of attack from minus 2 degrees to plus 30 degrees, nominal sideslip angles of minus 4 degrees to plus 10 degrees, and Reynolds numbers from 1.8 to 2.5 times one million per foot.

  20. Surface-Pressure and Flow-Visualization Data at Mach Number of 1.60 for Three 65 deg Delta Wings Varying in Leading-Edge Radius and Camber

    NASA Technical Reports Server (NTRS)

    McMillin, S. Naomi; Bryd, James E.; Parmar, Devendra S.; Bezos-OConnor, Gaudy M.; Forrest, Dana K.; Bowen, Susan

    1996-01-01

    An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.

  1. Surface-Pressure and Flow-Visualization Data at Mach Number of 1.60 for Three 65 deg Delta Wings Varying in Leading-Edge Radius and Camber

    NASA Technical Reports Server (NTRS)

    McMIllin, S. Naomi; Byrd, James E.; Parmar, Devendra S.; Bezos-O'Connor, Gaudy M.; Forrest, Dana K.; Bowen, Susan

    1996-01-01

    An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data are electronically stored on the CD-ROM. The data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.

  2. 3D flow visualization and tomographic particle image velocimetry for vortex breakdown over a non-slender delta wing

    NASA Astrophysics Data System (ADS)

    Wang, ChengYue; Gao, Qi; Wei, RunJie; Li, Tian; Wang, JinJun

    2016-06-01

    Volumetric measurement for the leading-edge vortex (LEV) breakdown of a delta wing has been conducted by three-dimensional (3D) flow visualization and tomographic particle image velocimetry (TPIV). The 3D flow visualization is employed to show the vortex structures, which was recorded by four cameras with high resolution. 3D dye streaklines of the visualization are reconstructed using a similar way of particle reconstruction in TPIV. Tomographic PIV is carried out at the same time using same cameras with the dye visualization. Q criterion is employed to identify the LEV. Results of tomographic PIV agree well with the reconstructed 3D dye streaklines, which proves the validity of the measurements. The time-averaged flow field based on TPIV is shown and described by sections of velocity and streamwise vorticity. Combining the two measurement methods sheds light on the complex structures of both bubble type and spiral type of breakdown. The breakdown position is recognized by investigating both the streaklines and TPIV velocity fields. Proper orthogonal decomposition is applied to extract a pair of conjugated helical instability modes from TPIV data. Therefore, the dominant frequency of the instability modes is obtained from the corresponding POD coefficients of the modes based on wavelet transform analysis.

  3. Numerical study of a delta planform with multiple jets in ground effect

    NASA Technical Reports Server (NTRS)

    Chawla, K.; Van Dalsem, W. R.; Rao, K. V.

    1989-01-01

    The flow past a 60-deg delta wing equipped with two thrust-reverser jets near the inboard trailing edge has been analyzed by numerical solution of the 3D thin-layer Navier-Stokes equations. An implicit, partially flux-split, approximately-factored Navier-Stokes solver coupled with a multiple grid embedding scheme has been adapted to this problem. Studies of the impact of numerical parameters (e.g., grid refinement and dissipation levels), and flow-field parameters such as the height of the delta wing above the ground plane and the jet size on the solution, were performed. Results of these numerical studies indicate some challenges in the accurate resolution of complex 3D free shear layers and jets. Nevertheless, flow features such as jet deformation and ground vortex formation observed in experimental flow visualizations are captured. Further, comparisons with experimental data confirm the ability to simulate the loss of wing-borne lift, commonly referred to 'suckdown, as the delta planform flies at slow speeds in close proximity to the ground. Detailed analysis of the numerical results has also given additional insight into the structure of the ground vortex and the mechanisms of lift loss.

  4. Flow field over the wing of a delta-wing fighter model with vortex control devices at Mach 0.6 to 1.2

    NASA Technical Reports Server (NTRS)

    Bare, E. Ann; Reubush, David E.; Haddad, Raymond C.

    1992-01-01

    As part of a cooperative research program between NASA, McDonnell Douglas Corporation, and Wright Research and Development Center, a flow field investigation was conducted on a 7.52 percent scale windtunnel model of an advanced fighter aircraft design. The investigation was conducted in the Langley 16 ft Transonic Tunnel at Mach numbers of 0.6, 0.9, and 1.2. Angle of attack was varied from -4 degrees to 30 degrees and the model was tested at angles of sideslip of 0, 5, and -5 degrees. Data for the over the wing flow field were obtained at four axial survey stations by the use of six 5 hole conical probes mounted on a survey mechanism. The wing leading edge primary vortex exerted the greatest influence in terms of total pressure loss on the over the wing flow field in the area surveyed. A number of vortex control devices were also investigated. They included two different apex flaps, wing leading edge vortex flaps, and small large wing fences. The vortex flap and both apex flaps were beneficial in controlling the wing leading edge primary vortex.

  5. The effect of high curing temperature on the reaction kinetics in MK/lime and MK-blended cement matrices at 60 deg. C

    SciTech Connect

    Rojas, Moises Frias; Sanchez de Rojas, M.I

    2003-05-01

    It is well known that the pozzolanic reaction between metakaolin (MK) and calcium hydroxide produces CSH, C{sub 2}ASH{sub 8} (stratlingite), C{sub 4}AH{sub 13} and C{sub 3}ASH{sub 6} (hydrogarnet). However, the presence or absence of these hydrated phases depends on different parameters, such as curing temperature, matrix used, etc. This paper shows the results of a study in order to know the effect of high curing temperature (60 deg. C) on the kinetics of the pozzolanic reaction in different matrices. MK/lime (calcium hydroxide) and MK-blended cement matrices were studied in samples stored and cured at 60 deg. C and up to 123 days of hydration. The nature, sequence and crystallinity of the hydrated phases were analysed using differential thermal analysis (DTA) and X-ray diffraction (XRD) techniques. Results showed that the sequence and formation of the hydrated phases was different in both matrices cured at 60 deg. C. In an MK/lime matrix, C{sub 2}ASH{sub 8}, C{sub 4}AH{sub 13} and C{sub 3}ASH{sub 6} were the main hydrated phases; while in an MK-blended cement, stratlingite was the sole hydrated phase issued from pozzolanic reaction. The DTA and XRD data also reveal an important fact: there is no evidence of the presence of hydrogarnet in blended cements.

  6. Investigation of Porous Gas-Heated Leading-Edge Section for Icing Protection of a Delta Wing

    NASA Technical Reports Server (NTRS)

    Bowden, Dean T.

    1955-01-01

    A tip section of a delta wing having an NACA 0004-65 airfoil section and a 600 leading-edge sweepback was equipped with a porous leading-edge section through which hot gas was 'bled for anti-icing. Heating rates for anti-icing were determined for a wide range of icing conditions. The effects of gas flow through the porous leading-edge section on airfoil pressure distribution and drag in dry air were investigated. The drag increase caused by an ice formation on the unheated airfoil was measured for several icing conditions. Experimental porous surface- to free-stream convective heat-transfer coefficients were obtained in dry air and compared with theory. Adequate icing protection was obtained at all icing conditions investigated. Savings in total gas-flow rate up to 42 percent may be obtained with no loss in anti-icing effectiveness by sealing half the upper-surface porous area. Gas flow through the leading-edge section had no appreciable effect on airfoil pressure distribution. The airfoil section drag increased slightly (5-percent average) with gas flow through the porous surface. A heavy glaze-ice formation produced after 10 minutes of icing caused an increase in section drag coefficient of 240 percent. Experimental convective heat-transfer coefficients obtained with hot-gas flow through the porous area in dry air and turbulent flow were 20 to 30 percent lower than the theoretical values for a solid surface under similar conditions. The transition region from laminar to turbulent flow moved forward as the ratio of gas velocity through the porous surface to air-stream velocity was increased.

  7. Supersonic aerodynamic characteristics of canard, tailless, and aft-tail configurations for 2 wing planforms

    NASA Technical Reports Server (NTRS)

    Covell, P. F.

    1985-01-01

    Aerodynamic characteristics of canard, tailless, and aft tail configurations were compared in tests on a general research model (generic fuselage without canopy, inlets, or vertical tails) at Mach 1.60 and 2.00 in the Langley Unitary Plan Wind Tunnel. Two uncambered wing planforms (trapezoidal with 44 deg leading edge sweep and delta with 60 deg leading edge sweep) were tested for each configuration. The relative merits of the configurations were also determined theoretically, to evaluate the capabilities of a linear theory code for such analyses. The canard and aft tail configurations have similar measured values for lift curve slope, maximum lift drag ratio, and zero lift drag. The stability decrease as Mach number increases is greatest for the tailless configuration and least for the canard configuration. Because of very limited accuracy in predicting the aerodynamic parameter increments between configurations, the linear theory code is not adequate for determining the relative merits of canard, tailless, and aft tail configurations.

  8. Comparisons of two-dimensional shock-expansion theory with experimental aerodynamic data for delta-planform wings at high supersonic speeds

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.

    1974-01-01

    An investigation has been conducted to explore the potential for optimizing airfoil shape at high supersonic speeds by utilizing the two-dimensional shock-expansion method. Theoretical and experimental force and moment coefficients are compared for four delta-planform semispan wings having a leading-edge sweep angle of 65 deg and incorporating modified diamond airfoils with a thickness-chord ratio of 0.06. The wings differ only in airfoil maximum-thickness position and camber. The experimental data are obtained at Mach numbers of 3.95 and 4.63 and at a Reynolds number of 9.84 million per meter. A relatively simple method is developed for predicting, in terms of lift-drag ratio, the optimum modified diamond airfoil at high supersonic and hypersonic speeds.

  9. Vortex flap flow reattachment line and subsonic longitudinal aerodynamic data on 50 deg to 74 deg Delta wings on common fuselage

    NASA Technical Reports Server (NTRS)

    Frink, N. T.; Huffman, J. K.; Johnson, T. D., Jr.

    1983-01-01

    Positions of the primary vortex flow reattachment line and longitudinal aerodynamic data were obtained at Mach number 0.3 for a systematic series of vortex flaps on delta wing body configurations with leading edge sweeps of 50, 58, 66, and 74 deg. The investigation was performed to study the parametric effects of wing sweep, vortex flap geometry and deflection, canards, and trailing edge flaps on the location of the primary vortex reattachment line relative to the flap hinge line. The vortex reattachment line was located via surface oil flow photographs taken at selected angles of attack. Force and moment measurements were taken over an angle of attack range of -1 deg to 22 deg at zero sideslip angle for many configurations to further establish the data base and to assess the aforementioned parametric effects on longitudinal aerodynamics. Both the flow reattachment and aerodynamic data are presented.

  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. Effects of Various Fillet Shapes on a 76/40 Double Delta Wing from Mach 0.18 to 0.7

    NASA Technical Reports Server (NTRS)

    Gonzalez, Hugo A.; McLachlan, Blair G.; Erickson, Gary E.; Bell, James H.

    2003-01-01

    The effects of linear, diamond, and parabolic fillets on a double delta wing were investigated in the NASA Langley 7 x 10 ft High Speed Tunnel from Mach 0.18 to 0.7 and angles of attack from 4 deg. to 42 deg. Force and moment, pneumatic pressures, pressure sensitive paint, and vapor screen flow visualization measurements were used to characterize the flow field and to determine longitudinal forces and moments. The fillets increased lift coefficient and reduced induced drag without significantly affecting pitching moment. Pressure sensitive paint showed the increase in lift is caused by an increase in suction and broadening of the vortex suction footprint. Vapor screen results showed the mixing and coalescing of the strake fillet and wing vortices causes the footprint to broaden.

  12. DELTAE

    SciTech Connect

    Ward, W.C.; Swift, G.W. )

    1993-11-01

    In thermoacoustic engines and refrigerators, and in many simple acoustic systems, a one dimensional wave equation determines the spatial dependence of the acoustic pressure and velocity. DELTAE numerically integrates such wave equations in the acoustic approximation, in gases or liquids, in user-defined geometries. Boundary conditions can include conventional acoustic boundary conditions of geometry and impedance, as well as temperature and thermal power in thermoacoustic systems. DELTAE can be used easily for apparatus ranging from simple duct networks and resonators to thermoacoustic engines refrigerators and combinations thereof. It can predict how a given apparatus will perform, or can allow the user to design an apparatus to achieve desired performance. DELTAE views systems as a series of segments; twenty segment types are supported. The purely acoustic segments include ducts and cones, and lumped impedances including compliances, series impedances, and endcaps. Electroacoustics tranducer segments can be defined using either frequency-independent coefficients or the conventional parameters of loudspeaker-style drivers: mass, spring constant, magnetic field strength, etc. Tranducers can be current driven, voltage driven, or connected to an electrical load impedance. Thermoacoustic segment geometries include parallel plates, circular and rectangular pores, and pin arrays. Side branches can be defined with fixed impedances, frequency-dependent radiation impedances, or as an auxiliary series of segments of any types. The user can select working fluids from among air, helium, neon, argon, hydrogen, deuterium, carbon dioxide, nitrogen, helium-argon mixtures, helium-xenon mixtures, liquid sodium, and eutectic sodium-potassium. Additional fluids and solids can be defined by the user.

  13. Space shuttle: Heat transfer investigation of the McDonnell-Douglas delta wing orbiter at a nominal Mach number of 10.5

    NASA Technical Reports Server (NTRS)

    Eaves, R. H.; Buchanan, T. D.

    1972-01-01

    Heat transfer tests for the delta wing orbiter were conducted in a hypervelocity wind tunnel. A 1.1 percent scale model was tested at a Mach number of approximately 10.5 over an angle of attack range from 10 to 60 degrees over a length Reynolds number range from 5 times 10 to the 6th power to 24 times 10 to the 6th power. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint. Limited pressure measurements were obtained.

  14. Effect of wing flexibility on the experimental aerodynamic characteristics of an oblique wing

    NASA Technical Reports Server (NTRS)

    Hopkins, E. J.; Yee, S. C.

    1977-01-01

    A solid-aluminum oblique wing was designed to deflect considerably under load so as to relieve the asymmetric spanwise stalling that is characteristic of this type of wing by creating washout on the trailing wing panel and washin on the leading wing panel. Experimental forces, and pitching, rolling and yawing moments were measured with the wing mounted on a body of revolution. In order to vary the dynamic pressure, measurements were made at several unit Reynolds numbers, and at Mach numbers. The wing was investigated when unswept (at subsonic Mach numbers only) and when swept 45 deg, 50 deg, and 60 deg. The wing was straight tapered in planform, had an aspect ratio of 7.9 (based on the unswept span), and a profile with a maximum thickness of 4 percent chord. The results substantiate the concept that an oblique wing designed with the proper amount of flexibility self relieves itself of asymmetric spanwise stalling and the associated nonlinear moment curves.

  15. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Double Delta Wing Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Gonzalez, Hugo A.

    2006-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

  16. Design and performance of a fixed, nonaccelerating, guide vane cascade that operates over an inlet flow angle range of 60 deg

    NASA Technical Reports Server (NTRS)

    Sanz, J. M.; Mcfarland, E. R.; Sanger, N. L.; Gelder, T. F.; Cavicchi, R. H.

    1984-01-01

    A unique set of wind tunnel guide vanes are designed with an inverse design code and analyzed with a panel method and an integral boundary layer code developed at the NASA Lewis Research Center. The fixed guide vanes, 80 feet long with 6-foot chord length, were designed for the NASA Ames 40 x 80/80 x 120 ft Wind Tunnel. Low subsonic flow is accepted over a 60 deg range of inlet angle from either the 40 x 80 leg or the 80 x 120 leg of the wind tunnel, and directed axially into the main leg of the tunnel where drive fans are located. Experimental tests of 1/10-scale models were conducted to verify design calculations.

  17. Piloted-simulation study of effects of vortex flaps on low-speed handling qualities of a Delta-wing airplane

    NASA Technical Reports Server (NTRS)

    Brandon, Jay M.; Brown, Philip W.; Wunschel, Alfred J.

    1987-01-01

    A piloted-simulation study was conducted to investigate the effects of vortex flaps on low-speed handling qualities of a delta-wing airplane. The simulation math model was developed from wind tunnel tests of a 0.15 scale model of the F-106B airplane. Pilot evaluations were conducted using a six-degree-of-freedom motion base simulator. The results of the investigation showed that the reduced static longitudinal stability caused by the vortex flaps significantly degraded handling qualities in the approach-to-landing task. Acceptable handling qualities could be achieved by limiting the aft center-of-gravity location, consequently reducing the operational envelope of the airplane. Further improvement were possible by modifying the flight control force-feel system to reduce pitch-control sensitivity.

  18. Effects of roughness size on the position of boundary-layer transition and on the aerodynamic characteristics of a 55 deg. swept delta wing at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Stallings, R. L., Jr.; Lamb, M.

    1977-01-01

    An experimental investigation was conducted to determine the effects of roughness size on the position of boundary layer transition and on the aerodynamic characteristics of a 55 deg swept delta wing model. Results are presented and discussed for wind tunnel tests conducted at free stream Mach numbers from 1.50 to 4.63, Reynolds numbers per meter from 3,300,000 to 1.6 x 10 to the 7th power, angles of attack from -8 to 16 deg, and roughness sizes ranging from 0.027 cm sand grit to 0.127 cm high cylinders. Comparisons were made with existing flat plate data. An approximate method was derived for predicting the drag of roughness elements used in boundary layer trips.

  19. Motion simulator study of longitudinal stability requirements for large delta wing transport airplanes during approach and landing with stability augmentation systems failed

    NASA Technical Reports Server (NTRS)

    Snyder, C. T.; Fry, E. B.; Drinkwater, F. J., III; Forrest, R. D.; Scott, B. C.; Benefield, T. D.

    1972-01-01

    A ground-based simulator investigation was conducted in preparation for and correlation with an-flight simulator program. The objective of these studies was to define minimum acceptable levels of static longitudinal stability for landing approach following stability augmentation systems failures. The airworthiness authorities are presently attempting to establish the requirements for civil transports with only the backup flight control system operating. Using a baseline configuration representative of a large delta wing transport, 20 different configurations, many representing negative static margins, were assessed by three research test pilots in 33 hours of piloted operation. Verification of the baseline model to be used in the TIFS experiment was provided by computed and piloted comparisons with a well-validated reference airplane simulation. Pilot comments and ratings are included, as well as preliminary tracking performance and workload data.

  20. Exploratory Investigation of Forebody Strakes for Yaw Control of a Generic Fighter with a Symmetric 60 deg Half-Angle Chine Forebody

    NASA Technical Reports Server (NTRS)

    Ross, Holly M.; ORourke, Matthew J.

    1997-01-01

    Forebody strakes were tested in a low-speed wind tunnel to determine their effectiveness producing yaw control on a generic fighter model with a symmetric 60 deg half-angle chine forebody. Previous studies conducted using smooth, conventionally shaped forebodies show that forebody strakes provide increased levels of yaw control at angles of attack where conventional rudders are ineffective. The chine forebody shape was chosen for this study because chine forebodies can be designed with lower radar cross section (RCS) values than smooth forebody shapes. Because the chine edges of the forebody would fix the point of flow separation, it was unknown if any effectiveness achieved could be modulated as was successfully done on the smooth forebody shapes. The results show that use of forebody strakes on a chine forebody produce high levels of yaw control, and when combined with the rudder effectiveness, significant yaw control is available for a large range of angles of attack. The strake effectiveness was very dependent on radial location. Very small strakes placed at the tip of the forebody were nearly as effective as very long strakes. An axial translation scheme provided almost linear increments of control effectiveness.

  1. 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.

  2. Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 1; Sharp Leading Edge; [conducted in the Langley National Transonic Facility (NTF)

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Luckring, James M.

    1996-01-01

    An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

  3. Wind-tunnel force and flow visualization data at Mach numbers from 1.6 to 4.63 for a series of bodies of revolution at angles of attack from minus 4 deg to 60 deg

    NASA Technical Reports Server (NTRS)

    Landrum, E. J.; Babb, C. D.

    1979-01-01

    Flow visualization and force data for a series of six bodies of revolution are presented without analysis. The data were obtained in the Langley Unitary Plan wind tunnel for angles of attack from -4 deg to 60 deg. The Reynolds number used for these tests was 6,600,000 per meter.

  4. The natural flow wing-design concept

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.

    1992-01-01

    A wing-design study was conducted on a 65 degree swept leading-edge delta wing in which the wing geometry was modified to take advantage of the naturally occurring flow that forms over a slender wing in a supersonic flow field. Three-dimensional nonlinear analysis methods were used in the study which was divided into three parts: preliminary design, initial design, and final design. In the preliminary design, the wing planform, the design conditions, and the near-conical wing-design concept were derived, and a baseline standard wing (conventional airfoil distribution) and a baseline near-conical wing were chosen. During the initial analysis, a full-potential flow solver was employed to determine the aerodynamic characteristics of the baseline standard delta wing and to investigate modifications to the airfoil thickness, leading-edge radius, airfoil maximum-thickness position, and wing upper to lower surface asymmetry on the baseline near-conical wing. The final design employed an Euler solver to analyze the best wing configurations found in the initial design and to extend the study of wing asymmetry to develop a more refined wing. Benefits resulting from each modification are discussed, and a final 'natural flow' wing geometry was designed that provides an improvement in aerodynamic performance compared with that of a baseline conventional uncambered wing, linear-theory cambered wing, and near-conical wing.

  5. Wind-tunnel pressure data at Mach numbers from 1.6 to 4.63 for a series of bodies of revolution at angles of attack from -4 deg to 60 deg

    NASA Technical Reports Server (NTRS)

    Landrum, E. J.

    1977-01-01

    The tabulated results of wind tunnel pressure tests are presented without analysis. The data were obtained for a series of six bodies of revolution at Mach numbers of 1.6, 2.3, 2.96, and 4.63 for angles of attack from -4 deg. to 60 deg. The Reynolds number used for these tests was 6.6 x 6/million per meter.

  6. Space shuttle: Pressure investigation of a space shuttle launch configuration consisting of a delta-wing orbiter and a swept-wing booster with canard and tip fans (M equals 0.6 to 1.3). Volume 1, part A: Booster data

    NASA Technical Reports Server (NTRS)

    Rampy, J. M.; Blackwell, K. L.; Gomillion, G. R.

    1973-01-01

    Wind tunnel tests to determine the pressure distribution on a space shuttle launch configuration consisting of a delta wing orbiter and a swept wing booster with canard and tip fins were conducted. Pressure data were obtained for the combined orbiter and booster and for the booster alone at Mach numbers from 0.6 to 1.3, angles of attack from minus 8 degrees to plus 10 degrees, and sideslip angles from minus 6 degrees to plus 6 degrees. Pressure data were also obtained for the booster alone without canard at Mach numbers of 0.9 and 1.1. The pressure taps were distributed primarily over the booster upper surface and the orbiter lower surface.

  7. Wake Measurement Downstream of a Hybrid Wing Body Model with Blown Flaps

    NASA Technical Reports Server (NTRS)

    Lin, John C.; Jones, Gregory S.; Allan, Brian G.; Westra, Bryan W.; Collins, Scott W.; Zeune, Cale H.

    2010-01-01

    Flow-field measurements were obtained in the wake of a full-span Hybrid Wing Body model with internally blown flaps. The test was performed at the NASA Langley 14 x 22 Foot Subsonic Tunnel at low speeds. Off-body measurements were obtained with a 7-hole probe rake survey system. Three model configurations were investigated. At 0deg angle of attack the surveys were completed with 0deg and 60deg flap deflections. At 10deg angle of attack the wake surveys were completed with a slat and a 60deg flap deflection. The 7-hole probe results further quantified two known swirling regions (downstream of the outboard flap edge and the inboard/outboard flap juncture) for the 60deg flap cases with blowing. Flowfield results and the general trends are very similar for the two blowing cases at nozzle pressure ratios of 1.37 and 1.56. High downwash velocities correlated with the enhanced lift for the 60deg flap cases with blowing. Jet-induced effects are the largest at the most inboard station for all (three) velocity components due in part to the larger inboard slot height. The experimental data are being used to improve computational tools for high-lift wings with integrated powered-lift technologies.

  8. Flow-Field Measurement of a Hybrid Wing Body Model with Blown Flaps

    NASA Technical Reports Server (NTRS)

    Lin, John C.; Jones, Gregory S.; Allan, Brian G.; Westra, Bryan W.; Collins, Scott W.; Zeune, Cal H.

    2008-01-01

    In this paper we describe flow-field measurements obtained in the wake of a full-span Hybrid Wing Body model with internally blown flaps. The test was performed at the NASA Langley 14 x 22 Foot Subsonic Tunnel at low speeds. Off-body measurements were obtained with a 7-hole probe rake survey system. Three model configurations were investigated. At 0deg angle of attack the surveys were completed with 0deg and 60deg flap deflections. At 10deg angle of attack the wake surveys were completed with a slat and a 60deg flap deflection. The 7-hole probe results further quantified two known swirling regions (downstream of the outboard flap edge and the inboard/outboard flap juncture) for the 60deg flap cases with blowing. Flow-field results and the general trends are very similar for the two blowing cases at nozzle pressure ratios of 1.37 and 1.56. High downwash velocities correlated with the enhanced lift for the 60deg flap cases with blowing. Jet-induced effects are the largest at the most inboard station for all (three) velocity components due in part to the larger inboard slot height. The experimental data are being used to improve computational tools for high-lift wings with integrated powered-lift technologies.

  9. A study of canard-wing interference using experimental pressure data at transonic speeds

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.; Washburn, K. E.

    1979-01-01

    The canard had an exposed area of 28.0 percent of the wing reference area and was located in the chord plane of the wing or in a position 18.5 percent of the wing mean geometric chord above or below the wing chord plane. The canard leading edge sweep was 51.7 deg and the wing leading-edge sweep was 60 deg. The results indicated that the direct canard downwash effects on the wing loading are limited to the forward half of the wing directly behind the canard. The wing leading-edge vortex is located farther forward for the wing in the presence of the canard than for the wing-alone configuration. The wake, from the canard located below the wing chord plane, physically interacts with the wing inboard surface and produces a substantial loss of wing lift. For the Mach number 0.70 case, the presence of the wing increased the loading on the canard for the higher angles of attack. However, at Mach numbers of 0.95 and 1.20, the presence of the wing had the unexpected result of unloading the canard.

  10. Aileron and Krueger nose flap effectiveness measured on an oblique wing

    NASA Technical Reports Server (NTRS)

    Hopkins, E. J.; Lovette, G. H.

    1977-01-01

    The effects that deflection of conventional ailerons have on the lateral/directional characteristics of an oblique wing mounted on top of a body of revolution were measured. The wing had an aspect ratio of 6 (based on the unswept wing span) and was tested at various sweep angles ranging from 0 deg to 60 deg. Krueger nose flaps mounted on drooped-nose flaps were investigated on the downstream wing panel (for sweep angles of 45 deg and 60 deg) in an attempt to create more symmetrical spanwise wing stalling at high lift coefficients. The tests were conducted over a Mach number range of 0.25 to 1.4 and an angle-of-attack range from -3 deg to 28 deg. With the oblique wing in the swept position, considerable difference in the aileron effectiveness was measured for positive or negative differential aileron deflections. With the wing swept 45 deg at a Mach number of 0.8 and above, an extension of the span of the Krueger nose flaps on the downstream wing panel did not improve the linearity of the pitching-, rolling-, and yawing-moment curves.

  11. Delta-9 desaturase from sharpshooters

    Technology Transfer Automated Retrieval System (TEKTRAN)

    Genomic analyses of several leafhoppers identified the first leafhopper delta-9 desaturase. Identification of important gene transcripts within insect pests permits them to be targeted with RNA interference, RNAi, strategies. The glassy-winged sharpshooter, GWSS (Hemiptera: Cicadellidae; Homalodis...

  12. Static and unsteady pressure measurements on a 50 degree clipped delta wing at M = 0.9. [conducted in the Langley Transonic Dynamics Tunnel

    NASA Technical Reports Server (NTRS)

    Hess, R. W.; Wynne, E. C.; Cazier, F. W.

    1982-01-01

    Pressures were measured with Freon as the test medium. Data taken at M = 0.9 is presented for static and oscillatory deflections of the trailing edge control surface and for the wing in pitch. Comparisons of the static measured data are made with results computed using the Bailey-Ballhaus small disturbance code.

  13. Assessment at full scale of exhaust nozzle to wing size on STOL-OTW acoustic characteristics

    NASA Technical Reports Server (NTRS)

    Vonglahn, U.; Grosbeck, D.

    1979-01-01

    On the basis of static aero/acoustic data obtained at model scale, the effect of exhaust nozzle size on flyover noise is evaluated at full scale for different STOL-OTW nozzle configurations. Three types of nozzles are evaluated: a circular/deflector nozzle mounted above the wing; a slot/deflector nozzle mounted on the wing; and a slot nozzle mounted on the wing. The nozzle exhaust plane location, measured from the wing leading edge, was varied from 10 to 46 percent of the wing chord (flaps retracted). Flap angles of 20 deg (takeoff) and 60 deg (approach) are included in the study. Initially, perceived noise levels (PNL) are calculated as a function flyover distance at 152m altitude. From these plots, static EPNL values (defined as flyover relative noise levels), are obtained as functions of nozzle size for equal aerodynamic performance (lift and thrust). The acoustic benefits attributable to nozzle size relative to a given wing chord size are assessed.

  14. Aerodynamic Characteristics over a Mach Number Range of 1.40 to 2.78 of a Rocket-Propelled Airplane Configuration having a Low 52.50 Delta Wing and an Unswept Horizontal Tail

    NASA Technical Reports Server (NTRS)

    Kehlet, Alan B.

    1961-01-01

    A free-flight investigation of an airplane configuration having a low 52.5 deg. delta wing and an unswept horizontal tail has been conducted over a Mach number range of 1.40 to 2.78. At a fixed tail setting of -3.0 deg., the trim lift coefficient and angle of attack varied from about 0.12 to 0.04 and 3.8 deg. to 2.0 deg., respectively. The base drag was approximately 5 percent of the total drag at trim lift. Lift-curve slope, static longitudinal stability, and damping in pitch were obtained only at Mach numbers of 2.59 t o 2.74. Theoretical calculations of lift-curve slope and aerodynamic-center location were in good agreement with experimental results.

  15. Static Wind-Tunnel and Radio-Controlled Flight Test Investigation of a Remotely Piloted Vehicle Having a Delta Wing Planform

    NASA Technical Reports Server (NTRS)

    Yip, Long P.; Fratello, David J.; Robelen, David B.; Makowiec, George M.

    1990-01-01

    At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the Flight Dynamics Branch at the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted in the Langley 12 foot Low Speed Wind Tunnel to identify and improve the stability and control characteristics of the vehicle. The wind tunnel test resulted in several configuration modifications which included increased elevator size, increased vertical tail size and tail moment arm, increased rudder size and aileron size, the addition of vertical wing tip fins, and the addition of leading-edge droops on the outboard wing panel to improve stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.

  16. Summary of low-speed longitudinal aerodynamics of two powered close-coupled wing-canard fighter configurations. [conducted in Langley C/STOL tunnel

    NASA Technical Reports Server (NTRS)

    Paulson, J. W., Jr.; Thomas, J. L.

    1979-01-01

    Investigations of the low speed longitudinal characteristics of two powered close coupled wing-canard fighter configurations are discussed. Data obtained at angles of attack from -2 deg to 42 deg, Mach numbers from 0.12 to 0.20, nozzle and flap deflections from 0 deg to 40 deg, and thrust coefficients from 0 to 2.0, to represent both high angle of attack subsonic maneuvering characteristics and conventional takeoff and landing characteristics are examined. Data obtained with the nozzles deflected either 60 deg or 90 deg and the flaps deflected 60 deg to represent vertical or short takeoff and landing characteristics are discussed.

  17. Flight-Test Evaluation of the Longitudinal Stability and Control Characteristics of 0.5-Scale Models of the Fairchild Lark Pilotless-Aircraft Configuration: Standard Configuration with Wing Flaps Deflected 60 Degrees and Model having Tail in Line with Wings, TED No. NACA 2387

    NASA Technical Reports Server (NTRS)

    Stone, David G.

    1947-01-01

    Flight tests were conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallop Island, Va., to determine the longitudinal control and stability characteristics of 0.5-scale models of the Fairchild Lark pilotless aircraft with the tail in line with the wings a d with the horizontal wing flaps deflected 60 deg. The data were obtained by the use of a telemeter and by radar tracking.

  18. Effect of canard location and size on canard-wing interference and aerodynamic center shift related to maneuvering aircraft at transonic speeds

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.

    1974-01-01

    A generalized wind-tunnel model, typical of highly maneuverable aircraft, was tested in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.70 to 1.20 to determine the effects of canard location and size on canard-wing interference effects and aerodynamic center shift at transonic speeds. The canards had exposed areas of 16.0 and 28.0 percent of the wing reference area and were located in the chord plane of the wing or in a position 18.5 percent of the wing mean geometric chord above or below the wing chord plane. Two different wing planforms were tested, one with leading-edge sweep of 60 deg and the other 44 deg; both wings had the same reference area and span. The results indicated that the largest benefits in lift and drag were obtained with the canard above the wing chord plane for both wings tested. The low canard configuration for the 60 deg swept wing proved to be more stable and produced a more linear pitching-moment curve than the high and coplanar canard configurations for the subsonic test Mach numbers.

  19. Effect of wing sweep, angle of attack, Reynolds number, and wing root fillet on the interference heating to the wing windward surface of an entry vehicle configuration

    NASA Technical Reports Server (NTRS)

    Clark, L. E.

    1972-01-01

    The phase-change-coating technique was used to study the interference heating to the windward surface of 14 deg, 25 deg, and 50 deg swept wings of an entry vehicle configuration. One wing root of each model was faired to the fuselage with a fillet. Tests were made at Mach 8 at angles of attack of 0 deg, 20 deg, 40 deg, and 60 deg and at free-stream Reynolds numbers based on model length of 0.47 and 1.7 million. Bow shock impingement heating was found to increase in magnitude and affected area with increasing angle of attack until at a higher angle of attack it decreases; this angle of attack is lower for a 50 deg swept wing. Wing root interference heating was found to increase with angle of attack up to 40 deg and then to remain approximately constant. Consequently, wing root interference heating becomes the major type of interference heating at large angles of attack, and this occurs at a lower angle of attack for the highest sweep angle. A wing leading-edge root fillet reduces the peak in wing root interference heating near the leading edge, and increasing Reynolds number increases the level of interference heating.

  20. Avian Wings

    NASA Technical Reports Server (NTRS)

    Liu, Tianshu; Kuykendoll, K.; Rhew, R.; Jones, S.

    2004-01-01

    This paper describes the avian wing geometry (Seagull, Merganser, Teal and Owl) extracted from non-contact surface measurements using a three-dimensional laser scanner. The geometric quantities, including the camber line and thickness distribution of airfoil, wing planform, chord distribution, and twist distribution, are given in convenient analytical expressions. Thus, the avian wing surfaces can be generated and the wing kinematics can be simulated. The aerodynamic characteristics of avian airfoils in steady inviscid flows are briefly discussed. The avian wing kinematics is recovered from videos of three level-flying birds (Crane, Seagull and Goose) based on a two-jointed arm model. A flapping seagull wing in the 3D physical space is re-constructed from the extracted wing geometry and kinematics.

  1. Assessment at full scale of nozzle/wing geometry effects on OTW aeroacoustic characteristics. [Over The Wing STOL engine configurations

    NASA Technical Reports Server (NTRS)

    Groesbeck, D.; Von Glahn, U.

    1979-01-01

    The effects on acoustic characteristics of nozzle type and location on a wing for STOL engine over-the-wing configurations are assessed at full scale on the basis of model-scale data. Three types of nozzle configurations are evaluated: a circular nozzle with external deflector mounted above the wing, a slot nozzle with external deflector mounted on the wing and a slot nozzle mounted on the wing. Nozzle exhaust plane locations with respect to the wing leading edge are varied from 10 to 46 percent chord (flaps retracted) with flap angles of 20 deg (take-off attitude) and 60 deg (approach attitude). Perceived noise levels (PNL) are calculated as a function of flyover distance at 152 m altitude. From these plots, static EPNL values, defined as flyover relative noise levels, are calculated and plotted as a function of lift and thrust ratios. From such plots the acoustic benefits attributable to variations in nozzle/deflector/wing geometry at full scale are assessed for equal aerodynamic performance.

  2. Studies of IRAS sources at high galactic latitudes. I - Source counts at /b/greater than 60 deg and evidence for a north-south anisotropy of cosmological significance

    NASA Technical Reports Server (NTRS)

    Rowan-Robinson, M.; Walker, D.; Chester, T.; Soifer, T.; Fairclough, J.

    1986-01-01

    A study of the IRAS sky at b with an absolute value greater than 60 deg is conducted. Source counts at 12, 25, 60 and 100 microns are presented, and it is shown that emission from interstellar dust at 100 microns is localized to a few small areas of tathe galactic polar caps. At 12 and 25 microns, the sky is dominated by stars; at 60 and 100 microns, by galaxies. Comparison with the minisurvey source counts indicates the 12and 25-micron source denstiy is lower at the present latitude than at a latitude whereby the absolute value of b equals 10-40 deg. Due to the greatly reduced effects of emission from interstellar dust, the 100 micron survey reaches a factor 1.6 deeper in flux at the present latitude than the minisurvey. An anisotropy significant at the 4-sigma level was found between the north and south galactic polar caps at 60 and 100 microns, after exclusion of the Virgo cluster and of the few remaining areas significantly affected by interstellar-dust emission. It is suggested that this anisotropy represents a cosmologically significant anisotropy in the galaxy distribution. The scale of associated inhomogeneity is of the order of at least 100(50/H)Mpc.

  3. A Preliminary Analysis of the Flying Qualities of the Consolidated Vultee MX-813 Delta-Wing Airplane Configuration at Transonic and Low Supersonic Speeds as Determined from Flights of Rocket-Powered Models

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L.

    1949-01-01

    A preliminary analysis of the flying qualities of the Consolidated Vultee MX-813 delta-wing airplane configuration has been made based on the results obtained from the first two 1/8 scale models flown at the NACA Pilotless Aircraft Research Station, Wallop's Island, VA. The Mach number range covered in the tests was from 0.9 to 1.2. The analysis indicates adequate elevator control for trim in level flight over the speed range investigated. Through the transonic range there is a mild trim change with a slight tucking-under tendency. The elevator control effectiveness in the supersonic range is reduced to about one-half the subsonic value although sufficient control for maneuvering is available as indicated by the fact that 10 deg elevator deflection produced 5g acceleration at Mach number of 1.2 at 40,000 feet.The elevator control forces are high and indicate the power required of the boost system. The damping. of the short-period oscillation is adequate at sea-level but is reduced at 40,000 feet. The directional stability appears adequate for the speed range and angles of attack covered.

  4. Subsonic dynamic stability characteristics of two close-coupled canard-wing configurations

    NASA Technical Reports Server (NTRS)

    Boyden, R. P.

    1978-01-01

    The pitch, yaw, and roll damping, as well as the oscillatory stability in pitch and in yaw, were measured for two canard wing configurations with wing sweeps of 44 deg and 60 deg. Tests were made at free stream Mach numbers of 0.3, 0.4, and 0.7 and for angles of attack from about -4 deg to 20 deg. The effects of various components such as the canard, nose strakes, wings, vertical tail, and horizontal tail were determined. The basic canard wing, vertical tail configurations generally had positive damping in pitch, yaw, and roll. The effect of the canard was generally beneficial except for its tendency to decrease the oscillatory directional stability.

  5. Joukowski wings

    NASA Technical Reports Server (NTRS)

    Margoulis, W

    1922-01-01

    To sum up, Professor Joukowski's theory of supporting wings renders it possible to calculate the coefficient of lift in terms of the angle of attack, and Prandtl's coefficient of induced drag and the correction of the angle of attack in terms of the disposition and aspect ratio of the wings.

  6. Wing pressure distributions from subsonic tests of a high-wing transport model. [in the Langley 14- by 22-Foot Subsonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Applin, Zachary T.; Gentry, Garl L., Jr.; Takallu, M. A.

    1995-01-01

    A wind tunnel investigation was conducted on a generic, high-wing transport model in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data that document effects of various model configurations and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a part-span, double-slotted trailing-edge flap system. The trailing-edge flap was tested at four different deflection angles (20 deg, 30 deg, 40 deg, and 60 deg). Four wing configurations were tested: cruise, flaps only, Krueger flap only, and high lift (Krueger flap and flaps deployed). Tests were conducted at free-stream dynamic pressures of 20 psf to 60 psf with corresponding chord Reynolds numbers of 1.22 x 10(exp 6) to 2.11 x 10(exp 6) and Mach numbers of 0.12 to 0.20. The angles of attack presented range from 0 deg to 20 deg and were determined by wing configuration. The angle of sideslip ranged from minus 20 deg to 20 deg. In general, pressure distributions were relatively insensitive to free-stream speed with exceptions primarily at high angles of attack or high flap deflections. Increasing trailing-edge Krueger flap significantly reduced peak suction pressures and steep gradients on the wing at high angles of attack. Installation of the empennage had no effect on wing pressure distributions. Unpowered engine nacelles reduced suction pressures on the wing and the flaps.

  7. Tabulated Pressure Data for a Series of Controls on a 40 Deg Sweptback Wing at Mach Numbers of 1.61 and 2.01

    NASA Technical Reports Server (NTRS)

    Lord, D. R.

    1957-01-01

    An investigation has been made at Mach numbers of 1.61 and 2.01 and Reynolds numbers of 1.7 x l0(exp 6) and 3.6 x l0(exp 6) to determine the pressure distributions over a swept wing with a series of 14 control configurations. The wing had 40 deg of sweep of the quarter-chord line, an aspect ratio of 3.1, and a taper ratio of 0.4. Measurements were made at angles of attack from 0 deg to +/- 15 deg for control deflections from -60 deg to 60 deg. This report contains tabulated pressure data for the complete range of test conditions.

  8. Scapular Winging

    PubMed Central

    Gooding, Benjamin W. T.; Geoghegan, John M.; Wallace, W. Angus; Manning, Paul A.

    2013-01-01

    This review explores the causes of scapula winging, with overview of the relevant anatomy, proposed aetiology and treatment. Particular focus is given to lesions of the long thoracic nerve, which is reported to be the most common aetiological factor.

  9. Subsonic roll damping of a model with swept-back and swept-forward wings

    NASA Technical Reports Server (NTRS)

    Boyden, R. P.

    1978-01-01

    The aerodynamic roll damping and the yawing moment due to roll rate characteristics were investigated at subsonic speeds for a model with either sweptback or swept forward wings. The tests were made in the Langley high speed 7 by 10 foot tunnel for Mach numbers between 0.3 and 0.7. The configuration with a 60 deg sweptback wing had positive damping in roll up to the maximum test angle of attack of almost 20 deg. The 32 deg swept forward wing configuration had positive damping in roll at the lower angles of attack, but there was a decrease in damping and negative damping in roll was measured at the highest angles of attack.

  10. Subsonic longitudinal and lateral aerodynamic characteristics for a systematic series of strake-wing configurations

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    1979-01-01

    A systematic wind tunnel study was conducted in the Langley 7 by 10 foot high speed tunnel to help establish a parametric data base of the longitudinal and lateral aerodynamic characteristics for configurations incorporating strake-wing geometries indicative of current and proposed maneuvering aircraft. The configurations employed combinations of strakes with reflexed planforms having exposed spans of 10%, 20%, and 30% of the reference wing span and wings with trapezoidal planforms having leading edge sweep angles of approximately 30, 40, 44, 50, and 60 deg. Tests were conducted at Mach numbers ranging from 0.3 to 0.8 and at angles of attack from approximately -4 to 48 deg at zero sideslip.

  11. Effect of canard vertical location, size, and deflection on canard-wing interference at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.; Ray, E. J.; Washburn, K. E.

    1978-01-01

    A generalized close-coupled canard-wing configuration was tested in a high speed 7 by 10 foot tunnel at Mach numbers of 0.40, 0.70, and 0.85 over an angle-of-attack range from -4 deg to 24 deg. Studies were made to determine the effects of canard vertical location, size, and deflection and wing leading-edge sweep on the longitudinal characteristics of the basic configuration. The two wings tested had thin symmetrical circular-arc airfoil sections with characteristically sharp leading edges swept at 60 deg and 44 deg. Two balances which allow separation of the canard-forebody contribution from the total forces and moments were used in this study.

  12. Survey of wing and flap lower-surface temperatures and pressures during full-scale ground tests of an externally blown flap system

    NASA Technical Reports Server (NTRS)

    Hughes, D. L.

    1972-01-01

    Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power and engine/wing positions. The tests were made with no airflow over the wing. The leading-edge wing sweep angle was fixed at 26 deg, the angle of incidence between the engine and the wing was fixed at 3 deg, and the tests were conducted with the flap retracted, extended and deflected 35 deg, and extended and deflected 60 deg. The integrated local pressures on the undersurface of the flap produced loads approximately three times as great at the 60 deg flap position as at the 35 deg flap position. With both nozzle configurations, more than 90 percent of the integrated pressure loads were contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline. The maximum temperature recorded on the flaps was 218 C (424 F) for the conical nozzle and 180 C (356 F) for the daisy nozzle.

  13. Wing rock suppression using forebody vortex control

    NASA Technical Reports Server (NTRS)

    Ng, T. T.; Ong, L. Y.; Suarez, C. J.; Malcolm, G. N.

    1991-01-01

    Static and free-to-roll tests were conducted in a water tunnel with a configuration that consisted of a highly-slender forebody and 78-deg sweep delta wings. Flow visualization was performed and the roll angle histories were obtained. The fluid mechanisms governing the wing rock of this configuration were identified. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetries had to be induced at the same time. On the other hand, alternating pulsed blowing on the left and right sides of the forebody was demonstrated to be potentially an effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

  14. Computed Aeroelastic Motions Of Wings In Transonic Flows

    NASA Technical Reports Server (NTRS)

    Guruswamy, Guru P.; Obayashi, Shigeru

    1995-01-01

    Report describes computational simulations of aeroelastic motions of delta and swept wings in transonic flows. Study directed toward understanding aerodynamic behavior and enhancing maneuverability of fighter airplanes equipped with such wings. Also has implications for gas pumps and turbines, in which flows near tips of vanes and blades reach supersonic speeds.

  15. Aerodynamic characteristics of a distinct wing-body configuration at Mach 6: Experiment, theory, and the hypersonic isolation principle

    NASA Technical Reports Server (NTRS)

    Penland, J. A.; Pittman, J. L.

    1985-01-01

    An experimental investigation has been conducted to determine the effect of wing leading edge sweep and wing translation on the aerodynamic characteristics of a wing body configuration at a free stream Mach number of about 6 and Reynolds number (based on body length) of 17.9 x 10 to the 6th power. Seven wings with leading edge sweep angles from -20 deg to 60 deg were tested on a common body over an angle of attack range from -12 deg to 10 deg. All wings had a common span, aspect ratio, taper ratio, planform area, and thickness ratio. Wings were translated longitudinally on the body to make tests possible with the total and exposed mean aerodynamic chords located at a fixed body station. Aerodynamic forces were found to be independent of wing sweep and translation, and pitching moments were constant when the exposed wing mean aerodynamic chord was located at a fixed body station. Thus, the Hypersonic Isolation Principle was verified. Theory applied with tangent wedge pressures on the wing and tangent cone pressures on the body provided excellent predictions of aerodynamic force coefficients but poor estimates of moment coefficients.

  16. Inflatable wing

    DOEpatents

    Priddy, Tommy G.

    1988-01-01

    An inflatable wing is formed from a pair of tapered, conical inflatable tubes in bonded tangential contact with each other. The tubes are further connected together by means of top and bottom reinforcement boards having corresponding longitudinal edges lying in the same central diametral plane passing through the associated tube. The reinforcement boards are made of a stiff reinforcement material, such as Kevlar, collapsible in a direction parallel to the spanwise wing axis upon deflation of the tubes. The stiff reinforcement material cooperates with the inflated tubes to impart structural I-beam characteristics to the composite structure for transferring inflation pressure-induced tensile stress from the tubes to the reinforcement boards. A plurality of rigid hoops shaped to provide airfoil definition are spaced from each other along the spanwise axis and are connected to the top and bottom reinforcement boards. Tension lines are employed for stabilizing the hoops along the trailing and leading edges thereof.

  17. Nile Delta

    Atmospheric Science Data Center

    2013-04-15

    article title:  The Nile River Delta     View Larger Image ... of eastern Africa. At the apex of the fertile Nile River Delta is the Egyptian capital city of Cairo. To the west are the Great Pyramids ...

  18. Experimental and numerical analysis of the wing rock characteristics of a 'wing-body-tail' configuration

    NASA Technical Reports Server (NTRS)

    Suarez, Carlos J.; Smith, Brooke C.; Malcolm, Gerald N.

    1993-01-01

    Free-to-roll wind tunnel tests were conducted and a computer simulation exercise was performed in an effort to investigate in detail the mechanism of wing rock on a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. In the wind tunnel test, the roll angle and wing surface pressures were measured during the wing rock motion. A limit cycle oscillation was observed for angles of attack between 22 deg and 30 deg. In general, the wind tunnel test confirmed that the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. The variation of roll acceleration (determined from the second derivative of the roll angle time history) with roll angle clearly showed the energy balance necessary to sustain the limit cycle oscillation. Pressure measurements on the wing revealed the hysteresis of the wing rock process. First, second and nth order models for the aerodynamic damping were developed and examined with a one degree of freedom computer simulation. Very good agreement with the observed behavior from the wind tunnel was obtained.

  19. Theoretical study of aerodynamic characteristics of wings having vortex flow

    NASA Technical Reports Server (NTRS)

    Reddy, C. S.

    1979-01-01

    The aerodynamic characteristics of slender wings having separation induced vortex flows are investigated by employing three different computer codes--free vortex sheet, quasi vortex lattice, and suction analogy methods. Their capabilities and limitations are examined, and modifications are discussed. Flat wings of different configurations: arrow, delta, and diamond shapes, as well as cambered delta wings, are studied. The effect of notch ratio on the load distributions and the longitudinal characteristics of a family of arrow and diamond wings is explored. The sectional lift coefficients and the accumulated span loadings are determined for an arrow wing and are seen to be unusual in comparison with the attached flow results. The theoretically predicted results are compared with the existing experimental values.

  20. Unsteady flow computation of oscillating flexible wings

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Chuang, H. Andrew; Salman, Ahmed A.

    1990-01-01

    The problem of unsteady flow around flexible wings is solved using the unsteady, compressible, thin-layer Navier-Stokes equations in conjunction with the unsteady, linearized, Navier-displacement equations. Starting with the initial shape of the wing, the Navier-Stokes equations are solved on an initial structured grid to obtain the steady-flow solution which is used for the initial conditions. The forced deformation motion of the wing boundaries is then applied, and the problem is solved accurately in time. During the time-accurate stepping, the Navier-displacement equations are used to solve for the grid deformation and sequently, the Navier-Stokes equations are used to solve for the flowfield. Two applications are presented; the first is for a pulsating oscillation of a bending-mode airfoil in transonic flow, and the second is for a bending-mode oscillation of a sharp-edged delta wing in supersonic flow.

  1. Supersonic aerodynamic characteristics of a lifting-body orbiter model with a blunted delta planform at Mach 2.30 to 4.60

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1972-01-01

    An investigation has been made in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a lifting-body orbiter model with a blunted delta planform. The model was tested at Mach numbers from 2.30 to 4.60, at nominal angles of attack from -4 deg to 60 deg and angles of sideslip from -4 deg to 10 deg, and at a Reynolds number of 2.5 million per foot.

  2. Aerodynamic and flowfield hysteresis of slender wing aircraft undergoing large-amplitude motions

    NASA Technical Reports Server (NTRS)

    Nelson, Robert C.; Arena, Andrew S., Jr.; Thompson, Scott A.

    1991-01-01

    The implication of maneuvers through large angles of incidence is discussed by examining the unsteady aerodynamic loads, surface pressures, vortical position, and breakdown on slender, flat plate delta wings. Two examples of large amplitude unsteady motions are presented. First, the unsteady characteristics of a 70 degree swept delta wing undergoing pitch oscillation from 0 to 60 degrees is examined. Data is presented that shows the relationship between vortex breakdown and the overshoot and undershoot of the aerodynamic loads and surface pressure distribution. The second example examines the leading edge vortical flow over an 80 degree swept wing undergoing a limit cycle roll oscillation commonly called wing rock.

  3. Flutter analysis of low aspect ratio wings

    NASA Technical Reports Server (NTRS)

    Parnell, L. A.

    1986-01-01

    Several very low aspect ratio flat plate wing configurations are analyzed for their aerodynamic instability (flutter) characteristics. All of the wings investigated are delta planforms with clipped tips, made of aluminum alloy plate and cantilevered from the supporting vehicle body. Results of both subsonic and supersonic NASTRAN aeroelastic analyses as well as those from another version of the program implementing the supersonic linearized aerodynamic theory are presented. Results are selectively compared with the experimental data; however, supersonic predictions of the Mach Box method in NASTRAN are found to be erratic and erroneous, requiring the use of a separate program.

  4. Experimental Aerodynamic Characteristics of an Oblique Wing for the F-8 OWRA

    NASA Technical Reports Server (NTRS)

    Kennelly, Robert A., Jr.; Carmichael, Ralph L.; Smith, Stephen C.; Strong, James M.; Kroo, Ilan M.

    1999-01-01

    An experimental investigation was conducted during June-July 1987 in the NASA Ames 11-Foot Transonic Wind Tunnel to study the aerodynamic performance and stability and control characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. This effort was part of the Oblique Wing Research Aircraft (OWRA) program performed in conjunction with Rockwell International. The Ames-designed, aspect ratio 10.47, tapered wing used specially designed supercritical airfoils with 0.14 thickness/chord ratio at the root and 0.12 at the 85% span location. The wing was tested at two different mounting heights above the fuselage. Performance and longitudinal stability data were obtained at sweep angles of 0deg, 30deg, 45deg, 60deg, and 65deg at Mach numbers ranging from 0.30 to 1.40. Reynolds number varied from 3.1 x 10(exp 6)to 5.2 x 10(exp 6), based on the reference chord length. Angle of attack was varied from -5deg to 18deg. The performance of this wing is compared with that of another oblique wing, designed by Rockwell International, which was tested as part of the same development program. Lateral-directional stability data were obtained for a limited combination of sweep angles and Mach numbers. Sideslip angle was varied from -5deg to +5deg. Landing flap performance was studied, as were the effects of cruise flap deflections to achieve roll trim and tailor wing camber for various flight conditions. Roll-control authority of the flaps and ailerons was measured. A novel, deflected wing tip was evaluated for roll-control authority at high sweep angles.

  5. View east, showing Northwest Wing (Wing 5) and rear elevations ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    View east, showing Northwest Wing (Wing 5) and rear elevations of facade and tis flaking wings (Wings 1 and 2) - Hospital for Sick Children, 1731 Bunker Hill Road, Northeast, Washington, District of Columbia, DC

  6. View east, showing Northwest Wing (Wing 5), west wall of ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    View east, showing Northwest Wing (Wing 5), west wall of the North Wing (Wing 2) and rear elevations of the facade and its flanking wings (Wings 1 and 2) - Hospital for Sick Children, 1731 Bunker Hill Road, Northeast, Washington, District of Columbia, DC

  7. Assessment at full scale of exhaust nozzle-to-wing size on STOL-OTW acoustic characteristics

    NASA Technical Reports Server (NTRS)

    Von Glahn, U.; Groesbeck, D.

    1979-01-01

    On the basis of static zero/acoustic data obtained at model scale, the effect of exhaust nozzle size on flyover noise is evaluated at full scale for different STOL-OTW nozzle configurations. Three types of nozzles are evaluated: a circular/deflector nozzle mounted above the wing, a slot/deflector nozzle mounted on the wing, and a slot nozzle mounted on the wing. The nozzle exhaust plane location, measured from the wing leading edge was varied from 10 to 46 percent of the wing chord (flaps retracted). Flap angles of 20 deg (takeoff) and 60 deg (approach) are included in the study. Initially, perceived noise levels (PNL) are calculated as a function of flyover distance at 152 m altitude. From these plots static EPNL values, defined as flyover relative noise levels, then are obtained as functions of nozzle size for equal aerodynamic performance (lift and thrust). On the basis of these calculations, the acoustic benefits attributable to nozzle size relative to a given wing chord size are assessed.

  8. Theory of wing rock

    NASA Technical Reports Server (NTRS)

    Hsu, C.-H.; Lan, C. E.

    1985-01-01

    Wing rock is one type of lateral-directional instabilities at high angles of attack. To predict wing rock characteristics and to design airplanes to avoid wing rock, parameters affecting wing rock characteristics must be known. A new nonlinear aerodynamic model is developed to investigate the main aerodynamic nonlinearities causing wing rock. In the present theory, the Beecham-Titchener asymptotic method is used to derive expressions for the limit-cycle amplitude and frequency of wing rock from nonlinear flight dynamics equations. The resulting expressions are capable of explaining the existence of wing rock for all types of aircraft. Wing rock is developed by negative or weakly positive roll damping, and sustained by nonlinear aerodynamic roll damping. Good agreement between theoretical and experimental results is obtained.

  9. Unsteady transonic flow analysis for low aspect ratio, pointed wings.

    NASA Technical Reports Server (NTRS)

    Kimble, K. R.; Ruo, S. Y.; Wu, J. M.; Liu, D. Y.

    1973-01-01

    Oswatitsch and Keune's parabolic method for steady transonic flow is applied and extended to thin slender wings oscillating in the sonic flow field. The parabolic constant for the wing was determined from the equivalent body of revolution. Laplace transform methods were used to derive the asymptotic equations for pressure coefficient, and the Adams-Sears iterative procedure was employed to solve the equations. A computer program was developed to find the pressure distributions, generalized force coefficients, and stability derivatives for delta, convex, and concave wing planforms.

  10. Slender wing theory including regions of embedded total pressure loss

    NASA Technical Reports Server (NTRS)

    Mccune, James E.; Tavares, T. Sean; Lee, Norman K. W.; Weissbein, David

    1988-01-01

    An aerodynamic theory of the flow about slender delta wings is described. The theory includes a treatment of the self-consistent development of the vortex wake patterns above the wing necessary to maintain smooth flow at the wing edges. The paper focuses especially on the formation within the wake of vortex 'cores' as embedded regions of total pressure loss, fed and maintained by umbilical vortex sheets emanating from the wing edges. Criteria are developed for determining the growing size and location of these cores, as well as the distribution and strength of the vorticity within them. In this paper, however, the possibility of vortex breakup is omitted. The aerodynamic consequences of the presence and evolution of the cores and the associated wake structure are illustrated and discussed. It is noted that wake history effects can have substantial influence on the distribution of normal force on the wing as well as on its magnitude.

  11. The aerodynamic design of the oblique flying wing supersonic transport

    NASA Technical Reports Server (NTRS)

    Vandervelden, Alexander J. M.; Kroo, Ilan

    1990-01-01

    The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

  12. Effects of Canard on the Flowfield over a Wing

    NASA Astrophysics Data System (ADS)

    Nayebzadeh, Arash

    2015-11-01

    Surface and flowfield pressure measurements have been done over delta wing/canard configuration in a variety of canard vertical and horizontal locations and angles of attack. The experimental model consisted of wing, canard and a body to accommodate pressure tubing and canard rotation mechanism. All the tests have been performed at subsonic velocities and the effect of canard were analyzed through comparison between surface and flowfield pressure distributions. It was found that vortex flow pattern over the wing is dominated mainly by canard vertical position and in some cases, by merging of canard and wing vortices. In addition, the pressure loss induced by canard vortex on the wing surface moves the wing vortex toward the leading edge. In the mid canard configuration, canard and wing vortices merge at x/c greater than 0.5 and as a result of this phenomenon, abrupt pressure loss induces more stable vortex flow over the wing. It is also shown that canard plays a vital role in vortex break down over the wing.

  13. Subsonic longitudinal and lateral-directional static aerodynamic characteristics for a close-coupled wing-canard model in both swept back and swept forward configurations

    NASA Technical Reports Server (NTRS)

    Huffman, J. K.; Fox, C. H., Jr.

    1978-01-01

    A general research fighter model was tested in the Langley 7 by 10-foot high speed tunnel at a Mach number of 0.3. The close-coupled wing-canard combination was tested with both lifting surfaces in a 60 deg swept back configuration and in a 32 deg swept forward configuration. The angle-of-attack range was from approximately -4 deg to 48 deg at sideslip angles of zero deg, -5 deg. The data is presented without analysis in order to expedite publication.

  14. Impingement of Droplets in 60 Deg Elbows with Potential Flow

    NASA Technical Reports Server (NTRS)

    Hacker, Paul T.; Saper, Paul G.; Kadow, Charles F.

    1956-01-01

    Trajectories were determined for water droplets or other aerosol particles in air flowing through 600 elbows especially designed for two-dimensional potential motion. The elbows were established by selecting as walls of each elbow two streamlines of a flow field produced by a complex potential function that establishes a two-dimensional flow around. a 600 bend. An unlimited number of elbows with slightly different shapes can be established by selecting different pairs of streamlines as walls. Some of these have a pocket on the outside wall. The elbows produced by the complex potential function are suitable for use in aircraft air-inlet ducts and have the following characteristics: (1) The resultant velocity at any point inside the elbow is always greater than zero but never exceeds the velocity at the entrance. (2) The air flow field at the entrance and exit is almost uniform and rectilinear. (3) The elbows are symmetrical with respect to the bisector of the angle of bend. These elbows should have lower pressure losses than bends of constant cross-sectional area. The droplet impingement data derived from the trajectories are presented along with equations so that collection efficiency, area, rate, and distribution of droplet impingement can be determined for any elbow defined by any pair of streamlines within a portion of the flow field established by the complex potential function. Coordinates for some typical streamlines of the flow field and velocity components for several points along these streamlines are presented in tabular form. A comparison of the 600 elbow with previous calculations for a comparable 90 elbow indicated that the impingement characteristics of the two elbows were very similar.

  15. Geometry effects on STOL engine-over-the-wing acoustics with 5.1 slot nozzles

    NASA Technical Reports Server (NTRS)

    Vonglahn, U.; Groesbeck, D.

    1975-01-01

    The correspondence of far field acoustic trends with changes in the characteristics of the flow field at the wing trailing edge caused by alterations in the nozzle-wing geometry were determined for several STOL-OTW configurations. Nozzle roof angles of 10 to 40 deg were tested with and without cutback of the nozzle sidewalls. Three wing chord sizes were used: baseline (33 cm with flaps retracted), 2/3-baseline, and 3/2-baseline. Flap deflection angles of 20 and 60 deg were used. The nozzle locations were at 21 and 46-percent of chord. With increasing wing size the jet noise shielding benefits increased. With increasing nozzle roof angle, the jet velocity at the trailing edge was decreased, causing a decrease in trailing-edge and fluctuating lift noise. Cutback of the nozzle sides improved flow attachment and reduced far-field noise. The best flow attachment and least trailing-edge noise generally were obtained with a 40 deg external deflector configuration and a cutback nozzle with a 40 deg roof angle.

  16. Flapping of Insectile Wings

    NASA Astrophysics Data System (ADS)

    Huang, Yangyang; Kanso, Eva

    2015-11-01

    Insects use flight muscles attached at the base of the wings to produce impressive wing flapping frequencies. Yet the effects of muscle stiffness on the performance of insect wings remain unclear. Here, we construct an insectile wing model, consisting of two rigid wings connected at their base by an elastic torsional spring and submerged in an oscillatory flow. The wing system is free to rotate and flap. We first explore the extent to which the flyer can withstand roll perturbations, then study its flapping behavior and performance as a function of spring stiffness. We find an optimal range of spring stiffness that results in large flapping amplitudes, high force generation and good storage of elastic energy. We conclude by conjecturing that insects may select and adjust the muscle spring stiffness to achieve desired movement. These findings may have significant implications on the design principles of wings in micro air-vehicles.

  17. Status review of a supersonically-biased fighter wing-design study

    NASA Technical Reports Server (NTRS)

    Wood, R. M.; Miller, D. S.; Hahne, D. E.; Niedling, L. G.; Klein, J. R.

    1983-01-01

    Results from an ongoing supersonically-biased fighter wing-design study are summarized. The study has been conducted to explore the effects of supersonic aerodynamic performance, transonic maneuvering, low-speed/high angle-of-attack characteristics, and airframe system integration requirements on fighter aircraft wing design. The approach adopted involves the theoretical and experimental investigation of four advanced aircraft configurations which differ only in wing geometry. Supersonic and low-speed/high angle-of-attack wind tunnel results have been obtained for 20 deg trapezoidal, 65 deg delta, 70/30-deg advanced cranked, and 70/66-deg advanced cranked wing configurations. The supersonic data show that the advanced cranked wings outperform the trapezoidal and delta wings at cruise and moderate lift conditions. Low-speed/high angle-of-attack results show that all wings have significant stability problems above an angle of attack of 20 deg. Aircraft sizing analysis results show that the advanced cranked-wing configurations are significantly lighter, based upon take-off gross weight, than either the trapezoidal or the delta wings.

  18. Forebody vortex control for suppressing wing rock on a highly-swept wing configuration

    NASA Technical Reports Server (NTRS)

    Suarez, Carlos J.; Kramer, Brian R.; Ayers, Bert; Malcolm, Gerald N.

    1992-01-01

    Free-to-roll tests were conducted in a wind tunnel with a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. A limit cycle oscillation was observed for angles of attack between 22 and 30 deg. In general, the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. Various blowing techniques were evaluated as means of wing rock suppression. Blowing tangentially aft from leeward side nozzles near the forebody tip can damp the roll motion at low blowing rates and stop it completely at higher blowing rates. At the high rates, significant vortex asymmetries are created, causing the model to stop at a non-zero roll angle. Forward blowing and alternating right/left pulsed blowing appear to be more efficient techniques for suppressing wing rock. The oscillations can be damped almost completely at lower blowing coefficients, and, apparently, no major vortex asymmetries are induced. Good agreement is observed between this study and previous water tunnel tests on the same configuration.

  19. Comparison of Theoretical Stresses and Deflections of Multicell Wings with Experimental Results Obtained from Plastic Models

    NASA Technical Reports Server (NTRS)

    Zender, George W

    1956-01-01

    The experimental deflections and stresses of six plastic multicell-wing models of unswept, delta, and swept plan form are presented and compared with previously published theoretical results obtained by the electrical analog method. The comparisons indicate that the theory is reliable except for the evaluation of stresses in the vicinity of the leading edge of delta wings and the leading and trailing edges of swept wings. The stresses in these regions are questionable, apparently because of simplifications employed in idealizing the actual structure for theoretical purposes and because of local effects of concentrated loads.

  20. The minimum drag of thin wings at supersonic speed according to Kogan's theory

    NASA Technical Reports Server (NTRS)

    Jones, Robert T.

    1989-01-01

    In the theory of thin lifting surfaces the minimum drag consistent with a given total lift occurs when the downwash, averaged between forward and reversed motion of the wing, has the same value at all points of the wing planform. In Kogan's theory the conditions for minimum drag are determined on the forward sloping characteristic surface touching the trailing edge of the wing and it is shown that such a surface plays the role of the Trefftz plane familiar in subsonic wing theory. This paper shows how Kogan's theory may be applied to determine the drag of elliptic wings at supersonic speed. It appears that such wings have lower drag than the conventional 'delta' wing.

  1. The Vortices Trapped above Low-aspect-ratio Wings

    NASA Astrophysics Data System (ADS)

    Tang, Jian

    2007-11-01

    A stationary vortex trapped above the nondelta, low-aspect-ratio wings was first obtained in 3D unsteady numerical simulation. Flow visualization was conducted in water-channel using hydrogen bubble. The results verify that there is a vortex trapped above the low-aspect-ratio wings and the stationary vortex is consisted of two semi-ball, anti-rotation vortices which are different from the leading edge vortices on the delta wing. This stationary vortex trapped above the nondelta, low-aspect-ratio wings is a new phenomenon, which is different from the leading edge vortex on the delta wing. The numerical results show that lift coefficient increase to 0.8 when incidence increases form 0^o to 30^o, the lift coefficient keeps this value up to 45^o--a very high stall angle. The numerical results indicate that the trapped vortex might be the source of the high stall angle of attack and nonlinear lift at high incidence. Accompanied with the low-aspect-ratio wing, the existence of the stationary vortex is thought to be related to the strong effects of tip vortices. Further experimental and numerical works have been undertaken, the results show that trapped vortices have variant shapes and different critical angels of attack.

  2. Slotted Aircraft Wing

    NASA Technical Reports Server (NTRS)

    McLean, James D. (Inventor); Witkowski, David P. (Inventor); Campbell, Richard L. (Inventor)

    2006-01-01

    A swept aircraft wing includes a leading airfoil element and a trailing airfoil element. At least one full-span slot is defined by the wing during at least one transonic condition of the wing. The full-span slot allows a portion of the air flowing along the lower surface of the leading airfoil element to split and flow over the upper surface of the trailing airfoil element so as to achieve a performance improvement in the transonic condition.

  3. Optimum hovering wing planform.

    PubMed

    Nabawy, Mostafa R A; Crowther, William J

    2016-10-01

    Theoretical analysis is used to identify the optimum wing planform of a flapping/revolving wing in hover. This solution is of interest as a benchmark to which hovering wing geometries driven by broader multidisciplinary evolutionary or engineering constraints can be compared. Furthermore, useful insights into the aerodynamic performance of untwisted hovering wings are delivered. It is shown that profile power is minimised by using an untwisted elliptical planform whereas induced power is minimised by a more highly tapered planform similar to that of a hummingbird. PMID:27329340

  4. Prediction of vortex flow characteristics of wings at subsonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Lamar, J. E.

    1975-01-01

    The leading-edge-suction analogy of Polhamus, which has been successful in the prediction of vortex lift characteristics on wings with pointed tips at subsonic and supersonic speeds, has recently been extended to account for the vortex flow characteristics for wings with side edges. Comparisons of experimental data and other currently used methods with the extended method are made for wings having side edges at subsonic and supersonic speeds. Recent data obtained for a low-aspect-ratio cropped-delta wing with various amounts of asymmetrical tip rake, simulating a roll control device, are also presented.

  5. A theory for lateral wing-tip blowing

    NASA Technical Reports Server (NTRS)

    Tavella, D.; Roberts, L.

    1985-01-01

    The concept of lateral blowing consists in utilizing thin jets of air, which are ejected in the spanwise direction from slots at the tips of straight and swept wings, or along the leading edges of delta wings, to generate aerodynamic forces without the assistance of deflecting solid surfaces. For weak intensities of blowing the so-generated forces could be used for roll and lateral control of aircraft. In this work a theory for this concept as applied to straight wings is presented, revealing the analytical relationship between blowing and aerodynamic forces. The approach is based on perturbing the span of an elliptically loaded wing. Scaling laws involving blowing intensity, aspect ratio, and angle of attack are derived and compared with experiments. It is concluded that this concept has potential as a novel roll and lateral control device.

  6. Wing planform effects at supersonic speeds for an advanced fighter configuration

    NASA Technical Reports Server (NTRS)

    Wood, R. M.; Miller, D. S.

    1984-01-01

    Four advanced fighter configurations, which differed in wing planform and airfoil shape, were investigated in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Supersonic data were obtained on the four uncambered wings, which were each attached to a single fighter fuselage. The fuselage geometry varied in cross-sectional shape and had two side-mounted, flow-through, half-axisymmetric inlets. Twin vertical tails were attached to the fuselage. The four planforms tested were a 65 deg delta wing, a combination of a 20 deg trapezoidal wing and a 45 deg horizontal tail, a 70 deg/30 deg cranked wing, and a 70 deg/66 deg crank wing, where the angle values refer to the leading-edge sweep angle of the lifting-surface planform. Planform effects on a single fuselage representative of an advanced fighter aircraft were studied. Results show that the highly swept cranked wings exceeded the aerodynamic performance levels, at low lift coefficients, of the 65 deg delta wing and the 20 deg trapezoidal wing at trimmed and untrimmed conditions.

  7. A comparison of the aerodynamic characteristics at transonic speeds of four wing-fuselage configurations as determined from different test techniques, 4 October 1960

    NASA Technical Reports Server (NTRS)

    Donlan, C. J.; Myers, B. C., II; Mattson, A. T.

    1976-01-01

    The high speed aerodynamic characteristics of a family of four wing-fuselage configurations of 0, 35, 45, and 60 deg sweepback were determined from transonic bump model tests that were conducted in the Langley high speed 7 by 10 foot tunnel; sting supported model tests were conducted in the Langley 8 foot high speed tunnel and in the Langley high speed 7 by 10 foot tunnel, and rocket model tests were conducted by the Langley Pilotless Aircraft Research Division. A complementary study of the effect of Mach number gradients and streamline curvature on bump results is also included. The qualitative data obtained from the various test facilities for the wing-fuselage configurations were in essential agreement as regards the relative effects of sweepback and Mach number except for drag at zero lift. Quantitatively, important differences were present.

  8. AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading-and Trailing-Edge Slot Blowing

    NASA Technical Reports Server (NTRS)

    Horne, Clifton; Burnside, Nathan J.

    2013-01-01

    Aeroacoustic measurements of the 11 % scale full-span AMELIA CESTOL model with leading- and trailing-edge slot blowing circulation control (CCW) wing were obtained during a recent test in the Arnold Engineering Development Center 40- by 80-Ft. Wind Tunnel at NASA Ames Research Center, Sound levels and spectra were acquired with seven in-flow microphones and a 48-element phased microphone array for a variety of vehicle configurations, CCW slot flow rates, and forward speeds, Corrections to the measurements and processing are in progress, however the data from selected configurations presented in this report confirm good measurement quality and dynamic range over the test conditions, Array beamform maps at 40 kts tunnel speed show that the trailing edge flap source is dominant for most frequencies at flap angles of 0deg and 60deg, The overall sound level for the 60deg flap was similar to the 0deg flap for most slot blowing rates forward of 90deg incidence, but was louder by up to 6 dB for downstream angles, At 100 kts, the in-flow microphone levels were louder than the sensor self-noise for the higher blowing rates, while passive and active background noise suppression methods for the microphone array revealed source levels as much as 20 dB lower than observed with the in-flow microphones,

  9. Mississippi Delta

    NASA Technical Reports Server (NTRS)

    2002-01-01

    The streamers of clouds draped over the Gulf of Mexico in this true-color MODIS image from February 27, 2002, suggest that a cold, dry wind was blowing southward over the United States and began to pick up moisture over the Gulf, causing these strips of clouds. That the clouds didn't pick up until some distance from the coastline allowed MODIS to get a perfect view of the dynamic Gulf Coast environment spanning (left to right) Texas, Louisiana, Mississippi, Alabama, and Florida's Western Panhandle. The Mississippi River runs roughly down the center of the image, and is joined in Louisiana by the Red River coming in from the northwest. Over the past 7000 years, the actual delta, where the main river channel empties into the Gulf, has wandered around what we now think of as the Louisiana coast. Considering all the sediment visible in this image, it's not hard to imagine that the river carries about 2.4 billion kilograms of sediment into the Gulf each year. Deposition of some of this sediment has been building up the current delta, called the Birdfoot Delta, for obvious reasons, for about 700 years. The coastal waters are alive with microscopic organisms called phytoplankton, which contain colorful pigments, including chlorophyll, for harvesting sunlight. Beyond the sediment plume off Louisiana, the waters are very dark, which could indicate that a large amount of chlorophyll is present, absorbing lots of sunlight and causing the water to appear dark. Farther south, the waters appear bright blue, which could be a signature of coccolithophores, which use highly reflective calcium carbonate to build scaly coverings for themselves. The brighter offshore waters could also be caused by a blue-green algae called Trichodesmium, an organism that can not only harness carbon dioxide for photosynthesis, but can also take nitrogen from the air and turn it into a form that can be used by living organisms. Credit: Jacques Descloitres, MODIS Land Rapid Response Team, NASA/GSFC

  10. Wing-Wake Interactions between Ipsilateral Wings in Dragonfly Flight

    NASA Astrophysics Data System (ADS)

    Dong, Haibo; Liang, Zongxian

    2009-11-01

    Bilateral and ipsilateral wing-wing interactions can be commonly observed in insect flights. As a representative example of ipsilateral wing-wing interaction, dragonflies in flight have been widely studied. An important fact is that the flow over their hindwings is affected by the presence of the forewings. Wake capture and phase-change play very important role on aerodynamic performance of the hindwings We present a direct numerical simulation of a modeled dragonfly (Aeshna juncea) in slow flight as studied in Azuma et al (JEB 1985). Realistic morphologies of wing, body, and kinematics are used for maximum including wing and body features of a dragonfly. This work aims to study the relations between wake-topology and aerodynamic performance due to wing-wing and wing-wake interactions of dragonfly ipsilateral wings. DNS results are also compared with Local Momentum Theory (Azuma et al).

  11. Effective Control of Computationally Simulated Wing Rock in Subsonic Flow

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Menzies, Margaret A.

    1997-01-01

    The unsteady compressible, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the delta wing rock phenomenon. The NS equations are solved time accurately, using the implicit, upwind, Roe flux-difference splitting, finite-volume scheme. The rigid-body dynamics equations are solved using a four-stage Runge-Kutta scheme. Once the wing reaches the limit-cycle response, an active control model using a mass injection system is applied from the wing surface to suppress the limit-cycle oscillation. The active control model is based on state feedback and the control law is established using pole placement techniques. The control law is based on the feedback of two states: the roll-angle and roll velocity. The primary model of the computational applications consists of a 80 deg swept, sharp edged, delta wing at 30 deg angle of attack in a freestream of Mach number 0.1 and Reynolds number of 0.4 x 10(exp 6). With a limit-cycle roll amplitude of 41.1 deg, the control model is applied, and the results show that within one and one half cycles of oscillation, the wing roll amplitude and velocity are brought to zero.

  12. Sonic-box method employing local Mach number for oscillating wings with thickness

    NASA Technical Reports Server (NTRS)

    Ruo, S. Y.

    1978-01-01

    A computer program was developed to account approximately for the effects of finite wing thickness in the transonic potential flow over an oscillating wing of finite span. The program is based on the original sonic-box program for planar wing which was previously extended to include the effects of the swept trailing edge and the thickness of the wing. Account for the nonuniform flow caused by finite thickness is made by application of the local linearization concept. The thickness effect, expressed in terms of the local Mach number, is included in the basic solution to replace the coordinate transformation method used in the earlier work. Calculations were made for a delta wing and a rectangular wing performing plunge and pitch oscillations, and the results were compared with those obtained from other methods. An input quide and a complete listing of the computer code are presented.

  13. Flying wings / flying fuselages

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.

    2001-01-01

    The present paper has documented the historical relationships between various classes of all lifting vehicles, which includes the flying wing, all wing, tailless, lifting body, and lifting fuselage. The diversity in vehicle focus was to ensure that all vehicle types that map have contributed to or been influenced by the development of the classical flying wing concept was investigated. The paper has provided context and perspective for present and future aircraft design studies that may employ the all lifting vehicle concept. The paper also demonstrated the benefit of developing an understanding of the past in order to obtain the required knowledge to create future concepts with significantly improved aerodynamic performance.

  14. Slotted Aircraft Wing

    NASA Technical Reports Server (NTRS)

    Vassberg, John C. (Inventor); Gea, Lie-Mine (Inventor); McLean, James D. (Inventor); Witowski, David P. (Inventor); Krist, Steven E. (Inventor); Campbell, Richard L. (Inventor)

    2006-01-01

    An aircraft wing includes a leading airfoil element and a trailing airfoil element. At least one slot is defined by the wing during at least one transonic condition of the wing. The slot may either extend spanwise along only a portion of the wingspan, or it may extend spanwise along the entire wingspan. In either case, the slot allows a portion of the air flowing along the lower surface of the leading airfoil element to split and flow over the upper surface of the trailing airfoil element so as to achieve a performance improvement in the transonic condition.

  15. Mississippi Delta

    NASA Technical Reports Server (NTRS)

    2002-01-01

    The Mississippi River delta teems with sediment deposited by the river as it flows into the Gulf of Mexico in this true-color image captured by MODIS on October 15, 2001. The sediment, which is marked by brown swirls in the Gulf, provides nutrients for the bloom of phytoplankton visible as blue-green swirls off the coastline. In the high-resolution image the city of Memphis can be seen in the southwest corner of Tennessee, which is just to left of center at the top of the image. The brown coloration that encompasses Memphis and either side of the river, as flows north to south along the left side of the image, is the river's flood plain. Also visible, in the upper-right hand corner of the image is the southern end of the Appalachian Mountains.

  16. Low subsonic aerodynamic characteristics of five irregular planform wings with systematically varying wing fillet geometry tested in the NASA/Ames 12 foot pressure tunnel (LA65)

    NASA Technical Reports Server (NTRS)

    Ball, J. W.; Watson, D. B.

    1976-01-01

    An experimental and analytical aerodynamic program to develop predesign guides for irregular planform wings (also referred to as cranked leading edge or double delta 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. Because subsonic and hypersonic conditions were the two prime areas of concern in the initial application of this program to optimize shuttle orbiter landing and entry characteristics, the study was designated the Subsonic/Hypersonic Irregular Planforms Study (SHIPS).

  17. Analysis of iced wings

    NASA Technical Reports Server (NTRS)

    Cebeci, T.; Chen, H. H.; Kaups, K.; Schimke, S.; Shin, J.

    1992-01-01

    A method for computing ice shapes along the leading edge of a wing and a method for predicting its aerodynamic performance degradation due to icing is described. Ice shapes are computed using an extension of the LEWICE code which was developed for airfoils. The aerodynamic properties of the iced wing are determined with an interactive scheme in which the solutions of the inviscid flow equations are obtained from a panel method and the solutions of the viscous flow equations are obtained from an inverse three-dimensional finite-difference boundary-layer method. A new interaction law is used to couple the inviscid and viscous flow solutions. The application of the LEWICE wing code to the calculation of ice shapes on a MS-317 swept wing shows good agreement with measurements. The interactive boundary-layer method is applied to a tapered ice wing in order to study the effect of icing on the aerodynamic properties of the wing at several angles of attack.

  18. Wind-tunnel investigation of a large-scale VTOL aircraft model with wing root and wing thrust augmentors. [Ames 40 by 80 foot wind tunnel

    NASA Technical Reports Server (NTRS)

    Aoyagi, K.; Aiken, T. N.

    1979-01-01

    Tests were conducted in the Ames 40 by 80 foot wind tunnel to determine the aerodynamic characteristics of a large-scale V/STOL aircraft model with thrust augmentors. The model had a double-delta wing of aspect ratio 1.65 with augmentors located in the wing root and the wing trailing edge. The supply air for the augmentor primary nozzles was provided by the YJ-97 turbojet engine. The airflow was apportioned approximately 74 percent to the wing root augmentor and 24 percent to wing augmentor. Results were obtained at several trailing-edge flap deflections with the nozzle jet-momentum coefficients ranging from 0 to 7.9. Three-component longitudinal data are presented with the agumentor operating with and without the horizontal tail. A limited amount of six component data are also presented.

  19. Aerodynamic control of NASP-type vehicles through vortex manipulation. Volume 3: Wing rock experiments

    NASA Technical Reports Server (NTRS)

    Suarez, Carlos J.; Smith, Brooke C.; Kramer, Brian R.; Ng, T. Terry; Ong, Lih-Yenn; Malcolm, Gerald N.

    1993-01-01

    Free-to-roll tests were conducted in water and wind tunnels in an effort to investigate the mechanisms of wing rock on a NASP-type vehicle. The configuration tested consisted of a highly-slender forebody and a 78 deg swept delta wing. In the water tunnel test, extensive flow visualization was performed and roll angle histories were obtained. In the wind tunnel test, the roll angle, forces and moments, and limited forebody and wing surface pressures were measured during the wing rock motion. A limit cycle oscillation was observed for angles of attack between 22 deg and 30 deg. In general, the experiments confirmed that the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. The variation of roll acceleration (determined from the second derivative of the roll angle time history) with roll angle clearly slowed the energy balance necessary to sustain the limit cycle oscillation. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetrices are created, causing the model to stop at a non-zero roll angle. On the other hand, alternating pulsed blowing on the left and right sides of the fore body was demonstrated to be a potentially effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

  20. An experimental study of the nonlinear dynamic phenomenon known as wing rock

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    An experimental investigation into the physical phenomena associated with limit cycle wing rock on slender delta wings has been conducted. The model used was a slender flat plate delta wing with 80-deg leading edge sweep. The investigation concentrated on three main areas: motion characteristics obtained from time history plots, static and dynamic flow visualization of vortex position, and static and dynamic flow visualization of vortex breakdown. The flow visualization studies are correlated with model motion to determine the relationship between vortex position and vortex breakdown with the dynamic rolling moments. Dynamic roll moment coefficient curves reveal rate-dependent hysteresis, which drives the motion. Vortex position correlated with time and model motion show a time lag in the normal position of the upward moving wing vortex. This time lag may be the mechanism responsible for the hysteresis. Vortex breakdown is shown to have a damping effect on the motion.

  1. Delta III—an evolutionary delta growth

    NASA Astrophysics Data System (ADS)

    Arvesen, R. J.; Simpson, J. S.

    1996-03-01

    In order to remain competitive in the future and expand the McDonnell Douglas Aerospace market share, MDA has developed an expendable launch system strategy that devices cost-effective launch systems from the Delta II with a growth vehicle configuration called Delta III. The Delta III evolves from the Delta II launch system through development of a larger payload fairing (4-meter diameter), new cryogenically propelled upper stage, new first stage fuel tank, and larger strap-on solid rocket motors. We are developing the Delta III using Integrated Product Development Teams that capitalize on the experience base that has led us to a world record breaking mission success of 49 consecutive Delta II missions. The Delta III first-launch capability is currently planned for the spring of 1998 in support of our first spacecraft customer, Hughes Space and Communications International.

  2. Nonplanar wing load-line and slender wing theory

    NASA Technical Reports Server (NTRS)

    Deyoung, J.

    1977-01-01

    Nonplanar load line, slender wing, elliptic wing, and infinite aspect ratio limit loading theories are developed. These are quasi two dimensional theories but satisfy wing boundary conditions at all points along the nonplanar spanwise extent of the wing. These methods are applicable for generalized configurations such as the laterally nonplanar wing, multiple nonplanar wings, or wing with multiple winglets of arbitrary shape. Two dimensional theory infers simplicity which is practical when analyzing complicated configurations. The lateral spanwise distribution of angle of attack can be that due to winglet or control surface deflection, wing twist, or induced angles due to multiwings, multiwinglets, ground, walls, jet or fuselage. In quasi two dimensional theory the induced angles due to these extra conditions are likewise determined for two dimensional flow. Equations are developed for the normal to surface induced velocity due to a nonplanar trailing vorticity distribution. Application examples are made using these methods.

  3. Prediction and control of slender-wing rock

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1992-01-01

    The unsteady Euler equations and the Euler equations of rigid-body dynamics, both written in the moving frame of reference, are sequentially solved to simulate the limit-cycle rock motion of slender delta wings. The governing equations of the fluid flow and the dynamics of the present multidisciplinary problem are solved using an implicit, approximately-factored, central-difference-like, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. For the control of wing-rock motion, leading-edge flaps are forced to oscillate anti-symmetrically at prescribed frequency and amplitude, which are tuned in order to suppress the rock motion. Since the computational grid deforms due to the leading-edge flaps motion, the grid is dynamically deformed using the Navier-displacement equations. Computational applications cover locally-conical and three-dimensional solutions for the wing-rock simulation and its control.

  4. Formulas Pertinent to the Calculation of Flow-Field Effects at Supersonic Speeds Due to Wing Thickness

    NASA Technical Reports Server (NTRS)

    Margolis, Kenneth; Elliott, Miriam H.

    1959-01-01

    Expressions based on linearized supersonic-flow theory are derived for the perturbation velocity potential in space due to wing thickness for rectangular wings with biconvex airfoil sections and for arrow, delta, and quadrilateral wings with wedge-type airfoil sections. The complete range of supersonic speeds is considered subject to a minor aspect-ratio-Mach number restriction for the rectangular plan form and to the condition that the trailing edge is supersonic for the sweptback wings. The formulas presented can be utilized in determining the induced-flow characteristics at any point in the field and are readily adaptable for either numerical computation or analytical determination of any velocity components desired.

  5. Flight test results from a supercritical mission adaptive wing with smooth variable camber

    NASA Technical Reports Server (NTRS)

    Powers, Sheryll Goecke; Webb, Lannie D.; Friend, Edward L.; Lokos, William A.

    1992-01-01

    The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-ideal wing camber shape for any flight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 deg, a Mach number of 0.85, leading and trailing-edge cambers (delta(sub LE/TE)) of 0/2 and 5/10, and angles of attack from 3.0 deg to 14.0 deg. For the well-behaved flow of the delta(sub LE/TE) = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the delta(sub LE/TE) = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.

  6. The characterization of tandem and corrugated wings

    NASA Astrophysics Data System (ADS)

    Lian, Yongsheng; Broering, Timothy; Hord, Kyle; Prater, Russell

    2014-02-01

    Dragonfly wings have two distinct features: a tandem configuration and wing corrugation. Both features have been extensively studied with the aim to understand the superior flight performance of dragonflies. In this paper we review recent development of tandem and corrugated wing aerodynamics. With regards to the tandem configuration, this review will focus on wing/wing and wing/vortex interactions at different flapping modes and wing spacing. In addition, the aerodynamics of tandem wings under gusty conditions will be reviewed and compared with isolated wings to demonstrate the gust resistance characteristics of flapping wings. Regarding corrugated wings, we review their structural and aerodynamic characteristics.

  7. Numerical study of delta wing leading edge blowing

    NASA Technical Reports Server (NTRS)

    Yeh, David; Tavella, Domingo; Roberts, Leonard

    1988-01-01

    Spanwise and tangential leading edge blowing as a means of controlling the position and strength of the leading edge vortices are studied by numerical solution of the three-dimensional Navier-Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of spanwise leading edge blowing at moderate angle of attack magnifies the size and strength of the leading edge vortices, and moves the vortex cores outboard and upward. The increase in lift primarily comes from the greater nonlinear vortex lift. However, spanwise blowing causes earlier vortex breakdown, thus decreasing the stall angle. The effects of tangential blowing at low to moderate angles of attack tend to reduce the pressure peaks associated with leading edge vortices and to increase the suction peak around the leading edge, so that the integrated value of the surface pressure remains about the same. Tangential leading edge blowing in post-stall conditions is shown to re-establish vortical flow and delay vortex bursting, thus increasing C sub L sub max and stall angle.

  8. Theory of wing rock

    NASA Technical Reports Server (NTRS)

    Hsu, C. H.; Lan, C. E.

    1984-01-01

    A theory is developed for predicting wing rock characteristics. From available data, it can be concluded that wing rock is triggered by flow asymmetries, developed by negative or weakly positive roll damping, and sustained by nonlinear aerodynamic roll damping. A new nonlinear aerodynamic model that includes all essential aerodynamic nonlinearities is developed. The Beecham-Titchener method is applied to obtain approximate analytic solutions for the amplitude and frequency of the limit cycle based on the three degree-of-freedom equations of motion. An iterative scheme is developed to calculate the average aerodynamic derivatives and dynamic characteristics at limit cycle conditions. Good agreement between theoretical and experimental results is obtained.

  9. Status of wing flutter

    NASA Technical Reports Server (NTRS)

    Kussner, H G

    1936-01-01

    This report presents a survey of previous theoretical and experimental investigations on wing flutter covering thirteen cases of flutter observed on airplanes. The direct cause of flutter is, in the majority of cases, attributable to (mass-) unbalanced ailerons. Under the conservative assumption that the flutter with the phase angle most favorable for excitation occurs only in two degrees of freedom, the lowest critical speed can be estimated from the data obtained on the oscillation bench. Corrective measures for increasing the critical speed and for definite avoidance of wing flutter, are discussed.

  10. Wind-tunnel roll-damping measurements of a winged space shuttle configuration in launch attitude

    NASA Technical Reports Server (NTRS)

    Hess, R. W.; Davenport, E. E.

    1973-01-01

    Ground-wind load studies were conducted on three model configurations to assess the importance of aeroelastic instabilities of erected space shuttle vehicles. Roll damping was measured on a fuselage-alone model, which had a D cross section, and a fuselage and tail surfaces in combination with either a clipped-delta wing or a low-sweep tapered wing as the primary lifting surface. The largest negative roll-damping coefficients were measured with the fuselage-alone configuration and were a function of wind azimuth. At the wind azimuths at which the wing-fuselage configuration was unstable, the negative roll-damping coefficients were a function of reduced frequency.

  11. Low-speed aerodynamic characteristics of a wing-canard configuration with underwing spanwise blowing on the trailing-edge flap system

    NASA Technical Reports Server (NTRS)

    Banks, Daniel W.; Paulson, John W., Jr.

    1987-01-01

    An investigation of the effects of spanwise blowing applied to the lower surface of a trailing-edge flap system on a wing-canard configuration has been conducted in the Langley 4- by 7-Meter Tunnel. The investigation studied spanwise-blowing angles of 30 deg., 45 deg., and 60 deg. measured from a perpendicular to the body center-line. The test conditions covered a range of free-stream dynamic pressures up to 50 psf for thrust coefficients up to 2.1 over a range of angles of attack from -2 deg. to 26 deg. Model height above the wind tunnel floor was varied from a height-to-span ratio of 1.70 down to 0.20 (a representative wheel touchdown height). The results indicate that blowing angles of 30 deg. and 45 deg. increase the induced-lift increment produced by spanwise blowing on the lower surface of a trailing-edge flap system. Increasing the blowing angle to 60 deg., in general, produces little further improvement.

  12. Calculation of tapered monoplane wings

    NASA Technical Reports Server (NTRS)

    Amstutz, E

    1930-01-01

    The tapered wing shape increases the lift in the middle of the wing and thus reduces the bending moment of the lifting forces in the plane of symmetry. Since this portion of the wing is the thickest, the stresses of the wing material are reduced and desirable space is provided for stowing the loads in the wing. This statically excellent form of construction, however, has aerodynamic disadvantages which must be carefully weighed, if failures are to be avoided. This treatise is devoted to the consideration of these problems.

  13. Autosomal Mutations Affecting Adhesion between Wing Surfaces in Drosophila Melanogaster

    PubMed Central

    Prout, M.; Damania, Z.; Soong, J.; Fristrom, D.; Fristrom, J. W.

    1997-01-01

    Integrins are evolutionarily conserved transmembrane α,β heterodimeric receptors involved in cell-to-matrix and cell-to-cell adhesions. In Drosophila the position-specific (PS) integrins mediate the formation and maintenance of junctions between muscle and epidermis and between the two epidermal wing surfaces. Besides integrins, other proteins are implicated in integrin-dependent adhesion. In Drosophila, somatic clones of mutations in PS integrin genes disrupt adhesion between wing surfaces to produce wing blisters. To identify other genes whose products function in adhesion between wing surfaces, we conducted a screen for autosomal mutations that produce blisters in somatic wing clones. We isolated 76 independent mutations in 25 complementation groups, 15 of which contain more than one allele. Chromosomal sites were determined by deficiency mapping, and genetic interactions with mutations in the β(PS) integrin gene myospheroid were investigated. Mutations in four known genes (blistered, Delta, dumpy and mastermind) were isolated. Mutations were isolated in three new genes (piopio, rhea and steamer duck) that affect myo-epidermal junctions or muscle function in embryos. Mutations in three other genes (kakapo, kiwi and moa) may also affect cell adhesion or muscle function at hatching. These new mutants provide valuable material for the study of integrin-dependent cell-to-cell adhesion. PMID:9136017

  14. delta-Hexachlorocyclohexane (delta-HCH)

    Integrated Risk Information System (IRIS)

    delta - Hexachlorocyclohexane ( delta - HCH ) ; CASRN 319 - 86 - 8 Human health assessment information on a chemical substance is included in the IRIS database only after a comprehensive review of toxicity data , as outlined in the IRIS assessment development process . Sections I ( Health Hazard Ass

  15. Low-speed tests of a high-aspect-ratio, supercritical-wing transport model equipped with a high-lift flap system in the Langley 4- by 7-meter and Ames 12-foot pressure tunnels

    NASA Technical Reports Server (NTRS)

    Morgan, H. L., Jr.; Kjelgaard, S. O.

    1983-01-01

    The Ames 12-Foot Pressure Tunnel was used to determine the effects of Reynolds number on the static longitudinal aerodynamic characteristics of an advanced, high-aspect-ratio, supercritical wing transport model equipped with a full span, leading edge slat and part span, double slotted, trailing edge flaps. The model had a wing span of 7.5 ft and was tested through a free stream Reynolds number range from 1.3 to 6.0 x 10 to 6th power per foot at a Mach number of 0.20. Prior to the Ames tests, an investigation was also conducted in the Langley 4 by 7 Meter Tunnel at a Reynolds number of 1.3 x 10 to 6th power per foot with the model mounted on an Ames strut support system and on the Langley sting support system to determine strut interference corrections. The data obtained from the Langley tests were also used to compare the aerodynamic charactertistics of the rather stiff, 7.5-ft-span steel wing model tested during this investigation and the larger, and rather flexible, 12-ft-span aluminum-wing model tested during a previous investigation. During the tests in both the Langley and Ames tunnels, the model was tested with six basic wing configurations: (1) cruise; (2) climb (slats only extended); (3) 15 deg take-off flaps; (4) 30 deg take-off flaps; (5) 45 deg landing flaps; and (6) 60 deg landing flaps.

  16. ACTE Wing Loads Analysis

    NASA Technical Reports Server (NTRS)

    Horn, Nicholas R.

    2015-01-01

    The Adaptive Compliant Trailing Edge (ACTE) project modified a Gulfstream III (GIII) aircraft with a new flexible flap that creates a seamless transition between the flap and the wing. As with any new modification, it is crucial to ensure that the aircraft will not become overstressed in flight. To test this, Star CCM a computational fluid dynamics (CFD) software program was used to calculate aerodynamic data for the aircraft at given flight conditions.

  17. Wing on a String

    ERIC Educational Resources Information Center

    Fitzgerald, Mike; Brand, Lance

    2004-01-01

    In this article, the authors present an activity that shows students how flight occurs. The "wing on a string" is a simple teacher-made frame that consists of PVC pipe, fishing line, and rubber bands--all readily available hardware store items. The only other materials/tools involved are a sheet of paper, some pieces of a soda straw, a stapler,…

  18. Flight investigation of the effect of tail configuration on stall, spin, and recovery characteristics of a low-wing general aviation research airplane

    NASA Technical Reports Server (NTRS)

    Stough, H. Paul, III; Patton, James M., Jr.; Sliwa, Steven M.

    1987-01-01

    Flight tests were performed to investigate the stall, spin, and recovery characteristics of a low-wing, single-engine, light airplane with four interchangeable tail configurations. The four tail configurations were evaluated for effects of varying mass distribution, center-of-gravity position, and control inputs. The airplane tended to roll-off at the stall. Variations in tail configuration produced spins ranging from 40 deg to 60 deg angle of attack and turn rates of about 145 to 208 deg/sec. Some unrecoverable flat spins were encountered which required use of the airplane spin chute for recovery. For recoverable spins, antispin rudder followed by forward wheel with ailerons centered provided the quickest spin recovery. The moderate spin modes agreed very well with those predicted from spin-tunnel model tests, however, the flat spin was at a lower angle of attack and a slower rotation rate than indicated by the model tests.

  19. Simulation and analysis of wing rock physics for a generic fighter model with three degrees-of-freedom

    NASA Astrophysics Data System (ADS)

    Saad, Ahmed Abdelaziz

    2000-12-01

    Modern fighter designs have been associated with lateral self-excited oscillations known as "Wing Rock". In this study, for the first time, wing rock is computationally simulated in three DoF: roll, sideslip, and vertical motion to study the effect of adding the sideslip and vertical motion. The results are for a generic fighter model consisting of a fore-body, a cropped delta wing, and a vertical fin. The effect of including the vertical fin is also studied. The interaction of aerodynamics and rigid-body dynamics during a single DoF wing rock for the wing-body configuration has been studied via snap shots of a cross-plane stagnation pressure distribution and tracing the instantaneous locations of vortex burst for an entire cycle of wing rock. An innovative explanation of the fluid mechanism that drives and sustains the motion has been introduced. The effect of adding the sideslip and vertical motion DoF to the simulations of the wing-body configuration was found to delay the onset and to reduce the amplitude of wing rock by about 50% with surprisingly no change in frequency. The wing rock simulation in three DoF was repeated for the full generic fighter model with the fin included. The aerodynamic effect of the fin was found to significantly delay the vortex burst on the upper surface of the wing. The net effect of the fin was found to augment the damping of the oscillations with significant increase in frequency.

  20. Aerodynamic cause of the asymmetric wing deformation of insect wings

    NASA Astrophysics Data System (ADS)

    Luo, Haoxiang; Tian, Fangbao; Song, Jialei; Lu, Xi-Yun

    2012-11-01

    Insect wings typically exhibit significant asymmetric deformation patterns, where the magnitude of deflection during upstroke is greater than during downstroke. Such a feature is beneficial for the aerodynamics since it reduces the projected wing area during upstroke and leads to less negative lift. Previously, this asymmetry has been mainly attributed to the directional bending stiffness in the wing structure, e.g., one-way hinge, or a pre-existing camber in the wing surface. In the present study, we demonstrate that the asymmetric pattern can also be caused by the asymmetric force due to the flow, while the wing structure and kinematics are symmetric. A two-dimensional translating/pitching wing in a free stream is used as the model, and the wing is represented by an elastic sheet with large displacement. The result shows that, interestingly, the wing experiences larger deformation during upstroke even though the aerodynamic force is greater during downstroke. The physical mechanism of the phenomenon can be explained by the modulating effect of the aerodynamic force on the timing of storage/release of the elastic energy in the wing. Supported by NSF (No. CBET-0954381).

  1. An experimental study of spanwise flow effects on lift generation in flapping wings

    NASA Astrophysics Data System (ADS)

    Hong, Youngsun

    Using a combination of force transducer measurement to quantify net lift force, a high frame rate camera to quantify and subtract inertial contributions, and Digital Particle Image Velocimetry (DPIV) to calculate aerodynamic contributions in the spanwise plane, the contribution of spanwise flow to the generation of lift force in wings undergoing a pure flapping motion in hover is shown as a function of flapping angle throughout the flapping cycle. When flapping a flat plate wing and a wing of identical wing area and aspect ratio, but cambered in span (both wings in hover with no change in pitch), the spanwise cambered wing was found to generate a greater mean lift force through the whole flap cycle under the same acceleration. However, depending on the angle in flapping arc, the spanwise cambered wing can generate less lift than the flat wing. Additionally, since the lift force generated by the wingtip vortex in the spanwise plane resulting from the flapping motion has yet to be directly quantified, the wingtip vortex is investigated to determine precisely how it augments the lift force through the various phases in the flapping motion. Vortices in the vicinity of the wingtip generate lift force in the spanwise plane of flapping wings. In classical fixed wing aerodynamics, the presence of wing tip vortices has been shown to increase the lift locally near the tip. Also, the impingement of large vortices on the upper surface of delta wings is considered to contribute largely to the lift force at higher angles of attack. This study determined that vortices in the spanwise plane (streamwise vorticity) generate lift in a similar manner in flapping wings. Using a mechanical ornithopter with wings fabricated in-house, vortices were identified at several different locations along the span of the wing, and at numerous different points throughout the flapping cycle under a variety of operating conditions. The lift generated by these spanwise planar oriented vortices was

  2. Fog spontaneously folds mosquito wings

    NASA Astrophysics Data System (ADS)

    Dickerson, Andrew K.; Liu, Xing; Zhu, Ting; Hu, David L.

    2015-02-01

    The flexibility of insect wings confers aerodynamic benefits, but can also present a hazard if exposed to fog or dew. Fog can cause water to accumulate on wings, bending them into tight taco shapes and rendering them useless for flight. In this combined experimental and theoretical study, we use high-speed video to film the spontaneous folding of isolated mosquito wings due to the evaporation of a water drop. We predict shapes of the deformed wing using two-dimensional elastica theory, considering both surface tension and Laplace pressure. We also recommend fold-resistant geometries for the wings of flapping micro-aerial vehicles. Our work reveals the mechanism of insect wing folding and provides a framework for further study of capillarity-driven folding in both natural and biomimetic systems at small scales.

  3. Wing flapping with minimum energy

    NASA Technical Reports Server (NTRS)

    Jones, R. T.

    1980-01-01

    A technique employed by Prandtl and Munk is adapted for the case of a wing in flapping motion to determine its lift distribution. The problem may be reduced to one of minimizing induced drag for a specified and periodically varying bending moment at the wing root. It is concluded that two wings in close tandem arrangement, moving in opposite phase, would eliminate the induced aerodynamic losses calculated

  4. Simulation of iced wing aerodynamics

    NASA Technical Reports Server (NTRS)

    Potapczuk, M. G.; Bragg, M. B.; Kwon, O. J.; Sankar, L. N.

    1991-01-01

    The sectional and total aerodynamic load characteristics of moderate aspect ratio wings with and without simulated glaze leading edge ice were studied both computationally, using a three dimensional, compressible Navier-Stokes solver, and experimentally. The wing has an untwisted, untapered planform shape with NACA 0012 airfoil section. The wing has an unswept and swept configuration with aspect ratios of 4.06 and 5.0. Comparisons of computed surface pressures and sectional loads with experimental data for identical configurations are given. The abrupt decrease in stall angle of attack for the wing, as a result of the leading edge ice formation, was demonstrated numerically and experimentally.

  5. Freight Wing Trailer Aerodynamics

    SciTech Connect

    Graham, Sean; Bigatel, Patrick

    2004-10-17

    Freight Wing Incorporated utilized the opportunity presented by this DOE category one Inventions and Innovations grant to successfully research, develop, test, patent, market, and sell innovative fuel and emissions saving aerodynamic attachments for the trucking industry. A great deal of past scientific research has demonstrated that streamlining box shaped semi-trailers can significantly reduce a truck's fuel consumption. However, significant design challenges have prevented past concepts from meeting industry needs. Market research early in this project revealed the demands of truck fleet operators regarding aerodynamic attachments. Products must not only save fuel, but cannot interfere with the operation of the truck, require significant maintenance, add significant weight, and must be extremely durable. Furthermore, SAE/TMC J1321 tests performed by a respected independent laboratory are necessary for large fleets to even consider purchase. Freight Wing used this information to create a system of three practical aerodynamic attachments for the front, rear and undercarriage of standard semi trailers. SAE/TMC J1321 Type II tests preformed by the Transportation Research Center (TRC) demonstrated a 7% improvement to fuel economy with all three products. If Freight Wing is successful in its continued efforts to gain market penetration, the energy and environmental savings would be considerable. Each truck outfitted saves approximately 1,100 gallons of fuel every 100,000 miles, which prevents over 12 tons of CO2 from entering the atmosphere. If all applicable trailers used the technology, the country could save approximately 1.8 billion gallons of diesel fuel, 18 million tons of emissions and 3.6 billion dollars annually.

  6. 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.

  7. Control of a swept wing tailless aircraft through wing morphing

    NASA Astrophysics Data System (ADS)

    Guiler, Richard W.

    Inspired by flight in nature, work done by Lippisch, the Hortens, and Northrop offered insight to achieving the efficiency of bird flight with swept-wing tailless aircraft. Tailless designs must incorporate aerodynamic compromises for control, which have inhibited potential advantages. A morphing mechanism, capable of changing the twist of wing and that can also provide pitch, roll and yaw control for a tailless swept wing aircraft is the first step to a series of morphing techniques, which will lead to more fluid, bird-like flight. This research focuses on investigating the design of a morphing wing to improve the flight characteristics of swept wing Horten type tailless aircraft. Free flight demonstrators, wind tunnel flow visualization, wind-tunnel force and moment data along with CFD studies have been used to evaluate the stability, control and efficiency of a morphing swept wing tailless aircraft. A wing morphing mechanism for the control of a swept wing tailless aircraft has been developed. This new control technique was experimentally and numerically compared to an existing elevon equipped tailless aircraft and has shown the potential for significant improvement in efficiency. The feasibility of this mechanism was also validated through flight testing of a flight weight version. In the process of comparing the Horten type elevon equipped aircraft and the morphing model, formal wind tunnel verification of wingtip induced thrust, found in Horten (Bell Shaped Lift distribution) type swept wing tailless aircraft was documented. A more complete physical understanding of the highly complex flow generated in the control region of the morphing tailless aircraft has been developed. CFD models indicate the possibility of the presence of a Leading Edge Vortex (LEV) on the control section morphing wing when the tip is twisted between +3.5 degrees and +7 degrees. The presence of this LEV causes a reduction of drag while lift is increased. Similar LEVs have been

  8. Beetle wings are inflatable origami

    NASA Astrophysics Data System (ADS)

    Chen, Rui; Ren, Jing; Ge, Siqin; Hu, David

    2015-11-01

    Beetles keep their wings folded and protected under a hard shell. In times of danger, they must unfold them rapidly in order for them to fly to escape. Moreover, they must do so across a range of body mass, from 1 mg to 10 grams. How can they unfold their wings so quickly? We use high-speed videography to record wing unfolding times, which we relate to the geometry of the network of blood vessels in the wing. Larger beetles have longer unfolding times. Modeling of the flow of blood through the veins successfully accounts for the wing unfolding speed of large beetles. However, smaller beetles have anomalously short unfolding times, suggesting they have lower blood viscosity or higher driving pressure. The use of hydraulics to unfold complex objects may have implications in the design of micro-flying air vehicles.

  9. Hydrodynamics of penguin wing models

    NASA Astrophysics Data System (ADS)

    Noca, Flavio; Cuong Duong, Nhut; Herpich, Jerome

    2010-11-01

    The three-dimensional kinematics of penguin wings were obtained from movie footage in aquariums. A 1:1 scale model of the penguin wing (with an identical planform but with a flat section profile and a rigid configuration) was actuated with a robotic arm in a water channel. The experiments were performed at a chord Reynolds number of about 10^4 (an order of magnitude lower than for the observed penguin). The dynamics of the wing were analyzed with force and flowfield measurements. The two main results are: 1. a net thrust on both the upstroke and downstroke movement; 2. the occurence of a leading edge vortex (LEV) along the wing span. The effects of section profile, wing flexibility, and a higher Reynolds number will be investigated in the future.

  10. On the nonlinear aerodynamic and stability characteristics of a generic chine-forebody slender-wing fighter configuration

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Brandon, Jay M.

    1987-01-01

    An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.

  11. Determination of mean camber surfaces for wings having uniform chordwise loading and arbitrary spanwise loading in subsonic flow

    NASA Technical Reports Server (NTRS)

    Katzoff, S; Faison, M Frances; Dubose, Hugh C

    1954-01-01

    The field of a uniformly loaded wing in subsonic flow is discussed in terms of the acceleration potential. It is shown that, for the design of such wings, the slope of the mean camber surface at any point can be determined by a line integration around the wing boundary. By an additional line integration around the wing boundary, this method is extended to include the case where the local section lift coefficient varies with spanwise location (the chordwise loading at every section still remaining uniform). For the uniformly loaded wing of polygonal plan form, the integrations necessary to determine the local slope of the surface and the further integration of the slopes to determine the ordinate can be done analytically. An outline of these integrations and the resulting formulas are included. Calculated results are given for a sweptback wing with uniform chordwise loading and a highly tapered spanwise loading, a uniformly loaded delta wing, a uniformly loaded sweptback wing, and the same sweptback wing with uniform chordwise loading but elliptical span load distribution.

  12. Parallel Nonlinear Aeroelastic Computation for Fighter Wings in the Transonic Region

    NASA Astrophysics Data System (ADS)

    Larsen, Bradley Robert

    In this dissertation, a parallel three-dimensional aeroelastic simulation is applied to current and next generation fighter aircraft wings. The computational model is a nonlinear fluid and structural mesh coupled using the Direct Eulerian-Langrangian method. This method attaches unique local coordinates to each node and connects the fluid mesh to the structure in such a way that a transformation preserved to the global coordinates. This allows the fluid and structure to be updated in the same time step and maintains spatial accuracy at their interface. The structural mesh is modeled using modified nonlinear von Karman finite elements and is discretized using the Galerkin finite element method. The fluid mesh also used the Galerkin finite element method to discretize the unsteady Euler equations. Computational results over a large range of Mach numbers and densities are presented for two candidate fighter wing models for transonic wing tunnel testing. The FX-35 is a trapezoidal wing based on the F-35A, and the F-Wing is a truncated delta wing similar to the F-16. Both wings exhibit a variety of flutter behaviors including strong bending-torsion flutter, limit-cycle oscillations, and essentially single degree-of-freedom responses.

  13. Aeroelastic Analysis of Aircraft: Wing and Wing/Fuselage Configurations

    NASA Technical Reports Server (NTRS)

    Chen, H. H.; Chang, K. C.; Tzong, T.; Cebeci, T.

    1997-01-01

    A previously developed interface method for coupling aerodynamics and structures is used to evaluate the aeroelastic effects for an advanced transport wing at cruise and under-cruise conditions. The calculated results are compared with wind tunnel test data. The capability of the interface method is also investigated for an MD-90 wing/fuselage configuration. In addition, an aircraft trim analysis is described and applied to wing configurations. The accuracy of turbulence models based on the algebraic eddy viscosity formulation of Cebeci and Smith is studied for airfoil flows at low Mach numbers by using methods based on the solutions of the boundary-layer and Navier-Stokes equations.

  14. Pen Branch delta expansion

    SciTech Connect

    Nelson, E.A.; Christensen, E.J.; Mackey, H.E.; Sharitz, R.R.; Jensen, J.R.; Hodgson, M.E.

    1984-02-01

    Since 1954, cooling water discharges from K Reactor ({anti X} = 370 cfs {at} 59 C) to Pen Branch have altered vegetation and deposited sediment in the Savannah River Swamp forming the Pen Branch delta. Currently, the delta covers over 300 acres and continues to expand at a rate of about 16 acres/yr. Examination of delta expansion can provide important information on environmental impacts to wetlands exposed to elevated temperature and flow conditions. To assess the current status and predict future expansion of the Pen Branch delta, historic aerial photographs were analyzed using both basic photo interpretation and computer techniques to provide the following information: (1) past and current expansion rates; (2) location and changes of impacted areas; (3) total acreage presently affected. Delta acreage changes were then compared to historic reactor discharge temperature and flow data to see if expansion rate variations could be related to reactor operations.

  15. Airplane wing vibrations due to atmospheric turbulence

    NASA Technical Reports Server (NTRS)

    Pastel, R. L.; Caruthers, J. E.; Frost, W.

    1981-01-01

    The magnitude of error introduced due to wing vibration when measuring atmospheric turbulence with a wind probe mounted at the wing tip was studied. It was also determined whether accelerometers mounted on the wing tip are needed to correct this error. A spectrum analysis approach is used to determine the error. Estimates of the B-57 wing characteristics are used to simulate the airplane wing, and von Karman's cross spectrum function is used to simulate atmospheric turbulence. It was found that wing vibration introduces large error in measured spectra of turbulence in the frequency's range close to the natural frequencies of the wing.

  16. Supersonic aerodynamic characteristics of a tail-control cruciform maneuverable missile with and without wings

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.

    1983-01-01

    The aerodynamic characteristics for a winged and wingless cruciform missile configuration were examined. The configuration had an ogive-cylinder body with a 3.5 caliber forebody; an overall length-to-diameter ratio of 11.667; and had cruciform tails that were trapezoidal in planform. Tests were made both with and without 72.9 degree cruciform delta wings. The investigation was made for Mach numbers from 1.50 to 4.63, roll attitudes of 0 degrees and 45 degrees, angles of attack from -4 degrees to 22 degrees, and tail control deflections from 10 degrees to -40 degrees.

  17. Computation of wing-vortex interaction in transonic flow using implicit finite difference algorithm

    NASA Technical Reports Server (NTRS)

    Srinivasan, G.; Steger, J. L.

    1981-01-01

    An implicit delta form finite difference algorithm for Euler equations in conservation law form was used in preliminary calculations of three dimensional wing vortex interaction. Both steady and unsteady transonic flow wing vortex interactions are computed. The computations themselves are meant to guide upcoming wind tunnel experiments of the same flow field. Various modifications to the numerical method that are intended to improve computational efficiency are also described and tested in both two and three dimensions. Combination of these methods can reduce the overall computational time by a factor of 4.

  18. An efficient coordinate transformation technique for unsteady, transonic aerodynamic analysis of low aspect-ratio wings

    NASA Technical Reports Server (NTRS)

    Guruswamy, G. P.; Goorjian, P. M.

    1984-01-01

    An efficient coordinate transformation technique is presented for constructing grids for unsteady, transonic aerodynamic computations for delta-type wings. The original shearing transformation yielded computations that were numerically unstable and this paper discusses the sources of those instabilities. The new shearing transformation yields computations that are stable, fast, and accurate. Comparisons of those two methods are shown for the flow over the F5 wing that demonstrate the new stability. Also, comparisons are made with experimental data that demonstrate the accuracy of the new method. The computations were made by using a time-accurate, finite-difference, alternating-direction-implicit (ADI) algorithm for the transonic small-disturbance potential equation.

  19. Wing rotation and lift in SUEX flapping wing mechanisms

    NASA Astrophysics Data System (ADS)

    Mateti, Kiron; Byrne-Dugan, Rory A.; Tadigadapa, Srinivas A.; Rahn, Christopher D.

    2013-01-01

    This research presents detailed modeling and experimental testing of wing rotation and lift in the LionFly, a low cost and mass producible flapping wing mechanism fabricated monolithically from SUEX dry film and powered by piezoelectric bimorph actuators. A flexure hinge along the span of the wing allows the wing to rotate in addition to flapping. A dynamic model including aerodynamics is developed and validated using experimental testing with a laser vibrometer in air and vacuum, stroboscopic photography and high definition image processing, and lift measurement. The 112 mg LionFly produces 46° flap and 44° rotation peak to peak with 12° phase lag, which generates a maximum average lift of 71 μN in response to an applied sinusoidal voltage of 75 V AC and 75 V DC at 37 Hz. Simulated wing trajectories accurately predict measured wing trajectories at small voltage amplitudes, but slightly underpredict amplitude and lift at high voltage amplitudes. By reducing the length of the actuator, reducing the mechanism amplification and tuning the rotational hinge stiffness, a redesigned device is simulated to produce a lift to weight ratio of 1.5.

  20. Delta hepatitis in Malaysia.

    PubMed

    Sinniah, M; Dimitrakakis, M; Tan, D S

    1986-06-01

    Sera from one hundred and fifty nine Malaysian individuals were screened for the prevalence of delta markers. These included 15 HBsAg positive homosexuals, 16 acute hepatitis B cases, 9 chronic hepatitis B patients, 13 healthy HBsAg carriers and 106 intravenous (i.v.) drug abusers, of whom 27 were positive for HBsAg only and the rest were anti-HBc IgG positive but HBsAg negative. The prevalence of delta markers in the homosexuals was found to be 6.7%, in the HBsAg positive drug abusers 17.8%, in acute hepatitis B cases 12.5%. No evidence of delta infection was detected in healthy HBsAg carriers, chronic hepatitis B cases and HBsAg negative i.v. drug abusers. With reference to i.v. drug abusers, the prevalence of delta markers was higher in Malays (23%) than in Chinese (7%) although the latter had a higher HBsAg carrier rate. Although the HBsAg carrier rate in the homosexuals was high, their delta prevalence rate was low as compared to drug abusers. In Malaysia, as in other non-endemic regions, hepatitis delta virus transmission appeared to occur mainly via the parenteral and sexual routes. This is the first time in Malaysia that a reservoir of delta infection has been demonstrated in certain groups of the population at high risk for hepatitis B. PMID:3787309

  1. Design of supercritical swept wings

    NASA Technical Reports Server (NTRS)

    Garabedian, P.; Mcfadden, G.

    1982-01-01

    Computational fluid dynamics are used to discuss problems inherent to transonic three-dimensional flow past supercritical swept wings. The formulation for a boundary value problem for the flow past the wing is provided, including consideration of weak shock waves and the use of parabolic coordinates. A swept wing code is developed which requires a mesh of 152 x 10 x 12 points and 200 time cycles. A formula for wave drag is calculated, based on the idea that the conservation form of the momentum equation becomes an entropy inequality measuring the drag, expressible in terms of a small-disturbance equation for a potential function in two dimensions. The entropy inequality has been incorporated in a two-dimensional code for the analysis of transonic flow over airfoils. A method of artificial viscosity is explored for optimum pressure distributions with design, and involves a free boundary problem considering speed over only a portion of the wing.

  2. Origin Story: Blended Wing Body

    NASA Video Gallery

    NASA is partnering with the Boeing Company, among others, to develop and test the blended wing body aircraft. The BWB has the potential to significantly reduce fuel use and noise. In this video, Bo...

  3. Embedded Wing Propulsion Conceptual Study

    NASA Technical Reports Server (NTRS)

    Kim, Hyun D.; Saunders, John D.

    2003-01-01

    As a part of distributed propulsion work under NASA's Revolutionary Aeropropulsion Concepts or RAC project, a new propulsion-airframe integrated vehicle concept called Embedded Wing Propulsion (EWP) is developed and examined through system and computational fluid dynamics (CFD) studies. The idea behind the concept is to fully integrate a propulsion system within a wing structure so that the aircraft takes full benefits of coupling of wing aerodynamics and the propulsion thrust stream. The objective of this study is to assess the feasibility of the EWP concept applied to large transport aircraft such as the Blended-Wing-Body aircraft. In this paper, some of early analysis and current status of the study are presented. In addition, other current activities of distributed propulsion under the RAC project are briefly discussed.

  4. Delta Scuti stars: Theory

    SciTech Connect

    Guzik, J.A.

    1998-03-01

    The purpose of asteroseismology is not only to derive the internal structure of individual stars from their observed oscillation frequencies, but also to test and extend one`s understanding of the physics of matter under the extremes of temperature, density, and pressure found in stellar interiors. In this review, the author hopes to point out what one can learn about the Sun by studying {delta} Scuti stars, as well as what one can learn about stars more massive or evolved than the Sun. He discusses some of the difficulties in theoretical approaches to asteroseismology for {delta} Scuti stars, using FG Vir, {delta} Scuti, and CD-24{degree} 7599 as examples.

  5. The Delta Clipper dream

    NASA Astrophysics Data System (ADS)

    Furniss, Tim

    1992-04-01

    A conceptual development status evaluation is presented for the SDIO's projected VTOL SSTOV, dubbed the 'Delta Clipper', which is envisioned as an alternative to the slowly developing NASP and the next-generation National Launch System. Delta Clipper program managers believe that the lightweight materials and structures entailed by the requisite empty/gross-weight ratio for an SSTOV are now available, precluding the airbreathing propulsion of such alternatives as HOTOL. The Delta Clipper could operate with a crew of two, or entirely unmanned. The 8-12 LH2/LOX engines employed are derived from the RL-10 engines of the Centaur launcher.

  6. Modeling river delta formation.

    PubMed

    Seybold, Hansjörg; Andrade, José S; Herrmann, Hans J

    2007-10-23

    A model to simulate the time evolution of river delta formation process is presented. It is based on the continuity equation for water and sediment flow and a phenomenological sedimentation/erosion law. Different delta types are reproduced by using different parameters and erosion rules. The structures of the calculated patterns are analyzed in space and time and compared with real data patterns. Furthermore, our model is capable of simulating the rich dynamics related to the switching of the mouth of the river delta. The simulation results are then compared with geological records for the Mississippi River. PMID:17940031

  7. Nile River Delta, Egypt

    NASA Technical Reports Server (NTRS)

    1984-01-01

    The Nile River Delta of Egypt (30.0N, 31.0E) irrigated by the Nile River and its many distributaries, is some of the richest farm land in the world and home to some 45 million people, over half of Egypt's population. The capital city of Cairo is at the apex of the delta. Just across the river from Cairo can be seen the ancient three big pyramids and sphinx at Giza and the Suez Canal is just to the right of the delta.

  8. Nile Delta, Egypt

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The Nile Delta of Egypt (30.0N, 31.0E) irrigated by the Nile River and its many distributaries, is some of the richest farm land in the world and home to some 45 million people, over half of Egypt's population of 57 million. The capital city of Cairo is at the apex of the delta in the middle of the scene. Across the river from Cairo can be seen the three big pyramids and sphinx at Giza and the Suez Canal is just to the right of the delta.

  9. Modeling river delta formation

    PubMed Central

    Seybold, Hansjörg; Andrade, José S.; Herrmann, Hans J.

    2007-01-01

    A model to simulate the time evolution of river delta formation process is presented. It is based on the continuity equation for water and sediment flow and a phenomenological sedimentation/erosion law. Different delta types are reproduced by using different parameters and erosion rules. The structures of the calculated patterns are analyzed in space and time and compared with real data patterns. Furthermore, our model is capable of simulating the rich dynamics related to the switching of the mouth of the river delta. The simulation results are then compared with geological records for the Mississippi River. PMID:17940031

  10. The buzzing of flies' wings

    NASA Astrophysics Data System (ADS)

    Easton, Don

    1999-02-01

    The object of this science-fair project was to determine the wing-beat frequency of common house flies from the frequency of the sound produced when they buzz their wings. The data produced a pattern that I have seen many times and felt sure that I knew what was going on. Like many interesting and seemingly simple phenomena, the longer that I look at this one the more intriguing and less explicable it becomes.

  11. Oblique-wing supersonic aircraft

    NASA Technical Reports Server (NTRS)

    Jones, R. T. (Inventor)

    1976-01-01

    An aircraft including a single fuselage having a main wing and a horizontal stabilizer airfoil pivotally attached at their centers to the fuselage is described. The pivotal attachments allow the airfoils to be yawed relative to the fuselage for high speed flight, and to be positioned at right angles with respect to the fuselage during takeoff, landing, and low speed flight. The main wing and the horizontal stabilizer are upwardly curved from their center pivotal connections towards their ends to form curvilinear dihedrals.

  12. Man made deltas.

    PubMed

    Maselli, Vittorio; Trincardi, Fabio

    2013-01-01

    The review of geochronological and historical data documents that the largest southern European deltas formed almost synchronously during two short intervals of enhanced anthropic pressure on landscapes, respectively during the Roman Empire and the Little Ice Age. These growth phases, that occurred under contrasting climatic regimes, were both followed by generalized delta retreat, driven by two markedly different reasons: after the Romans, the fall of the population and new afforestation let soil erosion in river catchments return to natural background levels; since the industrial revolution, instead, flow regulation through river dams overkill a still increasing sediment production in catchment basins. In this second case, furthermore, the effect of a reduced sediment flux to the coasts is amplified by the sinking of modern deltas, due to land subsidence and sea level rise, that hampers delta outbuilding and increases the vulnerability of coastal zone to marine erosion and flooding. PMID:23722597

  13. Man made deltas

    PubMed Central

    Maselli, Vittorio; Trincardi, Fabio

    2013-01-01

    The review of geochronological and historical data documents that the largest southern European deltas formed almost synchronously during two short intervals of enhanced anthropic pressure on landscapes, respectively during the Roman Empire and the Little Ice Age. These growth phases, that occurred under contrasting climatic regimes, were both followed by generalized delta retreat, driven by two markedly different reasons: after the Romans, the fall of the population and new afforestation let soil erosion in river catchments return to natural background levels; since the industrial revolution, instead, flow regulation through river dams overkill a still increasing sediment production in catchment basins. In this second case, furthermore, the effect of a reduced sediment flux to the coasts is amplified by the sinking of modern deltas, due to land subsidence and sea level rise, that hampers delta outbuilding and increases the vulnerability of coastal zone to marine erosion and flooding. PMID:23722597

  14. Federal Funding in the Delta.

    ERIC Educational Resources Information Center

    Reeder, Richard J.; Calhoun, Samuel D.

    2002-01-01

    The Lower Mississippi Delta region, especially the rural Delta, faces many economic challenges. The rural Delta has received much federal aid in basic income support and funding for human resource development, but less for community resource programs, which are important for economic development. Federal aid to the Delta is analyzed in terms of…

  15. Delta method, an empirical drag buildup technique

    NASA Technical Reports Server (NTRS)

    Feagin, R. C.; Morrison, W. D.

    1978-01-01

    An empirical drag correlation technique was developed from analysis of 19 subsonic and supersonic military aircraft and 15 advanced or supercritical airfoil configurations which can be applied in conceptual and advanced aircraft design activities. The Delta Method may be used for estimating the clean wing drag polar for cruise and maneuver conditions up to buffet onset, and to approximately Mach 2.0. This technique incorporates a unique capability of predicting the off-design performance of advanced or supercritical airfoil sections. The buffet onset limit may also be estimated. The method is applicable to wind tunnel models as well as to full scale configurations. This technique has been converted into a computer code for use on the IBM 360 and CDC 7600 computer facilities at NASA AMES. Results obtained using this method to predict known aircraft characteristics are good and agreement can be obtained within a degree of accuracy judged to be sufficient for the initial processes of preliminary design.

  16. 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.

  17. Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

    NASA Technical Reports Server (NTRS)

    Hartshorn, Fletcher

    2011-01-01

    Aerodynamic analysis on a business jet with a wing glove attached to one wing is presented and discussed. If a wing glove is placed over a portion of one wing, there will be asymmetries in the aircraft as well as overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to make sure the wing glove does not have a drastic effect on the aircraft flight characteristics. TRANAIR, a non-linear full potential solver was used to analyze a full aircraft, with and without a glove, at a variety of flight conditions and angles of attack and sideslip. Changes in the aircraft lift, drag and side force, along with roll, pitch and yawing moment are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove and its fairing are discussed. Results show that the glove used here does not present a drastic change in forces and moments on the aircraft, but an added torsional moment around the quarter-chord of the wing may be a cause for some structural concerns.

  18. Man made deltas?

    NASA Astrophysics Data System (ADS)

    Maselli, V.; Trincardi, F.

    2014-12-01

    During the last few millennia, southern European fluvio-deltaic systems have evolved in response to changes in the hydrological cycle, mostly driven by high-frequency climate oscillations and increasing anthropic pressure on natural landscapes. The review of geochronological and historical data documents that the bulk of the four largest northern Mediterranean and Black Sea deltas (Ebro, Rhone, Po and Danube) formed during two short and synchronous intervals during which anthropogenic land cover change was the main driver for enhanced sediment production. These two major growth phases occurred under contrasting climatic regimes and were both followed by generalized delta retreat, supporting the hypothesis of human-driven delta progradation. Delta retreat, in particular, was the consequence of reduced soil erosion for renewed afforestation after the fall of the Roman Empire, and of river dams construction that overkilled the still increasing sediment production in catchment basins since the Industrial Era. In this second case, in particular, the effect of a reduced sediment flux to the coasts is amplified by the sinking of modern deltas, due to land subsidence and sea level rise, that hampers delta outbuilding and increases the vulnerability of coastal zone to marine erosion and flooding.

  19. Dew-driven folding of insect wings

    NASA Astrophysics Data System (ADS)

    Dickerson, Andrew; Beadles, Sam; Clement, Courtney; Hu, David

    2013-11-01

    Small insect wings fold into tacos when exposed to dewfall or fog for extended times. Such shapes are tightly held together and require great force or long evaporation times for the wings to unfold. In this experimental investigation, we use time-lapse and high-speed videography on a mosquito wing exposed to fog to characterize the folding process from a flat wing to a taco. We observe a taco is formed through a series of processes involving wing bending, unbending, and subsequent tight folding of the wing following the sliding of the drop off the wing. We use a simplified 2D model to determine the forces coalescing drops exert on the wing, and present folding-resistant design suggestions for micro-aerial vehicle wings.

  20. Aerodynamic characteristics of scissor-wing geometries

    NASA Technical Reports Server (NTRS)

    Selberg, Bruce P.; Rokhsaz, Kamran; Housh, Clinton S.

    1991-01-01

    A scissor-wing configuration, consisting of two independently sweeping-wing surfaces, is compared with an equivalent fixed-wing geometry baseline over a wide Mach number range. The scissor-wing configuration is shown to have a higher total lift-to-drag ratio than the baseline in the subsonic region primarily due to the slightly higher aspect ratio of the unswept scissor wing. In the transonic region, the scissor wing is shown to have a higher lift-to-drag ratio than the baseline for values of lift coefficient greater than 0.35. It is also shown that, through the use of wing decalage, the lift of the two independent scissor wings can be equalized. In the supersonic regime, the zero lift wave drag of the scissor-wing at maximum sweep is shown to be 50 and 28 percent less than the zero lift wave drag of the baseline at Mach numbers 1.5 and 3.0, respectively. In addition, a pivot-wing configuration is introduced and compared with the scissor wing. The pivot-wing configuration is shown to have a slightly higher total lift-to-drag ratio than the scissor wing in the supersonic region due to the decreased zero lift wave drag of the pivot-wing configuration.

  1. Investigation into the Role of Dragonfly Wing Flexibility During Passive Wing Pitch Reversal

    NASA Astrophysics Data System (ADS)

    Bajwa, Yousaf; Williams, Ventress; Ren, Yan; Dong, Haibo; Flow Simulation Research Group Team

    2013-11-01

    Wing deformation is a characteristic part of flapping wing flight. In dragonflies, a torsion wave can be observed propagating from the tip to the root during stroke reversal. In this paper, we utilize high-speed photogrammetry and 3d surface reconstruction techniques to quantify wing deformation and kinematics of a dragonfly. We then use finite elements in the absolute nodal coordinate formulation to estimate strain energy in the wing during wing pitch reversal. We use this data to analyze the role of wing structure in facilitating wing rotation and bringing about the characteristic torsion wave. The influence of the elastic force in facilitating wing rotation is then compared with inertial and aerodynamic forces as well. A quantitative look into the variation of strain energy within the insect wing during wing rotation could lead to more efficient design of dynamic wing pitching mechanisms. Supported by NSF CBET-1343154.

  2. The aerodynamic effects of wing-wing interaction in flapping insect wings.

    PubMed

    Lehmann, Fritz-Olaf; Sane, Sanjay P; Dickinson, Michael

    2005-08-01

    We employed a dynamically scaled mechanical model of the small fruit fly Drosophila melanogaster (Reynolds number 100-200) to investigate force enhancement due to contralateral wing interactions during stroke reversal (the ;clap-and-fling'). The results suggest that lift enhancement during clap-and-fling requires an angular separation between the two wings of no more than 10-12 degrees . Within the limitations of the robotic apparatus, the clap-and-fling augmented total lift production by up to 17%, but depended strongly on stroke kinematics. The time course of the interaction between the wings was quite complex. For example, wing interaction attenuated total force during the initial part of the wing clap, but slightly enhanced force at the end of the clap phase. We measured two temporally transient peaks of both lift and drag enhancement during the fling phase: a prominent peak during the initial phase of the fling motion, which accounts for most of the benefit in lift production, and a smaller peak of force enhancement at the end fling when the wings started to move apart. A detailed digital particle image velocimetry (DPIV) analysis during clap-and-fling showed that the most obvious effect of the bilateral ;image' wing on flow occurs during the early phase of the fling, due to a strong fluid influx between the wings as they separate. The DPIV analysis revealed, moreover, that circulation induced by a leading edge vortex (LEV) during the early fling phase was smaller than predicted by inviscid two-dimensional analytical models, whereas circulation of LEV nearly matched the predictions of Weis-Fogh's inviscid model at late fling phase. In addition, the presence of the image wing presumably causes subtle modifications in both the wake capture and viscous forces. Collectively, these effects explain some of the changes in total force and lift production during the fling. Quite surprisingly, the effect of clap-and-fling is not restricted to the dorsal part of the

  3. Supersonic aerodynamic characteristics of a tail-control cruciform maneuverable missile with and without wings

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.; Fournier, R. H.

    1978-01-01

    The aerodynamic characteristics for a winged and a wingless cruciform missile are examined. The body was an ogive-cylinder with a 3.5 caliber forebody; an overall length-to-diameter ratio of 11.667; and has cruciform tails that were trapexoidal in planform. Tests were made both with and without 72.9 deg cruciform delta wings. The investigation was made for Mach numbers from 1.50 to 4.63, roll attitudes of 0 and 45 deg, angles of attack from -40 to 22 deg, and tail control deflections from 10 to -40 deg. The purpose is to determine the influence of the aerodynamic behavior on the design choice for maneuverable missiles intended primarily for air-to-air or surface-to-surface missions. The results indicate that the winged missile with its more linear aerodynamic characteristics and higher lift-curve slope, should provide the highest maneuverability over a large operational range.

  4. Three-dimensional canard-wing shape optimization in aircraft cruise and maneuver environments

    NASA Technical Reports Server (NTRS)

    De Silva, B. M. E.; Carmichael, R. L.

    1978-01-01

    This paper demonstrates a numerical technique for canard-wing shape optimization at two operating conditions. For purposes of simplicity, a mean surface wing paneling code is employed for the aerodynamic calculations. The optimization procedures are based on the method of feasible directions. The shape functions for describing the thickness, camber, and twist are based on polynomial representations. The primary design requirements imposed restrictions on the canard and wing volumes and on the lift coefficients at the operating conditions. Results indicate that significant improvements in minimum drag and lift-to-drag ratio are possible with reasonable aircraft geometries. Calculations were done for supersonic speeds with Mach numbers ranging from 1 to 6. Planforms were mainly of a delta shape with aspect ratio of 1.

  5. Aircraft wing structural detail design (wing, aileron, flaps, and subsystems)

    NASA Technical Reports Server (NTRS)

    Downs, Robert; Zable, Mike; Hughes, James; Heiser, Terry; Adrian, Kenneth

    1993-01-01

    The goal of this project was to design, in detail, the wing, flaps, and ailerons for a primary flight trainer. Integrated in this design are provisions for the fuel system, the electrical system, and the fuselage/cabin carry-through interface structure. This conceptual design displays the general arrangement of all major components in the wing structure, taking into consideration the requirements set forth by the appropriate sections of Federal Aviation Regulation Part 23 (FAR23) as well as those established in the statement of work.

  6. A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

    NASA Technical Reports Server (NTRS)

    Mehrotra, S. C.; Lan, C. E.

    1978-01-01

    The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

  7. A Summary of Numerous Strain-Gage Load Calibrations on Aircraft Wings and Tails in a Technological Format

    NASA Technical Reports Server (NTRS)

    Jenkins, Jerald M.; DeAngelis, V. Michael

    1997-01-01

    Fifteen aircraft structures that were calibrated for flight loads using strain gages are examined. The primary purpose of this paper is to document important examples of load calibrations on airplanes during the past four decades. The emphasis is placed on studying the physical procedures of calibrating strain-gaged structures and all the supporting analyses and computational techniques that have been used. The results and experiences obtained from actual data from 14 structures (on 13 airplanes and 1 laboratory test structure) are presented. This group of structures includes fins, tails, and wings with a wide variety of aspect ratios. Straight- wing, swept-wing, and delta-wing configurations are studied. Some of the structures have skin-dominant construction; others are spar-dominant. Anisotropic materials, heat shields, corrugated components, nonorthogonal primary structures, and truss-type structures are particular characteristics that are included.

  8. Subsonic aerodynamic and flutter characteristics of several wings calculated by the SOUSSA P1.1 panel method

    NASA Technical Reports Server (NTRS)

    Yates, E. C., Jr.; Cunningham, H. J.; Desmarais, R. N.; Silva, W. A.; Drobenko, B.

    1982-01-01

    Several applications of the steady, oscillatory, and unsteady subsonic and supersonic aerodynamics (SOUSSA) computer program to wings with steady and oscillatory motion, including flutter, are discussed. The program employs a generalized Green's function to the full, time-dependent potential-flow equation to obtain an integral equation for the velocity potential at any point in a flow, even points on a body or whole bodies in a flow. Aerodynamic calculations are provided for two rectangular wings, a clipped-tip delta wing, and two swept wings with and without a fuselage. The number and distribution of the finite element panels are varied in order to demonstrate the convergence of the results. The results are shown to be close to those of lifting-surface theory, and further applications with bodies having deformities, arbitrary shapes, motions, and deformations are indicated.

  9. Wind Tunnel Pressure Distribution Tests on a Series of Biplane Wing Models Part III Effects of Charges in Various Combinations of Stagger, Gap, Sweepback, and Decalage

    NASA Technical Reports Server (NTRS)

    Knight, Montgomery; Noyes, Richard W

    1929-01-01

    A concept for the calculation of the vortex lift of sharp-edge delta wings is presented and compared with experimental data. The concept is based on an analogy between the vortex lift and the leading-edge suction associated with the potential flow about the leading edge. This concept, when combined with potential-flow theory modified to include the nonlinearities associated with the exact boundary condition and the loss of the lift component of the leading-edge suction, provides excellent prediction of the total lift for a wide range of delta wings up to angles of attack of 20 degrees or greater.

  10. The Nichols Wing Cutting Equipment

    NASA Technical Reports Server (NTRS)

    Ford, James B

    1923-01-01

    Described here is wing cutting equipment for the economical production of metal wings for wind tunnel models. The machine will make any size of constant-section wing or strut up to one-sixth inch chord by 36-inch span and up to a thickness of one and one-quarter inches. It cuts a smooth, true model that is accurate to within two-thousandths of an inch on any ordinate. The holding jaws are so designed as to leave the model free of chip marks, and the only hand finishing necessary after the cutting is a rub with amunite to remove burrs. The actual change on ordinate in this finishing rub is less than .0002 inches.

  11. Aircraft wing structure detail design

    NASA Technical Reports Server (NTRS)

    Sager, Garrett L.; Roberts, Ron; Mallon, Bob; Alameri, Mohamed; Steinbach, Bill

    1993-01-01

    The provisions of this project call for the design of the structure of the wing and carry-through structure for the Viper primary trainer, which is to be certified as a utility category trainer under FAR part 23. The specific items to be designed in this statement of work were Front Spar, Rear Spar, Aileron Structure, Wing Skin, and Fuselage Carry-through Structure. In the design of these parts, provisions for the fuel system, electrical system, and control routing were required. Also, the total weight of the entire wing planform could not exceed 216 lbs. Since this aircraft is to be used as a primary trainer, and the SOW requires a useful life of 107 cycles, it was decided that all of the principle stresses in the structural members would be kept below 10 ksi. The only drawback to this approach is a weight penalty.

  12. Drosophila Cyclin G Is a Regulator of the Notch Signalling Pathway during Wing Development

    PubMed Central

    Nagel, Anja C.; Szawinski, Jutta; Zimmermann, Mirjam; Preiss, Anette

    2016-01-01

    Notch signalling regulates a multitude of differentiation processes during Drosophila development. For example, Notch activity is required for proper wing vein differentiation which is hampered in mutants of either the receptor Notch, the ligand Delta or the antagonist Hairless. Moreover, the Notch pathway is involved in several aspects of Drosophila oogenesis as well. We have identified Drosophila Cyclin G (CycG) as a molecular interaction partner of Hairless, the major antagonist in the Notch signalling pathway, in vitro and in vivo. Loss of CycG was shown before to cause female sterility and to disturb the architecture of the egg shell. Nevertheless, Notch dependent processes during oogenesis appeared largely unaffected in cycG mutant egg chambers. Loss of CycG modified the dominant wing phenotypes of Notch, Delta and Hairless mutants. Whereas the Notch loss of function phenotype was ameliorated by a loss of CycG, the phenotypes of either Notch gain of function or of Delta or Hairless loss of function were enhanced. In contrast, loss of CycG had only a minor effect on the wing vein phenotype of mutants affecting the EGFR signalling pathway emphasizing the specificity of the interaction of CycG and Notch pathway members. PMID:26963612

  13. Aerodynamic control with passively pitching wings

    NASA Astrophysics Data System (ADS)

    Gravish, Nick; Wood, Robert

    Flapping wings may pitch passively under aerodynamic and inertial loads. Such passive pitching is observed in flapping wing insect and robot flight. The effect of passive wing pitch on the control dynamics of flapping wing flight are unexplored. Here we demonstrate in simulation and experiment the critical role wing pitching plays in yaw control of a flapping wing robot. We study yaw torque generation by a flapping wing allowed to passively rotate in the pitch axis through a rotational spring. Yaw torque is generated through alternating fast and slow upstroke and and downstroke. Yaw torque sensitively depends on both the rotational spring force law and spring stiffness, and at a critical spring stiffness a bifurcation in the yaw torque control relationship occurs. Simulation and experiment reveal the dynamics of this bifurcation and demonstrate that anomalous yaw torque from passively pitching wings is the result of aerodynamic and inertial coupling between the pitching and stroke-plane dynamics.

  14. Results of experiments with slotted wings

    NASA Technical Reports Server (NTRS)

    Lachmann, G

    1924-01-01

    This report recounts some of the successful experiments conducted in England and Germany on slotted wings and ailerons. Wind tunnel test results are given and examples of aircraft constructed with these new wing components are described.

  15. Flexible-Wing-Based Micro Air Vehicles

    NASA Technical Reports Server (NTRS)

    Ifju, Peter G.; Jenkins, David A.; Ettinger, Scott; Lian, Yong-Sheng; Shyy, Wei; Waszak, Martin R.

    2002-01-01

    This paper documents the development and evaluation of an original flexible-wing-based Micro Air Vehicle (MAV) technology that reduces adverse effects of gusty wind conditions and unsteady aerodynamics, exhibits desirable flight stability, and enhances structural durability. The flexible wing concept has been demonstrated on aircraft with wingspans ranging from 18 inches to 5 inches. Salient features of the flexible-wing-based MAV, including the vehicle concept, flexible wing design, novel fabrication methods, aerodynamic assessment, and flight data analysis are presented.

  16. Free wing assembly for an aircraft

    NASA Technical Reports Server (NTRS)

    Wolowicz, C. H. (Inventor)

    1978-01-01

    A free wing is attached to a fuselage of an aircraft in a manner such that the wing is free to pivot about a spanwise axis forward of its aerodynamic center. The wing is angularly displaced about the axis by aerodynamic pitching moments, resulting from lift, and is trimmed through a use of a trimmable free stabilizer comprising a floating canard mounted on a strut rigidly connected to the wing and projected forward from it.

  17. Delta II Mars Pathfinder

    NASA Technical Reports Server (NTRS)

    1998-01-01

    Final preparations for lift off of the DELTA II Mars Pathfinder Rocket are shown. Activities include loading the liquid oxygen, completing the construction of the Rover, and placing the Rover into the Lander. After the countdown, important visual events include the launch of the Delta Rocket, burnout and separation of the three Solid Rocket Boosters, and the main engine cutoff. The cutoff of the main engine marks the beginning of the second stage engine. After the completion of the second stage, the third stage engine ignites and then cuts off. Once the third stage engine cuts off spacecraft separation occurs.

  18. X-Wing Research Vehicle

    NASA Technical Reports Server (NTRS)

    1986-01-01

    One of the most unusual experimental flight vehicles appearing at NASA's Ames-Dryden Flight Research Facility (later redesignated Dryden Flight Research Center) in the 1980s was the Rotor Systems Research Aircraft (RSRA) X-Wing aircraft, seen here on the ramp. The craft was developed originally and then modified by Sikorsky Aircraft for a joint NASA-Defense Advanced Research Projects Agency (DARPA) program and was rolled out 19 August 1986. Taxi tests and initial low-altitude flight tests without the main rotor attached were carried out at Dryden before the program was terminated in 1988. The unusual aircraft that resulted from the Ames Research Center/Army X-Wing Project was flown at the Ames-Dryden Flight Research Facility (now Dryden Flight Research Center), Edwards, California, beginning in the spring of 1984, with a follow-on program beginning in 1986. The program, was conceived to provide an efficient combination of the vertical lift characteristic of conventional helicopters and the high cruise speed of fixed-wing aircraft. It consisted of a hybrid vehicle called the NASA/Army Rotor Systems Research Aircraft (RSRA), which was equipped with advanced X-wing rotor systems. The program began in the early 1970s to investigate ways to increase the speed of rotor aircraft, as well as their performance, reliability, and safety . It also sought to reduce the noise, vibration, and maintenance costs of helicopters. Sikorsky Aircraft Division of United Technologies Laboratories built two RSRA aircraft. NASA's Langley Research Center, Hampton, Virginia, did some initial testing and transferred the program to Ames Research Center, Mountain View, California, for an extensive flight research program conducted by Ames and the Army. The purpose of the 1984 tests was to demonstrate the fixed-wing capability of the helicopter/airplane hybrid research vehicle and explore its flight envelope and flying qualities. These tests, flown by Ames pilot G. Warren Hall and Army Maj (soon

  19. [Winged scapula in lyme borreliosis].

    PubMed

    Rausch, V; Königshausen, M; Gessmann, J; Schildhauer, T A; Seybold, D

    2016-06-01

    Here we present the case of a young patient with one-sided winged scapula and lyme borreliosis. This disease can be very delimitating in daily life. If non-operative treatment fails, dynamic or static stabilization of the scapula can be a therapeutic option. PMID:26849378

  20. The Wings for Angels Project

    ERIC Educational Resources Information Center

    McMillan, Liberty; McMillan, Ellen; Ayers, Ann

    2012-01-01

    How can the spirits of critically ill children be raised? Alexis Weisel (co-president of the Monarch High School National Art Honor Society, 2010-2011) had this question in mind when she initiated and developed the Wings for Angels Project after hearing about the Believe in Tomorrow (BIT) organization through her art teacher, Ellen McMillan. The…

  1. On Wings: Aerodynamics of Eagles.

    ERIC Educational Resources Information Center

    Millson, David

    2000-01-01

    The Aerodynamics Wing Curriculum is a high school program that combines basic physics, aerodynamics, pre-engineering, 3D visualization, computer-assisted drafting, computer-assisted manufacturing, production, reengineering, and success in a 15-hour, 3-week classroom module. (JOW)

  2. Rotary wing aerodynamically generated noise

    NASA Technical Reports Server (NTRS)

    Schmitz, F. J.; Morse, H. A.

    1982-01-01

    The history and methodology of aerodynamic noise reduction in rotary wing aircraft are presented. Thickness noise during hover tests and blade vortex interaction noise are determined and predicted through the use of a variety of computer codes. The use of test facilities and scale models for data acquisition are discussed.

  3. Aerodynamic characteristics of three slender sharp-edge 74 degrees swept wings at subsonic, transonic, and supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Davenport, E. E.

    1974-01-01

    Slender sharp-edge wings having leading-edge sweep angles of 74 deg have been studied at Mach numbers from 0.60 to 2.80, at angles of attack from about minus 4 deg to 22 deg, and at angles of sideslip from 0 deg to 5 deg. The wings had delta, arrow, and diamond planforms. The experimental tests were made in the Langley 8-foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel test section number 1. The theoretical predictions were made using the theories of NASA TN D-3767 and NASA TN D-6243. The results of the study indicated that the lift and drag characteristics as affected by planform and Mach number could be reasonably well predicted for the delta wing in the subsonic and transonic Mach number range. In the supersonic range, the delta and diamond wings were about equally good in the degree of agreement between experiment and theory. In making drag-due-to-lift predictions the vortex lift effects must be taken into account if reasonable results are to be obtained at moderate or high lift coefficients.

  4. Lessons from KIPP Delta

    ERIC Educational Resources Information Center

    Maranto, Robert; Shuls, James V.

    2011-01-01

    KIPP Delta succeeds at its stated mission, probably because of its careful attention to culture building. What distinguishes this KIPP school is thoughtful work linking the daily processes of schooling to the goals of schooling, in this case success in college. Day to day tactics reflect broader themes: having a clear mission and hiring staff who…

  5. Delta Airlines LOFT training

    NASA Technical Reports Server (NTRS)

    Whitehead, J.

    1981-01-01

    A LOFT program was developed as part of the DC-9 training program which serves as a prototype for much of Delta's other aircraft training programs. The LOFT used differs little from the ideology presented in the Advisory Circular. Difficulty and experienced concerns regarding the effectiveness of LOFT as a complete training vehicle are explored.

  6. DELTA PHASE PLUTONIUM ALLOYS

    DOEpatents

    Cramer, E.M.; Ellinger, F.H.; Land. C.C.

    1960-03-22

    Delta-phase plutonium alloys were developed suitable for use as reactor fuels. The alloys consist of from 1 to 4 at.% zinc and the balance plutonium. The alloys have good neutronic, corrosion, and fabrication characteristics snd possess good dimensional characteristics throughout an operating temperature range from 300 to 490 deg C.

  7. Using delta-front bathymetry to understand river delta progradation

    NASA Astrophysics Data System (ADS)

    Shaw, J. B.; Mohrig, D. C.

    2010-12-01

    We investigate the delta-front bathymetry of the Wax Lake Delta in Louisiana, USA; a sand rich river delta prograding quickly (~100 m/yr) into a shallow (~2.5 m) basin. The delta-front is the zone separating the bottomset from the topset of the delta. Bottomset sedimentation covers the bed evenly whereas topset sediment transport is focused by flow through distributary channels. The delta front connects these two disparate transport regimes and has a profound effect on channel-network evolution and sedimentary structure of river deltas. Predictions of delta-front topography made by models of delta progradation have rarely been compared to the bathymetry of field-scale deltas. We have mapped 60 km2 of delta front bathymetry immediately seaward of two sub-aerial distributary channels. Subaqueous channels extend up to 2 km seaward of their subaerial portions. These channels lose definition at their distal ends through a combination of channel-bed shoaling and loss of bank relief. Little bathymetric relief is observed at the fronts of the subaqueous channels, calling into question the role of channel-mouth bars in generating the bifurcations observed in this delta-channel network. Near the subaerial to subaqueous transition, steep and eroding sidewalls transition to constructional banks with gentle grades. Grab samples of bed material have been collected throughout the study area in order to detect proximal to distal fining and to constrain the shear stresses connected with delta-front sedimentation. A better understanding of sediment transport in the delta front and its affiliated patterns of erosion and deposition is essential for progress in understanding how river deltas prograde and fill their basins.

  8. Polarization-Controllable Winged Nanocone Tip Antenna

    NASA Astrophysics Data System (ADS)

    Huttunen, Mikko J.; Mäkitalo, Jouni; Kauranen, Martti

    We propose winged metal nanocone optical antennas for efficient coupling of far-field radiation into the near-field of sharp metal tips. Unlike normal sharp metal tips, the winged nanocones require no oscillating field along the tip axis for the excitation of the tips. We calculate extinction spectra and local-field enhancements for single and winged gold nanocones and show that the field enhancements in the tips of the winged cones are due to a combination of particle plasmon resonances and a lightning-rod effect. We also propose that the winged nanocones could be used for optical far-field background suppression for tip-enhanced microscopy.

  9. F-8 oblique wing structural feasibility study

    NASA Technical Reports Server (NTRS)

    Koltko, E.; Katz, A.; Bell, M. A.; Smith, W. D.; Lauridia, R.; Overstreet, C. T.; Klapprott, C.; Orr, T. F.; Jobe, C. L.; Wyatt, F. G.

    1975-01-01

    The feasibility of fitting a rotating oblique wing on an F-8 aircraft to produce a full scale manned prototype capable of operating in the transonic and supersonic speed range was investigated. The strength, aeroelasticity, and fatigue life of such a prototype are analyzed. Concepts are developed for a new wing, a pivot, a skewing mechanism, control systems that operate through the pivot, and a wing support assembly that attaches in the F-8 wing cavity. The modification of the two-place NTF-8A aircraft to the oblique wing configuration is discussed.

  10. Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

    NASA Technical Reports Server (NTRS)

    Hartshorn, Fletcher

    2011-01-01

    Aerodynamic computational fluid dynamics analysis of a wing glove attached to one wing of a business jet is presented and discussed. A wing glove placed on only one wing will produce asymmetric aerodynamic effects that will result in overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to ensure that the wing glove does not have a significant effect on the aircraft flight characteristics. TRANAIR (Calmar Research Corporation, Cato, New York), a nonlinear full potential solver, and Star-CCM+ (CD-adapco, Melville, New York), a finite volume full Reynolds-averaged Navier-Stokes computational fluid dynamics solver, are used to analyze a full aircraft with and without the glove at a variety of flight conditions, aircraft configurations, and angles of attack and sideslip. Changes in the aircraft lift, drag, and side force along with roll, pitch, and yaw are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove are discussed. Results show that the glove produces only small changes in the aerodynamic forces and moments acting on the aircraft, most of which are insignificant.

  11. The Devil's in the Delta

    ERIC Educational Resources Information Center

    Luyben, William L.

    2007-01-01

    Students frequently confuse and incorrectly apply the several "deltas" that are used in chemical engineering. The deltas come in three different flavors: "out minus in", "big minus little" and "now versus then." The first applies to a change in a stream property as the stream flows through a process. For example, the "[delta]H" in an energy…

  12. Rotor/wing aerodynamic interactions in hover

    NASA Technical Reports Server (NTRS)

    Felker, F. F.; Light, J. S.

    1986-01-01

    An experimental and theoretical investigation of rotor/wing aerodynamic interactions in hover is described. The experimental investigation consisted of both a large-scale and small-scale test. A 0.658-scale, V-22 rotor and wing was used in the large-scale test. Wind download, wing surface pressure, rotor performance, and rotor downwash data from the large-scale test are presented. A small-scale experiment was conducted to determine how changes in the rotor/wing geometry affected the aerodynamic interactions. These geometry variations included the distance between the rotor and wing, wing incidence angle, and configurations both with the rotor axis at the tip of the wing (tilt rotor configuration) and with the rotor axis at the center of the wing (compound helicopter configuration). A wing with boundary-layer control was also tested to evaluate the effect of leading and trailing edge upper surface blowing on the wing download. A computationally efficient, semi-empirical theory was developed to predict the download on the wing. Finally, correlations between the theoretical predictions and test data are presented.

  13. AST Composite Wing Program: Executive Summary

    NASA Technical Reports Server (NTRS)

    Karal, Michael

    2001-01-01

    The Boeing Company demonstrated the application of stitched/resin infused (S/RFI) composite materials on commercial transport aircraft primary wing structures under the Advanced Subsonic technology (AST) Composite Wing contract. This report describes a weight trade study utilizing a wing torque box design applicable to a 220-passenger commercial aircraft and was used to verify the weight savings a S/RFI structure would offer compared to an identical aluminum wing box design. This trade study was performed in the AST Composite Wing program, and the overall weight savings are reported. Previous program work involved the design of a S/RFI-base-line wing box structural test component and its associated testing hardware. This detail structural design effort which is known as the "semi-span" in this report, was completed under a previous NASA contract. The full-scale wing design was based on a configuration for a MD-90-40X airplane, and the objective of this structural test component was to demonstrate the maturity of the S/RFI technology through the evaluation of a full-scale wing box/fuselage section structural test. However, scope reductions of the AST Composite Wing Program pre-vented the fabrication and evaluation of this wing box structure. Results obtained from the weight trade study, the full-scale test component design effort, fabrication, design development testing, and full-scale testing of the semi-span wing box are reported.

  14. Piezoelectrically actuated insect scale flapping wing

    NASA Astrophysics Data System (ADS)

    Mukherjee, Sujoy; Ganguli, Ranjan

    2010-04-01

    An energy method is used in order to derive the non-linear equations of motion of a smart flapping wing. Flapping wing is actuated from the root by a PZT unimorph in the piezofan configuration. Dynamic characteristics of the wing, having the same size as dragonfly Aeshna Multicolor, are analyzed using numerical simulations. It is shown that flapping angle variations of the smart flapping wing are similar to the actual dragonfly wing for a specific feasible voltage. An unsteady aerodynamic model based on modified strip theory is used to obtain the aerodynamic forces. It is found that the smart wing generates sufficient lift to support its own weight and carry a small payload. It is therefore a potential candidate for flapping wing of micro air vehicles.

  15. Locomotion by Tandem and Parallel Wings

    NASA Astrophysics Data System (ADS)

    Tanida, Yoshimichi

    A two-dimensional analysis was carried out on the locomotion by tandem and parallel wings in relation to the free flight of dragonflies and beetles, remarking the mutual interference between fore and hind wings. The results obtained are summarized as follows: In the case of tandem wings, (1)High thrust and propulsive efficiency can be achieved when the forewing oscillates with a definite phase lag behind the hindwing, as in the case of real dragonflies, (2)Somewhat smaller amplitude of hindwing leads to optimum condition for work sharing of two wings, (3)The hard forewing does not serve for the thrust and propulsive efficiency, whereas the hard hindwing does for the augmentation of them; In the case of parallel wings, (4)The hard wing placed near the soft wing acts nearly as an infinite plate, as for the ground effect, increasing both thrust and propulsive efficiency.

  16. From wakes to wings: using a multi-fidelity approach to design flapping wings

    NASA Astrophysics Data System (ADS)

    Willis, David J.; Salehipour, Hesam

    2014-06-01

    We present two potential flow based computational tools for the design and analysis of efficient, low Reynolds number flapping wings. Our approach starts with a series of wake-only momentum and energetics analyses. We have used and extended the classical wake-only approach to efficiently perform a large number of computations over the flapping parameter space. The method considers the balance of flight forces in the coupled prediction of wing flapping kinematics and flight energetics. Following the wake-only energetics analysis, a quasi-inverse doublet lattice method (qi-DLM) is applied to determine flapping wing shape including localized wing morphing and deformation. This local wing morphing prescribed so that the wing may achieve the desired, minimum power wake vorticity distribution defined by the wake-only analysis. In this paper we illustrate these methods and perform a preliminary study to assess the impact of wing taper, wing camber and wing twist variations on efficient flapping flight.

  17. DELTAS: A new Global Delta Sustainability Initiative (Invited)

    NASA Astrophysics Data System (ADS)

    Foufoula-Georgiou, E.

    2013-12-01

    Deltas are economic and environmental hotspots, food baskets for many nations, home to a large part of the world population, and hosts of exceptional biodiversity and rich ecosystems. Deltas, being at the land-water interface, are international, regional, and local transport hubs, thus providing the basis for intense economic activities. Yet, deltas are deteriorating at an alarming rate as 'victims' of human actions (e.g. water and sediment reduction due to upstream basin development), climatic impacts (e.g. sea level rise and flooding from rivers and intense tropical storms), and local exploration (e.g. sand or aggregates, groundwater and hydrocarbon extraction). Although many efforts exist on individual deltas around the world, a comprehensive global delta sustainability initiative that promotes awareness, science integration, data and knowledge sharing, and development of decision support tools for an effective dialogue between scientists, managers and policy makers is lacking. Recently, the international scientific community proposed to establish the International Year of Deltas (IYD) to serve as the beginning of such a Global Delta Sustainability Initiative. The IYD was proposed as a year to: (1) increase awareness and attention to the value and vulnerability of deltas worldwide; (2) promote and enhance international and regional cooperation at the scientific, policy, and stakeholder level; and (3) serve as a launching pad for a 10-year committed effort to understand deltas as complex socio-ecological systems and ensure preparedness in protecting and restoring them in a rapidly changing environment. In this talk, the vision for such an international coordinated effort on delta sustainability will be presented as developed by a large number of international experts and recently funded through the Belmont Forum International Opportunities Fund. Participating countries include: U.S., France, Germany, U.K., India, Japan, Netherlands, Norway, Brazil, Bangladesh

  18. Effect of wing inertia on hovering performance of flexible flapping wings

    NASA Astrophysics Data System (ADS)

    Yin, Bo; Luo, Haoxiang

    2010-11-01

    Insect wings in flight typically deform under the combined aerodynamic force and wing inertia; whichever is dominant depends on the mass ratio defined as m∗=ρsh/(ρfc), where ρsh is the surface density of the wing, ρf is the density of the air, and c is the characteristic length of the wing. To study the differences that the wing inertia makes in the aerodynamic performance of the deformable wing, a two-dimensional numerical study is applied to simulate the flow-structure interaction of a flapping wing during hovering flight. The wing section is modeled as an elastic plate, which may experience nonlinear deformations while flapping. The effect of the wing inertia on lift production, drag resistance, and power consumption is studied for a range of wing rigidity. It is found that both inertia-induced deformation and flow-induced deformation can enhance lift of the wing. However, the flow-induced deformation, which corresponds to the low-mass wing, produces less drag and leads to higher aerodynamic power efficiency. In addition, the wing deformation has a significant effect on the unsteady vortices around the wing. The implication of the findings on insect flight is discussed.

  19. Effects of Wing Platform on the Aerodynamic Performance of Finite-Span Flapping Wings

    NASA Astrophysics Data System (ADS)

    Yu, Meilin; Wang, Z. J.; Hu, Hui

    2010-11-01

    A numerical study is conducted to investigate the effects of wing platform on the aerodynamics performance of finite-span flapping wings. A three-dimensional high-order Navier-Stokes compressible flow solver was developed using the spectral difference method and dynamic grids. An AUSM^+-up Riemann solver was implemented to simulate the unsteady low Mach number flows over finite-span flapping wings with explicit third order Runge-Kutta time integration. The studied finite-span flapping wings, which include a rectangular flapping wing, an elliptic flapping wing and a bio-inspired flapping wing, have the same wing span, aspect ratio of the platform and the characteristics of the flapping motion (i.e., sinusoidal trajectory of the flapping wing tip, Strouhal number and reduced frequency). In the present study, the Strouhul number (Str) of the finite-span flapping wings was selected to be well within the optimal range usually used by flying insects and birds and swimming fishes (i.e., 0.2 < Str < 0.4). The effects of the wing platform on the aerodynamics performance of the finite-span flapping wings were elucidated in the terms of the evolutions and dynamic interaction between the leading edge vortices (LEV) and the wing tip vortices as well as the resultant aerodynamic forces (both lift and thrust) generated by the flapping wings.

  20. Wing spar stress charts and wing truss proportions

    NASA Technical Reports Server (NTRS)

    Warner, Edward P

    1926-01-01

    In order to simplify the calculation of beams continuous over three supports, a series of charts have been calculated giving the bending moments at all the critical points and the reactions at all supports for such members. Using these charts as a basis, calculations of equivalent bending moments, representing the total stresses acting in two bay-wing trusses of proportions varying over a wide range, have been determined, both with and without allowance for column effect. This leads finally to the determination of the best proportions for any particular truss or the best strut locations in any particular airplane. The ideal proportions are found to vary with the thickness of the wing section used, the aspect ratio, and the ratio of gap to chord.

  1. Computation of laminar mixed convection flow in a rectangular duct with wing-type built-in obstacles

    NASA Astrophysics Data System (ADS)

    Biswas, G.; Mitra, N. K.; Fiebig, M.

    1988-06-01

    The object of this investigation is numerical computation of laminar mixed convection flows and heat transfer in a rectangular channel whose bottom plate has been punched out in form of a delta wing. The geometrical configuration represents proposed model of a part of gas-liquid, fin-pipe cross flow heat exchanger (gas side). Enhancement of heat transfer between the gas and the channel wall is evidenced. This augmentation is primarily due to the longitudinal vortices generated by the wing. Mixed convection condition is characterized by buoyancy induced secondary flows which increase the vortex strength and improve the heat transfer still further.

  2. Martian deltas: Morphology and distribution

    NASA Technical Reports Server (NTRS)

    Rice, J. W., Jr.; Scott, D. H.

    1993-01-01

    Recent detailed mapping has revealed numerous examples of Martian deltas. The location and morphology of these deltas are described. Factors that contribute to delta morphology are river regime, coastal processes, structural stability, and climate. The largest delta systems on Mars are located near the mouths of Maja, Maumee, Vedra, Ma'adim, Kasei, and Brazos Valles. There are also several smaller-scale deltas emplaced near channel mouths situated in Ismenius Lacus, Memnonia, and Arabia. Delta morphology was used to reconstruct type, quantity, and sediment load size transported by the debouching channel systems. Methods initially developed for terrestrial systems were used to gain information on the relationships between Martian delta morphology, river regime, and coastal processes.

  3. Understanding pesticides in California's Delta

    USGS Publications Warehouse

    Kuivila, Kathryn M.; Orlando, James L.

    2012-01-01

    The Sacramento-San Joaquin River Delta (Delta) is the hub of California’s water system and also an important habitat for imperiled fish and wildlife. Aquatic organisms are exposed to mixtures of pesticides that flow through the maze of Delta water channels from sources including agricultural, landscape, and urban pest-control applications. While we do not know all of the effects pesticides have on the ecosystem, there is evidence that they cause some damage to organisms in the Delta. Decades of USGS research have provided a good understanding of when, where, and how pesticides enter the Delta. However, pesticide use is continually changing. New field studies and methods are needed so that scientists can analyze which pesticides are present in the Delta, and at what concentrations, enabling them to estimate exposure and ultimate effects on organisms. Continuing research will provide resource managers and stakeholders with crucial information to manage the Delta wisely.

  4. Evaluation of Blended Wing-Body Combinations with Curved Plan Forms at Mach Numbers Up to 3.50

    NASA Technical Reports Server (NTRS)

    Holdaway, George H.; Mellenthin, Jack A.

    1960-01-01

    This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge.

  5. LS Pegasi: A Low-Inclination SW Sextantis-Type Cataclysmic Binary with High-Velocity Balmer Emission-Line Wings

    NASA Astrophysics Data System (ADS)

    Taylor, Cynthia J.; Thorstensen, John R.; Patterson, Joseph

    1999-02-01

    We present time-resolved spectroscopy and photometry of the bright cataclysmic variable LS Peg (=S193; V~13.0-Szkody et al.). The Balmer lines exhibit broad, asymmetric wings Doppler-shifted by about 2000 km s^-1 at the edges, while the He I lines show phase-dependent absorption features strikingly similar to SW Sextantis stars, as well as emission through most of the phase. The C III/N III emission blend does not show any phase dependence. From velocities of Hα emission lines, we determine an orbital period of 0.174774+/-0.000003 days (=4.1946 hr), which agrees with Szkody's value of approximately 4.2 hr. No stable photometric signal was found at the orbital period. A noncoherent quasi-periodic photometric signal was seen at a period of 20.7+/-0.3 minutes. The high-velocity Balmer wings most probably arise from a stream reimpact point close to the white dwarf. We present simulated spectra based on a kinematic model similar to the modified disk-overflow scenario of Hellier & Robinson. The models reproduce the broad line wings, though some other details are unexplained. Using an estimate of dynamical phase based on the model, we show that the phasing of the emission- and absorption-line variations is consistent with that in (eclipsing) SW Sex stars. We therefore identify LS Peg as a low-inclination SW Sex star. Our model suggests i=30^deg, and the observed absence of any photometric signal at the orbital frequency establishes i<60^deg. This constraint puts a severe strain on interpretations of the SW Sex phenomenon which rely on disk structures lying slightly out of the orbital plane.

  6. Numerical simulation of a powered-lift landing, tracking flow features using overset grids, and simulation of high lift devices on a fighter-lift-and-control wing

    NASA Technical Reports Server (NTRS)

    Chawla, Kalpana

    1993-01-01

    Attached as appendices to this report are documents describing work performed on the simulation of a landing powered-lift delta wing, the tracking of flow features using overset grids, and the simulation of flaps on the Wright Patterson Lab's fighter-lift-and-control (FLAC) wing. Numerical simulation of a powered-lift landing includes the computation of flow about a delta wing at four fixed heights as well as a simulated landing, in which the delta wing descends toward the ground. Comparison of computed and experimental lift coefficients indicates that the simulations capture the qualitative trends in lift-loss encountered by thrust-vectoring aircraft operating in ground effect. Power spectra of temporal variations of pressure indicate computed vortex shedding frequencies close to the jet exit are in the experimentally observed frequency range; the power spectra of pressure also provide insights into the mechanisms of lift oscillations. Also, a method for using overset grids to track dynamic flow features is described and the method is validated by tracking a moving shock and vortices shed behind a circular cylinder. Finally, Chimera gridding strategies were used to develop pressure coefficient contours for the FLAC wing for a Mach no. of 0.18 and Reynolds no. of 2.5 million.

  7. Aerodynamic effects of flexibility in flapping wings

    PubMed Central

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

    2010-01-01

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

  8. Aerodynamic-structural study of canard wing, dual wing, and conventional wing systems for general aviation applications

    NASA Technical Reports Server (NTRS)

    Selberg, B. P.; Cronin, D. L.

    1985-01-01

    An analytical aerodynamic-structural airplane configuration study was conducted to assess performance gains achievable through advanced design concepts. The mission specification was for 350 mph, range of 1500 st. mi., at altitudes between 30,000 and 40,000 ft. Two payload classes were studied - 1200 lb (6 passengers) and 2400 lb (12 passengers). The configurations analyzed included canard wings, closely coupled dual wings, swept forward - swept rearward wings, joined wings, and conventional wing tail arrangements. The results illustrate substantial performance gains possible with the dual wing configuration. These gains result from weight savings due to predicted structural efficiencies. The need for further studies of structural efficiencies for the various advanced configurations was highlighted.

  9. Wing-wake interaction reduces power consumption in insect tandem wings

    NASA Astrophysics Data System (ADS)

    Lehmann, Fritz-Olaf

    Insects are capable of a remarkable diversity of flight techniques. Dragonflies, in particular, are notable for their powerful aerial manoeuvres and endurance during prey catching or territory flights. While most insects such as flies, bees and wasps either reduced their hinds wings or mechanically coupled fore and hind wings, dragonflies have maintained two independent-controlled pairs of wings throughout their evolution. An extraordinary feature of dragonfly wing kinematics is wing phasing, the shift in flapping phase between the fore and hind wing periods. Wing phasing has previously been associated with an increase in thrust production, readiness for manoeuvrability and hunting performance. Recent studies have shown that wing phasing in tandem wings produces a twofold modulation in hind wing lift, but slightly reduces the maximum combined lift of fore and hind wings, compared to two wings flapping in isolation. Despite this disadvantage, however, wing phasing is effective in improving aerodynamic efficiency during flight by the removal of kinetic energy from the wake. Computational analyses demonstrate that this increase in flight efficiency may save up to 22% aerodynamic power expenditure compared to insects flapping only two wings. In terms of engineering, energetic benefits in four-wing flapping are of substantial interest in the field of biomimetic aircraft design, because the performance of man-made air vehicles is often limited by high-power expenditure rather than by lift production. This manuscript provides a summary on power expenditures and aerodynamic efficiency in flapping tandem wings by investigating wing phasing in a dynamically scaled robotic model of a hovering dragonfly.

  10. Wing-wake interaction reduces power consumption in insect tandem wings

    NASA Astrophysics Data System (ADS)

    Lehmann, Fritz-Olaf

    2009-05-01

    Insects are capable of a remarkable diversity of flight techniques. Dragonflies, in particular, are notable for their powerful aerial manoeuvres and endurance during prey catching or territory flights. While most insects such as flies, bees and wasps either reduced their hinds wings or mechanically coupled fore and hind wings, dragonflies have maintained two independent-controlled pairs of wings throughout their evolution. An extraordinary feature of dragonfly wing kinematics is wing phasing, the shift in flapping phase between the fore and hind wing periods. Wing phasing has previously been associated with an increase in thrust production, readiness for manoeuvrability and hunting performance. Recent studies have shown that wing phasing in tandem wings produces a twofold modulation in hind wing lift, but slightly reduces the maximum combined lift of fore and hind wings, compared to two wings flapping in isolation. Despite this disadvantage, however, wing phasing is effective in improving aerodynamic efficiency during flight by the removal of kinetic energy from the wake. Computational analyses demonstrate that this increase in flight efficiency may save up to 22% aerodynamic power expenditure compared to insects flapping only two wings. In terms of engineering, energetic benefits in four-wing flapping are of substantial interest in the field of biomimetic aircraft design, because the performance of man-made air vehicles is often limited by high-power expenditure rather than by lift production. This manuscript provides a summary on power expenditures and aerodynamic efficiency in flapping tandem wings by investigating wing phasing in a dynamically scaled robotic model of a hovering dragonfly.

  11. Effects of wing deformation on aerodynamic performance of a revolving insect wing

    NASA Astrophysics Data System (ADS)

    Noda, Ryusuke; Nakata, Toshiyuki; Liu, Hao

    2014-12-01

    Flexible wings of insects and bio-inspired micro air vehicles generally deform remarkably during flapping flight owing to aerodynamic and inertial forces, which is of highly nonlinear fluid-structure interaction (FSI) problems. To elucidate the novel mechanisms associated with flexible wing aerodynamics in the low Reynolds number regime, we have built up a FSI model of a hawkmoth wing undergoing revolving and made an investigation on the effects of flexible wing deformation on aerodynamic performance of the revolving wing model. To take into account the characteristics of flapping wing kinematics we designed a kinematic model for the revolving wing in two-fold: acceleration and steady rotation, which are based on hovering wing kinematics of hawkmoth, Manduca sexta. Our results show that both aerodynamic and inertial forces demonstrate a pronounced increase during acceleration phase, which results in a significant wing deformation. While the aerodynamic force turns to reduce after the wing acceleration terminates due to the burst and detachment of leading-edge vortices (LEVs), the dynamic wing deformation seem to delay the burst of LEVs and hence to augment the aerodynamic force during and even after the acceleration. During the phase of steady rotation, the flexible wing model generates more vertical force at higher angles of attack (40°-60°) but less horizontal force than those of a rigid wing model. This is because the wing twist in spanwise owing to aerodynamic forces results in a reduction in the effective angle of attack at wing tip, which leads to enhancing the aerodynamics performance by increasing the vertical force while reducing the horizontal force. Moreover, our results point out the importance of the fluid-structure interaction in evaluating flexible wing aerodynamics: the wing deformation does play a significant role in enhancing the aerodynamic performances but works differently during acceleration and steady rotation, which is mainly induced by

  12. Flexible Wing Model for Structural Sizing and Multidisciplinary Design Optimization of a Strut-Braced Wing

    NASA Technical Reports Server (NTRS)

    Gern, Frank H.; Naghshineh, Amir H.; Sulaeman, Erwin; Kapania, Rakesh K.; Haftka, Raphael T.

    2000-01-01

    This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.

  13. Mississippi River Delta

    NASA Technical Reports Server (NTRS)

    2002-01-01

    As the Mississippi River enters the Gulf of Mexico, it loses energy and dumps its load of sediment that it has carried on its journey through the mid continent. This pile of sediment, or mud, accumulates over the years building up the delta front. As one part of the delta becomes clogged with sediment, the delta front will migrate in search of new areas to grow. The area shown on this image is the currently active delta front of the Mississippi. The migratory nature of the delta forms natural traps for oil. Most of the land in the image consists of mud flats and marsh lands. There is little human settlement in this area due to the instability of the sediments. The main shipping channel of the Mississippi River is the broad stripe running northwest to southeast.

    This image was acquired on May 24, 2001 by the Advanced Spaceborne Thermal Emission and Reflection Radiometer (ASTER) on NASA's Terra satellite. With its 14 spectral bands from the visible to the thermal infrared wavelength region, and its high spatial resolution of 15 to 90 meters (about 50 to 300 feet), ASTER will image Earth for the next 6 years to map and monitor the changing surface of our planet.

    ASTER is one of five Earth-observing instruments launched December 18,1999, on NASA's Terra satellite. The instrument was built by Japan's Ministry of Economy, Trade and Industry. A joint U.S./Japan science team is responsible for validation and calibration of the instrument and the data products. Dr. Anne Kahle at NASA's Jet Propulsion Laboratory, Pasadena, California, is the U.S. Science team leader; Bjorn Eng of JPL is the project manager. The Terra mission is part of NASA's Earth Science Enterprise, a long-term research and technology program designed to examine Earth's land, oceans, atmosphere, ice and life as a total integrated system.

    The broad spectral coverage and high spectral resolution of ASTER will provide scientists in numerous disciplines with critical information for surface mapping

  14. Methodologies for reproducing in-flight loads of aircraft wings on the ground and predicting their response to battle-induced damage

    NASA Astrophysics Data System (ADS)

    Bou-Mosleh, Charbel Fouad

    Survivability of an aircraft in combat is achieved by not getting hit or by withstanding the effects of some suffered hits. Combat damage is described by the removal of one or more portions of the wing or any other flight control surface. To determine whether a wing will survive a specific damage, the structural and aerodynamic response of the wing should be predicted and tested. The response of wings to battle-induced damage is currently addressed through live-fire testing on the ground. The loading methodology used in these live-fire tests does not reproduce the loads encountered during flight, and does not account for the changes in structural stiffness and mass of the wing after damage infliction. In addition, current live-fire tests fail to address the changes in the aerodynamic performance of the wing caused by the battle-induced damage. To better address the structural response of aircraft wings to combat damage, this thesis investigates a concept for an alternative loading methodology that exploits recent advances in nonlinear aeroelastic simulations and smart material actuators. The main idea behind this concept is to accurately predict the stress states of the wing before, during, and after sustaining a hit, for a given flight condition, and reproduce them on the ground by loading the spars and ribs of the wings with programmable actuators and/or a few external tethers. Mathematically, this entails solving an optimization problem to determine the locations and gains of the actuators. Two different types of actuators are investigated: 1D actuators or actuators with tension/compression capability and bimorph bender actuators. The potential of the investigated loading methodology is evaluated for "slender" wings (ARW-2 wing) and for "delta" wings (HSCT and F-16 wing) at a transonic flight condition. The obtained numerical results suggest that the investigated loading methodology can reproduce a desired stress state fairly accurately using external tethers

  15. Rotor/Wing Interactions in Hover

    NASA Technical Reports Server (NTRS)

    Young, Larry A.; Derby, Michael R.

    2002-01-01

    Hover predictions of tiltrotor aircraft are hampered by the lack of accurate and computationally efficient models for rotor/wing interactional aerodynamics. This paper summarizes the development of an approximate, potential flow solution for the rotor-on-rotor and wing-on-rotor interactions. This analysis is based on actuator disk and vortex theory and the method of images. The analysis is applicable for out-of-ground-effect predictions. The analysis is particularly suited for aircraft preliminary design studies. Flow field predictions from this simple analytical model are validated against experimental data from previous studies. The paper concludes with an analytical assessment of the influence of rotor-on-rotor and wing-on-rotor interactions. This assessment examines the effect of rotor-to-wing offset distance, wing sweep, wing span, and flaperon incidence angle on tiltrotor inflow and performance.

  16. Constraints on the wing morphology of pterosaurs

    PubMed Central

    Palmer, Colin; Dyke, Gareth

    2012-01-01

    Animals that fly must be able to do so over a huge range of aerodynamic conditions, determined by weather, wind speed and the nature of their environment. No single parameter can be used to determine—let alone measure—optimum flight performance as it relates to wing shape. Reconstructing the wings of the extinct pterosaurs has therefore proved especially problematic: these Mesozoic flying reptiles had a soft-tissue membranous flight surface that is rarely preserved in the fossil record. Here, we review basic mechanical and aerodynamic constraints that influenced the wing shape of pterosaurs, and, building on this, present a series of theoretical modelling results. These results allow us to predict the most likely wing shapes that could have been employed by these ancient reptiles, and further show that a combination of anterior sweep and a reflexed proximal wing section provides an aerodynamically balanced and efficient theoretical pterosaur wing shape, with clear benefits for their flight stability. PMID:21957137

  17. Evaluation of installed performance of a wing-tip-mounted pusher turboprop on a semispan wing

    NASA Technical Reports Server (NTRS)

    Patterson, James C., Jr.; Bartlett, Glynn R.

    1987-01-01

    An exploratory investigation has been conducted at the Langley Research Center to determine the effect of a wing-tip-mounted pusher turboprop on the aerodynamic characteristics of a semispan wing. Tests were conducted on a semispan model with an upswept, untapered wing and an airdriven motor that powered an SR-2 high-speed propeller located on the tip of the wing as a pusher propeller. All tests were conducted at a Mach number of 0.70 over an angle-of-attack range from approximately -2 to 4 deg at a Reynolds number of 3.82 x 10 to the 6th based on the wing reference chord of 13 in. The data indicate that, as a result of locating the propeller behind the wing trailing edge at the wing tip in the crossflow of the wing-tip vortex, it is possible to improve propeller performance and simultaneously reduce the lift-induced drag.

  18. Elements of the Wing Section Theory and of the Wing Theory

    NASA Technical Reports Server (NTRS)

    Munk, Max M.

    1979-01-01

    Results are presented of the theory of wings and of wing sections which are of immediate practical value. They are proven and demonstrated by the use of the simple conceptions of kinetic energy and momentum only.

  19. Natural processes in delta restoration: application to the Mississippi Delta.

    PubMed

    Paola, Chris; Twilley, Robert R; Edmonds, Douglas A; Kim, Wonsuck; Mohrig, David; Parker, Gary; Viparelli, Enrica; Voller, Vaughan R

    2011-01-01

    Restoration of river deltas involves diverting sediment and water from major channels into adjoining drowned areas, where the sediment can build new land and provide a platform for regenerating wetland ecosystems. Except for local engineered structures at the points of diversion, restoration mainly relies on natural delta-building processes. Present understanding of such processes is sufficient to provide a basis for determining the feasibility of restoration projects through quantitative estimates of land-building rates and sustainable wetland area under different scenarios of sediment supply, subsidence, and sea-level rise. We are not yet to the point of being able to predict the evolution of a restored delta in detail. Predictions of delta evolution are based on field studies of active deltas, deltas in mine-tailings ponds, experimental deltas, and countless natural experiments contained in the stratigraphic record. These studies provide input for a variety of mechanistic delta models, ranging from radially averaged formulations to more detailed models that can resolve channels, topography, and ecosystem processes. Especially exciting areas for future research include understanding the mechanisms by which deltaic channel networks self-organize, grow, and distribute sediment and nutrients over the delta surface and coupling these to ecosystem processes, especially the interplay of topography, network geometry, and ecosystem dynamics. PMID:21329199

  20. Incompressible flutter characteristics of representative aircraft wings

    NASA Technical Reports Server (NTRS)

    Wilts, C H

    1958-01-01

    This report gives the results of a detailed study of the flutter characteristics of four representative aircraft wings. This study was made using the electric-analog computer at the California Institute of Technology. During the course of this investigation eight important parameters of each wing were varied and, in addition, the effects of mass, inertia, pitching spring, and location of a concentrated mass were investigated for all four wings and at several sweepback angles.

  1. Pioneer Launch on Delta Vehicle

    NASA Technical Reports Server (NTRS)

    1969-01-01

    NASA launches the last in the series of interplanetary Pioneer spacecraft, Pioneer 10 from Cape Kennedy, Florida. The long-tank Delta launch vehicle placed the spacecraft in a solar orbit along the path of Earth's orbit. The spacecraft then passed inside and outside Earth's orbit, alternately speeding up and slowing down relative to Earth. The Delta launch vehicle family started development in 1959. The Delta was composed of parts from the Thor, an intermediate-range ballistic missile, as its first stage, and the Vanguard as its second. The first Delta was launched from Cape Canaveral on May 13, 1960 and was powerful enough to deliver a 100-pound spacecraft into geostationary transfer orbit. Delta has been used to launch civil, commercial, and military satellites into orbit. For more information about Delta, please see Chapter 3 in Roger Launius and Dennis Jenkins' book To Reach the High Frontier published by The University Press of Kentucky in 2002.

  2. Delta in Eberswalde

    NASA Technical Reports Server (NTRS)

    2006-01-01

    This HiRISE image covers a portion of a delta that partially fills Eberswalde crater in Margaritifer Sinus. The delta was first recognized and mapped using MOC images that revealed various features whose presence required sustained flow and deposition into a lake that once occupied the crater. The HiRISE image resolves meter-scale features that record the migration of channels and delta distributaries as the delta grew over time. Differences in grain-size of sediments within the environments on the delta enable differential erosion of the deposits. As a result, coarser channel deposits are slightly more resistant and stand in relief relative to finer-grained over-bank and more easily eroded distal delta deposits. Close examination of the relict channel deposits confirms the presence of some meter-size blocks that were likely too coarse to have been transported by water flowing within the channels. These blocks may be formed of the sand and gravel that more likely moved along the channels that was lithified and eroded. Numerous meter-scale polygonal structures are common on many surfaces, but mostly those associated with more quiescent depositional environments removed from the channels. The polygons could be the result of deposition of fine-grained sediments that were either exposed and desiccated (dried out), rich in clays that shrunk when the water was removed, turned into rock and then fractured and eroded, or some combination of these processes.

    Image PSP_001336_1560 was taken by the High Resolution Imaging Science Experiment (HiRISE) camera onboard the Mars Reconnaissance Orbiter spacecraft on November 8, 2006. The complete image is centered at -23.8 degrees latitude, 326.4 degrees East longitude. The range to the target site was 256.3 km (160.2 miles). At this distance the image scale is 25.6 cm/pixel (with 1 x 1 binning) so objects 77 cm across are resolved. The image shown here has been map-projected to 25 cm/pixel and north is up. The image was

  3. Optimal redesign study of the harm wing

    NASA Technical Reports Server (NTRS)

    Mcintosh, S. C., Jr.; Weynand, M. E.

    1984-01-01

    The purpose of this project was to investigate the use of optimization techniques to improve the flutter margins of the HARM AGM-88A wing. The missile has four cruciform wings, located near mid-fuselage, that are actuated in pairs symmetrically and antisymmetrically to provide pitch, yaw, and roll control. The wings have a solid stainless steel forward section and a stainless steel crushed-honeycomb aft section. The wing restraint stiffness is dependent upon wing pitch amplitude and varies from a low value near neutral pitch attitude to a much higher value at off-neutral pitch attitudes, where aerodynamic loads lock out any free play in the control system. The most critical condition for flutter is the low-stiffness condition in which the wings are moved symmetrically. Although a tendency toward limit-cycle flutter is controlled in the current design by controller logic, wing redesign to improve this situation is attractive because it can be accomplished as a retrofit. In view of the exploratory nature of the study, it was decided to apply the optimization to a wing-only model, validated by comparison with results obtained by Texas Instruments (TI). Any wing designs that looked promising were to be evaluated at TI with more complicated models, including body modes. The optimization work was performed by McIntosh Structural Dynamics, Inc. (MSD) under a contract from TI.

  4. High performance forward swept wing aircraft

    NASA Technical Reports Server (NTRS)

    Koenig, David G. (Inventor); Aoyagi, Kiyoshi (Inventor); Dudley, Michael R. (Inventor); Schmidt, Susan B. (Inventor)

    1988-01-01

    A high performance aircraft capable of subsonic, transonic and supersonic speeds employs a forward swept wing planform and at least one first and second solution ejector located on the inboard section of the wing. A high degree of flow control on the inboard sections of the wing is achieved along with improved maneuverability and control of pitch, roll and yaw. Lift loss is delayed to higher angles of attack than in conventional aircraft. In one embodiment the ejectors may be advantageously positioned spanwise on the wing while the ductwork is kept to a minimum.

  5. The function of resilin in honeybee wings.

    PubMed

    Ma, Yun; Ning, Jian Guo; Ren, Hui Lan; Zhang, Peng Fei; Zhao, Hong Yan

    2015-07-01

    The present work aimed to reveal morphological characteristics of worker honeybee (Apis mellifera) wings and demonstrate the function of resilin on camber changes during flapping flight. Detailed morphological investigation of the wings showed that different surface characteristics appear on the dorsal and ventral side of the honeybee wings and the linking structure connecting the forewing and hindwing plays an indispensable role in honeybee flapping flight. Resilin stripes were found on both the dorsal and ventral side of the wings, and resilin patches mostly existed on the ventral side. On the basis of resilin distribution, five flexion lines and three cambered types around the lines of passive deformation of the coupled-wing profile were obtained, which defined the deformation mechanism of the wing along the chord, i.e. concave, flat plate and convex. From a movie obtained using high-speed photography from three orthogonal views of free flight in honeybees, periodic changes of the coupled-wing profile were acquired and further demonstrated that the deformation mechanism is a fundamental property for variable deformed shapes of the wing profile during flapping flight, and, in particular, the flat wing profile achieves a nice transition between downstrokes and upstrokes. PMID:25987733

  6. A magnetic fluid microdevice using insect wings

    NASA Astrophysics Data System (ADS)

    Sudo, S.; Tsuyuki, K.; Yano, T.; Takagi, K.

    2008-05-01

    A magnetic fluid microdevice using Diptera insect wings is proposed and constructed. The magnetic fluid device is composed of insect wings, a small permanent magnet, coil, and kerosene-based magnetic fluid. First, the structural properties of insect wings are studied through measurements of certain morphological parameters. Secondly, the novel type of microwind energy converter is constructed. Thirdly, the power generation characteristics of the magnetic fluid microdevice using insect wings are examined. It is found that the output power is roughly proportional to the cube of the airflow velocity.

  7. Subtractive Structural Modification of Morpho Butterfly Wings.

    PubMed

    Shen, Qingchen; He, Jiaqing; Ni, Mengtian; Song, Chengyi; Zhou, Lingye; Hu, Hang; Zhang, Ruoxi; Luo, Zhen; Wang, Ge; Tao, Peng; Deng, Tao; Shang, Wen

    2015-11-11

    Different from studies of butterfly wings through additive modification, this work for the first time studies the property change of butterfly wings through subtractive modification using oxygen plasma etching. The controlled modification of butterfly wings through such subtractive process results in gradual change of the optical properties, and helps the further understanding of structural optimization through natural evolution. The brilliant color of Morpho butterfly wings is originated from the hierarchical nanostructure on the wing scales. Such nanoarchitecture has attracted a lot of research effort, including the study of its optical properties, its potential use in sensing and infrared imaging, and also the use of such structure as template for the fabrication of high-performance photocatalytic materials. The controlled subtractive processes provide a new path to modify such nanoarchitecture and its optical property. Distinct from previous studies on the optical property of the Morpho wing structure, this study provides additional experimental evidence for the origination of the optical property of the natural butterfly wing scales. The study also offers a facile approach to generate new 3D nanostructures using butterfly wings as the templates and may lead to simpler structure models for large-scale man-made structures than those offered by original butterfly wings. PMID:26397977

  8. Wing Damage Effects on Dragonfly's maneuverability

    NASA Astrophysics Data System (ADS)

    Ning, Zhe; Gai, Kuo; Zeyghami, Samane; Dong, Haibo; Flow Simulation Research Group (FSRG) Team

    2011-11-01

    In this work, how the insect flight behavior contributes to its adaptability to limited performance condition is studied through a combined experimental and computational study. High speed photogrammetry is used to collect the data of dragonflies' takeoffs with intact and damaged wings along the chord and span separately. Then the effect of the spanwise and chordwise damage on the dragonfly wing is investigated. Results show that both changes have different effects on the wing and body kinematics and the merit of maneuverability. Two theories will be introduced to explain the wing damage tolerance behavior of the dragonfly flight. This work is supported by NSF CBET-1055949.

  9. Veins improve fracture toughness of insect wings.

    PubMed

    Dirks, Jan-Henning; Taylor, David

    2012-01-01

    During the lifetime of a flying insect, its wings are subjected to mechanical forces and deformations for millions of cycles. Defects in the micrometre thin membranes or veins may reduce the insect's flight performance. How do insects prevent crack related material failure in their wings and what role does the characteristic vein pattern play? Fracture toughness is a parameter, which characterises a material's resistance to crack propagation. Our results show that, compared to other body parts, the hind wing membrane of the migratory locust S. gregaria itself is not exceptionally tough (1.04±0.25 MPa√m). However, the cross veins increase the wing's toughness by 50% by acting as barriers to crack propagation. Using fracture mechanics, we show that the morphological spacing of most wing veins matches the critical crack length of the material (1132 µm). This finding directly demonstrates how the biomechanical properties and the morphology of locust wings are functionally correlated in locusts, providing a mechanically 'optimal' solution with high toughness and low weight. The vein pattern found in insect wings thus might inspire the design of more durable and lightweight artificial 'venous' wings for micro-air-vehicles. Using the vein spacing as indicator, our approach might also provide a basis to estimate the wing properties of endangered or extinct insect species. PMID:22927966

  10. The supercritical profile of the supercritical wing

    NASA Technical Reports Server (NTRS)

    Wagner, O.

    1981-01-01

    The profile wing design for supercritical structures is discussed. Emphasis is placed on the flow of air surrounding the wing and variations in flow fields are examined. Modifications to the profile for flight below transonic level are presented that increase the uplift pressure and permit the achievement of critical Mach numbers on the order of 0.85. The uplift pressure along the upper side of the profile is compared for a classical and a Peaky profile. A comparison of classical and supercritical wing cross sections indicates a flatter upper side, a large nose radius, and a thicker profile to the supercritical wing.

  11. Definition of the unsteady vortex flow over a wing/body configuration

    NASA Technical Reports Server (NTRS)

    Liou, S. G.; Debry, B.; Lenakos, J.; Caplin, J.; Komerath, N. M.

    1991-01-01

    A problem of current interest in computational aerodynamics is the prediction of unsteady vortex flows over aircraft at high angles of attack. A six-month experimental effort was conducted at the John H. Harper Wind Tunnel to acquire qualitative and quantitative information on the unsteady vortex flow over a generic wing-body configuration at high angles of attack. A double-delta flat-plate wing with beveled edges was combined with a slender sharp-nosed body-of-revolution fuselage to form the generic configuration. This configuration produces a strong attached leading edge vortex on the wing, as well as sharply-peaked flow velocity spectra above the wing. While it thus produces flows with several well-defined features of current interest, the model was designed for efficiency of representation in computational codes. A moderate number of surface pressure ports and two unsteady pressure sensors were used to study the pressure distribution over the wing and body surface at high angles of attack; the unsteady pressure sensing did not succeed because of inadequate signal-to-noise ratio. A pulsed copper vapor laser sheet was used to visualize the vortex flow over the model, and vortex trajectories, burst locations, mutual induction of vortex systems from the forebody, strake, and wing, were quantified. Laser Doppler velocimetry was used to quantify all 3 components of the time-average velocity in 3 data planes perpendicular to the freestream direction. Statistics of the instantaneous velocity were used to study intermittency and fluctuation intensity. Hot-film anemometry was used to study the fluctuation energy content in the velocity field, and the spectra of these fluctuations. In addition, a successful attempt was made to measure velocity spectra, component by component, using laser velocimetry, and these were compared with spectra measured by hot-film anemometry at several locations.

  12. Pegasus Rocket Wing and PHYSX Glove Undergoes Stress Loads Testing

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The Pegasus Hypersonic Experiment (PHYSX) Project's Pegasus rocket wing with attached PHYSX glove rests after load-tests at Scaled Composites, Inc., in Mojave, California, in January 1997. Technicians slowly filled water bags beneath the wing, to create the pressure, or 'wing-loading,' required to determine whether the wing could withstand its design limit for stress. The wing sits in a wooden triangular frame which serves as the test-rig, mounted to the floor atop the waterbags. Pegasus is an air-launched space booster produced by Orbital Sciences Corporation and Hercules Aerospace Company (initially; later, Alliant Tech Systems) to provide small satellite users with a cost-effective, flexible, and reliable method for placing payloads into low earth orbit. Pegasus has been used to launch a number of satellites and the PHYSX experiment. That experiment consisted of a smooth glove installed on the first-stage delta wing of the Pegasus. The glove was used to gather data at speeds of up to Mach 8 and at altitudes approaching 200,000 feet. The flight took place on October 22, 1998. The PHYSX experiment focused on determining where boundary-layer transition occurs on the glove and on identifying the flow mechanism causing transition over the glove. Data from this flight-research effort included temperature, heat transfer, pressure measurements, airflow, and trajectory reconstruction. Hypersonic flight-research programs are an approach to validate design methods for hypersonic vehicles (those that fly more than five times the speed of sound, or Mach 5). Dryden Flight Research Center, Edwards, California, provided overall management of the glove experiment, glove design, and buildup. Dryden also was responsible for conducting the flight tests. Langley Research Center, Hampton, Virginia, was responsible for the design of the aerodynamic glove as well as development of sensor and instrumentation systems for the glove. Other participating NASA centers included Ames Research

  13. Projection Moire Interferometry Measurements of Micro Air Vehicle Wings

    NASA Technical Reports Server (NTRS)

    Fleming, Gary A.; Bartram, Scott M.; Waszak, Martin R.; Jenkins, Luther N.

    2001-01-01

    Projection Moire Interferometry (PMI) has been used to measure the structural deformation of micro air vehicle (MAV) wings during a series of wind tunnel tests. The MAV wings had a highly flexible wing structure, generically reminiscent of a bat s wing, which resulted in significant changes in wing shape as a function of MAV angle-of-attack and simulated flight speed. This flow-adaptable wing deformation is thought to provide enhanced vehicle stability and wind gust alleviation compared to rigid wing designs. Investigation of the potential aerodynamic benefits of a flexible MAV wing required measurement of the wing shape under aerodynamic loads. PMI was used to quantify the aerodynamically induced changes in wing shape for three MAV wings having different structural designs and stiffness characteristics. This paper describes the PMI technique, its application to MAV testing, and presents a portion of the PMI data acquired for the three different MAV wings tested.

  14. Projection moire interferometry measurements of micro air vehicle wings

    NASA Astrophysics Data System (ADS)

    Fleming, Gary A.; Bartram, Scott M.; Waszak, Martin R.; Jenkins, Luther N.

    2001-11-01

    Projection Moire Interferometry (PMI) has been used to measure the structural deformation of micro air vehicle (MAV) wings during a series of wind tunnel tests. The MAV wings had a highly flexible wing structure, generically reminiscent of a bat's wing, which resulted in significant changes in wing shape as a function of MAV angle-of-attack and simulated flight speed. This flow-adaptable wing deformation is thought to provide enhanced vehicle stability and wind gust alleviation compared to rigid wing designs. Investigation of the potential aerodynamic benefits of a flexible MAV wing required measurement of the wing shape under aerodynamic loads. PMI was used to quantify the aerodynamically induced changes in wing shape for three MAV wings having different structural designs and stiffness characteristics. This paper describes the PMI technique, its application to MAV testing, and presents a portion of the PMI data acquired for the three different MAV wings tested.

  15. Topological structures of vortex flow on a flying wing aircraft, controlled by a nanosecond pulse discharge plasma actuator

    NASA Astrophysics Data System (ADS)

    Du, Hai; Shi, Zhiwei; Cheng, Keming; Wei, Dechen; Li, Zheng; Zhou, Danjie; He, Haibo; Yao, Junkai; He, Chengjun

    2016-06-01

    Vortex control is a thriving research area, particularly in relation to flying wing or delta wing aircraft. This paper presents the topological structures of vortex flow on a flying wing aircraft controlled by a nanosecond plasma dielectric barrier discharge actuator. Experiments, including oil flow visualization and two-dimensional particle image velocimetry (PIV), were conducted in a wind tunnel with a Reynolds number of 0.5 × 106. Both oil and PIV results show that the vortex can be controlled. Oil topological structures on the aircraft surface coincide with spatial PIV flow structures. Both indicate vortex convergence and enhancement when the plasma discharge is switched on, leading to a reduced region of separated flow.

  16. Extension of leading-edge-suction analogy to wings with separated flow around the side edges at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Lamar, J. E.

    1974-01-01

    A method for determining the lift, drag, and pitching moment for wings which have separated flow at the leading and side edges with subsequently reattached flow downstream and inboard is presented. Limiting values of the contribution to lift of the side-edge reattached flow are determined for rectangular wings. The general behavior of this contribution is computed for rectangular, cropped-delta, cropped-diamond, and cropped-arrow wings. Comparisons of the results of the method and experiment indicate reasonably good correlation of the lift, drag, and pitching moment for a wide planform range. The agreement of the method with experiment was as good as, or better than, that obtained by other methods. The procedure is computerized and is available from COSMIC as NASA Langley computer program A0313.

  17. Holden Crater Delta

    NASA Technical Reports Server (NTRS)

    2006-01-01

    [figure removed for brevity, see original site] Context image for PIA03694 Holden Crater Delta

    This fan-shaped delta deposit is located in Holden Crater.

    Image information: VIS instrument. Latitude -27.3N, Longitude 324.5E. 17 meter/pixel resolution.

    Note: this THEMIS visual image has not been radiometrically nor geometrically calibrated for this preliminary release. An empirical correction has been performed to remove instrumental effects. A linear shift has been applied in the cross-track and down-track direction to approximate spacecraft and planetary motion. Fully calibrated and geometrically projected images will be released through the Planetary Data System in accordance with Project policies at a later time.

    NASA's Jet Propulsion Laboratory manages the 2001 Mars Odyssey mission for NASA's Office of Space Science, Washington, D.C. The Thermal Emission Imaging System (THEMIS) was developed by Arizona State University, Tempe, in collaboration with Raytheon Santa Barbara Remote Sensing. The THEMIS investigation is led by Dr. Philip Christensen at Arizona State University. Lockheed Martin Astronautics, Denver, is the prime contractor for the Odyssey project, and developed and built the orbiter. Mission operations are conducted jointly from Lockheed Martin and from JPL, a division of the California Institute of Technology in Pasadena.

  18. Shanghai Delta Complex

    SciTech Connect

    Hart, R.E.; Hoffman, P.F.; Parker, R.W.

    1988-01-01

    The upper Eocene Yegua Formation expands dramatically across a regional system of growth faults into an area generally 12-15 km wide, extending at least from the western edge of the Houston sale dome basin to the San Marcos arch. Within this area, the expanded Yegua trend has yielded, since 1982, at least seven noteworthy discoveries: Toro Grande and Lost Bridge fields in Jackson County, and Black Owl, Shanghai, Shanghai East, El Campo, and Phase Four fields in Wharton County. During each of several postulated Yegua sea level drops, this flexure became a focal point for deltaic deposition of excellent reservoir-quality sands. Shanghai, Shanghai East, and El Campo fields are located within what the writers have labeled the ''Shanghai delta complex.'' Integration of seismic and well data in this vicinity shows a marked increase in the expansion indices of growth faults, and moderately thick progradational sand sequences have accumulated immediately downthrow. This structural-stratigraphic pattern, as well as internal bedding characteristics and other lithologic data observed, is believed typical of deltas deposited along the Yegua shelf margin.

  19. Thermostatted delta f

    SciTech Connect

    Krommes, J.A.

    2000-01-18

    The delta f simulation method is revisited. Statistical coarse-graining is used to rigorously derive the equation for the fluctuation delta f in the particle distribution. It is argued that completely collisionless simulation is incompatible with the achievement of true statistically steady states with nonzero turbulent fluxes because the variance of the particle weights w grows with time. To ensure such steady states, it is shown that for dynamically collisionless situations a generalized thermostat or W-stat may be used in lieu of a full collision operator to absorb the flow of entropy to unresolved fine scales in velocity space. The simplest W-stat can be implemented as a self-consistently determined, time-dependent damping applied to w. A precise kinematic analogy to thermostatted nonequilibrium molecular dynamics (NEMD) is pointed out, and the justification of W-stats for simulations of turbulence is discussed. An extrapolation procedure is proposed such that the long-time, steady-state, collisionless flux can be deduced from several short W-statted runs with large effective collisionality, and a numerical demonstration is given.

  20. Space shuttle: Static stability and control investigation of NR/GD delta wing booster (B-20) and delta wing orbiter (134D), volume 4

    NASA Technical Reports Server (NTRS)

    Allen, E. C., Jr.; Eder, F. W.

    1972-01-01

    Test results of booster and orbiter models of various component buildup configurations are reported. Dataset Collation Sheets, which give a complete summary of the configurations, are presented along with a description of the test facility. Data reduction procedures are described.

  1. On wings and keels (2)

    NASA Astrophysics Data System (ADS)

    Slooff, J. W.

    1985-05-01

    The physical mechanisms governing the hydrodynamics of sailing yacht keels and the parameters that, through these mechanisms, determine keel performance are discussed. It is concluded that due to the presence of the free water surface optimum keel shapes differ from optimum shapes for aircraft wings. Utilizing computational fluid dynamics analysis and optimization it is found that the performance of conventional keels can be improved significantly by reducing taper or even applying inverse taper (upside-down keel) and that decisive improvements in performance can be realized through keels with winglets.

  2. The Aerodynamic Aspect of Wing-fuselage Fillets

    NASA Technical Reports Server (NTRS)

    Muttray, H

    1935-01-01

    Model tests prove the feasibility of enhancing the aerodynamic qualities of wing-fuselage fillets by appropriate design of fuselage and wing roots. Abrupt changes from maximum fuselage height to wing chord must be avoided and every longitudinal section of fuselage and wing roots must be so faired and arranged as to preserve the original lift distribution of the continuous wing. Adapting the fuselage to the curvilinear circulation of the wing affords further improvement. The polars of such arrangements are almost the same as those of the "wing alone," thus voiding the superiority of the high-wing type airplane known with conventional design.

  3. Effect of wing flexibility on phasing of tandem wings in forward flight

    NASA Astrophysics Data System (ADS)

    Naidu, Vishal; Young, John; Lai, Joseph

    2014-11-01

    The dragonfly with two pairs of wings in tandem uses different phases between the wing pairs to suit the needs of the flight. Previous studies to understand the effect of phasing in forward flight are based on rigid wings. This is in contrast to the highly flexible dragonfly wings, with varying spanwise and chordwise flexibility. Here, we study flexible flapping wing simulations using Fluid Structure Interaction (FSI) in forward flight, at an advance ratio of 0.3 and Reynolds number of approximately 1300. The FSI simulations are carried out for phase 90° (hindwing leading), 0° (in-phase) and 180° (anti-phase). The performance of flexible wings will be compared with that of the rigid wings and the effect of flexibility will be discussed. PhD Student.

  4. Wing Deployment Sequence #2: The deployable, inflatable wing technology demonstrator experiment airc

    NASA Technical Reports Server (NTRS)

    2001-01-01

    Wing Deployment Sequence #2: The deployable, inflatable wing technology demonstrator experiment aircraft's wings continue deploying following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, California. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.

  5. Wing Deployment Sequence #1: The deployable, inflatable wing technology demonstrator experiment airc

    NASA Technical Reports Server (NTRS)

    2001-01-01

    Wing Deployment Sequence #1: The deployable, inflatable wing technology demonstrator experiment aircraft's wings begin deploying following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, California. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.

  6. Wing Deployment Sequence #3: The deployable, inflatable wing technology demonstrator experiment airc

    NASA Technical Reports Server (NTRS)

    2001-01-01

    Wing Deployment Sequence #3: The deployable, inflatable wing technology demonstrator experiment aircraft's wings fully deployed during flight following separation from its carrier aircraft during a flight conducted by the NASA Dryden Flight Research Center, Edwards, Californiaornia. The inflatable wing project represented a basic flight research effort by Dryden personnel. Three successful flights of the I2000 inflatable wing aircraft occurred. During the flights, the team air-launched the radio-controlled (R/C) I2000 from an R/C utility airplane at an altitude of 800-1000 feet. As the I2000 separated from the carrier aircraft, its inflatable wings 'popped-out,' deploying rapidly via an on-board nitrogen bottle. The aircraft remained stable as it transitioned from wingless to winged flight. The unpowered I2000 glided down to a smooth landing under complete control.

  7. Advanced wing design survivability testing and results

    NASA Technical Reports Server (NTRS)

    Bruno, J.; Tobias, M.

    1992-01-01

    Composite wings on current operational aircraft are conservatively designed to account for stress/strain concentrations, and to assure specified damage tolerance. The technology that can lead to improved composite wing structures and associated structural efficiency is to increase design ultimate strain levels beyond their current limit of 3500 to 4000 micro-in/in to 6000 micro-in/in without sacrificing structural integrity, durability, damage tolerance, or survivability. Grumman, under the sponsorship of the Naval Air Development Center (NADC), has developed a high-strain composite wing design for a subsonic aircraft wing using novel and innovative design concepts and manufacturing methods, while maintaining a state-of-the-art fiber/resin system. The current advanced wing design effort addressed a tactical subsonic aircraft wing using previously developed, high-strain wing design concepts in conjunction with newer/emerging fiber and polymer matrix composite (PMC) materials to achieve the same goals, while reducing complexity. Two categories of advanced PMC materials were evaluated: toughened thermosets; and engineered thermoplastics. Advanced PMC materials offer the technological opportunity to take maximum advantage of improved material properties, physical characteristics, and tailorability to increase performance and survivability over current composite structure. Damage tolerance and survivability to various threats, in addition to structural integrity and durability, were key technical issues addressed during this study, and evaluated through test. This paper focuses on the live-fire testing, and the results performed to experimentally evaluate the survivability of the advanced wing design.

  8. Computer Code Aids Design Of Wings

    NASA Technical Reports Server (NTRS)

    Carlson, Harry W.; Darden, Christine M.

    1993-01-01

    AERO2S computer code developed to aid design engineers in selection and evaluation of aerodynamically efficient wing/canard and wing/horizontal-tail configurations that includes simple hinged-flap systems. Code rapidly estimates longitudinal aerodynamic characteristics of conceptual airplane lifting-surface arrangements. Developed in FORTRAN V on CDC 6000 computer system, and ported to MS-DOS environment.

  9. Biaxial mechanical characterization of bat wing skin.

    PubMed

    Skulborstad, A J; Swartz, S M; Goulbourne, N C

    2015-06-01

    The highly flexible and stretchable wing skin of bats, together with the skeletal structure and musculature, enables large changes in wing shape during flight. Such compliance distinguishes bat wings from those of all other flying animals. Although several studies have investigated the aerodynamics and kinematics of bats, few have examined the complex histology and mechanical response of the wing skin. This work presents the first biaxial characterization of the local deformation, mechanical properties, and fiber kinematics of bat wing skin. Analysis of these data has provided insight into the relationships among the structural morphology, mechanical properties, and functionality of wing skin. Large spatial variations in tissue deformation and non-negligible fiber strains in the cross-fiber direction for both chordwise and spanwise fibers indicate fibers should be modeled as two-dimensional elements. The macroscopic constitutive behavior was anisotropic and nonlinear, with very low spanwise and chordwise stiffness (hundreds of kilopascals) in the toe region of the stress-strain curve. The structural arrangement of the fibers and matrix facilitates a low energy mechanism for wing deployment and extension, and we fabricate examples of skins capturing this mechanism. We propose a comprehensive deformation map for the entire loading regime. The results of this work underscore the importance of biaxial field approaches for soft heterogeneous tissue, and provide a foundation for development of bio-inspired skins to probe the effects of the wing skin properties on aerodynamic performance. PMID:25895436

  10. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  11. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  12. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  13. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  14. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  15. Numerical investigation of insect wing fracture behaviour.

    PubMed

    Rajabi, H; Darvizeh, A; Shafiei, A; Taylor, D; Dirks, J-H

    2015-01-01

    The wings of insects are extremely light-weight biological composites with exceptional biomechanical properties. In the recent years, numerical simulations have become a very powerful tool to answer experimentally inaccessible questions on the biomechanics of insect flight. However, many of the presented models require a sophisticated balance of biomechanical material parameters, many of which are not yet available. In this article we show the first numerical simulations of crack propagation in insect wings. We have used a combination of the maximum-principal stress theory, the traction separation law and basic biomechanical properties of cuticle to develop simple yet accurate finite element (FE) models of locust wings. The numerical results of simulated tensile tests on wing samples are in very good qualitative, and interestingly, also in excellent quantitative agreement with previously obtained experimental data. Our study further supports the idea that the cross-veins in insect wings act as barriers against crack propagation and consequently play a dominant role in toughening the whole wing structure. The use of numerical simulations also allowed us to combine experimental data with previously inaccessible data, such as the distribution of the first principal stress through the wing membrane and the veins. A closer look at the stress-distribution within the wings might help to better understand fracture-toughening mechanisms and also to design more durable biomimetic micro-air vehicles. PMID:25468669

  16. Unsteady vortex-dominated flow around wings with oscillating leading-edge flaps

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1991-01-01

    Unsteady vortex-dominated flow around delta wings with oscillating leading-edge flaps represents an important classs of problems for supermaneuverability and flow control of advanced aircraft. The problem is solved using time accurate integration of the unsteady, compressible, thin-layer Navier-Stokes equations in conjunction with the unsteady, linearized, Navier-displacement equations. Starting with an initial configuration of the wing and its flaps, the Navier-Stokes equations are solved on an initial structured grid for the steady flow. The forced oscillation of the flaps is then applied, and the problem is solved accurately in time. The Navier-displacement equations are solved for the grid deformation and the Navier-Stokes equations are solved for the flowfield. Symmetric and anti-symmetric flaps oscillations are presented to study the effect of the flaps oscillation on the leading-edge vortical flow.

  17. Some Divergence Characteristics of Low-Aspect-Ratio Wings at Transonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Woolston, Donald S.; Gibson, Frederick W.; Cunningham, Herbert J.

    1960-01-01

    The problem of chordwise, or camber, divergence at transonic and supersonic speeds is treated with primary emphasis on slender delta wings having a cantilever support at the trailing edge. Experimental and analytical results are presented for four wing models having apex half-angles of 5 deg, 10 deg, 15 deg, and 20 deg. A Mach number range from 0.8 to 7.3 is covered. The analytical results include calculations based on small-aspect-ratio theory, lifting-surface theory, and strip theory. A closed-form solution of the equilibrium equation is given, which is based on low-aspect-ratio theory but which applies only to certain stiffness distributions. Also presented is an iterative procedure for use with other aerodynamic theories and with arbitrary stiffness distribution.

  18. Delta Electroproduction in 12-C

    SciTech Connect

    Steven McLauchlan

    2003-01-31

    The Delta-nucleus potential is a crucial element in the understanding of the nuclear system. Previous electroexcitation measurements in the delta region reported a Q2 dependence of the delta mass indicating that this potential is dependent on the momentum of the delta. Such a dependence is not observed for protons and neutrons in the nuclear medium. This thesis presents the experimental study of the electroexcitation of the delta resonance in 12C, performed using the high energy electron beam at the Thomas Jefferson National Accelerator Facility, and the near 4(pie) acceptance detector CLAS that enables the detection of the full reaction final state. Inclusive, semi inclusive, and exclusive cross sections were measured with an incident electron beam energy of 1.162GeV over the Q2 range 0.175-0.475 (GeV/c)2. A Q2 dependence of the delta mass was only observed in the exclusive measurements indicating that the delta-nucleus potential is affected by the momentum of the delta.

  19. High speed flow past wings

    NASA Technical Reports Server (NTRS)

    Norstrud, H.

    1973-01-01

    The analytical solution to the transonic small perturbation equation which describes steady compressible flow past finite wings at subsonic speeds can be expressed as a nonlinear integral equation with the perturbation velocity potential as the unknown function. This known formulation is substituted by a system of nonlinear algebraic equations to which various methods are applicable for its solution. Due to the presence of mathematical discontinuities in the flow solutions, however, a main computational difficulty was to ensure uniqueness of the solutions when local velocities on the wing exceeded the speed of sound. For continuous solutions this was achieved by embedding the algebraic system in an one-parameter operator homotopy in order to apply the method of parametric differentiation. The solution to the initial system of equations appears then as a solution to a Cauchy problem where the initial condition is related to the accompanying incompressible flow solution. In using this technique, however, a continuous dependence of the solution development on the initial data is lost when the solution reaches the minimum bifurcation point. A steepest descent iteration technique was therefore, added to the computational scheme for the calculation of discontinuous flow solutions. Results for purely subsonic flows and supersonic flows with and without compression shocks are given and compared with other available theoretical solutions.

  20. Swept wing ice accretion modeling

    NASA Technical Reports Server (NTRS)

    Potapczuk, Mark G.; Bidwell, Colin S.

    1990-01-01

    An effort to develop a three-dimensional modeling method was initiated. This first step towards creation of a complete aircraft icing simulation code builds on previously developed methods for calculating three-dimensional flow fields and particle trajectories combined with a two-dimensional ice accretion calculation along coordinate locations corresponding to streamlines. This work is a demonstration of the types of calculations necessary to predict a three-dimensional ice accretion. Results of calculations using the 3-D method for a MS-317 swept wing geometry are projected onto a 2-D plane normal to the wing leading edge and compared to 2-D results for the same geometry. It is anticipated that many modifications will be made to this approach, however, this effort will lay the groundwork for future modeling efforts. Results indicate that the flow field over the surface and the particle trajectories differed for the two calculations. This led to lower collection efficiencies, convective heat transfer coefficients, freezing fractions, and ultimately ice accumulation for the 3-D calculation.

  1. Collective Flow Enhancement by Tandem Flapping Wings.

    PubMed

    Gravish, Nick; Peters, Jacob M; Combes, Stacey A; Wood, Robert J

    2015-10-30

    We examine the fluid-mechanical interactions that occur between arrays of flapping wings when operating in close proximity at a moderate Reynolds number (Re≈100-1000). Pairs of flapping wings are oscillated sinusoidally at frequency f, amplitude θ_{M}, phase offset ϕ, and wing separation distance D^{*}, and outflow speed v^{*} is measured. At a fixed separation distance, v^{*} is sensitive to both f and ϕ, and we observe both constructive and destructive interference in airspeed. v^{*} is maximized at an optimum phase offset, ϕ_{max}, which varies with wing separation distance, D^{*}. We propose a model of collective flow interactions between flapping wings based on vortex advection, which reproduces our experimental data. PMID:26565499

  2. Collective Flow Enhancement by Tandem Flapping Wings

    NASA Astrophysics Data System (ADS)

    Gravish, Nick; Peters, Jacob M.; Combes, Stacey A.; Wood, Robert J.

    2015-10-01

    We examine the fluid-mechanical interactions that occur between arrays of flapping wings when operating in close proximity at a moderate Reynolds number (Re ≈100 - 1000 ). Pairs of flapping wings are oscillated sinusoidally at frequency f , amplitude θM, phase offset ϕ , and wing separation distance D*, and outflow speed v* is measured. At a fixed separation distance, v* is sensitive to both f and ϕ , and we observe both constructive and destructive interference in airspeed. v* is maximized at an optimum phase offset, ϕmax, which varies with wing separation distance, D*. We propose a model of collective flow interactions between flapping wings based on vortex advection, which reproduces our experimental data.

  3. General theory of thin wing sections

    NASA Technical Reports Server (NTRS)

    Munk, Max M

    1923-01-01

    This report contains a new, simple method of calculating the air forces to which thin wings are subjected at small angles of attack, if their curvature is not too great. Two simple integrals are the result. They contain only the coordinates of the wing section. The first integral gives the angle of attack at which the lift of the wing is zero, the second integral gives the moment experienced by the wing when its angle is zero. The two constants thus obtained are sufficient to determine the lift and moment for any other angle of attack. This with the theory of the aerodynamical induction, and with our empirical knowledge of the drag due to friction, the results are valuable for actual wings also. A particular result obtained is the calculation of the elevator effect. (author)

  4. Design of a transonically profiled wing

    NASA Technical Reports Server (NTRS)

    Kiekebusch, B.

    1978-01-01

    The application of well known design concepts with the combined use of thick transonic profiles to aircraft wing design was investigated. Optimization in terms of weight and operational costs was emphasized. It is shown that the usual design criteria and concepts are too restricted and do not sufficiently represent the physical processes over the wing. Suggestions are made for improving this situation, and a design example given. Compared with a wing design according to previously used criteria, the new design is found to be superior in the most important functions. It is concluded that an isobar concept adjusted to the planform in conjunction with an 'organically' designed wing will lead to the weight optimum solutions of wing profiles.

  5. X-wing potential for Navy applications

    NASA Technical Reports Server (NTRS)

    Linden, Arthur W.; Biggers, James C.

    1987-01-01

    The X-wing provides a VTOL aircraft which has a low disc loading hover capability, similar to a conventional helicopter, combined with a high subsonic cruise speed capability. As a result, it hovers with low fuel flow rates which make extended hover duration missions practical. Its low hover power requirements also permit hovering and low speed flight on only one engine in a high speed twin engine aircraft design. The NASA DARPA Sikorsky RSRA X-wing program developed flightworthy X-wing hardware. All design activity and the majority of its component fabrication is completed. A design study was performed on an X-wing concept demonstrator aircraft which is based on the RSRA X-wing components, combined with two MTE engines and a new fuselage.

  6. An explanation of the instability of the free vortex cores occurring over delta winds with raised edges

    NASA Technical Reports Server (NTRS)

    Ludwieg, H.

    1980-01-01

    By rolling up the surfaces of discontinuity originating from the leading edge of delta wings, free vortex cores are formed above the wing. In case of greater angles of incidence, the flow in these vortex cores shows an instability which abruptly produces strong turbulence. In the present paper an explanation is given of this instability being a "frictionless instability" of the vortex core flow by increasing helical interference vortices. The occurring vortex core flows are calculated and investigated for stability by means of a stability criterion concerning flows with helical streamlines given by H. Ludwieg.

  7. Colorado River Delta

    NASA Technical Reports Server (NTRS)

    2008-01-01

    The Colorado River ends its 2330 km journey in the Gulf of Mexico in Baja California. The heavy use of the river as an irrigation source for the Imperial Valley has dessicated the lower course of the river in Mexico such that it no longer consistently reaches the sea. Prior to the mid 20th century, the Colorado River Delta provided a rich estuarine marshland that is now essentially desiccated, but nonetheless is an important ecological resource.

    The image was acquired May 29, 2006, covers an area of 44.3 x 57.5 km, and is located at 32.1 degrees north latitude, 115.1 degrees west longitude.

    The U.S. science team is located at NASA's Jet Propulsion Laboratory, Pasadena, Calif. The Terra mission is part of NASA's Science Mission Directorate.

  8. Mackenzie River Delta, Canada

    NASA Technical Reports Server (NTRS)

    2007-01-01

    The Mackenzie River in the Northwest Territories, Canada, with its headstreams the Peace and Finley, is the longest river in North America at 4241 km, and drains an area of 1,805,000 square km. The large marshy delta provides habitat for migrating Snow Geese, Tundra Swans, Brant, and other waterfowl. The estuary is a calving area for Beluga whales. The Mackenzie (previously the Disappointment River) was named after Alexander Mackenzie who travelled the river while trying to reach the Pacific in 1789.

    The image was acquired on August 4, 2005, covers an area of 55.8 x 55.8 km, and is located at 68.6 degrees north latitude, 134.7 degrees west longitude.

    The U.S. science team is located at NASA's Jet Propulsion Laboratory, Pasadena, Calif. The Terra mission is part of NASA's Science Mission Directorate.

  9. Numerical computation of viscous flow around bodies and wings moving at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Tannehill, J. C.

    1984-01-01

    Research in aerodynamics is discussed. The development of equilibrium air curve fits; computation of hypersonic rarefield leading edge flows; computation of 2-D and 3-D blunt body laminar flows with an impinging shock; development of a two-dimensional or axisymmetric real gas blunt body code; a study of an over-relaxation procedure forthe MacCormack finite-difference scheme; computation of 2-D blunt body turbulent flows with an impinging shock; computation of supersonic viscous flow over delta wings at high angles of attack; and computation of the Space Shuttle Orbiter flowfield are discussed.

  10. Design, fabrication, and characterization of multifunctional wings to harvest solar energy in flapping wing air vehicles

    NASA Astrophysics Data System (ADS)

    Perez-Rosado, Ariel; Gehlhar, Rachel D.; Nolen, Savannah; Gupta, Satyandra K.; Bruck, Hugh A.

    2015-06-01

    Currently, flapping wing unmanned aerial vehicles (a.k.a., ornithopters or robotic birds) sustain very short duration flight due to limited on-board energy storage capacity. Therefore, energy harvesting elements, such as flexible solar cells, need to be used as materials in critical components, such as wing structures, to increase operational performance. In this paper, we describe a layered fabrication method that was developed for realizing multifunctional composite wings for a unique robotic bird we developed, known as Robo Raven, by creating compliant wing structure from flexible solar cells. The deformed wing shape and aerodynamic lift/thrust loads were characterized throughout the flapping cycle to understand wing mechanics. A multifunctional performance analysis was developed to understand how integration of solar cells into the wings influences flight performance under two different operating conditions: (1) directly powering wings to increase operation time, and (2) recharging batteries to eliminate need for external charging sources. The experimental data is then used in the analysis to identify a performance index for assessing benefits of multifunctional compliant wing structures. The resulting platform, Robo Raven III, was the first demonstration of a robotic bird that flew using energy harvested from solar cells. We developed three different versions of the wing design to validate the multifunctional performance analysis. It was also determined that residual thrust correlated to shear deformation of the wing induced by torsional twist, while biaxial strain related to change in aerodynamic shape correlated to lift. It was also found that shear deformation of the solar cells induced changes in power output directly correlating to thrust generation associated with torsional deformation. Thus, it was determined that multifunctional solar cell wings may be capable of three functions: (1) lightweight and flexible structure to generate aerodynamic forces, (2

  11. Aerodynamics of high frequency flapping wings

    NASA Astrophysics Data System (ADS)

    Hu, Zheng; Roll, Jesse; Cheng, Bo; Deng, Xinyan

    2010-11-01

    We investigated the aerodynamic performance of high frequency flapping wings using a 2.5 gram robotic insect mechanism developed in our lab. The mechanism flaps up to 65Hz with a pair of man-made wing mounted with 10cm wingtip-to-wingtip span. The mean aerodynamic lift force was measured by a lever platform, and the flow velocity and vorticity were measured using a stereo DPIV system in the frontal, parasagittal, and horizontal planes. Both near field (leading edge vortex) and far field flow (induced flow) were measured with instantaneous and phase-averaged results. Systematic experiments were performed on the man-made wings, cicada and hawk moth wings due to their similar size, frequency and Reynolds number. For insect wings, we used both dry and freshly-cut wings. The aerodynamic force increase with flapping frequency and the man-made wing generates more than 4 grams of lift at 35Hz with 3 volt input. Here we present the experimental results and the major differences in their aerodynamic performances.

  12. Insect Wing Displacement Measurement Using Digital Holography

    SciTech Connect

    Aguayo, Daniel D.; Mendoza Santoyo, Fernando; Torre I, Manuel H. de la; Caloca Mendez, Cristian I.

    2008-04-15

    Insects in flight have been studied with optical non destructive techniques with the purpose of using meaningful results in aerodynamics. With the availability of high resolution and large dynamic range CCD sensors the so called interferometric digital holographic technique was used to measure the surface displacement of in flight insect wings, such as butterflies. The wings were illuminated with a continuous wave Verdi laser at 532 nm, and observed with a CCD Pixelfly camera that acquire images at a rate of 11.5 frames per second at a resolution of 1392x1024 pixels and 12 Bit dynamic range. At this frame rate digital holograms of the wings were captured and processed in the usual manner, namely, each individual hologram is Fourier processed in order to find the amplitude and phase corresponding to the digital hologram. The wings displacement is obtained when subtraction between two digital holograms is performed for two different wings position, a feature applied to all consecutive frames recorded. The result of subtracting is seen as a wrapped phase fringe pattern directly related to the wing displacement. The experimental data for different butterfly flying conditions and exposure times are shown as wire mesh plots in a movie of the wings displacement.

  13. Periodic and chaotic flapping of insectile wings

    NASA Astrophysics Data System (ADS)

    Huang, Y.; Kanso, E.

    2015-11-01

    Insects use flight muscles attached at the base of the wings to produce impressive wing flapping frequencies. The maximum power output of these flight muscles is insufficient to maintain such wing oscillations unless there is good elastic storage of energy in the insect flight system. Here, we explore the intrinsic self-oscillatory behavior of an insectile wing model, consisting of two rigid wings connected at their base by an elastic torsional spring. We study the wings behavior as a function of the total energy and spring stiffness. Three types of behavior are identified: end-over-end rotation, chaotic motion, and periodic flapping. Interestingly, the region of periodic flapping decreases as energy increases but is favored as stiffness increases. These findings are consistent with the fact that insect wings and flight muscles are stiff. They further imply that, by adjusting their muscle stiffness to the energy level at which they are operating, insects can maintain periodic flapping mechanically for a range of operating conditions.

  14. Nanostructured Antireflective and Thermoisolative Cicada Wings.

    PubMed

    Morikawa, Junko; Ryu, Meguya; Seniutinas, Gediminas; Balčytis, Armandas; Maximova, Ksenia; Wang, Xuewen; Zamengo, Massimiliano; Ivanova, Elena P; Juodkazis, Saulius

    2016-05-10

    Inter-related mechanical, thermal, and optical macroscopic properties of biomaterials are defined at the nanoscale by their constituent structures and patterns, which underpin complex functions of an entire bio-object. Here, the temperature diffusivity of a cicada (Cyclochila australasiae) wing with nanotextured surfaces was measured using two complementary techniques: a direct contact method and IR imaging. The 4-6-μm-thick wing section was shown to have a thermal diffusivity of α⊥ = (0.71 ± 0.15) × 10(-7) m(2)/s, as measured by the contact temperature wave method along the thickness of the wing; it corresponds to the inherent thermal property of the cuticle. The in-plane thermal diffusivity value of the wing was determined by IR imaging and was considerably larger at α∥ = (3.6 ± 0.2) × 10(-7) m(2)/s as a result of heat transport via air. Optical properties of wings covered with nanospikes were numerically simulated using an accurate 3D model of the wing pattern and showed that light is concentrated between spikes where intensity is enhanced by up to 3- to 4-fold. The closely packed pattern of nanospikes reduces the reflectivity of the wing throughout the visible light spectrum and over a wide range of incident angles, hence acting as an antireflection coating. PMID:27101865

  15. Integrated technology wing design study

    NASA Technical Reports Server (NTRS)

    Hays, A. P.; Beck, W. E.; Morita, W. H.; Penrose, B. J.; Skarshaug, R. E.; Wainfan, B. S.

    1984-01-01

    The technology development costs and associated benefits in applying advanced technology associated with the design of a new wing for a new or derivative trijet with a capacity for 350 passengers and maximum range of 8519 km, entering service in 1990 were studied. The areas of technology are: (1) airfoil technology; (2) planform parameters; (3) high lift; (4) pitch active control system; (5) all electric systems; (6) E to 3rd power propulsion; (7) airframe/propulsion integration; (8) graphite/epoxy composites; (9) advanced aluminum alloys; (10) titanium alloys; and (11) silicon carbide/aluminum composites. These technologies were applied to the reference aircraft configuration. Payoffs were determined for block fuel reductions and net value of technology. These technologies are ranked for the ratio of net value of technology (NVT) to technology development costs.

  16. Artificial delta growth

    NASA Astrophysics Data System (ADS)

    Mikeš, Daniel

    2010-05-01

    A deltaic sedimentary system has a point source; sediment is carried over the delta plain by distributary channels away from the point source and deposited at the delta front by distributary mouth bars. The established methods to describe such a sedimentary system are "bedding analysis", "facies analysis", and "basin analysis". We shall call the ambient conditions "input" and the rock record "output". There exist a number of methods to deduce input from output, e.g. "Sequence stratigraphy" (a.o. Vail et al. 1977, Catuneanu et al. 2009), "Shoreline trajectory" (a.o. Helland-Hansen & Martinsen 1996, Helland-Hansen & Hampson 2009) on the one hand and the complex use of established techniques on the other (a.o. Miall & Miall 2001, Miall & Miall 2002). None of these deductive methods seems to be sufficient. I claim that the common errors in all these attempts are the following: (1) a sedimentary system is four-dimensional (3+1) and a lesser dimensional analysis is insufficient; (2) a sedimentary system is complex and any empirical/deductive analysis is non-unique. The proper approach to the problem is therefore the theoretical/inductive analysis. To that end we performed six scenarios of a scaled version of a passive margin delta in a flume tank. The scenarios have identical stepwise tectonic subsidence and semi-cyclic sealevel, but different supply curves, i.e. supply is: constant, highly-frequent, proportional to sealevel, inversely proportional to sealevel, lagging to sealevel, ahead of sealevel. The preliminary results are indicative. Lobe-switching occurs frequently and hence locally sedimentation occurs shortly and hiatuses are substantial; therefore events in 2D (+1) cross-sections don't correlate temporally. The number of sedimentary cycles disequals the number of sealevel cycles. Lobe-switching and stepwise tectonic subsidence cause onlap/transgression. Erosional unconformities are local diachronous events, whereas maximum flooding surfaces are regional

  17. Delta spots and great flares

    NASA Technical Reports Server (NTRS)

    Zirin, Harold; Liggett, Margaret A.

    1987-01-01

    The development of delta spots and the great flares they produce are reviewed based on 18 years of observations. Delta groups are found to develop in three ways: (1) by the eruption of a single complex active region formed below the surface; (2) by the eruption of large satellite spots near a large older spot; and (3) by the collision of spots of opposite polarity from different dipoles. It is shown that the present sample of 21 delta spots never separate once they lock together, and that the driving force for the shear is spot motion. Indicators for the prediction of the occurrence of great flares are identified.

  18. Novel Control Effectors for Truss Braced Wing

    NASA Technical Reports Server (NTRS)

    White, Edward V.; Kapania, Rakesh K.; Joshi, Shiv

    2015-01-01

    At cruise flight conditions very high aspect ratio/low sweep truss braced wings (TBW) may be subject to design requirements that distinguish them from more highly swept cantilevered wings. High aspect ratio, short chord length and relative thinness of the airfoil sections all contribute to relatively low wing torsional stiffness. This may lead to aeroelastic issues such as aileron reversal and low flutter margins. In order to counteract these issues, high aspect ratio/low sweep wings may need to carry additional high speed control effectors to operate when outboard ailerons are in reversal and/or must carry additional structural weight to enhance torsional stiffness. The novel control effector evaluated in this study is a variable sweep raked wing tip with an aileron control surface. Forward sweep of the tip allows the aileron to align closely with the torsional axis of the wing and operate in a conventional fashion. Aft sweep of the tip creates a large moment arm from the aileron to the wing torsional axis greatly enhancing aileron reversal. The novelty comes from using this enhanced and controllable aileron reversal effect to provide roll control authority by acting as a servo tab and providing roll control through intentional twist of the wing. In this case the reduced torsional stiffness of the wing becomes an advantage to be exploited. The study results show that the novel control effector concept does provide roll control as described, but only for a restricted class of TBW aircraft configurations. For the configuration studied (long range, dual aisle, Mach 0.85 cruise) the novel control effector provides significant benefits including up to 12% reduction in fuel burn.

  19. 14 CFR 23.697 - Wing flap controls.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Wing flap controls. 23.697 Section 23.697... Systems § 23.697 Wing flap controls. (a) Each wing flap control must be designed so that, when the flap... with § 23.145(b)(3) necessitates wing flap retraction to positions that are not fully retracted,...

  20. 14 CFR 23.697 - Wing flap controls.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Wing flap controls. 23.697 Section 23.697... Systems § 23.697 Wing flap controls. (a) Each wing flap control must be designed so that, when the flap... with § 23.145(b)(3) necessitates wing flap retraction to positions that are not fully retracted,...

  1. 14 CFR 23.697 - Wing flap controls.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Wing flap controls. 23.697 Section 23.697... Systems § 23.697 Wing flap controls. (a) Each wing flap control must be designed so that, when the flap... with § 23.145(b)(3) necessitates wing flap retraction to positions that are not fully retracted,...

  2. 14 CFR 23.697 - Wing flap controls.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Wing flap controls. 23.697 Section 23.697... Systems § 23.697 Wing flap controls. (a) Each wing flap control must be designed so that, when the flap... with § 23.145(b)(3) necessitates wing flap retraction to positions that are not fully retracted,...

  3. The plane problem of the flapping wing

    NASA Technical Reports Server (NTRS)

    Birnbaum, Walter

    1954-01-01

    In connection with an earlier report on the lifting vortex sheet which forms the basis of the following investigations this will show how the methods developed there are also suitable for dealing with the air forces for a wing with a circulation variable with time. The theory of a propulsive wing flapping up and down periodically in the manner of a bird's wing is developed. This study shows how the lift and its moment result as a function of the flapping motion, what thrust is attainable, and how high is the degree of efficiency of this flapping propulsion unit if the air friction is disregarded.

  4. Sensitivity Analysis of Wing Aeroelastic Responses

    NASA Technical Reports Server (NTRS)

    Issac, Jason Cherian

    1995-01-01

    Design for prevention of aeroelastic instability (that is, the critical speeds leading to aeroelastic instability lie outside the operating range) is an integral part of the wing design process. Availability of the sensitivity derivatives of the various critical speeds with respect to shape parameters of the wing could be very useful to a designer in the initial design phase, when several design changes are made and the shape of the final configuration is not yet frozen. These derivatives are also indispensable for a gradient-based optimization with aeroelastic constraints. In this study, flutter characteristic of a typical section in subsonic compressible flow is examined using a state-space unsteady aerodynamic representation. The sensitivity of the flutter speed of the typical section with respect to its mass and stiffness parameters, namely, mass ratio, static unbalance, radius of gyration, bending frequency, and torsional frequency is calculated analytically. A strip theory formulation is newly developed to represent the unsteady aerodynamic forces on a wing. This is coupled with an equivalent plate structural model and solved as an eigenvalue problem to determine the critical speed of the wing. Flutter analysis of the wing is also carried out using a lifting-surface subsonic kernel function aerodynamic theory (FAST) and an equivalent plate structural model. Finite element modeling of the wing is done using NASTRAN so that wing structures made of spars and ribs and top and bottom wing skins could be analyzed. The free vibration modes of the wing obtained from NASTRAN are input into FAST to compute the flutter speed. An equivalent plate model which incorporates first-order shear deformation theory is then examined so it can be used to model thick wings, where shear deformations are important. The sensitivity of natural frequencies to changes in shape parameters is obtained using ADIFOR. A simple optimization effort is made towards obtaining a minimum weight

  5. Moveable Leading Edge Device for a Wing

    NASA Technical Reports Server (NTRS)

    Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

    2013-01-01

    A method and apparatus for managing a flight control surface system. A leading edge section on a wing of an aircraft is extended into a deployed position. A deformable section connects the leading edge section to a trailing section. The deformable section changes from a deformed shape to an original shape when the leading edge section is moved into the deployed position. The leading edge section on the wing is moved from the deployed position to an undeployed position. The deformable section changes to the deformed shape inside of the wing.

  6. A Miniature Controllable Flapping Wing Robot

    NASA Astrophysics Data System (ADS)

    Arabagi, Veaceslav Gheorghe

    The agility and miniature size of nature's flapping wing fliers has long baffled researchers, inspiring biological studies, aerodynamic simulations, and attempts to engineer their robotic replicas. Flapping wing flight is characterized by complex reciprocating wing kinematics, transient aerodynamic effects, and very small body lengths. These characteristics render robotic flapping wing aerial vehicles ideal for surveillance and defense applications, search and rescue missions, and environment monitoring, where their ability to hover and high maneuverability is immensely beneficial. One of the many difficulties in creating flapping wing based miniature robotic aerial vehicles lies in generating a proper wing trajectory that would result in sufficient lift forces for hovering and maneuvering. Since design of a flapping wing system is a balance between overall weight and the number of actuated inputs, we take the approach of having minimal controlled inputs, allowing passive behavior wherever possible. Hence, we propose a completely passive wing pitch reversal design that relies on wing inertial dynamics, an elastic energy storage mechanism, and low Reynolds number aerodynamic effects. Theoretical models, compiling previous research on piezoelectric actuators, four-bar transmissions, and aerodynamics effects, are developed and used as basis for a complete numerical simulation. Limitations of the model are discussed in comparison to experimental results obtained from a working prototype of the proposed passive pitch reversal flapping wing mechanism. Given that the mechanism is under-actuated, methods to control lift force generation by actively varying system parameters are proposed, discussed, and tested experimentally. A dual wing aerial platform is developed based on the passive pitch reversal wing concept. Design considerations are presented, favoring controllability and structural rigidity of the final platform. Finite element analysis and experimental

  7. Aerodynamic shape optimization of wing and wing-body configurations using control theory

    NASA Technical Reports Server (NTRS)

    Reuther, James; Jameson, Antony

    1995-01-01

    This paper describes the implementation of optimization techniques based on control theory for wing and wing-body design. In previous studies it was shown that control theory could be used to devise an effective optimization procedure for airfoils and wings in which the shape and the surrounding body-fitted mesh are both generated analytically, and the control is the mapping function. Recently, the method has been implemented for both potential flows and flows governed by the Euler equations using an alternative formulation which employs numerically generated grids, so that it can more easily be extended to treat general configurations. Here results are presented both for the optimization of a swept wing using an analytic mapping, and for the optimization of wing and wing-body configurations using a general mesh.

  8. The effect of over-the-wing nacelles on wing-body aerodynamics

    NASA Technical Reports Server (NTRS)

    Reubush, D. E.

    1978-01-01

    An investigation was conducted in the Langley 16-foot transonic tunnel to further study benefits in climb and cruise performance due to blowing the jet over the wing for a transport-type wing-body configuration. In this investigation a wing-body model/powered-nacelle test rig combination was tested at Mach numbers of 0.5 and 0.8 at angles of attack from -2 to 4 deg and jet total-pressure ratios from jet off to 3 or 4 (depending on Mach number) for a variety of nacelle locations relative to the wing. Results from this investigation show that positioning of the nacelles can have very large effects on the wing-body drag (nacelles were nonmetric). Some positions yielded much higher drag than the baseline wing-body while others yielded drag which was somewhat lower than the baseline.

  9. Thin tailored composite wing for civil tiltrotor

    NASA Technical Reports Server (NTRS)

    Rais-Rohani, Masoud

    1994-01-01

    The tiltrotor aircraft is a flight vehicle which combines the efficient low speed (i.e., take-off, landing, and hover) characteristics of a helicopter with the efficient cruise speed of a turboprop airplane. A well-known example of such vehicle is the Bell-Boeing V-22 Osprey. The high cruise speed and range constraints placed on the civil tiltrotor require a relatively thin wing to increase the drag-divergence Mach number which translates into lower compressibility drag. It is required to reduce the wing maximum thickness-to-chord ratio t/c from 23% (i.e., V-22 wing) to 18%. While a reduction in wing thickness results in improved aerodynamic efficiency, it has an adverse effect on the wing structure and it tends to reduce structural stiffness. If ignored, the reduction in wing stiffness leads to susceptibility to aeroelastic and dynamic instabilities which may consequently cause a catastrophic failure. By taking advantage of the directional stiffness characteristics of composite materials the wing structure may be tailored to have the necessary stiffness, at a lower thickness, while keeping the weight low. The goal of this study is to design a wing structure for minimum weight subject to structural, dynamic and aeroelastic constraints. The structural constraints are in terms of strength and buckling allowables. The dynamic constraints are in terms of wing natural frequencies in vertical and horizontal bending and torsion. The aeroelastic constraints are in terms of frequency placement of the wing structure relative to those of the rotor system. The wing-rotor-pylon aeroelastic and dynamic interactions are limited in this design study by holding the cruise speed, rotor-pylon system, and wing geometric attributes fixed. To assure that the wing-rotor stability margins are maintained a more rigorous analysis based on a detailed model of the rotor system will need to ensue following the design study. The skin-stringer-rib type architecture is used for the wing

  10. Active Dihedral Control System for a Torisionally Flexible Wing

    NASA Technical Reports Server (NTRS)

    Kendall, Greg T. (Inventor); Lisoski, Derek L. (Inventor); Morgan, Walter R. (Inventor); Griecci, John A. (Inventor)

    2015-01-01

    A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

  11. On the Minimum Induced Drag of Wings

    NASA Technical Reports Server (NTRS)

    Bowers, Albion H.

    2007-01-01

    This viewgraph presentation reviews the minimum induced drag of wings. The topics include: 1) The History of Spanload Development of the optimum spanload Winglets and their implications; 2) Horten Sailplanes; and 3) Flight Mechanics & Adverse yaw.

  12. Left-Wing Extremism: The Current Threat

    SciTech Connect

    Karl A. Seger

    2001-04-30

    Left-wing extremism is ''alive and well'' both in the US and internationally. Although the current domestic terrorist threat within the U. S. is focused on right-wing extremists, left-wing extremists are also active and have several objectives. Leftist extremists also pose an espionage threat to U.S. interests. While the threat to the U.S. government from leftist extremists has decreased in the past decade, it has not disappeared. There are individuals and organizations within the U.S. who maintain the same ideology that resulted in the growth of left-wing terrorism in this country in the 1970s and 1980s. Some of the leaders from that era are still communicating from Cuba with their followers in the U.S., and new leaders and groups are emerging.

  13. Coriolis effects enhance lift on revolving wings.

    PubMed

    Jardin, T; David, L

    2015-03-01

    At high angles of attack, an aircraft wing stalls. This dreaded event is characterized by the development of a leading edge vortex on the upper surface of the wing, followed by its shedding which causes a drastic drop in the aerodynamic lift. At similar angles of attack, the leading edge vortex on an insect wing or an autorotating seed membrane remains robustly attached, ensuring high sustained lift. What are the mechanisms responsible for both leading edge vortex attachment and high lift generation on revolving wings? We review the three main hypotheses that attempt to explain this specificity and, using direct numerical simulations of the Navier-Stokes equations, we show that the latter originates in Coriolis effects. PMID:25871040

  14. Measurements of Supersonic Wing Tip Vortices

    NASA Technical Reports Server (NTRS)

    Smart, Michael K.; Kalkhoran, Iraj M.; Benston, James

    1994-01-01

    An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.

  15. Territoriality in the Red-winged Blackbird

    ERIC Educational Resources Information Center

    Newhouse, Chris

    1977-01-01

    Reports findings on research in Red-winged Blackbird territoriality and describes the educational potential of use of similar studies in the classroom. Territorial mapping and observational techniques are explained. (CS)

  16. Wing planform optimization via an adjoint method

    NASA Astrophysics Data System (ADS)

    Leoviriyakit, Kasidit

    This dissertation focuses on the problem of wing planform optimization for transonic aircraft based on flow simulation using Computational Fluid Dynamics (CFD) combined with an adjoint-gradient based numerical optimization procedure. The adjoint method, traditionally used for wing section design has been extended to cover planform variations and to compute the sensitivities of the structural weight of both the wing section and planform variations. The two relevant disciplines accounted for are the aerodynamics and structural weight. A simplified structural weight model is used for the optimization. Results of a variety of long range transports indicate that significant improvement in both aerodynamics and structures can be achieved simultaneously. The proof-of-concept optimal results indicate large improvements for both drag and structural weight. The work is an "enabling step" towards a realistic automated wing designed by a computer.

  17. The Design of Airplane Wing Ribs

    NASA Technical Reports Server (NTRS)

    Newlin, J A; Trayer, George W

    1931-01-01

    The purpose of this investigation was to obtain information for use in the design of truss and plywood forms, particularly with reference to wing ribs. Tests were made on many designs of wing ribs, comparing different types in various sizes. Many tests were also made on parallel-chord specimens of truss and plywood forms in place of the actual ribs and on parts of wing ribs, such as truss diagonals and sections of cap strips. It was found that for ribs of any size or proportions, when they were designed to obtain a well-balanced construction and were carefully manufactured, distinct types are of various efficiencies; the efficiency is based on the strength per unit of weight. In all types of ribs the heavier are the stronger per unit of weight. Reductions in the weight of wing ribs are accompanied even in efficient designs by a much greater proportional reduction in strength.

  18. Coriolis effects enhance lift on revolving wings

    NASA Astrophysics Data System (ADS)

    Jardin, T.; David, L.

    2015-03-01

    At high angles of attack, an aircraft wing stalls. This dreaded event is characterized by the development of a leading edge vortex on the upper surface of the wing, followed by its shedding which causes a drastic drop in the aerodynamic lift. At similar angles of attack, the leading edge vortex on an insect wing or an autorotating seed membrane remains robustly attached, ensuring high sustained lift. What are the mechanisms responsible for both leading edge vortex attachment and high lift generation on revolving wings? We review the three main hypotheses that attempt to explain this specificity and, using direct numerical simulations of the Navier-Stokes equations, we show that the latter originates in Coriolis effects.

  19. Mallard age and sex determination from wings

    USGS Publications Warehouse

    Carney, S.M.; Geis, A.D.

    1960-01-01

    This paper describes characters on the wing plumage of the mallard that indicate age and sex. A key outlines a logical order in which to check age and sex characters on wings. This method was tested and found to be more than 95 percent reliable, although it was found that considerable practice and training with known-age specimens was required to achieve this level of accuracy....The implications of this technique and the sampling procedure it permits are discussed. Wing collections could provide information on production, and, if coupled with a banding program could permit seasonal population estimates to be calculated. In addition, representative samples of wings would provide data to check the reliability of several other waterfowl surveys.

  20. Improvement of Aerodynamic Performance of the Aero-Train by Controlling Wing-Wing Interaction Using Single-Slotted Flap

    NASA Astrophysics Data System (ADS)

    Yoon, Dong-Hee; Kohama, Yasuaki; Kikuchi, Satoshi; Kato, Takuma

    Aero-train is a new driving concept using aerodynamic technology under development by the Kohama Laboratory, Institute of Fluid Science, Tohoku University. It employs the wing-in-ground effect to enable travel at high speeds over land. Aero-train makes use of the ground effects of lift and side force between the wings and a U-shaped guideway for stability. The main wings have vertical wings at the tips, which are arranged in tandem to regulate the roll and yaw stability in the U-shaped guideway. However, the vertical wings deteriorate the lift-to-drag ratio of the Aero-train by aerodynamic interaction with the main wings. The present study was performed to improve the aerodynamic performance of the Aero-train by controlling wing-wing interaction. Installation of a single-slotted flap on the wings considerably improved the aerodynamic performance of the wings.

  1. Butterfly wing color: A photonic crystal demonstration

    NASA Astrophysics Data System (ADS)

    Proietti Zaccaria, Remo

    2016-01-01

    We have theoretically modeled the optical behavior of a natural occurring photonic crystal, as defined by the geometrical characteristics of the Teinopalpus Imperialis butterfly. In particular, following a genetic algorithm approach, we demonstrate how its wings follow a triclinic crystal geometry with a tetrahedron unit base. By performing both photonic band analysis and transmission/reflection simulations, we are able to explain the characteristic colors emerging by the butterfly wings, thus confirming their crystal form.

  2. Numerical simulation of swept-wing flows

    NASA Technical Reports Server (NTRS)

    Reed, Helen L.

    1991-01-01

    The transition process characteristics of flows over swept wings were computationally modelled. The crossflow instability and crossflow/T-S wave interaction are analyzed through the numerical solution of the full three dimensional Navier-Stokes equations including unsteadiness, curvature, and sweep. The leading-edge region of a swept wing is considered in a three-dimensional spatial simulation with random disturbances as the initial conditions.

  3. Experiments with suction-type wings

    NASA Technical Reports Server (NTRS)

    Schrenk, O

    1935-01-01

    The present report collects the investigations of the past years which, while not as yet intended for use in construction, show different possibilities for the building of a suction-type wing and at the same time present some basic explanations concerning the problem of suction. Experiments and results with a thick wing profile are detailed as well as boundary layer removal by suction and sink action. Experiments with flap profiles are also included.

  4. The oscillating wing with aerodynamically balanced elevator

    NASA Technical Reports Server (NTRS)

    Kussner, H G; Schwartz, I

    1941-01-01

    The two-dimensional problem of the oscillating wing with aerodynamically balanced elevator is treated in the manner that the wing is replaced by a plate with bends and stages and the airfoil section by a mean line consisting of one or more straights. The computed formulas and tables permit, on these premises, the prediction of the pressure distribution and of the aerodynamic reactions of oscillating elevators and tabs with any position of elevator hinge in respect to elevator leading edge.

  5. Selection on Wing Allometry in Drosophila Melanogaster

    PubMed Central

    Weber, K. E.

    1990-01-01

    Five bivariate distributions of wing dimensions of Drosophila melanogaster were measured, in flies 1) subjected to four defined environmental regimes during development, 2) taken directly from nature in seven U.S. states, 3) selected in ten populations for change in wing form, and 4) sampled from 21 long inbred wild-type lines. Environmental stresses during development altered both wing size and the ratios of wing dimensions, but regardless of treatment all wing dimensions fell near a common allometric baseline in each bivariate distribution. The wings of wild-caught flies from seven widely separated localities, and of their laboratory-reared offspring, also fell along the same baselines. However, when flies were selected divergently for lateral offset from these developmental baselines, response to selection was rapid in every case. The mean divergence in offset between oppositely selected lines was 14.68 SD of the base population offset, after only 15 generations of selection at 20%. Measurements of 21 isofemale lines, founded from wild-caught flies and maintained in small populations for at least 22 years, showed large reductions in phenotypic variance of offsets within lines, but a large increase in the variance among lines. The variance of means of isofemale lines within collection localities was ten times the variance of means among localities of newly established wild lines. These observations show that much additive genetic variance exists for individual dimensions within the wing, such that bivariate developmental patterns can be changed in any direction by selection or by drift. The relative invariance of the allometric baselines of wing morphology in nature is most easily explained as the result of continuous natural selection around a local optimum of functional design. PMID:2127580

  6. Integrated technology wing study (oral presentation)

    NASA Technical Reports Server (NTRS)

    1981-01-01

    The design of a plan for a commercial transport manufacturer to integrate advanced technology into a new wing for a derivative and/or new aircraft that could enter service in the late 1980s to early 1990s time period is proposed. The development of a new wing for a derivative or a new long range commercial aircraft and the incorporation of cost effective technologies are studied. The decision provides guidelines for the best allocation of research funds.

  7. Calculations Of Transonic Flow About A Wing

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.; Gundy, Karen L.; Flores, Jolen; Chaderjian, Neal; Kaynak, Univer; Thomas, Scott D.

    1988-01-01

    Report describes calculations of transonic airflows about wing in wind tunnel. Basic equations of flow used in study are Reynolds-averaged Navier-Stokes equations in strong conservation-law form. Equations of flow incorporated into finite-difference computer code called TNS (Transonic Navier-Stokes). Computational grid generated by solution of partial differential equations yielding smooth meshes conforming to surfaces of wing and wind tunnel.

  8. Determination of the characteristics of tapered wings

    NASA Technical Reports Server (NTRS)

    Anderson, Raymond F

    1937-01-01

    This report presents tables and charts for use in determining the characteristics of tapered wings. Theoretical factors are given from which the following characteristics of tapered wings may be found: the span lift distribution, the induced-angle-of attack distribution, the lift-curve slope, the angle of zero lift, the induced drag, the aerodynamic-center position, and the pitching moment about the aerodynamic center.

  9. Modal control of an oblique wing aircraft

    NASA Technical Reports Server (NTRS)

    Phillips, James D.

    1989-01-01

    A linear modal control algorithm is applied to the NASA Oblique Wing Research Aircraft (OWRA). The control law is evaluated using a detailed nonlinear flight simulation. It is shown that the modal control law attenuates the coupling and nonlinear aerodynamics of the oblique wing and remains stable during control saturation caused by large command inputs or large external disturbances. The technique controls each natural mode independently allowing single-input/single-output techniques to be applied to multiple-input/multiple-output systems.

  10. Aircraft noise propagation. [sound diffraction by wings

    NASA Technical Reports Server (NTRS)

    Hadden, W. J.; Pierce, A. D.

    1978-01-01

    Sound diffraction experiments conducted at NASA Langley Research Center to study the acoustical implications of the engine over wing configuration (noise-shielding by wing) and to provide a data base for assessing various theoretical approaches to the problem of aircraft noise reduction are described. Topics explored include the theory of sound diffraction around screens and wedges; the scattering of spherical waves by rectangular patches; plane wave diffraction by a wedge with finite impedence; and the effects of ambient flow and distribution sources.

  11. Numerical study of the trailing vortex of a wing with wing-tip blowing

    NASA Technical Reports Server (NTRS)

    Lim, Hock-Bin

    1994-01-01

    Trailing vortices generated by lifting surfaces such as helicopter rotor blades, ship propellers, fixed wings, and canard control surfaces are known to be the source of noise, vibration, cavitation, degradation of performance, and other hazardous problems. Controlling these vortices is, therefore, of practical interest. The formation and behavior of the trailing vortices are studied in the present research. In addition, wing-tip blowing concepts employing axial blowing and spanwise blowing are studied to determine their effectiveness in controlling these vortices and their effects on the performance of the wing. The 3D, unsteady, thin-layer compressible Navier-Stokes equations are solved using a time-accurate, implicit, finite difference scheme that employs LU-ADI factorization. The wing-tip blowing is simulated using the actuator plane concept, thereby, not requiring resolution of the jet slot geometry. Furthermore, the solution blanking feature of the chimera scheme is used to simplify the parametric study procedure for the wing-tip blowing. Computed results are shown to compare favorably with experimental measurements. It is found that axial wing-tip blowing, although delaying the rolling-up of the trailing vortices and the near-field behavior of the flowfield, does not dissipate the circulation strength of the trailing vortex farther downstream. Spanwise wing-tip blowing has the effect of displacing the trailing vortices outboard and upward. The increased 'wing-span' due to the spanwise wing-tip blowing has the effect of lift augmentation on the wing and the strengthening of the trailing vortices. Secondary trailing vortices are created at high spanwise wing-tip blowing intensities.

  12. Theoretical damping in roll and rolling moment due to differential wing incidence for slender cruciform wings and wing-body combinations

    NASA Technical Reports Server (NTRS)

    Adams, Gaynor J; DUGAN DUANE W

    1952-01-01

    A method of analysis based on slender-wing theory is developed to investigate the characteristics in roll of slender cruciform wings and wing-body combinations. The method makes use of the conformal mapping processes of classical hydrodynamics which transform the region outside a circle and the region outside an arbitrary arrangement of line segments intersecting at the origin. The method of analysis may be utilized to solve other slender cruciform wing-body problems involving arbitrarily assigned boundary conditions. (author)

  13. Nonlinear Aerodynamics and the Design of Wing Tips

    NASA Technical Reports Server (NTRS)

    Kroo, Ilan

    1991-01-01

    The analysis and design of wing tips for fixed wing and rotary wing aircraft still remains part art, part science. Although the design of airfoil sections and basic planform geometry is well developed, the tip regions require more detailed consideration. This is important because of the strong impact of wing tip flow on wing drag; although the tip region constitutes a small portion of the wing, its effect on the drag can be significant. The induced drag of a wing is, for a given lift and speed, inversely proportional to the square of the wing span. Concepts are proposed as a means of reducing drag. Modern computational methods provide a tool for studying these issues in greater detail. The purpose of the current research program is to improve the understanding of the fundamental issues involved in the design of wing tips and to develop the range of computational and experimental tools needed for further study of these ideas.

  14. Imaging and Laser Spectroscopy Investigation of Insect Wings

    NASA Astrophysics Data System (ADS)

    Shiver, Tegan; Lawhead, Carlos; Anderson, Josiah; Cooper, Nathan; Ujj, Laszlo; Pall Life Sciences Collaboration

    2014-03-01

    Measuring the surface morphology and chemical composition of insect wings is important to understand the extreme mechanical properties and the biophysical functionalities of the wings. We have measured the image of the membrane of the cicada (genus Tibicen) wing with the help of Scanning Electron Microscopy (SEM). The results confirm the existing periodic structure of the wing measured previously. The SEM imaging can be used to measure the surface morphology of any insect species wings. The physical surface structure of the cicada wing is an example of a new class of biomaterials that can kill bacteria on contact. In order to identify the chemical composition of the wing, we have measured the vibrational spectra of the wing's membrane (Raman and CARS). The measured spectra are consistent with the original assumption that the wing membrane is composed of protein, wax, and chitin. The results of these studies can be used to make artificial materials in the future.

  15. Functional Gustatory Role of Chemoreceptors in Drosophila Wings.

    PubMed

    Raad, Hussein; Ferveur, Jean-François; Ledger, Neil; Capovilla, Maria; Robichon, Alain

    2016-05-17

    Neuroanatomical evidence argues for the presence of taste sensilla in Drosophila wings; however, the taste physiology of insect wings remains hypothetical, and a comprehensive link to mechanical functions, such as flight, wing flapping, and grooming, is lacking. Our data show that the sensilla of the Drosophila anterior wing margin respond to both sweet and bitter molecules through an increase in cytosolic Ca(2+) levels. Conversely, genetically modified flies presenting a wing-specific reduction in chemosensory cells show severe defects in both wing taste signaling and the exploratory guidance associated with chemodetection. In Drosophila, the chemodetection machinery includes mechanical grooming, which facilitates the contact between tastants and wing chemoreceptors, and the vibrations of flapping wings that nebulize volatile molecules as carboxylic acids. Together, these data demonstrate that the Drosophila wing chemosensory sensilla are a functional taste organ and that they may have a role in the exploration of ecological niches. PMID:27160896

  16. Effect of wing flexibility in dragonfly hovering flight

    NASA Astrophysics Data System (ADS)

    Naidu, Vishal; Young, John; Lai, Joseph

    2011-11-01

    Dragonflies have two pairs of tandem wings, which can be operated independently. Most studies on tandem wings are based on rigid wings, which is in strong contradiction to the natural, flexible dragonfly wings. The effect of wing flexibility in tandem wings is little known. We carry out a comparative, computational study between rigid and flexible, dragonfly shaped wings for hovering flight. In rigid wings during downstroke, a leading edge vortex (LEV) is formed on the upper surface, which forms a low pressure zone. This conical LEV joins the tip vortex and shortly after the mid downstroke when the wing starts to rotate, these vortices are gradually shed resulting in a drop in lift. The vortex system creates a net downwards momentum in the form of a jet. The flexible wings while in motion deform due to aerodynamic and inertial forces. Since there is a strong interaction between wing deformation and air flow around the deformed wings, flexible wing simulations are carried out using a two way fluid structure interaction. The effect of wing flexibility on the flow structure and the subsequent effect on the aerodynamic forces will be studied and presented.

  17. Numerical solutions of Navier-Stokes equations for a Butler wing

    NASA Technical Reports Server (NTRS)

    Abolhassani, J. S.; Tiwari, S. N.

    1985-01-01

    The flow field is simulated on the surface of a given delta wing (Butler wing) at zero incident in a uniform stream. The simulation is done by integrating a set of flow field equations. This set of equations governs the unsteady, viscous, compressible, heat conducting flow of an ideal gas. The equations are written in curvilinear coordinates so that the wing surface is represented accurately. These equations are solved by the finite difference method, and results obtained for high-speed freestream conditions are compared with theoretical and experimental results. In this study, the Navier-Stokes equations are solved numerically. These equations are unsteady, compressible, viscous, and three-dimensional without neglecting any terms. The time dependency of the governing equations allows the solution to progress naturally for an arbitrary initial initial guess to an asymptotic steady state, if one exists. The equations are transformed from physical coordinates to the computational coordinates, allowing the solution of the governing equations in a rectangular parallel-piped domain. The equations are solved by the MacCormack time-split technique which is vectorized and programmed to run on the CDC VPS 32 computer.

  18. The DELTA Synchrotron Light Interferometer

    SciTech Connect

    Berges, U.

    2004-05-12

    Synchrotron radiation sources like DELTA, the Dortmund Electron Accelerator, a third generation synchrotron light source, need an optical monitoring system to measure the beam size at different points of the ring with high resolution and accuracy. These measurements also allow an investigation of the emittance of the storage ring, an important working parameter for the efficiency of working beamlines with experiments using the synchrotron radiation. The resolution limits of the different types of optical synchrotron light monitors at DELTA are investigated. The minimum measurable beamsize with the normal synchrotron light monitor using visible light at DELTA is about 80 {mu}m. Due to this a synchrotron light interferometer was built up and tested at DELTA. The interferometer uses the same beamline in the visible range. The minimum measurable beamsize is with about 8 {mu}m one order of magnitude smaller. This resolution is sufficient for the expected small vertical beamsizes at DELTA. The electron beamsize and emittance were measured with both systems at different electron beam energies of the storage ring. The theoretical values of the present optics are smaller than the measured emittance. So possible reasons for beam movements are investigated.

  19. An Investigation of the Effects of Sweep on the Characteristics of a High-Aspect-Ratio Wing in the Langley 8-Foot High-Speed Tunnel

    NASA Technical Reports Server (NTRS)

    Whitcomb, Richard T.

    1947-01-01

    An untwisted wing, which when unswept has an NACA 65-210 section, an aspect ratio of 9.0 and a taper ration of 2.5:1.0, has been tested with no sweep, and 30 deg and 45 deg of sweepback and sweepforward in conjunction with a typical fuselage at Mach numbers from 0.60 to 0.96 at angles of attack generally between -2 deg and 10 deg in the Langley 8-foot high-speed tunnel. Sweep was obtained by rotating the wing semispans about a point in the plane of symmetry. The normal-force, pitching-moment, profile-drag, and loading characteristics for the wings have been obtained from pressure measurements and wake surveys. The results indicate that the wings with +/-30 deg of sweep experienced the severe changes in characteristics associated with the presence of a shock at higher Mach numbers then did the wing without sweep. The differences between the Mach numbers at which the changes occurred for the wings with +/-30 deg sweep and no sweep were generally slightly less than the factor 1/cosDelta(sub r) times the Mach numbers at which the changes occurred for the unswept wing, Delta(sub r) being the sweep angle. The wings with +/-45 deg of sweep did not experience the changes in the characteristics associated with the presence of shock at an angle of attack of 2 deg at Mach numbers up to the highest test value. The magnitudes of changes in the normal-force and pitching-moment coefficients that occurred were less for the wing with 30 deg of sweep than for the unswept wing. The use of sweepforward was superior to sweepback in delaying and reducing the changes in the normal-force coefficients, but was inferior in delaying and reducing the changes in the profile-drag coefficients. Increasing the Mach number to the highest test values had little effect on the positions of the center of loads on the various configurations for the probable design load conditions.

  20. Wing Torsional Stiffness Tests of the Active Aeroelastic Wing F/A-18 Airplane

    NASA Technical Reports Server (NTRS)

    Lokos, William A.; Olney, Candida D.; Crawford, Natalie D.; Stauf, Rick; Reichenbach, Eric Y.

    2002-01-01

    The left wing of the Active Aeroelastic Wing (AAW) F/A-18 airplane has been ground-load-tested to quantify its torsional stiffness. The test has been performed at the NASA Dryden Flight Research Center in November 1996, and again in April 2001 after a wing skin modification was performed. The primary objectives of these tests were to characterize the wing behavior before the first flight, and provide a before-and-after measurement of the torsional stiffness. Two streamwise load couples have been applied. The wing skin modification is shown to have more torsional flexibility than the original configuration has. Additionally, structural hysteresis is shown to be reduced by the skin modification. Data comparisons show good repeatability between the tests.

  1. Experimental investigation of a flapping wing model

    NASA Astrophysics Data System (ADS)

    Hubel, Tatjana Y.; Tropea, Cameron

    2009-05-01

    The main objective of this research study was to investigate the aerodynamic forces of an avian flapping wing model system. The model size and the flow conditions were chosen to approximate the flight of a goose. Direct force measurements, using a three-component balance, and PIV flow field measurements parallel and perpendicular to the oncoming flow, were performed in a wind tunnel at Reynolds numbers between 28,000 and 141,000 (3-15 m/s), throughout a range of reduced frequencies between 0.04 and 0.20. The appropriateness of quasi-steady assumptions used to compare 2D, time-averaged particle image velocimetry (PIV) measurements in the wake with direct force measurements was evaluated. The vertical force coefficient for flapping wings was typically significantly higher than the maximum coefficient of the fixed wing, implying the influence of unsteady effects, such as delayed stall, even at low reduced frequencies. This puts the validity of the quasi-steady assumption into question. The (local) change in circulation over the wing beat cycle and the circulation distribution along the wingspan were obtained from the measurements in the tip and transverse vortex planes. Flow separation could be observed in the distribution of the circulation, and while the circulation derived from the wake measurements failed to agree exactly with the absolute value of the circulation, the change in circulation over the wing beat cycle was in excellent agreement for low and moderate reduced frequencies. The comparison between the PIV measurements in the two perpendicular planes and the direct force balance measurements, show that within certain limitations the wake visualization is a powerful tool to gain insight into force generation and the flow behavior on flapping wings over the wing beat cycle.

  2. Experimental investigation of a flapping wing model

    NASA Astrophysics Data System (ADS)

    Hubel, Tatjana Y.; Tropea, Cameron

    The main objective of this research study was to investigate the aerodynamic forces of an avian flapping wing model system. The model size and the flow conditions were chosen to approximate the flight of a goose. Direct force measurements, using a three-component balance, and PIV flow field measurements parallel and perpendicular to the oncoming flow, were performed in a wind tunnel at Reynolds numbers between 28,000 and 141,000 (3-15 m/s), throughout a range of reduced frequencies between 0.04 and 0.20. The appropriateness of quasi-steady assumptions used to compare 2D, time-averaged particle image velocimetry (PIV) measurements in the wake with direct force measurements was evaluated. The vertical force coefficient for flapping wings was typically significantly higher than the maximum coefficient of the fixed wing, implying the influence of unsteady effects, such as delayed stall, even at low reduced frequencies. This puts the validity of the quasi-steady assumption into question. The (local) change in circulation over the wing beat cycle and the circulation distribution along the wingspan were obtained from the measurements in the tip and transverse vortex planes. Flow separation could be observed in the distribution of the circulation, and while the circulation derived from the wake measurements failed to agree exactly with the absolute value of the circulation, the change in circulation over the wing beat cycle was in excellent agreement for low and moderate reduced frequencies. The comparison between the PIV measurements in the two perpendicular planes and the direct force balance measurements, show that within certain limitations the wake visualization is a powerful tool to gain insight into force generation and the flow behavior on flapping wings over the wing beat cycle.

  3. An experimental study of the effect of pitch rate on delta wing aerodynamics and stability

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.

    1988-01-01

    The final report for the research conducted under this grant (NAG1-641) are contained in the two documents attached as Apendices A and B. The first is the presentation made to NASA Langley personnel on 10 December, 1987, which gave a brief analysis of the experiments. The second is a copy of an AIAA paper given in June 1988, which describes in detail the test setup, data acquisition and reduction, and results obtained.

  4. Flow field of flexible flapping wings

    NASA Astrophysics Data System (ADS)

    Sallstrom, Erik

    The agility and maneuverability of natural fliers would be desirable to incorporate into engineered micro air vehicles (MAVs). However, there is still much for engineers to learn about flapping flight in order to understand how such vehicles can be built for efficient flying. The goal of this study is to develop a methodology for capturing high quality flow field data around flexible flapping wings in a hover environment and to interpret it to gain a better understanding of how aerodynamic forces are generated. The flow field data was captured using particle image velocimetry (PIV) and required that measurements be taken around a repeatable flapping motion to obtain phase-averaged data that could be studied throughout the flapping cycle. Therefore, the study includes the development of flapping devices with a simple repeatable single degree of freedom flapping motion. The acquired flow field data has been examined qualitatively and quantitatively to investigate the mechanisms behind force production in hovering flight and to relate it to observations in previous research. Specifically, the flow fields have been investigated around a rigid wing and several carbon fiber reinforced flexible membrane wings. Throughout the whole study the wings were actuated with either a sinusoidal or a semi-linear flapping motion. The semi-linear flapping motion holds the commanded angular velocity nearly constant through half of each half-stroke while the sinusoidal motion is always either accelerating or decelerating. The flow fields were investigated by examining vorticity and vortex structures, using the Q criterion as the definition for the latter, in two and three dimensions. The measurements were combined with wing deflection measurements to demonstrate some of the key links in how the fluid-structure interactions generated aerodynamic forces. The flow fields were also used to calculate the forces generated by the flapping wings using momentum balance methods which yielded

  5. BMI Sandwich Wing Box Analysis and Test

    NASA Technical Reports Server (NTRS)

    Palm, Tod; Mahler, Mary; Shah, Chandu; Rouse, Marshall; Bush, Harold; Wu, Chauncey; Small, William J.

    2000-01-01

    A composite sandwich single bay wing box test article was developed by Northrop Grumman and tested recently at NASA Langley Research Center. The objectives for the wing box development effort were to provide a demonstration article for manufacturing scale up of structural concepts related to a high speed transport wing, and to validate the structural performance of the design. The box concept consisted of highly loaded composite sandwich wing skins, with moderately loaded composite sandwich spars. The dimensions of the box were chosen to represent a single bay of the main wing box, with a spar spacing of 30 inches, height of 20 inches constant depth, and length of 64 inches. The bismaleimide facesheet laminates and titanium honeycomb core chosen for this task are high temperature materials able to sustain a 300F service temperature. The completed test article is shown in Figure 1. The tests at NASA Langley demonstrated the structures ability to sustain axial tension and compression loads in excess of 20,000 lb/in, and to maintain integrity in the thermal environment. Test procedures, analysis failure predictions, and test results are presented.

  6. Mission adaptive wing soars at NASA Facility

    NASA Technical Reports Server (NTRS)

    Rahn, D.; Reinertson, L.

    1986-01-01

    Research pilots have flown the Mission Adaptive Wing (MAW) aircraft, a highly modified F-111 jet fighter, from subsonic speeds up to Mach 1.4 in initial flight tests. The inital test flights are clearing the envelope with the wings flexed at various curvatures. This process allows further research data to be safely gathered so that designers of future variable camber wing aircraft have the best information possible. The altitude envelope was cleared from 27,500 down to 7,500 feet where denser air can cause more stress on the aircraft. Testing with the aircraft was conducted with wing sweep angles of 26 and 58 degrees. At the conclusion of the performance tests in the manual configuration, the system will be reconfigured for automatic mode tests. The limited automatic modes include maneuver camber control where the wings are deflected automatically to the best lift versus drag combination for a particular speed; cruise camber control which can help protect the aircraft from high G stresses; and maneuver enhancement/gust alleviation which is designed to improve the aircraft's up and down movement response to pilot commands and reduce the aircraft response to turbulence.

  7. Flapping wing aerodynamics: from insects to vertebrates.

    PubMed

    Chin, Diana D; Lentink, David

    2016-04-01

    More than a million insects and approximately 11,000 vertebrates utilize flapping wings to fly. However, flapping flight has only been studied in a few of these species, so many challenges remain in understanding this form of locomotion. Five key aerodynamic mechanisms have been identified for insect flight. Among these is the leading edge vortex, which is a convergent solution to avoid stall for insects, bats and birds. The roles of the other mechanisms - added mass, clap and fling, rotational circulation and wing-wake interactions - have not yet been thoroughly studied in the context of vertebrate flight. Further challenges to understanding bat and bird flight are posed by the complex, dynamic wing morphologies of these species and the more turbulent airflow generated by their wings compared with that observed during insect flight. Nevertheless, three dimensionless numbers that combine key flow, morphological and kinematic parameters - the Reynolds number, Rossby number and advance ratio - govern flapping wing aerodynamics for both insects and vertebrates. These numbers can thus be used to organize an integrative framework for studying and comparing animal flapping flight. Here, we provide a roadmap for developing such a framework, highlighting the aerodynamic mechanisms that remain to be quantified and compared across species. Ultimately, incorporating complex flight maneuvers, environmental effects and developmental stages into this framework will also be essential to advancing our understanding of the biomechanics, movement ecology and evolution of animal flight. PMID:27030773

  8. Ring Wing for an underwater missile

    NASA Astrophysics Data System (ADS)

    August, Henry; Carapezza, Edward

    Hughes Aircraft has performed exploratory wind tunnel studies of compressed carriage missile designs having extendable Ring Wing and wrap-around tail control surfaces. These force and moment data indicate that significant improvements in a missile's lift and aerodynamic efficiency can be realized. Low speed test results of these data were used to estimate potential underwater improved hydrodynamic characteristics that a Ring Wing and wrap-around tails can bring to an advanced torpedo design. Estimates of improved underwater flight performance of a heavyweight torpedo (4000 lbs.) having an extendable Ring Wing and wrap-around tails were made. The compressed volume design of this underwater missile is consistent with tube-launch constraints and techniques. Study results of this novel Ring Wing torpedo design include extended flight performance in range and endurance due to lowered speeds capable of sustaining underwater level flight. Correspondingly, reduced radiated noise for enhanced stealth qualities is projected. At high speeds, greater maneuverability and aimpoint selection can be realized by a Ring Wing underwater missile.

  9. A Flight Comparison of Conventional Ailerons on a Rectangular Wing and of Conventional and Floating Wing-Tip Ailerons on a Tapered Wing

    NASA Technical Reports Server (NTRS)

    Soule, H A; Gracey, W

    1938-01-01

    Report presents the results of flight tests comparing the relative effectiveness of conventional ailerons of the same size on wings of rectangular and tapered plan forms made with a Fairchild 22 airplane. Information is included comparing conventional and floating wing-tip ailerons on a tapered wing. The results showed that the conventional ailerons were somewhat more effective on the tapered than on the rectangular wing. The difference, however, was so small as to be imperceptible to the pilots. The floating wing-tip ailerons were only half as effective as the conventional ailerons and, for this reason, were considered unsatisfactory.

  10. Reynolds number effects on the aerodynamic characteristics of irregular planform wings at Mach number 0.3. [in the Ames 12 ft pressure wind tunnel

    NASA Technical Reports Server (NTRS)

    Kruse, R. L.; Lovette, G. H.; Spencer, B., Jr.

    1977-01-01

    The subsonic aerodynamic characteristics of a series of irregular planform wings were studied in wind tunnel tests conducted at M = 0.3 over a range of Reynolds numbers from 1.6 million to 26 million/m. The five basic wing planforms varied from a trapezoidal to a delta shape. Leading edge extensions, added to the basic shape, varied in approximately 5 deg increments from the wing leading edge sweep-back angle to a maximum 80 deg. Most of the tests were conducted using an NACA 0008 airfoil section with grit boundary layer trips. Tests were also conducted using an NACA 0012 airfoil section and an 8% thick wedge. In addition, the effect of free transition (no grit) was investigated. A body was used on all models.

  11. Principle of bio-inspired insect wing rotational hinge design

    NASA Astrophysics Data System (ADS)

    Fei, Fan

    A principle for designing and fabricating bio-inspired miniature artificial insect flapping wing using flexure rotational hinge design is presented. A systematic approach of selecting rotational hinge stiffness value is proposed. Based on the understanding of flapping wing aerodynamics, a dynamic simulation is constructed using the established quasi-steady model and the wing design. Simulations were performed to gain insight on how different parameters affect the wing rotational response. Based on system resonance a model to predict the optimal rotational hinge stiffness based on given wing parameter and flapping wing kinematic is proposed. By varying different wing parameters, the proposed method is shown to be applicable to a wide range of wing designs with different sizes and shapes. With the selected hinge stiffness value, aspects of the rotational joint design is discussed and an integrated wing-hinge structure design using laminated carbon fiber and polymer film is presented. Manufacturing process of such composite structure is developed to achieve high accuracy and repeatability. The yielded hinge stiffness is verified by measurements. To validate the proposed model, flapping wing experiments were conducted. A flapping actuation set up is built using DC motor and a controller is implemented on a microcontroller to track desired wing stroke kinematic. Wing stroke and rotation kinematic were extracted using a high speed camera and the lift generation is evaluated. A total of 49 flapping experiments were presented, experimental data shows good correlation with the model's prediction. With the wing rotational hinge stiffness designed so that the rotational resonant frequency is twice as the stroke frequency, the resulting wing rotation generates near optimal lift. With further simulation, the proposed model shows low sensitivity to wing parameter variation. As a result, giving a design parameter of a flapping wing robot platform, the proposed principle can

  12. Design and demonstration of a small expandable morphing wing

    NASA Astrophysics Data System (ADS)

    Heryawan, Yudi; Park, Hoon C.; Goo, Nam S.; Yoon, Kwang J.; Byun, Yung H.

    2005-05-01

    In this paper, we present design, manufacturing, and wind tunnel test for a small-scale expandable morphing wing. The wing is separated into inner and outer wings as a typical bird wing. The part from leading edge of the wing chord is made of carbon composite strip and balsa. The remaining part is covered with curved thin carbon fiber composite mimicking wing feathers. The expandable wing is driven by a small DC motor, reduction gear, and fiber reinforced composite linkages. Rotation of the motor is switched to push-pull linear motion by a screw and the linear motion of the screw is transferred to linkages to create wing expansion and folding motions. The wing can change its aspect ratio from 4.7 to 8.5 in about 2 seconds and the speed can be controlled. Two LIPCAs (Lightweight Piezo-Composite Actuators) are attached under the inner wing section and activated on the expanded wing state to modify camber of the wing. In the wind tunnel test, change of lift, drag, and pitching moment during wing expansion have been investigated for various angles of attack. The LIPCA activation has created significant additional lift.

  13. Gliding swifts attain laminar flow over rough wings.

    PubMed

    Lentink, David; de Kat, Roeland

    2014-01-01

    Swifts are among the most aerodynamically refined gliding birds. However, the overlapping vanes and protruding shafts of their primary feathers make swift wings remarkably rough for their size. Wing roughness height is 1-2% of chord length on the upper surface--10,000 times rougher than sailplane wings. Sailplanes depend on extreme wing smoothness to increase the area of laminar flow on the wing surface and minimize drag for extended glides. To understand why the swift does not rely on smooth wings, we used a stethoscope to map laminar flow over preserved wings in a low-turbulence wind tunnel. By combining laminar area, lift, and drag measurements, we show that average area of laminar flow on swift wings is 69% (n = 3; std 13%) of their total area during glides that maximize flight distance and duration--similar to high-performance sailplanes. Our aerodynamic analysis indicates that swifts attain laminar flow over their rough wings because their wing size is comparable to the distance the air travels (after a roughness-induced perturbation) before it transitions from laminar to turbulent. To interpret the function of swift wing roughness, we simulated its effect on smooth model wings using physical models. This manipulation shows that laminar flow is reduced and drag increased at high speeds. At the speeds at which swifts cruise, however, swift-like roughness prolongs laminar flow and reduces drag. This feature gives small birds with rudimentary wings an edge during the evolution of glide performance. PMID:24964089

  14. Lift on Flexible and Rigid Cambered Wings at High Incidence

    NASA Astrophysics Data System (ADS)

    Jones, Anya; Mancini, Peter; Granlund, Kenneth; Ol, Michael

    2014-11-01

    The effects of camber and camber change due to elastic deflection of a membrane wing were investigated for wings in rectilinear translation with parameter variations in wing incidence and acceleration. Direct force and moment measurements were performed on a rigid flat plate wing, rigid cambered wings, and a membrane wing. Features in the force histories were further examined via flow visualization by planar laser illumination of fluorescent dye. Below 10 degrees of incidence, Wagner's approximation accurately predicts the time-evolution of lift for the rigid wings. At higher incidence, flow separation results in force transients, and the effect of wing camber is no longer additive. Both the rigid flat plate and rigid cambered wings reach peak lift at a 35 degree angle of attack, whereas the flexible wing experiences stall delay and reaches peak lift at 50 degrees. Due to the aeroelasticity of the flexible membrane, flow over the suction surface remains attached for much higher incidence angles than for the rigid wings. For incidence angles less than 30 degrees, the peak lift of the flexible wing is lower than that of its rigid counterparts. Beyond 30 degrees, the flexible wing experiences an aeroelastically induced stall delay that allows lift to exceed the rigid analogs. This work was supported by the Air Force Office of Scientific Research (AFOSR) Summer Faculty Fellowship Program and the U.S. Army Research Laboratory under the Micro Autonomous Systems and Technology (MAST) program.

  15. Gliding Swifts Attain Laminar Flow over Rough Wings

    PubMed Central

    Lentink, David; de Kat, Roeland

    2014-01-01

    Swifts are among the most aerodynamically refined gliding birds. However, the overlapping vanes and protruding shafts of their primary feathers make swift wings remarkably rough for their size. Wing roughness height is 1–2% of chord length on the upper surface—10,000 times rougher than sailplane wings. Sailplanes depend on extreme wing smoothness to increase the area of laminar flow on the wing surface and minimize drag for extended glides. To understand why the swift does not rely on smooth wings, we used a stethoscope to map laminar flow over preserved wings in a low-turbulence wind tunnel. By combining laminar area, lift, and drag measurements, we show that average area of laminar flow on swift wings is 69% (n = 3; std 13%) of their total area during glides that maximize flight distance and duration—similar to high-performance sailplanes. Our aerodynamic analysis indicates that swifts attain laminar flow over their rough wings because their wing size is comparable to the distance the air travels (after a roughness-induced perturbation) before it transitions from laminar to turbulent. To interpret the function of swift wing roughness, we simulated its effect on smooth model wings using physical models. This manipulation shows that laminar flow is reduced and drag increased at high speeds. At the speeds at which swifts cruise, however, swift-like roughness prolongs laminar flow and reduces drag. This feature gives small birds with rudimentary wings an edge during the evolution of glide performance. PMID:24964089

  16. Investigating the Force Production of Functionally-Graded Flexible Wings in Flapping Wing Flight

    NASA Astrophysics Data System (ADS)

    Mudbhari, Durlav; Erdogan, Malcolm; He, Kai; Bateman, Daniel; Lipkis, Rory; Moored, Keith

    2015-11-01

    Birds, insects and bats oscillate their wings to propel themselves over long distances and to maneuver with unprecedented agility. A key element to achieve their impressive aerodynamic performance is the flexibility of their wings. Numerous studies have shown that homogeneously flexible wings can enhance force production, propulsive efficiency and lift efficiency. Yet, animal wings are not homogenously flexible, but instead have varying material properties. The aim of this study is to characterize the force production and energetics of functionally-graded flexible wings. A partially-flexible wing composed of a rigid section and a flexible section is used as a first-order model of functionally-graded materials. The flexion occurs in the spanwise direction and it is affected by the spanwise flexion ratio, that is, the ratio of the length of the rigid section compared to the total span length. By varying the flexion ratio as well as the material properties of the flexible section, the study aims to examine the force production and energetics of flapping flight with functionally-graded flexible wings. Supported by the Office of Naval Research under Program Director Dr. Bob Brizzolara, MURI grant number N00014-14-1-0533.

  17. Oblique wing transonic transport configuration development

    NASA Technical Reports Server (NTRS)

    1977-01-01

    Studies of transport aircraft designed for boom-free supersonic flight show the variable sweep oblique wing to be the most efficient configuration for flight at low supersonic speeds. Use of this concept leads to a configuration that is lighter, quieter, and more fuel efficient than symmetric aircraft designed for the same mission. Aerodynamic structural, weight, aeroelastic and flight control studies show the oblique wing concept to be technically feasible. Investigations are reported for wing planform and thickness, pivot design and weight estimation, engine cycle (bypass ratio), and climb, descent and reserve fuel. Results are incorporated into a final configuration. Performance, weight, and balance characteristics are evaluated. Flight control requirements are reviewed, and areas in which further research is needed are identified.

  18. Computational fluid dynamics of airfoils and wings

    NASA Technical Reports Server (NTRS)

    Garabedian, P.; Mcfadden, G.

    1982-01-01

    It is pointed out that transonic flow is one of the fields where computational fluid dynamics turns out to be most effective. Codes for the design and analysis of supercritical airfoils and wings have become standard tools of the aircraft industry. The present investigation is concerned with mathematical models and theorems which account for some of the progress that has been made. The most successful aerodynamics codes are those for the analysis of flow at off-design conditions where weak shock waves appear. A major breakthrough was achieved by Murman and Cole (1971), who conceived of a retarded difference scheme which incorporates artificial viscosity to capture shocks in the supersonic zone. This concept has been used to develop codes for the analysis of transonic flow past a swept wing. Attention is given to the trailing edge and the boundary layer, entropy inequalities and wave drag, shockless airfoils, and the inverse swept wing code.

  19. General Potential Theory of Arbitrary Wing Sections

    NASA Technical Reports Server (NTRS)

    Theodorsen, T.; Garrick, I. E.

    1979-01-01

    The problem of determining the two dimensional potential flow around wing sections of any shape is examined. The problem is condensed into the compact form of an integral equation capable of yielding numerical solutions by a direct process. An attempt is made to analyze and coordinate the results of earlier studies relating to properties of wing sections. The existing approximate theory of thin wing sections and the Joukowski theory with its numerous generalizations are reduced to special cases of the general theory of arbitrary sections, permitting a clearer perspective of the entire field. The method which permits the determination of the velocity at any point of an arbitrary section and the associated lift and moments is described. The method is also discussed in terms for developing new shapes of preassigned aerodynamical properties.

  20. Wing-Body Aeroelasticity on Parallel Computers

    NASA Technical Reports Server (NTRS)

    Guruswamy, Guru P.; Byun, Chansup

    1996-01-01

    This article presents a procedure for computing the aeroelasticity of wing-body configurations on multiple-instruction, multiple-data parallel computers. In this procedure, fluids are modeled using Euler equations discretized by a finite difference method, and structures are modeled using finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. A parallel integration scheme is used to compute aeroelastic responses by solving the coupled fluid and structural equations concurrently while keeping modularity of each discipline. The present procedure is validated by computing the aeroelastic response of a wing and comparing with experiment. Aeroelastic computations are illustrated for a high speed civil transport type wing-body configuration.