Science.gov

Sample records for low-aspect-ratio airplane wings

  1. 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.

  2. Winglets on low aspect ratio wings

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Liaw, Paul

    1987-01-01

    The drag reduction potentially available from the use of winglets at the tips of low aspect ratio (1.75-2.67) wings with pronounced (45-60 deg) leading edge sweep is assessed numerically for the case of a cruise design point at Mach of 0.8 and a lift coefficient of 0.3. Both wing-winglet and wing-alone design geometries are derived from a linear-theory, minimum induced drag design methodology. Relative performance is evaluated with a nonlinear extended small disturbance potential flow analysis code. Predicted lift coefficient/pressure drag coefficient increases at equal lift for the wing-winglet configurations over the wing-alone planform are of the order of 14.6-15.8, when boundary layer interaction is included.

  3. Aerodynamic tests of a low aspect ratio tapered wing with various flaps, for use on tailless airplanes

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Sanders, Robert

    1933-01-01

    Wind tunnel tests were made of a model wing having an aspect ratio of 3 and a tapered plan form with a straight trailing edge. The model had the Clark Y airfoil section throughout it's entire span and had no washout, depending on a trailing-edge flap for longitudinal balance and control. The flap had a constant chord and was divided into four equal portions along the span. The tests were made with the entire flap deflected to obtain longitudinal control and balance, and also with the inner portions deflected alone, and with the outer portions deflected alone. It was found that the simple wing with no washout or change of basic section along the span has aerodynamic characteristics well suited for use on tailless airplanes. A higher lift coefficient was obtained with the full-span flap deflected as a unit to give longitudinal balance than was obtained with either the inner or the outer portions of the flap deflected.

  4. Vortex Interaction on Low Aspect Ratio Membrane Wings

    NASA Astrophysics Data System (ADS)

    Waldman, Rye M.; Breuer, Kenneth S.

    2013-11-01

    Inspired by the flight of bats and by recent interest in Micro Air Vehicles, we present measurements on the steady and unsteady behavior of low aspect ratio membrane wings. We conduct wind tunnel experiments with coupled force, kinematic, and flow field measurements, both on the wing and in the near wake. Membrane wings interact strongly with the vortices shed from the leading- and trailing-edges and the wing tips, and the details of the membrane support play an important role in the fluid-structure interaction. Membranes that are supported at the wing tip exhibit less membrane flutter, more coherent tip vortices, and enhanced lift. The interior wake can exhibit organized spanwise vortex shedding, and shows little influence from the tip vortex. In contrast, membranes with an unsupported wing tip show exaggerated static deformation, significant membrane fluttering and a diffuse, unsteady tip vortex. The unsteady tip vortex modifies the behavior of the interior wake, disrupting the wake coherence.

  5. Unsteady Flow Structure on Low Aspect Ratio Wings

    DTIC Science & Technology

    2011-01-06

    4 EFFECT OF PITCH RATE ON NEAR-SURFACE TOPOLOGY ON A DELTA WING The near-surface flow structure and topology on a delta wing of...34Investigation of Flow Structure on a Pitching Delta Wing of Moderate Sweep Angle using Stereoscopic Particle Image Velocimetry", October, 2008. * All...for inducement of flow reattachment, were in the range fC/U = 1 to 2. The present configuration is a substantial departure from a flat delta wing

  6. Stability Characteristics of Low Reynolds Number, Low Aspect Ratio Wings

    NASA Astrophysics Data System (ADS)

    Shields, Matthew; Mohseni, Kamran

    2010-11-01

    The recent interest in Micro Aerial Vehicles (MAVs) has led to the development of many different aircraft; however, little progress has been made in understanding the physics of MAV flow. MAVs aerodynamics is affected by low Reynolds number flow and low aspect ratios. As a result nonlinear effects due to tip vortices are quite important. We have developed a new experimental setup for measuring stability derivatives in a small wind tunnel. Using a four degree of freedom actuation system, a model can be placed in the test section and maneuvered in such a way to isolate the flow components responsible for creating stability derivatives. Accurate measurements of the aerodynamic loading can then be used to compute these values. Initial testing was conducted primarily on a series of flat plates of different aspect ratios. In addition, the CU MAV was tested as a specific case study. Test results indicate that some of the cross coupled stability derivatives, ignored for larger aircrafts, are on the same order of magnitude as standard derivatives and thus can not be ignored in the derivation of the linear equations of motion for a micro aerial vehicle. As a result, a more general set of equations of motion are derived based upon experimentally obtained stability derivatives.

  7. High-Lift Capability of Low Aspect Ratio Wings Utilizing Circulation Control and Upper Surface Blowing

    DTIC Science & Technology

    1980-07-01

    the Upper Surface Blowing (USB) and the Circulation Control Wing (CCW). Both concepts use the Coanda effect as a means of augmenting aerodynamic lift...USB), and a unique combination of the two (CCW/USB). Wing tip sails were used as a means of increasing th(, effective aspect ratio of these wings...wing tip sails are effective in reducing the induced drag of these powered- lift low aspect ratio wings under high-lift conditions. The induced drag

  8. Characteristics of Low-Aspect-Ratio Wings at Supercritical Mach Numbers

    NASA Technical Reports Server (NTRS)

    Stack, John; Lindsey, W F

    1949-01-01

    The separation of the flow over wings precipitated by the compression shock that forms as speeds are increased into the supercritical Mach number range has imposed serious difficulties in the improvement of aircraft performance. Three difficulties rise principally as a consequence of the rapid drag rise and the loss of lift that causes serious stability changes when the wing shock-stalls. Favorable relieving effects due to the three-dimensional flow around the tips were obtained and these effects were of such magnitude that it is indicated that low-aspect-ratio wings offer a possible solution of the problems encountered.

  9. Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

    NASA Technical Reports Server (NTRS)

    Mansour, N. N.

    1984-01-01

    The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

  10. Numerical simulation of transonic separated flows over low-aspect ratio wings

    NASA Technical Reports Server (NTRS)

    Kaynak, U.; Holst, T. L.; Sorenson, R. L.; Cantwell, B. J.

    1986-01-01

    Transonic flow fields about a low-aspect-ratio advanced technology wing have been computed using a viscous/inviscid zonal approach. The flow field near the wing where viscous effects are important was solved using the 'Reynolds-Averaged Navier-Stokes Equations' in 'thin-layer' form. The Euler equations were used to determine the flow field in regions away from the wing where viscous effects are insignificant. A zonal grid using an H-H topology was generated around the wing by first solving a set of Poisson's equations for the global grid. This grid was then subdivided into separate zones of viscous or inviscid flow as suggested by the flow physics. A series of flow cases were computed and compared with corresponding sets of experimental data. All cases showed good agreement with experiment in terms of the pressure field. Also, a good correlation between computed separated surface flow and experimental oil flow was obtained.

  11. Computational design of low aspect ratio wing-winglets for transonic wind-tunnel testing

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1989-01-01

    A computational design has been performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three planforms has been selected to be constructed as a wind tunnel model for testing in the NASA LaRC 7 x 10 High Speed Wind Tunnel. A design point of M = 0.8, CL approx = 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. This report summarizes the design process and the predicted transonic performance for each configuration.

  12. Computational design of low aspect ratio wing-winglet configurations for transonic wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1988-01-01

    A computational design has been performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three planforms has been selected to be constructed as a wind tunnel model for testing in the NASA LaRC 7 x 10 High Speed Wind Tunnel. A design point of M = 0.8, CL approx = 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. This report summarizes the design process and the predicted transonic performance for each configuration.

  13. Boundary-layer measurements on a transonic low-aspect ratio wing

    NASA Technical Reports Server (NTRS)

    Keener, Earl R.

    1985-01-01

    Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.

  14. Effects of fluid behavior around low aspect ratio, low Reynolds number wings on aerodynamic stability

    NASA Astrophysics Data System (ADS)

    Shields, Matthew; Mohseni, Kamran

    2011-11-01

    The innovation of micro aerial vehicles (MAVs) has brought to attention the unique flow regime associated with low aspect ratio (LAR), low Reynolds number fliers. The dominant effects of developing tip vortices and leading edge vortices create a fundamentally different flow regime than that of conventional aircraft. An improved knowledge of low aspect ratio, low Reynolds number aerodynamics can be greatly beneficial for future MAV design. A little investigated but vital aspect of LAR aerodynamics is the behavior of the fluid as the wing yaws. Flow visualization experiments undertaken in the group for the canonical case of varying AR flat plates indicate that the propagation of the tip vortex keeps the flow attached over the upstream portion of the wing, while the downstream vortex is convected away from the wing. This induces asymmetric, destabilizing loading on the wing which has been observed to adversely affect MAV flight. In addition, experimental load measurements indicate significant nonlinearities in forces and moments which can be attributed to the development and propagation of these vortical structures. A non-dimensional analysis of the rigid body equations of motion indicates that these nonlinearities create dependencies which dramatically change the conventional linearization process. These flow phenomena are investigated with intent to apply to future MAV design.

  15. Some low-speed flutter characteristics of simple low-aspect-ratio delta wing models

    NASA Technical Reports Server (NTRS)

    Doggett, Robert V., Jr.; Soistmann, David L.

    1989-01-01

    Some results from a combined experimental and analytical study of the low-speed flutter characteristics of low-aspect-ratio delta wings are presented. Data are presented which show the effects of sweep angle on the flutter characteristics of some simple plate models of constant planform area. The range of sweep angles studied was from 30 to 72 degrees. In addition, flutter results are presented for two 30 deg-sweep clipped-delta wing models. Further, results are presented that show the effects of root clamping (percentage length of the root chord that is cantilevered) for a 45 deg-sweep delta wing. The experimental data are compared with analytical results obtained by using kernel function and doublet lattice subsonic unsteady lifting surface theories.

  16. 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.

  17. Properties of low-aspect-ratio pointed wings at speeds below and above the speed of sound

    NASA Technical Reports Server (NTRS)

    Jones, Robert T

    1946-01-01

    Low-aspect-ratio wings having pointed plan forms are treated on the assumption that the flow potentials in planes at right angles to the long axis of the airfoils are similar to the corresponding two-dimensional potentials.

  18. Computational design of low aspect ratio wing-winglet configurations for transonic wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1989-01-01

    Computational designs were performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three configurations was selected to be constructed as a wind tunnel model for testing in the NASA LaRC 8-foot transonic pressure tunnel. A design point of M = 0.8, C(sub L) is approximate or = to 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 deg and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. The design process and the predicted transonic performance are summarized for each configuration. In addition, a companion low-speed design study was conducted, using one of the transonic design wing-winglet planforms but with different camber and thickness distributions. A low-speed wind tunnel model was constructed to match this low-speed design geometry, and force coefficient data were obtained for the model at speeds of 100 to 150 ft/sec. Measured drag coefficient reductions were of the same order of magnitude as those predicted by numerical subsonic performance predictions.

  19. Navier-Stokes computations of vortical flows over low aspect ratio wings

    NASA Technical Reports Server (NTRS)

    Thomas, J. L.; Taylor, S. L.; Anderson, W. K.

    1987-01-01

    An upwind-biased finite-volume algorithm is applied to the low-speed flow over a low aspect ratio delta wing from zero to forty degrees angle of attack. The differencing is second-order accurate spatially, and a multigrid algorithm is used to promote convergence to the steady state. The results compare well with the detailed experiments of Hummel (1983) and others for a Re(L) of 0.95 x 10 to the 6th. The predicted maximum lift coefficient of 1.10 at thirty-five degrees angle of attack agrees closely with the measured maximum lift of 1.06 at thirty-three degrees. At forty degrees angle of attack, a bubble type of vortex breakdown is evident in the computations, extending from 0.6 of the root chord to just downstream of the trailing edge.

  20. Pressure-distribution measurements on a transonic low-aspect ratio wing

    NASA Technical Reports Server (NTRS)

    Keener, E. R.

    1985-01-01

    Experimental surface pressure distributions and oil flow photographs are presented for a 0.90 m semispan model of NASA/Lockheed Wing C, a generic transonic, supercritical, low aspect ratio, highly 3-dimensional configuration. This wing was tested at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96, and a Reynolds number range from 3.4 x 1,000,000 to 10 x 1,000,000. Pressures were measured with both the tunnel floor and ceiling suction slots open for most of the tests but taped closed for some tests to simulate solid walls. A comparison is made with the measured pressures from a small model in high Reynolds number facility and with predicted pressures using two three dimesional, transonic full potential flow wing codes: design code FLO22 (nonconservative) and TWING code (conservative). At the given design condition, a small region of flow separation occurred. At a Mach number of 0.82 the flow was unseparated and the surface flow angles were less than 10 deg, indicating that the boundary layer flow was not 3-D. Evidence indicate that wings that are optimized for mild shock waves and mild pressure recovery gradients generally have small 3-D boundary layer flow at design conditions for unseparated flow.

  1. Unsteady flow structure and loading of a pitching low-aspect-ratio wing

    NASA Astrophysics Data System (ADS)

    Visbal, Miguel R.

    2017-02-01

    This study addresses the flow structure and unsteady loading arising over a pitching low-aspect-ratio rectangular wing under low-Reynolds-number conditions of interest in small unmanned aerial vehicle operation and gust interactions. Simulations are performed employing a high-fidelity computational approach capable of accurately capturing the complex unsteady transitional flows. The wing is pitched about its quarter-chord axis to a maximum incidence of 45∘ over time intervals ranging from four to 16 convective time scales. The Reynolds number based on the wing chord varied from 103 to 4 ×104 . For the highest pitch rate, good agreement between the computed three-dimensional (3D) flow structure and recent experimental measurements is demonstrated. The 3D dynamic stall process is characterized by the formation of an initially spanwise-oriented leading-edge vortex which evolves into an arch-type structure with legs anchored to the wing surface. The normal vorticity in the arch vortex legs establishes a low-pressure region and swirling pattern on the wing surface. A distinct characteristic of the arch vortex is its upstream propagation and persistence over the wing, postulated to be the result of the self-induced velocity of the vortex and its image underneath the plate. Increasing either pitch rate or Reynolds number promotes a more coherent arch vortex and circulation pattern, and delays the onset of stall to a higher angle of attack. Even for the lowest pitch rate considered, a significant increase in maximum lift is achieved relative to the static situation.

  2. Aerodynamics of low aspect ratio wings at low Reynolds numbers with applications to micro air vehicle design

    NASA Astrophysics Data System (ADS)

    Torres, Gabriel Eduardo

    2002-09-01

    The recent interest in the development of small UAVs and micro air vehicles has revealed a need for a more thorough understanding of the aerodynamics of small airplanes flying at low speeds. In response to this need, the present work presents a study of the lift, drag, and pitching moment characteristics of wings of low aspect ratio operating at low Reynolds numbers. Wind tunnel tests of wings with aspect ratios between 0.5 and 2.0, four distinct planforms, thickness-to-chord ratios of ≈2%, and 5-to-1 elliptical leading edges have been conducted as part of this research. The Reynolds numbers considered were in the range of 70,000 to 200,000. Analysis of the data includes comparison of lift-curve slope, induced-drag coefficient, and maximum lift coefficient with theory, discussion of aerodynamic center, and the effects of Reynolds number, camber, and leading-edge shape. As an example of an application of this wind tunnel data, the experimental results are implemented within an aircraft performance prediction procedure. This procedure is incorporated into a genetic algorithm optimization program that identifies near-optimum MAV configurations given certain requirements and constraints. Results obtained by use of this optimization procedure have revealed that useful design tools can be developed based on the experimental database.

  3. Strain-gage bridge calibration and flight loads measurements on a low-aspect-ratio thin wing

    NASA Technical Reports Server (NTRS)

    Peele, E. L.; Eckstrom, C. V.

    1975-01-01

    Strain-gage bridges were used to make in-flight measurements of bending moment, shear, and torque loads on a low-aspect-ratio, thin, swept wing having a full depth honeycomb sandwich type structure. Standard regression analysis techniques were employed in the calibration of the strain bridges. Comparison of the measured loads with theoretical loads are included.

  4. Hydrodynamic Performance of a Dual-Slotted Circulation Control Wing of Low-Aspect Ratio

    DTIC Science & Technology

    2006-12-01

    experimental effort was undertaken to document the performance of a low-aspect ratio hydrofoil with trailing edge blowing across a Coanda surface in a...CCFoil, upper slot blowing, Coanda surface, tip vortex, and lighting vortex at 80% span shown...62 Figure 70. Surface cavitation on Coanda Surface, close-up of the slot flow, incpetion behind slot lip face

  5. Characteristics of a Dual-Slotted Circulation Control Wing of Low Aspect Ratio Intended for Naval Hydrodynamic Applications

    DTIC Science & Technology

    2004-01-01

    water tunnel to explore the application of Coanda -effect circulation control to low aspect ratio wings. The facility was the Large Cavitation Channel...with the effects on conven- tional lift-due-to-angle-of-attack, and cavitation in the Coanda wall jet region does not result in jet detachment or an...2) cavitation properties of the trailing edge Coanda wall jet, (3) attributes of dual slots, including wake-filling. The answers to these

  6. Properties of low-aspect-ratio pointed wings at speeds below and above the speed of sound

    NASA Technical Reports Server (NTRS)

    Jones, R. T.

    1976-01-01

    Low aspect ratio wings having pointed planforms are treated on the assumption that the flow potentials in planes at right angles of the long axis of the airfoils are similar to the corresponding two dimensional potentials. For the limiting case of small angles of attack and low aspect ratios the theory brings out the following significant properties: (1) the lift of a slender pointed airfoil moving in the direction of its long axis depends on the increase in width of the sections in a downstream direction; (2) spanwise loading of such an airfoil is independent of planform and approaches the distribution giving a minimum induced drag; and (3) lift distribution of a pointed airfoil traveling point-foremost is relatively unaffected by the compressibility of the air below or above the speed of sound.

  7. Properties of low-aspect-ratio pointed wings at speeds below and above the speed of sound

    NASA Technical Reports Server (NTRS)

    Jones, R. T.

    1976-01-01

    Low aspect ratio wings having pointed planforms are treated on the assumption that the flow potentials in planes at right angles of the long axis of the airfoils are similar to the corresponding two dimensional potentials. For the limiting case of small angles of attack and low aspect ratios the theory brings out the following significant properties: (1) the lift of a slender pointed airfoil moving in the direction of its long axis depends on the increase in width of the sections in a downstream direction; (2) spanwise loading of such an airfoil is independent of planform and approaches the distribution giving a minimum induced drag; and (3) lift distribution of a pointed airfoil traveling point-foremost is relatively unaffected by the compressibility of the air below or above the speed of sound.

  8. 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.

  9. Self-induced roll oscillations of low-aspect-ratio rectangular wings

    NASA Technical Reports Server (NTRS)

    Levin, Daniel; Katz, Joseph

    1990-01-01

    Experimental investigation of small aspect ratio rectangular wings mounted on a free-to-roll sting balance indicated that self-induced roll oscillations are possible when the aspect ratio of such wings is less than 0.5. The oscillations are probably driven by the periodic changes in the location and strength of the side edge vortices, as it has been shown for the 'wing rock' motion of delta wings, where similar changes in the leading edge vortex strength and position cause the roll oscillations. During the roll oscillation cycle the roll angle, normal force, and the side force were recorded and presented for three wings with aspect ratios of 0.25, 0.35, and 0.47. This data indicates that the lift loss during roll oscillations of rectangular wings is less than what was measured for similar delta wings. Also, the flow field of such slender rectangular wings at high angles of attack is more complicated due to the additional leading edge vortex, when compared with the flow field over slender delta wings.

  10. Effect of torsional stiffness and inertia on the dynamics of low aspect ratio flapping wings.

    PubMed

    Xiao, Qing; Hu, Jianxin; Liu, Hao

    2014-03-01

    Micro air vehicle-motivated aerodynamics in biological flight has been an important subject in the past decade. Inspired by the novel flapping wing mechanisms in insects, birds and bats, we have carried out a numerical study systematically investigating a three-dimensional flapping rigid wing with passively actuated lateral and rotational motion. Distinguishing it from the limited existing studies, this work performs a systematic examination on the effects of wing aspect ratio (AR = 1.0 to infinity), inertia (density ratio σ = 4-32), torsional stiffness (frequency ratio F = 1.5-10 and infinity) and pivot point (from chord-center to leading edge) on the dynamics response of a low AR rectangular wing under an initial zero speed flow field condition. The simulation results show that the symmetry breakdown of the flapping wing results in a forward/backward motion with a rotational pitching. When the wing reaches its stable periodic state, the induced pitching frequency is identical to its forced flapping frequency. However, depending on various kinematic and dynamic system parameters, (i.e. flapping frequency, density ratio and pitching axis), the lateral induced velocity shows a number of different oscillating frequencies. Furthermore, compared with a one degree of freedom (DoF) wing in the lateral direction only, the propulsion performance of such a two DoF wing relies very much on the magnitude of torsional stiffness adding on the pivot point, as well as its pitching axis. In all cases examined here, thrust force and moment generated by a long span wing is larger than that of a short wing, which is remarkably linked to the strong reverse von Kármán vortex street formed in the wake of a wing.

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

    NASA Technical Reports Server (NTRS)

    Nielsen, J. N.

    1986-01-01

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

  12. A model for roll stall and the inherent stability modes of low aspect ratio wings at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Shields, Matt

    The development of Micro Aerial Vehicles has been hindered by the poor understanding of the aerodynamic loading and stability and control properties of the low Reynolds number regime in which the inherent low aspect ratio (LAR) wings operate. This thesis experimentally evaluates the static and damping aerodynamic stability derivatives to provide a complete aerodynamic model for canonical flat plate wings of aspect ratios near unity at Reynolds numbers under 1 x 105. This permits the complete functionality of the aerodynamic forces and moments to be expressed and the equations of motion to solved, thereby identifying the inherent stability properties of the wing. This provides a basis for characterizing the stability of full vehicles. The influence of the tip vortices during sideslip perturbations is found to induce a loading condition referred to as roll stall, a significant roll moment created by the spanwise induced velocity asymmetry related to the displacement of the vortex cores relative to the wing. Roll stall is manifested by a linearly increasing roll moment with low to moderate angles of attack and a subsequent stall event similar to a lift polar; this behavior is not experienced by conventional (high aspect ratio) wings. The resulting large magnitude of the roll stability derivative, Cl,beta and lack of roll damping, Cl ,rho, create significant modal responses of the lateral state variables; a linear model used to evaluate these modes is shown to accurately reflect the solution obtained by numerically integrating the nonlinear equations. An unstable Dutch roll mode dominates the behavior of the wing for small perturbations from equilibrium, and in the presence of angle of attack oscillations a previously unconsidered coupled mode, referred to as roll resonance, is seen develop and drive the bank angle? away from equilibrium. Roll resonance requires a linear time variant (LTV) model to capture the behavior of the bank angle, which is attributed to the

  13. Properties of low-aspect-ratio pointed wings at speeds below and above the speed of sound

    NASA Technical Reports Server (NTRS)

    Jones, Robert T

    1946-01-01

    Low-aspect-ratio wings having pointed plan forms are treated on the assumption that the flow potentials in planes at right angles to the long axis of the airfoils are similar to the corresponding two-dimensional potentials. For the limiting case of small angles of attack and low aspect ratios the theory brings out the following significant properties: (1) The lift of a slender, pointed airfoil moving in the direction of its long axis depends on the increase in width of the sections in a downstream direction. Sections behind the section of maximum width develop no lift. (2) The spanwise loading of such an airfoil is independent of the plan form and approaches the distribution giving a minimum induced drag. (3) The lift distribution of a pointed airfoil travelling point-foremost is relatively unaffected by the compressibility of the air below or above the speed of sound. A best of a triangular airfoil at a Mach number of 1.75 verified the theoretical values of lift and center of pressure.

  14. A vortex model for forces and moments on low-aspect-ratio wings in side-slip with experimental validation.

    PubMed

    DeVoria, Adam C; Mohseni, Kamran

    2017-02-01

    This paper studies low-aspect-ratio ([Formula: see text]) rectangular wings at high incidence and in side-slip. The main objective is to incorporate the effects of high angle of attack and side-slip into a simplified vortex model for the forces and moments. Experiments are also performed and are used to validate assumptions made in the model. The model asymptotes to the potential flow result of classical aerodynamics for an infinite aspect ratio. The [Formula: see text] → 0 limit of a rectangular wing is considered with slender body theory, where the side-edge vortices merge into a vortex doublet. Hence, the velocity fields transition from being dominated by a spanwise vorticity monopole ([Formula: see text] ≫ 1) to a streamwise vorticity dipole ([Formula: see text] ∼ 1). We theoretically derive a spanwise loading distribution that is parabolic instead of elliptic, and this physically represents the additional circulation around the wing that is associated with reattached flow. This is a fundamental feature of wings with a broad-facing leading edge. The experimental measurements of the spanwise circulation closely approximate a parabolic distribution. The vortex model yields very agreeable comparison with direct measurement of the lift and drag, and the roll moment prediction is acceptable for [Formula: see text] ≤ 1 prior to the roll stall angle and up to side-slip angles of 20°.

  15. A vortex model for forces and moments on low-aspect-ratio wings in side-slip with experimental validation

    NASA Astrophysics Data System (ADS)

    DeVoria, Adam C.; Mohseni, Kamran

    2017-02-01

    This paper studies low-aspect-ratio () rectangular wings at high incidence and in side-slip. The main objective is to incorporate the effects of high angle of attack and side-slip into a simplified vortex model for the forces and moments. Experiments are also performed and are used to validate assumptions made in the model. The model asymptotes to the potential flow result of classical aerodynamics for an infinite aspect ratio. The → 0 limit of a rectangular wing is considered with slender body theory, where the side-edge vortices merge into a vortex doublet. Hence, the velocity fields transition from being dominated by a spanwise vorticity monopole ( ≫ 1) to a streamwise vorticity dipole ( ˜ 1). We theoretically derive a spanwise loading distribution that is parabolic instead of elliptic, and this physically represents the additional circulation around the wing that is associated with reattached flow. This is a fundamental feature of wings with a broad-facing leading edge. The experimental measurements of the spanwise circulation closely approximate a parabolic distribution. The vortex model yields very agreeable comparison with direct measurement of the lift and drag, and the roll moment prediction is acceptable for ≤ 1 prior to the roll stall angle and up to side-slip angles of 20°.

  16. Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.

    1974-01-01

    Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.

  17. Spectral element simulations of unsteady flow over a 3D, low aspect-ratio semi-circular wing

    NASA Astrophysics Data System (ADS)

    Kandala, Sriharsha; Rempfer, Dietmar

    2011-11-01

    Numerical simulations of unsteady 3D flow over a low-aspect-ratio semi-circular wing are performed using a spectral element method. Specsolve, a parallel spectral element solver currently under development at IIT, is used for the simulation. The solution is represented locally as a tensor product of Legendre polynomials and C0-continuity is enforced between adjacent domains. A BDF/EXT scheme is used for temporal integration. The fractional step method is used for computing velocity and pressure. The code incorporates a FDM (fast diagonalization method) based overlapping Schwarz preconditioner for the consistent Poisson operator and an algebraic multigrid based coarse grid solver (P.F. Fischer et al, J. Phys.: Conf. Ser.(125) 012076, 2008) for pressure. The simulation replicates the conditions of the active flow control experiment (D. Williams et al, AIAA paper, 2010-4969). The Reynolds number based on chord length and free-stream velocity is about 68000. Different angles of attack, encompassing both pre-stall and post-stall regimes, are considered. These results are compared with data from the experiment and numerical simulations based on Lattice Boltzmann method (G. Brès et al, AIAA paper, 2010-4713).

  18. Effect of Aspect Ratio on the Low-Speed Lateral Control Characteristics of Untapered Low-Aspect-Ratio Wings Equipped with Flap and with Retractable Ailerons

    NASA Technical Reports Server (NTRS)

    Fischel, Jack; Naeseth, Rodger L; Hagerman, John R; O'Hare, William M

    1952-01-01

    A low-speed wind-tunnel investigation was made to determine the lateral control characteristics of a series of untapered low-aspect-ratio wings. Sealed flap ailerons of various spans and spanwise locations were investigated on unswept wings of aspect ratios 1.13, 1.13, 4.13, and 6.13; and various projections of 0.60-semispan retractable ailerons were investigated on the unsweptback wings of aspect ratios 1.13, 2.13, and 4.13 and on a 45 degree sweptback wing. The retractable ailerons investigated on the unswept wings spanned the outboard stations of each wing; whereas the plain and stepped retractable ailerons investigated on the sweptback wing were located at various spanwise stations. Design charts based on experimental results are presented for estimating the flap aileron effectiveness for low-aspect-ratio, untapered, unswept.

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

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  20. Supersonic aerodynamic characteristics of a low-aspect-ratio missile model with wing and tail controls and with tails in line and interdigitated

    NASA Technical Reports Server (NTRS)

    Graves, E. B.

    1972-01-01

    A study has been made to determine the aerodynamic characteristics of a low-aspect ratio cruciform missile model with all-movable wings and tails. The configuration was tested at Mach numbers from 1.50 to 4.63 with the wings in the vertical and horizontal planes and with the wings in a 45 deg roll plane with tails in line and interdigitated.

  1. Flight loads measurements obtained from calibrated strain-gage bridges mounted externally on the skin of a low-aspect-ratio wing

    NASA Technical Reports Server (NTRS)

    Eckstrom, C. V.

    1976-01-01

    Flight-test measurements of wingloads (shear, bending moment, and torque) were obtained by means of strain-gage bridges mounted on the exterior surface of a low-aspect-ratio, thin, swept wing which had a structural skin, full-depth honeycomb core, sandwich construction. Details concerning the strain-gage bridges, the calibration procedures used, and the flight-test results are presented along with some pressure measurements and theoretical calculations for comparison purposes.

  2. Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of a Fighter Model Employing a Low-Aspect-Ratio Unswept Wing and a Horizontal Tail Mounted Well Above the Wing Plane - Longitudinal Stability and Control

    NASA Technical Reports Server (NTRS)

    Smith, Williard G.

    1954-01-01

    Experimental results showing the static longitudinal-stability and control characteristics of a model of a fighter airplane employing a low-aspect-ratio unswept wing and an all-movable horizontal tail are presented. The investigation was made over a Mach number range from 0.60 to 0.90 and from 1.35 to 1.90 at a constant Reynolds number of 2.40 million, based on the wing mean aerodynamic chord. Because of the location of the horizontal tail at the tip of the vertical tail, interference was noted between the vertical tail and the horizontal tail and between the wing and the horizontal tail. This interference produced a positive pitching-moment coefficient at zero lift throughout the Mach number range of the tests, reduced the change in stability with increasing lift coefficient of the wing at moderate lift coefficients in the subsonic speed range, and reduced the stability at low lift coefficients at high supersonic speeds. The lift and pitching-moment effectiveness of the all movable tail was unaffected by the interference effects and was constant throughout the lift-coefficient range of the tests at each Mach number except 1.90.

  3. Space shuttle: Longitudinal aerodynamic characteristics of low aspect ratio wing configurations in ground effect for a moving and stationary ground surface

    NASA Technical Reports Server (NTRS)

    Romere, P. O.; Chambliss, E. B.

    1972-01-01

    A 0.05-scale model of the NASA-MSC Orbiter 040A Configuration was tested. Test duration was approximately 80 hours during which the model was tested in and out of ground effect with a stationary and moving ground belt. Model height from ground plane surface was varied from one and one-half wing span to landing touchdown while angle of attack varied from -4 to 20 degrees. Eleven effectiveness and alternate configuration geometries were tested to insure complete analysis of low aspect ratio wing aircraft in the presence of ground effect. Test Mach number was approximately 0.067 with a corresponding dynamic pressure value of 6.5 psf.

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

    NASA Technical Reports Server (NTRS)

    Robinson, Ross B

    1957-01-01

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

  5. The effects of winglets on low aspect ratio wings at supersonic Mach numbers. M.S. Thesis Report Feb. 1989 - Apr. 1991

    NASA Technical Reports Server (NTRS)

    Keenan, James A.; Kuhlman, John M.

    1991-01-01

    A computational study was conducted on two wings, of aspect ratios 1.244 and 1.865, each having 65 degree leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A Mach number of 1.62 was selected as the design value. The winglets studied were parametrically varied in alignment, length, sweep, camber, thickness, and dihedral angle to determine which geometry had the best predicted performance. For the computational analysis, an available Euler marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan. The first wing with winglet used NACA 1402 airfoils for the base wing and was shown to have lift-to-pressure drag ratios within 0.136 percent to 0.360 percent of the NACA 1402 wing-alone. The other base wing was a natural flow wing which was previously designed specifically for a Mach number of 1.62. The results obtained showed that the natural wing-alone had a slightly higher lift-to-pressure drag than the natural wing with winglets.

  6. High-Reynolds-Number Test of a 5-Percent-Thick Low-Aspect-Ratio Semispan Wing in the Langley 0.3-Meter Transonic Cryogenic Tunnel: Wing Pressure Distributions

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Lawing, Pierce L.

    1990-01-01

    A high Reynolds number test of a 5 percent thick low aspect ratio semispan wing was conducted in the adaptive wall test section of the Langley 0.3 m Transonic Cryogenic Tunnel. The model tested had a planform and a NACA 64A-105 airfoil section that is similar to that of the pressure instrumented canard on the X-29 experimental aircraft. Chordwise pressure data for Mach numbers of 0.3, 0.7, and 0.9 were measured for an angle-of-attack range of -4 to 15 deg. The associated Reynolds numbers, based on the geometric mean chord, encompass most of the flight regime of the canard. This test was a free transition investigation. A summary of the wing pressures are presented without analysis as well as adapted test section top and bottom wall pressure signatures. However, the presented graphical data indicate Reynolds number dependent complex leading edge separation phenomena. This data set supplements the existing high Reynolds number database and are useful for computational codes comparison.

  7. Influence of optimized leading-edge deflection and geometric anhedral on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration. [langley 7 by 10 foot tunnel

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Huffman, J. K.

    1979-01-01

    An investigation conducted in the Langley 7 by 10 foot tunnel to determine the influence of an optimized leading-edge deflection on the low speed aerodynamic performance of a configuration with a low aspect ratio, highly swept wing. The sensitivity of the lateral stability derivative to geometric anhedral was also studied. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to spanwise variation of unwash, the resulting optimized leading edge was a smooth, continuously warped surface for which the deflection varied from 16 deg at the side of body to 50 deg at the wing tip. For the particular configuration studied, levels of leading-edge suction on the order of 90 percent were achieved. The results of tests conducted to determine the sensitivity of the lateral stability derivative to geometric anhedral indicate values which are in reasonable agreement with estimates provided by simple vortex-lattice theories.

  8. Effects of Sweep and Thickness on the Static Longitudinal Aerodynamic Characteristics of a Series of Thin, Low-aspect-ratio, Highly Tapered Wings at Transonic Speeds : Transonic-bump Method

    NASA Technical Reports Server (NTRS)

    Fournier, Paul G; Few, Albert G , Jr

    1954-01-01

    An investigation by the transonic-bump technique of the static longitudinal aerodynamic characteristics of a series of thin, low-aspect-ratio, highly tapered wings has been made in the Langley high-speed 7- by 10-foot tunnel. The Mach number range extended from about 0.60 to 1.18, with corresponding Reynolds numbers ranging from about 0.75 x 10(6) to 0.95 x 10(6). The angle of attack range was from -10 degrees to approximately 32 degrees.The effects on drag and lift-drag ratio of a variation in sweep angle from -14.03 degrees to 45 degrees with respect to the quarter-chord line for wings of 3-percent-chord thickness was found to be small in comparison to the effects of a variation in thickness from 2 percent chord to 4.5 percent chord for wings with 14.03 degree sweepback. For the range of variables considered, variations in plan form were considerably more important with regard to longitudinal stability characteristics than the variations in thickness. For the series of basic wings having an aspect ratio of 4, the most hearly linear pitching-moment characteristics were obtained with 26.57 degree of sweepback of the quarter-chord line. However, for the modified series of wings (obtained by clipping the tips of the original wings parallel to the plane of symmetry to give an aspect ratio of 3 and a taper ratio of 0.143), the most nearly linear pitching-moment characteristics were obtained with 36.87 degrees of sweepback. By decreasing the thickness-to-chord ratios from 0.03 to 0.02, a large increase in lift-curve slope was obtained for both the basic and modified wings. All of the wings of both series had fairly large inward shifts of the lateral center-of-pressure location (indicative of tip stalling) with increasing lift coefficient, except those wings having minimum sweepback angles.

  9. A low aspect ratio tokamak transmutation system

    NASA Astrophysics Data System (ADS)

    Qiu, L. J.; Wu, Y. C.; Xiao, B. J.; Xu, Q.; Huang, Q. Y.; Wu, B.; Chen, Y. X.; Xu, W. N.; Chen, Y. P.; Liu, X. P.

    2000-03-01

    A low aspect ratio tokamak transmutation system is proposed as an alternative application of fusion energy on the basis of a review of previous studies. This system includes: (1) a low aspect ratio tokamak as fusion neutron driver, (2) a radioactivity-clean nuclear power system as blanket, and (3) a novel concept of liquid metal centre conductor post as part of the toroidal field coils. In the conceptual design, a driver of 100 MW fusion power under 1 MW/m2 neutron wall loading can transmute the amount of high level waste (including minor actinides and fission products) produced by ten standard pressurized water reactors of 1 GW electrical power output. Meanwhile, the system can produce tritium on a self-sustaining basis and an output of about 2 GW of electrical energy. After 30 years of operation, the biological hazard potential level of the whole system will decrease by two orders of magnitude.

  10. On virial analysis at low aspect ratio

    SciTech Connect

    Bongard, Michael W.; Barr, Jayson L.; Fonck, Raymond J.; Reusch, Joshua A.; Thome, Kathreen E.

    2016-07-28

    The validity of virial analysis to infer global MHD equilibrium poloidal beta βp and internal inductance ℓi from external magnetics measurements is examined for low aspect ratio configurations with A < 2. Numerical equilibrium studies at varied aspect ratio are utilized to validate the technique at finite aspect ratio. The effect of applying high-A approximations to low-A experimental data is quantified and demonstrates significant over-estimation of stored energy (factors of 2–10) in spherical tokamak geometry. Experimental approximations to equilibrium-dependent volume integral terms in the analysis are evaluated at low-A. Highly paramagnetic configurations are found to be inadequately represented through the virial mean radius parameter RT. Alternate formulations for inferring βp and ℓi that are independent of RT to avoid this difficulty are presented for the static isotropic limit. Lastly, these formulations are suitable for fast estimation of tokamak stored energy components at low aspect ratio using virial analysis.

  11. On virial analysis at low aspect ratio

    DOE PAGES

    Bongard, Michael W.; Barr, Jayson L.; Fonck, Raymond J.; ...

    2016-07-28

    The validity of virial analysis to infer global MHD equilibrium poloidal beta βp and internal inductance ℓi from external magnetics measurements is examined for low aspect ratio configurations with A < 2. Numerical equilibrium studies at varied aspect ratio are utilized to validate the technique at finite aspect ratio. The effect of applying high-A approximations to low-A experimental data is quantified and demonstrates significant over-estimation of stored energy (factors of 2–10) in spherical tokamak geometry. Experimental approximations to equilibrium-dependent volume integral terms in the analysis are evaluated at low-A. Highly paramagnetic configurations are found to be inadequately represented through themore » virial mean radius parameter RT. Alternate formulations for inferring βp and ℓi that are independent of RT to avoid this difficulty are presented for the static isotropic limit. Lastly, these formulations are suitable for fast estimation of tokamak stored energy components at low aspect ratio using virial analysis.« less

  12. On virial analysis at low aspect ratio

    SciTech Connect

    Bongard, Michael W.; Barr, Jayson L.; Fonck, Raymond J.; Reusch, Joshua A.; Thome, Kathreen E.

    2016-07-28

    The validity of virial analysis to infer global MHD equilibrium poloidal beta βp and internal inductance ℓi from external magnetics measurements is examined for low aspect ratio configurations with A < 2. Numerical equilibrium studies at varied aspect ratio are utilized to validate the technique at finite aspect ratio. The effect of applying high-A approximations to low-A experimental data is quantified and demonstrates significant over-estimation of stored energy (factors of 2–10) in spherical tokamak geometry. Experimental approximations to equilibrium-dependent volume integral terms in the analysis are evaluated at low-A. Highly paramagnetic configurations are found to be inadequately represented through the virial mean radius parameter RT. Alternate formulations for inferring βp and ℓi that are independent of RT to avoid this difficulty are presented for the static isotropic limit. Lastly, these formulations are suitable for fast estimation of tokamak stored energy components at low aspect ratio using virial analysis.

  13. Omniclassical Diffusion in Low Aspect Ratio Tokamaks

    SciTech Connect

    H.E. Mynick; R.B. White; D.A. Gates

    2004-03-19

    Recently reported numerical results for axisymmetric devices with low aspect ratio A found radial transport enhanced over the expected neoclassical value by a factor of 2 to 3. In this paper, we provide an explanation for this enhancement. Transport theory in toroidal devices usually assumes large A, and that the ratio B{sub p}/B{sub t} of the poloidal to the toroidal magnetic field is small. These assumptions result in transport which, in the low collision limit, is dominated by banana orbits, giving the largest collisionless excursion of a particle from an initial flux surface. However in a small aspect ratio device one may have B{sub p}/B{sub t} {approx} 1, and the gyroradius may be larger than the banana excursion. Here, we develop an approximate analytic transport theory valid for devices with arbitrary A. For low A, we find that the enhanced transport, referred to as omniclassical, is a combination of neoclassical and properly generalized classical effects, which become dominant in the low-A, B{sub p}/B{sub t} {approx} 1 regime. Good agreement of the analytic theory with numerical simulations is obtained.

  14. A parametric study of planform and aeroelastic effects on aerodynamic center, alpha- and q-stability derivatives. Appendix E: Procedures used to determine the structural representation for idealized low aspect ratio two spar fighter wings

    NASA Technical Reports Server (NTRS)

    Roskam, J.; Lan, C.; Smith, H.; Gibson, G.

    1972-01-01

    An explanation is presented of the method used to locate the elastic axis and the method to determine the EI and GJ distributions along the elastic axes of wings with a 2-spar (front and rear) construction or a single torque-box construction.

  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. An Airplane Design having a Wing with Fuselage Attached to Each Tip

    NASA Technical Reports Server (NTRS)

    Spearman, Leroy M.

    2001-01-01

    This paper describes the conceptual design of an airplane having a low aspect ratio wing with fuselages that are attached to each wing tip. The concept is proposed for a high-capacity transport as an alternate to progressively increasing the size of a conventional transport design having a single fuselage with cantilevered wing panels attached to the sides and tail surfaces attached at the rear. Progressively increasing the size of conventional single body designs may lead to problems in some area's such as manufacturing, ground-handling and aerodynamic behavior. A limited review will be presented of some past work related to means of relieving some size constraints through the use of multiple bodies. Recent low-speed wind-tunnel tests have been made of models representative of the inboard-wing concept. These models have a low aspect ratio wing with a fuselage attached to each tip. Results from these tests, which included force measurements, surface pressure measurements, and wake surveys, will be presented herein.

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

    NASA Technical Reports Server (NTRS)

    Brown, Clarence A , Jr

    1957-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Brown, C. A., Jr.

    1957-01-01

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

  19. Equilibrium studies for low-aspect-ratio torsatrons

    SciTech Connect

    Garcia, L.; Carreras, B.A.; Harris, J.H.; Hicks, H.R.; Lynch, V.E.

    1983-10-01

    Three-dimensional MHD equilibrium studies have been done for low aspect ratio torsatrons. These studies cover continuous-coil and modular torsatrons. The equilibrium properties improve with aspect ratio, as is expected. The role played by the vertical field has also been investigated. Outward shifts cause the flux surfaces to deteriorate. Inward shifts improve the flux surfaces at the expense of stability properties.

  20. Longitudinal Stability and Control Characteristics as Determined by the Rocket-Model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-Aspect-Ratio Wing Having Trailing-Edge Flap Controls for a Mach Number Range of 0.7 to 1.8

    NASA Technical Reports Server (NTRS)

    Baber, H. T., Jr.; Moul, M. T.

    1955-01-01

    Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle-of-attack range of this test (0 deg to 8 deg ). The aerodynamic-center location for angles of attack near 5 deg remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near O deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of O deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle-of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

  1. Longitudinal Stability and Control Characteristics as Determined by the Rocket-model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-aspect-ratio Wing Having Trailing-edge Flap Controls for a Mach Number Range of 0.7 to 1.

    NASA Technical Reports Server (NTRS)

    Baber, Hal T , Jr; Moul, Martin T

    1955-01-01

    Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle - of-attack range of this test (0 deg to 8 deg). The aerodynamic-center location for angles of attack near 50 remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near 0 deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of 0 deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle -of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

  2. Evolution of turbulent jets in low aspect ratio containers

    NASA Astrophysics Data System (ADS)

    Pol, S.; Nath, C.; Gest, D.; Voropayev, S.; Fernando, H. J. S.; Webb, S.

    2009-11-01

    The evolution of homogeneous and buoyant turbulent jets released into a low aspect ratio (width/height) container was investigated experimentally using PIV, MSCT probing and digital imaging. The motivation was to understand mixing process occurring in U.S. Strategic Petroleum Reserves (SPR), where crude oil is stored in salt caverns of low aspect ratio. During maintenance or filling, oil is introduced as a jet from the top of the caverns. This study is focussed on mean and turbulent flow characteristics as well as global flow instability and periodic oscillations intrinsic to jets in low aspect ratio containers. Scaling arguments were advanced for salient flow parameters, which included the characteristic length (container width D) and velocity (for homogeneous jets, J^1/2D, where J is the momentum flux at the jet exit) scales. For buoyant jets, the buoyancy flux B needs to be introduced as an additional parameter. Such jet flows do not reach a steady state, but bifurcate periodically with a frequency scale J^1/2/ D^2 while enhancing global mixing.

  3. Compressional Alfvin Eigenmode Dispersion in Low Aspect Ratio Plasmas

    SciTech Connect

    N.N. Gorelenkov; C.Z. Cheng; E. Fredrickson

    2002-01-29

    Recent observations of new fast ion beam driven instabilities in MHz frequency range in National Spherical Torus experiments (NSTX) are suggested to be Compressional Alfvin Eigenmodes (CAEs). A new theory of CAEs applicable to low aspect ratio toroidal plasmas is developed based on the ballooning representation for the poloidal dependence of the perturbed quantities. In agreement with observations, the analytical theory predicts that CAEs are discrete modes with frequencies correlated with the characteristic Alfvin velocity of the plasma. Plasma equilibrium structure is essential to determine accurately the dispersion of CAEs. The mode structure is localized in both the minor radius and the poloidal directions on the low magnetic field side.

  4. Dean Flow Dynamics in Low-Aspect Ratio Spiral Microchannels

    PubMed Central

    Nivedita, Nivedita; Ligrani, Phillip; Papautsky, Ian

    2017-01-01

    A wide range of microfluidic cell-sorting devices has emerged in recent years, based on both passive and active methods of separation. Curvilinear channel geometries are often used in these systems due to presence of secondary flows, which can provide high throughput and sorting efficiency. Most of these devices are designed on the assumption of two counter rotating Dean vortices present in the curved rectangular channels and existing in the state of steady rotation and amplitude. In this work, we investigate these secondary flows in low aspect ratio spiral rectangular microchannels and define their development with respect to the channel aspect ratio and Dean number. This work is the first to experimentally and numerically investigate Dean flows in microchannels for Re > 100, and show presence of secondary Dean vortices beyond a critical Dean number. We further demonstrate the impact of these multiple vortices on particle and cell focusing. Ultimately, this work offers new insights into secondary flow instabilities for low-aspect ratio, spiral microchannels, with improved flow models for design of more precise and efficient microfluidic devices for applications such as cell sorting and micromixing. PMID:28281579

  5. Reduced MHD Equations For Low Aspect Ratio Tokamaks

    NASA Astrophysics Data System (ADS)

    Callen, J. D.; Hegna, C. C.; Kruger, S. E.

    1997-11-01

    The usual derivations [1] of reduced MHD equations for describing tokamak plasmas use a large aspect ratio expansion to remove, to lowest order, the compressional Alfven wave response that enforces the axisymmetric ideal MHD (Grad-Shafranov) equilibrium. For very low aspect ratio tokamaks such an approach is invalid, even though [2] the radially localized compressional Alfven waves that enforce ideal MHD equilibrium still have higher frequencies than the shear Alfven waves that appear in reduced MHD descriptions. We are beginning to explore and will discuss a multiple time scale approach that does not explicitly require a large aspect ratio expansion in order to develop a reduced MHD description appropriate for low aspect ratio tokamak plasmas. The relationship of this approach to geometric-properties-based formulations of the ideal MHD energy principle [3] will also be discussed. [1] H.R. Strauss, Phys. Fluids 19, 134 (1976); 20, 1354 (1977). [2] R.D. Hazeltine and J.D. Meiss, Plasma Confinement (Addison-Wesley, Redwood City, CA, 1992), Chapter 7. [3] R.L. Dewar, D.A. Monticello, W.N.-C. Sy, Phys. Fluids 27, 1723 (1984); J.M. Greene, Phys. Plasmas 3, 8(1996).

  6. Dean Flow Dynamics in Low-Aspect Ratio Spiral Microchannels

    NASA Astrophysics Data System (ADS)

    Nivedita, Nivedita; Ligrani, Phillip; Papautsky, Ian

    2017-03-01

    A wide range of microfluidic cell-sorting devices has emerged in recent years, based on both passive and active methods of separation. Curvilinear channel geometries are often used in these systems due to presence of secondary flows, which can provide high throughput and sorting efficiency. Most of these devices are designed on the assumption of two counter rotating Dean vortices present in the curved rectangular channels and existing in the state of steady rotation and amplitude. In this work, we investigate these secondary flows in low aspect ratio spiral rectangular microchannels and define their development with respect to the channel aspect ratio and Dean number. This work is the first to experimentally and numerically investigate Dean flows in microchannels for Re > 100, and show presence of secondary Dean vortices beyond a critical Dean number. We further demonstrate the impact of these multiple vortices on particle and cell focusing. Ultimately, this work offers new insights into secondary flow instabilities for low-aspect ratio, spiral microchannels, with improved flow models for design of more precise and efficient microfluidic devices for applications such as cell sorting and micromixing.

  7. All Metal Iron Core For A Low Aspect Ratio Tokamak

    SciTech Connect

    D.A. Gates, C. Jun, I. Zatz, A. Zolfaghari

    2010-06-02

    A novel concept for incorporating a iron core transformer within a axisymmetric toroidal plasma containment device with a high neutron flux is described. This design enables conceptual design of low aspect ratio devices which employ standard transformer-driven plasma startup by using all-metal high resistance separators between the toroidal field windings. This design avoids the inherent problems of a multiturn air core transformer which will inevitably suffer from strong neutron bombardment and hence lose the integrity of its insulation, both through long term material degradation and short term neutron- induced conductivity.. A full 3-dimensional model of the concept has been developed within the MAXWELL program and the resultant loop voltage calculated. The utility of the result is found to be dependent on the resistivity of the high resistance separators. Useful loop voltage time histories have been obtained using achievable resistivities.

  8. Modular low aspect ratio-high beta torsatron

    DOEpatents

    Sheffield, George V.; Furth, Harold P.

    1984-02-07

    A fusion reactor device in which the toroidal magnetic field and at least a portion of the poloidal magnetic field are provided by a single set of modular coils. The coils are arranged on the surface of a low aspect ratio toroid in planes having the cylindrical coordinate relationship .phi.=.phi..sub.i +kz where k is a constant equal to each coil's pitch and .phi..sub.i is the toroidal angle at which the i'th coil intersects the z=o plane. The device may be described as a modular, high beta torsation whose screw symmetry is pointed along the systems major (z) axis. The toroid defined by the modular coils preferably has a racetrack minor cross section. When vertical field coils and preferably a toroidal plasma current are provided for magnetic field surface closure within the toroid, a vacuum magnetic field of racetrack shaped minor cross section with improved stability and beta valves is obtained.

  9. Modular low-aspect-ratio high-beta torsatron

    DOEpatents

    Sheffield, G.V.

    1982-04-01

    A fusion-reactor device is described which the toroidal magnetic field and at least a portion of the poloidal magnetic field are provided by a single set of modular coils. The coils are arranged on the surface of a low-aspect-ratio toroid in planed having the cylindrical coordinate relationship phi = phi/sub i/ + kz, where k is a constant equal to each coil's pitch and phi/sub i/ is the toroidal angle at which the i'th coil intersects the z = o plane. The toroid defined by the modular coils preferably has a race track minor cross section. When vertical field coils and, preferably, a toroidal plasma current are provided for magnetic-field-surface closure within the toroid, a vacuum magnetic field of racetrack-shaped minor cross section with improved stability and beta valves is obtained.

  10. Prospects and status of low-aspect-ratio tokamaks

    SciTech Connect

    Peng, Y.K.M.

    1994-12-31

    The prospects for the low-aspect-ratio (A) tokamak to fulfill the requirements of viable fusion power plants are considered relative to the present status in data and modeling. Desirable physics and design features for an attractive Blanket Test Facility and power reactors are estimated for low-A tokamaks based on calculations improved with the latest data from small pioneering experiments. While these experiments have confirmed some of the recent predictions for low-A, they also identify the remaining issues that require verification before reliable projections can be made for these deuterium-tritium applications. The results show that the low-A regime of small size, modest field, and high current offers a path complementary to the standard and high A tokamaks in developing the full potential of fusion power.

  11. Overview of the Pegasus Extremely Low-Aspect Ratio Tokamak

    NASA Astrophysics Data System (ADS)

    Fonck, R.; Garstka, G.; Intrator, T.; Lewicki, B.; Thorson, T.; Toonen, R.; Tritz, K. L.; White, B.; Winz, G.

    1996-11-01

    Pegasus is a new experiment designed to explore the potential of Extremely Low Aspect Ratio Tokamaks (ELART) at very high toroidal β. Ohmic induction for plasma startup will be followed by ohmic sustainment initially and noninductive RF current drive in the future. Plasma parameters are projected to be Ip <= 0.3 MA, Bt < 0.2T, <β> ≈ 5-40 % or higher, A=1.1-2, R=0.2-0.4 m, and P_RF <= 2MW. Goals of the program include: demonstrate high-β spherical tokamak operation in the near term; examine the stability, n=0 stability properties at high elongation and low- A, confinement and scaling characteristics at A <= 1.25; and extend high power ST operation to the extrema of A <= 1.1. Hollow current profiles should be accessible in Pegasus using a fast current ramp during formation plus off-axis FWCD in the longer term. Recent changes to the design include: increased vacuum vessel height to allow for divertor operation with an internal X-point plus increased accessible elongations (i.,e., κ <= 3.7 at A = 1.25); additional coils for X-point control; and elimination of toroidal gaps in favor of a resistive vacuum vessel. Initial operation will emphasize ohmic access to high- β, followed by high power RF heating.

  12. Stability of local modes in low-aspect-ratio stellarators

    SciTech Connect

    Garcia, L.; Carreras, B.A.; Dominguez, N.

    1989-01-01

    The studies of MHD stability properties of stellarator-type configurations are complicated by the three-dimensional (3-D) nature of the equilibrium of these configurations. Usually, the unstable modes are pressure-driven and very localized around the rational surface. The 3-D Mercier stability criterion for local instabilities is therefore adequate for most of the stability studies. The method has been implemented in different 3-D equilibrium codes and is widely employed in stability calculations for stellarators. We use an average method based on vacuum flux coordinates to study the equilibrium of low-aspect-ratio configurations. The specific vacuum flux coordinates (/rho/,theta,/var phi/) are those described by Boozer. The radial variable is /rho/, while theta and /var phi/ are, respectively, poloidal and toroidal-like angles. The Jacobian is D = B/sub v//sup 2///rho/B/sub 0/g/sub B/. The only vacuum magnetic field covariant component different from zero is B/sub /var phi// = g/sub B/, while B/sub 0/..pi../rho//sup 2/ is the toroidal flux. 7 refs., 1 fig.

  13. The Aneutronic Rodless Ultra Low Aspect Ratio Tokamak

    NASA Astrophysics Data System (ADS)

    Ribeiro, Celso

    2016-10-01

    The replacement of the metal centre-post in spherical tokamaks (STs) by a plasma centre-post (PCP, the TF current carrier) is the ideal scenario for a ST reactor. A simple rodless ultra low aspect-ratio tokamak (RULART) using a screw-pinch PCP ECR-assisted with an external solenoid has been proposed in the most compact RULART [Ribeiro C, SOFE-15]. There the solenoid provided the stabilizing field for the PCP and the toroidal electrical field for the tokamak start-up, which will stabilize further the PCP, acting as stabilizing closed conducting surface. Relative low TF will be required. The compactness (high ratio of plasma-spherical vessel volume) may provide passive stabilization and easier access to L-H mode transition. It is presented here: 1) stability analysis of the PCP (initially MHD stable due to the hollow J profile); 2) tokamak equilibrium simulations, and 3) potential use for aneutronic reactions studies via pairs of proton p and boron 11B ion beams in He plasmas. The beams' line-of-sights sufficiently miss the sources of each other, thus allowing a near maximum relative velocities and reactivity. The reactions should occur close to the PCP mid-plane. Some born alphas should cross the PCP and be dragged by the ion flow (higher momentum exchange) towards the anode but escape directly to a direct electricity converter. Others will reach evenly the vessel directly or via thermal diffusion (favourable heating by the large excursion 2a), leading to the lowest power wall load possible. This might be a potential hybrid direct-steam cycle conversion reactor scheme, nearly aneutronic, and with no ash or particle retention problems, as opposed to the D-T thermal reaction proposals.

  14. Buffet characteristics of the F-8 supercritical wing airplane

    NASA Technical Reports Server (NTRS)

    Deangelis, V. M.; Monaghan, R. C.

    1977-01-01

    The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane's structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator.

  15. Numerical Investigation of a Low Aspect Ratio Flat Plate

    DTIC Science & Technology

    2011-05-19

    M.R., ―High-Fidelity Simulations of Transitional Flow Over Pitching Airfoils,‖ AIAA Paper 2009- 3693 . Henshaw, W.D., and Petersson, N. A., ―A Split...Amplitude Pitch Problem: Airfoils, Plates and Wings,‖ AIAA Paper 2009-3686, 2009. OL, M.V., Altman, A., Eldredge, J.D., Garmann, D ., and Lian, Y...incompressible flow solver. AIAA J. 39,253 -260, 2001. Thomas, P. D ., and Lombard, C. K., ―Geometric Conservation Law and its Application to Flow

  16. Development of tailless and all-wing gliders and airplanes

    NASA Technical Reports Server (NTRS)

    Lademann, Robert W E

    1932-01-01

    Tailless airplanes are characterized by having all their control surfaces, especially the elevator, incorporated in the wings. This paper provides a discussion of the history of their development and current state of development.

  17. Development of a Composite Tailoring Technique for Airplane Wing

    NASA Technical Reports Server (NTRS)

    Chattopadhyay, Aditi; Jha, Ratneshwar

    1996-01-01

    Development of a new composite beam modeling technique to represent the principal load-carrying member in the wing is reported along with the development of a formal design optimization procedure to investigate the effect of composite tailoring on aeroelastic stability and structural characteristics of airplane wings. The developed procedure is used to perform design optimization studies on realistic airplane configurations to investigate the various aeroelastic/structural/dynamic design issues.

  18. Effects of Winglets on the Drag of a Low-Aspect-Ratio Configuration

    NASA Technical Reports Server (NTRS)

    Smith, Leigh Ann; Campbell, Richard L.

    1996-01-01

    A wind-tunnel investigation has been performed to determine the effect of winglets on the induced drag of a low-aspect-ratio wing configuration at Mach numbers between 0.30 and 0.85 and a nominal angle-of-attack range from -2 deg to 20 deg. Results of the tests at the cruise lift coefficient showed significant increases in lift-drag ratio for the winglet configuration relative to a wing-alone configuration designed for the same lift coefficient and Mach number. Further, even larger increases in lift-drag ratio were observed at lift coefficients above the design value at all Mach numbers tested. The addition of these winglets had a negligible effect on the static lateral-directional stability characteristics of the configuration. No tests were made to determine the effect of these winglets at supersonic Mach numbers, where increases in drag caused by winglets might be more significant. Computational analyses were also performed for the two configurations studied. Linear and small-disturbance formulations were used. The codes were found to give reasonable performance estimates sufficient for predicting changes of this magnitude.

  19. Numerical Simulations of Natural and Actuated Flow over a 3D, Low-Aspect-Ratio Airfoil

    NASA Astrophysics Data System (ADS)

    Brès, Guillaume; Williams, David; Colonius, Tim

    2010-11-01

    Numerical simulations of the unsteady flow over a low-aspect-ratio, low Reynolds number semi-circular planform wing are performed using Lattice Boltzmann method. The simulations exactly match the flow conditions and the detailed geometry from previous wind-tunnel experiments, including the flow actuators installed internally along the leading edge of the wing. To reproduce the pulsed-blowing actuation used in the experiment, a single pulsed square wave forcing is imposed in the simulations as a mass flow boundary condition in the actuators. Three angles of attack, with the active flow control both on and off, are investigated. For both mean and unsteady lift and drag, the numerical simulations show good agreement with the experiments. In particular, the transient increase in lift after the forcing is turned off is well captured in the simulations. Both PIV measurements and transient numerical results indicate that this behavior is associated with the advection of large vortical structures generated by the flow actuation at the leading edge.

  20. Collisional Transport in a Low Aspect Ratio Tokamak -- Beyond the Drift Kinetic Formalism

    SciTech Connect

    D.A. Gates; R.B. White

    2004-01-28

    Calculations of collisional thermal and particle diffusivities in toroidal magnetic plasma confinement devices order the toroidal gyroradius to be small relative to the poloidal gyroradius. This ordering is central to what is usually referred to as neoclassical transport theory. This ordering is incorrect at low aspect ratio, where it can often be the case that the toroidal gyroradius is larger than the poloidal gyroradius. We calculate the correction to the particle and thermal diffusivities at low aspect ratio by comparing the diffusivities as determined by a full orbit code (which we refer to as omni-classical diffusion) with those from a gyroaveraged orbit code (neoclassical diffusion). In typical low aspect ratio devices the omni-classical diffusion can be up to 2.5 times the calculated neoclassical value. We discuss the implications of this work on the analysis of collisional transport in low aspect ratio magnetic confinement experiments.

  1. Designs Studies of Low-Aspect Ratio Quasi-Omnigenous Stellarators

    SciTech Connect

    Batchelor, D.B.; Carreras, B.A.; Hirshman, S.P.; Lynch, V.E.; Sanchez, R.; Spong, D.A.; Whitson, J.C.

    1998-10-15

    Significant progress has been made in the development of new modest-size compact stellarator devices that could test optimization principles for the design of a more attractive reactor. These are 3 and 4 field period low-aspect-ratio quasi-omnigenous (QO) stellarators based on an optimization method that targets improved confinement, stability, ease of coil design, low-aspect-ratio, and low bootstrap current.

  2. Experiments on Low Aspect Ratio Hydroplanes to Measure Lift under Static and Dynamic Conditions

    DTIC Science & Technology

    1990-03-01

    ON LOW ASPECT RATIO HYDROPLANES TO MEASURE LIFT UNDER STATIC AND DYNAMIC CONDITIONS By B Ward A R J M Lloyd Summary This Technical Memorandum...describes experiments in the Circulating Water Channel to measure lift forces on low aspect ratio hydroplanes under static and dynamic conditions. Empirical...6 11. Conclusions 6 References 7 Figure 1. Three NACA 0020 Hydroplanes Figure 2. Hydroplane on Servo Figures 3 and 4. Experiment in Circulating Water

  3. Effects of Large Wing-Tip Masses on Oscillatory Stability of Wing Bending Coupled with Airplane Pitch

    NASA Technical Reports Server (NTRS)

    Higdon, Donald T.

    1959-01-01

    An examination of oscillatory stability for a straight-winged airplane with large concentrated wing-tip masses was made using wing-bending and airplane-pitching degrees of freedom and considering only quasi-steady aerodynamic forces. It was found that instability caused by coupling of airplane pitching and wing bending occurred for large ratios of effective wing-tip mass to total airplane mass and for coupled wing-bending frequencies near or below the uncoupled pitching frequency. Boundaries for this instability are given in terms of two quantities: (1) the ratio of effective tip mass to airplane mass, which can be estimated, and (2) the ratio of the coupled bending frequency to the uncoupled pitch frequency, which can be measured in flight. These boundaries are presented for various values of several airplane parameters.

  4. Theory and Observations of High Frequency Alfven Eigenmodes in Low Aspect Ratio Plasma

    SciTech Connect

    N.N. Gorelenkov; E. Fredrickson; E. Belova; C.Z. Cheng; D. Gates; S. Kaye; R. White

    2003-06-27

    New observations of sub-cyclotron frequency instability in low aspect ratio plasma in National Spherical Torus Experiments (NSTX) are reported. The frequencies of observed instabilities correlate with the characteristic Alfven velocity of the plasma. A theory of localized Compressional Alfven Eigenmodes (CAE) and Global shear Alfven Eigenmodes (GAE) in low aspect ratio plasma is presented to explain the observed high frequency instabilities. CAE's/GAE's are driven by the velocity space gradient of energetic super-Alfvenic beam ions via Doppler shifted cyclotron resonances. One of the main damping mechanisms of GAE's, the continuum damping, is treated perturbatively within the framework of ideal MHD. Properties of these cyclotron instabilities ions are presented.

  5. Stresses Produced in Airplane Wings by Gusts

    NASA Technical Reports Server (NTRS)

    Kussner, Hans Georg

    1932-01-01

    Accurate prediction of gust stress being out of the question because of the multiplicity of the free air movements, the exploration of gust stress is restricted to static method which must be based upon: 1) stress measurements in free flight; 2) check of design specifications of approved type airplanes. With these empirical data the stress must be compared which can be computed for a gust of known intensity and structure. This "maximum gust" then must be so defined as to cover the whole ambit of empiricism and thus serve as prediction for new airplane designs.

  6. 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.

  7. A Novel Demountable TF Joint Design for Low Aspect Ratio Spherical Torus Tokamaks

    SciTech Connect

    Robert D. Woolley

    2009-06-11

    A novel shaped design for the radial conductors and demountable electrical joints connecting inner and outer legs of copper TF system conductors in low aspect ratio tokamaks is described and analysis results are presented. Specially shaped designs can optimize profiles of electrical current density, magnetic force, heating, and mechanical stress.

  8. A Novel Demountable TF Joint Design for Low Aspect Ratio Spherical Torus Tokamaks

    SciTech Connect

    R.D. Woolley

    2009-05-29

    A novel shaped design for the radial conductors and demountable electrical joints connecting inner and outer legs of copper TF system conductors in low aspect ratio tokamaks is described and analysis results are presented. Specially shaped designs can optimize profiles of electrical current density, magnetic force, heating, and mechanical stress.

  9. Fluid-Structure Interaction of Oscillating Low Aspect Ratio Wings at Low Reynolds Numbers

    DTIC Science & Technology

    2010-03-01

    illustration of the rig is provided in Figure 1. Flow measurements were obtained using a Digital Particle Image Velocimetry (DPIV) system. The TSI...increased frequency, even so, no thrust is produced within the frequency range tested. The phase-averaged vorticity and velocity magnitude, at the...Reynolds numbers tested. The switch from drag to thrust occurs within the operational range of Strouhal numbers, at Stc≈1.1-1.25. Figure 11 shows the

  10. Modulation of rotation-induced lift force for cell filtration in a low aspect ratio microchannel

    PubMed Central

    Zhou, Jian; Giridhar, Premkumar Vummidi; Kasper, Susan; Papautsky, Ian

    2014-01-01

    Cell filtration is a critical step in sample preparation in many bioapplications. Herein, we report on a simple, filter-free, microfluidic platform based on hydrodynamic inertial migration. Our approach builds on the concept of two-stage inertial migration which permits precise prediction of microparticle position within the microchannel. Our design manipulates equilibrium positions of larger microparticles by modulating rotation-induced lift force in a low aspect ratio microchannel. Here, we demonstrate filtration of microparticles with extreme efficiency (>99%). Using multiple prostate cell lines (LNCaP and human prostate epithelial tumor cells), we show filtration from spiked blood, with 3-fold concentration and >83% viability. Results of a proliferation assay show normal cell division and suggest no negative effects on intrinsic properties. Considering the planar low-aspect-ratio structure and predictable focusing, we envision promising applications and easy integration with existing lab-on-a-chip systems. PMID:25379097

  11. High-β equilibrium and ballooning stability of the low aspect ratio CNT stellarator

    DOE PAGES

    Hammond, K. C.; Lazerson, S. A.; Volpe, F. A.

    2017-04-07

    In the paper, the existence and ballooning-stability of low aspect ratio stellarator equilibria is predicted for the Columbia Neutral Torus (CNT) with the aid of 3D numerical tools. In addition to having a low aspect ratio, CNT is characterized by a low magnetic field and small plasma volume. Also, highly overdense plasmas were recently heated in CNT by means of microwaves. These characteristics suggest that CNT might attain relatively high values of plasma beta and thus be of use in the experimental study of stellarator stability to high-beta instabilities such as ballooning modes. As a first step in that direction,more » here the ballooning stability limit is found numerically. Depending on the particular magnetic configuration we expect volume-averaged β limits in the range 0.9%–3.0%, and possibly higher, and observe indications of a second region of ballooning stability. As the aspect ratio is reduced, stability is found to increase in some configurations and decrease in others. Energy-balance estimates using stellarator scaling laws indicate that the lower β limit may be attainable with overdense heating at powers of 40 to 100 kW. The present study serves the additional purpose of testing VMEC and other stellarator codes at high values of β and at low aspect ratios. For this reason, the study was carried out both for free boundary, for maximum fidelity to experiment, as well as with a fixed boundary, as a numerical test.« less

  12. Particle Trajectory-Dependent Ionic Current Blockade in Low-Aspect-Ratio Pores.

    PubMed

    Tsutsui, Makusu; He, Yuhui; Yokota, Kazumichi; Arima, Akihide; Hongo, Sadato; Taniguchi, Masateru; Washio, Takashi; Kawai, Tomoji

    2016-01-26

    Resistive pulse sensing with nanopores having a low thickness-to-diameter aspect-ratio structure is expected to enable high-spatial-resolution analysis of nanoscale objects in a liquid. Here we investigated the sensing capability of low-aspect-ratio pore sensors by monitoring the ionic current blockades during translocation of polymeric nanobeads. We detected numerous small current spikes due to partial occlusion of the pore orifice by particles diffusing therein reflecting the expansive electrical sensing zone of the low-aspect-ratio pores. We also found wide variations in the ion current line-shapes in the particle capture stage suggesting random incident angle of the particles drawn into the pore. In sharp contrast, the ionic profiles were highly reproducible in the post-translocation regime by virtue of the spatial confinement in the pore that effectively constricts the stochastic capture dynamics into a well-defined ballistic motion. These results, together with multiphysics simulations, indicate that the resistive pulse height is highly dependent on the nanoscopic single-particle trajectories involved in ultrathin pore sensors. The present finding indicates the importance of regulating the translocation pathways of analytes in low-aspect-ratio pores for improving the discriminability toward single-bioparticle tomography in liquid.

  13. Low aspect ratio micropores for single-particle and single-cell analysis.

    PubMed

    Goyal, Gaurav; Mulero, Rafael; Ali, Jamel; Darvish, Armin; Kim, Min Jun

    2015-05-01

    This paper describes microparticle and bacterial translocation studies using low aspect ratio solid-state micropores. Micropores, 5 μm in diameter, were fabricated in 200 nm thick free-standing silicon nitride membranes, resulting in pores with an extremely low aspect ratio, nominally 0.04. For microparticle translocation experiments, sulfonated polystyrene microparticles and magnetic microbeads in size range of 1-4 μm were used. Using the microparticle translocation characteristics, we find that particle translocations result in a change only in the pore's geometrical resistance while the access resistance remains constant. Furthermore, we demonstrate the ability of our micropore to probe high-resolution shape information of translocating analytes using concatenated magnetic microspheres. Distinct current drop peaks were observed for each microsphere of the multibead architecture. For bacterial translocation experiments, nonflagellated Escherichia coli (strain HCB 5) and wild type flagellated Salmonella typhimurium (strain SJW1103) were used. Distinct current signatures for the two bacteria were obtained and this difference in translocation behavior was attributed to different surface protein distributions on the bacteria. Our findings may help in developing low aspect ratio pores for high-resolution microparticle characterization and single-cell analysis.

  14. 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.

  15. Airplane wing leading edge variable camber flap

    NASA Technical Reports Server (NTRS)

    Cole, J. B.

    1980-01-01

    The invention and design of an aerodynamic high lift device which provided a solution to an aircraft performance problem are described. The performance problem of converting a high speed cruise airfoil into a low speed aerodynamic shape that would provide landing and take-off characteristics superior to those available with contemporary high lift devices are addressed. The need for an improved wing leading edge device that would complement the high lift performance of a triple slotted trailing edge flap is examined. The mechanical and structural aspects of the variable camber flap are discussed and the aerodynamic performance aspects only as they relate to the invention and design of the device are presented.

  16. Small-scale fixed wing airplane software verification flight test

    NASA Astrophysics Data System (ADS)

    Miller, Natasha R.

    The increased demand for micro Unmanned Air Vehicles (UAV) driven by military requirements, commercial use, and academia is creating a need for the ability to quickly and accurately conduct low Reynolds Number aircraft design. There exist several open source software programs that are free or inexpensive that can be used for large scale aircraft design, but few software programs target the realm of low Reynolds Number flight. XFLR5 is an open source, free to download, software program that attempts to take into consideration viscous effects that occur at low Reynolds Number in airfoil design, 3D wing design, and 3D airplane design. An off the shelf, remote control airplane was used as a test bed to model in XFLR5 and then compared to flight test collected data. Flight test focused on the stability modes of the 3D plane, specifically the phugoid mode. Design and execution of the flight tests were accomplished for the RC airplane using methodology from full scale military airplane test procedures. Results from flight test were not conclusive in determining the accuracy of the XFLR5 software program. There were several sources of uncertainty that did not allow for a full analysis of the flight test results. An off the shelf drone autopilot was used as a data collection device for flight testing. The precision and accuracy of the autopilot is unknown. Potential future work should investigate flight test methods for small scale UAV flight.

  17. The right wing of the LEFT airplane

    NASA Technical Reports Server (NTRS)

    Powell, Arthur G.

    1987-01-01

    The NASA Leading-Edge Flight Test (LEFT) program addressed the environmental issues which were potential problems in the application of Laminar Flow Control (LFC) to transport aircraft. These included contamination of the LFC surface due to dirt, rain, insect remains, snow, and ice, in the critical leading-edge region. Douglas Aircraft Company designed and built a test article which was mounted on the right wing of the C-140 JetStar aircraft. The test article featured a retractable leading-edge high-lift shield for contamination protection and suction through perforations on the upper surface for LFC. Following a period of developmental flight testing, the aircraft entered simulated airline service, which included exposure to airborne insects, heavy rain, snow, and icing conditions both in the air and on the ground. During the roughly 3 years of flight testing, the test article has consistently demonstrated laminar flow in cruising flight. The experience with the LEFT experiment was summarized with emphasis on significant test findings. The following items were discussed: test article design and features; suction distribution; instrumentation and transition point reckoning; problems and fixes; system performance and maintenance requirements.

  18. Performance of two-stage fan having low-aspect-ratio first-stage rotor blading

    NASA Technical Reports Server (NTRS)

    Urasek, D. C.; Gorrell, W. T.; Cunnan, W. S.

    1979-01-01

    The NASA two stage fan was tested with a low aspect ratio first stage rotor having no midspan dampers. At design speed the fan achieved an adiabatic design efficiency of 0.846, and peak efficiencies for the first stage and rotor of 0.870 and 0.906, respectively. Peak efficiency occurred very close to the stall line. In an attempt to improve stall margin, the fan was retested with circumferentially grooved casing treatment and with a series of stator blade resets. Results showed no improvement in stall margin with casing treatment but increased to 8 percent with stator blade reset.

  19. The Flow Field Downstream of a Dynamic Low Aspect Ratio Circular Cylinder: A Parametric Study

    NASA Astrophysics Data System (ADS)

    Gildersleeve, Samantha; Dan, Clingman; Amitay, Michael

    2015-11-01

    Flow past a static, low aspect ratio cylinder (pin) has shown the formation of vortical structures, namely the horseshoe and arch-type vortex. These vortical structures may have substantial effects in controlling flow separation over airfoils. In the present experiments, the flow field associated with a low aspect ratio cylinder as it interacts with a laminar boundary layer under static and dynamic conditions was investigated through a parametric study over a flat plate. As a result of the pin being actuated in the wall-normal direction, the structures formed in the wake of the pin were seen to be a strong function of actuation amplitude, driving frequency, and aspect ratio of the cylinder. The study was conducted at a Reynolds number of 1875, based on the local boundary layer thickness, with a free stream velocity of 10 m/s. SPIV data were collected for two aspect ratios of 0.75 and 1.125, actuation amplitudes of 6.7% and 16.7%, and driving frequencies of 175 Hz and 350 Hz. Results indicate that the presence and interactions between vortical structures are altered in comparison to the static case and suggest increased large-scale mixing when the pin is driven at the shedding frequency (350 Hz). Supported by the Boeing Company.

  20. Three-dimensional wake topology and propulsive performance of low-aspect-ratio pitching-rolling plates

    NASA Astrophysics Data System (ADS)

    Li, Chengyu; Dong, Haibo

    2016-07-01

    The wake topology and propulsive performance of low-aspect-ratio plates undergoing a pitching-rolling motion in a uniform stream were numerically investigated by an in-house immersed-boundary-method-based incompressible Navier-Stokes equation solver. A detailed analysis of the vortical structures indicated that the pitching-rolling plate produced double-loop vortices with alternating signs from its trailing edge every half period. These vortices then shed and further evolved into interconnected "double-C"-shaped vortex rings, which eventually formed a bifurcating wake pattern in the downstream. As the wake convected downstream, there was a slight deflection in the spanwise direction to the plate tip, and the contained vortex ring size gradually increased. In addition, the analysis of the propulsive performance indicated that the shedding process of the double-loop vortices led to two peaks in the lift and thrust force production per half cycle. The observation of the double peaks in the force production is in agreement with previous flapping wing studies. Simulations were also used to examine the variations in the wake structures and propulsive performance of the plates over a range of major parameters. The aforementioned vortex structures were found to be quite robust over a range of Strouhal numbers, Reynolds numbers, and plate aspect ratios.

  1. On the evolution of the wake structure produced by a low-aspect-ratio pitching panel.

    PubMed

    Buchholz, James H J; Smits, Alexander J

    2005-12-05

    Flow visualization is used to interrogate the wake structure produced by a rigid flat panel of aspect ratio (span/chord) 0.54 pitching in a free stream at a Strouhal number of 0.23. At such a low aspect ratio, the streamwise vorticity generated by the plate tends to dominate the formation of the wake. Nevertheless, the wake has the appearance of a three-dimensional von Kármán vortex street, as observed in a wide range of other experiments, and consists of horseshoe vortices of alternating sign shed twice per flapping cycle. The legs of each horseshoe interact with the two subsequent horseshoes in an opposite-sign, then like-sign interaction in which they become entrained. A detailed vortex skeleton model is proposed for the wake formation.

  2. Hyperthermia in low aspect-ratio magnetic nanotubes for biomedical applications

    NASA Astrophysics Data System (ADS)

    Gutierrez-Guzman, D. F.; Lizardi, L. I.; Otálora, J. A.; Landeros, P.

    2017-03-01

    A simple model for the magnetization reversal process of low aspect-ratio ferromagnetic nanotubes (MNTs) is presented. Because of advantages over other geometries, these structures are interesting for biomedical applications, such as magnetic hyperthermia cancer therapy, where the heat released during magnetic reversal is used to destroy tumors. For example, the tubular geometry provides two independent functional surfaces that may be selectively manipulated and also gives a storage cavity. Owing to their large surface to weight ratio and low mass density, MNTs are not decanted by gravity. We calculated magnetic phase diagrams, energy barriers, nucleation fields, and the amount of dissipated heat and specific absorption rate for magnetite nanotubes. The geometrical parameters were varied, and simple formulae were used to optimize the tube response under alternating excitation, as required for magnetic hyperthermia applications.

  3. Aerodynamic and heat transfer analysis of the low aspect ratio turbine

    NASA Astrophysics Data System (ADS)

    Sharma, O. P.; Nguyen, P.; Ni, R. H.; Rhie, C. M.; White, J. A.

    1987-06-01

    The available two- and three-dimensional codes are used to estimate external heat loads and aerodynamic characteristics of a highly loaded turbine stage in order to demonstrate state-of-the-art methodologies in turbine design. By using data for a low aspect ratio turbine, it is found that a three-dimensional multistage Euler code gives good averall predictions for the turbine stage, yielding good estimates of the stage pressure ratio, mass flow, and exit gas angles. The nozzle vane loading distribution is well predicted by both the three-dimensional multistage Euler and three-dimensional Navier-Stokes codes. The vane airfoil surface Stanton number distributions, however, are underpredicted by both two- and three-dimensional boundary value analysis.

  4. Stability of low aspect ratio inverted flags and rods in a uniform flow

    NASA Astrophysics Data System (ADS)

    Huertas-Cerdeira, Cecilia; Sader, John E.; Gharib, Morteza

    2016-11-01

    Cantilevered elastic plates and rods in an inverted configuration, where the leading edge is free to move and the trailing edge is clamped, undergo complex dynamics when subjected to a uniform flow. The stability of low aspect ratio inverted plates and rods is theoretically examined, showing that it is markedly different from that of their large aspect ratio counterpart. In the limit of zero aspect ratio, the undeflected equilibrium position is found to be stable for all wind speeds. A saddle-node bifurcation emerges at finite wind speed, giving rise to a strongly deflected stable and a weakly deflected unstable equilibria. This theory is compared to experimental measurements, where good agreement is found. This research was supported by a Grant of the Gordon and Betty Moore Foundation, the Australian Research Council Grants scheme and a "la Caixa" Fellowship Grant for Post-Graduate Studies of "la Caixa" Banking Foundation.

  5. On the evolution of the wake structure produced by a low-aspect-ratio pitching panel

    PubMed Central

    BUCHHOLZ, JAMES H. J.; SMITS, ALEXANDER J.

    2009-01-01

    Flow visualization is used to interrogate the wake structure produced by a rigid flat panel of aspect ratio (span/chord) 0.54 pitching in a free stream at a Strouhal number of 0.23. At such a low aspect ratio, the streamwise vorticity generated by the plate tends to dominate the formation of the wake. Nevertheless, the wake has the appearance of a three-dimensional von Kármán vortex street, as observed in a wide range of other experiments, and consists of horseshoe vortices of alternating sign shed twice per flapping cycle. The legs of each horseshoe interact with the two subsequent horseshoes in an opposite-sign, then like-sign interaction in which they become entrained. A detailed vortex skeleton model is proposed for the wake formation. PMID:19746198

  6. Acquisition and reduction of blade-mounted pressure transducer data from a Low Aspect Ratio Fan

    NASA Astrophysics Data System (ADS)

    Russler, Patrick M.

    1995-02-01

    This report details the acquisition and reduction of blade-mounted, high-response, pressure transducer data. These data were acquired during the Augmented Damping of Low Aspect Ratio Fans (ADLARF) test conducted at the Compressor Research Facility (CRF) located at Wright-Patterson A.F.B. , Ohio. This report, which is exclusively concerned with the acquisition and digitizing of the blade-mounted data, is intended to compliment other related reports by documenting the data acquisition and reduction procedures. The primary goal of this work is to detail the methodology by which unsteady blade forces and momentum can he determined using blade-mounted pressure transducer data. The secondary goal is to use these data to show how inlet distortion and resulting unsteady forces affect the blade resonance of high-speed fans. By achieving the primary goal in this report, it is hoped that the secondary goal can be better achieved using data from future tests.

  7. Permutation entropy and statistical complexity in characterising low-aspect-ratio reversed-field pinch plasma

    NASA Astrophysics Data System (ADS)

    Onchi, T.; Fujisawa, A.; Sanpei, A.; Himura, H.; Masamune, S.; RELAX Team

    2017-05-01

    Permutation entropy and statistical complexity are measures for complex time series. The Bandt-Pompe methodology evaluates probability distribution using permutation. The method is robust and effective to quantify information of time series data. Statistical complexity is the product of Jensen-Shannon divergence and permutation entropy. These physical parameters are introduced to analyse time series of emission and magnetic fluctuations in low-aspect-ratio reversed-field pinch (RFP) plasma. The observed time-series data aggregates in a region of the plane, the so-called C-H plane, determined by entropy versus complexity. The C-H plane is a representation space used for distinguishing periodic, chaos, stochastic and noisy processes of time series data. The characteristics of the emissions and magnetic fluctuation change under different RFP-plasma conditions. The statistical complexities of soft x-ray emissions and magnetic fluctuations depend on the relationships between reversal and pinch parameters.

  8. Advanced Fuels Reactor using Aneutronic Rodless Ultra Low Aspect Ratio Tokamak Hydrogenic Plasmas

    NASA Astrophysics Data System (ADS)

    Ribeiro, Celso

    2015-11-01

    The use of advanced fuels for fusion reactor is conventionally envisaged for field reversed configuration (FRC) devices. It is proposed here a preliminary study about the use of these fuels but on an aneutronic Rodless Ultra Low Aspect Ratio (RULART) hydrogenic plasmas. The idea is to inject micro-size boron pellets vertically at the inboard side (HFS, where TF is very high and the tokamak electron temperature is relatively low because of profile), synchronised with a proton NBI pointed to this region. Therefore, p-B reactions should occur and alpha particles produced. These pellets will act as an edge-like disturbance only (cp. killer pellet, although the vertical HFS should make this less critical, since the unablated part should appear in the bottom of the device). The boron cloud will appear at midplance, possibly as a MARFE-look like. Scaling of the p-B reactions by varying the NBI energy should be compared with the predictions of nuclear physics. This could be an alternative to the FRC approach, without the difficulties of the optimization of the FRC low confinement time. Instead, a robust good tokamak confinement with high local HFS TF (enhanced due to the ultra low aspect ratio and low pitch angle) is used. The plasma central post makes the RULART concept attractive because of the proximity of NBI path and also because a fraction of born alphas will cross the plasma post and dragged into it in the direction of the central plasma post current, escaping vertically into a hole in the bias plate and reaching the direct electricity converter, such as in the FRC concept.

  9. Implications of in-vitro dosimetry on toxicological ranking of low aspect ratio engineered nanomaterials

    PubMed Central

    Pal, Anoop K.; Bello, Dhimiter; Cohen, Joel; Demokritou, Philip

    2016-01-01

    In-vitro high throughput screening platforms based on mechanistic injury pathways are been used for hazard assessment of engineered nanomaterials (ENM). Toxicity screening and other in vitro nanotoxicology assessment efforts in essence compare and rank nanomaterials relative to each other. We hypothesize that this ranking of ENM is susceptible to dispersion and dosimetry protocols, which continue to be poorly standardized. Our objective was to quantitate the impact of dosimetry on toxicity ranking of ENM. A set of eight well-characterized and diverse low aspect ratio ENMs, were utilized. The recently developed at Harvard in-vitro dosimetry platform, which includes preparation of fairly monodispersed suspensions, measurement of the effective density of formed agglomerates in culture media and fate and transport modeling was used for calculating the effective dose delivered to cells as a function of time. Changes in the dose-response relationships between the administered and delivered dose were investigated with two representative endpoints, cell viability and IL-8 production, in the human monocytic THP-1 cells. The slopes of administered/delivered dose-response relationships changed 1:4.94 times and were ENM-dependent. The overall relative ranking of ENM intrinsic toxicity also changed considerably, matching notably better the in vivo inflammation data (R2 0.97 vs. 0.64). This standardized dispersion and dosimetry methodology presented here is generalizable to low aspect ratio ENMs. Our findings further reinforce the need to reanalyze and reinterpret in-vitro ENM hazard ranking data published in the nanotoxicology literature in the light of dispersion and dosimetry considerations (or lack thereof) and to adopt these protocols in future in vitro nanotoxicology testing. PMID:25672815

  10. Implications of in vitro dosimetry on toxicological ranking of low aspect ratio engineered nanomaterials.

    PubMed

    Pal, Anoop K; Bello, Dhimiter; Cohen, Joel; Demokritou, Philip

    2015-01-01

    In vitro high throughput screening platforms based on mechanistic injury pathways are been used for hazard assessment of engineered nanomaterials (ENM). Toxicity screening and other in vitro nanotoxicology assessment efforts in essence compare and rank nanomaterials relative to each other. We hypothesize that this ranking of ENM is susceptible to dispersion and dosimetry protocols, which continue to be poorly standardized. Our objective was to quantitate the impact of dosimetry on toxicity ranking of ENM. A set of eight well-characterized and diverse low aspect ratio ENMs, were utilized. The recently developed in vitro dosimetry platform at Harvard, which includes preparation of fairly monodispersed suspensions, measurement of the effective density of formed agglomerates in culture media and fate and transport modeling was used for calculating the effective dose delivered to cells as a function of time. Changes in the dose-response relationships between the administered and delivered dose were investigated with two representative endpoints, cell viability and IL-8 production, in the human monocytic THP-1 cells. The slopes of administered/delivered dose-response relationships changed 1:4.94 times and were ENM-dependent. The overall relative ranking of ENM intrinsic toxicity also changed considerably, matching notably better the in vivo inflammation data (R(2 )= 0.97 versus 0.64). This standardized dispersion and dosimetry methodology presented here is generalizable to low aspect ratio ENMs. Our findings further reinforce the need to reanalyze and reinterpret in vitro ENM hazard ranking data published in the nanotoxicology literature in the light of dispersion and dosimetry considerations (or lack thereof) and to adopt these protocols in future in vitro nanotoxicology testing.

  11. An Investigation of the Aerodynamic Characteristics of an Airplane Equipped with Several Different Sets of Wings

    NASA Technical Reports Server (NTRS)

    Crowley, J W , Jr; Green, M W

    1929-01-01

    This investigation was conducted by the National Advisory Committee for Aeronautics at Langley Field, Va., at the request of the Army Air Corps, for the purpose of comparing the full scale lift and drag characteristics of an airplane equipped with several sets of wings of commonly used airfoil sections. A Sperry Messenger Airplane with wings of R.A.F.-15, U.S.A.-5, U.S.A.-27, and Gottingen 387 airfoil sections was flown and the lift and drag characteristics of the airplane with each set of wings were determined by means of glide tests. The results are presented in tabular and curve form. (author)

  12. High-Speed Wind-Tunnel Tests of a 1/16-Scale Model of the D-558 Research Airplane Lift and Drag Characteristics of the D-558-1 and Various Wing and Tail Configurations

    NASA Technical Reports Server (NTRS)

    Wright, John D.; Loving, Donald L.

    1953-01-01

    Tests were made in the Langley 8-foot high-speed tunnel to investigate the aerodynamic characteristics of the D-558-1 airplane and various wing and tail configurations on the D-558-1 fuselage. The various wing and tail configurations were tested to determine the aerodynamic effects of aspect ratio and sweep for suitable use on the second phase of the D-558 project (D-558-2). The tests were conducted through a speed range from a Mach number of 0.40 to approximately 0.94.This part of the investigation includes the lift and drag results available for the configurations tested at this rate. The D-558-1 results indicated that the lift force break would occur at a Mach number of 0.85 with some reduction in lift at speeds above this Mach number. Tests indicated that the airplane will have satisfactory lift and drag characteristics up to and including its design Mach number of 0.85. The 35deg sweptback, 35deg swept-forward, and low-aspect-ratio (2.0) wing configurations all showed pronounced improvements in maintaining lift throughout the Mach number range tested and in increasing the critical speeds above the D-558-1 value &itical to critical Mach numbers on the order of 0.9. Insofar as lift and drag characteristics are concerned level flight at speeds approaching the velocity of sound appears practical if swept or low-aspect-ratio configurations similar to those tested are used.

  13. Calculations of the Dynamic Stress of Several Airplane Wings in Various Gusts

    NASA Technical Reports Server (NTRS)

    Pierce, Harold B.

    1948-01-01

    A series of calculations of the dynamic response of airplane wings to gusts were made with the purpose of showing the relative response of a reference airplane, the DC-3 airplane, and of newer types of airplanes represented by the DC-4, DC-6, and L-49 airplanes. Additional calculations were made for the DC-6 airplane to show the effects of speed and altitude. On the basis of the method of calculation used and the conditions selected for analysis, it is indicated that: 1) The newer airplanes show appreciably greater dynamic stress in gusts then does the reference airplane; 2) Increasing the forward speed or the operating altitude results in an increase of the dynamic stress ratio for the gust with a gradient distance of 10 chords.

  14. Some problems on the lift and rolling moment of airplane wings

    NASA Technical Reports Server (NTRS)

    Scarborough, James B

    1925-01-01

    This report deals with the application of the airfoil and twisted wing theory to the calculation of the lift and rolling moment of airplane wings. Most of the results arrived at are strictly true only for wings of elliptic plan form. The investigation aims to give some indications of the accuracy with which the results can be applied to the wing forms in actual use.

  15. Pressure Distribution over the Wings of an MB-3 Airplane in Flight

    NASA Technical Reports Server (NTRS)

    Norton, F H

    1925-01-01

    This investigation was carried out to determine the distribution of load over the wings of a high speed airplane under all conditions of flight. In particular it was desired to find the pressure distribution during level flight, over the portions of the wings in the slipstream and, during violent maneuvers, over the entire wing surface. The method used consisted in connecting a number of holes in the surface of the wings to recording multiple manometers mounted in the fuselage of the airplane. In this way simultaneous records could be taken on all of the holes for any desired length of time. (author)

  16. Characteristics and Origin of Martian Low-Aspect-Ratio Layered Ejecta (LARLE) Craters

    NASA Astrophysics Data System (ADS)

    Barlow, Nadine G.; Boyce, J. M.

    2013-10-01

    An unusual crater morphology is found primarily at high latitudes on Mars. These craters display an extensive outer deposit beyond the normal layered ejecta blanket. This outer deposit extends up to 20 crater radii from the rim, terminates in a sinuous flame-like edge, and is extremely thin, leading to a low aspect ratio (A = thickness/length). These craters are thus called Low-Aspect-Ratio Layered Ejecta (LARLE) craters. We have conducted a survey of all LARLE craters 1-km-diameter and larger on Mars. We find 139 LARLE craters ranging in diameter from 1.0 to 12.2 km with a median of 2.8 km. Most (97%) are found poleward of 35N and 40S, with the remainder primarily found in the equatorial Medusae Fossae Formation. The surfaces of the freshest LARLE layers commonly exhibit radial, curvilinear ridges and dune-like landforms, and the LARLE deposit typically drapes over pre-existing terrain. We propose that the LARLE deposit is formed by a different mechanism than that responsible for the normal layered ejecta patterns. We suggest that impact into relatively-thick fine-grained ice-rich mantles enhances the formation of a base surge that is deposited after formation of the inner layered ejecta deposits. This base surge is similar to the density-driven, turbulent cloud of suspended fine-grained particles produced by impact erosion and mobilization of the surrounding surface material by ejecta from shallow-depth-of-burst nuclear and high-explosion craters. We have applied a base surge equation developed for terrestrial explosive events to two fresh LARLE craters. After adjustment of the equation for Martian conditions, it predicts runout distances that are within 99% of the observed values. All Martian craters likely produce a base surge during formation, but the presence of the obvious LARLE deposit is attributed to crater formation in thick, fine-grained, sedimentary deposits. These sediments are the source of the extra particulate debris incorporated into and deposited

  17. Flight-determined aerodynamic derivatives of the AD-1 oblique-wing research airplane

    NASA Technical Reports Server (NTRS)

    Sim, A. G.; Curry, R. E.

    1984-01-01

    The AD-1 is a variable-sweep oblique-wing research airplane that exhibits unconventional stability and control characteristics. In this report, flight-determined and predicted stability and control derivatives for the AD-1 airplane are compared. The predictions are based on both wind tunnel and computational results. A final best estimate of derivatives is presented.

  18. Physics Basis for High-Beta, Low-Aspect-Ratio Stellarator Experiments

    SciTech Connect

    A. Brooks; A.H. Reiman; G.H. Neilson; M.C. Zarnstorff; et al

    1999-11-01

    High-beta, low-aspect-ratio (compact) stellarators are promising solutions to the problem of developing a magnetic plasma configuration for magnetic fusion power plants that can be sustained in steady-state without disrupting. These concepts combine features of stellarators and advanced tokamaks and have aspect ratios similar to those of tokamaks (2-4). They are based on computed plasma configurations that are shaped in three dimensions to provide desired stability and transport properties. Experiments are planned as part of a program to develop this concept. A beta = 4% quasi-axisymmetric plasma configuration has been evaluated for the National Compact Stellarator Experiment (NCSX). It has a substantial bootstrap current and is shaped to stabilize ballooning, external kink, vertical, and neoclassical tearing modes without feedback or close-fitting conductors. Quasi-omnigeneous plasma configurations stable to ballooning modes at beta = 4% have been evaluated for the Quasi-Omnigeneous Stellarator (QOS) experiment. These equilibria have relatively low bootstrap currents and are insensitive to changes in beta. Coil configurations have been calculated that reconstruct these plasma configurations, preserving their important physics properties. Theory- and experiment-based confinement analyses are used to evaluate the technical capabilities needed to reach target plasma conditions. The physics basis for these complementary experiments is described.

  19. Design of a Low Aspect Ratio Transonic Compressor Stage Using CFD Techniques

    NASA Technical Reports Server (NTRS)

    Sanger, Nelson L.

    1994-01-01

    A transonic compressor stage has been designed for the Naval Postgraduate School Turbopropulsion Laboratory. The design relied heavily on CFD techniques while minimizing conventional empirical design methods. The low aspect ratio (1.2) rotor has been designed for a specific head ratio of .25 and a tip relative inlet Mach number of 1.3. Overall stage pressure ratio is 1.56. The rotor was designed using an Euler code augmented by a distributed body force model to account for viscous effects. This provided a relatively quick-running design tool, and was used for both rotor and stator calculations. The initial stator sections were sized using a compressible, cascade panel code. In addition to being used as a case study for teaching purposes, the compressor stage will be used as a research stage. Detailed measurements, including non-intrusive LDV, will be compared with the design computations, and with the results of other CFD codes, as a means of assessing and improving the computational codes as design tools.

  20. Evaluation of a low aspect ratio small axial compressor stage, volume 1

    NASA Technical Reports Server (NTRS)

    Sawyer, C. W., III

    1977-01-01

    A program was conducted to evaluate the effects of scaling, tip clearance, and IGV reset on the performance of a low aspect ratio compressor stage. Stage design was obtained by scaling an existing single stage compressor by a linear factor of 0.304. The design objective was to maintain the meanline velocity field of the base machine in the smaller size. Adjustments were made to account for predicted blockage differences and to chord lengths and airfoil edge radii to obtain reasonable blade geometries. Meanline velocity diagrams of the base stage were not maintained at the scaled size. At design speed and flowrate the scaled stage achieved a pressure ratio of 1.423, adiabatic efficiency of 0.822, and surge margin of 18.5%. The corresponding performance parameters for the base stage were 1.480, 0.872, and 25.2%, respectively. The base stage demonstrated a peak efficiency at design speed of 0.872; the scaled stage achieved a level of 0.838. When the scaled stage rotor and stator tip clearances were doubled, the stage achieved a pressure ratio of 1.413, efficiency of 0.799, and surge margin of 16.0% at the design flowrate. The peak stage efficiency at design speed was 0.825 with the increased clearance. Increased prewhirl lowered the stage pressure ratio as expected. Stage efficiency was maintained with ten degrees of increased prewhirl and then decreased substantially with ten additional degrees of reset.

  1. Energy confinement scaling in the low aspect ratio National Spherical Torus Experiment (NSTX)

    SciTech Connect

    Kaye, S. M.; Bell, M. G.; Bell, R. E.; Fredrickson, E. D.; LeBlanc, B. P.; Lee, K. C.; Lynch, S.; Sabbagh, S. A.

    2006-10-01

    Statistical and systematic studies have been conducted in order to develop an understanding of the parametric dependences of both the global and thermal energy confinement times at low aspect ratio in high power National Spherical Torus Experiment (NSTX) discharges. The global and thermal confinement times of both L- and H-mode discharges can exceed values given by H-mode scalings developed for conventional aspect ratio. Results of systematic scans in the H-mode indicate that the confinement times exhibit a nearly linear dependence on plasma current and a power degradation weaker than that observed at conventional aspect ratio. In addition, the dependence on the toroidal magnetic field is stronger than that seen in conventional aspect ratio tokamaks. Also, this latter trend is evident in statistical analyses of the available dataset. These statistical studies also indicate a weaker parametric dependence on plasma current than found in the systematic scans, due to correlations among the predictor variables. Regressions based on engineering variables, when transformed to dimensionless physics variables, indicate that the dependence of BτE on βt can range from being negative to null. Regressions based directly on the dimensionless physics variables are inexact because of large correlations among these variables. Scatter in the confinement data, at otherwise fixed operating parameters, is found to be due to variations in ELM activity, low frequency density fluctuations and plasma shaping.

  2. Coaxial helicity injection in open-flux low-aspect-ratio toroidal discharges

    SciTech Connect

    Redd, A. J.; Jarboe, T. R.; Nelson, B. A.; O'Neill, R. G.; Smith, R. J.

    2007-11-15

    Open-flux low-aspect-ratio toroidal discharges generated and sustained by coaxial helicity injection (CHI) in the Helicity Injected Torus device (HIT-II) are described. The discharges in this study are flux tubes directly connected to the CHI electrodes, with poloidal flux less than or equal to the CHI injector flux, and no possibility of a significant closed-flux plasma core. Theoretically derived scalings for the dependence of CHI injector current on the toroidal field current and magnitude of the injector flux are experimentally confirmed, and empirical models are developed for the poloidal magnetic field and toroidal plasma current in open-flux discharges. In particular, the toroidal plasma current is independent of the toroidal magnetic field, both theoretically and empirically. Variations in injector flux geometry demonstrate that the CHI injector current leaves the electrode surfaces at the flux strike points, and that the relative width of the CHI injector determines whether the dominant observed relaxation mechanism is a harmonic mode at the plasma edge or reconnection near the CHI electrodes. In the case of an effective interelectrode distance approximately equal to the device minor radius, the ratio of toroidal plasma current to CHI injector current is maximized. Global magnetic equilibrium quantities and local magnetic measurements are consistent with modelling these open-flux discharges as thin current sheets connected to the CHI electrodes.

  3. Comparing nonlinear MHD simulations of low-aspect-ratio RFPs to RELAX experiments

    NASA Astrophysics Data System (ADS)

    McCollam, K. J.; den Hartog, D. J.; Jacobson, C. M.; Sovinec, C. R.; Masamune, S.; Sanpei, A.

    2016-10-01

    Standard reversed-field pinch (RFP) plasmas provide a nonlinear dynamical system as a validation domain for numerical MHD simulation codes, with applications in general toroidal confinement scenarios including tokamaks. Using the NIMROD code, we simulate the nonlinear evolution of RFP plasmas similar to those in the RELAX experiment. The experiment's modest Lundquist numbers S (as low as a few times 104) make closely matching MHD simulations tractable given present computing resources. Its low aspect ratio ( 2) motivates a comparison study using cylindrical and toroidal geometries in NIMROD. We present initial results from nonlinear single-fluid runs at S =104 for both geometries and a range of equilibrium parameters, which preliminarily show that the magnetic fluctuations are roughly similar between the two geometries and between simulation and experiment, though there appear to be some qualitative differences in their temporal evolution. Runs at higher S are planned. This work is supported by the U.S. DOE and by the Japan Society for the Promotion of Science.

  4. Coaxial helicity injection in open-flux low-aspect-ratio toroidal discharges

    NASA Astrophysics Data System (ADS)

    Redd, A. J.; Jarboe, T. R.; Nelson, B. A.; O'Neill, R. G.; Smith, R. J.

    2007-11-01

    Open-flux low-aspect-ratio toroidal discharges generated and sustained by coaxial helicity injection (CHI) in the Helicity Injected Torus device (HIT-II) are described. The discharges in this study are flux tubes directly connected to the CHI electrodes, with poloidal flux less than or equal to the CHI injector flux, and no possibility of a significant closed-flux plasma core. Theoretically derived scalings for the dependence of CHI injector current on the toroidal field current and magnitude of the injector flux are experimentally confirmed, and empirical models are developed for the poloidal magnetic field and toroidal plasma current in open-flux discharges. In particular, the toroidal plasma current is independent of the toroidal magnetic field, both theoretically and empirically. Variations in injector flux geometry demonstrate that the CHI injector current leaves the electrode surfaces at the flux strike points, and that the relative width of the CHI injector determines whether the dominant observed relaxation mechanism is a harmonic mode at the plasma edge or reconnection near the CHI electrodes. In the case of an effective interelectrode distance approximately equal to the device minor radius, the ratio of toroidal plasma current to CHI injector current is maximized. Global magnetic equilibrium quantities and local magnetic measurements are consistent with modelling these open-flux discharges as thin current sheets connected to the CHI electrodes.

  5. Development of a low-aspect ratio fin for flight research experiments

    NASA Technical Reports Server (NTRS)

    Richwine, David M.; Delfrate, John H.

    1994-01-01

    A second-generation flight test fixture, developed at NASA Dryden Flight Research Center, offers a generic testbed for aerodynamic and fluid mechanics research. The new fixture, a low-aspect ratio vertical fin shape mounted on the centerline of an F-15B aircraft lower fuselage, is designed for flight research at Mach numbers up to 2.0. The new fixture is a composite structure with a modular configuration and removable components for functional flexibility. This report describes the multidisciplinary design and analysis approach used to develop the fixture. The approach integrates conservative assumptions with simple analysis techniques to minimize the time and cost associated with its development. Presented are the principal disciplines required for this effort, which include aerodynamics, structures, stability, and operational considerations. In addition, preliminary results from the first phase of flight testing are presented. Acceptable directional stability and flow quality are documented and show agreement with predictions. Future envelope expansion activities will minimize current limitations so that the fixture can be used for a wide variety of high-speed aerodynamic and fluid mechanics research experiments.

  6. Flight comparison of the transonic agility of the F-111A airplane and the F-111 supercritical wing airplane

    NASA Technical Reports Server (NTRS)

    Friend, E. L.; Sakamoto, G. M.

    1978-01-01

    A flight research program was conducted to investigate the improvements in maneuverability of an F-111A airplane equipped with a supercritical wing. In this configuration the aircraft is known as the F-111 TACT (transonic aircraft technology) airplane. The variable-wing-sweep feature permitted an evaluation of the supercritical wing in many configurations. The primary emphasis was placed on the transonic Mach number region, which is considered to be the principal air combat arena for fighter aircraft. An agility study was undertaken to assess the maneuverability of the F-111A aircraft with a supercritical wing at both design and off-design conditions. The evaluation included an assessment of aerodynamic and maneuver performance in conjunction with an evaluation of precision controllability during tailchase gunsight tracking tasks.

  7. Determination of the Profile Drag of an Airplane Wing in Flight at High Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Bicknell, Joseph

    1939-01-01

    Flight tests were made to determine the profile-drag coefficients of a portion of the original wing surface of an all-metal airplane and of a portion of the wing made aerodynamically smooth and more nearly fair than the original section. The wing section was approximately the NACA 2414.5. The tests were carried out over a range of airplane speeds giving a maximum Reynolds number of 15,000,000. Tests were also carried out to locate the point of transition from laminar to turbulent boundary layer and to determine the velocity distribution along the upper surface of the wing. The profile-drag coefficients of the original and of the smooth wing portions at a Reynolds number of 15,000,000 were 0.0102 and 0.0068, respectively; i.e., the surface irregularities on the original wing increased the profile-drag coefficient 50 percent above that of the smooth wing.

  8. Application of a performance modeling technique to an airplane with variable sweep wings

    NASA Technical Reports Server (NTRS)

    Redin, P. C.

    1981-01-01

    A performance modeling concept previously applied to an F-104F G and a YF-12C airplane was applied to an F-111A airplane. This application extended the concept to an airplane with variable sweep wings. The performance model adequately matched flight test data for maneuvers flown at different wing sweep angles at maximum afterburning and intermediate power settings. For maneuvers flown at less than intermediate power, including dynamic maneuvers, the performance model was not validated because the method used to correlate model and in-flight power setting was not adequate. Individual dynamic maneuvers were matched sucessfully by using adjustments unique to each maneuver.

  9. The Kizilkaya ignimbrite — an unusual low-aspect-ratio ignimbrite from Cappadocia, central Turkey

    NASA Astrophysics Data System (ADS)

    Schumacher, R.; Mues-Schumacher, U.

    1996-01-01

    The 4.3-m.y.-old medium-volume low-aspect-ratio Kizilkaya ignimbrite (50-100 km 3 DRE) is one of the most widespread in the Cappadocian Volcanic Province covering about 8500-10,600 km 2. The ignimbrite rests on a relatively fine-grained fan of Plinian pumice-fall deposit ( Md of 1.0-1.80 mm in proximal locations). The eruptive center was located in the Misli plain northeast of Nigde, as deduced from thickness and grain-size variations of the fall deposit, flow direction indicators, welding patterns of the ignimbrite and the distribution of certain types ofxenoliths. The massive ignimbrite, generally about 15 m thick, covers a paleoplain throughout at least two thirds of its areal extent. It comprizes two flow units, identified by local pumice enrichment in the upper part of the lower unit. The ignimbrite is completely welded in many places. In other places, the lower flow unit is non-welded, particularly where the initial pumice-fall deposit was eroded, a fine-grained ground layer was deposited, and undulating or cross-laminations with antidunes were developed. The ground layer was derived from the ignimbrite ground-mass by loss of fines < 250-500 μm. Depositional characteristics indicate that the ignimbrite was emplaced as high-concentration flows with relatively low velocity and low heat loss during runout. Local development of a ground layer and internal bedding structures indicate local increased turbulence only within individual flow portions due to agitated fluidization from engulfed air. The degree of welding of the lower flow unit was controlled by this turbulence and is not related to thickness variations.

  10. Leading-edge vortex burst on a low-aspect-ratio rotating flat plate

    NASA Astrophysics Data System (ADS)

    Medina, Albert; Jones, Anya R.

    2016-08-01

    This study experimentally investigates the phenomenon of leading-edge-vortex burst on rotating flat plate wings. An aspect-ratio-2 wing was driven in pure rotation at a Reynolds number of Re=2500 . Of primary interest is the evolution of the leading-edge vortex along the wing span over a single-revolution wing stroke. Direct force measurements of the lift produced by the wing revealed a single global lift maximum relatively early in the wing stroke. Stereoscopic particle image velocimetry was applied to several chordwise planes to quantify the structure and strength of the leading-edge vortex and its effect on lift production. This analysis revealed opposite-sign vorticity entrainment into the core of the leading-edge vortex, originating from a layer of secondary vorticity along the wing surface. Coincident with the lift peak, there emerged both a concentration of opposite vorticity in the leading-edge-vortex core, as well as axial flow stagnation within the leading-edge-vortex core. Planar control volume analysis was performed at the midspan to quantify the contributions of vorticity transport mechanisms to the leading-edge-vortex circulation. The rate of circulation annihilation by opposite-signed vorticity entrainment was found to be minimal during peak lift production, where convection balanced the flux of vorticity resulting in stagnation and eventually reversal of axial flow. Finally, vortex burst was found to be correlated with swirl number, where bursting occurs at a swirl threshold of Sw<0.6 .

  11. Analysis of Nonplanar Wing-tip-mounted Lifting Surfaces on Low-speed Airplanes

    NASA Technical Reports Server (NTRS)

    Vandam, C. P.; Roskam, J.

    1983-01-01

    Nonplanar wing tip mounted lifting surfaces reduce lift induced drag substantially. Winglets, which are small, nearly vertical, winglike surfaces, are an example of these devices. To achieve reduction in lift induced drag, winglets produce significant side forces. Consequently, these surfaces can seriously affect airplane lateral directional aerodynamic characteristics. Therefore, the effects of nonplanar wing tip mounted surfaces on the lateral directional stability and control of low speed general aviation airplanes were studied. The study consists of a theoretical and an experimental, in flight investigation. The experimental investigation involves flight tests of winglets on an agricultural airplane. Results of these tests demonstrate the significant influence of winglets on airplane lateral directional aerodynamic characteristics. It is shown that good correlations exist between experimental data and theoretically predicted results. In addition, a lifting surface method was used to perform a parametric study of the effects of various winglet parameters on lateral directional stability derivatives of general aviation type wings.

  12. Effect of the Surface Condition of a Wing on the Aerodynamic Characteristics of an Airplane

    NASA Technical Reports Server (NTRS)

    Defrance, S J

    1934-01-01

    In order to determine the effect of the surface conditions of a wing on the aerodynamic characteristics of an airplane, tests were conducted in the N.A.C.A. full-scale wind tunnel on the Fairchild F-22 airplane first with normal commercial finish of wing surface and later with the same wing polished. Comparison of the characteristics of the airplane with the two surface conditions shows that the polish caused a negligible change in the lift curve, but reduced the minimum drag coefficient by 0.001. This reduction in drag if applied to an airplane with a given speed of 200 miles per hour and a minimum drag coefficient of 0.025 would increase the speed only 2.9 miles per hour, but if the speed remained the same, the power would be reduced 4 percent.

  13. Approximations for column effect in airplane wing spars

    NASA Technical Reports Server (NTRS)

    Warner, Edward P; Short, Mac

    1927-01-01

    The significance attaching to "column effect" in airplane wing spars has been increasingly realized with the passage of time, but exact computations of the corrections to bending moment curves resulting from the existence of end loads are frequently omitted because of the additional labor involved in an analysis by rigorously correct methods. The present report represents an attempt to provide for approximate column effect corrections that can be graphically or otherwise expressed so as to be applied with a minimum of labor. Curves are plotted giving approximate values of the correction factors for single and two bay trusses of varying proportions and with various relationships between axial and lateral loads. It is further shown from an analysis of those curves that rough but useful approximations can be obtained from Perry's formula for corrected bending moment, with the assumed distance between points of inflection arbitrarily modified in accordance with rules given in the report. The discussion of general rules of variation of bending stress with axial load is accompanied by a study of the best distribution of the points of support along a spar for various conditions of loading.

  14. Experimental investigation of low aspect ratio, large amplitude, aeroelastic energy harvesting systems

    NASA Astrophysics Data System (ADS)

    Kirschmeier, Benjamin; Summerour, Jacob; Bryant, Matthew

    2017-04-01

    Interest in clean, stable, and renewable energy harvesting devices has increased dramatically with the volatility of petroleum markets. Specifically, research in aero/hydro kinetic devices has created numerous new horizontal and vertical axis wind turbines, and oscillating wing turbines. Oscillating wing turbines (OWTs) differ from their wind turbine cousins by having a rectangular swept area compared to a circular swept area. The OWT systems also possess a lower tip speed that reduces the overall noise produced by the system. OWTs have undergone significant computational analysis to uncover the underlying flow physics that can drive the system to high efficiencies for single wing oscillations. When two of these devices are placed in tandem configuration, i.e. one placed downstream of the other, they either can constructively or destructively interact. When constructive interactions occurred, they enhance the system efficiency to greater than that of two devices on their own. A new experimental design investigates the dependency of interaction modes on the pitch stiffness of the downstream wing. The experimental results demonstrated that interaction modes are functions of convective time scale and downstream wing pitch stiffness. Heterogeneous combinations of pitch stiffness exhibited constructive and destructive lock-in phenomena whereas the homogeneous combination exhibited only destructive interactions.

  15. Public Data Set: H-mode Plasmas at Very Low Aspect Ratio on the Pegasus Toroidal Experiment

    DOE Data Explorer

    Thome, Kathreen E. [University of Wisconsin-Madison; Oak Ridge Associated Universities] (ORCID:0000000248013922); Bongard, Michael W. [University of Wisconsin-Madison] (ORCID:0000000231609746); Barr, Jayson L. [University of Wisconsin-Madison] (ORCID:0000000177685931); Bodner, Grant M. [University of Wisconsin-Madison] (ORCID:0000000324979172); Burke, Marcus G. [University of Wisconsin-Madison] (ORCID:0000000176193724); Fonck, Raymond J. [University of Wisconsin-Madison] (ORCID:0000000294386762); Kriete, David M. [University of Wisconsin-Madison] (ORCID:0000000236572911); Perry, Justin M. [University of Wisconsin-Madison] (ORCID:0000000171228609); Reusch, Joshua A. [University of Wisconsin-Madison] (ORCID:0000000284249422); Schlossberg, David J. [University of Wisconsin-Madison] (ORCID:0000000287139448)

    2016-09-30

    This data set contains openly-documented, machine readable digital research data corresponding to figures published in K.E. Thome et al., 'H-mode Plasmas at Very Low Aspect Ratio on the Pegasus Toroidal Experiment,' Nucl. Fusion 57, 022018 (2017).

  16. Aerodynamic Loads at Mach Numbers from 0.70 to 2.22 on a Airplane Model Having a Wing and Canard of Triangular Plan Form and Either Single or Twin Vertical Tails

    NASA Technical Reports Server (NTRS)

    Peterson, Victor L.; Menees, Gene P.

    1961-01-01

    Results of an investigation of the aerodynamic loads on a canard airplane model are presented without detailed analysis for the Mach number range of 0.70 t o 2.22. The model consisted of a triangular wing and canard of aspect ratio 2 mounted on a Sears-Haack body of fineness ratio 12.5 and either a single body-mounted vertical tail or twin wing mounted vertical tails of low aspect ratio and sweptback plan form. The body, right wing panel, single vertical tail, and left twin vertical tail were instrumented for measuring pressures. Data were obtained for angles of attack ranging from -4 degrees to +16 degrees, nominal canard deflection angles of 0 degrees and 10 degrees, and angles of sideslip of 0 degrees and 5.3 degrees. The Reynolds number was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. Selected portions of the data are presented in graphical form and attention is directed to some of the results of the investigation. All of the experimental results have been tabulated in the form of pressure coefficients and integrations of the pressure coefficients and are available as supplements to this paper. A brief summary of the contents of the tabular material is given.

  17. Flight Tests of A 1/8-Scale Model of the Bell D-188A Jet VTOL Airplane

    NASA Technical Reports Server (NTRS)

    Smith, Charles C., Jr.

    1959-01-01

    The Bell D-188A VTOL airplane is a horizontal-attitude VTOL fighter with tilting engine nacelles at the tips of a low-aspect-ratio unswept wing and additional engines in the fuselage. The model could be flown smoothly in hovering and transition flight. In forward flight the model could be flown smoothly at the lower angles of attack but experienced an uncontrollable directional divergence at angles of attack above about 16 deg.

  18. A modified T-value method for selection of strain gages for measuring loads on a low aspect ratio wing

    NASA Technical Reports Server (NTRS)

    Tang, M. H.; Sheldon, R. G.

    1980-01-01

    A technique which may be useful for selecting strain gages for use in load equations is described. The technique is an adaptation of the previously used T-value method and is applied to a multispar structure. The technique, called the modified T-value method, is used to reduce the number of strain gages used in a load equation from twelve to two. A parallel reduction is made by calculating relative equation accuracies from three applied load distributions. The equations developed from the modified T-value method proved to be accurate more consistently than the T-value method.

  19. Pressure distribution on wing ribs of the VE-7 and TS airplanes in flight

    NASA Technical Reports Server (NTRS)

    Rhode, R V

    1927-01-01

    This paper is the first of a series of notes, each of which presents the complete results of pressure distribution tests made by the National Advisory Committee for Aeronautics, on single-wing ribs of the VE-7 and TS airplanes for a particular condition of flight. The level flight results are presented here in the form of curves and show the comparison between the pressure distribution over a representative thin wing, R.A.F.-15, and a moderately thick wing, U.S.A.-27, throughout the range of angle of attack.

  20. An analysis of life expectancy of airplane wings in normal cruising flight

    NASA Technical Reports Server (NTRS)

    Putnam, Abbott A

    1945-01-01

    In order to provide a basis for judging the relative importance of wing failure by fatigue and by single intense gusts, an analysis of wing life for normal cruising flight was made based on data on the frequency of atmospheric gusts. The independent variables considered in the analysis included stress-concentration factor, stress-load relation, wing loading, design and cruising speeds, design gust velocity, and airplane size. Several methods for estimating fatigue life from gust frequencies are discussed. The procedure selected for the analysis is believed to be simple and reasonably accurate, though slightly conservative.

  1. Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

    NASA Technical Reports Server (NTRS)

    Clarke, Robert; Allen, Michael J.; Dibley, Ryan P.; Gera, Joseph; Hodgkinson, John

    2005-01-01

    Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active Aeroelastic Wing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions.

  2. Distribution of Temperatures over an Airplane Wing with Reference to the Phenomena of Ice Formation

    NASA Technical Reports Server (NTRS)

    Brun, Edmond

    1938-01-01

    The results obtained from the present study of temperature distribution over an airplane wing afford means for making the following statements as regards the conditions of ice accretion and the use of a thermic anti-icer or de-icer: 1) Ice can form on a wing only when the temperature is below or hovering around zero. 2) The thermic effects produced on contact of the air with the moving wing rather oppose ice accretion. 3) The thermic procedure in the fight against ice accretion on the wing consists in electrical heating of the leading edge. 4) It seems that the formation of ice on the wing ought to be accompanied by a temperature rise which brings the accretion to 0 degrees. 5) If the thermic effects of friction favor the operation of the thermic anti-icer, the functioning of the de-icer is facilitated by the release of heat which accompanies the deposit of ice.

  3. Efficient light trapping in low aspect-ratio honeycomb nanobowl surface texturing for crystalline silicon solar cell applications

    NASA Astrophysics Data System (ADS)

    Gao, Pingqi; Wang, Hongzhe; Sun, Zixu; Han, Weiqiang; Li, Junshuai; Ye, Jichun

    2013-12-01

    We report a significant reflection reduction over a broadband light spectrum in crystalline silicon via introduction of low aspect-ratio honeycomb nanobowl front surface textures. A restructuration technique is developed to shape nanopores into nanobowls, enabling excellent impedance matching and efficient mode coupling. As a result, an overall reflection down to 2% in the spectrum range of 400-1 100 nm wavelength is achieved. In comparison to nanopores-structured light-trapping configurations, the nanobowls-textures have much smaller parasitic surface area, which mitigates the surface recombination losses. The texturing technique offers a promising approach to high efficiency c-Si thin-film solar cells.

  4. Exploration of low-aspect-ratio tokamak regimes in the CDX-U and TS-3 devices

    SciTech Connect

    Hwang, Y.S.; Yamada, M.; Jones, T.G.

    1994-12-31

    In the low-aspect-ratio tokamak regime, a lower q(a) regime (i.e. q(a) {le} 5, A = R/a {approx} 1.5) has been explored in CDX-U, and the ultra-low-aspect-ratio tokamak regime (1.05 {le} A {le} 1.5) has been explored in TS-3. Using a relatively low toroidal magnetic field, plasma discharges with I{sub p} {le} 53 kA, and q(a) {ge} 4 [q{sub cyl}(a) {ge}1] have been obtained in CDX-U. Low q(a), Ohmic plasmas in CDX-U show increasing MHD activity as the edge safety factor is lowered. These modes appear to reduce the current ramp-up rate and, at present, limit the access to even lower q(a) regimes. An experiment carried out in the ULART regime (A {approx} 1.05 {minus} 1.5) on the TS-3 device identifies a threshold of q(a) {ge} 3 with q{sub cyl}(a) < 1 for stability of global tilt/shift modes.

  5. Analysis of Effect of Rolling Pull-Outs on Wing and Aileron Loads of a Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Pearson, Henry A.; Aiken, William S.

    1946-01-01

    An analysis was made to determine the effect of rolling pull-out maneuvers on the wing and aileron loads of a typical fighter airplane, the P-47B. The results obtained indicate that higher loads are imposed upon wings and ailerons because of the rolling pull-out maneuver, than would be obtained by application of the loading requirements to which the airplane was designed. An increase of 102 lb or 15 percent of wing weight would be required if the wing were designed for rolling pull-out maneuver. It was also determined that the requirements by which the aileron was originally designed were inadequate.

  6. Measurement and Analysis of Wing and Tail Buffeting Loads on a Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Huston, Wilber B; Skopinski, T H

    1955-01-01

    The buffeting loads measured on the wing and tail of a fighter airplane during 194 maneuvers are given in tabular form, along with the associated flight conditions. Measurements were made at altitudes of 30,000 to 10,000 feet and at speeds up to a Mach number of 0.8. Least-squares methods have been used for a preliminary analysis of the data. The agreement between the results of this analysis and the loads measured in stalls is sufficiently good to suggest the examination of the buffeting of other airplanes on the same basis.

  7. Oblique-wing research airplane motion simulation with decoupling control laws

    NASA Technical Reports Server (NTRS)

    Kempel, Robert W.; Mc Neill, Walter E.; Maine, Trindel A.

    1988-01-01

    A large piloted vertical motion simulator was used to assess the performance of a preliminary decoupling control law for an early version of the F-8 oblique wing research demonstrator airplane. Evaluations were performed for five discrete flight conditions, ranging from low-altitude subsonic Mach numbers to moderate-altitude supersonic Mach numbers. Asymmetric sideforce as a function of angle of attack was found to be the primary cause of both the lateral acceleration noted in pitch and the tendency to roll into left turns and out of right turns. The flight control system was shown to be effective in generally decoupling the airplane and reducing the lateral acceleration in pitch maneuvers.

  8. Crash tests of four identical high-wing single-engine airplanes

    NASA Technical Reports Server (NTRS)

    Vaughan, V. L., Jr.; Hayduk, R. J.

    1980-01-01

    Four identical four place, high wing, single engine airplane specimens with nominal masses of 1043 kg were crash tested at the Langley Impact Dynamics Research Facility under controlled free flight conditions. These tests were conducted with nominal velocities of 25 m/sec along the flight path angles, ground contact pitch angles, and roll angles. Three of the airplane specimens were crashed on a concrete surface; one was crashed on soil. Crash tests revealed that on a hard landing, the main landing gear absorbed about twice the energy for which the gear was designed but sprang back, tending to tip the airplane up to its nose. On concrete surfaces, the airplane impacted and remained in the impact attitude. On soil, the airplane flipped over on its back. The crash impact on the nose of the airplane, whether on soil or concrete, caused massive structural crushing of the forward fuselage. The liveable volume was maintained in both the hard landing and the nose down specimens but was not maintained in the roll impact and nose down on soil specimens.

  9. Free-Spinning Wind-Tunnel Tests of a Low-Wing Monoplane with Systematic Changes in Wings and Tails V : Effect of Airplane Relative Density

    NASA Technical Reports Server (NTRS)

    Seidman, Oscar; Neihouse, A I

    1940-01-01

    The reported tests are a continuation of an NACA investigation being made in the free-spinning wind tunnel to determine the effects of independent variations in load distribution, wing and tail arrangement, and control disposition on the spin characteristics of airplanes. The standard series of tests was repeated to determine the effect of airplane relative density. Tests were made at values of the relative-density parameter of 6.8, 8.4 (basic), and 12.0; and the results were analyzed. The tested variations in the relative-density parameter may be considered either as variations in the wing loading of an airplane spun at a given altitude, with the radii of gyration kept constant, or as a variation of the altitude at which the spin takes place for a given airplane. The lower values of the relative-density parameter correspond to the lower wing loadings or to the lower altitudes of the spin.

  10. Recent Loads Calibration Experience With a Delta Wing Airplane

    NASA Technical Reports Server (NTRS)

    Jenkins, Jerald M.; Kuhl, Albert E.

    1977-01-01

    Aircraft which are designed for supersonic and hypersonic flight are evolving with delta wing configurations. An integral part of the evolution of all new aircraft is the flight test phase. Included in the flight test phase is an effort to identify and evaluate the loads environment of the aircraft. The most effective way of examining the loads environment is to utilize calibrated strain gages to provide load magnitudes. Using strain gage data to accomplish this has turned out to be anything but a straightforward task. The delta wing configuration has turned out to be a very difficult type of wing structure to calibrate. Elevated structural temperatures result in thermal effects which contaminate strain gage data being used to deduce flight loads. The concept of thermally calibrating a strain gage system is an approach to solving this problem. This paper will address how these problems were approached on a program directed toward measuring loads on the wing of a large, flexible supersonic aircraft. Structural configurations typical of high-speed delta wing aircraft will be examined. The temperature environment will be examined to see how it induces thermal stresses which subsequently cause errors in loads equations used to deduce the flight loads.

  11. Flight evaluation of the transonic stability and control characteristics of an airplane incorporating a supercritical wing

    NASA Technical Reports Server (NTRS)

    Matheny, N. W.; Gatlin, D. H.

    1978-01-01

    A TF-8A airplane was equipped with a transport type supercritical wing and fuselage fairings to evaluate predicted performance improvements for cruise at transonic speeds. A comparison of aerodynamic derivatives extracted from flight and wind tunnel data showed that static longitudinal stability, effective dihedral, and aileron effectiveness, were higher than predicted. The static directional stability derivative was slower than predicted. The airplane's handling qualities were acceptable with the stability augmentation system on. The unaugmented airplane exhibited some adverse lateral directional characteristics that involved low Dutch roll damping and low roll control power at high angles of attack and roll control power that was greater than satisfactory for transport aircraft at cruise conditions. Longitudinally, the aircraft exhibited a mild pitchup tendency. Leading edge vortex generators delayed the onset of flow separation, moving the pitchup point to a higher lift coefficient and reducing its severity.

  12. Parameters Governing the Wake Structure of a Low-Aspect-Ratio Pitching Panel at ReC=640

    NASA Astrophysics Data System (ADS)

    Buchholz, James; Smits, Alexander

    2008-11-01

    Measurements of the wake structure and thrust performance of a rigid rectangular low-aspect-ratio panel pitching about its leading edge have been previously reported (Buchholz, J.H.J and Smits, A.J, J. Fluid Mech, 603, pp. 331--365). Such a simplified propulsor has proven to be a useful platform for the investigation of the physics of aquatic animal swimming. Wake visualizations at ReC= 640 yielded vortex skeleton models that were used to understand the structure of the wake at ReC= O(10^4); however, at this low Reynolds number, only a single value of panel aspect ratio and pitching amplitude were investigated, for three representative Strouhal numbers. In the present work, we investigate variations in aspect ratio and pitching amplitude, and consider additional pitching frequencies in order to further elucidate parameters governing the structure of the wake at ReC= 640.

  13. Wind-tunnel investigation of basic aerodynamic characteristics of a supercritical-wing research airplane configuration

    NASA Technical Reports Server (NTRS)

    Bartlett, D. W.; Re, R. J.

    1972-01-01

    Transonic pressure tunnel and transonic tunnel tests were performed to determine the aerodynamic characteristics of a 0.087 scale model of a supercritical wing research airplane configuration at Mach numbers from 0.25 to 1.30. The investigation included tests to determine the basic longitudinal aerodynamic characteristics, the lateral-directional aerodynamic characteristics for sideslip angles of 0 deg and + or - 2.5 deg, and the effects of Reynolds number and aeroelasticity.

  14. High-Speed Tests of a Model Twin-Engine Low-Wing Transport Airplane

    NASA Technical Reports Server (NTRS)

    Becker, John V; LEONARD LLOYD H

    1942-01-01

    Report presents the results of force tests made of a 1/8-scale model of a twin-engine low-wing transport airplane in the NACA 8-foot high-speed tunnel to investigate compressibility and interference effects of speeds up to 450 miles per hour. In addition to tests of the standard arrangement of the model, tests were made with several modifications designed to reduce the drag and to increase the critical speed.

  15. An analysis of the effects of aeroelasticity on static longitudinal stability and control of a swept-back-wing airplane

    NASA Technical Reports Server (NTRS)

    Skoog, Richard B

    1951-01-01

    A theoretical analysis of the effects of aeroelasticity on the stick-fixed static longitudinal stability and elevator angle required for balance of an airplane is presented together with calculated effects for a swept-wing bomber of relatively high flexibility. Although large changes in stability due to certain parameters are indicated for the example airplane, the over-all stability change after considering all parameters was quite small, compared to the individual effects, due to the counterbalancing of wing and tail contributions. The effect of flexibility on longitudinal control for the example airplane was found to be of little real importance.

  16. Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

    NASA Technical Reports Server (NTRS)

    Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert

    2005-01-01

    The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

  17. Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

    NASA Technical Reports Server (NTRS)

    Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert

    2005-01-01

    The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

  18. Dynamic Ground Effect for a Cranked Arrow Wing Airplane

    NASA Technical Reports Server (NTRS)

    Curry, Robert E.

    1997-01-01

    Flight-determined ground effect characteristics for an F-16XL airplane are presented and correlated with wind tunnel predictions and similar flight results from other aircraft. Maneuvers were conducted at a variety of flightpath angles. Conventional ground effect flight test methods were used, with the exception that space positioning data were obtained using the differential global positioning system (DGPS). Accuracy of the DGPS was similar to that of optical tracking methods, but it was operationally more attractive. The dynamic flight determined lift and drag coefficient increments were measurably lower than steady-state wind-tunnel predictions. This relationship is consistent with the results of other aircraft for which similar data are available. Trends in the flight measured lift increments caused by ground effect as a function of flightpath angle were evident but weakly correlated. An engineering model of dynamic ground effect was developed based on linear aerodynamic theory and super-positioning of flows. This model was applied to the F-16XL data set and to previously published data for an F-15 airplane. In both cases, the model provided an engineering estimate of the ratio between the steady-state and dynamic data sets.

  19. An analysis of the effects of aeroelasticity on static longitudinal stability and control of a swept-wing airplane

    NASA Technical Reports Server (NTRS)

    Skoog, Richard B

    1957-01-01

    A theoretical analysis has been made of the effects of aeroelasticity on the static longitudinal stability and elevator angle required for balance of an airplane. The analysis is based on the familiar stability equation expressing the contribution of wing and tail to longitudinal stability. Effects of wing, tail, and fuselage flexibility are considered. Calculated effects are shown for a swept-wing bomber of relatively high flexibility.

  20. Development of a composite tailoring procedure for airplane wing

    NASA Technical Reports Server (NTRS)

    Chattopadhyay, Aditi; Zhang, Sen

    1995-01-01

    The development of a composite wing box section using a higher order-theory is proposed for accurate and efficient estimation of both static and dynamic responses. The theory includes the effect of through-the-thickness transverse shear deformations which is important in laminated composites and is ignored in the classical approach. The box beam analysis is integrated with an aeroelastic analysis to investigate the effect of composite tailoring using a formal design optimization technique. A hybrid optimization procedure is proposed for addressing both continuous and discrete design variables.

  1. Investigation of Vortical Flow Patterns in the Near Field of a Dynamic Low-Aspect-Ratio Cylinder

    NASA Astrophysics Data System (ADS)

    Gildersleeve, Samantha; Amitay, Michael

    2016-11-01

    The flowfield and associated flow structures of a low-aspect-ratio cylindrical pin were investigated experimentally in the near-field as the pin underwent wall-normal periodic oscillations. Under dynamic conditions, the pin is driven at the natural wake shedding frequency with an amplitude of 33% of its mean height. Additionally, a static pin was also tested at various mean heights of 0.5, 1.0, and 1.5 times the local boundary layer thickness to explore the effect of the mean height on the flowfield. Three-dimensional flowfields were reconstructed and analyzed from SPIV measurements where data were collected along streamwise planes for several spanwise locations under static and dynamic conditions. The study focuses on the incoming boundary layer as it interacts with the pin, as well as two main vortical formations: the arch-type vortex and the horseshoe vortex. Under dynamic conditions, the upstream boundary layer is thinner, relative to the baseline, and the downwash in the wake increases, resulting in a reduced wake deficit. These results indicate enhanced strength of the aforementioned vortical flow patterns under dynamic conditions. The flow structures in the near-field of the static/dynamic cylinder will be discussed in further detail. Supported by The Boeing Company.

  2. Analytical modeling of seismic wave scattered from a 2D fracture simulated by a low-aspect ratio elliptical cylinder

    NASA Astrophysics Data System (ADS)

    Chen, T.; Wang, P.; Fehler, M.; Zhang, Y.; Burns, D.

    2009-12-01

    Localizing subsurface fractures and estimating their mechanical parameters and geometric properties are very important in oil and gas industry as well as geothermal energy research. It is essential to quantitatively understand how the elastic wave propagation is affected by these fractures. In this paper, an analytical expression for the scattered P- and SV waves from a 2D fracture is formulated based on a normal mode method, where the 2D fracture is modeled by a low-aspect ratio elliptical cylinder. The scatter function of this 2D fracture are expressed in terms of the incident angle, the orientation and aspect ratio of the fracture as well as the elastic impedance contrast between the surrounding medium and the inhomogeneity inside the fracture. Results from this analytical solution match well with those from a finite-difference approach. Solutions of this analytical model at two limiting cases (a circular cylinder with aspect ratio equal to one and a strip with aspect ratio equal to zero) are also compared to analytical solutions directly derived for the circular cylinder and strip by other studies.

  3. Crash tests of three identical low-wing single-engine airplane

    NASA Technical Reports Server (NTRS)

    Castle, C. B.; Alfaro-Bou, E.

    1983-01-01

    Three identical four place, low wing single engine airplane specimens with nominal masses of 1043 kg were crash tested under controlled free flight conditions. The tests were conducted at the same nominal velocity of 25 m/sec along the flight path. Two airplanes were crashed on a concrete surface (at 10 and 30 deg pitch angles), and one was crashed on soil (at a -30 deg pitch angle). The three tests revealed that the specimen in the -30 deg test on soil sustained massive structural damage in the engine compartment and fire wall. Also, the highest longitudinal cabin floor accelerations occurred in this test. Severe damage, but of lesser magnitude, occurred in the -30 deg test on concrete. The highest normal cabin floor accelerations occurred in this test. The least structural damage and lowest accelerations occurred in the 10 deg test on concrete.

  4. Kinetic effects in the conversion of fast waves in pre-heated, low aspect ratio tokamak plasmas

    NASA Astrophysics Data System (ADS)

    Kommoshvili, K.; Cuperman, S.; Bruma, C.

    2003-03-01

    Kinetic effects in the conversion of fast waves to Alfvèn waves and their subsequent deposition in low aspect ratio (spherical) tokamaks (LARTs) have been investigated theoretically. More specifically, we have considered the consequences of incorporation of kinetic effects in the electron parallel (to the ambient magnetic field) dynamics derived by following the drift-tearing mode analysis of Chen et al (Chen L, Rutherford P H and Tang W M 1977 Phys. Rev. Lett. 39 460), and particle-conserving Krook collision operator for the passing electrons involved (Mett R R and Mahajan S M 1992 Phys. Fluids B 4 2885). The perpendicular plasma dynamics is described by a quite general resistive two-fluid (2F) model based dielectric tensor-operator (Cuperman S, Bruma C and Komoshvili K 2002 Solution of the resistive 2F wave equations for Alfvènic modes in spherical tokamak plasmas J. Plasma Phys. accepted for publication). The full-wave electromagnetic equations, formulated in terms of the vector and scalar potentials, have been solved by the aid of an advanced finite elements numerical code (Sewell G 1993 Adv. Eng. Software 17 105). Detailed solutions of the full-wave equations are obtained and compared with those corresponding to a pure resistive 2F model, this, for the illustrative pre-heated START-type device (Sykes 1994). Our results quantitatively confirm the general theory of the conversion of fast waves with subsequent power dissipation for the conditions of spherical tokamaks thus providing the required auxilliary energy source for the succesful operation of LARTs. Moreover, these results indicate the absolute necessity of using a full model for the parallel electron dynamics, i.e. including both kinetic and collisional effects.

  5. A simple method for increasing the lift of airplane wings by means of flaps

    NASA Technical Reports Server (NTRS)

    Gruschwitz, Eugen; Schrenk, Oskar

    1933-01-01

    Aerodynamic considerations led us, not long ago, to investigate a device which seemed to promise a contribution to the problem of reducing the landing speed of an airplane. We have subsequently learned that similar devices had already been proposed and investigated by others, but it seems advisable, nevertheless, to report our results. The problem is to create, in landing, a region of turbulence on the lower side of the wing near the trailing edge by some obstacle to the air flow. The devices tested by us consisted of flaps of varying chord and position, the chord s being equal to the distance of the pivot from the trailing edge.

  6. The Strength of One-Piece Solid, Build-Up and Laminated Wood Airplane Wing Beams

    NASA Technical Reports Server (NTRS)

    Nelson, John H

    1920-01-01

    The purpose of this report is to summarize the results of all wood airplane wing beams tested to date in the Bureau of Standards Laboratory in order that the various kinds of wood and methods of construction may be compared. All beams tested were of an I section and the majority were somewhat similar in size and cross section to the front wing beam of the Curtiss JN-4 machine. Construction methods may be classed as (1) solid beams cut from solid stock; (2) three-piece beams, built up of three pieces, web and flanges glued together by a tongue-and-groove joint and (3) laminated beams built up of thin laminations of wood glued together.

  7. Aerodynamic characteristics at Mach 6 of a wing-body concept for a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Dillon, J. L.; Pittman, J. L.

    1978-01-01

    The static aerodynamic characteristics of a 1/30 scale model of a wing-body concept for a high speed research airplane were investigated in the Langley 20 inch Mach six tunnel. The investigation consisted of configuration buildup from the basic body by adding a wing, center vertical tail, three-module scramjet, and six-module scramjet engine. The test Mach number was six at a Reynolds number, based on model fuselage length, of about 13,700,000. The test angle-of-attack range was 4 to 20 D at constant angles of sideslip of 0, 2, and 4 deg. The elevons were deflected from 10 to -15 D for pitch control. Roll and yaw control were investigated. Experimental aerodynamic characteristics are compared with analytical elements.

  8. A critical evaluation of a three-dimensional Navier-Stokes method as a tool to calculate transonic flows inside a low-aspect-ratio compressor

    NASA Technical Reports Server (NTRS)

    Hah, Chunill; Puterbaugh, Steven L.

    1992-01-01

    A numerical study to evaluate a three-dimensional Navier-Stokes method as a tool to predict the detailed flow field inside a low-aspect-ratio compressor at various operating conditions was conducted. The details of the flow structure inside a low aspect ratio compressor (three-dimensional shock structure, shock-boundary layer interaction, and tip leakage vortex) and the overall aerodynamic performance at design and off-design conditions are numerically analyzed and the results are compared with the available experimental data. The flow field inside a state-of-the-art transonic compressor is used for the purpose of the evaluation.

  9. Power reduction and the radial limit of stall delay in revolving wings of different aspect ratio

    PubMed Central

    Kruyt, Jan W.; van Heijst, GertJan F.; Altshuler, Douglas L.; Lentink, David

    2015-01-01

    Airplanes and helicopters use high aspect ratio wings to reduce the power required to fly, but must operate at low angle of attack to prevent flow separation and stall. Animals capable of slow sustained flight, such as hummingbirds, have low aspect ratio wings and flap their wings at high angle of attack without stalling. Instead, they generate an attached vortex along the leading edge of the wing that elevates lift. Previous studies have demonstrated that this vortex and high lift can be reproduced by revolving the animal wing at the same angle of attack. How do flapping and revolving animal wings delay stall and reduce power? It has been hypothesized that stall delay derives from having a short radial distance between the shoulder joint and wing tip, measured in chord lengths. This non-dimensional measure of wing length represents the relative magnitude of inertial forces versus rotational accelerations operating in the boundary layer of revolving and flapping wings. Here we show for a suite of aspect ratios, which represent both animal and aircraft wings, that the attachment of the leading edge vortex on a revolving wing is determined by wing aspect ratio, defined with respect to the centre of revolution. At high angle of attack, the vortex remains attached when the local radius is shorter than four chord lengths and separates outboard on higher aspect ratio wings. This radial stall limit explains why revolving high aspect ratio wings (of helicopters) require less power compared with low aspect ratio wings (of hummingbirds) at low angle of attack and vice versa at high angle of attack. PMID:25788539

  10. Power reduction and the radial limit of stall delay in revolving wings of different aspect ratio.

    PubMed

    Kruyt, Jan W; van Heijst, GertJan F; Altshuler, Douglas L; Lentink, David

    2015-04-06

    Airplanes and helicopters use high aspect ratio wings to reduce the power required to fly, but must operate at low angle of attack to prevent flow separation and stall. Animals capable of slow sustained flight, such as hummingbirds, have low aspect ratio wings and flap their wings at high angle of attack without stalling. Instead, they generate an attached vortex along the leading edge of the wing that elevates lift. Previous studies have demonstrated that this vortex and high lift can be reproduced by revolving the animal wing at the same angle of attack. How do flapping and revolving animal wings delay stall and reduce power? It has been hypothesized that stall delay derives from having a short radial distance between the shoulder joint and wing tip, measured in chord lengths. This non-dimensional measure of wing length represents the relative magnitude of inertial forces versus rotational accelerations operating in the boundary layer of revolving and flapping wings. Here we show for a suite of aspect ratios, which represent both animal and aircraft wings, that the attachment of the leading edge vortex on a revolving wing is determined by wing aspect ratio, defined with respect to the centre of revolution. At high angle of attack, the vortex remains attached when the local radius is shorter than four chord lengths and separates outboard on higher aspect ratio wings. This radial stall limit explains why revolving high aspect ratio wings (of helicopters) require less power compared with low aspect ratio wings (of hummingbirds) at low angle of attack and vice versa at high angle of attack.

  11. Pressure distribution for the wing of the YAV-8B airplane; with and without pylons

    NASA Technical Reports Server (NTRS)

    Saltzman, Edwin J.; Delfrate, John H.; Sabsay, Catherine M.; Yarger, Jill M.

    1992-01-01

    Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.

  12. Public Data Set: Continuous, Edge Localized Ion Heating During Non-Solenoidal Plasma Startup and Sustainment in a Low Aspect Ratio Tokamak

    DOE Data Explorer

    Burke, Marcus G. [University of Wisconsin-Madison] (ORCID:0000000176193724); Barr, Jayson L. [University of Wisconsin-Madison] (ORCID:0000000177685931); Bongard, Michael W. [University of Wisconsin-Madison] (ORCID:0000000231609746); Fonck, Raymond J. [University of Wisconsin-Madison] (ORCID:0000000294386762); Hinson, Edward T. [University of Wisconsin-Madison] (ORCID:000000019713140X); Perry, Justin M. [University of Wisconsin-Madison] (ORCID:0000000171228609); Reusch, Joshua A. [University of Wisconsin-Madison] (ORCID:0000000284249422); Schlossberg, David J. [University of Wisconsin-Madison] (ORCID:0000000287139448)

    2017-05-16

    This public data set contains openly-documented, machine readable digital research data corresponding to figures published in M.G. Burke et. al., 'Continuous, Edge Localized Ion Heating During Non-Solenoidal Plasma Startup and Sustainment in a Low Aspect Ratio Tokamak,' Nucl. Fusion 57, 076010 (2017).

  13. Feasibility of wing shielding of the airplane interior from the shock noise generated by supersonic tip speed propellers

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.

    1978-01-01

    A high tip speed turboprop is being considered as a future energy conservative airplane. The high tip speed of the propeller, combined with the speed of the airplane, results in supersonic relative flow on the propeller tips. These supersonic blade sections could generate noise that is a cabin environment problem. The feasibility of using wing shielding to lessen the impact of this supersonic propeller noise was investigated. An analytical model is chosen which considers that shock waves are associated with the propeller tip flow and indicates how they would be prevented from impinging on the airplane fuselage. An example calculation is performed where a swept wing is used to shield the fuselage from significant portions of the propeller shock waves.

  14. Effects of wing leading-edge deflection on low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration. [wind tunnel tests

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    Static force tests were conducted in the Langley V/STOL tunnel at a Reynolds number (based on the mean aerodynamic chord) of about 2.0 x 10 to the 6th power for an angle-of-attack range from about - 10 deg to 17 deg and angles of sideslip of 0 and + or - 5 deg. Limited flow visualization studies were also conducted in order to provide a qualitative assessment of leading-edge upwash characteristics.

  15. Preliminary study of a large span-distributed-load flying-wing cargo airplane concept

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.

    1978-01-01

    An aircraft capable of transporting containerized cargo over intercontinental distances is analyzed. The specifications for payload weight, density, and dimensions in essence configure the wing and establish unusually low values of wing loading and aspect ratio. The structural weight comprises only about 18 percent of the design maximum gross weight. Although the geometric aspect ratio is 4.53, the winglet effect of the wing-tip-mounted vertical tails, increase the effective aspect ratio to approximately 7.9. Sufficient control power to handle the large rolling moment of inertia dictates a relatively high minimum approach velocity of 315 km/hr (170 knots). The airplane has acceptable spiral, Dutch roll, and roll-damping modes. A hardened stability augmentation system is required. The most significant noise source is that of the airframe. However, for both take-off and approach, the levels are below the FAR-36 limit of 108 db. The design mission fuel efficiency is approximately 50 percent greater than that of the most advanced, currently operational, large freighter aircraft. The direct operating cost is significantly lower than that of current freighters, the advantage increasing as fuel price increases.

  16. Residual strength and crack propagation tests on C-130 airplane center wings with service-imposed fatigue damage

    NASA Technical Reports Server (NTRS)

    Snider, H. L.; Reeder, F. L.; Dirkin, W. J.

    1972-01-01

    Fourteen C-130 airplane center wings, each containing service-imposed fatigue damage resulting from 4000 to 13,000 accumulated flight hours, were tested to determine their fatigue crack propagation and static residual strength characteristics. Eight wings were subjected to a two-step constant amplitude fatigue test prior to static testing. Cracks up to 30 inches long were generated in these tests. Residual static strengths of these wings ranged from 56 to 87 percent of limit load. The remaining six wings containing cracks up to 4 inches long were statically tested as received from field service. Residual static strengths of these wings ranged from 98 to 117 percent of limit load. Damage-tolerant structural design features such as fastener holes, stringers, doublers around door cutouts, and spanwise panel splices proved to be effective in retarding crack propagation.

  17. Longitudinal Characteristics of a Semispan Model of the Grumman Airplane Design 83 having a Sweptback Wing and of the Model with a Straight Wing as Determined from Wing-Flow Tests at Transonic Speeds, TED No. NACA DE337

    NASA Technical Reports Server (NTRS)

    Silsby, Norman S.; Kennedy, Robert M.

    1947-01-01

    An investigation has been made by the NACA wing-flow method to provide information on the relative longitudinal characteristics of a straight and sweptback wing in the transonic speed range. Tests were made of a semispan model of the Grumman airplane design 83 (XFlOF) incorporating a wing swept back 42.5deg with reference to quarter-chord line and also of the model with the swept wing replaced by a straight wing similar to that of the XF9F airplane. The airfoil sections were symmetrical 64l-series, with thickness ratios of 12 percent for the straight wing and 10 percent for the sweptback wing parallel to the stream direction. Measurements were made of normal force, chord force, and pitching moment at various angles of attack with the two wings both with and without the empennage, and with the fuselage alone. The tests covered a range of effective Mach numbers at the wing of the model from 0.65 to 1.10.

  18. The Effect of Various Wing-Gun Installations on the Aerodynamic Characteristics of an Airplane Model Equipped with an NACA Low-Drag Wing, Special Report

    NASA Technical Reports Server (NTRS)

    Muse, Thomas C.

    1941-01-01

    An investigation was made in the NACA 19-foot pressure wind tunnel to determine the effect of various win-gun installation on the aerodynamic characteristics of a model with an NACA low-drag wing. Measurements were made of lift and drag over an angle-of-attack range and for several values of dynamic pressure on a four-tenths scale model of a high-speed airplane equipped with the low-drag wing and with various wing-gun installations. Two installations were tested: one in which the blast tube and part of the gun barrel protrude ahead of the wing and another in which the guns is mounted wholly within the wing. Two types of openings for the latter installation were tested. For each installation three simulated guns were mounted in each wing. The results are given in the form of nondimensional coefficients. The installations tested appear to have little effect on the maximum-lift coefficient of the model. However, the drag coefficient shows a definite change. The least adverse effect was obtained with the completely internal mounting and small nose entrance. The results indicate that a properly designed wing-gun installation will have very little adverse effect on the aerodynamic characteristics of the low-drag wing.

  19. An in-flight investigation of ground effect on a forward-swept wing airplane

    NASA Technical Reports Server (NTRS)

    Curry, Robert E.; Moulton, Bryan J.; Kresse, John

    1989-01-01

    A limited flight experiment was conducted to document the ground-effect characteristics of the X-29A research airplane. This vehicle has an aerodynamic platform which includes a forward-swept wing and close-coupled, variable incidence canard. The flight-test program obtained results for errors in the airdata measurement and for incremental normal force and pitching moment caused by ground effect. Correlations with wind-tunnel and computational analyses were made. The results are discussed with respect to the dynamic nature of the flight measurements, similar data from other configurations, and pilot comments. The ground-effect results are necessary to obtain an accurate interpretation of the vehicle's landing characteristics. The flight data can also be used in the development of many modern aircraft systems such as autoland and piloted simulations.

  20. Tests of Aluminum-alloy Stiffened-sheet Specimens Cut from an Airplane Wing

    NASA Technical Reports Server (NTRS)

    Holt, Marshall

    1943-01-01

    The specimens used in the present tests were cut from an actual airplane wing of the stressed-skin type. The specimens thus obtained were not representative of the usual type of laboratory specimens because the stiffeners were not exactly parallel nor evenly spaced and, in one case, the skin consisted of pieces of sheet of different thicknesses. The test data obtained indicate that the buckling strain of stiffened curved sheet can be computed with reasonable accuracy by the equation given by Wenzek. The ultimate loads of the specimens when tested as flat sheet were within +/-11 percent of the product of the compressive yield strength and the cross-sectional area of the stiffeners. A rivet spacing equal to 98 times the sheet thickness was a source of weakness, and rivet spacings up to 36 times the sheet thickness appeared satisfactory.

  1. Measurements in Flight of the Pressure Distribution on the Right Wing of a Pursuit-Type Airplane at Several Values of Mach Number

    NASA Technical Reports Server (NTRS)

    Clousing, Lawrence A; Turner, William N; Rolls, L Stewart

    1946-01-01

    Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.

  2. Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

    NASA Technical Reports Server (NTRS)

    Voracek, David

    2007-01-01

    A viewgraph presentation of flight tests performed on the F/A active aeroelastic wing airplane is shown. The topics include: 1) F/A-18 AAW Airplane; 2) F/A-18 AAW Control Surfaces; 3) Flight Test Background; 4) Roll Control Effectiveness Regions; 5) AAW Design Test Points; 6) AAW Phase I Test Maneuvers; 7) OBES Pitch Doublets; 8) OBES Roll Doublets; 9) AAW Aileron Flexibility; 10) Phase I - Lessons Learned; 11) Control Law Development and Verification & Validation Testing; 12) AAW Phase II RFCS Envelopes; 13) AAW 1-g Phase II Flight Test; 14) Region I - Subsonic 1-g Rolls; 15) Region I - Subsonic 1-g 360 Roll; 16) Region II - Supersonic 1-g Rolls; 17) Region II - Supersonic 1-g 360 Roll; 18) Region III - Subsonic 1-g Rolls; 19) Roll Axis HOS/LOS Comparison Region II - Supersonic (open-loop); 20) Roll Axis HOS/LOS Comparison Region II - Supersonic (closed-loop); 21) AAW Phase II Elevated-g Flight Test; 22) Region I - Subsonic 4-g RPO; and 23) Phase II - Lessons Learned

  3. Comparison of theoretical and flight-measured local flow aerodynamics for a low-aspect-ratio fin

    NASA Technical Reports Server (NTRS)

    Johnson, J. B.; Sandlin, D. R.

    1984-01-01

    Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes, a two dimensional transonic code called Code H, and a three dimensional subsonic and supersonic code call wing-body. Pressure distributions generated by the codes for the flight test fixture as well as boundary layer displacement thickness generated by the two dimensional code were compared to the flight test data. The two dimensional code pressure distributions compared well except at the minimum pressure point and trailing edge. Shock locations compared well except at high transonic speeds. The three dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture. The two dimensional code does not predict displacement thickness of the flight test fixture well.

  4. Flight investigation of the effects of an outboard wing-leading-edge modification on stall/spin characteristics of a low-wing, single-engine, T-tail light airplane

    NASA Technical Reports Server (NTRS)

    Stough, H. Paul, III; Dicarlo, Daniel J.; Patton, James M., Jr.

    1987-01-01

    Flight tests were performed to investigate the change in stall/spin characteristics due to the addition of an outboard wing-leading-edge modification to a four-place, low-wing, single-engine, T-tail, general aviation research airplane. Stalls and attempted spins were performed for various weights, center of gravity positions, power settings, flap deflections, and landing-gear positions. Both stall behavior and wind resistance were improved compared with the baseline airplane. The latter would readily spin for all combinations of power settings, flap deflections, and aileron inputs, but the modified airplane did not spin at idle power or with flaps extended. With maximum power and flaps retracted, the modified airplane did enter spins with abused loadings or for certain combinations of maneuver and control input. The modified airplane tended to spin at a higher angle of attack than the baseline airplane.

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

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  6. In-flight total forces, moments and static aeroelastic characteristics of an oblique-wing research airplane

    NASA Technical Reports Server (NTRS)

    Curry, R. E.; Sim, A. G.

    1984-01-01

    A low-speed flight investigation has provided total force and moment coefficients and aeroelastic effects for the AD-1 oblique-wing research airplane. The results were interpreted and compared with predictions that were based on wind tunnel data. An assessment has been made of the aeroelastic wing bending design criteria. Lateral-directional trim requirements caused by asymmetry were determined. At angles of attack near stall, flow visualization indicated viscous flow separation and spanwise vortex flow. These effects were also apparent in the force and moment data.

  7. Pressure distribution on wing ribs of the VE-7 and TS airplanes in flight Part II : pull-ups

    NASA Technical Reports Server (NTRS)

    Rhode, R V

    1928-01-01

    This paper is the second of a series of notes, each of which presents the complete results of pressure distribution tests made at Langley Field by the National Advisory Committee for Aeronautics, on wing and tail ribs of the VE-7 and TS airplanes for a particular maneuver of flight. The results for pull-ups are presented in the form of curves which show the variation of pressure distribution, total loads, normal acceleration and center of pressure with respect to time.

  8. Spin-Tunnel Investigation of the Spinning Characteristics of Typical Single-Engine General Aviation Airplane Designs. 1. Low-Wing Model A: Effects of Tail Configurations

    NASA Technical Reports Server (NTRS)

    Burk, S. M., Jr.; Bowman, J. S., Jr.; White, W. L.

    1977-01-01

    The effects of tail design on spin and recovery were investigated in a spin tunnel. A 1/11-scale model of a research airplane which represents a typical low-wing, single engine, light general aviation airplane was used. A tail design criterion for satisfactory spin recovery for light airplanes was evaluated. The effects of other geometric design features on the spin and recovery characteristics were also determined. Results indicate that the existing tail design criterion for light airplanes, which uses the tail damping power factor as a parameter, cannot be used to predict spin-recovery characteristics.

  9. Geometric Model for a Parametric Study of the Blended-Wing-Body Airplane

    NASA Technical Reports Server (NTRS)

    Mastin, C. Wayne; Smith, Robert E.; Sadrehaghighi, Ideen; Wiese, Micharl R.

    1996-01-01

    A parametric model is presented for the blended-wing-body airplane, one concept being proposed for the next generation of large subsonic transports. The model is defined in terms of a small set of parameters which facilitates analysis and optimization during the conceptual design process. The model is generated from a preliminary CAD geometry. From this geometry, airfoil cross sections are cut at selected locations and fitted with analytic curves. The airfoils are then used as boundaries for surfaces defined as the solution of partial differential equations. Both the airfoil curves and the surfaces are generated with free parameters selected to give a good representation of the original geometry. The original surface is compared with the parametric model, and solutions of the Euler equations for compressible flow are computed for both geometries. The parametric model is a good approximation of the CAD model and the computed solutions are qualitatively similar. An optimal NURBS approximation is constructed and can be used by a CAD model for further refinement or modification of the original geometry.

  10. Active Aeroelastic Wing Aerodynamic Model Development and Validation for a Modified F/A-18A Airplane

    NASA Technical Reports Server (NTRS)

    Cumming, Stephen B.; Diebler, Corey G.

    2005-01-01

    A new aerodynamic model has been developed and validated for a modified F/A-18A airplane used for the Active Aeroelastic Wing (AAW) research program. The goal of the program was to demonstrate the advantages of using the inherent flexibility of an aircraft to enhance its performance. The research airplane was an F/A-18A with wings modified to reduce stiffness and a new control system to increase control authority. There have been two flight phases. Data gathered from the first flight phase were used to create the new aerodynamic model. A maximum-likelihood output-error parameter estimation technique was used to obtain stability and control derivatives. The derivatives were incorporated into the National Aeronautics and Space Administration F-18 simulation, validated, and used to develop new AAW control laws. The second phase of flights was used to evaluate the handling qualities of the AAW airplane and the control law design process, and to further test the accuracy of the new model. The flight test envelope covered Mach numbers between 0.85 and 1.30 and dynamic pressures from 600 to 1250 pound-force per square foot. The results presented in this report demonstrate that a thorough parameter identification analysis can be used to improve upon models that were developed using other means. This report describes the parameter estimation technique used, details the validation techniques, discusses differences between previously existing F/A-18 models, and presents results from the second phase of research flights.

  11. Effects of discontinuous drooped wing leading-edge modifications on the spinning characteristics of a low-wing general aviation airplane

    NASA Technical Reports Server (NTRS)

    Dicarlo, D. J.; Stough, H. P., III; Patton, J. M., Jr.

    1980-01-01

    Wind tunnel and flight tests were conducted to determine the effects of several discontinuous drooped wing leading-edge configurations on the spinning characteristics of a light, single-engine, low-wing research airplane. Particular emphasis was placed on the identification of modifications which would improve the spinning characteristics. The spanwise length of a discontinuous outboard droop was varied and several additional inboard segments were added to determine the influence of such leading-edge configurations on the spin behavior. Results of the study indicated that the use of only the discontinuous outboard droop, over a specific spanwise area, was most effective towards improving spin and spin recovery characteristics, whereas the segmented configurations having both inboard and outboard droop exhibited a tendency to enter a flat spin.

  12. Experimental and analytical study on the flutter and gust response characteristics of a torsion-free-wing airplane model. [in the Langley transonic dynamics tunnel

    NASA Technical Reports Server (NTRS)

    Murphy, A. C.

    1981-01-01

    Experimental data and correlative analytical results on the flutter and gust response characteristics of a torsion-free-wing (TFW) fighter airplane model are presented. TFW consists of a combined wing/boom/canard surface and was tested with the TFW free to pivot in pitch and with the TFW locked to the fuselage. Flutter and gust response characteristics were measured in the Langley Transonic Dynamics Tunnel with the complete airplane model mounted on a cable mount system that provided a near free flying condition. Although the lowest flutter dynamic pressure was measured for the wing free configuration, it was only about 20 deg less than that for the wing locked configuration. However, no appreciable alleviation of the gust response was measured by freeing the wing.

  13. Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Matranga, Gene J.; Armstrong, Neil A.

    1959-01-01

    A series of landings was performed with a straight-wing airplane to evaluate the effect of low lift-drag ratios on approach and landing characteristics. Landings with a peak lift-drag ratio as low as 3 were performed by altering the airplane configuration (extending speed brakes, flaps, and gear and reducing throttle setting). As lift-drag ratio was reduced, it was necessary either to make the landing pattern tighter or to increase initial altitude, or both. At the lowest lift-drag ratio the pilots believed a 270 deg overhead pattern was advisable because of the greater ease afforded in visually positioning the airplane. The values of the pertinent flare parameters increased with the reduction of lift-drag ratio. These parameters included time required for final flare; speed change during final flare; and altitude, glide slope, indicated airspeed, and vertical velocity at initiation of final flare. The pilots believed that the tolerable limit was reached with this airplane in the present configuration, and that if, because of a further reduction in lift-drag ratio, more severe approaches than those experienced in this program were attempted, additional aids would be required to determine the flare-initiation point.

  14. Pressure Distribution over a Wing and Tail Rib of a VE-7 and of a TS Airplane in Flight

    NASA Technical Reports Server (NTRS)

    Crowley, J W , Jr

    1928-01-01

    This investigation was made to determine the pressure distribution over a rib of the wing and over a rib of the horizontal tail surface of an airplane in flight and to obtain information as to the time correlation of the loads occurring on these ribs. Two airplanes, VE-7 and TS, were selected in order to obtain the information for a thin and a thick wing section. In each case the pressure distribution was recorded for the full range of angle of attack in level flight and throughout violent maneuvers. The results show: (a) that the present rib load specifications in use by the Army Air Corps and the Bureau of Aeronautics, Navy Department, are in fair agreement with the loads actually occurring in flight, but could be slightly improved; (b) that there appears to be no definite sequence in which wing and tail surface ribs reach their respective maximum loads in different maneuvers; (c) that in accelerated flight, at air speeds less than or equal to 60 per cent of the maximum speed, the accelerations measured agree very closely with the theoretically possible maximum accelerations. In maneuvers at higher air speeds the observed accelerations were smaller than those theoretically possible. (author)

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

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  16. Flight Tests of a Model of a High-wing Transport Vertical-take-off Airplane with Tilting Wing and Propellers and with Jet Controls at the Rear of the Fuselage for Pitch and Yaw Control

    NASA Technical Reports Server (NTRS)

    Lovell, Powell M , Jr; Parlett, Lysle P

    1957-01-01

    An investigation of the stability and control of a high-wing transport vertical-take-off airplane with four engines during constant-altitude transitions from hovering to normal forward flight was conducted with a remotely controlled free-flight model. The model had four propellers distributed along the wing with the thrust axes in the wing chord plane. The wing could be rotated to 90 degrees incidence so that the propeller thrust axes were vertical for hovering flight. An air jet at the rear of the fuselage provided pitch and yaw control for hovering and low-speed flight.

  17. Flight determined lift and drag characteristics of an F-8 airplane modified with a supercritical wing with comparison to wind-tunnel results

    NASA Technical Reports Server (NTRS)

    Pyle, J. S.; Steers, L. L.

    1975-01-01

    Flight measurements obtained with a TF-8A airplane modified with a supercritical wing are presented for altitudes from 7.6 kilometers (25,000 feet) to 13.7 kilometers (45,000 feet), Mach numbers from 0.6 to 1.2, and Reynolds numbers from 0.8 x 10 to the 7th power to 2.3 x 10 to the 7th power. Flight results for the airplane with and without area-rule fuselage fairings are compared. The techniques used to determine the lift and drag characteristics of the airplane are discussed. Flight data are compared with wind-tunnel model results, where applicable.

  18. Spin tests of a single-engine, high-wing light airplane

    NASA Technical Reports Server (NTRS)

    Stewart, E. C.; Suit, W. T.; Moul, T. M.; Brown, P. W.

    1982-01-01

    The airplane has a relatively steep spin mode (low angle of attack) with a high load factor and high velocity. The airplane recovers almost immediately after any deviation from the prospin control positions, except for one maneuver with reduced flexibility in the elevator control system.

  19. Wind-tunnel investigation of a full-scale general aviation airplane equipped with an advanced natural laminar flow wing

    NASA Technical Reports Server (NTRS)

    Murri, Daniel G.; Jordan, Frank L., Jr.

    1987-01-01

    An investigation was conducted in the Langley 30- by 60-Foot Wind Tunnel to evaluate the performance, stability, and control characteristics of a full-scale general aviation airplane equipped with an advanced laminar flow wing. The study focused on the effects of natural laminar flow and advanced boundary layer transition on performance, stability, and control, and also on the effects of several wing leading edge modifications on the stall/departure resistance of the configuration. Data were measured over an angle-of-attack range from -6 to 40 deg and an angle-of-sideslip range from -6 to 20 deg. The Reynolds number was varied from 1.4 to 2.4 x 10 to the 6th power based on the mean aerodynamic chord. Additional measurements were made using hot-film and sublimating chemical techniques to determine the condition of the wing boundary layer, and wool tufts were used to study the wing stall characteristics. The investigation showed that large regions of natural laminar flow existed on the wing which would significantly enhance cruise performance. Also, because of the characteristics of the airfoil section, artificially tripping the wing boundary layer to a turbulent condition did not significantly effect the lift, stability, and control characteristics. The addition of a leading-edge droop arrangement was found to increase the stall angle of attack at the wingtips and, therefore, was considered to be effective in improving the stall/departure resistance of the configuration. Also the addition of the droop arrangement resulted in only minor increases in drag.

  20. Time-resolved stereo PIV measurements of the horseshoe vortex system at multiple locations in a low-aspect-ratio pin-fin array

    NASA Astrophysics Data System (ADS)

    Anderson, Corey D.; Lynch, Stephen P.

    2016-01-01

    Pin-fin arrays are a type of cooling feature found in heat exchangers, with elements (generally cylindrical or square) that span between two endwalls. Flow around the pin-fins generates highly turbulent mixing that increases convective heat transfer from the pins to the cooling flow. At the junction of a pin-fin and the endwall, a complex flow known as the horseshoe vortex (HSV) system is present. Although the HSV is a well-studied phenomenon, its behavior is not understood in the highly turbulent flow of a pin-fin array. Furthermore, the presence of close confining endwalls for low-aspect-ratio (short) pin-fins may have an impact on HSV dynamics. The present study utilized time-resolved stereo particle image velocimetry to examine the fluid dynamics of the HSV system in rows 1, 3, and 5 of a low-aspect-ratio pin-fin array, for a range of Reynolds numbers. In the first row, instantaneous flowfields indicated a clearly defined HSV at the leading edge, with dynamics similar to previous studies of bluff-body junction flows. The time-averaged HSV system moved closer to the pin with increasing Reynolds number, with more concentrated vorticity and turbulent kinetic energy (TKE). For downstream rows, there was a significant increase in the amount of mid-channel vorticity, with levels on the same order as the value in the core of the HSV. The time-averaged HSV system in downstream rows showed minimal variation with respect to either Reynolds number or row location. Regions of maximum streamwise and wall-normal turbulent fluctuations around the HSV were a result of its quasiperiodic oscillation between so-called backflow and zero-flow modes, which were present even in downstream rows despite the extremely high mid-channel turbulence. In the downstream rows, normalized TKE across the entire field of view decreased with increased Reynolds number, likely due to dissipation rates proportionally outpacing increases in mean channel velocity and Reynolds number. The flowfield

  1. On the Generation of Transport Suppression Barriers by Externally Driven Alfvén Waves in D-Shaped, Low Aspect Ratio Tokamaks

    NASA Astrophysics Data System (ADS)

    Bruma, C.; Cuperman, S.; Komoshvili, K.

    We investigate quantitatively the possibility of turbulence suppression through the generation of transport barriers in pre-heated low aspect ratio tokamaks (LARTs) by the sheared electric fields generated by externally driven rf waves in the frequency range ωA ≃ ω < ωci (ωA and ωci are, respectively, the Alfvén and ion cyclotron frequencies). To this aim the following sequential steps are followed: (1) Solutions of the resistive two-fluid model full wave equation for a realistic LART configuration (D-shape cross-section and aspect ratio, R/a ≿ 1, as well as suitably located low field side, LFS, antenna) upon using a quite general dielectric tensor operator; (2) Calculation of the ponderomotive forces and their magnetic surface averages; (3) Solution of a strongly non-linear differential equation for the quasi-stationary radial electric field, including the particle orbit squeezing effects, based on the results of steps (1) and (2); and (4) Calculation of the radial flow shear, S⊥, for both banana and potato collisional regimes.

  2. Single-stage experimental evaluation of low aspect ratio, highly loaded blading for compressors. Part 9: Stage F and stage G, volume 1

    NASA Technical Reports Server (NTRS)

    Cheatham, J. G.; Smith, J. D.; Wright, D. L.

    1976-01-01

    Two single-stage, 0.77 hub/tip ratio axial-flow compressors were tested to evaluate the effectiveness of low aspect ratio blading as a means of obtaining higher stage loadings. One compressor, designated Stage F, was comprised of circular arc blading with an aspect ratio of 0.9 for both the rotor and stator. This compressor was tested with uniform inlet flow, hub radial, tip radial, and 180 deg arc circumferential inlet distortion. The second compressor, designated Stage G, was comprised of multiple circular arc blading with an aspect ratio of 1.0 for both the rotor and stator. This compressor was tested with uniform inlet flow only. Design rotor tip speeds for Rotor F and Rotor G were 285 m/sec (934 ft/sec) and 327 m/sec (1,074 ft/sec) respectively. Both stages operated at high loading levels with adequate efficiency and operating range. The peak efficiencies and corresponding average stage diffusion factors for Stages F and G at design rotor speed were 86.4% and 84.1% and 0.59 and 0.55 respectively. The surge margin at peak efficiency for Stage F was 12.6% and the corresponding value for Stage G was 16.5%. Both stages experienced a loss in efficiency with increasing rotor speed; however, the multiple circular arc rotor delayed the characteristic loss in efficiency within increasing Mach number to higher Mach number.

  3. Aerodynamic Characteristics and Flying Qualities of a Tailless Triangular-wing Airplane Configuration as Obtained from Flights of Rocket-propelled Models at Transonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L; Stevens, Joseph E; Norris, Harry P

    1956-01-01

    A flight investigation of rocket-powered models of a tailless triangular-wing airplane configuration was made through the transonic and low supersonic speed range at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. An analysis of the aerodynamic coefficients, stability derivatives, and flying qualities based on the results obtained from the successful flight tests of three models is presented.

  4. Simulator study of the stall departure characteristics of a light general aviation airplane with and without a wing-leading-edge modification

    NASA Technical Reports Server (NTRS)

    Riley, D. R.

    1985-01-01

    A six-degree-of-freedom nonlinear simulation was developed for a two-place, single-engine, low-wing general aviation airplane for the stall and initial departure regions of flight. Two configurations, one with and one without an outboard wing-leading-edge modification, were modeled. The math models developed are presented simulation predictions and flight-test data for validation purposes and simulation results for the two configurations for various maneuvers and power settings are compared to show the beneficial influence of adding the wing-leading-edge modification.

  5. Low-speed aerodynamic characteristics of a hypersonic research airplane concept having a 70 deg swept delta wing

    NASA Technical Reports Server (NTRS)

    Creel, T. R., Jr.; Penland, J. A.

    1974-01-01

    An experimental investigation of the low-speed static longitudinal, lateral and directional stability characteristics of a hypersonic research airplane concept having a 70 deg swept delta wing was conducted in a low-speed tunnel with a 12-foot (3.66 meter) octagonal test section. Aircraft component variations included: (1) fuselage shape modifications, (2) tip fins, (3) center vertical fin, (4) wing camber, and (5) wing planform. This investigation was conducted at a dynamic pressure of 262.4 Pa (5.48 psf), a Mach number of 0.06, and a Reynolds number of 2.24 million, based on body length. Tests were conducted through an angle-of-attack range of 0 deg to 30 deg with elevon deflections from +5.0 deg to minus 30.0 deg. The complete configuration exhibited positive static longitudinal, lateral and directional stability up to angles of attack of at least 20 deg and was trimmable to lift coefficients of at least 0.70 with elevon deflections of minus 30 deg.

  6. Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

    NASA Technical Reports Server (NTRS)

    Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.

    1992-01-01

    Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

  7. Design of a Large Span-Distributed Load Flying-Wing Cargo Airplane

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.; Quartero, C. B.

    1977-01-01

    The design and operation of very large, long-range, subsonic cargo aircraft are considered. A design concept which distributes the payload along the wingspan to counterbalance the aerodynamic loads, with a resultant decrease in the in-flight wing bending moments and shear forces, is described. The decreased loading of the wing structure, coupled with the very thick wing housing the cargo, results in a relatively low overall structural weight in comparison to that of conventional aircraft.

  8. General problem of the airplane

    NASA Technical Reports Server (NTRS)

    Richard, Maurice; Richard, Paul

    1922-01-01

    A series of equations relating to airplanes are given and examples listed. Some of the equations listed include: the speed, altitude and carrying capacity of various airplanes; weight of an airplane; weight of various parts of an airplane; the polars of the wings; speeds of airplanes; radius of action.

  9. The Effects of a Highly Cambered Low-Drag Wing and of Auxiliary Flaps on the High-Speed Aerodynamic Characteristics of a Twin-Engine Pursuit Airplane Model

    NASA Technical Reports Server (NTRS)

    Ganzer, Victor M

    1944-01-01

    Results are presented for tests of two wings, an NACA 230-series wing and a highly-cambered NACA 66-series wing on a twin-engine pursuit airplane. Auxiliary control flaps were tested in combinations with each wing. Data showing comparison of high-speed aerodynamic characteristics of the model when equipped with each wing, the effect of the auxiliary control flaps on aerodynamic characteristics, and elevator effectiveness for the model with the 66-series wing are presented. High-speed aerodynamic characteristics of the model were improved with the 66-series wing.

  10. Study of belly-flaps to enhance lift and pitching moment coefficient of a Blended-Wing-Body airplane in landing and takeoff configuration

    NASA Astrophysics Data System (ADS)

    Staelens, Yann Daniel

    During the first century of flight few major changes have been made to the configuration of subsonic airplanes. A distinct fuselage with wings, a tail, engines and a landing gear persists as the dominant arrangement. During WWII some companies developed tailless all-wing airplanes. However the concept failed to advance till the late 80's when the B-2, the only flying wing to enter production to date, illustrated its benefits at least for a stealth platform. The advent of the Blended-Wing-Body (BWB) addresses the historical shortcomings of all-wing designs, specifically poor volume utility and excess wetted area as a result. The BWB is now poised to become the new standard for large subsonic airplanes. Major aerospace companies are studying the concept for next generation of passenger airplanes. But there are still challenges. One is the BWB's short control lever-arm pitch. This affects rotation and go-around performances. This study presents a possible solution by using a novel type of control surface, a belly-flap, on the under side of the wing to enhance its lift and pitching moment coefficient during landing, go-around and takeoff. Increases of up to 30% in lift-off CL and 8% in positive pitching moment have been achieved during wind tunnel tests on a generic BWB-model with a belly-flap. These aerodynamic improvements when used in a mathematical simulation of landing, go-around and takeoff procedure were showing reduction in landing-field-length by up to 22%, in takeoff-field-length by up to 8% and in loss in altitude between initiation of rotation and actual rotation during go-around by up to 21.5%.

  11. Configuration heating for a hypersonic research airplane concept having a 70 deg swept double-delta wing

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.

    1978-01-01

    The heating on a candidate hypersonic research airplane configuration has been examined experimentally at Mach 6 by the phase-change-paint technique. The configuration has a double-delta wing with tip fins. Phase-change-paint diagrams give heating data for the model top, side, and bottom, with and without deflected elevons for an angle-of-attack range of 0 deg to 24 deg. Nominal Reynolds numbers are on the order of 15,000,000 with supplementary data at length Reynolds number of 4,000,000, which moves the model from the predominantly turbulent into the predominantly laminar regime. Also, intermediate Reynolds numbers were investigated on the lee side for one angle of attack.

  12. Design of a large span-distributed load flying-wing cargo airplane with laminar flow control

    NASA Technical Reports Server (NTRS)

    Lovell, W. A.; Price, J. E.; Quartero, C. B.; Turriziani, R. V.; Washburn, G. F.

    1978-01-01

    A design study was conducted to add laminar flow control to a previously design span-distributed load airplane while maintaining constant range and payload. With laminar flow control applied to 100 percent of the wing and vertical tail chords, the empty weight increased by 4.2 percent, the drag decreased by 27.4 percent, the required engine thrust decreased by 14.8 percent, and the fuel consumption decreased by 21.8 percent. When laminar flow control was applied to a lesser extent of the chord (approximately 80 percent), the empty weight increased by 3.4 percent, the drag decreased by 20.0 percent, the required engine thrust decreased by 13.0 percent, and the fuel consumption decreased by 16.2 percent. In both cases the required take-off gross weight of the aircraft was less than the original turbulent aircraft.

  13. Crash tests of four low-wing twin-engine airplanes with truss-reinforced fuselage structure

    NASA Technical Reports Server (NTRS)

    Williams, M. S.; Fasanella, E. L.

    1982-01-01

    Four six-place, low-wing, twin-engine, general aviation airplane test specimens were crash tested under controlled free flight conditions. All airplanes were impacted on a concrete test surface at a nomial flight path velocity of 27 m/sec. Two tests were conducted at a -15 deg flight path angle (0 deg pitch angle and 15 deg pitch angle), and two were conducted at a -30 deg flight path angle (-30 deg pitch angle). The average acceleration time histories (crash pulses) in the cabin area for each principal direction were calculated for each crash test. In addition, the peak floor accelerations were calculated for each test as a function of aircraft fuselage longitudinal station number. Anthropomorphic dummy accelerations were analyzed using the dynamic response index and severity index (SI) models. Parameters affecting the dummy restraint system were studied; these parameters included the effect of no upper torso restraint, measurement of the amount of inertia-reel strap pullout before locking, measurement of dummy chest forward motion, and loads in the restraints. With the SI model, the dummies with no shoulder harness received head impacts above the concussive threshold.

  14. The dynamic-response characteristics of a 35 degree swept-wing airplane as determined from flight measurements

    NASA Technical Reports Server (NTRS)

    Triplett, William C; Brown, Stuart C; Smith, G Allan

    1955-01-01

    The longitudinal and lateral-directional dynamic-response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind-tunnel data. Flights were made at an altitude of 35,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directional data were also obtained at 10,000 feet. The flight consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations. Experimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest effect on the dynamic response of the airplane. The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test.

  15. An Estimation of the Flying Qualities of the Kaiser Fleetwing All-Wing Airplane from Tests of a 1/7-Scale Model, TED No. NACA 2340

    NASA Technical Reports Server (NTRS)

    Brewer, Gerald W.

    1946-01-01

    An investigation of a 1/7-scale powered model of the Kaiser Fleetwing all-wing airplane was made in the Langley full-scale tunnel to provide data for an estimation of the flying qualities of the airplane. The analysis of the stability and control characteristics of the airplane has been made as closely as possible in accordance with the requirements of the Bureau of Aeronautics, Navy Department's specifications, and a summary of the more significant conclusions is presented as follows. With the normal center of gravity located at 20 percent of the mean aerodynamic chord, the airplane will have adequate static longitudinal stability, elevator fixed, for all flight conditions except for low-power operation at low speeds where the stability will be about neutral. There will not be sufficient down-elevator deflection available for trim above speeds of about 130 miles per hour. It is probable that the reduction in the up-elevator deflections required for trim will be accompanied by reduced elevator hinge moments for low-power operation at low flight speeds. The static directional stability for this airplane will be low for all rudder-fixed or rudder-free flight conditions. The maximum rudder deflection of 30 deg will trim only about 15 deg yaw for most flight conditions and only 10 deg yaw for the condition with low power at low speeds. Also, at low powers and low speeds, it is estimated that the rudders will not trim the total adverse yaw resulting from an abrupt aileron roll using maximum aileron deflection. The airplane will meet the requirements for stability and control for asymmetric power operation with one outboard engine inoperative. The airplane would have no tendency for directional divergence but would probably be spirally unstable, with rudders fixed. The static lateral stability of the airplane will probably be about neutral for the high-speed flight conditions and will be only slightly increased for the low-power operation in low-speed flight. The

  16. A study of the problem of designing airplanes with satisfactory inherent damping of the dutch roll oscillation

    NASA Technical Reports Server (NTRS)

    Campbell, John P; Mckinney, Marion O , Jr

    1954-01-01

    Considerable interest has recently been shown in means of obtaining satisfactory stability of the dutch roll oscillation for modern high-performance airplanes without resort to complicated artificial stabilizing devices. One approach to this problem is to lay out the airplane in the earliest stages of design so that it will have the greatest practicable inherent stability of the lateral oscillation. The present report presents some preliminary results of a theoretical analysis to determine the design features that appear most promising in providing adequate inherent stability. These preliminary results cover the case of fighter airplanes at subsonic speeds. The investigation indicated that it is possible to design fighter airplanes to have substantially better inherent stability than most current designs. Since the use of low-aspect-ratio swept-back wings is largely responsible for poor dutch roll stability, it is important to design the airplane with the maximum aspect ratio and minimum sweep that will permit attainment of the desired performance. The radius of gyration in roll should be kept as low as possible and the nose-up inclination of the principal longitudinal axis of inertia should be made as great as practicable. (author)

  17. Flight Investigation of Wing-gun Fairings on a Fighter Type Airplane

    NASA Technical Reports Server (NTRS)

    White, M D

    1941-01-01

    Description is given of flight tests conducted on gun fairings, designed to correct the detrimental effects of the projecting and submerged wing guns on an F4F-3 fighter. It was found that the installation of unfaired guns on a clean wing resulted in a premature stall that increased the stalling speed in the carrier-approach and landing conditions of flight by suitably fairing the guns, it was possible to reduce the stalling speeds to values approaching very nearly the clean-wing values.

  18. Development and First Results of a new Airplane Based Fixed Wing Electromagnetic Induction Sea Ice Thickness Sounder

    NASA Astrophysics Data System (ADS)

    Rabenstein, L.; Lobach, J.; Haas, C.

    2007-12-01

    Regular observation of Arctic and Antarctic sea ice thickness is of high importance for a better understanding of processes of climate change in polar regions. For regular and accurate observations of polar sea ice thickness a long range airborne device is necessary. Airborne electromagnetic induction (AEM) sounding was found to be an ideal method for accurate and wide area sea ice thickness measurements. As a consequence of five years of successful helicopter electromagnetic (HEM) sea ice thickness measurements and to overcome helicopter range restrictions, the Alfred Wegener Institute (AWI) constructed a new airplane based fixed wing EM system. The first test flights were carried out in 2006 over the North Sea and in April 2007 in Svalbard, where the system's performance was proven under arctic conditions. The system operates in frequency domain with 1990 Hz and a vertical coplanar coil configuration. Thus the system produces a horizontal dipole. The coils are mounted beneath the wings with a separation of 11.6 meters. The airplane, a Dornier 228, is also equipped with a laser altimeter to determine the altitude of the instrument with an accuracy of 2cm. The compensation of the transmitter signal at the receiver coil is done electronically. Flights over open sea are used for the calibration of the system, because the ocean functions as a homogeneous half space with well known conductivity. A data acquisition computer records four voltages with a sample rate of 10 Hz. These are the reference voltage of the transmitter, the compensated and raw receiver voltages and the compensation signal. The laser height is recorded with a sample rate of 100 Hz to account for surface roughness. EM instruments for sea ice thickness sounding should have a vertical resolution of 10cm but due to the electrical noise caused by the airplane engines this was not easy to achieve. To account for the noise a time average filter is used. Alternatively, in order to keep the original

  19. Integration effects of pylon geometry on a high-wing transport airplane

    NASA Technical Reports Server (NTRS)

    Carlson, John R.; Lamb, Milton

    1989-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the installation effects of a series of pylons that had differing cross-sectional shapes on the pressure distributions and aerodynamic characteristics of a 1/24-scale high wing transport. The tests were conducted at Mach numbers at 0.70 and 0.80 at angles of attack from -3 degrees to 4 degrees with the pylons tested at various toe angles between 5 degrees inboard and 5 degrees outboard. Results of this study indicate that the installed drag was lowest for the pylons with a compression pylon type design which kept the flow under the wing in the pylon/wing junction comparable to the clean wing velocities.

  20. Exploratory study of the effects of wing-leading-edge modifications on the stall/spin behavior of a light general aviation airplane

    NASA Technical Reports Server (NTRS)

    1979-01-01

    Configurations with full-span and segmented leading-edge flaps and full-span and segmented leading-edge droop were tested. Studies were conducted with wind-tunnel models, with an outdoor radio-controlled model, and with a full-scale airplane. Results show that wing-leading-edge modifications can produce large effects on stall/spin characteristics, particularly on spin resistance. One outboard wing-leading-edge modification tested significantly improved lateral stability at stall, spin resistance, and developed spin characteristics.

  1. 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.

  2. Experimental and Predicted Longitudinal and Lateral-Directional Response Characteristics of a Large Flexible 35 Degree Swept-Wing Airplane at an Altitude of 35,000 Feet

    NASA Technical Reports Server (NTRS)

    Cole, Henry A , Jr; Brown, Stuart C; Holleman, Euclid C

    1957-01-01

    Measured and predicted dynamic response characteristics of a large flexible swept-wing airplane to control surface inputs are presented for flight conditions of 0.6 to 0.85 Mach number at an altitude of 35,000 feet. The report is divided into two parts. The first part deals with the response of the airplane to elevator control inputs with principal responses contained in a band of frequencies including the longitudinal short-period mode and several symmetrical structural modes. The second part deals with the response of the airplane to aileron and rudder control inputs with principal responses contained in a band of frequencies including the dutch roll mode, the rolling mode, and three antisymmetrical structural modes.

  3. Blowing-Type Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Kelly, Mark W; Anderson, Seth B; Innis, Robert C

    1958-01-01

    A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.

  4. 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.

  5. Design aspects of long range supersonic LFC airplanes with highly swept wings. [laminar flow control

    NASA Technical Reports Server (NTRS)

    Pfenninger, W.; Vemuru, C. S.

    1990-01-01

    Studies on supersonic long-range LFC (laminar flow control) aircraft were performed with the aim of maximizing L/D and alleviating sonic boom during supersonic cruise. It is found that configurations with highly swept LFC wings of very high structural aspect ratio, with the sweep increasing toward the wing root and braced externally by wide chord laminarized struts, appear especially promising. In the supersonic cruise design condition the wing upper surface isobars are swept such that the flow in the direction normal to them is transonic with embedded supersonic zones and practically shock-free over most of the span, with M-perpendicular equal to the two-dimensional design values of advanced SC LFC airfoils, e.g., of the X-787 or X-6 type.

  6. The Effect of Blunt-Trailing-Edge Modifications on the High-Speed Stability and Control Characteristics of a Swept-Wing Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Sadoff, Melvin; Matteson, Frederick H.; Van Dyke, Rudolph D., Jr.

    1954-01-01

    An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

  7. Gust-Response Analysis of an Airplane Including Wing Bending Flexibility

    DTIC Science & Technology

    1952-08-01

    reference paper which is listed therein as “ Greidanus , J. H.: ‘The Loading of firplane Structures by Symmetrical Gusts’ (in Dutch with an English summary...Arbitrary Shape. NACA TN 2036, 1950. 9. Greidanus , J. H., and Van de Vooren, A. I.: Gust Load Coefficients . for Wing and Tail Surfaces of an Aeroplane. Rep

  8. Free-spinning-tunnel Investigation of a 1/30 Scale Model of a Twin-jet-swept-wing Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Bowman, James S., Jr.; Healy, Frederick M.

    1960-01-01

    An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics of a 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat-type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be satisfactory for emergency spin recovery.

  9. Measurements of the Buffeting Loads on the Wing and Horizontal Tail of a 1/4-scale Model of the X-1E Airplane

    NASA Technical Reports Server (NTRS)

    Rainey, A Gerald; Igoe, William B

    1958-01-01

    The buffeting loads acting on the wing and horizontal tail of a 1/4-scale model of the X-1E airplane have been measured in the Langley 16-foot transonic tunnel in the Mach number range from 0.40 to 0.90. When the buffeting loads were reduced to a nondimensional aerodynamic coefficient of buffeting intensity, it was found that the maximum buffeting intensity of the horizontal tail was about twice as large as that of the wing. Comparison of power spectra of buffeting loads acting on the horizontal tail of the airplaneand of the model indicated that the model horizontal tail, which was of conventional force-test-model design, responded in an entirely different mode than did the airplane.This result implied that if quantitative extrapolation of model data to flight conditions were desired a dynamically scaled model of the rearward portion of the fuselage and empennage would be required. A study of the sources of horizontal-tail buffeting of the model indicated that the wing wake contributed a large part of the total buffeting load. At one condition it was found that removal of the wing wake would reduce the buffeting loads on the horizontal tail to about one-third of the original value.

  10. Wind-Tunnel Investigation of Some Effects of Wing Sweep and Horizontal-Tail Height on the Static Stability of an Airplane Model at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Fisher, Lewis R.; Williams, James L.

    1958-01-01

    A research model of an airplane with a configuration suitable for supersonic flight was tested at transonic speeds in order to establish the effects on longitudinal and lateral stability of certain changes in both wing sweep and height of the horizontal tail. Two wings of aspect ratio 3 and taper ratio 0.15, one having the quarter-chord line swept back 30 deg and the other 45 deg, were each tested with the horizontal tail of the model in a low and in a high position. One configuration was also tested with fuselage strakes. The tests were made at Mach numbers from 0.60 to 1.17 and Reynolds numbers from 1.9 x 10(exp 6) to 2.6 x 10(exp 6). The results indicated that a low horizontal-tail position (below the wing-chord plane) gave positive longitudinal stability for the model for all angles of attack used (angles of attack up to 24 deg); whereas, a higher tail position (above the wing-chord plane) resulted in a large reduction in stability at moderate angles of attack. With the higher horizontal tail, the 30 deg-swept-wing model had somewhat more stability than the 45 deg-swept-wing model at subsonic Mach numbers. With the lower tail, the 45 deg-swept-wing model had slightly more stability at all Mach numbers. The model with the 30 deg swept wing had greater directional stability with the tail in the higher rather than the lower position, but the opposite was true for the 45 deg-swept-wing model. The directional stability decreased sharply at high angles of attack; this characteristic was alleviated by the use of fuselage strakes which, however, proved to be detrimental to the longitudinal stability of the model tested.

  11. Automated design optimization of supersonic airplane wing structures under dynamic constraints

    NASA Technical Reports Server (NTRS)

    Fox, R. L.; Miura, H.; Rao, S. S.

    1972-01-01

    The problems of the preliminary and first level detail design of supersonic aircraft wings are stated as mathematical programs and solved using automated optimum design techniques. The problem is approached in two phases: the first is a simplified equivalent plate model in which the envelope, planform and structural parameters are varied to produce a design, the second is a finite element model with fixed configuration in which the material distribution is varied. Constraints include flutter, aeroelastically computed stresses and deflections, natural frequency and a variety of geometric limitations.

  12. Automated design optimization of supersonic airplane wing structures under dynamic constraints.

    NASA Technical Reports Server (NTRS)

    Fox, R. L.; Miura, H.; Rao, S. S.

    1972-01-01

    The problems of the preliminary and first level detail design of supersonic aircraft wings are stated as mathematical programs and solved using automated optimum design techniques. The problem is approached in two phases: the first is a simplified equivalent plate model in which the envelope, plan form and structural parameters are varied to produce a design, the second is a finite element model with fixed configuration in which the material distribution is varied. Constraints include flutter, aeroelastically computed stresses and deflections, natural frequency and a variety of geometric limitations. The Phase I objective is a combination of weight and drag while Phase II is a weight minimization.

  13. Summary of Low-Lift Drag and Directional Stability Data from Rocket Models of the Douglas XF4D-1 Airplane with and without External Stores and Rocket Packets at Mach Numbers from 0.8 to 1.38 TED No. NACA DE-349

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L.; Blanchard, Willard S.; Hastings, Earl C., Jr.

    1952-01-01

    At the request of the Bureau of Aeronautics, Department of the Navy, an investigation at transonic and low supersonic speeds of the drag and longitudinal trim characteristics of the Douglas XF4D-1 airplane is being conducted by the Langley Pilotless Aircraft Research Division. The Douglas XF4D-1 is a jet-propelled, low-aspect-ratio, swept-wing, tailless, interceptor-type airplane designed to fly at low supersonic speeds. As a part of this investigation, flight tests were made using rocket- propelled 1/10- scale models to determine the effect of the addition of 10 external stores and rocket packets on the drag at low lift coefficients. In addition to these data, some qualitative values of the directional stability parameter C(sub n beta) and duct total-pressure recovery are also presented.

  14. A blended wing body airplane with a close-coupled, tilting tail

    NASA Astrophysics Data System (ADS)

    Nasir, R. E. M.; Mazlan, N. S. C.; Ali, Z. M.; Wisnoe, W.; Kuntjoro, W.

    2016-10-01

    This paper highlights a novel approach to stabilizing and controlling pitch and yaw motion via a set of horizontal tail that can act as elevator and rudder. The tail is incorporated into a new design of blended wing body (BWB) aircraft, known as Baseline-V, located just aft of the trailing edge of its inboard wing. The proposed close-coupled tail is equipped with elevators that deflect in unison, and can tilt - an unusual means of tilting where if starboard side is tilted downward at k degree, and then the portside must be tilted upward at k degree too. A wind tunnel experiment is conducted to investigate aerodynamics and static stability of Baseline-V BWB aircraft. The model is being tested at actual flight speed of 15 m/s (54 km/h) with varying angle of attack for five elevator angle cases at zero tilt angle and varying sideslip angle for four tilt angle cases at one fixed elevator angle. The result shows that the aircraft's highest lift-to-drag ratio is 32. It is also found that Baseline-V is statically stable in pitch and yaw but has no clear indication in terms of roll stability.

  15. Grid generation and inviscid flow computation about cranked-winged airplane geometries

    NASA Technical Reports Server (NTRS)

    Eriksson, L.-E.; Smith, R. E.; Wiese, M. R.; Farr, N.

    1987-01-01

    An algebraic grid generation procedure that defines a patched multiple-block grid system suitable for fighter-type aircraft geometries with fuselage and engine inlet, canard or horizontal tail, cranked delta wing and vertical fin has been developed. The grid generation is based on transfinite interpolation and requires little computational power. A finite-volume Euler solver using explicit Runge-Kutta time-stepping has been adapted to this grid system and implemented on the VPS-32 vector processor with a high degree of vectorization. Grids are presented for an experimental aircraft with fuselage, canard, 70-20-cranked wing, and vertical fin. Computed inviscid compressible flow solutions are presented for Mach 2 at 3.79, 7 and 10 deg angles of attack. Conmparisons of the 3.79 deg computed solutions are made with available full-potential flow and Euler flow solutions on the same configuration but with another grid system. The occurrence of an unsteady solution in the 10 deg angle of attack case is discussed.

  16. Grid generation and inviscid flow computation about cranked-winged airplane geometries

    NASA Technical Reports Server (NTRS)

    Eriksson, L.-E.; Smith, R. E.; Wiese, M. R.; Farr, N.

    1987-01-01

    An algebraic grid generation procedure that defines a patched multiple-block grid system suitable for fighter-type aircraft geometries with fuselage and engine inlet, canard or horizontal tail, cranked delta wing and vertical fin has been developed. The grid generation is based on transfinite interpolation and requires little computational power. A finite-volume Euler solver using explicit Runge-Kutta time-stepping has been adapted to this grid system and implemented on the VPS-32 vector processor with a high degree of vectorization. Grids are presented for an experimental aircraft with fuselage, canard, 70-20-cranked wing, and vertical fin. Computed inviscid compressible flow solutions are presented for Mach 2 at 3.79, 7 and 10 deg angles of attack. Conmparisons of the 3.79 deg computed solutions are made with available full-potential flow and Euler flow solutions on the same configuration but with another grid system. The occurrence of an unsteady solution in the 10 deg angle of attack case is discussed.

  17. Flight Calibration of four airspeed systems on a swept-wing airplane at Mach numbers up to 1.04 by the NACA radar-phototheodolite method

    NASA Technical Reports Server (NTRS)

    Thompson, Jim Rogers; Bray, Richard S; COOPER GEORGE E

    1950-01-01

    The calibrations of four airspeed systems installed in a North American F-86A airplane have been determined in flight at Mach numbers up to 1.04 by the NACA radar-phototheodolite method. The variation of the static-pressure error per unit indicated impact pressure is presented for three systems typical of those currently in use in flight research, a nose boom and two different wing-tip booms, and for the standard service system installed in the airplane. A limited amount of information on the effect of airplane normal-force coefficient on the static-pressure error is included. The results are compared with available theory and with results from wind-tunnel tests of the airspeed heads alone. Of the systems investigated, a nose-boom installation was found to be most suitable for research use at transonic and low supersonic speeds because it provided the greatest sensitivity of the indicated Mach number to a unit change in true Mach number at very high subsonic speeds, and because it was least sensitive to changes in airplane normal-force coefficient. The static-pressure error of the nose-boom system was small and constant above a Mach number of 1.03 after passage of the fuselage bow shock wave over the airspeed head.

  18. Wave formation on a liquid layer for de-icing airplane wings

    NASA Technical Reports Server (NTRS)

    Yih, Chia-Shun

    1990-01-01

    Wave formation on a thin liquid layer used for deicing aircraft wings is investigated by studying the stability of airflow over a liquid-coated flat plate at zero angle of incidence. The ratio of the viscosity of the liquid to that of air is very high (over 500,000), and the Reynolds number based on liquid depth and air viscosity is of the order of a few thousand. Under these circumstances the analysis gives two formulas, in closed form, for the growth rate and phase velocity of the waves in terms of the wavenumber and other relevant parameters. The expected wavenumber is that for which the growth rate is the maximum. The instability is one in which the viscosity difference between the two fluids (air and liquid) plays the dominant role, and is of the kind found by Yih (1967).

  19. An Analysis of the Tracking Performances of Two Straight-wing and Two Swept-wing Fighter Airplanes with Fixed Sights in a Standardized Test Maneuver

    NASA Technical Reports Server (NTRS)

    Ziff, Howard L; Rathert, George A; Gadeberg, Burnett L

    1953-01-01

    Standard air-to-air-gunnery tracking runs were conducted with F-51H, F8F-1, F-86A, and F-86E airplanes equipped with fixed gunsights. The tracking performances were documented over the normal operating range of altitude, Mach number, and normal acceleration factor for each airplane. The sources of error were studied by statistical analyses of the aim wander.

  20. Wind-tunnel measurements of the chordwise pressure distribution and profile drag of a research airplane model incorporating a 17-percent-thick supercritical wing

    NASA Technical Reports Server (NTRS)

    Ferris, J. C.

    1973-01-01

    The Langley 8-foot transonic pressure tunnel to determine the wing chordwise pressure distribution for a 0.09-scale model of a research airplane incorporating a 17-percent-thick supercritical wing. Airfoil profile drag was determined from wake pressure measurements at the 42-percent-semispan wing station. The investigation was conducted at Mach numbers from 0.30 to 0.80 over an angle-of-attack range sufficient to include buffet onset. The Reynolds number based on the mean geometric chord varied from 2 x 10 to the 6th power at Mach number 0.30 to 3.33 x 10 to the 6th power at Mach number 0.65 and was maintained at a constant value of 3.86 x 10 to the 6th power at Mach numbers from 0.70 to 0.80. Pressure coefficients for four wing semispan stations and wing-section normal-force and pitching-moment coefficients for two semispan stations are presented in tabular form over the Mach number range from 0.30 to 0.80. Plotted chordwise pressure distributions and wake profiles are given for a selected range of section normal-force coefficients over the same Mach number range.

  1. Study of the Mutual Interaction Between a Wing Wake and an Encountering Airplane

    NASA Technical Reports Server (NTRS)

    Walden, A. B.; vanDam, C. P.

    1996-01-01

    In an effort to increase airport productivity, several wind-tunnel and flight-test programs are currently underway to determine safe reductions in separation standards between aircraft. These programs are designed to study numerous concepts from the characteristics and detection of wake vortices to the wake-vortex encounter phenomenon. As part of this latter effort, computational tools are being developed and utilized as a means of modeling and verifying wake-vortex hazard encounters. The objective of this study is to assess the ability of PMARC, a low-order potential-flow panel method, to predict the forces and moments imposed on a following business-jet configuration by a vortex interaction. Other issues addressed include the investigation of several wake models and their ability to predict wake shape and trajectory, the validity of the velocity field imposed on the following configuration, modeling techniques and the effect of the high-lift system and the empennage. Comparisons with wind-tunnel data reveal that PMARC predicts the characteristics for the clean wing-body following configuration fairly well. Non-linear effects produced by the addition of the high-lift system and empennage, however, are not so well predicted.

  2. Effect of Wing Loading and Altitude on Lateral Stability and Control Characteristics of an Airplane as Determined by Tests of a Model in the Free-Flight Tunnel

    DTIC Science & Technology

    1943-06-01

    varied from 0° to 11.5° and the vertical-tall area from 3.5 to 14 percent of the-wing area. IHTHODUOTION Recent military airplane design trends have...National Advisory Committoc for Aoronautics, Langley Jield," Va. 13 ’ REPEBIHOES 1, Bamber, Millard J.: Effect of Some PreBent-Day Air- plane Design ... Trends on Requirements for Lateral Stability. T.H. Ho. 814, HACA, 1941. 2. Bryant, L. V., and Fugsley, A. G.: The Lateral Sta- bility of Highly

  3. A study of the use of experimental stability derivatives in the calculation of the lateral disturbed motions of a swept-wing airplane and comparison with flight results

    NASA Technical Reports Server (NTRS)

    Bird, John D; Jaquet, Byron M

    1951-01-01

    An investigation was made to determine the accuracy with which the lateral flight motions of a swept-wing airplane could be predicted from experimental stability derivatives, determined in the 6-foot-diameter rolling-flow test section and 6 by 6-foot curved-flow test section of the Langley stability tunnel. In addition, determination of the significance of including the nonlinear aerodynamic effects of sideslip in the calculations of the motions was desired. All experimental aerodynamic data necessary for prediction of the lateral flight motions are presented along with a number of comparisons between flight and calculated motions caused by rudder and aileron disturbances.

  4. Experimental low-speed and calculated high-speed aerodynamic characteristics of a hypersonic research airplane concept having a 65 deg swept delta wing

    NASA Technical Reports Server (NTRS)

    Penland, J. A.; Creel, T. R., Jr.; Howard, F. G.

    1974-01-01

    An experimental wind-tunnel investigation has been carried out to determine the static longitudinal, lateral, and directional stability and control characteristics of a model of a large-body, delta-wing hypersonic research airplane concept at low speed. This investigation was conducted at a dynamic pressure of 239.4 Pa (5 psf) and a Reynolds number, based on fuselage length, of 2 million. The configuration variables included vertical fins, engine modules, canards, and a canopy. The aerodynamic results of a computer study at Mach numbers of 3 to 12 are presented.

  5. Effects of Canard 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.

    1961-01-01

    An investigation has been conducted at low subsonic speeds to study the effects of canard planform and wing-leading-edge modification on the longitudinal aerodynamic characteristics of a general research canard airplane configuration. The basic wing of the model 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 wing included addition of full-span and partial-span leading-edge chord-extensions. Two canard planforms were employed in the study; one was a 60 deg sweptback delta planform and the other was a trapezoidal planform similar to that of the basic wing. Modifications to these canards included addition of a full-span leading-edge chord-extension to the trapezoidal planform and a fence to the delta planform. For the basic-wing-trapezoidal-canard configuration, rather abrupt increases in stability occurred at about 12 deg angle of attack. A slight pitch-up tendency occurred for the delta-canard configuration at approximately 8 deg angle of attack. A comparison of the longitudinal control effectiveness for the basic-wing-trapezoidal-canard combination and for the basic-wing-delta-canard combination indicates higher values of control effectiveness at law angles of attack for the trapezoidal canard. The control effectiveness for the delta-canard configuration, however, is seen to hold up for higher canard deflections and to higher angles of attack. Use of a full-span chord-extension deflected approximately 30 deg on the trapezoidal canard greatly improved the control characteristics of this configuration and enabled a sizeable increase in trim lift to be realized.

  6. Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of the Static and Dynamic Stability Derivatives of an Airplane Model with an Unswept Wing and a High Horizontal Tail

    NASA Technical Reports Server (NTRS)

    Lessing, Henry C.; Butler, James K.

    1959-01-01

    Results are presented of a wind-tunnel investigation to evaluate the static and dynamic stability derivatives of a model with a low-aspect-ratio unswept wing and a high horizontal tail. In addition to results for the complete model, results were also obtained of the body alone, body and wing, and body and tail. Data were obtained in the Mach number range from 0.65 to 2.2, at a Reynolds number of 2 million based on the wing mean aerodynamic chord. The angle-of-attack range for most of the data was -11.5 deg to 18 deg. A limited amount of data was obtained with fixed transition. A correspondence between the damping in pitch and the static stability, previously noted in other investigations, was also observed in the present results. The effect observed was that a decrease (or increase) in the static stability was accompanied by an increase (or decrease) in the damping in pitch. A similar correspondence was observed between the damping in yaw and the static-directional stability. Results from similar tests of the same model configuration in two other facilities over different speed ranges are presented for comparison. It was found that most of the results from the three investigations correlated reasonably well. Estimates of the rotary derivatives were made using available procedures. Comparison with the experimental results indicates the need for development of more precise estimation procedures.

  7. Wind-tunnel investigation of aerodynamic characteristics and wing pressure distributions of an airplane with variable-sweep wings modified for laminar flow

    NASA Technical Reports Server (NTRS)

    Hallissy, James B.; Phillips, Pamela S.

    1989-01-01

    A wind tunnel test was conducted to evaluate the aerodynamic characteristics and wing pressure distributions of a variable wing sweep aircraft having wing panels that are modified to promote laminar flow. The modified wing section shapes were incorporated over most of the exposed outer wing panel span and were obtained by extending the leading edge and adding thickness to the existing wing upper surface forward of 60 percent chord. Two different wing configurations, one each for Mach numbers 0.7 and 0.8, were tested on the model simultaneously, with one wing configuration on the left side and the other on the right. The tests were conducted at Mach numbers 0.20 to 0.90 for wing sweep angles of 20, 25, 30, and 35 degrees. Longitudinal, lateral and directional aerodynamic characteristics of the modified and baseline configurations, and selected pressure distributions for the modified configurations, are presented in graphical form without analysis. A tabulation of the pressure data for the modified configuration is available as microfiche.

  8. 77 FR 16492 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-03-21

    ... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; Airbus Airplanes... airplanes. This proposed AD was prompted by reports that some nuts installed on the wing, including on... the airplane wings being impaired. DATES: We must receive comments on this proposed AD by May 7,...

  9. Coil configurations for low aspect ratio stellerators

    SciTech Connect

    Rome, J.; Morris, R.; Hirshman, S.; Fowler, R.; Merkel, P.

    1989-01-01

    Using the NESCOIL code, it is possible to find a surface current distribution which can create a given last closed flux surface. Thus, almost any stellarator configuration can be created with either helical or modular currents. In particular, we have succeeded in finding current distributions which generate the optimized ATF-II configuration described in Hirshman's paper. As the aspect ratio of the configuration is decreased, or the plasma-to-coil distance is increased, the harmonic content of the surface currents is increased. This makes it a challenge to cut the distribution into either discrete helical or modular coils which are attractive from an engineering point of view. Several approaches to this problem are discussed. There have been some hints of promising configurations, but to date, none of them are satisfactory. 2 refs., 7 figs.

  10. The Airplane Experiment.

    ERIC Educational Resources Information Center

    Larson, Lee; Grant, Roderick

    1991-01-01

    Presents an experiment to investigate centripetal force and acceleration that utilizes an airplane suspended on a string from a spring balance. Investigates the possibility that lift on the wings of the airplane accounts for the differences between calculated tension and measured tension on the string. (MDH)

  11. The Airplane Experiment.

    ERIC Educational Resources Information Center

    Larson, Lee; Grant, Roderick

    1991-01-01

    Presents an experiment to investigate centripetal force and acceleration that utilizes an airplane suspended on a string from a spring balance. Investigates the possibility that lift on the wings of the airplane accounts for the differences between calculated tension and measured tension on the string. (MDH)

  12. Wind-tunnel and Flight Investigations of the Use of Leading-Edge Area Suction for the Purpose of Increasing the Maximum Lift Coefficient of a 35 Degree Swept-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Holzhauser, Curt A; Bray, Richard S

    1956-01-01

    An investigation was undertaken to determine the increase in maximum lift coefficient that could be obtained by applying area suction near the leading edge of a wing. This investigation was performed first with a 35 degree swept-wing model in the wind tunnel, and then with an operational 35 degree swept-wing airplane which was modified in accord with the wind-tunnel results. The wind-tunnel and flight tests indicated that the maximum lift coefficient was increased more than 50 percent by the use of area suction. Good agreement was obtained in the comparison of the wind-tunnel results with those measured in flight.

  13. Stability Characteristics of Two Missiles of Fineness Ratios 12 and 18 with Six Rectangular Fins of Very Low Aspect Ratio Over a Mach Number Range of 1.4 to 3.2

    NASA Technical Reports Server (NTRS)

    Henning, Allen B.

    1959-01-01

    Two rocket-propelled missiles have been test flown by the Langley Pilotless Aircraft Research Division in order to study the stability characteristics of a body with six rectangular fins of very low aspect ratio. The fins, which had exposed aspect ratios of approximately o.o4 and 0.02 per fin, were mounted on bodies of fineness ratios of 12 and 18, respectively. Each body had a nose with a fineness ratio of 3.5 and a cylindrical afterbody. The body and the fin chord of the model having a fineness ratio of 12 were extended the length of 6 body diameters to produce the model with a fineness ratio of 18. The missiles were disturbed in flight by pulse rockets in order to obtain the stability data. The tests were performed over a Mach number range of 1.4 to 3.2 and a Reynolds number range of 2 x 10(exp 6) to 21 x l0(exp 6). The results of these tests indicate that these configurations with the long rectangular fins of very low aspect ratio showed little induced roll" with the missile of highest fineness ratio and longest fin chord exhibiting the least amount. Extending the body and fin chord of the shorter missile six body diameters and thereby increasing the fin area approximately 115 percent increased the lift-curve slope based on body cross-sectional area approximately 40 to 55 percent, increased the dynamic stability by a substantial amount, and increased the drag from 14 to 33 percent throughout the comparable Mach number range. The center-of-pressure location of both missiles remained constant over the Mach number range.

  14. 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.

  15. Exhaust-nozzle characterisitcs for a twin-jet variable-wing-sweep fighter airplane model at Mach numbers to 2.2

    NASA Technical Reports Server (NTRS)

    Reubush, D. E.; Mercer, C. E.

    1974-01-01

    A wind-tunnel investigation has been conducted to determine the exhaust-nozzle aerodynamic and propulsive characteristics for a twin-jet variable-wing-sweep fighter airplane model. The powered model was tested in the Langley 16-foot transonic tunnel and in the Langley 4-foot supersonic pressure tunnel at Mach numbers to 2.2 and at angles of attack from about minus 2 to 6 deg. Compressed air was used to simulate the nozzle exhaust flow at values of jet total-pressure ratio from approximately 1 (jet off) to about 21. Effects of configuration variables such as speed-brake deflection, store installation, and boundary-layer thickness on the the nozzle characteristics were also investigated.

  16. Wind-tunnel investigation of the effect of power and flaps on the static lateral characteristics of a single-engine low-wing airplane model

    NASA Technical Reports Server (NTRS)

    Tamburello, Vito; Weil, Joseph

    1947-01-01

    Tests were made in the Langley 7- by 10-foot tunnel to determine the lateral-stability characteristics with and without power of a model of a typical low-wing single-engine airplane with flaps neutral, with a full-span single slotted flap, and with a full-span double slotted flap. Power decreased the dihedral effect regardless of flap condition, and the double-slotted flap configuration showed the most marked decrease. The usual effect of power in increasing the directional stability was also shown. Deflection of the single slotted flap produced negative dihedral effect, but increased the directional stability. The effects of deflecting the double slotted flap were erratic and marked changes in both effective dihedral and directional stability occurred. The addition of the tail surfaces always contributed directional stability and generally produced positive dihedral effect.

  17. Effects of landing gear, speed brake and protuberances on the longitudinal aerodynamic characteristics of an NASA supercritical-wing research airplane model

    NASA Technical Reports Server (NTRS)

    Bartlett, D. W.; Sangiorgio, G.

    1975-01-01

    An investigation was conducted in the Langley Research Center 8-foot transonic pressure tunnel to determine the effects of the landing gear, speed brake and the major airplane protuberances on the longitudinal aerodynamic characteristics of an 0.087-scale model of the TF-8A supercritical-wing research airplane. For the effects of the landing gear and speed brake, tests were conducted at Mach numbers of 0.25 and 0.35 with a flap deflection of 20 degrees and a horizontal-tail angle of -10 degrees. These conditions simulated those required for take-off and landing. The effects of the protuberances were determined with the model configured for cruise (i.e., horizontal-tail angle of -2.5 degrees and no other control deflection), and these tests were conducted at Mach numbers from 0.50 to 1.00. The angle-of-attack range for all tests varied from about -5 degrees to 12 degrees.

  18. Reliable formulae for estimating airplane performance and the effects of changes in weight, wing area, or power

    NASA Technical Reports Server (NTRS)

    Diehl, Walter S

    1924-01-01

    This report contains the derivation and the verification of formulae for predicting the speed range ratio, the initial rate of climb, and the absolute ceiling of an airplane. Curves used in the computation are given in NACA-TR-171. Standard formulae for service ceiling, time of climb, cruising range, and endurance are also given in the conventional forms.

  19. 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.

  20. Investigation at High Subsonic Speeds of the Static Longitudinal and Lateral Stability Characteristics of Two Canard Airplane Configurations

    NASA Technical Reports Server (NTRS)

    Sleeman, William C., Jr.

    1957-01-01

    The present investigation was conducted in the Langley high-speed 7-by 10-foot tunnel to determine the static longitudinal and lateral stability characteristics at high subsonic speeds of two canard airplane configurations previously tested at supersonic speeds. The Mach number range of this investigation extended from 0.60 to 0.94 and a maximum angle-of-attack range of -2dewg to 24deg was obtained at the lowest test Mach number. Two wing plan forms of equal area were studied in the present tests; one was a 60deg delta wing and the other was a trapezoid wing having an aspect ratio of 3, taper ratio of 0.143, and an unswept 80-percent-chord line. The canard control had a trapezoidal plan form and its area was approximately 11.5 percent of the wing area. The model also had a low-aspect-ratio highly swept vertical tail and twin ventral fins. The longitudinal control characteristics of the models were consistent with past experience at low speed on canard configurations in that stalling of the canard surface occurred at moderate and high control deflections for moderate values of angle of attack. This stalling could impose appreciable limitations on the maximum trim-lift coefficient attainable. The control effectiveness and maximum value of trim-lift was significantly increased by addition of a body flap having a conical shape and located slightly behind the canard surface on the bottom of the body. Addition of the canard surface at 0deg deflection had relatively little effect on overall directional stability of the delta-wing configuration; however, deflection of the canard surface from 0deg to 10deg had a large favorable effect on directional stability at high angles of attack for both the trapezoid- and delta-wing configurations.

  1. Structural response to discrete and continuous gusts of an airplane having wing bending flexibility and a correlation of calculated and flight results

    NASA Technical Reports Server (NTRS)

    Houbolt, John C; Kordes, Eldon E

    1954-01-01

    An analysis is made of the structural response to gusts of an airplane having the degrees of freedom of vertical motion and wing bending flexibility and basic parameters are established. A convenient and accurate numerical solution of the response equations is developed for the case of discrete-gust encounter, an exact solution is made for the simpler case of continuous-sinusoidal-gust encounter, and the procedure is outlined for treating the more realistic condition of continuous random atmospheric turbulence, based on the methods of generalized harmonic analysis. Correlation studies between flight and calculated results are then given to evaluate the influence of wing bending flexibility on the structural response to gusts of two twin-engine transports and one four-engine bomber. It is shown that calculated results obtained by means of a discrete-gust approach reveal the general nature of the flexibility effects and lead to qualitative correlation with flight results. In contrast, calculations by means of the continuous-turbulence approach show good quantitative correlation with flight results and indicate a much greater degree of resolution of the flexibility effects.

  2. Aerodynamic characteristics of a hypersonic research airplane concept having a 70 degree swept double delta wing at Mach numbers from 1.50 to 2.86

    NASA Technical Reports Server (NTRS)

    Penland, J. A.; Fournier, R. H.; Marcum, D. C., Jr.

    1975-01-01

    An experimental investigation of the static longitudinal, lateral, and directional stability characteristics of a hypersonic research airplane concept having a 70 deg swept double-delta wing was conducted in the Langley unitary plan wind tunnel. The configuration variables included wing planform, tip fins, center fin, and scramjet engine modules. The investigation was conducted at Mach numbers from 1.50 to 2.86 and at a constant Reynolds number, based on fuselage length, of 3,330,000. Tests were conducted through an angle-of-attack range from about -4 deg to 24 deg with angles of sideslip of 0 deg and 3 deg and at elevon deflections of 0, -10, and -20 deg. The complete configuration was trimmable up to angles of attack of about 22 deg with the exception of regions at low angles of attack where positive elevon deflections should provide trim capability. The angle-of-attack range for which static longitudinal stability also exists was reduced at the higher Mach numbers due to the tendency of the complete configuration to pitch up at the higher angles of attack. The complete configuration was statically stable directionally up to trimmed angles of attack of at least 20 deg for all Mach numbers M with the exception of a region near 4 deg at M = 2.86 and exhibited positive effective dihedral at all positive trimmed angles of attack.

  3. Tabulated pressure measurements of a NASA supercritical-wing research airplane model with and without fuselage area-rule additions at Mach 0.25 to 1.00

    NASA Technical Reports Server (NTRS)

    Harris, C. D.; Bartlett, D. W.

    1972-01-01

    Basic pressure measurements were made on a 0.087-scale model of a supercritical wing research airplane in the Langley 8 foot transonic pressure tunnel at Mach numbers from 0.25 to 1.00 to determine the effects on the local aerodynamic loads over the wing and rear fuselage of area-rule additions to the sides of the fuselage. In addition, pressure measurements over the surface of the area-rule additions themselves were obtained at angles of sideslip of approximately - 5 deg, 0 deg, and 5 deg to aid in the structural design of the additions. Except for representative figures, results are presented in tabular form without analysis.

  4. Preliminary Results Obtained from Flight Test of a 1/7-Scale Rocket-Powered Model of the Grumman XF10F Airplane Configuration in the Swept-Wing Condition, TED No. NACA DE 354

    NASA Technical Reports Server (NTRS)

    Gardner, William N.

    1951-01-01

    A flight investigation of a 1/7-scale rocket-powered model of the XF10F Grumman XFl0F airplane in the swept-wing configuration has been made. The purpose of this test was to determine the static longitudinal stability, damping in pitch, and longitudinal control effectiveness of the airplane with the center of gravity at 20 percent of the wing mean aerodynamic chord. Only a small amount of data was obtained from the test because, immediately after booster separation at a Mach number of 0.88, the configuration was directionally unstable and diverged in sideslip. Simultaneous with the sideslip divergence, the model became longitudinally unstable at 3 degree angle of attack and -6 degree sideslip and diverged in pitch to a high angle of attack. During the pitch-up the free-floating horizontal tail became unstable at 5 degree angle of attack and the tail drifted against its positive deflection limit.

  5. Summary and Analysis of Horizontal-Tail Contribution to Longitudinal Stability of Swept-Wing Airplanes at Low Speeds

    NASA Technical Reports Server (NTRS)

    Neely, Robert H.; Griner, Roland F.

    1959-01-01

    Air-flow characteristics behind wings and wing-body combinations are described and are related to the downwash at specific tall locations for unseparated and separated flow conditions. The effects of various parameters and control devices on the air-flow characteristics and tail contribution are analyzed and demonstrated. An attempt has been made to summarize certain data by empirical correlation or theoretical means in a form useful for design. The experimental data herein were obtained mostly at Reynolds numbers greater than 4 x 10(exp 6) and at Mach numbers less than 0.25.

  6. Amphibious Airplane

    NASA Technical Reports Server (NTRS)

    1986-01-01

    The airplane pictured is the new Air Shark I, a four-place amphibian that makes extensive use of composite materials and cruises at close to 200 miles per hour under power from a 200-horsepower engine. Air Shark I is a "homebuilt" airplane, assembled from a kit of parts and components furnished by Freedom Master Corporation, Satellite Beach, Florida. The airplane incorporates considerable NASA technology and its construction benefited from research assistance provided by Kennedy Space Center (KSC) In designing the Shark, company president Arthur M. Lueck was able to draw on NASA's aeronautical technology bank through KSC's computerized "recon" library. As a result of his work at KSC, the wing of the Air Shark I is a new airfoil developed by Langley Research Center for light aircraft. In addition, Lueck opted for NASA-developed "winglets," vertical extensions of the wing that reduce drag by smoothing air turbulence at the wingtips. The NASA technology bank also contributed to the hull design. Lueck is considering application of NASA laminar flow technology-means of smoothing the airflow over wing and fuselage-to later models for further improvement of the Shark's aerodynamic efficiency. A materials engineer, Lueck employed his own expertise in designing and selecting the materials for the composite segments, which include all structural members, exposed surfaces and many control components. The materials are fiber reinforced plastics, or FRP They offer a high strength-to-weight ratio, with a nominal strength rating about one and a half times that of structural steel. They provide other advantages: the materials can be easily molded into finished shapes without expensive tooling or machining, and they are highly corrosion resistant. The first homebuilt to be offered by Freedom Master, Air Shark I completed air and water testing in mid-1985 and the company launched production of kits.

  7. The development of an augmentor wing jet STOL research airplane (modified C-8A). Volume 1: Summary

    NASA Technical Reports Server (NTRS)

    Ashleman, R. H.; Kavdahl, H.

    1972-01-01

    A project to develop an experimental aircraft for use as an inflight demonstrator of the augmentor wing, short takeoff concept is discussed. The required modifications were made on a de Havilland C-8A aircraft. The modifications to the aircraft are explained and the performance of the modified aircraft is reported.

  8. Free-Flight Tests of 0.11-Scale North American F-100 Airplane Wings to Investigate the Possibility of Flutter in Transonic Speed Range at Varying Angles of Attack

    NASA Technical Reports Server (NTRS)

    O'Kelly, Burke R.

    1954-01-01

    Free-flight tests in the transonic speed range utilizing rocketpropelled models have been made on three pairs of 0.11-scale North American F-100 airplane wings having an aspect ratio of 3.47, a taper ratio of 0.308, 45 degree sweepback at the quarter-chord line, and thickness ratios of 31 and 5 percent to investigate the possibility of flutte r. Data from tests of two other rocket-propelled models which accidentally fluttered during a drag investigation of the North American F-100 airplane are also presented. The first set of wings (5 percent thick) was tested on a model which was disturbed in pitch by a moving tail and reached a maximum Mach number of 0.85. The wings encountered mild oscillations near the first - bending frequency at high lift coefficients. The second set of wings 9 percent thick was tested up to a maximum Mach number of 0.95 at (2) angles of attack provided by small rocket motors installed in the nose of the model. No oscillations resembling flutter were encountered during the coasting flight between separation from the booster and sustainer firing (Mach numbers from 0.86 to 0.82) or during the sustainer firing at accelerations of about 8g up to the maximum Mach number of the test (0.95). The third set of wings was similar to the first set and was tested up to a maximum Mach number of 1.24. A mild flutter at frequencies near the first-bending frequency of the wings was encountered between a Mach number of 1.15 and a Mach number of 1.06 during both accelerating and coasting flight. The two drag models, which were 0.ll-scale models of the North American F-100 airplane configuration, reached a maximum Mach number of 1.77. The wings of these models had bending and torsional frequencies which were 40 and 89 percent, respectively, of the calculated scaled frequencies of the full-scale 7-percent-thick wing. Both models experienced flutter of the same type as that experienced-by the third set of wings.

  9. Cross-field potential hill arisen eccentrically in toroidal electron cyclotron resonance plasmas in the Low Aspect ratio Torus Experiment device to regulate electron and ion flows from source to boundary

    NASA Astrophysics Data System (ADS)

    Kuroda, Kengoh; Wada, Manato; Uchida, Masaki; Tanaka, Hitoshi; Maekawa, Takashi

    2015-07-01

    We have investigated the electron and ion flows in toroidal electron cyclotron resonance (ECR) plasmas maintained by a 2.45 GHz microwave power around 1 kW under a simple toroidal field in the low aspect ratio torus experiment (LATE) device. We have found that a vertically uniform ridge of electron pressure that also constitutes the source belt of electron impact ionization is formed along just lower field side of the ECR layer and a cross-field potential hill ({{V}S}\\cong 30 V while {{T}e}\\cong 10 eV), eccentrically shifted toward the corner formed by the top panel and the ECR layer, arises. Combination of the hill-driven E× B drift and the vertical drift due to the field gradient and curvature, being referred to as vacuum toroidal field (VTF) drift, realizes steady flows of electrons and ions from the source to the boundary. In particular, the ions, of which VTF drift velocity is much slower than the electron VTF drift velocity near the source belt, are carried by the E× B drift around the hill to the vicinity of the top panel, where the ion VTF drift is enhanced on the steep down slope of potential toward the top panel. On the other hand the electron temperature strongly decreases in this area. Thus the carrier of VTF drift current is replaced from the electrons to the ions before the top panel, enabling the current circulation through the top and bottom panels and the vessel (electrons mainly to the bottom and ions mainly to the top) that keeps the charge neutrality very high. A few percent of electrons from the source turn around the hill by 360 degree and reentry the source belt from the high field side as seed electrons for the impact ionization, keeping the discharge stable.

  10. Full-scale-wind-tunnel Tests of a 35 Degree Sweptback Wing Airplane with High-velocity Blowing over the Training-edge Flaps

    NASA Technical Reports Server (NTRS)

    Kelley, Mark W; Tolhurst, William H JR

    1955-01-01

    A wind-tunnel investigation was made to determine the effects of ejecting high-velocity air near the leading edge of plain trailing-edge flaps on a 35 degree sweptback wing. The tests were made with flap deflections from 45 degrees to 85 degrees and with pressure ratios across the flap nozzles from sub-critical up to 2.9. A limited study of the effects of nozzle location and configuration on the efficiency of the flap was made. Measurements of the lift, drag, and pitching moment were made for Reynolds numbers from 5.8 to 10.1x10(6). Measurements were also made of the weight rate of flow, pressure, and temperature of the air supplied to the flap nozzles.The results show that blowing on the deflected flap produced large flap lift increments. The amount of air required to prevent flow separation on the flap was significantly less than that estimated from published two-dimensional data. When the amount of air ejected over the flap was just sufficient to prevent flow separation, the lift increment obtained agreed well with linear inviscid fluid theory up to flap deflections of 60 degrees. The flap lift increment at 85 degrees flap deflection was about 80 percent of that predicted theoretically.With larger amounts of air blown over the flap, these lift increments could be significantly increased. It was found that the performance of the flap was relatively insensitive to the location of the flap nozzle, to spacers in the nozzle, and to flow disturbances such as those caused by leading-edge slats or discontinuities on the wing or flap surfaces. Analysis of the results indicated that installation of this system on an F-86 airplane is feasible.

  11. Investigation in the 7-By-10 Foot Wind Tunnel of Ducts for Cooling Radiators Within an Airplane Wing, Special Report

    NASA Technical Reports Server (NTRS)

    Harris, Thomas A.; Recant, Isidore G.

    1938-01-01

    An investigation was made in the NACA 7- by 10-foot wind tunnel of a large-chord wing model with a duct to house a simulated radiator suitable for a liquid-cooled engine. The duct was expanded to reduce the radiator losses, and the installation of the duct and radiator was made entirely within the wing to reduce form and interference drag. The tests were made using a two-dimensional flow set-up with a full-span duct and radiator. Section aerodynamic characteristics of the basic airfoil are given and also curves showing the characteristics of the various duct-radiator combinations. An expression for efficiency, the primary criterion of merit of any duct, and the effect of the several design parameters of the duct-radiator arrangement are discussed. The problem of throttling is considered and a discussion of the power required for cooling is included. It was found that radiators could be mounted in the wing and efficiently pass enough air for cooling with duct outlets located at any point from 0.25c to 0.70c from the wing leading edge on the upper surface. The duct-inlet position was found to be critical and, for maximum efficiency, had to be at the stagnation point of the airfoil and to change with flight attitude. The flow could be efficiently throttled only by a simultaneous variation of duct inlet and outlet sizes and of inlet position. It was desirable to round both inlet and outlet lips. With certain arrangements of duct, the power required for cooling at high speed was a very low percentage of the engine power.

  12. Preliminary study of propulsion systems and airplane wing parameters for a US Navy subsonic V/STOL aircraft

    NASA Technical Reports Server (NTRS)

    Zola, C. L.; Fishbach, L. H.; Allen, J. L.

    1978-01-01

    Two V/STOL propulsion concepts were evaluated in a common aircraft configuration. One propulsion system consists of cross coupled turboshaft engines driving variable pitch fans. The other system is a gas coupled combination of turbojet gas generators and tip turbine fixed pitch fans. Evaluations were made of endurance at low altitude, low speed loiter with equal takeoff fuel loads. Effects of propulsion system sizing, bypass ratio, and aircraft wing planform parameters were investigated and compared. Shaft driven propulsion systems appear to result in better overall performance, although at higher installed weight, than gas systems.

  13. Subsonic Aerodynamic Characteristics of an Airplane Configuration with a 63 deg Sweptback Wing and Twin-Boom Tails

    NASA Technical Reports Server (NTRS)

    Savage, Howard F.; Edwards, George G.

    1959-01-01

    A wind-tunnel investigation has been conducted to determine the effects of an unconventional tail arrangement on the subsonic static longitudinal and lateral stability characteristics of a model having a 63 deg sweptback wing of aspect ratio 3.5 and a fuselage. Tail booms, extending rearward from approximately the midsemispan of each wing panel, supported independent tail assemblies well outboard of the usual position at the rear of the fuselage. The horizontal-tail surfaces had the leading edge swept back 45 deg and an aspect ratio of 2.4. The vertical tail surfaces were geometrically similar to one panel of the horizontal tail. For comparative purposes, the wing-body combination was also tested with conventional fuselage-mounted tail surfaces. The wind-tunnel tests were conducted at Mach numbers from 0.25 to 0.95 with a Reynolds number of 2,000,000, at a Mach number of 0.46 with a Reynolds number of 3,500,000, and at a Mach number of 0.20 with a Reynolds number of 7,000,000. The results of the investigation indicate that longitudinal stability existed to considerably higher lift coefficients for the outboard tail configuration than for the configuration with conventional tail. Wing fences were necessary with both configurations for the elimination of sudden changes in longitudinal stability at lift coefficients between 0.3 and 0.5. Sideslip angles up to 15 deg had only small effects upon the pitching-moment characteristics of the outboard tail configuration. There was an increase in the directional stability for the outboard tail configuration at the higher angles of attack as opposed to a decrease for the conventional tail configuration at most of the Mach numbers and Reynolds numbers of this investigation. The dihedral effect increased rapidly with increasing angle of attack for both the outboard and the conventional tail configurations but the increase was greater for the outboard tail configuration. The data indicate that the outboard tail is an effective

  14. The Aerodynamic Forces on Slender Plane- and Cruciform-Wing and Body Combinations

    NASA Technical Reports Server (NTRS)

    Spreiter, John R

    1950-01-01

    The load distribution, forces, and moments are calculated theoretically for inclined slender wing-body combinations consisting of a slender body of revolution and either a plane or cruciform arrangement of low-aspect-ratio pointed wings. The results are applicable at subsonic and transonic speeds, and at supersonic speeds, provided the entire wing-body combination lies near the center of the Mach cone.

  15. Summary of the Aerodynamic Characteristics and Flying Qualities Obtained from Flights of Rocket-Propelled Models of an Airplane Configuration Incorporating a Sweptback Inversely Tapered Wing at Transonic and Low-Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitcham, Grady L.; Blanchard, Willard S., Jr.

    1950-01-01

    Flight tests have been conducted on rocket-propelled models of an airplane configuration incorporating a sweptback wing with inverse taper to investigate the drag, stability, and control characteristics at transonic and supersonic speeds. The models were tested with a conventional tail arrangement in the Mach number range from 0.55 to 1.2. In addition to the various aerodynamic parameters obtained, the flying qualities were computed for a full-scale airplane with the center-of-gravity location at 18 percent of the mean aerodynamic chord. Also, included in this investigation are drag measurements made on relatively simple fixed-control models tested with both conventional and V-tail arrangements.

  16. Spin-tunnel investigation of the spinning characteristics of typical single-engine general aviation airplane designs. 2: Low-wing model A; tail parachute diameter and canopy distance for emergency spin recovery

    NASA Technical Reports Server (NTRS)

    Burk, S. M., Jr.; Bowman, J. S., Jr.; White, W. L.

    1977-01-01

    A spin tunnel study is reported on a scale model of a research airplane typical of low-wing, single-engine, light general aviation airplanes to determine the tail parachute diameter and canopy distance (riser length plus suspension-line length) required for energency spin recovery. Nine tail configurations were tested, resulting in a wide range of developed spin conditions, including steep spins and flat spins. The results indicate that the full-scale parachute diameter required for satisfactory recovery from the most critical conditions investigated is about 3.2 m and that the canopy distance, which was found to be critical for flat spins, should be between 4.6 and 6.1 m.

  17. Testing a Windmill Airplane ("autogiro")

    NASA Technical Reports Server (NTRS)

    Seiferth, R

    1927-01-01

    In order to clear up the matter ( In the Spanish report it was stated that the reference surface for the calculation of the coefficients c(sub a) and c(sub w) was the area of all four wings, instead of a single wing), the model of a windwill airplane was tested in the Gottingen wind tunnel.

  18. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps. Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical...

  19. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps. Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical...

  20. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps. Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical...

  1. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps. Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical...

  2. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps. Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical...

  3. Aerodynamic Characteristics in Pitch and Sideslip at High Subsonic Speeds of a 1/14-Scale Model of the Grumman XF104 Airplane with Wing Sweepback of 42.5 Degrees

    NASA Technical Reports Server (NTRS)

    Kuhn, Richard E.; Draper, John W.

    1953-01-01

    An investigation has been made at high subsonic speeds of the aerodynamic'characteristics in pitch and sideslip of a l/l4-scale model of the Grumman XF10F airplane with a wing sweepback angle of 42.5. The longitudinal stability characteristics (with the horizontal tail fixed) indicate a pitch-up near the stall; however, this was somewhat alleviated by the addition of fins to the side of the fuselage below the horizontal tail. The original model configuration became directionally unstable for small sideslip angles at Mach numbers above 0.8; however, the instability was eliminated by several different modifications.

  4. DOUGLAS XA3D-1 #413 AIRPLANE.

    NASA Image and Video Library

    1955-07-27

    DOUGLAS XA3D-1 #413 AIRPLANE MOUNTED IN THE NACA AMES RESEARCH CENTER'S 40X80_FOOT SUBSONIC WIND TUNNEL sweptback wing Testing the wing boundary layer control of the A3D in the 40 x 80 wind tunnel. Boundary layer control was added to increase the lift of the wing for aircraft carrier take off and landing.

  5. Development of spin resistance criteria for light general aviation airplanes

    NASA Technical Reports Server (NTRS)

    Dicarlo, D. J.; Stough, H. P., III; Glover, K. E.; Brown, P. W.; Patton, J. M., Jr.

    1986-01-01

    A brief history of stall/spin technology for light general aviation airplanes is presented. Criteria needed to describe desirable characteristics of a spin-resistant airplane and means to evaluate airplanes for compliance with the criteria have been developed. Initial results from limited flight tests of an experimental high-wing airplane indicated that the basic configuration would not meet the spin resistance criteria. Further tests with the airplane modified to enhance its spin resistance are planned.

  6. 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.

  7. Experimental transonic flutter characteristics of two 72 deg-sweep delta-wing models

    NASA Technical Reports Server (NTRS)

    Doggett, Robert V., Jr.; Soistmann, David L.; Spain, Charles V.; Parker, Ellen C.; Silva, Walter A.

    1989-01-01

    Transonic flutter boundaries are presented for two simple, 72 deg. sweep, low-aspect-ratio wing models. One model was an aspect-ratio 0.65 delta wing; the other model was an aspect-ratio 0.54 clipped-delta wing. Flutter boundaries for the delta wing are presented for the Mach number range of 0.56 to 1.22. Flutter boundaries for the clipped-delta wing are presented for the Mach number range of 0.72 to 0.95. Selected vibration characteristics of the models are also presented.

  8. 14 CFR 23.201 - Wings level stall.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Wings level stall. 23.201 Section 23.201... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Flight Stalls § 23.201 Wings level... airplane stalls. (b) The wings level stall characteristics must be demonstrated in flight as...

  9. 14 CFR 23.201 - Wings level stall.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Wings level stall. 23.201 Section 23.201... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Flight Stalls § 23.201 Wings level... airplane stalls. (b) The wings level stall characteristics must be demonstrated in flight as...

  10. 14 CFR 23.201 - Wings level stall.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Wings level stall. 23.201 Section 23.201... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Flight Stalls § 23.201 Wings level... unreversed use of the directional control, up to the time the airplane stalls. (b) The wings level...

  11. 14 CFR 23.201 - Wings level stall.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Wings level stall. 23.201 Section 23.201... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Flight Stalls § 23.201 Wings level... airplane stalls. (b) The wings level stall characteristics must be demonstrated in flight as...

  12. 14 CFR 23.201 - Wings level stall.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Wings level stall. 23.201 Section 23.201... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Flight Stalls § 23.201 Wings level... airplane stalls. (b) The wings level stall characteristics must be demonstrated in flight as...

  13. 78 FR 25377 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-05-01

    ... and -900ER series airplanes. This AD was prompted by a report of leaking fuel from the wing leading... drain path in the wing leading edge area, forward of the wing front spar, and doing all applicable... wing leading edge, and draining inboard and onto the engine exhaust nozzle, which could result in...

  14. Measured moments of inertia of 32 airplanes

    NASA Technical Reports Server (NTRS)

    Gracey, William

    1940-01-01

    A compilation of the experimentally determined moments of inertia of 32 airplanes is presented. The measurements were obtained at the laboratories of the naca by means of a pendulum method. The airplanes tested are representative of several types of aircraft of gross weight less than 10,000 pounds. The results are presented in coefficient as well as in dimensional form. An elementary analysis of the data disclosed the possibility of grouping the results according to wing type of the airplane, as low-wing monoplanes, parasol and high-wing monoplanes, and biplanes. The data are shown to provide a convenient means of rapidly estimating the moments of inertia of other airplanes. A three view drawing of each of the 32 airplanes is included.

  15. 77 FR 24367 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-04-24

    ... airplanes. This AD was prompted by reports of cracking in the forward lug wing of the aft bearing at rib 5... fit in the forward lug wing of the aft bearing at rib 5 of the MLG on the right-hand (RH) and left-hand (LH) wing. We are issuing this AD to prevent cracking of the forward lug wing of the aft...

  16. New Albatross commercial airplane "L 58"

    NASA Technical Reports Server (NTRS)

    Meyer, G

    1923-01-01

    The "L 58" is a monoplane with cantilever wings joined directly to the fuselage. It accordingly belongs to the new school of airplane construction, as founded and developed in Germany. A list of performance characteristics is included.

  17. Measurement of the handling characteristics of two light airplanes

    NASA Technical Reports Server (NTRS)

    1980-01-01

    A flight investigation of the handling characteristics of two single engine general aviation airplanes, one a high wing and the other a low wing, included a variety of measurements of different characteristics of the airplanes. The characteristics included those of the control systems, performance, longitudinal and lateral responses, and stall motions.

  18. An Investigation of the Free-Spinning and Recovery Characteristics of a 1/24-Scale Model of the Grumman F11F-1 Airplane with Alternate Nose Configurations with and without Wing Fuel Tanks, TED No. NACA AD 395

    NASA Technical Reports Server (NTRS)

    Bowman, James S., Jr.

    1958-01-01

    A supplementary investigation has been conducted in the langley 20-foot free-spinning tunnel on a l/24-scale model of the Grumman F11F-1 airplane to determine the spin and recovery characteristics with alternate nose configurations, the production version and the elongated APS-67 version, with and without empty and full wing tanks. When spins were obtained with either alternate nose configuration, they were oscillatory and recovery characteristics were considered unsatisfactory on the basis of the fact that very slow recoveries were indicated to be possible. The simultaneous extension of canards near the nose of the model with rudder reversal was effective in rapidly terminating the spin. The addition of empty wing tanks had little effect on the developed spin and recovery characteristics. The model did not spin erect with full wing tanks. For optimum recovery from inverted spins, the rudder should be reversed to 22O against the spin and simultaneously the flaperons should be moved with the developed spin; the stick should be held at or moved to full forward longitudinally. The minimum size parachute required to insure satisfactory recoveries in an emergency was found to be 12 feet in diameter (laid out flat) with a drag coefficient of 0.64 (based on the laid-out-flat diameter) and a towline length of 32 feet.

  19. Automated airplane surface generation

    SciTech Connect

    Smith, R.E.; Cordero, Y.; Jones, W.

    1996-12-31

    An efficient methodology and software axe presented for defining a class of airplane configurations. A small set of engineering design parameters and grid control parameters govern the process. The general airplane configuration has wing, fuselage, vertical tall, horizontal tail, and canard components. Wing, canard, and tail surface grids axe manifested by solving a fourth-order partial differential equation subject to Dirichlet and Neumann boundary conditions. The design variables are incorporated into the boundary conditions, and the solution is expressed as a Fourier series. The fuselage is described by an algebraic function with four design parameters. The computed surface grids are suitable for a wide range of Computational Fluid Dynamics simulation and configuration optimizations. Both batch and interactive software are discussed for applying the methodology.

  20. Wind-Tunnel Investigation at Low Speed of the Effects of Chordwise Wing Fences and Horizontal-Tail Position on the Static Longitudinal Stability Characteristics of an Airplane Model with a 35 Degree Sweptback Wing

    NASA Technical Reports Server (NTRS)

    Queijo, M J; Jaquet, Byron M; Wolhart, Walter D

    1954-01-01

    Low-speed tests of a model with a wing swept back 35 degrees at the 0.33-chord line and a horizontal tail located well above the extended wing-chord plane indicated static longitudinal instability at moderate angles of attack for all configurations tested. An investigation therefore was made to determine whether the longitudinal stability could be improved by the use of chordwise wing fences, by lowering the horizontal tail, or by a combination of both. The results of the investigation showed that the longitudinal stability characteristics of the model with slats retracted could be improved at moderate angles of attack by placing chordwise wing fences at a spanwise station of about 73 percent of the wing semispan from the plane of symmetry provided the nose of the fence extended slightly beyond or around the wing leading edge.

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

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  2. Lightplane Wing Design

    NASA Technical Reports Server (NTRS)

    1992-01-01

    Venture, a kit airplane designed and manufactured by Questair, is a high performance lightplane with excellent low speed characteristics and enhanced safety due to NASA technology incorporated in its unusual wing design. In 1987, North Carolina State graduate students and Langley Research Center spent seven months researching and analyzing the Venture. The result was a wing modification, improving control and providing more usable lift. The plane subsequently set 10 world speed records.

  3. 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.

  4. STOL Characteristics of a Propeller-Driven, Aspect-Ratio-10, Straight-Wing Airplane with Boundary-Layer Control Flaps, as Estimated from Large-Scale Wind-Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Weiberg, James A; Holzhauser, Curt A.

    1961-01-01

    A study is presented of the improvements in take-off and landing distances possible with a conventional propeller-driven transport-type airplane when the available lift is increased by propeller slipstream effects and by very effective trailing-edge flaps and ailerons. This study is based on wind-tunnel tests of a 45-foot span, powered model, with BLC on the trailing-edge flaps and controls. The data were applied to an assumed airplane with four propellers and a wing loading of 50 pounds per square foot. Also included is an examination of the stability and control problems that may result in the landing and take-off speed range of such a vehicle. The results indicated that the landing and take-off distances could be more than halved by the use of highly effective flaps in combination with large amounts of engine power to augment lift (STOL). At the lowest speeds considered (about 50 knots), adequate longitudinal stability was obtained but the lateral and directional stability were unsatisfactory. At these low speeds, the conventional aerodynamic control surfaces may not be able to cope with the forces and moments produced by symmetric, as well as asymmetric, engine operation. This problem was alleviated by BLC applied to the control surfaces.

  5. Effect of tip vortices on membrane vibration of flexible wings with different aspect ratios

    NASA Astrophysics Data System (ADS)

    Genç, Mustafa Serdar; Hakan Açikel, Halil; Demir, Hacımurat; Özden, Mustafa; Çağdaş, Mücahit; Isabekov, Iliasbek

    2016-03-01

    In this study, the effect of the aspect ratio on the aerodynamics characteristic of flexible membrane wings with different aspect ratios (AR = 1 and AR = 3) is experimentally investigated at Reynolds number of 25000. Time accurate measurements of membrane deformation using Digital Image Correlation system (DIC) is carried out while normal forces of the wing will be measured by helping a load-cell system and flow on the wing was visualized by means of smoke wire technic. The characteristics of high aspect ratio wings are shown to be affected by leading edge separation bubbles at low Reynolds number. It is concluded that the camber of membrane wing excites the separated shear layer and this situation increases the lift coefficient relatively more as compared to rigid wings. In membrane wings with low aspect ratio, unsteadiness included tip vortices and vortex shedding, and the combination of tip vortices and vortex shedding causes complex unsteady deformations of these membrane wings. The characteristic of high aspect ratio wings was shown to be affected by leading edge separation bubbles at low Reynolds numbers whereas the deformations of flexible wing with low aspect ratio affected by tip vortices and leading edge separation bubbles.

  6. Wing design for spin resistance

    NASA Technical Reports Server (NTRS)

    Stough, H. P., III; Dicarlo, D. J.; Glover, K. E.; Stewart, E. C.

    1984-01-01

    Use of a discontinuous outboard wing leading edge to improve stall/spin characteristics has been evaluated through wind-tunnel and flight tests. Addition of such a discontinuous outboard wing leading-edge droop design to three light airplanes having NACA 6-series airfoil sections produced significant improvements in stall characteristics and spin resistance. The increased spin resistance of the modified airplanes has been related to the difference in angle of attack between the outer wing panel stall and the maximum attainable angle of attack.

  7. 78 FR 76734 - Special Conditions: Bombardier Inc., Models BD-500-1A10 and BD-500-1A11 Series Airplanes; Seats...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-12-19

    ...: Alan Sinclair, FAA, Airplane and Cabin Safety Branch, ANM-115, Transport Airplane Directorate, Aircraft...- series airplanes are swept-wing monoplanes with a pressurized cabin. They share an identical supplier... is 41,000 feet for both model airplanes. The interior arrangements of the C-series airplanes...

  8. Determination of the Stability and Control Characteristics of a Tailless All-Wing Airplane Model with Sweepback in the Langley Free-Flight Tunnel

    NASA Technical Reports Server (NTRS)

    Seacord, Charles L.; Campbell, John P.

    1945-01-01

    Force and flight tests were performance on an all-wing model with windmilling propellers. Tests were conducted with deflected and retracted flaps, with and without auxiliary vertical tail surfaces, and with different centers of gravity and trim coefficients. Results indicate serious reduction of stick-fixed longitudinal stability because of wing-tip stalling at high lift coefficient. Directional stability without vertical tail is undesirably low. Low effective dihedral should be maintained. Elevator and rudder control system is satisfactory.

  9. A Flight Evaluation of the Longitudinal Stability Characteristics Associated with the Pitch-up of a Swept-Wing Airplane in Maneuvering Flight at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Anderson, Seth B; Bray, Richard S

    1955-01-01

    This report presents the results of flight measurements of longitudinal stability and control characteristics made on a swept-wing jet aircraft to determine the origin of the pitch-up encountered in maneuvering flight at transonic speeds. For this purpose measurements were made of elevator angle, tail angle of attack, and wing-fuselage pitching moments (obtained from measurements of the balancing tail loads).

  10. An experimental study of separated flow on a finite wing

    NASA Technical Reports Server (NTRS)

    Winkelmann, A. E.

    1981-01-01

    The flow field associated with the formation of a mushroom shaped trailing edge stall cell on a low-aspect-ratio (AR = 4.0) wing was investigated in a series of low speed wind tunnel tests (Reynolds number based on 15.2 cm chord = 480,000). Flow field surveys of the separation bubble and wake of a partially stalled and fully stalled wing were completed using a hot-wire probe, a split-film probe, and a directional sensitive pressure probe. A new color video display technique was developed to display the flow field survey data. Photographs were obtained of surface oil flow patterns and smoke flow visualization

  11. Full-Scale Wind-Tunnel Investigation of a Jet Flap in Conjunction with a Plain Flap with Blowing Boundary-Layer Control on a 35 deg Sweptback-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Aoyagi, Kiyoshi; Hickey, David H.

    1959-01-01

    Previous investigations have shown that increased blowing at the hinge-line radius of a plain flap will give flap lift increases above that realized with boundary-layer control. Other experiments and theory have shown that blowing from a wing trailing edge, through the jet flap effect, produced lift increases. The present investigation was made to determine whether blowing simultaneously at the hinge-line radius and trailing edge would be more effective than blowing separately at either location. The tests were made at a Reynolds number of 4.5 x 10(exp 6) with a 35 deg sweptback-wing airplane. For this report, only the lift data are presented. Of the three flap blowing arrangements tested, blowing distributed between the trailing edge and the hinge-line radius of a plain flap was found to be superior to blowing at either location separately at the plain flap deflections of interest. Comparison of estimated and experimental jet flap effectiveness was fair.

  12. Preliminary Results of an Investigation by the Wing-Flow Method of the Longitudinal Stability Characteristics of a 1/50-Scale Semispan Model of the McDonnell XP-88 Airplane

    NASA Technical Reports Server (NTRS)

    Crane, Harold L.

    1948-01-01

    This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.

  13. Longitudinal Stability and Control Characteristics of a Semispan Wind-Tunnel Model of the XF7U-1 Airplane and a Comparison with Complete-Model Wind-Tunnel Tests and Semispan-Model Wing-Flow Tests

    NASA Technical Reports Server (NTRS)

    Goodson, Kenneth W.; King, Thomas J., Jr.

    1949-01-01

    An investigation was conducted on an 0.08-scale semispan model of the Chance Vought XF7U-1 airplane in the Langley high-speed 7- by 10-foot tunnel in the Mach number range from 0.40 to 0.97. The results are compared with those obtained with an 0.08-scale sting-mounted complete model tested in the same tunnel and with an 0.026-scale semispan model tested by the wing-flow method. The lift-curve slopes obtained for the 0.08-scale semispan model and the 0.026-scale wing-flow model were in good agreement but both were generally lower than the values obtained for the sting model. The results of an unpublished investigation have shown that tunnel-wall boundary-layer and strut-leakage effects can came the difference noted between the lift-curve slopes of the sting and the semispan data. Fair agreement was obtained among the data of the three models as regard the variation of pitching-moment coefficients with lift coefficient. The agreement between the complete and the semispan models was more favorable with the vertical fine on, because the wall-boundary-layer and strut leakage effects were less severe. In the Mach number range between 0.94 and 0.97, ailavator-control reversal was indicated in the wing-flow data near zero lift; Whereas, these same trends were indicated in the larger scale semispan data at somewhat higher lift coefficients.

  14. Longitudinal Stability and Control Characteristics of a Semispan Model of the XF7U-1 Tailless Airplane at Transonic Speeds by the NACA Wing-Flow Method, TED No. NACA DE307

    NASA Technical Reports Server (NTRS)

    Sawyer, Richard H.; Trant, James P., Jr.

    1947-01-01

    An investigation was made by the NACA wing-flow method to determine the longitudinal stability and control characteristics at transonic speeds of a semispan model of the XF7U-1 tailless airplane. The 25-percent chord line of the wing of the model was swept back 35 deg. The airfoil sections of the wing perpendicular to the 25-percent chord line were 12 percent thick. Measurements were made of the normal force and pitching moment through an angle-of-attack range from about -3 deg to 14 deg for several ailavator deflections at Mach numbers from 0.65 to about 1.08. The results of the tests indicated no adverse effects of compressibility up to a Mach number of at least 0.85 at low normal-force coefficients and small ailavator deflections. Up to a Mach number of 0.85, the neutral point at low normal-force coefficients was at about 25 percent of the mean aerodynamic chord and moved rearward irregularly to 41 or 42 percent with a further increase in Mach number to about 1.05. For deflections up to -8.0 percent, the ailavator was effective in changing the pitching moment except at Mach numbers from 0.93 to about 1.0 where ineffectiveness or reversal was indicated for deflections and normal-force coefficients. With -13.2 deg deflection at normal-force coefficients above about 0.3, reversal of ailavator effectiveness occurred at Mach numbers as low as 0.81. A nose-down trim change, which began at a Mach number of about 0.85, together with the loss in effectiveness of the ailavator, indicated that with increase in the Mach number from about 0.95 to 1.05 an abrupt ailavator movement of 5 deg or 6 deg first up and then down would be required to maintain level flight.

  15. Chordwise and compressibility corrections for arbitrary planform slender wings

    NASA Technical Reports Server (NTRS)

    Levin, D.; Seginer, A.

    1982-01-01

    The Lomax and Sluder method for adapting slender-wing theory to delta or rectangular wings by making chordwise and compressibility corrections is extended to cover wings of any arbitrary planform in subsonic and supersonic flows. The numerical accuracy of the present work is better than that of the Lomax-Sluder results. Comparison of the results of this work with those of the vortex-lattice method and Kernel function method for a family of Gothic and arrowhead wings shows good agreement. A universal curve is proposed for the evaluation of the lift coefficient of a low aspect ratio wing of an arbitrary planform in subsonic flow. The location of the center of pressure can also be estimated.

  16. Chordwise and compressibility corrections for arbitrary planform slender wings

    NASA Technical Reports Server (NTRS)

    Levin, D.; Seginer, A.

    1982-01-01

    The Lomax and Sluder method for adapting slender-wing theory to delta or rectangular wings by making chordwise and compressibility corrections is extended to cover wings of any arbitrary planform in subsonic and supersonic flows. The numerical accuracy of the present work is better than that of the Lomax-Sluder results. Comparison of the results of this work with those of the vortex-lattice method and Kernel function method for a family of Gothic and arrowhead wings shows good agreement. A universal curve is proposed for the evaluation of the lift coefficient of a low aspect ratio wing of an arbitrary planform in subsonic flow. The location of the center of pressure can also be estimated.

  17. A comparison of flight and simulation data for three automatic landing system control laws for the Augmentor wing jet STOL research airplane

    NASA Technical Reports Server (NTRS)

    Feinreich, B.; Gevaert, G.

    1980-01-01

    Automatic flare and decrab control laws for conventional takeoff and landing aircraft were adapted to the unique requirements of the powered lift short takeoff and landing airplane. Three longitudinal autoland control laws were developed. Direct lift and direct drag control were used in the longitudinal axis. A fast time simulation was used for the control law synthesis, with emphasis on stochastic performance prediction and evaluation. Good correlation with flight test results was obtained.

  18. 77 FR 15638 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-03-16

    ... airplanes. That NPRM proposed to require modifying the fluid drain path in the wing leading edge area, forward of the wing front spar, and doing all applicable related investigative and corrective actions. That NPRM was prompted by a report of leaking fuel from the wing leading edge area at the inboard...

  19. The Ultra-Low Aspect Ratio Stellarator SCR-1

    NASA Astrophysics Data System (ADS)

    Ribeiro, Celso; Vargas, Ivan; Mora, Jaime; Zamora, Esteban; Asenjo, Jose; Ribas, Leonardo; Guadamuz, Saul

    2012-10-01

    The world most compact stellarator is currently being designed at the Costa Rica Institute of Technology (ITCR). The SCR-1(Stellarator of Costa Rica 1) is a 2-field period modular device with a circular cross-section vessel (Ro=0.238m, a=0.097m, Ro{/a≈ 2.5}, 0.014m3, {4mm} thickness 6061-T6 aluminum). The expected D-shaped high elongated plasma cross section has a maximum average radius of < a> ≈ 0.062m, leading to Ro/< a> ≥ 3.8. Such compactness was reached after a SCR-1 earlier proposal [1] was redesigned, both based on the low shear stellarator UST/1: Ro/< a> ≈ 6, ι =0.32/0.28 (core/edge) [2]. The set field at centre is 88mT produced by 12 copper modular coils, 8.7kA-turn each. This field is EC resonant at Ro with a 2.45GHz μ w, 1st harmonic, from 2/3kW magnetrons which will produce a second time-scale plasma pulse. The coil current will be produced by a bank of cell batteries. Poincar'{e} and EC deposition plots will be presented using COMSOL Multiphysics software. SCR-1 will be synergetic to the ST MEDUSA currently under donation to ITCR [3]. Both will benefit of the local new activities in technological plasmas.[4pt] [1] Barillas L et al., Proc.19th Int.Conf. Nucl.Eng., Japan, 2011[0pt] [2] Queral V, Stellarator News, 118, 2008[0pt] [3] Ribeiro C et al., 54th APS, Plasma Phys. Div., US, 2012

  20. Characterization of peeling modes in a low aspect ratio tokamak

    NASA Astrophysics Data System (ADS)

    Bongard, M. W.; Thome, K. E.; Barr, J. L.; Burke, M. G.; Fonck, R. J.; Hinson, E. T.; Redd, A. J.; Schlossberg, D. J.

    2014-11-01

    Peeling modes are observed at the plasma edge in the Pegasus Toroidal Experiment under conditions of high edge current density (Jedge ˜ 0.1 MA m-2) and low magnetic field (B ˜ 0.1 T) present at near-unity aspect ratio. Their macroscopic properties are measured using external Mirnov coil arrays, Langmuir probes and high-speed visible imaging. The modest edge parameters and short pulse lengths of Pegasus discharges permit direct measurement of the internal magnetic field structure with an insertable array of Hall-effect sensors, providing the current profile and its temporal evolution. Peeling modes generate coherent, edge-localized electromagnetic activity with low toroidal mode numbers n ⩽ 3 and high poloidal mode numbers, in agreement with theoretical expectations of a low-n external kink structure. Coherent MHD fluctuation amplitudes are found to be strongly dependent on the experimentally measured Jedge/B peeling instability drive, consistent with theory. Peeling modes nonlinearly generate ELM-like, field-aligned filamentary structures that detach from the edge and propagate radially outward. The KFIT equilibrium code is extended with an Akima spline profile parameterization and an improved model for induced toroidal wall current estimation to obtain a reconstruction during peeling activity with its current profile constrained by internal Hall measurements. It is used to test the analytic peeling stability criterion and numerically evaluate ideal MHD stability. Both approaches predict instability, in agreement with experiment, with the latter identifying an unstable external kink.

  1. Liquid crystalline tetrahedra and low-aspect-ratio organic materials

    NASA Astrophysics Data System (ADS)

    Pegenau, A.; Goering, P.; Diele, S.; Tschierske, C.

    1998-01-01

    3,4-Dialkoxybenzoates of pentaerythritol and structural related polyhydroxy compounds have been synthesized and investigated by polarizing microscopy, differential scanning calorimetry and some of them also by x-ray scattering. Despite the fact that these compounds are neither anisometric nor amphilphilic in a classical sense they can display columnar liquid-crystalline phases. It is proposed, that the segregation of the polar regions and the lipophilic alkyl chains into separate regions is mainly responsible for the formation of these mesophases.

  2. Aerodynamic performance of a transonic low aspect ratio turbine nozzle

    SciTech Connect

    Moustapha, S.H. . Turbine Aerodynamics); Carscallen, W.E. . Combustion and Fluids Engineering Lab.); McGeachy, J.D. . Dept. of Mechanical Engineering)

    1993-07-01

    This paper presents detailed information of the three-dimensional flow field in a realistic turbine nozzle with an aspect ratio of 0.65 and a turning angle of 76 deg. The nozzle has been tested in a large-scale planar cascade over a range of exit Mach numbers from 0.3 to 1.3. The experimental results are presented in the form of nozzle passage Mach number distributions and spanwise distribution of losses and exit flow angle. Details of the flow field inside the nozzle passage are examined by means of surface flow visualization and Schlieren pictures. The performance of the nozzle is compared to the data obtained for the same nozzle tested in an annular cascade and a stage environment. Excellent agreement is found between the measured pressure distribution and the prediction of a three-dimensional Euler flow solver.

  3. A Survey of Factors Affecting Blunt Leading-Edge Separation for Swept and Semi-Slender Wings

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2010-01-01

    A survey is presented of factors affecting blunt leading-edge separation for swept and semi-slender wings. This class of separation often results in the onset and progression of separation-induced vortical flow over a slender or semi-slender wing. The term semi-slender is used to distinguish wings with moderate sweeps and aspect ratios from the more traditional highly-swept, low-aspect-ratio slender wing. Emphasis is divided between a selection of results obtained through literature survey a section of results from some recent research projects primarily being coordinated through NATO s Research and Technology Organization (RTO). An aircraft context to these studies is included.

  4. Wind-Tunnel Tests of the 1/25-Scale Powered Model of the Martin JRM-1 Airplane. IV - Tests with Ground Board and with Modified Wing and Hull - TED No. NACA 232. Part 4; Tests with Ground Board and with Modified Wing and Hull, TED No. NACA 232

    NASA Technical Reports Server (NTRS)

    Lockwood, Vernard E.; Smith, Bernard J.

    1947-01-01

    Wind-tunnel tests were made of a 1/25 scale model of the Martin JRM-1 airplane to determine: (1) The longitudinal stability and control characteristics of the JRM-1 model near the water and lateral and directional stability characteristics with power while moving on the surface of the water, the latter being useful for the design of tip floats; (2) The stability and stalling characteristics of the wing with a modified airfoil contour; (3) Stability characteristics of a hull of larger design gross weight; The test results indicated that the elevator was powerful enough to trim the original model in a landing configuration at any lift coefficient within the specified range of centers of gravity. The ground-board tests for evaluating the aerodynamic forces and moments on an airplane in a simulated cross wind indicate a high dihedral effect in the presence of the ground board and, consequently, during low-speed taxying and take-off, large overturning moments would result which would have to be overcome by the tip floats.

  5. The evolution of airplanes

    NASA Astrophysics Data System (ADS)

    Bejan, A.; Charles, J. D.; Lorente, S.

    2014-07-01

    The prevailing view is that we cannot witness biological evolution because it occurred on a time scale immensely greater than our lifetime. Here, we show that we can witness evolution in our lifetime by watching the evolution of the flying human-and-machine species: the airplane. We document this evolution, and we also predict it based on a physics principle: the constructal law. We show that the airplanes must obey theoretical allometric rules that unite them with the birds and other animals. For example, the larger airplanes are faster, more efficient as vehicles, and have greater range. The engine mass is proportional to the body size: this scaling is analogous to animal design, where the mass of the motive organs (muscle, heart, lung) is proportional to the body size. Large or small, airplanes exhibit a proportionality between wing span and fuselage length, and between fuel load and body size. The animal-design counterparts of these features are evident. The view that emerges is that the evolution phenomenon is broader than biological evolution. The evolution of technology, river basins, and animal design is one phenomenon, and it belongs in physics.

  6. Theoretical-Numerical Study of Feasibility of Use of Winglets on Low Aspect Ration Wings at Subsonic and Transonic Mach Numbers to Reduce Drag

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Liaw, Paul; Cerney, Michael J.

    1988-01-01

    A numerical design study was conducted to assess the drag reduction potential of winglets installed on a series of low aspect ratio wings at a design point of M=0.8, C sub L=0.3. Wing-winglet and wing-alone design geometries were obtained for wings of aspect ratios between 1.75 and 2.67, having leading edge sweep angles between 45 and 60 deg. Winglet length was fixed at 15% of wing semispan. To assess the relative performance between wing-winglet and wing-alone configurations, the PPW nonlinear extended small disturbance potential flow code was utilized. This model has proven to yield plausible transonic flow field simulations for the series of low aspect ratio configurations selected. Predicted decreases in pressure drag coefficient for the wing-winglet configurations relative to the corresponding wing-alone planform are about 15% at the design point. Predicted decreases in wing-winglet total drag coefficient are about 12%, relative to the corresponding wing-alone design. Longer winglets (25% of the wing semispan) yielded decreases in the pressure drag of up to 22% and total drag of up to 16.4%. These predicted drag coefficient reductions are comparable to reductions already demonstrated by actual winglet designs installed on higher aspect ratio transport type aircraft.

  7. Drag measurements on a Junkers wing section : application of the Betz Method to the results of comparative tests made on a model and on an airplane in flight

    NASA Technical Reports Server (NTRS)

    Weidinger, Hanns

    1927-01-01

    The comparison of model tests in flight can be based on the result of such measurements. They are very important from the aerodynamical point of view, as they lead to useful conclusions regarding the behavior of the wing, its best shape and the conformity of theoretical and actual flow. Although there still remains a certain prejudice against such measurements, I have still attempted to make these comparative tests in order to inspire confidence in their reliability.

  8. Precision controllability of the F-15 airplane

    NASA Technical Reports Server (NTRS)

    Sisk, T. R.; Matheny, N. W.

    1979-01-01

    A flying qualities evaluation conducted on a preproduction F-15 airplane permitted an assessment to be made of its precision controllability in the high subsonic and low transonic flight regime over the allowable angle of attack range. Precision controllability, or gunsight tracking, studies were conducted in windup turn maneuvers with the gunsight in the caged pipper mode and depressed 70 mils. This evaluation showed the F-15 airplane to experience severe buffet and mild-to-moderate wing rock at the higher angles of attack. It showed the F-15 airplane radial tracking precision to vary from approximately 6 to 20 mils over the load factor range tested. Tracking in the presence of wing rock essentially doubled the radial tracking error generated at the lower angles of attack. The stability augmentation system affected the tracking precision of the F-15 airplane more than it did that of previous aircraft studied.

  9. Recent applications of the transonic wing analysis computer code, TWING

    NASA Technical Reports Server (NTRS)

    Subramanian, N. R.; Holst, T. L.; Thomas, S. D.

    1982-01-01

    An evaluation of the transonic-wing-analysis computer code TWING is given. TWING utilizes a fully implicit approximate factorization iteration scheme to solve the full potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations were analyzed, and the limits of applicability of this code was evaluated. Comparisons of computed results were made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative full potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.

  10. The efficient solution of transonic wing flow fields

    NASA Technical Reports Server (NTRS)

    Holst, T. L.; Subramanian, N. R.; Thomas, S. D.

    1983-01-01

    An evaluation of the transonic-wing-analysis computer code TWING is presented. TWING utilizes a fully implicit, approximate-factorization iteration scheme to solve the full-potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations have been analyzed, and comparisons of computed results have been made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative, full-potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations, including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.

  11. 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)

  12. 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.

  13. A study of commuter airplane design optimization

    NASA Technical Reports Server (NTRS)

    Roskam, J.; Wyatt, R. D.; Griswold, D. A.; Hammer, J. L.

    1977-01-01

    Problems of commuter airplane configuration design were studied to affect a minimization of direct operating costs. Factors considered were the minimization of fuselage drag, methods of wing design, and the estimated drag of an airplane submerged in a propellor slipstream; all design criteria were studied under a set of fixed performance, mission, and stability constraints. Configuration design data were assembled for application by a computerized design methodology program similar to the NASA-Ames General Aviation Synthesis Program.

  14. Shape, flapping and flexion: wing and fin design for forward flight.

    PubMed

    Combes, S A; Daniel, T L

    2001-06-01

    Both kinematics and morphology are critical determinants of performance in flapping flight. However, the functional consequences of changes in these traits are not yet well understood. Traditional aerodynamic studies of planform wing shape have suggested that high-aspect-ratio wings generate more force per area and perform more efficiently than low-aspect-ratio wings, but these analyses may neglect critical components of flapping flight such as unsteady fluid dynamics and wing or fin flexion. In this paper, we use an unsteady potential flow analysis that incorporates wing flexion to test predictions of optimal wing shape under varying degrees of unsteady motion and wing flexion. We focus on forward flapping flight and examine the effects of wing/fin morphology and movements on thrust generation and efficiency. We test the model by comparing our predictions with kinematic data derived from the aquatic flight of the ratfish Hydrolagus colliei. Our analyses show that aspect ratio and the proportion of area in the outer one-fifth of the wing can characterize wing shape in terms of aero- or hydrodynamic performance. By comparing the performance of wings that vary in these two parameters, we find that traditional predictions of optimal wing shape are valid only under limited circumstances (when flapping frequency is low, wings are stiff or wings are tapered at the tips). This indicates a complex relationship between locomotor traits and performance and helps explain the diversity of wing kinematics and morphologies observed in nature.

  15. Revolution in airplane construction? Grob G110: The first modern fiber glass composition airplane shortly before its maiden flight

    NASA Technical Reports Server (NTRS)

    Dorpinghaus, R.

    1982-01-01

    A single engine two passenger airplane, constructed completely from fiber reinforced plastic materials is introduced. The cockpit, controls, wing profile, and landing gear are discussed. Development of the airframe is also presented.

  16. Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint

    NASA Technical Reports Server (NTRS)

    Daum, Fred L.; Zalovcik, John A.

    1946-01-01

    Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

  17. An Investigation of the Aerodynamic Characteristics of an 0.08-Scale Model of the Chance Vought XF7U-1 Airplane in the Langley High-Speed 7- by 10-Foot Tunnel. Part V - Wing-Alone Tests and Effect of Modifications to the Vertical Fins, Speed Brakes, and Fuselage TED No. NACA DE308. Part V; Wing-Alone Tests and Effect of Modifications to the Vertical Fins, Speed Brakes, and Fuselage, TED No. NACA DE308

    NASA Technical Reports Server (NTRS)

    Kuhri, Richard E.; Myers, Boyd C., II

    1947-01-01

    Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane. The wing-alone tests and the effect of the various vertical-fin modifications, speed-brake modifications, and fuselage modifications on the aerodynamic characteristics in pitch and yaw are presented in the present paper with a limited analysis of the results. Also included are tuft studies of the flow for some of the modifications tested.

  18. Weight estimation techniques for composite airplanes in general aviation industry

    NASA Technical Reports Server (NTRS)

    Paramasivam, T.; Horn, W. J.; Ritter, J.

    1986-01-01

    Currently available weight estimation methods for general aviation airplanes were investigated. New equations with explicit material properties were developed for the weight estimation of aircraft components such as wing, fuselage and empennage. Regression analysis was applied to the basic equations for a data base of twelve airplanes to determine the coefficients. The resulting equations can be used to predict the component weights of either metallic or composite airplanes.

  19. Flight test results for an advanced technology light airplane

    NASA Technical Reports Server (NTRS)

    Kohlman, D. L.

    1977-01-01

    A single-engine light airplane was modified by the installation of a wing with reduced area, Fowler flaps, Kruger flaps, and spoilers. Flight test results show that zero-lift drag was reduced 13.8% and a trimmed maximum lift coefficient of 2.73 was achieved. Gust response was significantly reduced and excellent roll control was achieved with spoilers. Several design features employed in the new wings have excellent potential for incorporation in future light airplanes.

  20. 78 FR 14644 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-03-07

    ... fuel vapors, could result in a center wing fuel tank explosion, and consequent loss of the airplane... this regard. ] Request To Clarify Impacted Fuel Tank Boeing requested that the ``Summary'' and... fuel tank explosion, and consequent loss of the airplane,'' to clarify the area that might be...

  1. 78 FR 15658 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-03-12

    ... airplanes. This proposed AD was prompted by reports of spanwise cracks and corrosion in the wing center box... other airplanes, repetitive inspections for damage, cracking, and corrosion of the pressure seal; and repair if necessary. We are proposing this AD to detect and correct cracking and corrosion of the...

  2. 14 CFR 23.71 - Glide: Single-engine airplanes.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Glide: Single-engine airplanes. 23.71... Glide: Single-engine airplanes. The maximum horizontal distance traveled in still air, in nautical miles... with the engine inoperative, its propeller in the minimum drag position, and landing gear and wing...

  3. 14 CFR 23.71 - Glide: Single-engine airplanes.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Glide: Single-engine airplanes. 23.71... Glide: Single-engine airplanes. The maximum horizontal distance traveled in still air, in nautical miles... with the engine inoperative, its propeller in the minimum drag position, and landing gear and wing...

  4. 14 CFR 23.71 - Glide: Single-engine airplanes.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Glide: Single-engine airplanes. 23.71... Glide: Single-engine airplanes. The maximum horizontal distance traveled in still air, in nautical miles... with the engine inoperative, its propeller in the minimum drag position, and landing gear and wing...

  5. Preliminary Low-Speed Wind-Tunnel Investigation of Some Aspects of the Aerodynamic Problems Associated with Missiles Carried Externally in Positions Near Airplane Wings

    NASA Technical Reports Server (NTRS)

    Alford, William J., Jr.; Silvers, H. Norman; King, Thomas J., Jr.

    1954-01-01

    A low-speed wind-tunnel investigation has been made of some aspects of the aerodynamic problems associated with the use of air-to-air missiles when carried externally on aircraft. Measurements of the forces and moments on a missile model for a range of positions under the mid-semispan location of a 45deg sweptback wing indicated longitudinal and lateral forces with regard to both carriage and release of the missiles. Surveys of the characteristics of the flow field in the region likely to be traversed by the missiles showed abrupt gradients in both flow angularity and in local dynamic pressure. Through the use of aerodynamic data on the isolated missile and the measured flow-field characteristics, the longitudinal forces and moments acting on the missile while in the presence of the wing-fuselage combination could be estimated with fair accuracy. Although the lateral forces and moments predicted were qualitatively correct, there existed some large discrepancies in absolute magnitude.

  6. 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.

  7. 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.

  8. Longitudinal Stability Characteristics of a 1/40-Scale Model of a Proposed Configuration of the XF-91 Airplane Measured by the Wing-Flow Method

    NASA Technical Reports Server (NTRS)

    Crane, Harold L.; Beckhardt, Arnold R.

    1948-01-01

    This report presents the results of an investigation in the transonic speed range of the longitudinal stability characteristics of a proposed configuration for the Republic XF-91 airplane. The tests covered a Mach number range of 0.55 to 1.05 and a Reynolds number range from 400,000 to 1,375,000. Lift, pitching-moment, and rolling-moment characteristics of the half model and the hinge moments on the all-moving tail were measured. The downwash factor delta x epsilon / delta x alpha at the tail was determined from the pitching-moment data. A calculation of the elevator deflection and stick force required for trim was also made. It was found that the variation of force and moment coefficients was linear through the test angle-of-attack range of -1 deg to 8 deg at any Mach number; that the stability increased markedly at Mach numbers above 0.85; that the effectiveness of the tail in producing pitching moments decreased about one-third with increasing Mach numbers and that the value of the downwash factor, delta x epsilon / delta x alpha, at the tail decreased from about 0.35 at a Mach number of 0.85 to about zero at a Mach number near 0.95 and became slightly negative at higher Mach numbers. The calculated values of stick force per g and elevator deflection per g, assuming no aerodynamic balance, increased rapidly above a Mach number of 0.85.

  9. Airplane Ear

    MedlinePlus

    ... severe hearing loss Ringing in your ear (tinnitus) Spinning sensation (vertigo) Vomiting resulting from vertigo Bleeding from ... the back of the nasal cavity and the top of the throat meet (nasopharynx). When an airplane ...

  10. Giant airplanes

    NASA Technical Reports Server (NTRS)

    Warner, Edward P

    1922-01-01

    It is hardly possible for the most imaginative aeronautical enthusiast to look forward to a time when the airplane will have reached the dimensions commensurate with those already attained by the airship.

  11. 14 CFR 25.1403 - Wing icing detection lights.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Wing icing detection lights. 25.1403... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Equipment Lights § 25.1403 Wing icing... ice on the parts of the wings that are critical from the standpoint of ice accumulation....

  12. The Effects of Horizontal-Tail Location and Wing Modifications on the High-Speed Stability and Control Characteristics of a 01.17-Scale Model of the McDonnell XF2H-1 Airplane (TED No, NACA DE336)

    NASA Technical Reports Server (NTRS)

    Emerson, Horace F.; Axelson, John A.

    1949-01-01

    An additional series of high-speed wind-tunnel tests of a modified 0.17-scale model of the McDonnell XF2H-1 airplane was conducted to evaluate the effects of a reduction in the thickness-to-chord ratios of the tail planes, the displacement of the horizontal tail relative to the vertical tail, and the extension of the trailing edge of the wing. Two tail-intersection fairings designed to improve the flow at the tail were also tested. The pitching-moment characteristics of the model were improved slightly by the use of the thinner tail sections. Rearward or rearward and downward displacements of the horizontal tail increased the critical Mach number at the tail intersection from 0.725 to a maximum of 0.80, but caused an excessive change in pitching-moment coefficient at the higher Mach numbers. Extending the trailing edge of the wing did not improve the static longitudinal-stability characteristics, but increased the pitching-down tendency between 0.725 and 0.825 Mach numbers prior to the pitching-up tendency. The extended wing did, however, increase the Mach numbers at which these tendencies occurred. The increase in the Mach numbers of divergence and the tuft studies indicate a probable increase in the buffet limit of the prototype airplane. No perceptible improvement of flow at the tail intersection was observed with the two fairings tested on the forward tail configuration.

  13. Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Dillon, J. L.; Pittman, J. L.

    1977-01-01

    An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.

  14. 78 FR 53078 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-08-28

    ... and of the strut-to-wing attachment, and consequent separation of a strut and engine from the airplane...-to-wing attachment, which could result in consequent separation of a strut and engine from the... that section, Congress charges the FAA with promoting safe flight of civil aircraft in air commerce by...

  15. 78 FR 79599 - Airworthiness Directives; Various Aircraft Equipped With Wing Lift Struts

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-12-31

    ... 99-01-05 R1] RIN 2120-AA64 Airworthiness Directives; Various Aircraft Equipped With Wing Lift Struts... aircraft equipped with wing lift struts. The list of affected airplanes in the Applicability section is... wing lift struts for corrosion; repetitively inspecting the wing lift strut forks for cracks;...

  16. DOUGLAS XA3D-1 #413 AIRPLANE.

    NASA Image and Video Library

    1955-07-27

    DOUGLAS XA3D-1 #413 AIRPLANE MOUNTED IN THE NACA AMES RESEARCH CENTER'S 40X80_FOOT SUBSONIC WIND TUNNEL Testing the boundary layer control of the A3D in the 40 x 80 wind tunnel. Boundary layer control was added to increase the lift of the wing for take off from an aircraft carrier.

  17. Static longitudinal stability of "Ente" airplanes

    NASA Technical Reports Server (NTRS)

    Kiel, Heinrich Georg

    1931-01-01

    The stability conditions of Ente (duck) airplanes are investigated in this report. In developing the formulas, which afford an approximate solution, the unimportant effect of the height of the C.G. and the moment of the residual resistance are neglected. The effect of downwash from the forward horizontal empennage on the wing are also disregarded.

  18. Aerodynamic characteristics of a hypersonic research airplane concept having a 70 deg swept double-delta wing at Mach numbers from 0.80 to 1.20, with summary of data from 0.20 to 6.0. [Langley 8-ft transonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Penland, J. A.; Hallissy, J. B.; Dillon, J. L.

    1979-01-01

    The static longitudinal, lateral, and directional stability characteristics of a hypersonic research airplane concept having a 70 deg swept double-delta wing were investigated. Force tests were conducted in the Langley 8 foot transonic pressure tunnel for a Reynolds number (based on fuselage length) range of 6.30 x 10 to the 6th power to 7.03 x 10 to the 6th power, at angles of attack from about -4 deg to 23 deg, and at angles of sideslip of 0 deg and 5 deg. The configuration variables included the wing planform, tip fins, the center vertical tail, and scramjet engine modules. Variations of the more important aerodynamic parameters with Mach number for Mach numbers from 0.20 to 6.0 are summarized. A state-of-the-art example of theoretically predicting performance parameters and static longitudinal and directional stability over the Mach number range is included.

  19. Aerodynamics of Low Reynolds Number Rigid Flapping Wing Under Hover and Freestream Conditions

    NASA Astrophysics Data System (ADS)

    Trizila, Patrick Clark

    Micro air vehicles (MAVs) are defined by all spatial dimensions being less than 15 cm. Equipped with a video camera or a sensor, these vehicles could perform surveillance and reconnaissance with low rates of detection, or environmental and bio-chemical sensing at remote or otherwise hazardous locations. Its size makes the MAV easily transported and deployed as well as inexpensive and more expendable than alternatives, e.g. an airplane, a satellite, or a human. The ability to hover for an MAV is highly desirable in these contexts. The approach taken in the current studies is to numerically simulate the aerodynamics about flapping wings while controlling the kinematic motions and environmental conditions. Two complementary sets of tools were used in the investigations. Navier-Stokes solvers were used to obtain detailed fluid physics information, instantaneous force data, and to train the surrogate models. The surrogate models were used to estimate the average lift and power required over a flapping cycle while also providing information on the sensitivity of the kinematic variables, to identify trends in lift and power required as a function of the kinematic variables, and to construct a Pareto front showing the trade-offs between the competing objectives. Findings include i) an examination of the competing influences introduced by tip vortices, and it was seen that they could increase lift compared to their analogous 2D cases, counter to classical steady state theory. ii) The highest time averaged lift values were found during kinematics with high angles of attack during advanced rotation as they promoted LEV generation and subsequently took advantage of them during wake capture. iii) Kinematics with synchronized rotation and low angles of attack had surprisingly similar 2D and low-aspect-ratio force histories. iv) Modest environmental perturbations, those a fraction of the translational wing velocity, can have a profound impact on the resulting forces. Closely

  20. Wind-Tunnel Investigation of the Low-Speed Characteristics of a 1/8-Scale Model of the Republic XP-91 Airplane with a Vee and a Conventional Tail. Addendum - Characteristics with a Revised Conventional Tail and Drooped Wing Tips

    NASA Technical Reports Server (NTRS)

    Weiberg, James A.; Anderson, Warren E.

    1958-01-01

    Additional wind-tunnel tests were made of a 1/8-scale model of the Republic XP-91 airplane to determine its characteristics with various modifications. The modifications included a revised conventional tail, revised rocket arrangement, drooped wing tips, and revised landing gear and doors. Tests were also made to determine the effectiveness of the control surfaces of the model with the conventional tail and the effect of changing wing incidence and tail length. The revised rocket arrangement provided a considerable increase in the static directional stability contributed by the vee tail at small angles of yaw. The conventional tail provided a greater static directional stability than the vee tail without increasing the rolling moment due to sideslip. The rolling moment die to sideslip was considerable reduced by either drooped wing tips or open main landing-gear doors. The reduction in rolling moment due to sideslip resulting from the drooped tips was less with the landing-gear doors open than with the doors closed. A change in wing incidence from 0 degrees to 6 degrees reduced the elevator angle required for balance by approximately 6 degrees.

  1. 76 FR 27220 - Airworthiness Directives; Airbus Model A310 Series Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-05-11

    ... structural integrity of the wings. We are issuing this AD to require actions to correct the unsafe condition... (SBs which address structure fatigue related areas on the wing parts), until now part of the... wings. The required corrective actions are as follows, depending on airplane configuration: For...

  2. 77 FR 54800 - Airworthiness Directives; Diamond Aircraft Industries GmbH Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-09-06

    ... Aircraft Industries GmbH Models DA 42, DA 42 NG, and DA 42 M-NG airplanes. This AD results from mandatory... excessive voids in the adhesive joint between the center wing spars and the ] upper center wing skins. This... DA 42 NG, voids were detected in the adhesive joint between the centre wing spars and the...

  3. Celebrating 100 Years of Flight: Testing Wing Designs in Aircraft

    ERIC Educational Resources Information Center

    Pugalee, David K.; Nusinov, Chuck; Giersch, Chris; Royster, David; Pinelli, Thomas E.

    2005-01-01

    This article describes an investigation involving several designs of airplane wings in trial flight simulations based on a NASA CONNECT program. Students' experiences with data collection and interpretation are highlighted. (Contains 5 figures.)

  4. Celebrating 100 Years of Flight: Testing Wing Designs in Aircraft

    ERIC Educational Resources Information Center

    Pugalee, David K.; Nusinov, Chuck; Giersch, Chris; Royster, David; Pinelli, Thomas E.

    2005-01-01

    This article describes an investigation involving several designs of airplane wings in trial flight simulations based on a NASA CONNECT program. Students' experiences with data collection and interpretation are highlighted. (Contains 5 figures.)

  5. Effect of Various Modifications on Drag and Longitudinal Stability and Control Characteristics at Transonic Speeds of a Model of the XF7U-1 Tailless Airplane: NACA Wing-FLow Method, TED No. NACA DE 307

    NASA Technical Reports Server (NTRS)

    Sawyer, Richard H.; Trant, James P., Jr.

    1950-01-01

    An investigation was made by the NACA wing-flow method to determine the drag, pitching-moment, lift, and angle-of-attack characteristics at transonic speeds of various configurations of a semispan model of an early configuration of the XF7U-1 tailless airplane. The results of the tests indicated that for the basic configuration with undeflected ailavator, the zero-lift drag rise occurred at a Mach number of about 0.85 and that about a five-fold increase in drag occurred through the transonic speed range. The results of the tests also indicated that the drag increment produced by -8.0 degrees deflection of the ailavator increased with increase in normal-force coefficient and was smaller at speeds above than at speeds below the drag rise. The drag increment produced by 35 degree deflection of the speed brakes varied from 0.040 to 0.074 depending on the normal-force coefficient and Mach number. These values correspond to drag coefficients of about 0.40 and 0.75 based on speed-brake frontal area. Removal of the fin produced a small positive drag increment at a given normal-force coefficient at speeds during the drag rise. A large forward shift of the neutral-point location occurred at Mach numbers above about 0.90 upon removal of the fin, and also a considerable forward shift throughout the Mach number range occurred upon deflection of the speed brakes. Ailavator ineffectiveness or reversal at low deflections, similar to that determined in previous tests of the basic configuration of the model in the Mach number range from about 0.93 to 1.0, was found for the fin-off configuration and for the model equipped with skewed (more highly sweptback) hinge-line ailavators. With the speed brakes deflected, little or no loss in the incremental pitching moment produced by deflection of the ailavator from O degrees to -8.00 degrees occurred in the Mach number range from 0.85 to 1.0 in contrast to a considerable loss found in previous tests with the speed brakes off.

  6. Application of winglets and/or wing tip extensions with active load control on the Boeing 747

    NASA Technical Reports Server (NTRS)

    Allison, R. L.; Perkin, B. R.; Schoenman, R. L.

    1978-01-01

    The application of wing tip modifications and active control technology to the Boeing 747 airplane for the purpose of improving fuel efficiency is considered. Wing tip extensions, wing tip winglets, and the use of the outboard ailerons for active wing load alleviation are described. Modest performance improvements are indicated. A costs versus benefits approach is taken to decide which, if any, of the concepts warrant further development and flight test leading to possible incorporation into production airplanes.

  7. Simulation of Dynamics of a Flexible Miniature Airplane

    NASA Technical Reports Server (NTRS)

    Waszak, Martin R.

    2005-01-01

    A short report discusses selected aspects of the development of the University of Florida micro-aerial vehicle (UFMAV) basically, a miniature airplane that has a flexible wing and is representative of a new class of airplanes that would operate autonomously or under remote control and be used for surveillance and/or scientific observation. The flexibility of the wing is to be optimized such that passive deformation of the wing in the presence of aerodynamic disturbances would reduce the overall response of the airplane to disturbances, thereby rendering the airplane more stable as an observation platform. The aspect of the development emphasized in the report is that of computational simulation of dynamics of the UFMAV in flight, for the purpose of generating mathematical models for use in designing control systems for the airplane. The simulations are performed by use of data from a wind-tunnel test of the airplane in combination with commercial software, in which are codified a standard set of equations of motion of an airplane, and a set of mathematical routines to compute trim conditions and extract linear state space models.

  8. Precision controllability of the YF-17 airplane

    NASA Technical Reports Server (NTRS)

    Sisk, T. R.; Mataeny, N. W.

    1980-01-01

    A flying qualities evaluation conducted on the YF-17 airplane permitted assessment of its precision controllability in the transonic flight regime over the allowable angle of attack range. The precision controllability (tailchase tracking) study was conducted in constant-g and windup turn tracking maneuvers with the command augmentation system (CAS) on, automatic maneuver flaps, and the caged pipper gunsight depressed 70 mils. This study showed that the YF-17 airplane tracks essentially as well at 7 g's to 8 g's as earlier fighters did at 4 g's to 5 g's before they encountered wing rock. The pilots considered the YF-17 airplane one of the best tracking airplanes they had flown. Wing rock at the higher angles of attack degraded tracking precision, and lack of control harmony made precision controllability more difficult. The revised automatic maneuver flap schedule incorporated in the airplane at the time of the tests did not appear to be optimum. The largest tracking errors and greatest pilot workload occurred at high normal load factors at low angles of attack. The pilots reported that the high-g maneuvers caused some tunnel vision and that they found it difficult to think clearly after repeated maneuvers.

  9. Spinning Characteristics of Wings I : Rectangular Clark Y Monoplane Wing

    NASA Technical Reports Server (NTRS)

    Bamber, M J; Zimmerman, C H

    1936-01-01

    A series of wind tunnel tests of a rectangular Clark Y wing was made with the NACA spinning balance as part of a general program of research on airplane spinning. All six components of the aerodynamic force and moment were measured throughout the range of angles of attack, angles of sideslip, and values omega b/2v likely to be attained by a spinning airplane; the results were reduced to coefficient form. It is concluded that a conventional monoplane with a rectangular Clark y wing can be made to attain spinning equilibrium throughout a wide range of angles of attack but that provision of a yawing moment coefficient of -0.02 (against the spin) by the tail, fuselage, and interferences will insure against attainment of equilibrium in a steady spin.

  10. Features of owl wings that promote silent flight.

    PubMed

    Wagner, Hermann; Weger, Matthias; Klaas, Michael; Schröder, Wolfgang

    2017-02-06

    Owls are an order of birds of prey that are known for the development of a silent flight. We review here the morphological adaptations of owls leading to silent flight and discuss also aerodynamic properties of owl wings. We start with early observations (until 2005), and then turn to recent advances. The large wings of these birds, resulting in low wing loading and a low aspect ratio, contribute to noise reduction by allowing slow flight. The serrations on the leading edge of the wing and the velvet-like surface have an effect on noise reduction and also lead to an improvement of aerodynamic performance. The fringes at the inner feather vanes reduce noise by gliding into the grooves at the lower wing surface that are formed by barb shafts. The fringed trailing edge of the wing has been shown to reduce trailing edge noise. These adaptations to silent flight have been an inspiration for biologists and engineers for the development of devices with reduced noise production. Today several biomimetic applications such as a serrated pantograph or a fringed ventilator are available. Finally, we discuss unresolved questions and possible future directions.

  11. 14 CFR 25.1403 - Wing icing detection lights.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Wing icing detection lights. 25.1403... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Equipment Lights § 25.1403 Wing icing detection lights. Unless operations at night in known or forecast icing conditions are prohibited by an...

  12. 14 CFR 25.1403 - Wing icing detection lights.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Wing icing detection lights. 25.1403... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Equipment Lights § 25.1403 Wing icing detection lights. Unless operations at night in known or forecast icing conditions are prohibited by an...

  13. 14 CFR 25.1403 - Wing icing detection lights.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Wing icing detection lights. 25.1403... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Equipment Lights § 25.1403 Wing icing detection lights. Unless operations at night in known or forecast icing conditions are prohibited by...

  14. 14 CFR 25.1403 - Wing icing detection lights.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Wing icing detection lights. 25.1403... AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Equipment Lights § 25.1403 Wing icing detection lights. Unless operations at night in known or forecast icing conditions are prohibited by...

  15. 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.

  16. 78 FR 65233 - Special Conditions: Bombardier Inc., Models BD-500-1A10 and BD-500-1A11 Series Airplanes...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-10-31

    ... Safety Branch, ANM-115 Transport Airplane Directorate, Aircraft Certification Service, 1601 Lind Avenue... to as ``C-series''). The C- series airplanes are swept-wing monoplanes with a pressurized cabin. They... BD- 500-1A11 Series Airplanes; Fuselage Post-Crash Fire Survivability AGENCY: Federal...

  17. 78 FR 75287 - Special Conditions: Bombardier Inc., Models BD-500-1A10 and BD-500-1A11 Series Airplanes; Flight...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-12-11

    ..., ANM-111, Transport Airplane Directorate, Aircraft Certification Service, 1601 Lind Avenue SW., Renton...-series''). The C- series airplanes are swept-wing monoplanes with a pressurized cabin. They share an... BD- 500-1A11 Series Airplanes; Flight Envelope Protection: General Limiting Requirements...

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 75 FR 61999 - Airworthiness Directives; The Boeing Company Model 767 Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-10-07

    ... airplanes. This proposed AD results from reports of fuel leaks from certain drain locations of the slat... reports of fuel leaks from certain drain locations of the slat track housing near the engine exhaust... near the engine exhaust nozzles of the wings, which could result in a fire when the airplane is...

  5. 76 FR 62663 - Airworthiness Directives; The Boeing Company Model 767 Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-10-11

    ... structure. Continental believed the intent is to cold work the skin hole only for airplanes with titanium... working was meant for the wing skin holes for airplanes having titanium pitch load fittings. However, we have determined that the titanium fitting maintains an adequate level of safety if the cold...

  6. 75 FR 15357 - Airworthiness Directives; The Boeing Company Model 767 Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-03-29

    ...-to-wing upper link load path and possible separation of a strut and engine from the airplane during... possible separation of a strut and engine from the airplane during flight. Relevant Service Information We... flight of civil aircraft in air commerce by prescribing regulations for practices, methods, and...

  7. Membrane and adaptively-shaped wings for micro air vehicles

    NASA Astrophysics Data System (ADS)

    Lian, Yongsheng

    Micro air vehicles (MAVs), with wing span of 15 cm or less and flight speed around 10 m/s, have many applications in both civilian and military areas. The Reynolds number based on the given parameters is around 10 4, which often yields insufficient lift-to-drag ratio. Furthermore, one expects the unsteady effect to be noticeable for such flight vehicles. The flexible wing has been demonstrated to exhibit favorable characteristics such as passive adaptation to the flight; environment and delayed stall. The present study focuses on developing computational and modeling capabilities to better understand the MAV aerodynamics. Both flexible wings, utilizing membrane materials, and adaptively-shaped wings, utilizing piezo-actuated flaps, have been studied. In the adaptively-shaped wing study, we use piezo-actuated flaps to actively control the flow. We assess the impacts of the flap geometry, flapping amplitude, and turbulence; modeling on the flow structure with a parallel experimental effort. The membrane wing uses a passive control mechanism to delay the stall angle and to provide a smoother flight platform. Our study focuses on the mutual interactions between the membrane wing and its surrounding viscous flow. We compare the lift-to-drag ratio and the flow structure between the flexible wing and the corresponding rigid wing. We also investigate the aerodynamic characteristics associated with the low Reynolds number and low aspect ratio wing. To assist our study, we propose an automatic and efficient moving grid technique to facilitate the fluid and structure interaction computations; we also present a dynamic membrane model to study the intrinsic large deformation of the flexible membrane wing. Solutions obtained from the three-dimensional Navier-Stokes equations are presented to highlight, the salient features of the wing aerodynamics. Besides the aerodynamic study, we also perform shape optimization to improve the membrane wing performance. Since direct

  8. Airplane Balance

    NASA Technical Reports Server (NTRS)

    Huguet, L

    1921-01-01

    The authors argue that the center of gravity has a preponderating influence on the longitudinal stability of an airplane in flight, but that manufacturers, although aware of this influence, are still content to apply empirical rules to the balancing of their airplanes instead of conducting wind tunnel tests. The author examines the following points: 1) longitudinal stability, in flight, of a glider with coinciding centers; 2) the influence exercised on the stability of flight by the position of the axis of thrust with respect to the center of gravity and the whole of the glider; 3) the stability on the ground before taking off, and the influence of the position of the landing gear. 4) the influence of the elements of the glider on the balance, the possibility of sometimes correcting defective balance, and the valuable information given on this point by wind tunnel tests; 5) and a brief examination of the equilibrium of power in horizontal flight, where the conditions of stability peculiar to this kind of flight are added to previously existing conditions of the stability of the glider, and interfere in fixing the safety limits of certain evolutions.

  9. Assembly modes of dragonfly wings.

    PubMed

    Zhao, Hong-Xiao; Yin, Ya-Jun; Zhong, Zheng

    2011-12-01

    The assembly modes of dragonfly wings are observed through FEG-ESEM. Different from airplane wings, dragonfly wings are found to be assembled through smooth transition mode and global package mode. First, at the vein/membrane conjunctive site, the membrane is divided into upper and lower portions from the center layer and transited smoothly to the vein. Then the two portions pack the vein around and form the outer surface of the vein. Second, at the vein/spike conjunctive site, the vein and spike are connected smoothly into a triplet. Last, at the vein/membrane/spike conjunctive site, the membrane (i.e., the outer layer of the vein) transits smoothly to the spike, packs it around, and forms its outer layer. In short, the membrane looks like a closed coat packing the wing as a whole. The smooth transition mode and the global package mode are universal assembly modes in dragonfly wings. They provide us the references for better understanding of the functions of dragonfly wings and the bionic manufactures of the wings of flights with mini sizes.

  10. Supersonic airplane study and design

    NASA Technical Reports Server (NTRS)

    Cheung, Samson

    1993-01-01

    A supersonic airplane creates shocks which coalesce and form a classical N-wave on the ground, forming a double bang noise termed sonic boom. A recent supersonic commercial transport (the Concorde) has a loud sonic boom (over 100 PLdB) and low aerodynamic performance (cruise lift-drag ratio 7). To enhance the U.S. market share in supersonic transport, an airframer's market risk for a low-boom airplane has to be reduced. Computational fluid dynamics (CFD) is used to design airplanes to meet the dual constraints of low sonic boom and high aerodynamic performance. During the past year, a research effort was focused on three main topics. The first was to use the existing design tools, developed in past years, to design one of the low-boom wind-tunnel configurations (Ames Model 3) for testing at Ames Research Center in April 1993. The second was to use a Navier-Stokes code (Overflow) to support the Oblique-All-Wing (OAW) study at Ames. The third was to study an optimization technique applied on a Haack-Adams body to reduce aerodynamic drag.

  11. 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.

  12. Optico-photographic measurements of airplane deformations

    NASA Technical Reports Server (NTRS)

    Kussner, Hans Georg

    1931-01-01

    The deformation of aircraft wings is measured by photographically recording a series of bright shots on a moving paper band sensitive to light. Alternating deformations, especially vibrations, can thus be measured in operation, unaffected by inertia. A handy recording camera, the optograph, was developed by the static division of the D.V.L. (German Experimental Institute for Aeronautics) for the employment of this method of measurement on airplanes in flight.

  13. Spinning Characteristics of Wings III : a Rectangular and Tapered Clark Y Monoplane Wing with Rounded Tips

    NASA Technical Reports Server (NTRS)

    Bamber, M J; House, R O

    1937-01-01

    An investigation was made to determine the spinning characteristics of Clark Y monoplane wings with different plan forms. A rectangular wing and a wing tapered 5:2, both with rounded tips, were tested on the N.A.C.A. spinning balance in the 5-foot vertical wind tunnel. The aerodynamic characteristics of the models and a prediction of the angles of sideslip for steady spins are given. Also included is an estimate of the yawning moment that must be furnished by the parts of the airplane to balance the inertia couples and wing yawing moment for spinning equilibrium. The effects on the spin of changes in plan form and of variations of some of the important parameters are discussed and the results are compared with those for a rectangular wing with square tips. It is concluded that for a conventional monoplane using Clark Y wing the sideslip will be algebraically larger for the wing with the rounded tip than for the wing with the square tip and will be largest for the tapered wing. The effect of plan form on the spin will vary with the type of airplane; and the provision of a yawing-moment coefficient of -0.025 (i.e., opposing the spin) by the tail, fuselage, and interference effects will insure against the attainment of equilibrium on a steady spin for any of the plan forms tested and for any of the parameters used in the analysis.

  14. Supersonic aeroelastic instability results for a NASP-like wing model

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.

    1993-01-01

    Two wing-alone wind-tunnel models were tested in the NASA Langley Unitary Plan Wind Tunnel facility to study the static divergence behavior of such configurations and to provide a data base for correlation with supersonic analytical predictions. One model had a four percent maximum thickness airfoil and the other had an eight-percent maximum thickness airfoil. The wing models had low aspect ratios and highly swept leading edges. Results show that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneficial effect of increasing the divergence dynamic pressures. The calculations accurately predicted the trend of experimental divergence dynamic pressure with Mach number.

  15. Stability and control characteristics of an airplane model having a 45.1 degree swept-back wing with aspect ratio 2.50 and taper ratio 0.42 and a 42.8 degree swept-back horizontal tail with aspect ratio 3.87 and taper ratio 0.49

    NASA Technical Reports Server (NTRS)

    Schuldenfrei, Marvin; Comisarow, Paul; Goodson, Kenneth W

    1947-01-01

    Tests were made of an airplane model having a 45.1 degree swept-back wing with aspect ratio 2.50 and taper ratio 0.42 and a 42.8 degree swept-back horizontal tail with aspect ratio 3.87 and taper ratio 0.49 to determine its low-speed stability and control characteristics. The test Reynolds number was 2.87 x 10(6) based on a mean aerodynamic chord of 2.47 feet except for some of the aileron tests which were made at a Reynolds number of 2.05 x 10(6). With the horizontal tail located near the fuselage juncture on the vertical tail, model results indicated static longitudinal instability above a lift coefficient that was 0.15 below the lift coefficient at which stall occurred. Static longitudinal stability, however, was manifested throughout the life range with the horizontal tail located near the top of the vertical tail. The use of 10 degrees negative dihedral on the wing had little effect on the static longitudinal stability characteristics. Preliminary tests of the complete model revealed an undesirable flat spot in the yawing-moment curves at low angles of attack, the directional stability being neutral for yaw angles of plus-or-minus 2 degrees. This undesirable characteristic was improved by replacing the thick original vertical tail with a thin vertical tail and by flattening the top of the dorsal fairing.

  16. Aerodynamic Loads at Mach Numbers from 0.70 to 2.22 on an Airplane Model Having a Wing and Canard of Triangular Plan Form and Either Single or Twin Vertical Tails. Supplement 2; Tabulated Data for the Model with Twin Vertical Tails

    NASA Technical Reports Server (NTRS)

    Peterson, Victor L.; Menees, Gene P.

    1961-01-01

    Tabulated results of a wind-tunnel investigation of the aerodynamic loads on a canard airplane model with twin vertical tails are presented for Mach numbers from 0.70 to 2.22. The Reynolds number for the measurements was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. The results include local static-pressure coefficients measured on the wing, body, and one of the vertical tails for angles of attack from -4 degrees to 16 degree angles of sideslip of 0 degrees and 5.3 degrees, and nominal canard deflections of O degrees and 10 degrees. Also included are section force and moment coefficients obtained from integrations of the local pressures and model-component force and moment coefficients obtained from integrations of the section coefficients. Geometric details of the model are shown and the locations of the pressure orifices are shown. An index to the data contained herein is presented and definitions of nomenclature are given. Detailed descriptions of the model and experiments and a brief discussion of some of the results are given. Tabulated results of measurements of the aerodynamic loads on the same canard model but having a single vertical tail instead of twin vertical tails are presented.

  17. Aerodynamic Loads at Mach Numbers from 0.70 to 2.22 on an Airplane Model Having a Wing and Canard of Triangular Plan Form and Either Single or Twin Vertical Tails Supplement I-Tabulated Data for the Model with Single Vertical Tails. Supplement 1; Tabulated Data for the Model with Single Vertical Tail

    NASA Technical Reports Server (NTRS)

    Peterson, Victor L.; Menees, Gene P.

    1961-01-01

    Tabulated results of a wind-tunnel investigation of the aerodynamic loads on a canard airplane model with a single vertical tail are presented for Mach numbers from 0.70 to 2.22. The Reynolds number for the measurements was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. The results include local static pressure coefficients measured on the wing, body, and vertical tail for angles of attack from -4 deg to + 16 deg, angles of sideslip of 0 deg and 5.3 deg, vertical-tail settings of 0 deg and 5 deg, and nominal canard deflections of 0 deg and 10 deg. Also included are section force and moment coefficients obtained from integrations of the local pressures and model-component force and moment coefficients obtained from integrations of the section coefficients. Geometric details of the model and the locations of the pressure orifices are shown. An index to the data contained herein is presented and definitions of nomenclature are given.

  18. Axial flow effects on robustness of vortical structures about actively deflected wings in flapping flight

    NASA Astrophysics Data System (ADS)

    Medina, Albert; Kweon, Jihoon; Choi, Haecheon; Eldredge, Jeff D.

    2012-11-01

    Flapping wing flight has garnered much attention in the past decade driven by our desire to understand capabilities observed in nature and to develop agile small-scale aerial vehicles. Nature has demonstrated the breadth of maneuverability achievable by flapping wing flight. However, despite recent advances the role of wing flexibility remains poorly understood. In an effort to develop a deeper understanding of wing deflection effects and to explore novel approaches to increasing leading-edge vortex robustness, this three-dimensional computational study explores the aerodynamics of low aspect ratio plates, in hovering kinematics, with isolated flexion lines undergoing prescribed deflection. Major flexion lines, recognized as the primary avenue for deflection in biological fliers, are isolated here in two distinct configurations, resulting in deflection about the wing root and the wing tip, respectively. Of interest is the interaction between axial flow along the span and the vortical structures about the wing. It is proposed that the modes of deflection explored may provide a means of axial flow control for favorably promoting LEV robustness over a broad range of flapping conditions, and provide insight into the nature of flexibility in flapping wing flight. National Science Foundation, National Research Foundation of Korea.

  19. Performance of two load-limiting subfloor concepts in full-scale general aviation airplane crash tests

    NASA Technical Reports Server (NTRS)

    Carden, H. D.

    1984-01-01

    Three six-place, low wing, twin-engine general aviation airplane test specimens were crash tested at the langley Impact Dynamics research Facility under controlled free-flight conditions. One structurally unmodified airplane was the baseline airplane specimen for the test series. The other airplanes were structurally modified to incorporate load-limiting (energy-absorbing) subfloor concepts into the structure for full scale crash test evaluation and comparison to the unmodified airplane test results. Typically, the lowest floor accelerations and anthropomorphic dummy occupant responses, and the least seat crushing of standard and load-limiting seats, occurred in the modified load-limiting subfloor airplanes wherein the greatest structural crushing of the subfloor took place. The better performing of the two load-limiting subfloor concepts reduced the peak airplane floor accelerations at the pilot and four seat/occupant locations to -25 to -30 g's as compared to approximately -50 to -55 g's acceleration magnitude for the unmodified airplane structure.

  20. 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.

  1. Inflatable Wing project personnel prepare a deployable, inflatable wing technology demonstrator expe

    NASA Technical Reports Server (NTRS)

    2001-01-01

    Inflatable Wing project personnel prepare a deployable, inflatable wing technology demonstrator experiment flown 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.

  2. 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.

  3. 78 FR 64417 - Airworthiness Directives; Twin Commander Aircraft LLC Airplanes; Initial Regulatory Flexibility...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-10-29

    ... Commander Aircraft LLC Models 690, 690A, and 690B airplanes. AD 2013-09-05 requires inspection for cracking... window channels, aft cabin pressure web, external wing to fuselage fillets, and fasteners; repair or..., the vertical channels, the upper picture window channels, aft cabin pressure web, external wing...

  4. 77 FR 35890 - Airworthiness Directives; Diamond Aircraft Industries GmbH Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-06-15

    ... (AD) for certain Diamond Aircraft Industries GmbH Models DA 42, DA 42 NG, and DA 42 M-NG airplanes.... The MCAI describes the unsafe condition as excessive voids in the adhesive joint between the center... detected in the adhesive joint between the centre wing spars and the upper centre wing skins, between...

  5. Stability of airplanes

    NASA Technical Reports Server (NTRS)

    Warner, Edward P

    1922-01-01

    The author attempts to correct the misconception that piloting an airplane requires extraordinary skill and balance. He also tries to show that airplanes are extremely stable in flight. Some of the major points covered in this article include: automatic pilots, airplanes designed to be stable, and the reliance on mathematics to help in designing stable aircraft.

  6. Design of static reaction gantry for an ultralight airplane destruction test

    NASA Technical Reports Server (NTRS)

    Smith, H. W.

    1985-01-01

    The steel gantry superstructure needed to perform an airplane static test is described. Standard civil engineering design practices are used to react the loads generated by an airplane in flight. Reaction columns are mounted on a structural floor to carry the wing airloads and the downward acting fuselage loads are carried directly into the floor. The gantry can accommodate a general aviation airplane or rotorcraft. An immediate use for an ultralight airplane is shown as an example configuration of the four main steel frames.

  7. Presentation of structural component designs for the family of commuter airplanes

    NASA Technical Reports Server (NTRS)

    Russell, Mark; Haddad, Raphael; Creighton, Tom; Hendrich, Louis; Hensley, Doug; Morgan, Louise; Swift, Jerry

    1987-01-01

    The purpose is to present the implementation of structural commonality in the family of commuter airplanes. One of the main goals is implementation of structural commonality to as high a degree as possible. The structural layouts of those parts of the airplanes in which commonality is possible with all members of the family will be presented. The following airplane sections, which are common on all of the airplanes in the family, will be presented: common nose cone design; common wing torque box design; and common tail cone design. A proposed production and manufacturing breakdown is described. The advantages and disadvantages of implementing structural commonality and recommendations for further work will be discussed.

  8. 77 FR 42225 - Airworthiness Directives; PILATUS AIRCRAFT LTD. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-07-18

    ... specified in the Aircraft Maintenance Manual (AMM) under Chapter 4. Prompted by a crack found on one wing of... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; PILATUS AIRCRAFT... (AD) for all PILATUS AIRCRAFT LTD. Models PC 12, PC 12/45, PC 12/47, and PC 12/47E airplanes that...

  9. 78 FR 52875 - Airworthiness Directives; the Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-08-27

    ... airplane reaching its limit of validity (LOV) of the engineering data that support the established... the wing structure to support the limit load condition, and consequent loss of structural integrity of...: Data & Services Management, P.O. Box 3707, MC 2H-65, Seattle, WA 98124-2207; telephone 206-544- 5000...

  10. Nieuport-Delage Pursuit Airplane 48 C. 1. : "jockey" type

    NASA Technical Reports Server (NTRS)

    1927-01-01

    This is a light single-seat pursuit airplane with a tractor propeller actuated by a 12 cylinder V-type Hispano-Suiza engine giving 400 HP at 2000 R.P.M. This is a single winged aircraft capable of 273 km/h.

  11. AGATE concept model for a future general aviation airplane

    NASA Technical Reports Server (NTRS)

    1996-01-01

    Photo illustrates an advanced general aviation concept airplane. The pusher propeller - driven configuration seats 4 to 6 people (including pilot) in mid-wing, three-surface twin tail-boom configuration. The design concept incorporates natural laminar flow, ice protection, winglets and stall-departure-resistant flight dynamics.

  12. The airplane: A simulated commercial air transportation study

    NASA Technical Reports Server (NTRS)

    Dauteuil, Mark; Geniesse, Pete; Hunniford, Michael; Lawler, Kathleen; Quirk, Elena; Tognarelli, Michael

    1993-01-01

    The 'Airplane' is a moderate-range, 70 passenger aircraft. It is designed to serve demands for flights up to 10,000 feet and it cruises at 32 ft/s. The major drivers for the design of the Airplane are economic competitiveness, takeoff performance, and weight minimization. The Airplane is propelled by a single Astro 15 electric motor and a Zinger 12-8 propeller. The wing section is a Spica airfoil which, because of its flat bottom, provides simplicity in manufacturing and thus helps to cut costs. The wing is constructed of a single load bearing mainspar and shape-holding ribs coated with Monokote skin, lending to a light weight structural makeup. The fuselage houses the motor, flight deck and passenger compartments as well as the fuel and control actuating systems. The wing will be attached to the top of the fuselage as will the fuel and control actuator systems for easy disassembly and maintenance. The aircraft is maneuvered about its pitch axis by means of an aft elevator on the flat plate horizontal tail. The twin vertical tail surfaces are also flat plates and each features a rudder for both directional and roll control. Along with wing dihedral, the rudders will be used to roll the aircraft. The Airplane is less costly to operate at its own maximum range and capacity as well as at its maximum range and the HB-40's maximum capacity than the HB-40.

  13. 78 FR 39633 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-07-02

    ... physical damage to the trailing edge wedges on the leading edge wing slats, and repair if necessary. On.... This proposed AD would require a determination of the type of trailing edge wedges of the leading edge... airplanes. We are proposing this AD to prevent delamination of the trailing edge wedge of the leading edge...

  14. 78 FR 49221 - Airworthiness Directives; Piper Aircraft, Inc. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-08-13

    ... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; Piper Aircraft, Inc... Aircraft, Inc. Model J-2 airplanes equipped with wing lift struts. AD 99-26-19 currently requires... Aircraft, Inc., Customer Services, 2926 Piper Drive, Vero Beach, Florida 32960; telephone: (772) 567-4361...

  15. Preliminary design characteristics of a subsonic business jet concept employing an aspect ratio 25 strut braced wing

    NASA Technical Reports Server (NTRS)

    Turriziani, R. V.; Lovell, W. A.; Martin, G. L.; Price, J. E.; Swanson, E. E.; Washburn, G. F.

    1980-01-01

    The advantages of replacing the conventional wing on a transatlantic business jet with a larger, strut braced wing of aspect ratio 25 were evaluated. The lifting struts reduce both the induced drag and structural weight of the heavier, high aspect ratio wing. Compared to the conventional airplane, the strut braced wing design offers significantly higher lift to drag ratios achieved at higher lift coefficients and, consequently, a combination of lower speeds and higher altitudes. The strut braced wing airplane provides fuel savings with an attendant increase in construction costs.

  16. Development of an aerodyanmic theory capable of predicting surface loads on slender wings with vortex flow

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.; Johnson, F. T.

    1976-01-01

    The Boeing Commercial Airplane Company developed an inviscid three-dimensional lifting surface method that shows promise in being able to accurately predict loads, subsonic and supersonic, on wings with leading-edge separation and reattachment.

  17. Survey of research on unsteady aerodynamic loading of delta wings

    NASA Technical Reports Server (NTRS)

    Ashley, H.; Vaneck, T.; Katz, J.; Jarrah, M. A.

    1991-01-01

    For aeronautical applications, there has been recent interest in accurately determining the aerodynamic forces and moments experienced by low-aspect-ratio wings performing transient maneuvers which go to angles of attack as high as 90 deg. Focusing on the delta planform with sharp leading edges, the paper surveys experimental and theoretical investigations dealing with the associated unsteady flow phenomena. For maximum angles above a value between 30 and 40 deg, flow details and airloads are dominated by hysteresis in the 'bursting' instability of intense vortices which emanate from the leading edge. As examples of relevant test results, force and moment histories are presented for a model series with aspect ratios 1, 1.5 and 2. Influences of key parameters are discussed, notably those which measure unsteadiness. Comparisons are given with two theories: a paneling approximation that cannot capture bursting but clarifies other unsteady influences, and a simplified estimation scheme which uses measured bursting data.

  18. Development of aerodynamic prediction methods for irregular planform wings

    NASA Technical Reports Server (NTRS)

    Benepe, D. B., Sr.

    1983-01-01

    A set of empirical methods was developed to predict low-speed lift, drag and pitching-moment variations with angle of attack for a class of low aspect ratio irregular planform wings suitable for application to advanced aerospace vehicles. The data base, an extensive series of wind-tunnel tests accomplished by the Langley Research Center of the National Aeronautics and Space Administration, is summarized. The approaches used to analyze the wind tunnel data, the evaluation of previously existing methods, data correlation efforts, and the development of the selected methods are presented and discussed. A summary of the methods is also presented to document the equations, computational charts and design guides which have been programmed for digital computer solution. Comparisons of predictions and test data are presented which show that the new methods provide a significant improvement in capability for evaluating the landing characteristics of advanced aerospace vehicles during the preliminary design phase of the configuration development cycle.

  19. Stability derivatives of triangular wings at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Ribner, Herbert S; Malvestuto, Frank S , Jr

    1948-01-01

    The analysis of the stability derivatives of low-aspect-ratio triangular wings at subsonic and supersonic speeds, given in NACA TN no. 1423, is extended to apply to triangular wings having large vertex angles and traveling at supersonic speeds. The lift, rolling moment due to sideslip, and damping in roll and pitch for this more general case have been treated elsewhere on the basis of the theory of small disturbances. The surface potentials for angle of attack and rolling taken therefrom are used to obtain the several side-force and yawing-moment derivatives that depend on leading-edge suction, and a tentative value for the rolling moment due to yawing. The lift and moment due to downward acceleration are obtained on the basis of an unpublished unsteady-flow solution. All the known stability derivatives of the triangular wing at supersonic speeds, regardless of source, are summarized for convenience and presented with respect to both body axes and stability axes. The results are limited to Mach numbers for which the triangular wing is contained within the Mach cone for its vertex. The spanwise variation of Mach number in the case of yawing is neglected, although the effect must be of importance.

  20. Learjet Model 55 Wing Analysis with Landing Loads

    NASA Technical Reports Server (NTRS)

    Boroughs, R. R.

    1985-01-01

    The NASTRAN analysis was used to determine the impact of new landing loads on the Learjet Model 55 wing. These new landing loads were the result of a performance improvement effort to increase the landing weight of the aircraft to 18,000 lbs. from 17,000 lbs. and extend the life of the tires and brakes by incorporating larger tires and heavy duty brakes. Landing loads for the original 17,000 lb. airplane landing configuration were applied to the full airplane NASTRAN model. The analytical results were correlated with the strain gage data from the original landing load static tests. The landing loads for the 18,000 lb. airplane were applied to the full airplane NASTRAN model, and a comparison was made with the original Model 55 data. The results of this comparison enable Learjet to determine the difference in stress distribution in the wing due to these two different sets of landing loads.

  1. A review of current technical knowledge necessary to develop large scale wing-in-surface effect craft

    NASA Astrophysics Data System (ADS)

    Hooker, Stephan F.

    A comprehensive review is made of the conceptual development to date of large 'wing-in-ground' (WIG) aircraft suitable for large-payload/long range oceanic transport irrespective of the sea-states thus encountered. These WIG 'wingship' craft would be of a scale comparable to current cruise ships, while being capable of speeds fully one-half as great as those of existing airliners. Attention is given to the consequences of incorporating power-augmented ram propulsion schemes into these low aspect-ratio wingship configurations, as well as to extant methods for the prediction of the flight dynamics and efficiencies of such aircraft.

  2. Use of a hyperbolic grid generation scheme in simulating supersonic viscous flow about three-dimensional winged configuration

    NASA Technical Reports Server (NTRS)

    Rizk, Y. M.; Steger, J. L.; Chaussee, D. S.

    1985-01-01

    The present paper describes a numerical mesh generation technique to be used with an implicit finite difference method for simulating visous supersonic flow about low-aspect-ratio wing body configurations using a single grid strategy. The computational domain is segmented into multiple regions, with borders located in supersonic areas to avoid the otherwise costly interfacing procedure between adjacent segments. The numerical procedure is applied to calculate the turbulent flow around the shuttle orbiter and a canard projectile at supersonic free stream Mach number.

  3. Results of a Cyclic Load Test of an RB-47E Airplane

    NASA Technical Reports Server (NTRS)

    Huston, Wilber B.

    1959-01-01

    Results of a cyclic load test made by NASA on an EB-47E airplane are given. The test reported on is for one of three B-47 airplanes in a test program set up by the U. S. Air Force to evaluate the effect of wing structural reinforcements on fatigue life. As a result of crack development in the upper fuselage longerons of the other two airplanes in the program, a longeron and fuselage skin modification was incorporated early in the test. Fuselage strain-gage measurements made before and after the longeron modification and wing strain-gage measurements made only after wing reinforcement are summarized. The history of crack development and repair is given in detail. Testing was terminated one sequence short of the planned end of the program with the occurrence of a major crack in the lower right wing skin.

  4. AIRCL: A programmed system for generating NC tapes for airplane models. User's manual

    NASA Technical Reports Server (NTRS)

    Akgerman, N.; Billhardt, C. F.

    1981-01-01

    A computer program is presented which calculates the cutter location file needed to machine models of airplane wings or wing-fuselage combinations on numerically controlled machine tools. Input to the program is a data file consisting of coordinates on the fuselage and wing. From this data file, the program calculates tool offsets, determines the intersection between wing and fuselage tool paths, and generates additional information needed to machine the fuselage and/or wing. Output from the program can be post processed for use on a variety of milling machines. Information on program structure and methodology is given as well as the user's manual for implementation of the program.

  5. Statistical analysis of service stresses in aircraft wings

    NASA Technical Reports Server (NTRS)

    Kaul, Hans W

    1942-01-01

    On the wing structures of modern high speed aircraft, in particular, the comparatively high-service stresses and the consistently increasing number of hours of operation during the life of the separate airplane parts make the studies of strength requirement under recurrent stresses appear of major concern. The DVL has therefore made exhaustive studies for this structural group of airplanes, some results of which are reported here.

  6. The Factors that Determine the Minimum Speed of an Airplane

    NASA Technical Reports Server (NTRS)

    Norton, F H

    1921-01-01

    The author argues that because of a general misunderstanding of the principles of flight at low speed, there are a large number of airplanes that could be made to fly several miles per hour slower than at present by making slight modifications. In order to show how greatly the wing section affects the minimum speed, curves are plotted against various loadings. The disposition of wings on the airplane slightly affects the lift coefficient, and a few such cases are discussed. Another factor that has an effect on minimum speed is the extra lift exerted by the slip stream on the wings. Also discussed are procedures to be followed by the pilot, especially with regard to stick movements during low speed flight. Also covered are stalling, yaw, rolling moments, lateral control, and the effectiveness of ailerons and rudders.

  7. Advanced Subsonic Airplane Design and Economic Studies

    NASA Technical Reports Server (NTRS)

    Liebeck, Robert H.; Andrastek, Donald A.; Chau, Johnny; Girvin, Raquel; Lyon, Roger; Rawdon, Blaine K.; Scott, Paul W.; Wright, Robert A.

    1995-01-01

    A study was made to examine the effect of advanced technology engines on the performance of subsonic airplanes and provide a vision of the potential which these advanced engines offered. The year 2005 was selected as the entry-into-service (EIS) date for engine/airframe combination. A set of four airplane classes (passenger and design range combinations) that were envisioned to span the needs for the 2005 EIS period were defined. The airframes for all classes were designed and sized using 2005 EIS advanced technology. Two airplanes were designed and sized for each class: one using current technology (1995) engines to provide a baseline, and one using advanced technology (2005) engines. The resulting engine/airframe combinations were compared and evaluated on the basis on sensitivity to basic engine performance parameters (e.g. SFC and engine weight) as well as DOC+I. The advanced technology engines provided significant reductions in fuel burn, weight, and wing area. Average values were as follows: reduction in fuel burn = 18%, reduction in wing area = 7%, and reduction in TOGW = 9%. Average DOC+I reduction was 3.5% using the pricing model based on payload-range index and 5% using the pricing model based on airframe weight. Noise and emissions were not considered.

  8. Aerostructures Test Wing

    NASA Technical Reports Server (NTRS)

    Lind, RIck; Voracek, David F.; Doyle, Tim; Truax, Roger; Potter, Starr; Brenner, Marty; Voelker, Len; Freudinger, Larry; Stocjt. C (off)

    2003-01-01

    The Aerostructures Test Wing (ATW) was an apparatus used in a flight experiment during a program of research on aeroelastic instabilities. The ATW experiment was performed to study a specific instability known as flutter. Flutter is a destructive phenomenon caused by adverse coupling of structural dynamics and aerodynamics. The process of determining a flight envelope within which an aircraft will not experience flutter, known as flight flutter testing, is very dangerous and expensive because predictions of the instability are often unreliable. The ATW was a small-scale airplane wing that comprised an airfoil and boom (see upper part of Figure 1). For flight tests, the ATW was mounted on the F-15B/FTF-II testbed, which is a second-generation flight-test fixture described in Flight-Test Fixture for Aerodynamic Research (DRC- 95-27), NASA Tech Briefs, Vol. 19, No. 9, September 1995, page 84. The ATW was mounted horizontally on this fixture, and the entire assembly was attached to the undercarriage of the F-15B airplane (see lower part of Figure 1). The primary objective of the ATW project was to investigate traditional and advanced methodologies for predicting the onset of flutter. In particular, the ATW generated data that were used to evaluate a flutterometer. This particular flutterometer is an on-line computer program that uses method analysis to estimate worst-case flight conditions associated with flutter. This software was described in A Flutterometer Flight Test Tool NASA Tech Briefs, Vol. 23, No. 1, January 1999, page 52.

  9. Preliminary study of effects of winglets on wing flutter

    NASA Technical Reports Server (NTRS)

    Doggett, R. V., Jr.; Farmer, M. G.

    1976-01-01

    Some experimental flutter results are presented over a Mach number range from about 0.70 to 0.95 for a simple, swept, tapered, flat-plate wing model having a planform representative of subsonic transport airplanes and for the same wing model equipped with two different upper surface winglets. Both winglets had the same planform and area (about 2 percent of the basic-wing area); however, one weighed about 0.3 percent of the basic-wing weight, and the other weighed about 1.8 percent of the wing weight. The addition of the lighter winglet reduced the wing-flutter dynamic pressure by about 3 percent; the heavier winglet reduced the wing-flutter dynamic pressure by about 12 percent. The experimental flutter results are compared at a Mach number of 0.80 with analytical flutter results obtained by using doublet-lattice and lifting-surface (kernel-function) unsteady aerodynamic theories.

  10. Evolution of the revolutionary blended-wing-body

    NASA Technical Reports Server (NTRS)

    Liebeck, Robert H.; Page, Mark A.; Rawdon, Blaine K.

    1996-01-01

    The Blended-Wing-Body (BWB) airplane concept represents a potential revolution in subsonic transport efficiency for Very Large Airplanes (VLA's). NASA is sponsoring an advanced concept study to demonstrate feasibility and begin development of this new class of airplane. In this study, 800 passenger BWB and conventional configuration airplanes have been compared for a 7000 nautical mile design range, where both airplanes are based on technology keyed to 2015 entry into service. The BWB has been found to be superior to the conventional configuration in the following areas: Fuel Burn--31% lower, Takeoff Weight -- 1 3% lower, Operating Empty Weight -- 10% lower, Total Thrust -- 16% lower, and Lift/Drag --35% higher. The BWB advantage results from a double deck cabin that extends spanwise providing structural and aerodynamic overlap with the wing. This reduces the total wetted area of the airplane and allows a high aspect ratio to be achieved, since the deep and stiff centerbody provides efficient structural wingspan. Further synergy is realized through buried engines that ingest the wing's boundary layer, and thus reduce effective ram drag. Relaxed static stability allows optimal span loading, and an outboard leading-edge slat is the only high-lift system required.

  11. Flapping and fixed wing aerodynamics of low Reynolds number flight vehicles

    NASA Astrophysics Data System (ADS)

    Viieru, Dragos

    Lately, micro air vehicles (MAVs), with a maximum dimension of 15 cm and nominal flight speed around 10m/s, have attracted interest from scientific and engineering communities due to their potential to perform desirable flight missions and exhibit unconventional aerodynamics, control, and structural characteristics, compared to larger flight vehicles. Since MAVs operate at a Reynolds number of 105 or lower, the lift-to-drag ratio is noticeably lower than the larger manned flight vehicles. The light weight and low flight speed cause MAVs to be sensitive to wind gusts. The MAV's small overall dimensions result in low aspect ratio wings with strong wing tip vortices that further complicate the aerodynamics of such vehicles. In this work, two vehicle concepts are considered, namely, fixed wings with flexible structure aimed at passive shape control, and flapping wings aimed at enhancing aerodynamic performance using unsteady flow fields. A finite volume, pressure-based Navier-Stokes solver along with moving grid algorithms is employed to simulate the flow field. The coupled fluid-structural dynamics of the flexible wing is treated using a hyperelastic finite element structural model, the above-mentioned fluid solver via the moving grid technique, and the geometric conservation law. Three dimensional aerodynamics around a low aspect ratio wing for both rigid and flexible structures and fluid-structure interactions for flexible structures have been investigated. In the Reynolds numbers range of 7x10 4 to 9x104, the flexible wing exhibits self-initiated vibrations even in steady free-stream, and is found to have a similar performance to the identical rigid wing for modest angles of attack. For flapping wings, efforts are made to improve our understanding of the unsteady fluid physics related to the lift generation mechanism at low Reynolds numbers (75 to 1,700). Alternative moving grid algorithms, capable of handling the large movements of the boundaries (characteristic

  12. Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 1: Influence of airplane components for model D. [Langley spin tunnel tests

    NASA Technical Reports Server (NTRS)

    Ralston, J.

    1983-01-01

    The influence of airplane components, as well as wing location and tail length, on the rotational flow aerodynamics is discussed for a 1/6 scale general aviation airplane model. The airplane was tested in a built-up fashion (i.e., body, body-wing, body-wing-vertical, etc.) in the presence of two wing locations and two body lengths. Data were measured, using a rotary balance, over an angle-of-attack range of 8 deg to 90 deg, and for clockwise and counter-clockwise rotations covering an omega b/2V range of 0 to 0.9.

  13. Some Tests and Calculations Pertaining to the Dive Path and to Wing and Tail Loads in the Accident to Eastern Airlines C54B Airplane, NC-88814, Near Bainbridge, Maryland, May 30, 1947

    NASA Technical Reports Server (NTRS)

    Rhode, Richard V.; Stokke, Allen R.; Rogin, Leo

    1947-01-01

    Several dive paths were calculated for a C54 airplane starting from level flight at an altitude of 4000 feet and from an initial indicated airspeed of 200 miles per hour. The results show that, within the limits of the possible paths permitted by the evidence of the crash at Bainbridge, the speed of impact would be about 370 miles per hour and the time to crash would be between 12 1/2 and 15 1/2 seconds. Tail load calculations indicate that, with moderate negative acceleration of the airplane, the tail would fail near the end of the dive in a manner consistent in several important respects with the evidence. A number of tests were made of the elevator tab control system to determine whether the tab would move by an amount sufficient to have caused the observed dive if the stored energy in the tab control cable were suddenly released. The results of these tests indicated that the probable tab movement is such as to be capable of causing a dive similar to the one observed at Bainbridge.

  14. A Fundamental Study for Aerodynamic Characteristics of Supersonic Biplane Wing and Wing-Body Configurations

    NASA Astrophysics Data System (ADS)

    Odaka, Yusuke; Kusunose, Kazuhiro

    In order to develop a quiet supersonic transport, it is necessary to reduce shock waves around the transport. Shock waves, in general, are the cause of the airplane's sonic boom. Authors have been studying an aerodynamic feasibility of supersonic biplanes based on the concept of the Busemann biplane. In this paper, the three dimensional effect of wing geometries on their wave drags, including wing tip effects and the interference effects between the wing and a body (Wing-Body configurations) are investigated, using CFD code in Euler (inviscid) mode. As a result, we can conclude that the supersonic biplane wings at their design Mach number (M∞=1.7) are still capable of reducing wave drag significantly similar to that of the 2-D supersonic biplane.

  15. 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.

  16. Effect of drooped-nose flaps on the experimental force and moment characteristics of an oblique wing

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    Six-component experimental force and moment data are presented for a low aspect ratio, oblique wing equipped with drooped-nose flaps and mounted on top of a body of revolution. These flaps were investigated on the downstream wing panel with the nose drooped 5 deg, 10 deg, 20 deg, and 30 deg, and on both wing panels with the nose drooped 30 deg. It was to determine if such flaps would make the moment curves more linear by controlling the flow separation on the downstream wing panel at high lift coefficients. The wing was elliptical in planform and had an aspect ratio of 6.0 (based on the unswept wing span). The wing was tested at sweep angles of 45 deg and 50 deg throughout the Mach number range from 0.25 to 0.95. The drooped-nose flaps alone were not effective in making the moment curves more linear; however, a previous study showed that Kruger nose flaps improved the linearity of the moment curves when the Kruger flaps were used on only the downstream wing panel equipped with drooped-nose flaps deflected 5 deg.

  17. F-8 supercritical wing flight pressure, Boundary layer, and wake measurements and comparisons with wind tunnel data

    NASA Technical Reports Server (NTRS)

    Montoya, L. C.; Banner, R. D.

    1977-01-01

    Data for speeds from Mach 0.50 to Mach 0.99 are presented for configurations with and without fuselage area-rule additions, with and without leading-edge vortex generators, and with and without boundary-layer trips on the wing. The wing pressure coefficients are tabulated. Comparisons between the airplane and model data show that higher second velocity peaks occurred on the airplane wing than on the model wing. The differences were attributed to wind tunnel wall interference effects that caused too much rear camber to be designed into the wing. Optimum flow conditions on the outboard wing section occurred at Mach 0.98 at an angle of attack near 4 deg. The measured differences in section drag with and without boundary-layer trips on the wing suggested that a region of laminar flow existed on the outboard wing without trips.

  18. Lift-optimal aspect ratio of a revolving wing at low Reynolds number

    NASA Astrophysics Data System (ADS)

    Jardin, Thierry; Colonius, Tim

    2016-11-01

    Lentink & Dickinson (2009) showed that rotational acceleration stabilized the leading-edge vortex on revolving, low-aspect-ratio wings, and hypothesized that a Rossby number of around 3, which is achieved during each half-stroke for a variety of hovering insects, seeds, and birds, represents a convergent high-lift solution across a range of scales in nature. Subsequent work has verified that, in particular, the Coriolis acceleration is responsible for LEV stabilization. Implicit in these results is that there exists an optimal aspect ratio for wings revolving about their root, because it is otherwise unclear why, apart from possible physiological reasons, the convergent solution would not occur for an even lower Rossby number. We perform direct numerical simulations of the flow past revolving wings where we vary the aspect ratio and Rossby numbers independently by displacing the wing root from the axis of rotation. We show that the optimal lift coefficient represents a compromise between competing trends where the coefficient of lift increases monotonically with aspect ratio, holding Rossby number constant, but decreases monotonically with Rossby number, when holding aspect ratio constant. For wings revolving about their root, this favors wings of aspect ratio between 3 and 4. The authors gratefully acknowledge support from Fondation ISAE-Supaero.

  19. Airplane numerical simulation for the rapid prototyping process

    NASA Astrophysics Data System (ADS)

    Roysdon, Paul F.

    Airplane Numerical Simulation for the Rapid Prototyping Process is a comprehensive research investigation into the most up-to-date methods for airplane development and design. Uses of modern engineering software tools, like MatLab and Excel, are presented with examples of batch and optimization algorithms which combine the computing power of MatLab with robust aerodynamic tools like XFOIL and AVL. The resulting data is demonstrated in the development and use of a full non-linear six-degrees-of-freedom simulator. The applications for this numerical tool-box vary from un-manned aerial vehicles to first-order analysis of manned aircraft. A Blended-Wing-Body airplane is used for the analysis to demonstrate the flexibility of the code from classic wing-and-tail configurations to less common configurations like the blended-wing-body. This configuration has been shown to have superior aerodynamic performance -- in contrast to their classic wing-and-tube fuselage counterparts -- and have reduced sensitivity to aerodynamic flutter as well as potential for increased engine noise abatement. Of course without a classic tail elevator to damp the nose up pitching moment, and the vertical tail rudder to damp the yaw and possible rolling aerodynamics, the challenges in lateral roll and yaw stability, as well as pitching moment are not insignificant. This thesis work applies the tools necessary to perform the airplane development and optimization on a rapid basis, demonstrating the strength of this tool through examples and comparison of the results to similar airplane performance characteristics published in literature.

  20. Supercritical Wing Technology: A Progress Report on Flight Evaluations

    NASA Technical Reports Server (NTRS)

    1972-01-01

    The papers in this compilation were presented at the NASA Symposium on "Supercritical Wing Technology: A Progress Report on Flight Evaluation" held at the NASA Flight Research Center, Edwards, Calif., on February 29, 1972. The purpose of the symposium was to present timely information on flight results obtained with the F-8 and T-2C supercritical wing configurations, discuss comparisons with wind-tunnel predictions, and project [ ] flight programs planned for the F-8 and F-III (TACT) airplanes.

  1. 78 FR 73997 - Airworthiness Directives; Various Aircraft Equipped with Wing Lift Struts

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-12-10

    ... time for certain airplanes. AD 99-01-05 also required incorporating a ``NO STEP'' placard on the wing... and caused confusion. This AD clarifies the intent of the language in paragraph (c) of AD 99-01-05 and...-01-05 is being revised only to clarify language about how to maintain a sealed wing lift strut...

  2. Supersonic aeroelastic instability results for a NASP-like wing model

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.

    1993-01-01

    An experimental study and an analytical study have been conducted to examine static divergence for hypersonic-vehicle wing models at supersonic conditions. A supersonic test in the Langley Unitary Plan Wind Tunnel facility was conducted for two wind-tunnel models. These models were nearly identical with the exception of airfoil shape. One model had a four-percent maximum thickness airfoil and the other model had an eight-percent maximum thickness airfoil. The wing models had low-aspect ratios and highly swept leading edges. The all-movable wing models were supported by a single-pivot mechanism along the wing root. For both of the wind-tunnel models, configuration changes could be made in the wing-pivot location along the wing root and in the wing-pivot pitch stiffness. Three divergence conditions were measured for the four-percent thick airfoil model in the Mach number range of 2.6 to 3.6 and one divergence condition was measured for the eight-percent thick airfoil model at a Mach number of 2.9. Analytical divergence calculations were made for comparison with experimental results and to evaluate the parametric effects of wing-pivot stiffness, wing-pivot location, and airfoil thickness variations. These analyses showed that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneficial effect of increasing the divergence dynamic pressures. The calculations predicted the trend of experimental divergence dynamic pressure with Mach number accurately; however, the calculations were approximately 25 percent conservative with respect to dynamic pressure.

  3. 77 FR 7007 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-02-10

    ... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; Airbus Airplanes... airplanes; Model A330-300 series airplanes, Model A340-200 series airplanes; Model A340-300 series airplanes; Model A340-541 airplanes; and Model A340-642 airplanes. This proposed AD was prompted by reports...

  4. Investigation of wing shielding effects on CTOL engine noise

    NASA Technical Reports Server (NTRS)

    Bloomer, H. E.

    1979-01-01

    A full scale engine wing shielding investigation was conducted at the Lewis Research Center using a 97,900-N (22,000 lb) thrust turbofan engine and a simulated wing section sized around a conventional-take-off type four-engine narrow body airplane. Sound data were obtained for the wing placed at seven positions in a plane parallel to the engine axis, and were compared to data obtained without the wing at both take off and approach power. In addition, the engine was operated with and without extensive acoustic treatment, including a sonic inlet in order to evaluate wing shielding effectiveness with a highly suppressed engine. The wing shielding effectiveness was also calibrated using a 3.8 cm diam air nozzle as a second source. Results indicated that even though about 10 dB broad band shielding was achieved, the equivalent flyover noise reduction was less than 3.0 EPNdB for most configurations.

  5. Measured Engine Installation Effects of Four Civil Transport Airplanes

    NASA Technical Reports Server (NTRS)

    Senzig, David A.; Fleming, Gregg G.; Shepherd, Kevin P.

    2001-01-01

    The Federal Aviation Administration's Integrated Noise Model (INM) is one of the primary tools for land use planning around airports. The INM currently calculates airplane noise lateral attenuation using the methods contained in the Society of Automotive Engineer's Aerospace Information Report No. 1751 (SAE AIR 1751). Researchers have noted that improved lateral attenuation algorithms may improve airplane noise prediction. The authors of SAE AIR 1751 based existing methods on empirical data collected from flight tests using 1960s-technology airplanes with tail-mounted engines. To determine whether the SAE AIR 1751 methods are applicable for predicting the engine installation component of lateral attenuation for airplanes with wing-mounted engines, the National Aeronautics and Space Administration (NASA) sponsored a series of flight tests during September 2000 at their Wallops Flight Facility. Four airplanes, a Boeing 767-400, a Douglas DC-9, a Dassault Falcon 2000, and a Beech KingAir, were flown through a 20 microphone array. The airplanes were flown through the array at various power settings, flap settings, and altitudes to simulate take-off and arrival configurations. This paper presents the preliminary findings of this study.

  6. Automated optimum design of wing structures. Deterministic and probabilistic approaches

    NASA Technical Reports Server (NTRS)

    Rao, S. S.

    1982-01-01

    The automated optimum design of airplane wing structures subjected to multiple behavior constraints is described. The structural mass of the wing is considered the objective function. The maximum stress, wing tip deflection, root angle of attack, and flutter velocity during the pull up maneuver (static load), the natural frequencies of the wing structure, and the stresses induced in the wing structure due to landing and gust loads are suitably constrained. Both deterministic and probabilistic approaches are used for finding the stresses induced in the airplane wing structure due to landing and gust loads. A wing design is represented by a uniform beam with a cross section in the form of a hollow symmetric double wedge. The airfoil thickness and chord length are the design variables, and a graphical procedure is used to find the optimum solutions. A supersonic wing design is represented by finite elements. The thicknesses of the skin and the web and the cross sectional areas of the flanges are the design variables, and nonlinear programming techniques are used to find the optimum solution.

  7. Design and aerodynamic characteristics of a span morphing wing

    NASA Astrophysics Data System (ADS)

    Yu, Yuemin; Liu, Yanju; Leng, Jinsong

    2009-03-01

    Flight vehicles are often designed to function around a primary operating point such as an efficient cruise or a high maneuverability mode. Performance and efficiency deteriorate rapidly as the airplane moves towards other portions of the flight envelope. One solution to this quandary is to radically change the shape of the aircraft. This yields both improved efficiency and a larger flight envelope. This global shape change is an example of morphing aircraft . One concept of morphing is the span morphing wing in which the wingspan is varied to accommodate multiple flight regimes. This type of design allows for at least two discreet modes of the aircraft. The original configuration, in which the extensible portion of the wing is fully retracted, yields a high speed dash mode. Fully extending the wing provides the aircraft with a low speed mode tailored for fine tracking and loiter tasks. This paper discusses the design of a span morphing wing that permits a change in the aspect ratio while simultaneously supporting structural wing loads. The wing cross section is maintained by NACA 4412 rib sections . The span morphing wing was investigated in different configurations. The wing area and the aspect ratio of the span morphing wing increase as the wings pan increases. Computational aerodynamics are used to estimate the performance and dynamic characteristics of each wing shape of this span morphing wing as its wingspan is changed. Results show that in order to obtain the same lift, the conventional wing requires a larger angle of attach(AOA) than that of the span morphing wing.The lift of the span morphing wing increases as the wing span ,Mach number and AOA increases.

  8. Reduction of wing rock amplitudes using leading-edge vortex manipulations

    NASA Technical Reports Server (NTRS)

    Walton, James; Katz, Joseph

    1992-01-01

    A mechanically operated leading edge flap system was used to perturb leading edge vortex position on a free-to-roll double-delta wing. The motion of the flaps was synchronized with the wing rolling oscillations and the effect of the phase shift between the oscillations of the wing and the flaps was investigated. Experimental results indicated that this simple approach was effective in reducing the amplitude of the unintended rolling motion and its implementation to actual airplane configurations is rather simple.

  9. Tabulated pressure measurements on an executive-type jet transport model with a supercritical wing

    NASA Technical Reports Server (NTRS)

    Bartlett, D. W.

    1975-01-01

    A 1/9 scale model of an existing executive type jet transport refitted with a supercritical wing was tested on in the 8 foot transonic pressure tunnel. The supercritical wing had the same sweep as the original airplane wing but had maximum thickness chord ratios 33 percent larger at the mean geometric chord and almost 50 percent larger at the wing-fuselage juncture. Wing pressure distributions and fuselage pressure distributions in the vicinity of the left nacelle were measured at Mach numbers from 0.25 to 0.90 at angles of attack that generally varied from -2 deg to 10 deg. Results are presented in tabular form without analysis.

  10. MADCAT Aircraft Wings Optimize Their Shape For Efficient Flight

    NASA Image and Video Library

    2016-11-09

    The Mission Adaptive Digital Composites Aerostructures Technology (MADCAT) project is designing an aircraft wing that can change its shape to adapt to changing flight conditions. Constructed of lightweight lattice structures made of carbon fiber materials, the goal is to reduce drag, leading to more efficient airplanes.

  11. A flight-test evaluation of a go-around control system for a twin engine powered-lift STOL airplane

    NASA Technical Reports Server (NTRS)

    Watson, D. M.; Hardy, G. H.

    1983-01-01

    An automatic go-around control system was evaluated on the Augmentor Wing Jet Short Takeoff and Landing (STOL) Research Airplane (AWJSRA) as part of a study of an automatic landing system for a powered-lift STOL airplane. The results of the evaluation indicate that the go-around control system can successfully transition the airplane to a climb configuration from any initiation point during the glide-slope track or the flare maneuver prior to touchdown.

  12. Lateral-directional aerodynamic characteristics of light, twin-engine, propeller driven airplanes

    NASA Technical Reports Server (NTRS)

    Wolowicz, C. H.; Yancey, R. B.

    1972-01-01

    Analytical procedures and design data for predicting the lateral-directional static and dynamic stability and control characteristics of light, twin engine, propeller driven airplanes for propeller-off and power-on conditions are reported. Although the consideration of power effects is limited to twin engine airplanes, the propeller-off considerations are applicable to single engine airplanes as well. The procedures are applied to a twin engine, propeller driven, semi-low-wing airplane in the clean configuration through the linear lift range. The calculated derivative characteristics are compared with wind tunnel and flight data. Included in the calculated characteristics are the spiral mode, roll mode, and Dutch roll mode over the speed range of the airplane.

  13. Transonic wind-tunnel investigation of the maneuver potential of the NASA supercritical wing concept, phase 1

    NASA Technical Reports Server (NTRS)

    Hallissy, J. B.; Ayers, T. G.

    1977-01-01

    An investigation was conducted in the NASA Langley 8-foot transonic pressure tunnel at Mach numbers from 0.60 to 0.975 with a variable-wing-sweep airplane model in order to evaluate a series of wings designed to demonstrate the maneuver potential of the supercritical airfoil concept. Both conventional and supercritical wing designs for several planform configurations were investigated with wing sweep angles from 16.0 deg to 72.5 deg, depending on Mach number and wing configuration. The supercritical wing configuration showed significant improvements over the conventional configurations in drag-divergence Mach number and in drag level at transonic maneuver conditions.

  14. Semiempirical method for prediction of aerodynamic forces and moments on a steadily spinning light airplane

    NASA Technical Reports Server (NTRS)

    Pamadi, Bandu N.; Taylor, Lawrence W., Jr.

    1987-01-01

    A semi-empirical method is presented for the estimation of aerodynamic forces and moments acting on a steadily spinning (rotating) light airplane. The airplane is divided into wing, body, and tail surfaces. The effect of power is ignored. The strip theory is employed for each component of the spinning airplane to determine its contribution to the total aerodynamic coefficients. Then, increments to some of the coefficients which account for centrifugal effect are estimated. The results are compared to spin tunnel rotary balance test data.

  15. Study of small civil turbofan engines applicable to military trainer airplanes

    NASA Technical Reports Server (NTRS)

    Heldenbrand, R. W.; Merrill, G. L.; Burnett, G. A.

    1975-01-01

    Small turbofan engine design concepts were applied to military trainer airplanes to establish the potential for commonality between civil and military engines. Several trainer configurations were defined and studied. A ""best'' engine was defined for the trainer mission, and sensitivity analyses were performed to determine the effects on airplane size and efficiency of wing loading, power loading, configuration, aerodynamic quality, and engine quality. It is concluded that a small civil aircraft is applicable to military trainer airplanes. Aircraft designed with these engines are smaller, less costly, and more efficient than existing trainer aircraft.

  16. Annular wing

    NASA Technical Reports Server (NTRS)

    Walker, H. J. (Inventor)

    1981-01-01

    An annular wing particularly suited for use in supporting in flight an aircraft characterized by the absence of directional stabilizing surfaces is described. The wing comprises a rigid annular body of a substantially uniformly symmetrical configuration characterized by an annular positive lifting surface and cord line coincident with the segment of a line radiating along the surface of an inverted truncated cone. A decalage is established for the leading and trailing semicircular portions of the body, relative to instantaneous line of flight, and a dihedral for the laterally opposed semicircular portions of the body, relative to the line of flight. The direction of flight and climb angle or glide slope angle are established by selectively positioning the center of gravity of the wing ahead of the aerodynamic center along the radius coincident with an axis for a selected line of flight.

  17. Gordon Bennett Airplane Cup

    NASA Technical Reports Server (NTRS)

    Margoulis, W

    1921-01-01

    The characteristics of the airplanes built for the Gordon Bennet Airplane Cup race that took place on September 28, 1920 are described. The airplanes are discussed from a aerodynamical point of view, with a number of new details concerning the French machines. Also discussed is the regulation of future races. The author argues that there should be no limitations on the power of the aircraft engines. He reasons that in the present state of things, liberty with regard to engine power does not lead to a search for the most powerful engine, but for one which is reliable and light, thus leading to progress.

  18. MLS: Airplane system modeling

    NASA Technical Reports Server (NTRS)

    Thompson, A. D.; Stapleton, B. P.; Walen, D. B.; Rieder, P. F.; Moss, D. G.

    1981-01-01

    Analysis, modeling, and simulations were conducted as part of a multiyear investigation of the more important airplane-system-related items of the microwave landing system (MLS). Particular emphasis was placed upon the airplane RF system, including the antenna radiation distribution, the cabling options from the antenna to the receiver, and the overall impact of the airborne system gains and losses upon the direct-path signal structure. In addition, effort was expended toward determining the impact of the MLS upon the airplane flight management system and developing the initial stages of a fast-time MLS automatic control system simulation model. Results ot these studies are presented.

  19. The wake structure and thrust performance of a rigid low-aspect-ratio pitching panel

    PubMed Central

    BUCHHOLZ, JAMES H. J.; SMITS, ALEXANDER J.

    2009-01-01

    Thrust performance and wake structure were investigated for a rigid rectangular panel pitching about its leading edge in a free stream. For ReC = O(104), thrust coefficient was found to depend primarily on Strouhal number St and the aspect ratio of the panel AR. Propulsive efficiency was sensitive to aspect ratio only for AR less than 0.83; however, the magnitude of the peak efficiency of a given panel with variation in Strouhal number varied inversely with the amplitude to span ratio A/S, while the Strouhal number of optimum efficiency increased with increasing A/S. Peak efficiencies between 9 % and 21 % were measured. Wake structures corresponding to a subset of the thrust measurements were investigated using dye visualization and digital particle image velocimetry. In general, the wakes divided into two oblique jets; however, when operating at or near peak efficiency, the near wake in many cases represented a Kármán vortex street with the signs of the vortices reversed. The three-dimensional structure of the wakes was investigated in detail for AR = 0.54, A/S = 0.31 and ReC = 640. Three distinct wake structures were observed with variation in Strouhal number. For approximately 0.20 < St < 0.25, the main constituent of the wake was a horseshoe vortex shed by the tips and trailing edge of the panel. Streamwise variation in the circulation of the streamwise horseshoe legs was consistent with a spanwise shear layer bridging them. For St > 0.25, a reorganization of some of the spanwise vorticity yielded a bifurcating wake formed by trains of vortex rings connected to the tips of the horseshoes. For St > 0.5, an additional structure formed from a perturbation of the streamwise leg which caused a spanwise expansion. The wake model paradigm established here is robust with variation in Reynolds number and is consistent with structures observed for a wide variety of unsteady flows. Movies are available with the online version of the paper. PMID:19746195

  20. The wake structure and thrust performance of a rigid low-aspect-ratio pitching panel.

    PubMed

    Buchholz, James H J; Smits, Alexander J

    2008-04-30

    Thrust performance and wake structure were investigated for a rigid rectangular panel pitching about its leading edge in a free stream. For Re(C) = O(10(4)), thrust coefficient was found to depend primarily on Strouhal number St and the aspect ratio of the panel AR. Propulsive efficiency was sensitive to aspect ratio only for AR less than 0.83; however, the magnitude of the peak efficiency of a given panel with variation in Strouhal number varied inversely with the amplitude to span ratio A/S, while the Strouhal number of optimum efficiency increased with increasing A/S. Peak efficiencies between 9 % and 21 % were measured. Wake structures corresponding to a subset of the thrust measurements were investigated using dye visualization and digital particle image velocimetry. In general, the wakes divided into two oblique jets; however, when operating at or near peak efficiency, the near wake in many cases represented a Kármán vortex street with the signs of the vortices reversed. The three-dimensional structure of the wakes was investigated in detail for AR = 0.54, A/S = 0.31 and Re(C) = 640. Three distinct wake structures were observed with variation in Strouhal number. For approximately 0.20 < St < 0.25, the main constituent of the wake was a horseshoe vortex shed by the tips and trailing edge of the panel. Streamwise variation in the circulation of the streamwise horseshoe legs was consistent with a spanwise shear layer bridging them. For St > 0.25, a reorganization of some of the spanwise vorticity yielded a bifurcating wake formed by trains of vortex rings connected to the tips of the horseshoes. For St > 0.5, an additional structure formed from a perturbation of the streamwise leg which caused a spanwise expansion. The wake model paradigm established here is robust with variation in Reynolds number and is consistent with structures observed for a wide variety of unsteady flows. Movies are available with the online version of the paper.

  1. H-mode plasmas at very low aspect ratio on the Pegasus Toroidal Experiment

    NASA Astrophysics Data System (ADS)

    Thome, K. E.; Bongard, M. W.; Barr, J. L.; Bodner, G. M.; Burke, M. G.; Fonck, R. J.; Kriete, D. M.; Perry, J. M.; Reusch, J. A.; Schlossberg, D. J.

    2017-02-01

    H-mode is obtained at A˜ 1.2 in the Pegasus Toroidal Experiment via Ohmic heating, high-field-side fueling, and low edge recycling in both limited and diverted magnetic topologies. These H-mode plasmas show the formation of edge current and pressure pedestals and a doubling of the energy confinement time to {{H}98y,2}˜ 1 . The L-H power threshold {{P}\\text{LH}} increases with density, and there is no {{P}\\text{LH}} minimum observed in the attainable density space. The power threshold is equivalent in limited and diverted plasmas, consistent with the FM3 model. However, the measured {{P}\\text{LH}} is ˜ 15 × higher than that predicted by conventional International Tokamak Physics Activity (ITPA) scalings, and {{P}\\text{LH}}/{{P}\\text{ITPA08}} increases as A\\to 1 . Small ELMs are present at low input power {{P}\\text{IN}}˜ {{P}\\text{LH}} , with toroidal mode number n≤slant 4 . At {{P}\\text{IN}}\\gg {{P}\\text{LH}} , they transition to large ELMs with intermediate 5. The dominant-n component of a large ELM grows exponentially, while other components evolve nonlinearly and can damp prior to the crash. Direct measurements of the current profile in the pedestal region show that both ELM types exhibit a generation of a current-hole, followed by a pedestal recovery. Large ELMs are shown to further expel a current-carrying filament. Small ELM suppression via injection of low levels of helical current into the edge plasma region is also indicated.

  2. An experimental study of heat transfer and film cooling on low aspect ratio turbine nozzles

    NASA Astrophysics Data System (ADS)

    Takeishi, K.; Matsuura, M.; Aoki, S.; Sato, T.

    1989-06-01

    The effects of the three-dimensional flow field on the heat transfer and the film cooling on the endwall, suction and pressure surface of an airfoil were studied using a low speed, fully annular, low aspect h/c = 0.5 vane cascade. The predominant effects that the horseshoe vortex, secondary flow, and nozzle wake increases in the heat transfer and decreases in the film cooling on the suction vane surface and the endwall were clearly demonstrated. In addition, it was demonstrated that secondary flow has little effect on the pressure surface. Pertinent flow visualization of the flow passage was also carried out for better understanding of these complex phenomena. Heat transfer and film cooling on the fully annular vane passage surface is discussed.

  3. Low aspect ratio transonic rotors: Part 1. Baseline design and performance

    SciTech Connect

    Law, C.H. ); Wadia, A.R. )

    1993-04-01

    The analytical design and experimental test of a single-stage transonic axial-flow compressor are described. This design is the baseline of a compressor design study in which several blade design parameters have been systematically varied to determine their independent effects on compressor performance. The baseline design consisted of ruggedizing an existing compressor design that demonstrated outstanding aerodynamic performance, to correct some undesirable aeromechanical characteristics. The design study was performed by varying only one design parameter at a time, keeping the other design variables as close as possible to the baseline design. Specific design parameters of interest were those for which very few data were available to determine their sensitivity on compressor performance. This paper describes the baseline compressor design and its experimental performance. A detailed definition and flow analysis of the baseline design test point (used as the basis for all subsequent design variations) are provided.

  4. H-mode plasmas at very low aspect ratio on the Pegasus Toroidal Experiment

    DOE PAGES

    Thome, Kathleen E.; Bongard, Michael W.; Barr, Jayson L.; ...

    2016-09-30

    H-mode is obtained atmore » $$A\\sim 1.2$$ in the Pegasus Toroidal Experiment via Ohmic heating, high-field-side fueling, and low edge recycling in both limited and diverted magnetic topologies. These H-mode plasmas show the formation of edge current and pressure pedestals and a doubling of the energy confinement time to $${{H}_{98y,2}}\\sim 1$$ . The L–H power threshold $${{P}_{\\text{LH}}}$$ increases with density, and there is no $${{P}_{\\text{LH}}}$$ minimum observed in the attainable density space. The power threshold is equivalent in limited and diverted plasmas, consistent with the FM3 model. However, the measured $${{P}_{\\text{LH}}}$$ is $$\\sim 15\\,\\,\\times $$ higher than that predicted by conventional International Tokamak Physics Activity (ITPA) scalings, and $${{P}_{\\text{LH}}}/{{P}_{\\text{ITPA}08}}$$ increases as $$A\\to 1$$ . Small ELMs are present at low input power $${{P}_{\\text{IN}}}\\sim {{P}_{\\text{LH}}}$$ , with toroidal mode number $$n\\leqslant 4$$ . At $${{P}_{\\text{IN}}}\\gg {{P}_{\\text{LH}}}$$ , they transition to large ELMs with intermediate 5 < n < 15. The dominant-n component of a large ELM grows exponentially, while other components evolve nonlinearly and can damp prior to the crash. Direct measurements of the current profile in the pedestal region show that both ELM types exhibit a generation of a current-hole, followed by a pedestal recovery. Large ELMs are shown to further expel a current-carrying filament. Small ELM suppression via injection of low levels of helical current into the edge plasma region is also indicated.« less

  5. H-mode plasmas at very low aspect ratio on the Pegasus Toroidal Experiment

    SciTech Connect

    Thome, Kathleen E.; Bongard, Michael W.; Barr, Jayson L.; Bodner, Grant M.; Burke, Marcus G.; Fonck, Raymonf J.; Kriete, David M.; Perry, Justin M.; Reusch, Joshua A.; Schlossberg, David J.

    2016-09-30

    H-mode is obtained at $A\\sim 1.2$ in the Pegasus Toroidal Experiment via Ohmic heating, high-field-side fueling, and low edge recycling in both limited and diverted magnetic topologies. These H-mode plasmas show the formation of edge current and pressure pedestals and a doubling of the energy confinement time to ${{H}_{98y,2}}\\sim 1$ . The L–H power threshold ${{P}_{\\text{LH}}}$ increases with density, and there is no ${{P}_{\\text{LH}}}$ minimum observed in the attainable density space. The power threshold is equivalent in limited and diverted plasmas, consistent with the FM3 model. However, the measured ${{P}_{\\text{LH}}}$ is $\\sim 15\\,\\,\\times $ higher than that predicted by conventional International Tokamak Physics Activity (ITPA) scalings, and ${{P}_{\\text{LH}}}/{{P}_{\\text{ITPA}08}}$ increases as $A\\to 1$ . Small ELMs are present at low input power ${{P}_{\\text{IN}}}\\sim {{P}_{\\text{LH}}}$ , with toroidal mode number $n\\leqslant 4$ . At ${{P}_{\\text{IN}}}\\gg {{P}_{\\text{LH}}}$ , they transition to large ELMs with intermediate 5 < n < 15. The dominant-n component of a large ELM grows exponentially, while other components evolve nonlinearly and can damp prior to the crash. Direct measurements of the current profile in the pedestal region show that both ELM types exhibit a generation of a current-hole, followed by a pedestal recovery. Large ELMs are shown to further expel a current-carrying filament. Small ELM suppression via injection of low levels of helical current into the edge plasma region is also indicated.

  6. Marangoni-Benard convection in a low-aspect-ratio liquid layer

    NASA Technical Reports Server (NTRS)

    Duh, J. C.

    1994-01-01

    This paper presents the results of a study on the variation of the critical Marangoni number (Ma(sub c)) for the onset of Benard convection in a finite liquid layer bounded horizontally as well as from below. A direct-numerical-simulation procedure is devised to determine the Ma(sub c) for aspect ratios (Ar) ranging from 0.8 to 10. The results predict a strong increase of Ma(sub c) as Ar decreases to below 2. A dip of Ma(sub c) occurs between Ar = 1.45 and 1.3, which is accompanied by a pattern transition from a two-cell convection to a unicellular flow. For Ar above 4, the calculated Ma(sub c) shows little change and asymptotically approach a value of 116.15, with Biot number (Bi) equal to 1.

  7. Aerodynamic Effects on Mistuned Response of a High-Speed, Low Aspect Ratio Fan

    DTIC Science & Technology

    1998-10-01

    PLAYED AN IMPORTANT ROLE IN THE MISTUNED RESPONSE OF THE BLISK . 14. SUBJECT TERMS MISTUNING, MISTUNED RESPONSE, COMPRESSORS, FANS , AERODYNAMICS 17...Previous Tests on Current Test Article The blisk studied for this report has been previously tested under several programs. The first of these was the Fan ...aerodynamics play a particularly significant role in mistuning effects for modern compressor fans utilizing a blisk design. 59 5.0 REDUCED-ORDER MODELING The

  8. Shear Induced Alignment of Low Aspect Ratio Gold Nanorods in Newtonian Fluids.

    PubMed

    Xie, Donglin; Lista, Marco; Qiao, Greg G; Dunstan, Dave E

    2015-10-01

    The flow-induced alignment of small gold nanorods ranging in aspect ratio from 2.4 to 4.2 in aqueous sucrose solutions is reported. Optical absorption spectra have been measured over a range of shear rates using polarized incident light in an optically transparent quartz Couette cell. The measured spectral changes are directly attributed to the shear-induced anisotropy in the suspension due to particle alignment that saturates at Péclet number of around 200. The measured optical changes are reversible, indicating that the nanorods do not undergo aggregation during measurement. Numerical simulations show that the spectral shifts are consistent with the rods flipping between extreme orientations of the Jeffery's orbits and that the effect of the Brownian motion on the gold nanorods cannot be ignored even at large Péclet number.

  9. The flowfield and performance of a low aspect ratio unsteady propulsor

    NASA Astrophysics Data System (ADS)

    Buchholz, James H. J.

    Thrust performance and wake structure were investigated for a rigid rectangular panel pitching about its leading edge in a free stream. Thrust coefficient was found to depend primarily on Strouhal number St and the aspect ratio of the panel S/C. Propulsive efficiency was sensitive to aspect ratio only for S/C less than 0.83; however, the magnitude of the peak efficiency of a given panel with variation in Strouhal number varied inversely with the amplitude to span ratio A/S, while the Strouhal number of optimum efficiency increased with increasing A/S. Peak efficiencies between 9% and 21% were measured. Wake structures corresponding to a subset of the thrust measurements were investigated using dye visualization and Digital Particle Image Velocimetry. In general, the wakes divided into two oblique jets; however, when operating at or near peak efficiency, the near wake in many cases represented a von Karman vortex street with the signs of the vortices reversed. The three-dimensional structure of the wakes were investigated in detail for S/C = 0.54, A/S = 0.31, ReC = 640. Three distinct wake structures were observed with variation in Strouhal number. For approximately 0.20 < St < 0.25, the main constituent of the wake was a horseshoe vortex shed by the tips and trailing edge of the panel. Streamwise variation in the circulation of the streamwise horseshoe legs was consistent with a spanwise shear layer bridging them. For St > 0.25, a reorganization of some of the spanwise vorticity yielded a bifurcating wake formed by trains of vortex rings connected to the tips of the horseshoes. For St > 0.5 an additional structure formed from a perturbation of the streamwise leg which caused a spanwise expansion. The wake model paradigm established here is consistent with structures observed for a wide variety of unsteady flows.

  10. Prediction of Natural Convection Flow Pattern in Low-Aspect Ratio Enclosures.

    DTIC Science & Technology

    1982-05-26

    defined as- D Mass diffusivity g Acceleration due to gravity GrH Grashof number based on H, sgTH V GrL Grashof number based on L, BgATL 3 L~2- V H Height...to very small value with arbitrary but fixed Grashof number, GrH , by the matched asymptotic expansion method. Their results show that when A - 0, the...However, they could not indicate explicitly the upper limit of GrH within which their solution is valid. They just gave an approximate criterion by

  11. Stress intensity factors for low aspect ratio longitudinal surface cracks in thin pipes

    SciTech Connect

    Wang, X.; Lambert, S.B.

    1996-12-31

    Finite element calculations of stress intensity factors for longitudinal external and internal semi-elliptical surface cracks in thin pipes (R/t of 10) and external edge cracks in thin pipes (R/t of 10, 20, 40) are presented. Constant, linear, parabolic and cubic stress distributions were applied to the crack face. Comparisons of flat plate and thin pipe stress intensity factor results are made for the deepest and surface points of the surface cracks, and for edge cracks. Recommendations are given regarding the limits of applicability of flat plate solutions. For a pipe with R/t of 10, the stress intensity factors for shallow semi-elliptical cracks and edge cracks (a/t less than 0.4) can be approximated to within 6% using the corresponding solutions for cracks in flat plates. For pipes with R/t larger than 40, the stress intensity factors for a wide range of crack depths, 0 {le} a/t {le} 0.8, can be approximated with good accuracy using the corresponding solutions for cracks in flat plates.

  12. H-mode plasmas at very low aspect ratio on the Pegasus Toroidal Experiment

    SciTech Connect

    Thome, Kathleen E.; Bongard, Michael W.; Barr, Jayson L.; Bodner, Grant M.; Burke, Marcus G.; Fonck, Raymonf J.; Kriete, David M.; Perry, Justin M.; Reusch, Joshua A.; Schlossberg, David J.

    2016-09-30

    H-mode is obtained at $A\\sim 1.2$ in the Pegasus Toroidal Experiment via Ohmic heating, high-field-side fueling, and low edge recycling in both limited and diverted magnetic topologies. These H-mode plasmas show the formation of edge current and pressure pedestals and a doubling of the energy confinement time to ${{H}_{98y,2}}\\sim 1$ . The L–H power threshold ${{P}_{\\text{LH}}}$ increases with density, and there is no ${{P}_{\\text{LH}}}$ minimum observed in the attainable density space. The power threshold is equivalent in limited and diverted plasmas, consistent with the FM3 model. However, the measured ${{P}_{\\text{LH}}}$ is $\\sim 15\\,\\,\\times $ higher than that predicted by conventional International Tokamak Physics Activity (ITPA) scalings, and ${{P}_{\\text{LH}}}/{{P}_{\\text{ITPA}08}}$ increases as $A\\to 1$ . Small ELMs are present at low input power ${{P}_{\\text{IN}}}\\sim {{P}_{\\text{LH}}}$ , with toroidal mode number $n\\leqslant 4$ . At ${{P}_{\\text{IN}}}\\gg {{P}_{\\text{LH}}}$ , they transition to large ELMs with intermediate 5 < n < 15. The dominant-n component of a large ELM grows exponentially, while other components evolve nonlinearly and can damp prior to the crash. Direct measurements of the current profile in the pedestal region show that both ELM types exhibit a generation of a current-hole, followed by a pedestal recovery. Large ELMs are shown to further expel a current-carrying filament. Small ELM suppression via injection of low levels of helical current into the edge plasma region is also indicated.

  13. Factors Affecting Lateral Stability and Controllability

    NASA Technical Reports Server (NTRS)

    Campbell, John P; Toll, Thomas A

    1948-01-01

    The effects on dynamic lateral stability and controllability of some of the important aerodynamic and mass characteristics are discussed and methods are presented for estimating the various stability parameters to be used in the calculation of the dynamic lateral stability of airplanes with swept and low-aspect-ratio wings.

  14. Numerical prediction of airplane trailing vortices

    NASA Astrophysics Data System (ADS)

    Czech, M. J.; Crouch, J. D.; Miller, G. D.; Strelets, M.

    2004-11-01

    The accurate prediction of airplane trailing vortices is of great interest for both cruise conditions in conjunction with the formation of contrails as well as approach conditions for reasons of flight safety and active vortex control. A numerical approach is introduced based on a quasi-3D Reynolds-Averaged Navier-Stokes formulation with a one-equation turbulence model. The numerical results show good agreement with wind-tunnel data out to ten spans for a range of wing and tail loadings typical of commercial airplanes in a landing configuration. The results show a one-, two- and three-pair system in the near-field with only minor changes to the initial lift distribution. The CFD correctly predicts the strength, demise and position of the individual vortex pairs over a range of test cases. The approach is further extended by considering thrust effects. For cruise conditions, far field predictions show the entrainment of the jet plume into the wake and provide the potential for coupling with a micro-physics model to predict the formation and early evolution of contrails. Potential influences of configuration details on the plume entrainment are considered. This numerical method also offers an attractive approach for assessing active schemes designed to accelerate the break-up of airplane trailing vortices.

  15. Rapid Airplane Parametric Input Design(RAPID)

    NASA Technical Reports Server (NTRS)

    Smith, Robert E.; Bloor, Malcolm I. G.; Wilson, Michael J.; Thomas, Almuttil M.

    2004-01-01

    An efficient methodology is presented for defining a class of airplane configurations. Inclusive in this definition are surface grids, volume grids, and grid sensitivity. A small set of design parameters and grid control parameters govern the process. The general airplane configuration has wing, fuselage, vertical tail, horizontal tail, and canard components. The wing, tail, and canard components are manifested by solving a fourth-order partial differential equation subject to Dirichlet and Neumann boundary conditions. The design variables are incorporated into the boundary conditions, and the solution is expressed as a Fourier series. The fuselage has circular cross section, and the radius is an algebraic function of four design parameters and an independent computational variable. Volume grids are obtained through an application of the Control Point Form method. Grid sensitivity is obtained by applying the automatic differentiation precompiler ADIFOR to software for the grid generation. The computed surface grids, volume grids, and sensitivity derivatives are suitable for a wide range of Computational Fluid Dynamics simulation and configuration optimizations.

  16. Recent Load Calibrations Experience with the YF-12 Airplane

    NASA Technical Reports Server (NTRS)

    Jenkins, J. M.; Kuhl, A. E.

    1978-01-01

    The use of calibrated strain gages to measure wing loads on the YF-12A airplane is discussed as well as structural configurations relative to the thermal environment and resulting thermal stresses. A thermal calibration of the YF-12A is described to illustrate how contaminating thermal effects can be removed from loads equations. The relationship between ground load calibrations and flight measurements is examined for possible errors, and an analytical approach to accommodate such errors is presented.

  17. English airplane construction

    NASA Technical Reports Server (NTRS)

    Schwencke, D

    1930-01-01

    English airplane construction is presented with a particular emphasis on metal construction techniques. Steel rib and fuselage construction are discussed as well as the use of duralumin in construction.

  18. The Bristol "Badminton" Airplane

    NASA Technical Reports Server (NTRS)

    1926-01-01

    The Bristol Badminton, Type 99 airplane has a radial aircooled engine (a Bristol Jupiter 9 cylinder 450 HP.) and three fuel tanks. It is a single seat biplane weighing 1,840 lbs. empty and 2,460 lbs. loaded.

  19. 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. PMID:27582902

  20. Stall-proof Airplanes

    NASA Technical Reports Server (NTRS)

    Lachmann, G

    1927-01-01

    My lecture has to do with the following questions. Is the danger of stalling necessarily inherent in the airplane in its present form and structure, or can it be diminished or eliminated by suitable means? Do we possess such means or devices and how must they operate? In this connection I will devote special attention to the exhibition of stall-proof airplanes by Fokker under the auspices of the English Air Ministry, which took place in Croyden last April.

  1. Statical longitudinal stability of airplanes

    NASA Technical Reports Server (NTRS)

    Warner, Edward P

    1921-01-01

    This report, which is a continuation of the "Preliminary report on free flight testing" (report no. NACA-TR-70), presents a detailed theoretical analysis of statical stability with free and locked controls and also the results of many free flight test on several types of airplanes. In developing the theory of stability with locked controls an expression for pitching moment is derived in simple terms by considering the total moment as the sum of the moments due to wings and tail surface. This expression, when differentiated with respect to angle of incidence, enables an analysis to be made of the factors contributing to the pitching moment. The effects of slipstream and down wash are also considered and it is concluded that the C. G. Location has but slight effect or stability, and that stability is much improved by increasing the efficiency of the tail surfaces, which may be done by using an "inverted" tail plane. The results of free flight tests with locked controls are discussed at length and it is shown that the agreement between the experimental results and theory is very satisfactory. The theory of stability with free controls is not amendable to the simple mathematical treatment used in the case of locked controls, but a clear statement of the conditions enables several conclusions to be drawn, one of which is that the fixed tail surfaces should be much larger than the movable surfaces.

  2. 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.

  3. 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.

  4. Parametric study of variation in cargo-airplane performance related to progression from current to spanloader designs

    NASA Technical Reports Server (NTRS)

    Toll, T. A.

    1980-01-01

    A parametric analysis was made to investigate the relationship between current cargo airplanes and possible future designs that may differ greatly in both size and configuration. The method makes use of empirical scaling laws developed from statistical studies of data from current and advanced airplanes and, in addition, accounts for payload density, effects of span distributed load, and variations in tail area ratio. The method is believed to be particularly useful for exploratory studies of design and technology options for large airplanes. The analysis predicts somewhat more favorable variations of the ratios of payload to gross weight and block fuel to payload as the airplane size is increased than has been generally understood from interpretations of the cube-square law. In terms of these same ratios, large all wing (spanloader) designs show an advantage over wing-fuselage designs.

  5. Free-Spinning-Tunnel Investigation of a 1/30-Scale Model of the Grumman XF1OF-1 Airplane

    NASA Technical Reports Server (NTRS)

    Berman, Theodore

    1950-01-01

    An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/30 - scale model of the Grumman XFlOF-1 airplane to determine its spin and recovery characteristics. The investigation included erect and inverted spins for both the straight-wing and swept-wing configurations. Tests to determine the optimum size spin-recovery parachutes and the rudder forces required for recovery were also made. The results indicated that in the straight-wing configuration, satisfactory recoveries of the airplane will be obtained from erect and inverted spins by rudder reversal alone. In the swept-wing configuration recoveries will be unsatisfactory from erect spins. Unsweeping the wings during the spin and reversal of the rudder, however, will lead to eventual recovery. The test results also indicated that, if existing small ailerons are made deflectable through large angles, satisfactory recoveries will be obtained from erect spins in the swept-wing configuration by simultaneous movement of the rudder to against the spin and movement of the ailerons to with the spin. Normal-size ailerons deflected through a normal range would also be effective. Satisfactory recoveries by rudder reversal will be obtained from inverted spins in the swept-wing configuration. In the straight-wing configuration a 14.2-foot tail parachute or a 5.0-foot wing-tip parachute opened on the outer wing tip will effect satisfactory recovery of the airplane by parachute action alone; a 30.0-foot tail parachute or a 10.0-foot wing-tip parachute will be required for the swept-wing configuration. The forces required to fully reverse the rudder should be within the capabilities of the pilot.

  6. Lateral control required for satisfactory flying qualities based on flight tests of numerous airplanes

    NASA Technical Reports Server (NTRS)

    Gilruth, R R; Turner, W N

    1941-01-01

    Report presents the results of an analysis made of the aileron control characteristics of numerous airplanes tested in flight by the National Advisory Committee for Aeronautics. By the use of previously developed theory, the observed values of pb/2v for the various wing-aileron arrangements were examined to determine the effective section characteristics of the various aileron types.

  7. Vanguard 2C VTOL Airplane Tested in the Ames 40x80 Foot Wind Tunnel.

    NASA Image and Video Library

    1960-02-01

    Vanguard 2C vertical take-off and landing (VTOL) airplane, wind tunnel test. Front view from below, model 14 1/2 feet high disk off. Nasa Ames engineer Ralph Maki in photo. Variable height struts and ground plane, low pressure ratio, fan in wing. 02/01/1960.

  8. 75 FR 52255 - Airworthiness Directives; Air Tractor, Inc. Models AT-802 and AT-802A Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-08-25

    ...; AD 2010-17-18] RIN 2120-AA64 Airworthiness Directives; Air Tractor, Inc. Models AT-802 and AT- 802A... all Air Tractor, Inc. (Air Tractor) Models AT-802 and AT-802A airplanes. AD 2010-13-08 currently... both of the wing main spar lower caps at the center splice joint for cracks and repair or replace...

  9. 75 FR 35616 - Airworthiness Directives; Air Tractor, Inc. Models AT-802 and AT-802A Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-06-23

    .... Models AT-802 and AT- 802A Airplanes AGENCY: Federal Aviation Administration (FAA), Department of... supersede AD 2006-08-09, which applies to all Air Tractor, Inc. (Air Tractor) Models AT-802 and AT-802A... two outboard fastener holes in both of the wing main spar lower caps at the center splice joint...

  10. Study of the single body yawed-wing aircraft concept

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.; Nisbet, J. W.; Neuman, F. D.; Hamilton, E. J.; Murakami, J. K.; Mcbarron, J. P.; Kumasaka, K.

    1974-01-01

    Areas relating to the development and improvement of the single-fuselage, yawed-wing transonic transport concept were investigated. These included: (1) developing an alternate configuration with a simplified engine installation;(2) determining a structural design speed placard that would allow the engine-airframe match for optimum airplane performance; and (3) conducting an aeroelastic stability and control analysis of the yawed-wing configuration with a flexible wing. A two-engine, single-fuselage, yawed-wing configuration was developed that achieved the Mach 1.2 design mission at 5560 km (3000 nmi) and payload of 18,140 kg (40,000 lb) with a gross weight of 217,700 kg (480,000 lb). This airplane was slightly heavier than the aft-integrated four-engine configuration that had been developed in a previous study. A modified structural design speed placard, which was determined, resulted in a 6% to 8% reduction in the gross weight of the yawed-wing configurations. The dynamic stability characteristics of the single-fuselage yawed-wing configuration were found to be very dependent on the magnitude of the pitch/roll coupling, the static longitudinal stability, and the dihedral effect.

  11. Rapid Airplane Parametric Input Design (RAPID)

    NASA Technical Reports Server (NTRS)

    Smith, Robert E.

    1995-01-01

    RAPID is a methodology and software system to define a class of airplane configurations and directly evaluate surface grids, volume grids, and grid sensitivity on and about the configurations. A distinguishing characteristic which separates RAPID from other airplane surface modellers is that the output grids and grid sensitivity are directly applicable in CFD analysis. A small set of design parameters and grid control parameters govern the process which is incorporated into interactive software for 'real time' visual analysis and into batch software for the application of optimization technology. The computed surface grids and volume grids are suitable for a wide range of Computational Fluid Dynamics (CFD) simulation. The general airplane configuration has wing, fuselage, horizontal tail, and vertical tail components. The double-delta wing and tail components are manifested by solving a fourth order partial differential equation (PDE) subject to Dirichlet and Neumann boundary conditions. The design parameters are incorporated into the boundary conditions and therefore govern the shapes of the surfaces. The PDE solution yields a smooth transition between boundaries. Surface grids suitable for CFD calculation are created by establishing an H-type topology about the configuration and incorporating grid spacing functions in the PDE equation for the lifting components and the fuselage definition equations. User specified grid parameters govern the location and degree of grid concentration. A two-block volume grid about a configuration is calculated using the Control Point Form (CPF) technique. The interactive software, which runs on Silicon Graphics IRIS workstations, allows design parameters to be continuously varied and the resulting surface grid to be observed in real time. The batch software computes both the surface and volume grids and also computes the sensitivity of the output grid with respect to the input design parameters by applying the precompiler tool

  12. Rapid Airplane Parametric Input Design (RAPID)

    NASA Technical Reports Server (NTRS)

    Smith, Robert E.

    1995-01-01

    RAPID is a methodology and software system to define a class of airplane configurations and directly evaluate surface grids, volume grids, and grid sensitivity on and about the configurations. A distinguishing characteristic which separates RAPID from other airplane surface modellers is that the output grids and grid sensitivity are directly applicable in CFD analysis. A small set of design parameters and grid control parameters govern the process which is incorporated into interactive software for 'real time' visual analysis and into batch software for the application of optimization technology. The computed surface grids and volume grids are suitable for a wide range of Computational Fluid Dynamics (CFD) simulation. The general airplane configuration has wing, fuselage, horizontal tail, and vertical tail components. The double-delta wing and tail components are manifested by solving a fourth order partial differential equation (PDE) subject to Dirichlet and Neumann boundary conditions. The design parameters are incorporated into the boundary conditions and therefore govern the shapes of the surfaces. The PDE solution yields a smooth transition between boundaries. Surface grids suitable for CFD calculation are created by establishing an H-type topology about the configuration and incorporating grid spacing functions in the PDE equation for the lifting components and the fuselage definition equations. User specified grid parameters govern the location and degree of grid concentration. A two-block volume grid about a configuration is calculated using the Control Point Form (CPF) technique. The interactive software, which runs on Silicon Graphics IRIS workstations, allows design parameters to be continuously varied and the resulting surface grid to be observed in real time. The batch software computes both the surface and volume grids and also computes the sensitivity of the output grid with respect to the input design parameters by applying the precompiler tool

  13. A spin-recovery parachute system for light general aviation airplanes

    NASA Technical Reports Server (NTRS)

    Bradshaw, C. F.

    1980-01-01

    A tail-mounted spin-recovery parachute system has been designed and developed by the NASA Langley Research Center for use on light general aviation airplanes. The system was designed for use on typical airplane configurations, including low-wing, high-wing, single- and twin-engine designs. A mechanically triggered pyrotechnic slug gun is used to forcibly deploy a pilot parachute which extracts a bag that deploys a ring-slot spin-recovery parachute. The total system weighs 8.2 kg (18 lb). System design factors included airplane wake effects on parachute deployment, prevention of premature parachute deployment, positive parachute jettison, compact size, low weight, system reliability, and pilot and ground crew safety. Extensive ground tests were conducted to qualify the system. The recovery parachute has been used successfully in flight 17 times.

  14. A spin-recovery parachute system for light general-aviation airplanes

    NASA Technical Reports Server (NTRS)

    Bradshaw, C.

    1980-01-01

    A tail mounted spin recovery parachute system was designed and developed for use on light general aviation airplanes. The system was designed for use on typical airplane configurations, including low wing, high wing, single engine and twin engine designs. A mechanically triggered pyrotechnic slug gun is used to forcibly deploy a pilot parachute which extracts a bag that deploys a ring slot spin recovery parachute. The total system weighs 8.2 kg. System design factors included airplane wake effects on parachute deployment, prevention of premature parachute deployment, positive parachute jettison, compact size, low weight, system reliability, and pilot and ground crew safety. Extensive ground tests were conducted to qualify the system. The recovery parachute was used successfully in flight 17 times.

  15. Experiments on a Slotted Wing

    NASA Technical Reports Server (NTRS)

    Ruden, P

    1939-01-01

    The results of pressure distribution measurements that were made on a model wing section of a Fieseler F 5 R type airplane are presented. Comparison of those model tests with the corresponding flight tests indicates the limitations and also the advantages of wind tunnel investigations, the advantages being particularly that through the variety of measuring methods employed the more complicated flow conditions may also be clarified. A fact brought out in these tests is that even in the case of "well rounded" slots it is possible for a vortex to be set up at the slot entrance and this vortex is responsible for certain irregularities in the pressure distribution and in the efficiency of the slot.

  16. Wind-Tunnel Tests on Various Types of Dive Brakes Mounted in Proximity of the Leading Edge of the Wing

    NASA Technical Reports Server (NTRS)

    Lattanzi, Bernardino; Bellante, Erno

    1949-01-01

    The present report is concerned with a series of tests on a model airplane fitted with four types of dive flaps of various shapes, positions, and incidence located near the leading edge of the wing (from 5 to 20 percent of the wing chord). Tests were also made on a stub airfoil fitted with a ventral dive (located at 8 percent of the wing chord). The hinge moments of the dive flaps were measured.

  17. SCAR arrow-wing active flutter suppression system

    NASA Technical Reports Server (NTRS)

    Gordon, C. K.; Visor, O. E.

    1977-01-01

    The potential performance and direct operating cost benefits of an active flutter suppression system (FSS) for the NASA arrow-wing supersonic cruise configuration were determined. A FSS designed to increase the flutter speed of the baseline airplane 20 percent. A comparison was made of the performance and direct operating cost between the FSS equipped aircraft and a previously defined configuration with structural modifications to provide the same flutter speed. Control system synthesis and evaluation indicated that a FSS could provide the increase in flutter speed without degrading airplane reliability, safety, handling qualities, or ride quality, and without increasing repeated loads or hydraulic and electrical power capacity requirements.

  18. Surface pressure data for a supersonic-cruise airplane configuration at Mach numbers of 2.30, 2.96, 3.30

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.; Corlett, W. A.; Collins, I. K.

    1979-01-01

    The tabulated results of surface pressure tests conducted on the wing and fuselage of an airplane model in the Langley Unitary Plan wind tunnel are presented without analysis. The model tested was that of a supersonic-cruise airplane with a highly swept arrow-wing planform, two engine nacelles mounted beneath the wing, and outboard vertical tails. Data were obtained at Mach numbers of 2.30, 2.96, and 3.30 for angles of attack from -4 deg to 12 deg. The Reynolds number for these tests was 6,560,000 per meter.

  19. Tabulated pressure measurements on a large subsonic transport model airplane with high bypass ratio, powered, fan jet engines

    NASA Technical Reports Server (NTRS)

    Flechner, S. G.; Patterson, J. C., Jr.

    1972-01-01

    An experimental wind-tunnel investigation to determine the aerodynamic interference and the jet-wake interference associated with the wing, pylon, and high-bypass-ratio, powered, fan-jet model engines has been conducted on a typical high-wing logistics transport airplane configuration. Pressures were measured on the wing and pylons and on the surfaces of the engine fan cowl, turbine cowl, and plug. Combinations of wing, pylons, engines, and flow-through nacelles were tested, and the pressure coefficients are presented in tabular form. Tests were conducted at Mach numbers from 0.700 to 0.825 and angles of attack from -2 to 4 deg.

  20. Reduction of vertical-gust induced loads on elastic airplane model with slotted-interceptor-type surfaces

    NASA Astrophysics Data System (ADS)

    Mazutsky, A. Yu.

    2007-06-01

    Results of T-203 wind-tunnel tests of an elastic airplane model dynamically similar to Tu-204 aircraft equipped with a passive wing load reduction system are reported. The wing load reduction system comprised two auxiliary aerodynamic surfaces of the slotted-interceptor type installed symmetrically on the upper surface of the lifting wing at the boundary-layer height. It was found that, with the operating system, the additional bending moments induced by symmetric vertical discrete gusts in the wing root and at the fuselage mid-body decreased by 20%, with simultaneous increase observed in the flexure-pylon flutter speed.