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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. Low aspect ratio wings at high angles of attack

    NASA Technical Reports Server (NTRS)

    Stallings, R. L., Jr.

    1986-01-01

    A comprehensive evaluation is made of experimental data compiled to date for the flowfields and aerodynamic forces that occur at high angles of attack for low aspect ratio wings with delta, rectangular, clipped delta, and strake/wing planform geometries. Attention is given to wing leading edge-generated vortex breakdown, aspect ratio and compressibility effects, and strake vortex effects on main wing areas. Although the nonlinear effects created by a wing-body combination significantly alter wing-alone aerodynamics, the wing-alone data presented are vital to the development of prediction methodologies for large angle of attack aerodynamics.

  3. Low aspect ratio wings at high angles of attack

    NASA Technical Reports Server (NTRS)

    Stallings, Robert L., Jr.

    1992-01-01

    A survey is conducted of the results of investigations into the flowfields and aerodynamic forces associated with low aspect ratio wings at high angles of attack. Attention is given to criteria for the cataloging of these flowfields, the phenomenon of vortex breakdown, with varying wing incidence and Mach number, and the effects of aspect ratio and compressibility. The planforms treated are of rectangular, clipped-delta, and strake-wing combination geometries. Extensive graphic representations of performance trends with varying parameters are furnished.

  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. 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.

  6. 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 has been computed using 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. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, 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 wing 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 clearly observed by following the trajectory of particles released around the wing tip.

  7. 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.

  8. Aerodynamics of high-lift, low-aspect-ratio unswept wings

    NASA Technical Reports Server (NTRS)

    Katz, Joseph

    1989-01-01

    It is presently demonstrated that highly cambered, low aspect ratio airfoil configurations can be neither experimentally nor numerically developed on the basis of two-dimensional methods; the strong dependency of airfoil shape on aspect ratio requires the definition of a planform to be conducted first, in order to serve as input for three-dimensional optimization methods. It is found that even the simplified three-dimensional computational method presently employed can accelerate a multielement-airfoil wing development.

  9. 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.

  10. Effects of ground proximity on a low aspect ratio propulsive wing/canard configuration

    NASA Technical Reports Server (NTRS)

    Stewart, V. R.; Kemmerly, G. T.

    1987-01-01

    The effect of near proximity to the ground are investigated on a low aspect ratio propulsive wing/canard concept at STOL conditions. Data were obtained on a wing/body and wing/body/canard configuration at various heights above the ground, ranging from free air to approximately 1/4 of the mean aerodynamic chord (MAC) above the ground. The data presented and discussed include force and moment coefficients, surface pressure distributions, and downwash angles measured one MAC behind the wing. The test technique, model requirements, and special considerations required for testing these configurations are also discussed. Special model requirements included evenly distributed exit nozzle pressures along four separate nozzles of lengths of one and two feet with only one air supply to the model. Test techniques must recognize and deal with the ground boundary layer as well as the air supply pressure measurement and management.

  11. 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.

  12. Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.

    1990-01-01

    An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

  13. 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.

  14. 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.

  15. Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

    NASA Technical Reports Server (NTRS)

    Maki, Ralph L.

    1959-01-01

    Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

  16. 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.

  17. The immersed boundary projection method and its application to simulation and control of flows around low-aspect-ratio wings

    NASA Astrophysics Data System (ADS)

    Taira, Kunihiko

    First, we present a new formulation of the immersed boundary method that is algebraically identical to the traditional fractional step algorithm. This method, called the immersed boundary projection method, allows for the simulations of incompressible flows over arbitrarily shaped bodies under motion and/or deformation in both two and three dimensions. The no-slip condition along the immersed boundary is enforced simultaneously with the incompressibility constraint through a single projection. The boundary force is determined implicitly without any constitutive relations for the rigid body formulation, which in turn allows the use of high CFL numbers in our simulations compared to past methods. Next, the above immersed boundary projection method is used to analyze three-dimensional separated flows around low-aspect-ratio flat-plate wings. A number of simulations highlighting the unsteady nature of the separated flows are performed for Re=300 and 500 with various aspect ratios, angles of attack, and planform geometries. The aspect ratio and angle of attack are found to have a large influence on the stability of the wake profile and the force experienced by the low-aspect-ratio wing. At early times, following an impulsive start, topologies of the wake vortices are found to be the same across different aspect ratios and angles of attack. Behind low-aspect-ratio rectangular plates, leading-edge vortices form and eventually separate as hairpin vortices following the start-up. This phenomenon is found to be similar to dynamic stall observed behind pitching plates. The detached structure would then interact with the tip vortices, reducing the downward velocity induced by the tip vortices acting upon the leading-edge vortex. At large time, depending on the aspect ratio and angles of attack, the wakes reach one of the three states: (i) a steady state, (ii) a periodic unsteady state, or (iii) an aperiodic unsteady state. We have observed that the tip effects in three

  18. 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.

  19. 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.

  20. 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. PMID:24434625

  1. A theoretical investigation of the aerodynamics of low-aspect-ratio wings with partial leading-edge separation

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

    A numerical method is developed to predict distributed and total aerodynamic characteristics for low aspect-ratio wings with partial leading-edge separation. 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 mid-points 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. Comparison of the predicted results with complete leading-edge separation has shown reasonably good agreement. For cases with partial leading-edge separation, the lift is found to be highly nonlinear with angle of attack.

  2. 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

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Flight measurements of lifting pressures for a thin low-aspect-ratio wing at subsonic, transonic, and low supersonic speeds

    NASA Technical Reports Server (NTRS)

    Byrdsong, T. A.

    1977-01-01

    Pressure distributions in the form of differential pressure coefficients are presented for several wing chordwise and spanwise stations. Also presented are the results of limited analysis which show aircraft configuration effects, Mach number effects on the local wing loadings, comparisons of selected measured wing pressures with predicted pressures, and comparisons of wing loadings during right-turn and left-turn maneuvers.

  8. Powered low-aspect-ratio Wing In Ground effect (WIG) aerodynamic characteristics. [conducted in Langley V/STOL tunnel

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    A wing-in-ground effect configuration was investigated. The configuration used large diameter, low pressure ratio fans mounted about 0.76 wing chord ahead of the wing leading edge to achieve a power augmented ram wing during operation in ground effect. Tests of both in and out of ground effect aerodynamic transition characteristics from very low speeds to cruise speeds are described. The investigation provided a number of conclusions concerning the aerodynamic/propulsive performance interaction. While power augmented lift is required for low speed flight, there is a thrust loss when the efflux is trapped under the wing which reduced the effective thrust to weight available for acceleration by about a third of the installed thrust to weight ratio.

  9. Ground Effects on the Longitudinal Characteristics of Two Models with Wings Having Low Aspect Ratio and Pointed Tips

    NASA Technical Reports Server (NTRS)

    Buell, Donald A; Tinling, Bruce E

    1957-01-01

    Wind-tunnel tests were conducted to determine the ground effects on a tailless model with a wing of aspect ratio 2 and infinite taper, and on a tailed model with a triangular wing of aspect ratio 3, with flaps. Control-surface hinge moments were measured on the tailless model. The results are compared with the predictions of the theory of Tani, et al.

  10. 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.

  11. Extreme Low Aspect Ratio Stellarators

    NASA Astrophysics Data System (ADS)

    Moroz, Paul

    1997-11-01

    Recently proposed Spherical Stellarator (SS) concept [1] includes the devices with stellarator features and low aspect ratio, A <= 3.5, which is very unusual for stellarators (typical stellarators have A ≈ 7-10 or above). Strong bootstrap current and high-β equilibria are two distinguished elements of the SS concept leading to compact, steady-state, and efficient fusion reactor. Different coil configurations advantageous for the SS have been identified and analyzed [1-6]. In this report, we will present results on novel stellarator configurations which are unusual even for the SS approach. These are the extreme-low-aspect-ratio-stellarators (ELARS), with the aspect ratio A ≈ 1. We succeeded in finding ELARS configurations with extremely compact, modular, and simple design compatible with significant rotational transform (ι ≈ 0.1 - 0.15), large plasma volume, and good particle transport characteristics. [1] P.E. Moroz, Phys. Rev. Lett. 77, 651 (1996); [2] P.E. Moroz, Phys. Plasmas 3, 3055 (1996); [3] P.E. Moroz, D.B. Batchelor et al., Fusion Tech. 30, 1347 (1996); [4] P.E. Moroz, Stellarator News 48, 2 (1996); [5] P.E. Moroz, Plasma Phys. Reports 23, 502 (1997); [6] P.E. Moroz, Nucl. Fusion 37, No. 8 (1997). *Supported by DOE Grant No. DE-FG02-97ER54395.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. On virial analysis at low aspect ratio

    DOE PAGES

    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 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

  18. 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.

  19. Low-aspect-ratio optimization studies for ATF-II

    SciTech Connect

    Hirshman, S.P.; Morris, R.N.; Hedrick, C.L.; Lyon, J.F.; Rome, J.A.; Painter, S.L.; Van Rij, W.I.

    1989-01-01

    A numerical procedure for optimizing stellarator MHD and transport properties at finite ..beta.. is described. This method is applied to finding a low-aspect-ratio ATF-II configuration. 4 refs., 6 figs.

  20. 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.

  1. 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.

  2. 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.

  3. Advances in low aspect ratio stellarator coil design

    NASA Astrophysics Data System (ADS)

    Valanju Miner, P. M., Jr.; Hirshman, S. P.; Brooks, A.; Pomphrey, N.

    1999-11-01

    Recent advances have led to two new exciting low aspect ratio stellarator plasma configurations - QA: Quasi-Axisymmetric and QO: Quasi-Omnigeneous Stellarators. The difficult task of designing an optimal set of stellarator coils to produce these configurations is critical to the success of these new ideas. Since existing computational tools were found to be inadequate for these low aspect ratio configurations, new methods, consisting of singular value decomposition, surface displacement targeting, and Genetic Algorithms, have been developed. These new methods have led to successful designs of optimal, low current density coils which reproduce the target plasma configurations with good accuracy and flexibility while satisfying various engineering constraints. Results for modular and saddle coil sets for both QA and QO machines will be presented.

  4. 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.

  5. 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.

  6. Autonomous Deicing System For Airplane Wing

    NASA Technical Reports Server (NTRS)

    Hickman, G. A.; Gerardi, J. J.

    1993-01-01

    Prototype autonomous deicing system for airplane includes network of electronic and electromechanical modules at various locations in wings and connected to central data-processing unit. Small, integrated solid-state device, using long coils installed under leading edge, exciting small vibrations to detect ice and larger vibrations to knock ice off. In extension of concept, outputs of vibration sensors and other sensors used to detect rivet-line fractures, fatigue cracks, and other potentially dangerous defects.

  7. 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.

  8. Airplane Wings For Faster Climbing And Slower Landing

    NASA Technical Reports Server (NTRS)

    Cliff-Hovey, Susan

    1987-01-01

    Reshaped airfoils improve performance. Performances of general-aviation airplanes improved by modifying airfoil shapes. Equation used to determine new contour for each type of wing. Calculations straightforward enough to be done on hand calculator.

  9. 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.

  10. 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.

  11. 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.

  12. A possible method for preventing the autorotation of airplane wings

    NASA Technical Reports Server (NTRS)

    Schrenk, Oskar

    1930-01-01

    At the suggestion of Professor Betz, the following device was tested with the object of reducing the autorotational speed of airplane wings. The model of a normal wing with the Gottigen profile 420, 1 meter span and 0.2 meter chord was provided with a pair of symmetrical slots on the suction side, connected with each other inside the wing. The arrangement of the testing equipment and models are given and the effect of the slots can be seen in the experimental curves that are included.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. The Choice of Wing Sections for Airplanes

    NASA Technical Reports Server (NTRS)

    Warner, Edward P

    1921-01-01

    The subject of the choice of an airfoil section is by no means a closed one, and despite the impossibility of making a single rule serve, it is quite practicable to deduce in a strictly rational manner a series of rules and formulas which are capable of being of the greatest use if we but confine ourselves to the consideration of one element of performance at a time. There are seven such elements of performance which are here taken up in turn. The seven are of different relative importance in different types of airplanes. The seven elements are: maximum speed regardless of minimum; maximum speed for given minimum; maximum speed range ratio; maximum rate of climb; maximum absolute ceiling; maximum distance non-stop; and maximum duration non-stop.

  19. 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.

  20. 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.

  1. 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.

  2. Trapping and identifying single-nanoparticles using a low-aspect-ratio nanopore

    NASA Astrophysics Data System (ADS)

    Tsutsui, Makusu; Maeda, Yoichi; He, Yuhui; Hongo, Sadato; Ryuzaki, Sou; Kawano, Satoyuki; Kawai, Tomoji; Taniguchi, Masateru

    2013-07-01

    Manipulation of particles and molecules in fluid is a fundamental technology in biosensors. Here, we report electrical trapping and identification of single-nanoparticles using a low-aspect-ratio nanopore. Particle trapping and detrapping are implemented through a control of the cross-membrane electrophoretic voltage. This electrical method is found to enable placing an individual nanoparticle in vicinity of a lithographically-defined nanopore by virtue of the balance between the two counteracting factors, electrostatic and electroosmotic forces. We also demonstrate identification of trapped nanoparticles by the ionic current through the particle-pore gap space. This technique may find applications in electrode-embedded nanopore sensors.

  3. 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.

  4. 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.

  5. 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.

  6. 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

  7. Wake topology and hydrodynamic performance of low-aspect-ratio flapping foils

    NASA Astrophysics Data System (ADS)

    Dong, H.; Mittal, R.; Najjar, F. M.

    2006-11-01

    Numerical simulations are used to investigate the effect of aspect ratio on the wake topology and hydrodynamic performance of thin ellipsoidal flapping foils. The study is motivated by the quest to understand the hydrodynamics of fish pectoral fins. The simulations employ an immersed boundary method that allows us to simulate flows with complex moving boundaries on fixed Cartesian grids. A detailed analysis of the vortex topology shows that the wake of low-aspect-ratio flapping foils is dominated by two sets of interconnected vortex loops that evolve into distinct vortex rings as they convect downstream. The flow downstream of these flapping foils is characterized by two oblique jets and the implications of this characteristic on the hydrodynamic performance are examined. Simulations are also used to examine the thrust and propulsive efficiency of these foils over a range of Strouhal and Reynolds numbers as well as pitch-bias angles.

  8. Exploration of low-aspect-ratio tokamak regimes in CDX-U

    SciTech Connect

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

    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. Using a relatively low toroidal magnetic field, plasma discharges with I{sub p} {le}53 kA, and q(a) {ge} 4 [qcyl(a) {ge}1] have been obtained. Low q(a), ohmic plasmas in CDX-U show stronger MHD activity as the edge safety factor is reduced. Those MHD modes appear to reduce the current ramp-up rate and, at present, limit the access to even lower q(a) regimes.

  9. Global stability study of the ultra low aspect ratio tokamak, ULART

    SciTech Connect

    Yamada, M.; Pomphrey, N.; Morita, A.; Ono, Y.; Katsurai, M.

    1996-02-01

    By introducing a slender current carrying conductor through the geometric center axis of the Tokyo University TS-3 device, the authors have generated ultra low aspect ratio tokamak (ULART) configurations with aspect ratio as low as 1.05. In this extreme limit they study the transition of the spheromak (q{sub edge} = 0, I{sub tf} = 0) to a ULART plasma (q{sub edge} = 2--20). The authors investigate the global MHD characteristics of ULART by comparing theoretical results with the experimental data obtained. A small current in the center conductor (compared with the plasma current) is found to significantly improve the overall MHD stability characteristics of the formed plasmas by effectively stabilizing the global tilt/shift mode. Theoretical calculations of the threshold toroidal field current required for stability and the growth rates of the tilt/shift modes agree well with the TS-3 data.

  10. Transmutation of nuclear waste with a low-aspect-ratio tokamak neutron source

    NASA Astrophysics Data System (ADS)

    Hong, Bong Guen; Moon, Se Youn

    2014-10-01

    The transmutation characteristics of transuranics (TRUs) in a transmutation reactor based on a LAR (Low-aspect-ratio) tokamak as a neutron source are investigated. The optimum radial build of a transmutation reactor is found by using a coupled analysis of the tokamak systems and the neutron transport. The dependences of the transmutation characteristics on the aspect ratio A in the range of 1.5 to 2.5 and on the fusion power in the range of 150 to 500 MW are investigated. An equilibrium fuel cycle is developed for effective transmutation, and show that with one unit of the transmutation reactor based on the LAR tokamak producing fusion power in the range of a few hundred MWs, up to 3 PWRs (1.0 GWe capacity) can be supported with a burn-up fraction larger than 50%.

  11. MAST and the impact of low aspect ratio on tokamak physics

    NASA Astrophysics Data System (ADS)

    Lloyd, B.; Ahn, J.-W.; Akers, R. J.; Appel, L. C.; Applegate, D.; Axon, K. B.; Baranov, Y.; Brickley, C.; Bunting, C.; Buttery, R. J.; Carolan, P. G.; Challis, C.; Ciric, D.; Conway, N. J.; Cox, M.; Counsell, G. F.; Cunningham, G.; Darke, A.; Dnestrovskij, A.; Dowling, J.; Dudson, B.; Dunstan, M. R.; Field, A. R.; Gee, S.; Gryaznevich, M. P.; Helander, P.; Hender, T. C.; Hole, M.; Joiner, N.; Keeling, D.; Kirk, A.; Lehane, I. P.; Lott, F.; Maddison, G. P.; Manhood, S. J.; Martin, R.; McArdle, G. J.; McClements, K. G.; Meyer, H.; Morris, A. W.; Nelson, M.; O'Brien, M. R.; Patel, A.; Pinfold, T.; Preinhaelter, J.; Price, M. N.; Roach, C. M.; Rozhansky, V.; Saarelma, S.; Saveliev, A.; Scannell, R.; Sharapov, S.; Shevchenko, V.; Shibaev, S.; Stammers, K.; Storrs, J.; Sykes, A.; Tabasso, A.; Taylor, D.; Tournianski, M. R.; Turner, A.; Turri, G.; Valovic, M.; Volpe, F.; Voss, G.; Walsh, M. J.; Watkins, J. R.; Wilson, H. R.; Wisse, M.; MAST, the; NBI; ECRH Teams

    2004-12-01

    Low aspect ratio plasmas in devices such as the mega ampere spherical tokamak (MAST) are characterized by strong toroidicity, strong shaping and self fields, low magnetic field, high beta, large plasma flow and high intrinsic E × B flow shear. These characteristics have important effects on plasma behaviour, provide a stringent test of theories and scaling laws and offer new insight into underlying physical processes, often through the amplification of effects present in conventional tokamaks (e.g. impact of fuelling source and magnetic geometry on H-mode access). The enhancement of neoclassical effects makes MAST ideal for the study of particle pinch processes and neoclassical resistivity corrections, which can be assessed with unique accuracy. MAST data have an important influence on scaling laws for confinement and H-mode threshold power, exerting strong leverage on the form of these scaling laws (e.g. scaling with aspect ratio, beta, magnetic field, etc). The high intrinsic flow shear is conducive to transport barrier formation by turbulence suppression. Internal transport barriers are readily formed in MAST with both co- and counter-NBI, and electron and ion thermal diffusivities have been reduced to the ion neoclassical level. The strong variation in toroidal field (~ × 5 in MAST) between the inboard and outboard plasma edges, provides a useful test of edge models prompting, for example, a comparison of inboard and outboard scrape-off-layer transport to highlight magnetic field effects. Low aspect ratio plasmas are also an ideal testing ground for plasma instabilities, such as neoclassical tearing modes, edge localized modes (ELMs) and Alfvén eigenmodes, which are readily generated due to the supra-Alfvénic ion population. Examples of how MAST is providing new insights into such instabilities (e.g. ELM structure) are described.

  12. 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.

  13. 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)

  14. 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.

  15. Effects of airplane flexibility on wing bending strains in rough air

    NASA Technical Reports Server (NTRS)

    Coleman, Thomas L; Press, Harry; Shufflebarger, C C

    1957-01-01

    Some results on the effects of wing flexibility on wing bending strains as determined from flight tests of a Boeing B-29 and a Boeing B-47A airplane in rough air are presented. Results from an analytical study of the flexibility effects on the B-29 wing strains are compared with the experimental results. Both the experimental and calculated results are presented as frequency-response functions of the bending strains at various spanwise wing stations to gust disturbances. In addition, some indirect evidence of the effect of spanwise variations in turbulence on the response of the B-47A airplane is presented.

  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. 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.

  18. 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.

  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. Proposal for a risk banding framework for inhaled low aspect ratio nanoparticles based on physicochemical properties.

    PubMed

    Oosterwijk, Mattheus T T; Feber, Maaike Le; Burello, Enrico

    2016-08-01

    We present a conceptual framework that can be used to assign risk bands to inhaled low aspect ratio nanoparticles starting from exposure bands assigned to a specific exposure situation. The framework mimics a basic physiological scheme that captures the essential mechanisms of fate and toxicity of inhaled nanoparticles and is composed of several models and rules that estimate the result of the following processes: the deposition of particles in the respiratory tract, their (de-)agglomeration, lung burden and clearance, their diffusion through the lung mucus layer, translocation and cellular uptake and local and systemic toxicity. Each model is based on a set of particle's physicochemical properties, including the size and size distribution(s), the zeta potential (or net charge at a specific pH), the surface hydrophobicity or hydrophilicity, the conduction band energy (for metals, metal oxides, quantum dots, etc.) and the solubility at a specific pH. The framework takes the exposure bands as input and predicts, using the above-mentioned models, an internal dose band (module 1). Module 2 assigns a relative hazard ranking depending on the region of particle deposition in the respiratory tract, the likelihood of uptake and whether the toxicological effects are assumed to be local and/or systemic. By combining the results of Module 1 and 2, the framework provides a relative risk ranking. PMID:26763369

  1. Self-consistent Equilibrium Model of Low-aspect-ratio Toroidal Plasma with Energetic Beam Ions

    SciTech Connect

    E.V. Belova; N.N. Gorelenkov; C.Z. Cheng

    2003-04-09

    A theoretical model is developed which allows the self-consistent inclusion of the effects of energetic beam ions in equilibrium calculations of low-aspect-ratio toroidal devices. A two-component plasma is considered, where the energetic ions are treated using a kinetic Vlasov description, while a one-fluid magnetohydrodynamic description is used to represent the thermal plasma. The model allows for an anisotropic distribution function and a large Larmor radius of the beam ions. Numerical results are obtained for neutral-beam-heated plasmas in the National Spherical Torus Experiment (NSTX). Self-consistent equilibria with an anisotropic fast-ion distribution have been calculated for NSTX. It is shown for typical experimental parameters that the contribution of the energetic neutral-beam ions to the total current can be comparable to that of the background plasma, and that the kinetic modifications of the equilibrium can be significant. The range of validity of the finite-Larmor-radius expansion and of the reduced kinetic descriptions for the beam ions in NSTX is discussed. The calculated kinetic equilibria can be used for self-consistent numerical studies of beam-ion-driven instabilities in NSTX.

  2. Proposal for a risk banding framework for inhaled low aspect ratio nanoparticles based on physicochemical properties.

    PubMed

    Oosterwijk, Mattheus T T; Feber, Maaike Le; Burello, Enrico

    2016-08-01

    We present a conceptual framework that can be used to assign risk bands to inhaled low aspect ratio nanoparticles starting from exposure bands assigned to a specific exposure situation. The framework mimics a basic physiological scheme that captures the essential mechanisms of fate and toxicity of inhaled nanoparticles and is composed of several models and rules that estimate the result of the following processes: the deposition of particles in the respiratory tract, their (de-)agglomeration, lung burden and clearance, their diffusion through the lung mucus layer, translocation and cellular uptake and local and systemic toxicity. Each model is based on a set of particle's physicochemical properties, including the size and size distribution(s), the zeta potential (or net charge at a specific pH), the surface hydrophobicity or hydrophilicity, the conduction band energy (for metals, metal oxides, quantum dots, etc.) and the solubility at a specific pH. The framework takes the exposure bands as input and predicts, using the above-mentioned models, an internal dose band (module 1). Module 2 assigns a relative hazard ranking depending on the region of particle deposition in the respiratory tract, the likelihood of uptake and whether the toxicological effects are assumed to be local and/or systemic. By combining the results of Module 1 and 2, the framework provides a relative risk ranking.

  3. 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.

  4. Trapped Electron Stabilization of Ballooning Modes in Low Aspect Ratio Toroidal Plasmas

    SciTech Connect

    C.Z. Cheng and N.N. Gorelenkov

    2004-03-18

    The kinetic effects of trapped electron dynamics and finite gyroradii and magnetic drift motion of ions are shown to give rise to a large parallel electric field and hence a parallel current that greatly enhances the stabilizing effect of field line tension for ballooning modes in low aspect ratio toroidal plasmas. For large aspect ratio the stabilizing effect increases (reduces) the {beta}(= 2P/B{sup 2}) threshold for the first (second) stability of the kinetic ballooning mode (KBM) from the MHD {beta} threshold value by a factor proportional to the trapped electron density fraction. For small aspect ratio the stabilizing effect can greatly increase the {beta} threshold of the first stability of KBMs from the MHD {beta} threshold by S{sub c} {approx_equal} 1 + (n{sub e}/n{sub eu}){delta}, where n{sub e}/n{sub eu} is the ratio of the total electron density to the untrapped electron density, and {delta} depends on the trapped electron dynamics and finite gyroradii and magnetic drift motion of ions. If n{sub e}/n{sub eu} >> 1 as in the National Spherical Torus Experiment (NSTX) with an aspect ratio approximately equal to 1.4, the KBM should be stable for {beta} {le} 1 for finite magnetic shear. Therefore, unstable KBMs are expected only in the weak shear region near the radial location of the minimum of the safety factor in NSTX reverse shear discharges.

  5. 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.

  6. On the flow generated by rotating flat plates of low aspect ratio

    NASA Astrophysics Data System (ADS)

    DeVoria, Adam C.

    Low-aspect-ratio propulsors typically allow for high maneuverability at low-to-moderate speeds. This has made them the subject of much recent research aimed at employing such appendages on autonomous vehicles which are required to navigate tumultuous environments. This experimental investigation focuses on the fluid dynamic aspects associated with overly-simplified versions of such biologically-inspired propulsors. In doing so, fundamental contributions are made to the research area. The unsteady, three-dimensional flow of a low-aspect-ratio, trapezoidal flat plate undergoing rotation from rest at a 90° angle of attack and Reynolds numbers of O(103) is investigated experimentally. The objectives are to develop a straightforward protocol for vortex saturation, and to understand the effects of the root-to-tip flow for different velocity programs. The experiments are conducted in a glass-walled tank, and digital particle image velocimetry is used to obtain planar velocity measurements. A formation-parameter definition is investigated and is found to reasonably predict the state corresponding to the pinch-off of the initial tip vortex across the velocity programs tested. The flow in the region near the tip is relatively insensitive to Reynolds number over the range studied. The component normal to the plate is unaffected by total rotational amplitude while the tangential component has dependence on this angle. Also, an estimate of the first tip-vortex pinch-off time is obtained from the near-tip velocity data and agrees very well with values estimated using circulation. The angle of incidence of the bulk root-to-tip flow relative to the plate normal becomes more oblique with increasing rotational amplitude. Accordingly, the peak magnitude of the tangential velocity is also increased and as a result advects fluid momentum away from the plate at a higher rate. The more oblique impingement of the root-to-tip flow for increasing rotational amplitude is shown to have a

  7. 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 .

  8. 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.

  9. 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); 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)

    2016-08-05

    This public data set contains openly-documented, machine readable digital research data accompanying 'H-mode Plasmas at Very Low Aspect Ratio on the Pegasus Toroidal Experiment' by K.E. Thome et al., accepted for publication in Nuclear Fusion.

  10. 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.

  11. An analysis of available data on effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes

    NASA Technical Reports Server (NTRS)

    Wollner, Bertram C

    1949-01-01

    Available information on the effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes is analyzed. The effects of wing and nacelle incidence, horizontal andvertical position of wing and nacelle, fuselage shape, wing section and filleting are considered. Where sufficient data were unavailable to determine the distribution of the air load, the change in lift caused by interference between wing and fuselage was found. This increment is affected to the greatest extent by vertical wing position.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. Extension of the beam theory for polymer bio-transducers with low aspect ratios and viscoelastic characteristics

    NASA Astrophysics Data System (ADS)

    Du, Ping; Lin, I.-Kuan; Lu, Hongbing; Zhang, Xin

    2010-09-01

    Polydimethylsiloxane (PDMS)-based micropillars (or microcantilevers) have been used as bio-transducers for measuring cellular forces on the order of pN to µN. The measurement accuracy of these sensitive devices depends on appropriate modeling to convert the micropillar deformations into the corresponding reaction forces. The traditional approach to calculating the reaction force is based on the Euler beam theory with consideration of a linear elastic slender beam for the micropillar. However, the low aspect ratio in geometry of PDMS micropillars does not satisfy the slender beam requirement. Consequently, the Timoshenko beam theory, appropriate for a beam with a low aspect ratio, should be used. In addition, the inherently time-dependent behavior in PDMS has to be considered for accurate force conversion. In this paper, the Timoshenko beam theory, along with the consideration of viscoelastic behavior of PDMS, was used to model the mechanical response of micropillars. The viscoelastic behavior of PDMS was characterized by stress relaxation nanoindentation using a circular flat punch. A correction procedure was developed to determine the load-displacement relationship with consideration of ramp loading. The relaxation function was extracted and described by a generalized Maxwell model. The bending of rectangular micropillars was performed by a wedge indenter tip. The viscoelastic Timoshenko beam formula was used to calculate the mechanical response of the micropillar, and the results were compared with measurement data. The calculated reaction forces agreed well with the experimental data at three different loading rates. A parametric study was conducted to evaluate the accuracy of the viscoelastic Timoshenko beam model by comparing the reaction forces calculated from the elastic Euler beam, elastic Timoshenko beam and viscoelastic Euler beam models at various aspect ratios and loading rates. The extension of modeling from the elastic Euler beam theory to the

  2. 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.

  3. 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.

  4. 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.

  5. The Effect of Relative Submergence and Shape on the Wake of a Low-Aspect-Ratio Wall-Mounted body

    NASA Astrophysics Data System (ADS)

    Hajimirzaie, Seyed Mohammad; Wojcik, Craig; Buchholz, James

    2010-11-01

    Wall-mounted bodies in boundary layer flows are ubiquitous in nature and engineering applications and significantly enhance momentum and scalar transport in their vicinity. In this experimental study we evaluate the role of relative submergence (the ratio of flow depth to obstacle height) and shape on the wakes around four different wall-mounted obstacles. We consider four obstacle geometries: semi-ellipsoids with the major and minor axes of the base ellipses aligned in the streamwise and transverse directions, and two cylinders with matching aspect ratios D/H (where D is the maximum transverse dimension and H is the obstacle height). The aspect ratios considered are 0.67 and 0.89. DPIV was used to interrogate the flow. Streamwise structures observed in the mean wake include tip, base, and horseshoe vortex pairs as well as additional structures apparently not previously observed. The presence of a base vortex for such low-aspect-ratio obstacles is unexpected, and its strength increases with decreasing relative submergence. We will discuss hypotheses on the mechanisms of generation of the base and tertiary structures and their interconnection with the rest of the vortex skeleton.

  6. Flight measurements of buffet characteristics of the F-104 airplane for selected wing-flap deflections

    NASA Technical Reports Server (NTRS)

    Friend, E. L.; Sefic, W. J.

    1972-01-01

    A flight program was conducted on the F-104 airplane to investigate the effects of moderate deflections of wing leading- and trailing-edge flaps on the buffet characteristics at subsonic and transonic Mach numbers. Selected deflections of the wing leading and trailing-edge flaps, individually and in combination, were used to assess buffet onset, intensity, and frequency; lift curves; and wing-rock characteristics for each configuration. Increased deflection of the trailing-edge flap delayed the buffet onset and buffet intensity rise to a significantly higher airplane normal-force coefficient. Deflection of the leading-edge flap produced some delay in buffet onset and the resulting intensity rise at low subsonic speeds. Increased deflection of the trailing-edge flap provided appreciable lift increments in the angle-of-attack range covered, whereas the leading-edge flap provided lift increments only at high angles-of-attack. The pilots appreciated the increased maneuvering envelope provided by the flaps because of the improved turn capability.

  7. 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.

  8. 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

    2015-04-01

    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

  9. 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

    2015-04-01

    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.

  10. 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

  11. Summary of Results Obtained in Full-Scale Tunnel Investigation of the Ryan Flex-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Johnson, Joseph L., Jr.; Hassell, James L., Jr.

    1962-01-01

    The performance and static stability and control characteristics of the Ryan Flex-Wing airplane were determined in an investigation conducted in the Langley full-scale tunnel through an angle-of-attack range of the keel from about 14 to 44 deg. for power-on and -off conditions. Comparisons of the wind-tunnel data with flight-test data obtained with the same airplane by the Ryan Aeronautical Company were made in a number of cases.

  12. 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.

  13. 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.

  14. 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.

  15. Observation of neoclassical impurity transport in Ohmically heated plasmas of CDX-U low aspect ratio tokamak

    NASA Astrophysics Data System (ADS)

    Soukhanovskii, V. A.; Finkenthal, M.; Moos, H. W.; Stutman, D.; Munsat, T.; Jones, B.; Hoffman, D.; Kaita, R.; Majeski, R.

    2002-11-01

    High β, good confinement and stability properties of the low aspect ratio tokamaks, or spherical tori (ST), have been predicted theoretically and preliminarily confirmed in several large experiments recently. This paper reports on impurity transport experiments carried out in ohmically heated plasmas of the small spherical torus CDX-U with the aspect ratio of A≃1.5. Vacuum ultraviolet and soft x-ray multichannel spectroscopic diagnostics are used to measure intrinsic carbon, oxygen and radiated power radial brightness profiles in plasmas with Te(0)≃60-80 eV and ne(0)≃2×1013 cm-3. The measurements are performed in both magnetohydrodynamically dominated and quiescent phase of the plasmas. The properties of the observed low m/n modes, sawtooth oscillations, and ST-specific reconnection events are discussed in the context of particle transport. The measured impurity profiles are modelled using one-dimensional impurity transport code MIST and a collisional-radiative package CRMLIN. Impurity diffusion of 0.2 m2 s-1≤D≤0.6 m2 s-1 and convection velocity of v≃4-6 m s-1 are inferred from the modelling. These transport coefficients are very close to the neoclassical theory predictions obtained with the FORCEBAL code, which uses analytical plasma viscosity expressions valid for an arbitrary aspect ratio geometry. Neoclassical analysis indicates that both carbon and oxygen are in the collisional regime, and the Pfirsch-Schluter flux is the major fraction of the impurity flux. The causes of the observed strong non-diffusive transport are discussed, and it is concluded that the ∇ni/ni term, resulting from highly peaked ion density profile, makes the largest contribution to the inward pinch. Present analysis suggests that drift wave turbulence is reduced in CDX-U ohmically heated discharges within at least r/a≤0.4, however more refined measurements are needed to interpret the results in the framework of ST ion transport.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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

  2. 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.

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

    NASA Technical Reports Server (NTRS)

    Johnson, J. Blair; Sandlin, Doral R.

    1986-01-01

    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 called wing-body. Pressure distributions generated by the codes for the flight test fixture, as well as compared with the flight-measured data. The two-dimensional code pressure distributions compared well except at the minimum pressure point and the trailing edge. Shock locations compared well except at high transonic speeds. However, the two-dimensional code did not adequately predict the displacement thickness of the flight test fixture. The three-dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture.

  4. Whirling Arm Tests on the Effect of Ground Proximity to an Airplane Wing

    NASA Technical Reports Server (NTRS)

    Long, M. E.

    1944-01-01

    This report gives the results of tests on a rectangular wing model with a 20% full spun split flap, conducted on the whirling arm at the Daniel Guggenheim Airship Institute in Akron, Ohio. The effect of a ground board on the lift and pitching moment was measured. The ground board consisted of an inclined ramp rising up in the test channel to a level floor extending for some distance parallel to the model path. The path of the wing model with respect to the ground board accordingly represented with comparative exactness an airplane coming in for a landing. The ground clearances over the level portion of the board varied from 0 6 to 1,6 chord lengths. Results are given in the standard dimensionless coefficients plotted versus angle of attack for a particular ground clearance. The effect of the ground board is to increase the lift coefficient for a given angle of attack all the way up the stall. The magnitude of the increase varies both with the ground clearance and the angle of attack. The effect on the pitching moment coefficient is not so readily apparent due to experimental difficulties but, in general, the diving moment increases over the ground board. This effect is apparent principally at the high angles of attack. An exception to this effect occurs with flaps deflected at the lowest ground clearance (0.6 chords). Here the diving moment decreases over the ground board.

  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. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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)

  12. 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.

  13. Flight-measured buffet characteristics of a supercritical wing and a conventional wing on a variable-sweep airplane

    NASA Technical Reports Server (NTRS)

    Monaghan, R. C.

    1978-01-01

    Windup-turn maneuvers were performed to assess the buffet characteristics of the F-111A aircraft and the same aircraft with a supercritical wing, which is referred to as the F-111 transonic aircraft technology (TACT) aircraft. Data were gathered at wing sweep angles of 26, 35, and 58 deg for Mach numbers from 0.60 to 0.95. Wingtip accelerometer data were the primary source of buffet information. The analysis was supported by wing strain-gage and pressure data taken in flight, and by oil-flow photographs taken during tests of a wind tunnel model. In the transonic speed range, the overall buffet characteristics of the aircraft having a supercritical wing are significantly improved over those of the aircraft having a conventional wing.

  14. 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.

  15. 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.

  16. 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.

  17. 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

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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%.

  3. Some studies on the aerodynamic effect of the gap between airplane wings and fuselages

    NASA Technical Reports Server (NTRS)

    Ober, Shatswell

    1929-01-01

    The general result indicated by this study is that if desirable from any viewpoint the gap between wing and fuselage may be closed without detrimental aerodynamic effects, and with a given monoplane there is less drag if the wing is directly on top of the fuselage than if it is parasol.

  4. 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.

  5. Dynamic Longitudinal and Directional Stability Derivatives for a 45 deg. Sweptback-Wing Airplane Model at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Bielat, Ralph P.; Wiley, Harleth G.

    1959-01-01

    An investigation was made at transonic speeds to determine some of the dynamic stability derivatives of a 45 deg. sweptback-wing airplane model. The model was sting mounted and was rigidly forced to perform a single-degree-of-freedom angular oscillation in pitch or yaw of +/- 2 deg. The investigation was made for angles of attack alpha, from -4 deg. to 14 deg. throughout most of the transonic speed range for values of reduced-frequency parameter from 0.015 to 0.040 based on wing mean aerodynamic chord and from 0.04 to 0.14 based on wing span. The results show that reduced frequency had only a small effect on the damping-in-pitch derivative and the oscillatory longitudinal stability derivative for all Mach numbers M and angles of attack with the exception of the values of damping coefficient near M = 1.03 and alpha = 8 deg. to 14 deg. In this region, the damping coefficient changed rapidly with reduced frequency and negative values of damping coefficient were measured at low values of reduced frequency. This abrupt variation of pitch damping with reduced frequency was a characteristic of the complete model or wing-body-vertical-tail combination. The damping-in-pitch derivative varied considerably with alpha and M for the horizontal-tail-on and horizontal-tail-off configurations, and the damping was relatively high at angles of attack corresponding to the onset of pitch-up for both configurations. The damping-in-yaw derivative was generally independent of reduced frequency and M at alpha = -4 deg. to 4 deg. At alpha = 8 deg. to 14 deg., the damping derivative increased with an increase in reduced frequency and alpha for the configurations having the wing, whereas the damping derivative was either independent of or decreased with increase in reduced frequency for the configuration without the wing. The oscillatory directional stability derivative for all configurations generally decreased with an increase in the reduced-frequency parameter, and, in some instances

  6. 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

  7. 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.

  8. 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)

  9. Effect of the Proximity of the Wing First-Bending Frequency and the Short-Period Frequency on the Airplane Dynamic-Response Factor

    NASA Technical Reports Server (NTRS)

    Huss, Carl R.; Donegan, James J.

    1959-01-01

    The results are presented in the form of preliminary design charts which give a comparison between the dynamic-response factors of the semi-rigid case and the airplane longitudinal short-period case and between the dynamic-response factors of the semi-rigid case and the steady-state value of the airplane longitudinal short-period response. These charts can be used to estimate the first-order effects of the addition of a wing-bending degree of freedom on the short-period dynamic-response factor and on the maximum dynamic-response factor when compared with the steady-state response of the system.

  10. 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.

  11. Flight tests of a radio-controlled airplane mode with a free-wing, free-canard configuration

    NASA Technical Reports Server (NTRS)

    Gee, S. W.

    1978-01-01

    Flight characteristics, controllability, and potential operating problems were investigated in a radio-controlled airplane model in which the wing is so attached to the fuselage that it is free to pivot about a spanwise axis forward of its aerodynamic center and is subject only to aerodynamic pitching moments imposed by lift and drag forces and a control surface. A simple technique of flying the test vehicle in formation with a pickup truck was used to obtain trim data. The test vehicle was flown through a series of maneuvers designed to permit evaluation of certain characteristics by observation. The free-wing free-canard concept was determined to be workable. Stall/spin characteristics were considered to be excellent, and no effect on longitudinal stability was observed when center of gravity changes were made. Several problems were encountered during the early stages of flight testing, such as aerodynamic lockup of the free canard and excessive control sensitivity. Lack of onboard instrumentation precluded any conclusions about gust alleviation or ride qualities.

  12. Aerodynamic characteristics of an 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)

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

    1972-01-01

    Transonic pressure tunnel tests at Mach numbers from 0.25 to 1.00 were performed to determine the effects of area-rule additions to the sides of the fuselage on the aerodynamic characteristics of a 0.087 scale model of an NASA supercritical-wing research airplane. Presented are the longitudinal aerodynamic force and moment characteristics for horizontal-tail deflection angles of -2.5 deg and -5 deg with the side fuselage area-rule additions on and off the model. The effects of the side fuselage area-rule additions on selected wing and fuselage pressure distributions at near-cruise conditions are also presented.

  13. Increase in the maximum lift of an airplane wing due to a sudden increase in its effective angle of attack resulting from a gust

    NASA Technical Reports Server (NTRS)

    Kramer, Max

    1932-01-01

    Wind-tunnel tests are described, in which the angle of attack of a wing model was suddenly increased (producing the effect of a vertical gust) and the resulting forces were measured. It was found that the maximum lift coefficient increases in proportion to the rate of increase in the angle of attack. This fact is important for the determination of the gust stresses of airplanes with low wing loading. The results of the calculation of the corrective factor are given for a high-performance glider and a light sport plane of conventional type.

  14. 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.

  15. Transonic flutter and gust-response tests and analyses of a wind-tunnel model of a torsion free wing airplane

    NASA Technical Reports Server (NTRS)

    Ruhlin, C. L.; Murphy, A. C.

    1981-01-01

    An exploratory study of a 1/5.5 size, complete airplane version of a torsion free wing (TFW) fighter aircraft was conducted. The TFW consisted of a wing/boom/canard assembly on each fuselage side that was interconnected by a common pivot shaft so that the TFW could rotate freely in pitch. The effect of the TFW was evaluated by comparing data obtained with the TFW free and the TFW locked to the fuselage. With the model mounted on cables to simulate an airplane free flying condition, flutter boundaries were measured at Mach number (M) from 0.85 to 1.0 and gust responses at M = 0.65 and 0.90. The critical flutter mode for the TFW free configuration was found experimentally to occur at M = 0.95 and had the rigid TFW pitch mode as its apparent aerodynamic driver.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. The Static-Pressure Error of a Wing Airspeed Installation of the McDonnell XF-88 Airplane in Dives to Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Goodman, Harold R.

    1949-01-01

    Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35 deg along the 0.22 chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8-percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Results of the measurements indicate that the static-pressure error, within the accuracy of measurement, is negligible from a Mach number of 0.65 to a Mach number of about 0.97. With a further increase in Mach number, the static-pressure error increases rapidly; at the highest Mach number attained in these tests (about M = 1.038), the error increases to about 8 percent of the impact pressure. Above a Mach number of about 0.975, the recorded Mach number remains substantially constant with increasing true Mach number; the installation is of no value between a Mach number of about 0.975 and at least 1.038, as the true Mach number cannot be obtained from the recorded Mach number in this range. Previously published data have shown that at 0.96 chord ahead of the wing tip of the straight-wing X-l airplanes, a rapid rise of position error started at a Mach number of about 0.8. In the case of the XF-88 airplane, this rise of position error was delayed, presumably by the sweep of the wing, to a Mach number of about 0.97.

  3. 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.

  4. 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.

  5. Stability and Control Characteristics of a Complete Airplane Model Having a Wing with Quarter-chord Line Swept Back 40 Degrees, Aspect Ratio 2.50, and Taper Ratio 0.42

    NASA Technical Reports Server (NTRS)

    Schulderfrei, Marvin; Comisarow, Paul; Goodson, Kenneth W

    1951-01-01

    An investigation has been made of a complete airplane model having a wing with the quarter-chord line swept back 40 degrees, aspect ratio 2.50, and taper ratio 0.42 to determine its low-speed stability and control characteristics. The longitudinal stability investigation included stabilizer and tail-off tests with different wing dihedral angles (Gamma = 0 degrees and Gamma = -10 degrees) over an angle-of-attack range for the cruising and landing configurations and tests. with a high horizontal-tail location (Gamma = -10 degrees) for the cruising configuration. Tests were made of the wing alone and to determine the effect of wing end plates in pitch. Lateral stability characteristics were determined for the airplane with different geometric wing dihedrals, with end plates, and with several dorsal modifications. Tests were made with ailerons and spoilers to determine control characteristics.

  6. 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.

  7. Inlet Noise Reduction by Shielding for the Blended-Wing-Body Airplane

    NASA Technical Reports Server (NTRS)

    Clark, Lorenzo R.; Gerhold, Carl H.

    1999-01-01

    Noise shielding benefits associated with an advanced unconventional subsonic transport concept, the Blended-Wing-Body, were studied using a 4- percent scale, 3-engine nacelle model. The study was conducted in the Anechoic Noise Research Facility at NASA Langley Research Center. A high- frequency, wideband point source was placed inside the nacelles of the center engine and one of the side engines in order to simulate broadband engine noise. The sound field of the model was measured with a rotating microphone array that was moved to various stations along the model axis and with a fixed array of microphones that was erected behind the model. Ten rotating microphones were traversed a total of 22 degrees in 2-degree increments. Seven fixed microphones covered an arc that extended from a point in the exhaust exit plane of the center engine (and directly below its centerline) to a point 30 degrees above the jet centerline. While no attempt was made to simulate the noise emission characteristics of an aircraft engine, the model source was intended to radiate sound in a frequency range encompassing 1, 2, and 3 times the blade passage of a typical full-scale engine. In this study, the Blended-Wing-Body model was found to provide significant shielding of inlet noise. In particular, noise radiated downward into the forward sector was reduced by 20 to 25 dB overall in the full-scale frequencies from 2000 to 4000 Hz, decreasing to 10 dB or less at the lower frequencies. Also, it was observed that noise associated with the exhaust radiates into the sector directly below the model downstream to reduce shielding efficiency.

  8. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  10. 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.

  11. 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.

  12. The Effect of the Wings of Single Engine Airplanes on Propulsive Efficiency as Shown by Full Scale Wind Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Wood, Donald H

    1929-01-01

    An investigation was conducted to determine the effect of the wings on propulsive efficiency. The wings are shown to cause a reduction of 1 percent to 3 percent in propulsive efficiency, which is about the same for monoplane as well as biplane wings.

  13. 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.

  14. An Analysis of Flight-Test Measurements of the Wing Structural Deformations in Rough Air of a Large Flexible Swept-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Murrow, Harold N.

    1959-01-01

    An analysis is made of wing deflection and streamwise twist measurements in rough-air flight of a large flexible swept-wing bomber. Random-process techniques are employed in analyzing the data in order to describe the magnitude and characteristics of the wing deflection and twist responses to rough air. Power spectra and frequency-response functions for the wing deflection and twist responses at several spanwise stations are presented. The frequency-response functions describe direct and absolute response characteristics to turbulence and provide a convenient basis for assessing analytic calculation techniques. The wing deformations in rough air are compared with the expected deformations for quasi-static loadings of the same magnitude, and the amplifications are determined. The results obtained indicate that generally the deflections are amplified by a small amount, while the streamwise twists are amplified by factors of the order of 2.0. The magnitudes of both the deflection velocities and the twist angles are shown to have significant effects on the local angles of attack at the various stations and provide the main source of aerodynamic loading, particularly at frequencies in the vicinity of the first wing-vibration mode.

  15. 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.

  16. 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)

  17. 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.

  18. A study of the accuracy of wing calculations based on different schemes

    NASA Astrophysics Data System (ADS)

    Putilin, S. I.; Savchenko, V. T.

    Solutions obtained for the same flow problem using different vortex and computation point distribution laws are presented. The problem considered here is the plane problem for a plate near a circular contour, which is relevant to the design of wings with optimal load distribution and low aspect ratio wings in a bounded fluid. Ways to improve the accuracy of the results are discussed.

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

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  20. Flight investigation of stall, spin and recovery characteristics of a low-wing, single-engine, T-tail light airplane

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

    Flight tests were performed to investigate the stall, spin, and recovery characteristics of a four-place, low-wing, single-engine, T-tail, general aviation research airplane at an aft center-of-gravity position. Most stalls resulted in roll-offs. Spins were oscillatory in roll and pitch at 43 deg angle of attack; the magnitude of the oscillations was determined by aileron position. Power, flap deflection, and landing gear position did not affect the angle of attack to the spin. Antispin rudder followed by forward wheel with ailerons neutral produced the fastest and most consistent recoveries but the initial application of recovery controls did not always stop a spin.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Effects of Wing Flaps and Wing Duct Inlet on the Lift and Stalling Characteristics of a 1/4-Scale Partial-Span Model of the Republic XF-12 Airplane in the Langley 19-Foot Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    Graham, Robert R.; Martina, Albert P.; Salmi, Reino J.

    1946-01-01

    An investigation was conducted in the Langley 19-foot pressure tunnel to determine the lift, drag, pitching-moment and stalling characteristics fo a 1/4 -scale partial-span model of the left wing of the Republic XF-12 airplane. The effects of a duct inlet, located between the nacelles at the leading edge of the wing, on those characteristics were also investigated. The Reynolds numbers for the investigation covered a range from 4,500,000 to 8,600,000. The results of the investigation indicated that maximum lift coefficients of 1.36, 1.71, and 2.11 were measured on the model with flaps neutral and deflected 20 deg and 55 deg, respectively at a Reynolds number of 8,600,000. When the duct inlet was replaced by a basic airfoil nose the flap-neutral maximum-lift coefficient was increased from 1.36 to 1.41. The results also showed that at maximum lift with flaps neutral or deflected 55 deg. most of the area between the nacelles were stalled while only small areas on other portions of the model were stalled; when the duct inlet was replaced by the basic airfoil nose the stall was delayed to a slightly higher angle of attack but the nature of the stall was relatively unaffected.

  6. 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... skin and rear spar upper chord of the wing center box, which could result in loss of the airplane wing... proposed AD. Discussion We have received reports of spanwise cracks and corrosion in the wing center...

  7. 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.

  8. 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.

  9. 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.

  10. Effect of Wing Height and Dihedral on the Lateral Stability Characteristics at Low Lift of a 45 Deg Swept-Wing Airplane Configuration as Obtained from Time-Vector Analyses of Rocket-Propelled-Model Flights at Mach Numbers from 0.7 to 1.3

    NASA Technical Reports Server (NTRS)

    Gillis, Clarence L.; Chapman, Rowe, Jr.

    1956-01-01

    Lateral-stability flight tests were made over the Mach number range from 0.7 to 1.3 of models of three airplane configurations having 45deg sweptback wings. One model had a high wing; one, a low wing; and one, a high wing with cathedral. The models were otherwise identical. The lateral oscillations of the models resulting from intermittent yawing disturbances were interpreted in terms of full-scale airplane flying qualities and were further analyzed by the time-vector method to obtain values of the lateral stability derivatives. The effects of changes i n wing height on the static sideslip derivatives were fairly constant in the speed range investigated and agreed well with estimated values based on subsonic wind-tunnel tests. Effects of geometric dihedral on the rolling moment due to sideslip agreed well with theoretical and other experimental results and with a theoretical relation involving the damping in roll. The damping in roll, when compared with theoretical and other experimental results, shared good agreement at supersonic speeds but was somewhat higher at a Mach number of 1.0 and at subsonic speeds. The damping in yaw shared no large changes in the transonic region.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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...

  16. 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...

  17. 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...

  18. Aerodynamic performance of membrane wings with adaptive compliance

    NASA Astrophysics Data System (ADS)

    Curet, Oscar M.; Carrere, Alexander; Pande, Arjun; Breuer, Kenneth S.

    2012-11-01

    Some flying animals use wing membranes with adaptive compliance to control their aerodynamic performance. In this work we characterize the mechanical properties and aerodynamic performance of a low aspect ratio membrane wing composed of a dielectric film supported on a rigid frame. We test the wing model in a wind tunnel. When a fixed voltage is applied across the wing membrane the camber increases, accompanied by a small increase in lift (less than 2%). However, lift is significantly increased when the wing is forced with an oscillating field at specific frequencies that correspond to the characteristic vortex shedding frequency. We present the results concerning the kinematics and aerodynamic performance of the adaptive wing membrane and the coupling between the vortex shedding and the forced modulation of elastic modulus.

  19. 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.

  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. 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.

  3. 78 FR 28729 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-05-16

    ..., 2012) Nord Wind Airlines reported the status of compliance of its airplanes with the NPRM (77 FR 65506... Airplanes AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Final rule. SUMMARY: We are... series airplanes. That AD currently requires modifying the nacelle strut and wing structure,...

  4. 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

  5. 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.

  6. 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.

  7. 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).

  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 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.

  10. 78 FR 8054 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-02-05

    ...) inspection for cracking of the area around the fasteners of the landing plate of the wing bottom skin panel... the inspection of the area around the fasteners of the landing plate of the wing bottom skin panel... A310 series airplanes. The existing AD currently requires repetitive inspections for fatigue...

  11. 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.

  12. 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.

  13. Safeguards Against Flutter of Airplanes

    NASA Technical Reports Server (NTRS)

    deVries, Gerhard

    1956-01-01

    This report is a complilation of practical rules, derived at the same time from theory and from experience, intended to guide the aeronautical engineer in the design of flutter-free airplanes. Rules applicable to the wing, the ailerons, flaps, tabs,tail surfaces, and fuselage are discussed.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  2. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  3. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  4. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  5. 14 CFR 25.457 - Wing flaps.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Wing flaps. 25.457 Section 25.457 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Structure Control Surface and System Loads § 25.457 Wing flaps....

  6. 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.

  7. 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.

  8. 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... ice on the parts of the wings that are critical from the standpoint of ice accumulation....

  9. 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....

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 77 FR 15291 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-03-15

    ...We propose to adopt a new airworthiness directive (AD) for all Airbus Model A300 B4-603, B4-605R, and B4-622R airplanes; Model A300 C4-605R Variant F airplanes; and Model A300 F4-600R series airplanes. This proposed AD was prompted by a report that chafing was detected between the autopilot electrical wiring conduit and the wing bottom skin. This proposed AD would require modifying the wiring......

  15. 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.

  16. 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

  17. 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... ``Discussion'' sections in the NPRM (77 FR 51720, August 27, 2012) be revised to add the text, ``in the center... ignition source in the center wing tank, which in combination with flammable fuel vapors could result in...

  18. 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.)

  19. Wing-Fuselage Interference, Tail Buffeting, and Air Flow About the Tail of a Low-Wing Monoplane

    NASA Technical Reports Server (NTRS)

    White, James A; Hood, Manley J

    1935-01-01

    This report presents the results of wind tunnel tests on a Mcdonnell Douglas airplane to determine the wing-fuselage interference of a low-wing monoplane. The tests included a study of tail buffeting and the air flow in the region of the tail. The airplane was tested with and without the propeller slipstream, both in the original condition and with several devices designed to reduce or eliminate tail buffeting. The devices used were wing-fuselage fillets, a NACA cowling, reflexed trailing edge of the wing, and stub auxiliary airfoils.

  20. 77 FR 16492 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-03-21

    ... Policies and Procedures (44 FR 11034, February 26, 1979); 3. Will not affect intrastate aviation in Alaska... airplanes. This proposed AD was prompted by reports that some nuts installed on the wing, including on... certain nuts are installed or cracked, and replacing the affected nuts if necessary. We are proposing...

  1. 75 FR 6865 - Airworthiness Directives; The Boeing Company Model 737-700 (IGW) Series Airplanes Equipped With...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-02-12

    ... either pneumatic air pressure supplied from the airplane, or electrical power from the airplane to power... pneumatic air pressure to empty into the airplane center wing tank. All auxiliary tanks use some type of... Review, Flammability Reduction and Maintenance and Inspection Requirements'' (66 FR 23086, May 7,...

  2. 77 FR 23385 - Airworthiness Directives; Fokker Services B.V. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-04-19

    ... center wing fuel tank. We are issuing this AD to prevent failure of an in-flight engine re- light... indication), and, for certain airplanes, modifying the transfer logic of the center wing fuel tank. You may... the automatic fuel transfer system: Modify the transfer logic of the center wing fuel tank,...

  3. 76 FR 35336 - Airworthiness Directives; The Boeing Company Model 767 Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-06-17

    ... assemblies of the slat track housing of the wings. For certain other airplanes, the existing AD requires repetitive inspections of the drain tube assemblies of the slat track housing of the wings to find... housing, which could let fuel drain from the main fuel tanks into the dry bay area of the wings and...

  4. Unsteady flow over flexible wings at different low Reynolds numbers

    NASA Astrophysics Data System (ADS)

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

    2016-03-01

    In this study, unsteady flow around flexible membrane wing which had aspect ratio of 1 (AR=1) was investigated experimentally at various Reynolds numbers (Re = 25000 and Re = 50000). Smoke-wire technique for flow visualization over the flexible membrane wing was utilized in the experiments. Digital Image Correlation system (DIC) was used for measuring deformation of AR = 1 flexible membrane wing. Instantaneous deformation measurements of membrane wing were combined with the flow field measurements. In low aspect ratio flexible membrane wings, unsteadiness includes tip vortices and vortex shedding, and the combination of tip vortices. In these types of wings, complex unsteady deformations occurred due to vortex shedding. The results showed that the increasing angle of attack results in increase of membrane deformation. Moreover, it was concluded that analysis of the instantaneous deformation revealed chordwise and spanwise, modes which were due to the shedding of leading-edge vortices as well as tip vortices. Consequently, vibrational mode decreased and maximum standard deviation location approached to the trailing edge by reason of increasing angle of attack.

  5. 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.

  6. 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.

  7. 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...

  8. 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...

  9. 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...

  10. 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.

  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. Flow Structure on a Flapping Wing: Quasi-Steady Limit

    NASA Astrophysics Data System (ADS)

    Ozen, Cem; Rockwell, Donald

    2011-11-01

    The flapping motion of an insect wing typically involves quasi-steady motion between extremes of unsteady motion. This investigation characterizes the flow structure for the quasi-steady limit via a rotating wing in the form of a thin rectangular plate having a low aspect ratio (AR =1). Particle Image Velocimetry (PIV) is employed, in order to gain insight into the effects of centripetal and Coriolis forces. Vorticity, velocity and streamline patterns are used to describe the overall flow structure with an emphasis on the leading-edge vortex. A stable leading-edge vortex is maintained over effective angles of attack from 30° to 75° and it is observed that at each angle of attack the flow structure remains relatively same over the Reynolds number range from 3,600 to 14,500. The dimensionless circulation of the leading edge vortex is found to be proportional to the effective angle of attack. Quasi-three-dimensional construction of the flow structure is used to identify the different regimes along the span of the wing which is then complemented by patterns on cross flow planes to demonstrate the influence of root and tip swirls on the spanwise flow. The rotating wing results are also compared with the equivalent of translating wing to further illustrate the effects of the rotation.

  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. 75 FR 6821 - Airworthiness Directives; The Boeing Company Model 747-400, 747-400D, and 747-400F Series Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-02-11

    ...), would protect the fuel densitometer for the horizontal stabilizer tank (HST) and the center wing tank...) For all airplanes: Install the HSP in the center wing tank, in accordance with the Accomplishment... system (FQIS) of the center fuel tank and, for certain airplanes, the horizontal stabilizer fuel...

  15. An evaluation of aerodynamics modeling of spinning light airplanes

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    This paper extends the application of the modified strip theory for wing body combination of a spinning light airplane reported earlier. In addition, to account for the contribution of the tail plane, the shielding effect on vertical tail under steady state spin condition is modeled from basic aerodynamic considerations. The results of this combined analysis, presented for some light airplane configurations, are shown to be in good agreement with spin tunnel rotary balance test data.

  16. 75 FR 50853 - Special Conditions: Cirrus Design Corporation Model SF50 Airplane; Function and Reliability Testing

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-08-18

    ... airplanes. 1. Function and Reliability Testing. Flight tests: In place of 14 CFR 21.35(b)(2), the following... Airplane; Function and Reliability Testing AGENCY: Federal Aviation Administration (FAA), DOT. ACTION... ``Vision'' Jet. The SF50 is a low- wing, five-plus-two-place (2 children), single-engine...

  17. 75 FR 29962 - Special Conditions: Cirrus Design Corporation Model SF50 Airplane; Function and Reliability Testing

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-05-28

    ... SF50 airplanes. 1. Function and Reliability Testing Flight tests: In place of 14 CFR part 21.35(b)(2... Airplane; Function and Reliability Testing AGENCY: Federal Aviation Administration (FAA), DOT. ACTION... low-wing, five-plus-two-place (2 children), single-engine turbofan- powered aircraft. It...

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Effects of aspect ratio on flapping wing aerodynamics in animal flight

    NASA Astrophysics Data System (ADS)

    Fu, Jun-Jiang; Hefler, Csaba; Qiu, Hui-He; Shyy, Wei

    2014-12-01

    Morphology as well as kinematics is a critical determinant of performance in flapping flight. To understand the effects of the structural traits on aerodynamics of bio-flyers, three rectangular wings with aspect ratios (AR) of 1, 2, and 4 performing hovering-like sinusoidal kinematics at wingtip based Reynolds number of 5 300 are experimentally investigated. Flow structures on sectional cuts along the wing span are compared. Stronger K-H instability is found on the leading edge vortex of wings with higher aspect ratios. Vortex bursting only appears on the outer spanwise locations of high-aspect-ratio wings. The vortex bursting on high-aspect-ratio wings is perhaps one of the reasons why bio-flyers normally have low-aspect-ratio wings. Quantitative analysis exhibits larger dimensionless circulation of the leading edge vortex (LEV) over higher aspect ratio wings except when vortex bursting happens. The average dimensionless circulation of AR1 and AR2 along the span almost equals the dimensionless circulation at the 50% span. The flow structure and the circulation analysis show that the sinusoidal kinematics suppresses breakdown of the LEV compared with simplified flapping kinematics used in similar studies. The Reynolds number effect results on AR4 show that in the current Re range, the overall flow structure is not sensitive to Reynolds number.

  8. 76 FR 69163 - Airworthiness Directives; Fokker Services B.V. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-11-08

    ... airplanes modifying the transfer logic of the center wing fuel tank. You may obtain further information by... the transfer logic of the center wing fuel tank, in accordance with the Accomplishment Instructions of..., contact Fokker Services B.V., Technical Services Dept., P.O. Box 231, 2150 AE Nieuw-Vennep,...

  9. 76 FR 54093 - Airworthiness Directives; Bombardier, Inc. Model DHC-8-400 Series Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-08-31

    ... attachments. * * * * * The unsafe condition could result in separation of the wing from the airplane during... comments on this AD by October 17, 2011. ADDRESSES: You may send comments by any of the following methods... hardware if necessary]. The unsafe condition could result in separation of the wing from the...

  10. Joukowski wings

    NASA Technical Reports Server (NTRS)

    Margoulis, W

    1922-01-01

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

  11. 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-...

  12. 77 FR 42455 - Airworthiness Directives; Piper Aircraft, Inc. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-07-19

    ... airplane's wing root, is not at a high risk of an inadvertent in-flight engine shutdown and loss of flight... operators Relocate the magneto switch from the 3.5 work-hours x $85 $125 $422.50 $124,637.50 port side... and Procedures (44 FR 11034, February 26, 1979), (3) Will not affect intrastate aviation in...

  13. 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.

  14. 77 FR 40826 - Airworthiness Directives; Bombardier, Inc. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-07-11

    ... 12866; 2. Is not a ``significant rule'' under the DOT Regulatory Policies and Procedures (44 FR 11034... 1000) airplanes. This proposed AD was prompted by a report that certain wing-to-fuselage attachment...-to- fuselage attachment joints and replacement if necessary. We are proposing this AD to prevent...

  15. 77 FR 31483 - Airworthiness Directives; Gulfstream Aerospace Corporation Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-05-29

    ... Policies and Procedures (44 FR 11034, February 26, 1979), (3) Will not affect intrastate aviation in Alaska...-to-fuselage attachment, and repair if necessary. This AD also requires, for certain other airplanes... fitting at the wing-to-fuselage attachment, and repair if necessary. This AD was prompted by a report...

  16. 78 FR 46538 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-08-01

    ... airplanes. This proposed AD was prompted by reports indicating that a standard fuel tank access door was located where an impact-resistant access door was required, and stencils were missing from some impact-resistant access doors. This proposed AD would require an inspection of the left- and right-hand wing...

  17. 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.

  18. 77 FR 65506 - Airworthiness Directives; The Boeing Company Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-10-29

    ... modification of the nacelle strut and wing structure, and repair of any damage found during the modification... Commercial Airplanes, Attention: Data & Services Management, P.O. Box 3707, MC 2H-65, Seattle, WA 98124-2207... proposed AD. Discussion On August 29, 2003, we issued AD 2003-18-05, Amendment 39-13296 (68 FR...

  19. 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.

  20. The Light Airplane

    NASA Technical Reports Server (NTRS)

    Driggs, Ivan H.

    1925-01-01

    This report begins with a review and analysis of the work being done to develop light airplanes in the U.S. and abroad. A technical discussion of the construction and innovations in light airplanes is then presented.

  1. 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.

  2. A correlation of results of flight investigation with results of an analytical study of effects of wing flexibility on wing strains due to gusts

    NASA Technical Reports Server (NTRS)

    Shufflebarger, C C; Payne, Chester B; Cahen, George L

    1958-01-01

    An analytical study of the effects of wing flexibility on wing strains due to gusts has been made for four spanwise stations of a four-engine bomber airplane, and the results have been correlated with results of a previous flight investigation.

  3. 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.

  4. Static Longitudinal Characteristics at High Subsonic Speeds of a Complete Airplane Model with a Highly Tapered Wing having the 0.80 Chord Line Unswept and with Several Tail Configurations

    NASA Technical Reports Server (NTRS)

    Goodson, Kenneth W.

    1961-01-01

    An investigation was made at high subsonic speeds of a complete model having a highly tapered wing and several tail configurations. The basic aspect-ratio-4.00 wing had zero taper and an unswept 0.80 chord line. Several aspect-ratio modifications to the basic wing were made by clipping off portions of the wing tips. The complete model was tested with a chord-plane tail, a T-tail, and a biplane tail (combined T-tail and chord-plane tail). The model was tested in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92. The data show that, when reduced to the same static margin, all the tail configurations tested on the model provided fairly good stability characteristics, the biplane tail giving the.best overall characteristics as regards pitching-moment linearity. Changes in static margin at zero lift coefficient with Mach number were small for the model with these tails over the Mach number range investigated.

  5. 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.

  6. 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.

  7. Quasi-Steady Limit of Flow Structure on Flapping Wing in Mean Flow

    NASA Astrophysics Data System (ADS)

    Bross, Matthew; Ozen, Cem; Rockwell, Donald

    2012-11-01

    A limiting case of flapping motion of a wing (low aspect ratio plate) in presence of incident steady flow is compared to a rotating wing in quiescent fluid, in order to clarify the effect of advance ratio J (ratio of free-stream velocity to tangential velocity of wing) on the structure of the leading-edge vortex. Stereoscopic particle image velocimetry leads to patterns of vorticity, velocity contours, and streamlines. For each value of J, the effective angle of attack is held constant at 45°, while the wing rotates from rest through 270°. While at rest, the wing at high angle of attack in the presence of a steady free-stream gives rise to fully stalled flow. After the onset of rotation, the fully stalled region very quickly transforms to a stable leading edge vortex. Despite the change in advance ratio, the development of the flow structure around the wing throughout the rotation maneuver is similar, especially in the leading edge vortex region, as evidenced by patterns of streamline topology. To further demonstrate the effect of J, three-dimensional representations of of spanwise-oriented vorticity, spanwise velocity, and Q were constructed for hovering flight and forward flight.

  8. Spanwise gradients in flow speed and leading edge vortex attachment on low Reynolds number wings

    NASA Astrophysics Data System (ADS)

    Jardin, Thierry; David, Laurent

    2014-11-01

    It is now accepted that the aerodynamic performance of low aspect ratio revolving wings, such as insect wings or maple seed membranes, largely relies on sustained leading edge vortex attachment. However, the mechanisms responsible for this sustained attachment are still poorly understood. Here, we compute the Navier-Stokes solution of the flow around a finite wing (i) subjected to a uniform oncoming flow, (ii) subjected to a spanwise varying oncoming flow and (iii) revolving about its root. Therefore, we are able to isolate the mechanisms associated with the spanwise gradient of the local wing speed from those associated with centrifugal and Coriolis effects. We show that over flapping amplitudes typical of insect flight the spanwise gradient of the local wing speed may suffice in maintaining leading edge vortex attachment. We correlate this result with the development of spanwise flow and we evaluate the sensitivity of such a mechanism to the Reynolds number. It is noted, however, that leading edge vortex attachment through the spanwise gradient of the local wing speed does not promote large lift, which ultimately arises from centrifugal and Coriolis effects.

  9. Results of Measurements of Maximum Lift and Buffeting Intensities Obtained During Flight Investigation of the Northrop X-4 Research Airplane

    NASA Technical Reports Server (NTRS)

    Baker, Thomas F

    1953-01-01

    The variation of the intensity of buffeting experienced throughout the operational region of the semitailless Northrop X-4 airplane and the values of maximum and peak normal-force coefficients in the Mach number range from 0.42 to 0.92 have been determined. The results are compared with data obtained with the swept-wing Douglas D-558-II airplane.

  10. 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.

  11. Winged pipelaying

    SciTech Connect

    Ayers, R.R.; Kopp, F.

    1988-12-06

    This patent describes an apparatus for towing at least one submerged pipeline above-seabed comprising: tow means attached to the pipeline; and at least one wing attached to the pipeline and positioned to provide lifting force to the pipeline when the pipeline is being towed, the wing being rotatable from a substantially perpendicular alignment to a substantially perpendicular alignment to a substantially lateral alignment with the pipeline in a non-towing mode.

  12. 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.

  13. 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.

  14. 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.

  15. The oblique wing-research aircraft

    NASA Technical Reports Server (NTRS)

    Andrews, W. H.

    1980-01-01

    The AD-1 airplane was designed as a low cost, low speed manned research tool to evaluate the flying qualities of the oblique wing concept. The airplane is constructed primarily of foam and fiberglass and incorporates simplicity in terms of the onboard systems. There are no hydraulics, the control system is cable and torque tube, and the electrical systems consist of engine driven generators which power the battery for engine start, cockpit gages, trim motors, and the onboard data system. The propulsion systems consist of two Microturbo TRS-18 engines sea level trust rated at 220 pounds. The airplane weighs approximately 2100 pounds and has a performance potential in the range of 200 knots and an altitude of 15,000 feet.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. AD-1 oblique wing research aircraft pilot evaluation program

    NASA Technical Reports Server (NTRS)

    Painter, W. D.

    1983-01-01

    A flight test program of a low cost, low speed, manned, oblique wing research airplane was conducted at the NASA Dryden Flight Research Facility in cooperation with NASA Ames Research Center between 1979 and 1982. When the principal purpose of the test program was completed, which was to demonstrate the flight and handling characteristics of the configuration, particularly in wing-sweep-angle ranges from 45 to 60 deg, a pilot evaluation program was conducted to obtain a qualification evaluation of the flying qualities of an oblique wing aircraft. These results were documented for use in future studies of such aircraft.

  2. Design charts relating to the stalling of tapered wings

    NASA Technical Reports Server (NTRS)

    Soule, H A; Anderson, R F

    1940-01-01

    An aid in airplane design, charts have been prepared to show the effects of wing taper, thickness ratio, and Reynolds number on the spanwise location of the initial stalling point. Means of improving poor stalling characteristics resulting from certain combinations of the variables have also been considered; additional figures illustrate the influence of camber increase to the wing tips, washout, central sharp leading edges, and wing-tip slots on the stalling characteristics. Data are included from which the drag increases resulting from the use of these means can be computed. The application of the data to a specific problem is illustrated by an example.

  3. 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

  4. A family of supersonic airplanes: Technical and economic feasibility

    NASA Technical Reports Server (NTRS)

    Neumann, F. D.; Whitten, J. W.

    1980-01-01

    To improve the prospects for success in the market place, the family approach is essential to the design of future supersonic airplanes. The evolution from a basic supersonic airplane to a family could follow historic patterns, with one exception: substantial changes in passenger carrying capacity will be difficult by the conventional fuselage "doughnut" approach so successfully used on the cylindrical fuselage of subsonic airplanes. The primary reasons for this difference include the requirement for highly integrated "area ruled" configurations, to give the desired high supersonic aerodynamic efficiency, and other physical limitations such as takeoff and landing rotation. A concept for a supersonic airplane family that could effectively solve the variable range and passenger capacity problem provides for modification of the fuselage cross section that makes it possible to build a family of three airplanes with four, five, and six abreast passenger seating. This is done by replacing or modifying portions of the fuselage. All airplanes share the same wing, engines, and major subsystems. Only small sections of the fuselage would be different, and aerodynamic efficiency need not be compromised.

  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. Jet noise of an augmentor wing-advanced supersonic transport

    NASA Technical Reports Server (NTRS)

    Franciscus, L.

    1972-01-01

    A preliminary mission study was made of the range and jet noise of an advanced supersonic transport (AST) employing an augmentor wing and four duct burning turbofan engines. The airplane weight and aerodynamic characteristics of the Boeing 2707-300 airplane with a gross weight of 750,000 pounds and 234 passengers was used for the study. Engine thrust was fixed at 58,000 pounds per engine and engine size was increased to obtain the required thrust at reduced power settings for jet noise reduction. Turbofan engine core noise was reduced to FAR 36 noise levels and lower by proper selection of turbine inlet temperature, bypass ratio and fan pressure ratio. The study showed that an augmentor wing can reduce the bypass jet noise sufficiently so that total noise levels below FAR 36 can be attained without significant range penalties if the augmentor wing can be designed without severe weight and performance penalties.

  7. Thin tailored composite wing for civil tiltrotor

    NASA Technical Reports Server (NTRS)

    Rais-Rohani, Masoud

    1994-01-01

    The tiltrotor aircraft is a flight vehicle which combines the efficient low speed (i.e., take-off, landing, and hover) characteristics of a helicopter with the efficient cruise speed of a turboprop airplane. A well-known example of such vehicle is the Bell-Boeing V-22 Osprey. The high cruise speed and range constraints placed on the civil tiltrotor require a relatively thin wing to increase the drag-divergence Mach number which translates into lower compressibility drag. It is required to reduce the wing maximum thickness-to-chord ratio t/c from 23% (i.e., V-22 wing) to 18%. While a reduction in wing thickness results in improved aerodynamic efficiency, it has an adverse effect on the wing structure and it tends to reduce structural stiffness. If ignored, the reduction in wing stiffness leads to susceptibility to aeroelastic and dynamic instabilities which may consequently cause a catastrophic failure. By taking advantage of the directional stiffness characteristics of composite materials the wing structure may be tailored to have the necessary stiffness, at a lower thickness, while keeping the weight low. The goal of this study is to design a wing structure for minimum weight subject to structural, dynamic and aeroelastic constraints. The structural constraints are in terms of strength and buckling allowables. The dynamic constraints are in terms of wing natural frequencies in vertical and horizontal bending and torsion. The aeroelastic constraints are in terms of frequency placement of the wing structure relative to those of the rotor system. The wing-rotor-pylon aeroelastic and dynamic interactions are limited in this design study by holding the cruise speed, rotor-pylon system, and wing geometric attributes fixed. To assure that the wing-rotor stability margins are maintained a more rigorous analysis based on a detailed model of the rotor system will need to ensue following the design study. The skin-stringer-rib type architecture is used for the wing

  8. Metal Airplane Construction

    NASA Technical Reports Server (NTRS)

    1926-01-01

    It has long been thought that metal construction of airplanes would involve an increase in weight as compared with wood construction. Recent experience has shown that such is not the case. This report describes the materials used, treatment of, and characteristics of metal airplane construction.

  9. 75 FR 60614 - Airworthiness Directives; The Boeing Company Model 747-400, 747-400D, and 747-400F Series Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2010-10-01

    .... (1) For all airplanes: Install the HSP in the center wing tank, in accordance with the Accomplishment... the fuel quantity indicating system (FQIS) of the center fuel tank and, for certain airplanes, the... source inside the center or horizontal stabilizer fuel tanks. An ignition source, in combination...

  10. 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.

  11. Vortex Formation in the Starting Flow of Rotating Low-Aspect-Ratio Plates

    NASA Astrophysics Data System (ADS)

    Devoria, Adam; Ringuette, Matthew

    2010-11-01

    We investigate the unsteady flow of fish fin-like plates accelerating from rest through various angular velocity profiles. The objective is to gain an understanding of the connection among the prescribed kinematics and resulting vortex formation; a relationship which has not currently been thoroughly explored. The root-to-tip flow that is induced by the plate motion is expected to have significant effects on the vortex formation. Additionally, different plate shapes are studied to compare the effects of geometrical changes. The experiments are conducted in a water tank, and the plates have a fixed axis of rotation. Digital particle image velocimetry (DPIV) is used to measure the flow velocity in a symmetry plane through the plates. Vorticity and circulation are subsequently computed and vortices are distinguished from surrounding flow structures using vortex identification schemes. Carefully incorporating these techniques will aid the development of scaling laws to characterize the vortex formation with maximum attainable vortex strength.

  12. 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.

  13. 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

  14. Features of spherical torus plasmas of ultra-low aspect ratio and large elongation

    NASA Astrophysics Data System (ADS)

    Peng, Y. K.

    High beta, good confinement and steady-state operation in compact configuration and at modest field have long been major goals of magnetic fusion energy research. Accomplishing these in a single device will permit cost-effective and attractive embodiments of future fusion reactors. The introduction of the spherical torus concept was so motivated to a large degree. An equally important motivation is its prospect of reducing cost and time of fusion research and development. In comparison with alternate confinement concepts of RFP and spheromak, a spherical torus experiment is expected to be similar in compactness, low field, and high beta; but with better, tokamak-like confinement times by more than an order of magnitude.

  15. Subcyclotron Instability of Alfven Eigenmodes due to Energetic Ions in Low Aspect Ratio Plasmas

    SciTech Connect

    N.N. Gorelenko; E. Fredrickson; E. Belova; C.Z. Cheng; D. Gates; R. White

    2003-08-21

    High-frequency modes with frequencies below the fundamental cyclotron frequency of thermal ions were observed in the National Spherical Torus Experiment (NSTX). Based on the measured spectrum of high-frequency modes they are identified as Compressional Alfven Eigenmodes (CAEs) and Global Alfven Eigenmodes (GAEs). CAEs have similar time evolution as plasma parameters change, while GAEs may intersect due to q-profile relaxation. A theory has been developed to study the properties of these modes. Both types of instabilities are driven by the tangential neutral-beam injection in NSTX. Beam ions excite CAEs/GAEs through the Doppler-shifted cyclotron resonance. The main source for the drive is the velocity space anisotropy of the beam ion distribution function. Simulations of the effect CAEs/GAEs may have on plasma ions indicate that these modes may provide a channel for efficient energy transfer from fast ions directly to thermal ions.

  16. 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. PMID:19746195

  17. Progress towards Steady State at Low Aspect Ratio on the National Spherical Torus Experiment (NSTX)

    SciTech Connect

    D.A. Gates, J. Menard, R. Maingi, S. Kaye, S.A. Sabbagh, S. Diem, J.R.Wilson, M.G. Bell, R.E. Bell, J. Ferron, E.D. Fredrickson, C.E. Kessel, B.P. LeBlanc, F. Levinton, J. Manickam, D. Mueller, R. Raman, T. Stevenson, D. Stutman, G. Taylor, K. Tritz, H. Yu, and the NSTX Research Team

    2007-11-08

    Modifications to the plasma control capabilities and poloidal field coils of the National Spherical Torus Experiment (NSTX) have enabled a significant enhancement in shaping capability which has led to the transient achievement of a record shape factor (S ≡ q95 (Iρ/αΒτ)) of ~41 (MA m-1 Τ-1) simultaneous with a record plasma elongation of κ ≡ β /α ~ 3. This result was obtained using isoflux control and real-time equilibrium reconstruction. Achieving high shape factor together with tolerable divertor loading is an important result for future ST burning plasma experiments as exemplified by studies for future ST reactor concepts, as well as neutron producing devices, which rely on achieving high shape factors in order to achieve steady state operation while maintaining MHD stability. Statistical evidence is presented which demonstrates the expected correlation between increased shaping and improved plasma performance.

  18. Electron Bernstein Wave Experiments in the Low Aspect ratio Torus Experiment Device

    NASA Astrophysics Data System (ADS)

    Noguchi, Yuto; Fukunaga, Tadahiko; Uchida, Masaki; Tanaka, Hitoshi; Maekawa, Takashi

    2012-10-01

    Two kinds of EBW experiments have been done on LATE using 2.45 GHz microwaves obliquely injected using open circular waveguide launchers. The vacuum vessel is a cylinder with a diameter of 100 cm and a height of 100 cm, and an 11.4 cm center-post for center conductors for the Bt field with no central solenoid. In the first experiment we inject a relatively high power of 40-60 kW from three 20kW-magnetrons, which non-inductively starts up the discharge and finally maintains a toroidal current of 10 kA with a central plasma density 10 times the plasma cutoff density by EBW heating and current drive at the first propagation band. A wave injection with the optimal polarization for mode-conversion to EBW via one launcher (other two launchers being with usual linear polarization parallel to the mid-plane) improves plasma performance in terms of density and current. In the second experiment we investigate various basic characteristics of EBW, including the mode-conversion and subsequent propagation and absorption of EBW. Here, an 1.5 GHz microwave at 10 W is injected via OXB scheme into an ECR plasma maintained by a 2.45 GHz microwave at 1 kW. The wave field for 1.5GHz microwave is 2-dimensionally measured on the mid-plane using probes attached to a two-joint robot arm.

  19. X-Wing Research Vehicle

    NASA Technical Reports Server (NTRS)

    1986-01-01

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

  20. An analysis of the stability of an airplane with free controls

    NASA Technical Reports Server (NTRS)

    Jones, Robert T; Cohen, Doris

    1941-01-01

    Report presents the results of an investigation made of the essentials to the stability of an airplane with free control surfaces. Calculations are based on typical airplane characteristics with certain factors varied to cover a range of current designs. Stability charts are included to show the limiting values of the aerodynamic hinge moments and the weight hinge moments of the control surfaces for various positions of the center of gravity of the airplane and for control systems with various moments of inertia. The effects of reducing the chord and of eliminating the floating tendency of the surface, of changing the wing loading, and of decreasing the radius of gyration of the airplane are also indicated. An investigation has also been made of the nature of the motion of the airplane with controls free and of the modes of instability that may occur.

  1. 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.

  2. 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.

  3. Development and design of flexible Fowler flaps for an adaptive wing

    NASA Astrophysics Data System (ADS)

    Monner, Hans P.; Hanselka, Holger; Breitbach, Elmar J.

    1998-06-01

    Civil transport airplanes fly with fixed geometry wings optimized only for one design point described by altitude, Mach number and airplane weight. These parameters vary continuously during flight, to which means the wing geometry seldom is optimal. According to aerodynamic investigations a chordwide variation of the wing camber leads to improvements in operational flexibility, buffet boundaries and performance resulting in reduction of fuel consumption. A spanwise differential camber variation allows to gain control over spanwise lift distributions reducing wing root bending moments. This paper describes the design of flexible Fowler flaps for an adaptive wing to be used in civil transport aircraft that allows both a chordwise as well as spanwise differential camber variation during flight. Since both lower and upper skins are flexed by active ribs, the camber variation is achieved with a smooth contour and without any additional gaps.

  4. Flow Structure on a Wing Due to Unsteady Pitch-Up and Rotation Maneuvers

    NASA Astrophysics Data System (ADS)

    Bross, Matthew; Yilmaz, Turgut; Rockwell, Donald

    2013-11-01

    The flow structure along a rectangular (low aspect ratio) wing undergoing pure pitch-up, pitch-up with rotation, and pure rotation is characterized as a function of dimensionless convective time τ during each maneuver. Quantitative imaging via angular displacement stereo particle image velocimetry was used to determine the three-dimensional velocity field, thereby allowing analysis of the effects of different wing kinematics via representations of Q-criterion, vorticity flux, and velocity and vorticity contours. Despite the difference in wing kinematics, interactions between leading-edge and tip vortices persist across all values of τ. The three-dimensional flow structure involves a symmetric pattern along the wing during pure pitch-up and transforms to a conical leading-edge vortex in conjunction with a tip vortex that extends into the wake for both pitch-up with rotation and pure rotation. This observation suggests that rotational motion has a greater influence than pitching motion in establishing the form and scale of the leading-edge vortex. Finally, sectional images of the flow structure arising from combined pitch-up and rotation were acquired at three different pitch rates relative to a given rate of pure rotation at fixed angle of attack.

  5. Fluid-structure Interaction of Rigid and Flexible Wings in Ground Effect

    NASA Astrophysics Data System (ADS)

    Bleischwitz, Robert; de Kat, Roeland; Ganapathisubramani, Bharathram

    2015-11-01

    Inspired by trawling bats, combining flexible membrane wings and the vicinity of the ground, an experimental wind tunnel study is conducted at Re = 56,000 to determine the fluid-structure-ground interaction of rectangular, perimeter reinforced low aspect ratio (AR = 2) membrane wings in free flight and ground effect conditions. The pitch angle is varied between 10° <= α <=25° . Flexible membrane wings are compared with rigid flat plates. Instantaneous lift and drag forces are simultaneously recorded with membrane and flow dynamics (Digital-Image-Correlation + Particle-Image-Velocimetry). The focus of this study involves coupling effects of membrane mode shapes (chordwise + spanwise) and flow structures changing with angle of attack and height over ground. A POD analysis of the flow, membrane vibrations and forces should help to identify aerodynamic beneficial vibration shapes and their impact on flow features such as leading edge and tip vortices. The knowledge is seen to be essential for efficient usage of MAVs with membrane wings in and out of ground effect. PhD student.

  6. 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.

  7. Notes on the Construction and Testing of Model Airplanes

    NASA Technical Reports Server (NTRS)

    Diehl, Walter S

    1922-01-01

    Here, it is shown that the construction of an airplane model can and should be simplified in order to obtain the most reliable test data. General requirements for model construction are given, keeping in mind that the general purpose of wind tunnel tests on a model airplane is to obtain the aerodynamic characteristics, the static balance, and the efficiency of controls for the particular combination of wings, tail surfaces, fuselage, and landing gear employed in the design. These parts must be exact scale reproductions. Any appreciable variation from scale reproduction must be in the remaining parts of the model, i.e., struts, wires, fittings, control horns, radiators, engines, and the various attachments found exposed to the wind in special airplanes. Interplane bracing is discussed in some detail.

  8. Rotor systems research aircraft airplane configuration flight-test results

    NASA Technical Reports Server (NTRS)

    Painter, W. D.; Erickson, R. E.

    1984-01-01

    The Rotor Systems Research Aircraft (RSRA) has been undergoing ground and flight tests by Ames Research Center since late 1979, primarily as a compound aircraft. The purpose was to train pilots and to check out and develop the design flight envelope established by the Sikorsky Aircraft Company. This paper reviews the preparation and flight test of the RSRA in the airplane, or fixed-wing, configuration and discusses the results of that test.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. General airplane performance

    NASA Technical Reports Server (NTRS)

    Rockfeller, W C

    1939-01-01

    Equations have been developed for the analysis of the performance of the ideal airplane, leading to an approximate physical interpretation of the performance problem. The basic sea-level airplane parameters have been generalized to altitude parameters and a new parameter has been introduced and physically interpreted. The performance analysis for actual airplanes has been obtained in terms of the equivalent ideal airplane in order that the charts developed for use in practical calculations will for the most part apply to any type of engine-propeller combination and system of control, the only additional material required consisting of the actual engine and propeller curves for propulsion unit. Finally, a more exact method for the calculation of the climb characteristics for the constant-speed controllable propeller is presented in the appendix.

  14. 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.

  15. Short revolving wings enable hovering animals to avoid stall and reduce drag

    NASA Astrophysics Data System (ADS)

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

    2014-11-01

    Long and slender wings reduce the drag of airplanes, helicopters, and gliding animals, which operate at low angle of attack (incidence). Remarkably, there is no evidence for such influence of wing aspect ratio on the energetics of hovering animals that operate their wings at much higher incidence. High incidence causes aircraft wings to stall, hovering animals avoid stall by generating an attached vortex along the leading edge of their wings that elevates lift. Hypotheses that explain this capability include the necessity for a short radial distance between the shoulder joint and wing tip, measured in chord lengths, instead of the long tip-to-tip distance that elevates aircraft performance. This stems from how hovering animals revolve their wings around a joint, a condition for which the precise effect of aspect ratio on stall performance is unknown. Here we show that the attachment of the leading edge vortex is determined by wing aspect ratio with respect to the center of rotation-for a suite of aspect ratios that represent both animal and aircraft wings. The vortex remains attached when the local radius is shorter than 4 chord lengths, and separates outboard on more slender wings. Like most other hovering animals, hummingbirds have wing aspect ratios between 3 and 4, much stubbier than helicopters. Our results show this makes their wings robust against flow separation, which reduces drag below values obtained with more slender wings. This revises our understanding of how aspect ratio improves performance at low Reynolds numbers.

  16. 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.

  17. Preliminary Full-Scale Wind-Tunnel Investigation of Wing Ducts for Radiators, Special Report

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe; Nickle, F. R.

    1938-01-01

    Wing ducts for liquid-cooled engine radiators have been investigated in the N.A.C.A. full-scale wind tunnel on a large model airplane. Th e tests were made to determine the relative merits of several types of duct and radiator installations for an airplane of a particular des ign. In the test program the principal duct dimensions were system atically varied, and the results are therefore somewhat applicable to the general problems of wing duct design, although they should be co nsidered as preliminary and only indicative of the inherent possibil ities.

  18. 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.

  19. 77 FR 18141 - Airworthiness Directives; Fokker Services B.V. Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-03-27

    ...) fuel quantity indication system (FQIS) probe and the bottom of the tank structure. This condition, if not corrected, could, under certain conditions, result in an ignition source in the tank vapor space... require, for all airplanes, applying sealant below the FQIS probes in the wing tanks; and for...

  20. 78 FR 36084 - Special Conditions: Cirrus Design Corporation Model SF50 Airplane; Function and Reliability...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2013-06-17

    ... Airplane; Function and Reliability Testing; Withdrawal AGENCY: Federal Aviation Administration (FAA), DOT...: This special condition published on August 18, 2010 at 75 FR 50853 is withdrawn, effective June 17... ``Vision'' Jet. The SF50 is a low-wing, five-plus-two-place (2 children), single-engine...

  1. Damage tolerant composite wing panels for transport aircraft

    NASA Technical Reports Server (NTRS)

    Smith, Peter J.; Wilson, Robert D.; Gibbins, M. N.

    1985-01-01

    Commercial aircraft advanced composite wing surface panels were tested for durability and damage tolerance. The wing of a fuel-efficient, 200-passenger airplane for 1990 delivery was sized using grahite-epoxy materials. The damage tolerance program was structured to allow a systematic progression from material evaluations to the optimized large panel verification tests. The program included coupon testing to evaluate toughened material systems, static and fatigue tests of compression coupons with varying amounts of impact damage, element tests of three-stiffener panels to evaluate upper wing panel design concepts, and the wing structure damage environment was studied. A series of technology demonstration tests of large compression panels is performed. A repair investigation is included in the final large panel test.

  2. Comparison of predicted and measured drag for a single-engine airplane

    NASA Technical Reports Server (NTRS)

    Ward, D. T.; Taylor, F. C.; Doo, J. T. P.

    1985-01-01

    Renewed interest in natural laminar flow (NLF) has rekindled designers' concerns that manufacturing deviations, (loss of surface contours or other surface imperfections) may destroy the effectiveness of NLF for an operational airplane. This paper reports on experimental research that compares predicted and measured boundary layer transition, total drag, and two-dimensional drag coefficients for three different wing surface conditions on an airplane typical of general aviation manufacturing technology. The three flight test phases included: (1) assessment of an unpainted airframe, (2) flight tests of the same airplane after painstakingly filling and sanding the wings to design contours, and (3) similar measurements after this airplane was painted. In each flight phase, transition locations were monitored using either sublimating chemicals or pigmented oil. As expected, total drag changes were difficult to measure. Two-dimensional drag coefficients were estimated using the Eppler-Somers code and measured with a wake rake in a method very similar to Jones' pitot traverse method. The net change in two-dimensional drag was approximately 20 counts between the unpainted airplane and the 'hand-smoothed' airplane for typical cruise flight conditions.

  3. 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.

  4. 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.

  5. 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.

  6. Flutter clearance of the F-14 variable-sweep transition flight experiment airplane, phase 1

    NASA Technical Reports Server (NTRS)

    Kehoe, Michael W.

    1987-01-01

    An F-14 airplane was modified to become the test bed aircraft for the variable sweep transition flight experiment (VSTFE) program. The latter is a laminar flow program designed to measure the effects of wing sweep on boundary layer transition from laminar to turbulent flow. The airplane was modified by adding an upper surface foam-fiberglass glove over a portion of the left wing. Ground vibration and flight flutter testing were accomplished to clear a sufficient flight envelope to conduct the laminar flow experiments. Flight test data indicated satisfactory damping levels and damping trends for the elastic structural modes of the airplane. The data presented include frequency and damping as functions of Mach number.

  7. F-16XL Airplane Research Paint Scheme Aerials

    NASA Technical Reports Server (NTRS)

    1993-01-01

    The NASA F-16XL research airplane is shown in flight over Eastern Shore of Virginia During the Cranked-Arrow Wing Aerodynamics Project (CAWAP). The aircraft is shown in the research paint scheme utilized for on-surface flow visualization. During CAWAP, tufts, oil, and liquid crystals were used to document the flow patterns on the upper surfaces of the wings. The images were recorded by onboard video cameras and hand-held video and still cameras in a chase aircraft. The visual images are being correlated with on board measurements of surface pressures, boundary-layer velocity profiles, and skin friction. Ultimately, the in-flight data will be combined with wind-tunnel results and computational-fluid-dynamics solutions to document the aerodynamics of cranked-arrow wings. CAWAP is a cooperative venture of the Langley Research Center and the Dryden Flight Research Center.

  8. Nonlinear analysis and enhancement of wing-based piezoaeroelastic energy harvesters

    NASA Astrophysics Data System (ADS)

    Abdelkefi, A.; Ghommem, M.; Nuhait, A. O.; Hajj, M. R.

    2014-01-01

    We investigate the level of harvested power from aeroelastic vibrations for an elastically mounted wing supported by nonlinear springs. The energy is harvested by attaching a piezoelectric transducer to the plunge degree of freedom. The considered wing has a low-aspect ratio and hence three dimensional aerodynamic effects cannot be neglected. To this end, the three dimensional unsteady vortex lattice method for the prediction of the unsteady aerodynamic loads is developed. A strong coupling scheme that is based on Hamming's fourth-order predictor-corrector method and accounts for the interaction between the aerodynamic loads and the motion of the wing is employed. The effects of the electrical load resistance, nonlinear torsional spring and eccentricity between the elastic axis and the gravity axis on the level of the harvested power, pitch and plunge amplitudes are investigated for a range of operating wind speeds. The results show that there is a specific wind speed beyond which the pitch motion does not pick any further energy from the incident flow. As such, the displacement in the plunge direction grows significantly and causes enhanced energy harvesting. The results also show that the nonlinear torsional spring plays an important role in enhancing the level of the harvested power. Furthermore, the harvested power can be increased by an order of magnitude by properly choosing the eccentricity and the load resistance. This analysis is helpful in designing piezoaeroelastic energy harvesters that can operate optimally at specific wind speeds.

  9. Ground vibration test of F-16 airplane with initial decoupler pylon

    NASA Technical Reports Server (NTRS)

    Cazier, F. W., Jr.; Kehoe, M. W.

    1984-01-01

    A ground vibration test was conducted on an F-16 airplane loaded on each wing with a 370-gal tank mounted on a standard pylon, a GBU-8 store mounted on a decoupler pylon, and an AIM-9J missile mounted on a wing-tip launcher. The decoupler pylon is a passive wing/store flutter-suppression device. The test was conducted prior to initial flight tests to determine the modal frequencies, mode shapes, and structural damping coefficients. The data presented include frequency response plots, force effect plots, and limited mode shape data.

  10. 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

  11. A finite volume method for calculating transonic potential flow around wings from the pressure minimum integral

    NASA Technical Reports Server (NTRS)

    Eberle, A.

    1978-01-01

    Analysis of the pressure minimum integral in the calculation of three-dimensional potential flow around wings makes it possible to use non-rectangular mesh networks for distributing the three-dimensional potential into discrete points. The method is comparatively easily expanded to the treatment of realistic airplane configurations. Shock-pressure affected pressure distributions on any wings are determined with accuracy using this method.

  12. Inflatable wing

    SciTech Connect

    Priddy, T.G.

    1988-02-16

    An inflatable aerodynamic wing structure is described comprising: (a) an airfoil having at least two air-tight inflatable tubular enclosure means made of a first flexible material and extending along a spanwise axis; (b) top and bottom reinforcement member means made of a second stiff fabric material and connecting at least two air-tight inflatable tubular enclosure means together for transfer of inflation pressure-induced tensile stress from the enclosure means to the top and bottom reinforcement member means; (c) rigid hoops shaped to provide airfoil definition and spaced from each other along the spanwise axis and extending generally perpendicular thereto, the air-tight inflatable tubular enclosure means extending through the airfoil definition hoops and fastened thereto through the top and bottom reinforcement member means, the rigid hoops collapsing into each other for stacked stowage upon deflation of the enclosure means; and (d) means for forming an airfoil outer surface, made of a third thin, flexible and collapsible material, about substantially the entire tubular enclosure means and the top and bottom reinforcement member means, such that the area of a cross-section of the tubular enclosure means is much smaller than the area of a cross-section of the airfoil outer surface.

  13. Aerodynamic design optimization of a fuel efficient high-performance, single-engine, business airplane

    NASA Technical Reports Server (NTRS)

    Holmes, B. J.

    1980-01-01

    A design study has been conducted to optimize a single-engine airplane for a high-performance cruise mission. The mission analyzed included a cruise speed of about 300 knots, a cruise range of about 1300 nautical miles, and a six-passenger payload (5340 N (1200 lb)). The purpose of the study is to investigate the combinations of wing design, engine, and operating altitude required for the mission. The results show that these mission performance characteristics can be achieved with fuel efficiencies competitive with present-day high-performance, single- and twin-engine, business airplanes. It is noted that relaxation of the present Federal Aviation Regulation, Part 23, stall-speed requirement for single-engine airplanes facilitates the optimization of the airplane for fuel efficiency.

  14. Evaluation of an aeroelastic model technique for predicting airplane buffet loads

    NASA Technical Reports Server (NTRS)

    Hanson, P. W.

    1973-01-01

    A wind-tunnel technique which makes use of a dynamically scaled aeroelastic model to predict full-scale airplane buffet loads during buffet boundary penetration is evaluated. A 1/8-scale flutter model of a fighter airplane with remotely controllable variable-sweep wings and trimming surfaces was used for the evaluation. The model was flown on a cable-mount system which permitted high lift forces comparable to those in maneuvering flight. Bending moments and accelerations due to buffet were measured on the flutter model and compared with those measured on the full-scale airplane in an independent flight buffet research study. It is concluded that the technique can provide valuable information on airplane buffet load characteristics not available from any other source except flight test.

  15. Ubiquitous Supercritical Wing Design Cuts Billions in Fuel Costs

    NASA Technical Reports Server (NTRS)

    2015-01-01

    A Langley Research Center engineer’s work in the 1960s and ’70s to develop a wing with better performance near the speed of sound resulted in a significant increase in subsonic efficiency. The design was shared with industry. Today, Renton, Washington-based Boeing Commercial Airplanes, as well as most other plane manufacturers, apply it to all their aircraft, saving the airline industry billions of dollars in fuel every year.

  16. Aerodynamic Characteristics of a Portion of the Horizontal Tail from a Douglas C-74 Airplane with Fabric-Covered Elevators

    NASA Technical Reports Server (NTRS)

    Perone, Angelo; Berthold, Cecil L.

    1947-01-01

    A Douglas C-74 airplane, during a test dive at about 0.525 Mach number, experienced uncontrollable longitudinal oscillations sufficient to cause shedding of the outer wing panels and the subsequent crash of the airplane. Tests of a section of the horizontal tail plane from a C-74 airplane were conducted in the Ames 16-foot high-speed wind tunnel to investigate the possibility of the tail as a contributing factor to the accident. The results of the investigations of fabric-covered elevators in various conditions of surface deformation are presented in this report.

  17. Flight evaluation of an advanced technology light twin-engine airplane (ATLIT)

    NASA Technical Reports Server (NTRS)

    Holmes, B. J.

    1977-01-01

    Project organization and execution, airplane description and performance predictions, and the results of the flight evaluation of an advanced technology light twin engine airplane (ATLIT) are presented. The ATLIT is a Piper PA-34-200 Seneca I modified by the installation of new wings incorporating the GA(W)-1 (Whitcomb) airfoil, reduced wing area, roll control spoilers, and full span Fowler flaps. The conclusions for the ATLIT evaluation are based on complete stall and roll flight test results and partial performance test results. The Stalling and rolling characteristics met design expectations. Climb performance was penalized by extensive flow separation in the region of the wing body juncture. Cruise performance was found to be penalized by a large value of zero lift drag. Calculations showed that, with proper attention to construction details, the improvements in span efficiency and zero lift drag would permit the realization of the predicted increases in cruising and maximum rate of climb performance.

  18. Flutter Clearance of the F-14A Variable-Sweep Transition Flight Experiment Airplane, Phase 2

    NASA Technical Reports Server (NTRS)

    Freudinger, Lawrence C.; Kehoe, Michael W.

    1990-01-01

    An F-14A aircraft was modified for use as the test-bed aircraft for the variable-sweep transition flight experiment (VSTFE) program. The VSTFE program was a laminar flow research program designed to measure the effects of wing sweep on laminar flow. The airplane was modified by adding an upper surface foam and fiberglass glove to the right wing. An existing left wing glove had been added for the previous phase of the program. Ground vibration and flight flutter testing were accomplished to verify the absence of aeroelastic instabilities within a flight envelope of Mach 0.9 or 450 knots, calibrated airspeed, whichever was less. Flight test data indicated satisfactory damping levels and trends for the elastic structural modes of the airplane. Ground vibration test data are presented along with in-flight frequency and damping estimates, time histories and power spectral densities of in-flight sensors, and pressure distribution data.

  19. Aircraft wing weight build-up methodology with modification for materials and construction techniques

    NASA Technical Reports Server (NTRS)

    York, P.; Labell, R. W.

    1980-01-01

    An aircraft wing weight estimating method based on a component buildup technique is described. A simplified analytically derived beam model, modified by a regression analysis, is used to estimate the wing box weight, utilizing a data base of 50 actual airplane wing weights. Factors representing materials and methods of construction were derived and incorporated into the basic wing box equations. Weight penalties to the wing box for fuel, engines, landing gear, stores and fold or pivot are also included. Methods for estimating the weight of additional items (secondary structure, control surfaces) have the option of using details available at the design stage (i.e., wing box area, flap area) or default values based on actual aircraft from the data base.

  20. F-16XL ship #1 wing close-up showing boundary layer detection Preston tubes

    NASA Technical Reports Server (NTRS)

    1995-01-01

    This photo shows the boundary layer Preston tubes mounted on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

  1. F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

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

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

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

  3. 14 CFR 36.7 - Acoustical change: Transport category large airplanes and jet airplanes.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... airplanes and jet airplanes. 36.7 Section 36.7 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... § 36.7 Acoustical change: Transport category large airplanes and jet airplanes. (a) Applicability. This section applies to all transport category large airplanes and jet airplanes for which an acoustical...

  4. 14 CFR 36.7 - Acoustical change: Transport category large airplanes and jet airplanes.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... airplanes and jet airplanes. 36.7 Section 36.7 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... § 36.7 Acoustical change: Transport category large airplanes and jet airplanes. (a) Applicability. This section applies to all transport category large airplanes and jet airplanes for which an acoustical...

  5. Static Aeroelastic Effects of Formation Flight for Slender Unswept Wings

    NASA Technical Reports Server (NTRS)

    Hanson, Curtis E.

    2009-01-01

    The static aeroelastic equilibrium equations for slender, straight wings are modified to incorporate the effects of aerodynamically-coupled formation flight. A system of equations is developed by applying trim constraints and is solved for component lift distribution, trim angle-of-attack, and trim aileron deflection. The trim values are then used to calculate the elastic twist distribution of the wing box. This system of equations is applied to a formation of two gliders in trimmed flight. Structural and aerodynamic properties are assumed for the gliders, and solutions are calculated for flexible and rigid wings in solo and formation flight. It is shown for a sample application of two gliders in formation flight, that formation disturbances produce greater twist in the wingtip immersed in the vortex than for either the opposing wingtip or the wings of a similar airplane in solo flight. Changes in the lift distribution, resulting from wing twist, increase the performance benefits of formation flight. A flexible wing in formation flight will require greater aileron deflection to achieve roll trim than a rigid wing.

  6. Comparative Analysis of Interference Pathloss Coupling Patterns on B-737 VS. B757 Airplanes

    NASA Technical Reports Server (NTRS)

    Jafri, Madiha; Ely, Jay; Vahala, Linda

    2005-01-01

    Portable wireless technology provides many benefits to modern day travelers. Over the years however, numerous reports have cited portable electronic devices (PEDs) as a possible cause of electromagnetic interference (EMI) to aircraft navigation and communication radio systems. PEDs may act as transmitters, both intentional and unintentional, and their signals may be detected by the various radio receiver antennas installed on the aircraft. Measurement of the radiated field coupling between passenger cabin locations and aircraft communication and navigation receivers, via their antennas is defined herein as interference path loss (IPL). IPL data is required for assessing the threat of PEDs to aircraft radios, and is very dependent upon airplane size, the interfering transmitter position within the airplane, and the location of the particular antenna for the aircraft system of concern. NASA Langley Research Center, Eagles Wings Inc., and United Airlines personnel performed extensive IPL measurements on several Boeing 737 airplanes. In the Spring of 2004, extensive IPL measurements were also taken on several Boeing 757 airplanes under a cooperative agreement between NASA Langley Research Center and Delta Airlines. The objective of this paper is to analyze IPL measurement data, to better understand the impact on coupling levels based on the different locations of the aircraft radio antennas on B-757 and B-737 airplanes, and to provide a basis for future fuzzy logic modeling of airplane IPL. This effort will build upon previous fuzzy modeling of IPL data for B-737 airplane data.

  7. Blended Buffet-Load-Alleviation System for Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Moses, Robert W.

    2005-01-01

    The capability of modern fighter airplanes to sustain flight at high angles of attack and/or moderate angles of sideslip often results in immersion of part of such an airplane in unsteady, separated, vortical flow emanating from its forebody or wings. The flows from these surfaces become turbulent and separated during flight under these conditions. These flows contain significant levels of energy over a frequency band coincident with that of low-order structural vibration modes of wings, fins, and control surfaces. The unsteady pressures applied to these lifting surfaces as a result of the turbulent flows are commonly denoted buffet loads, and the resulting vibrations of the affected structures are known as buffeting. Prolonged exposure to buffet loads has resulted in fatigue of structures on several airplanes. Damage to airplanes caused by buffeting has led to redesigns of airplane structures and increased support costs for the United States Air Force and Navy as well as the armed forces of other countries. Time spent inspecting, repairing, and replacing structures adversely affects availability of aircraft for missions. A blend of rudder-control and piezoelectric- actuator engineering concepts was selected as a basis for the design of a vertical-tail buffet-load-alleviation system for the F/A-18 airplane. In this system, the rudder actuator is used to control the response of the first tail vibrational mode (bending at a frequency near 15 Hz), while directional patch piezoelectric actuators are used to control the second tail vibrational mode (tip torsion at a frequency near 45 Hz). This blend of two types of actuator utilizes the most effective features of each. An analytical model of the aeroservoelastic behavior of the airplane equipped with this system was validated by good agreement with measured results from a full-scale ground test, flight-test measurement of buffet response, and an in-flight commanded rudder frequency sweep. The overall performance of the

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

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

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

  9. KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

    NASA Technical Reports Server (NTRS)

    Montoya, L. C.; Jacobs, P.; Flechner, S.; Sims, R.

    1982-01-01

    A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

  10. Automatic Stability of Airplanes

    NASA Technical Reports Server (NTRS)

    Haus, FR

    1932-01-01

    It is endeavored in this report to give a full outline of the problem of airplane stability and to classify the proposed solutions systematically. Longitudinal stability, which can be studied separately, is considered first. The combination of lateral and directional stabilities, which cannot be separated, will be dealt with later.

  11. Measurements of Rudder Moments on an Airplane During Flight

    NASA Technical Reports Server (NTRS)

    Heidelberg, Ing V

    1921-01-01

    Tests indicated that: 1) C airplanes with two struts are extremely susceptible to aileron maneuvers, slight alterations of the aileron sufficing to compensate great unequalized moments; 2) great unequalized moments can be produced or neutralized by the unequalized alternation of the angle of attack below the outer and inner struts. Adjustment below the outer strut is the more effective of the two. 3) When a load of bombs is suspended beyond the center of the airplane, below the wings, the bombs need not be dropped simultaneously. 4) The propeller wash of a wide open engine has considerable influence on the position and operation of the elevator. The elevator is more susceptible in flight with the engine running than in gliding flight. 5) Adjustable tail planes are not advisable for D airplanes, nor for the C type, but they are, on the other hand, to be recommended for large size and giant airplanes in which the center of gravity changes during flight. 6) The aileron values obtained by wind tunnel measurements are about 10 percent too low, though otherwise applicable. For the elevator, the results of such measurements should be taken as mean values between flight with the engine running and gliding flight.

  12. LANN wing design

    NASA Technical Reports Server (NTRS)

    Firth, G. C.

    1983-01-01

    The LANN wing is the result of a joint effort between Lockheed, the Air Force, NASA, and the Netherlands to measure unsteady pressures at transonic speeds. It is a moderate-aspect-ratio transport wing configuration. The wing was machined from NITRONIC 40 and has 12 percent thick supercritical airfoil sections.

  13. Development of a Composite Tailoring Procedure for Airplane Wings

    NASA Technical Reports Server (NTRS)

    Chattopadhyay, Aditi

    2000-01-01

    The quest for finding optimum solutions to engineering problems has existed for a long time. In modern times, the development of optimization as a branch of applied mathematics is regarded to have originated in the works of Newton, Bernoulli and Euler. Venkayya has presented a historical perspective on optimization in [1]. The term 'optimization' is defined by Ashley [2] as a procedure "...which attempts to choose the variables in a design process so as formally to achieve the best value of some performance index while not violating any of the associated conditions or constraints". Ashley presented an extensive review of practical applications of optimization in the aeronautical field till about 1980 [2]. It was noted that there existed an enormous amount of published literature in the field of optimization, but its practical applications in industry were very limited. Over the past 15 years, though, optimization has been widely applied to address practical problems in aerospace design [3-5]. The design of high performance aerospace systems is a complex task. It involves the integration of several disciplines such as aerodynamics, structural analysis, dynamics, and aeroelasticity. The problem involves multiple objectives and constraints pertaining to the design criteria associated with each of these disciplines. Many important trade-offs exist between the parameters involved which are used to define the different disciplines. Therefore, the development of multidisciplinary design optimization (MDO) techniques, in which different disciplines and design parameters are coupled into a closed loop numerical procedure, seems appropriate to address such a complex problem. The importance of MDO in successful design of aerospace systems has been long recognized. Recent developments in this field have been surveyed by Sobieszczanski-Sobieski and Haftka [6].

  14. Curve Balls, Airplane Wings, and Prairie Dog Holes.

    ERIC Educational Resources Information Center

    Barnes, George B.

    1984-01-01

    Describes activities involving Bernoulli's principle which allows students to experience the difference between knowledge and scientific understanding. Explanations for each of the activities (using such materials as wooden spools, straws, soda bottles and table tennis balls) and explanations of phenomena in terms of Bernoulli's are provided. (BC)

  15. The Design of Plywood Webs for Airplane Wing Beams

    NASA Technical Reports Server (NTRS)

    Trayer, George W

    1931-01-01

    This report deals with the design of plywood webs for wooden box beams to obtain maximum strength per unit weight. A method of arriving at the most efficient and economical web thickness, and hence the most suitable unit shear stress, is presented and working stresses in shear for various types of webs and species of plywood are given. The questions of diaphragm spacing and required glue area between the webs and flange are also discussed.

  16. The Effect of the Slipstream on an Airplane Wing

    NASA Technical Reports Server (NTRS)

    Franke, A; Weinig, F

    1939-01-01

    The conditions which must be met at the slipstream boundary are developed, after which it is shown with the aid of the reflection method how these limiting conditions may be complied with in the case of an airfoil in a propeller slipstream in horizontal flow as well as for the propeller in yaw and with allowance for the slipstream rotation. In connection herewith, it is shown how the effective angles of attack and the circulation distribution with due regard to slipstream effect can be predicted and what inferences may be drawn therefrom for the distribution of lift, drag, and pitching moment.

  17. Distribution of Structural Weight of Wing Along the Span

    NASA Technical Reports Server (NTRS)

    Savelyev, V. V.

    1946-01-01

    In the present report the true weight distribution law of the wing structure along the span is investigated. It is shown that the triangular distribution and that based on the proportionality to the chords do not correspond to the actual weight distribution, On the basis of extensive data on wings of the CAHI type airplane formulas are obtained from which it is possible to determine the true diagram of the structural weight distribution along the span from a knowledge of only the geometrical dimensions of the wing. At the end of the paper data are presented showing how the structural weight is distributed between the straight center portion and the tapered portion as a function of their areas.

  18. A Theoretical Investigation of the Dynamic Lateral Stability Characteristics of the MX-838 (XB-51) Airplane

    NASA Technical Reports Server (NTRS)

    Paulson, Jon W.

    1948-01-01

    At the request of the Air Material Command, U. S. Air Force, a theoretical study has been made of the dynamic lateral stability characteristics of the MX-838 (XB-51) airplane. The calculations included the determination of the neutral-oscillatory-stability boundary (R = 0), the period and time to damp to one-half amplitude of the lateral oscillation, end the time to damp to one-half amplitude for the spiral mode. Factors varied in the investigation were lift coefficient, wing incidence, wing loading, and altitude. The results of the investigation showed that the lateral oscillation of the airplane is unstable below a lift coefficient of 1.2 with flaps . deflected 40deg but is stable over the entire speed range with flaps deflected 20deg or 0deg. The results showed that satisfactory oscillatory stability can probably be obtained for all lift coefficients with the proper variation of flap deflection and wing incidence with airspeed. Reducing the positive wing incidence improved the oscillatory stability characteristics. The airplane is spirally unstable for most conditions but the instability is mild and the Air Force requirements are easily met.

  19. Flapping of Insectile Wings

    NASA Astrophysics Data System (ADS)

    Huang, Yangyang; Kanso, Eva

    2015-11-01

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

  20. Exploring Venus by Solar Airplane

    NASA Technical Reports Server (NTRS)

    Landis, Geoffrey A.

    2001-01-01

    A solar-powered airplane is proposed to explore the atmospheric environment of Venus. Venus has several advantages for a solar airplane. At the top of the cloud level, the solar intensity is comparable to or greater than terrestrial solar intensities. The Earthlike atmospheric pressure means that the power required for flight is lower for Venus than that of Mars, and the slow rotation of Venus allows an airplane to be designed for continuous sunlight, with no energy storage needed for night-time flight. These factors mean that Venus is perhaps the easiest planet in the solar system for flight of a long-duration solar airplane.

  1. Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane

    NASA Technical Reports Server (NTRS)

    Mason, Homer P.

    1953-01-01

    Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

  2. High transonic speed transport aircraft study. [aerodynamic characteristics of single-fuselage, yawed-wing configuration

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.; Neumann, F. D.; Nisbet, J. W.; Mulally, A. R.; Murakami, J. K.; Noble, E. C.; Mcbarron, J. P.; Stalter, J. L.; Gimmestad, D. W.; Sussman, M. B.

    1973-01-01

    An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range=5560 Km (3000 nmi), payload-18 143 kg (40 000lb), Mach=1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211 828 Kg (467 000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226 796 Kg (500 000 lb). Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-phase development plan is recommended to establish the full potential of the yawed-wing concept.

  3. Wind-Tunnel Investigation of Subsonic Longitudinal Aerodynamic Characteristics of a Tiltable-Wing Vertical-Take-Off-and-Landing Supersonic Bomber Configuration Including Turbojet Power Effects

    NASA Technical Reports Server (NTRS)

    Thompson, Robert F.; Vogler, Raymond D.; Moseley, William C., Jr.

    1959-01-01

    Jet-powered model tests were made to determine the low-speed longitudinal aerodynamic characteristics of a vertical-take-off and-landing supersonic bomber configuration. The configuration has an unique engine-wing arrangement wherein six large turbojet engines (three on each side of the fuselage) are buried in a low-aspect-ratio wing which is tilted into the vertical plane for take-off. An essentially two-dimensional variable inlet, spanning the leading edge of each wing semispan, provides air for the engines. Jet flow conditions were simulated for a range of military (nonafterburner) and afterburner turbojet-powered flight at subsonic speeds. Three horizontal tails were tested at a station down-stream of the jet exit and at three heights above the jet axes. A semi-span model was used and test parameters covered wing-fuselage incidence angles from 0 deg to 15 deg, wing angles of attack from -4 deg to 36 deg, a variable range of horizontal-tail incidence angles, and some variations in power simulation conditions. Results show that, with all horizontal tails tested, there were large variations in static stability throughout the lift range. When the wing and fuselage were alined, the model was statically stable throughout the test range only with the largest tail tested (tail span of 1.25 wing span) and only when the tail was located in the low test position which placed the tail nearest to the undeflected jet. For transition flight conditions, none of the tail configurations provided satisfactory longitudinal stability or trim throughout the lift range. Jet flow was destabilizing for most of the test conditions, and varying the jet-exit flow conditions at a constant thrust coefficient had little effect on the stability of this model. Wing leading-edge simulation had some important effects on the longitudinal aerodynamic characteristics.

  4. Analysis of the effects of wing interference on the tail contributions to the rolling derivatives

    NASA Technical Reports Server (NTRS)

    Michael, William H , Jr

    1952-01-01

    An analysis of the effects of wing interference on the tail contributions to the rolling stability derivatives of complete airplane configurations is made by calculating the angularity of the air stream at the vertical tail due to rolling and determining the resulting forces and moments. Some of the important factors which affect the resultant angularity on the vertical tail are wing aspect ratio and sweepback, vertical-tail span, and considerations associated with angle of attack and airplane geometry. Some calculated sidewash results for a limited range of plan forms and vertical-tail sizes are presented. Equations taking into account the sidewash results are given for determining the tail contributions to the rolling derivatives. Comparisons of estimated and experimental results indicate that a consideration of wing interference effects improves the estimated values of the tail contributions to the rolling derivatives and that fair agreement with available experimental data is obtained.

  5. 77 FR 9166 - Airworthiness Directives; Airplanes Originally Manufactured by Lockheed for the Military as P2V...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-02-16

    ... Policies and Procedures (44 FR 11034, February 26, 1979), (3) Will not affect intrastate aviation in Alaska... prompted by a report of a significant crack in the principle wing structure on a Neptune Aviation Service... structure on a Neptune Aviation Service, Inc. Model SP-2H (P2V-7) airplane. This condition, if not...

  6. Flight Measurements of the Flying Qualities of a Lockheed P-80A Airplane (Army No. 44-85099) - Stalling Characteristics

    NASA Technical Reports Server (NTRS)

    Anderson, Seth B.; Cooper, George E.

    1947-01-01

    This report contains the flight-test results of the stalling characteristics measured during the flying-qualities investigation of the Lockheed P-8OA airplane (Army No. 44-85099). The tests were conducted in straight and turning flight with and without wing-tip tanks. These tests showed satisfactory stalling characteristics and adequate stall warning for all configurations and conditions tested.

  7. Mechanical control of airplanes

    NASA Technical Reports Server (NTRS)

    Boykow, H

    1929-01-01

    Before undertaking a detailed description of an automatic-control mechanism, I will state briefly the fundamental conditions for such devices. These are: 1) it must be sensitive at one or more reference values; 2) it must stop the angular motions of the airplane not produced by the pilot; and 3) it must be possible to switch it off and on by a simple hand lever.

  8. Strength calculations on airplanes

    NASA Technical Reports Server (NTRS)

    Baumann, A

    1925-01-01

    Every strength calculation, including those on airplanes, must be preceded by a determination of the forces to be taken into account. In the following discussion, it will be assumed that the magnitudes of these forces are known and that it is only a question of how, on the basis of these known forces, to meet the prescribed conditions on the one hand and the practical requirements on the other.

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

    NASA Technical Reports Server (NTRS)

    Jenkins, Jerald M.; DeAngelis, V. Michael

    1997-01-01

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

  10. Airplane dopes and doping

    NASA Technical Reports Server (NTRS)

    Smith, W H

    1919-01-01

    Cellulose acetate and cellulose nitrate are the important constituents of airplane dopes in use at the present time, but planes were treated with other materials in the experimental stages of flying. The above compounds belong to the class of colloids and are of value because they produce a shrinking action on the fabric when drying out of solution, rendering it drum tight. Other colloids possessing the same property have been proposed and tried. In the first stages of the development of dope, however, shrinkage was not considered. The fabric was treated merely to render it waterproof. The first airplanes constructed were covered with cotton fabric stretched as tightly as possible over the winds, fuselage, etc., and flying was possible only in fine weather. The necessity of an airplane which would fly under all weather conditions at once became apparent. Then followed experiments with rubberized fabrics, fabrics treated with glue rendered insoluble by formaldehyde or bichromate, fabrics treated with drying and nondrying oils, shellac, casein, etc. It was found that fabrics treated as above lost their tension in damp weather, and the oil from the motor penetrated the proofing material and weakened the fabric. For the most part the film of material lacked durability. Cellulose nitrate lacquers, however were found to be more satisfactory under varying weather conditions, added less weight to the planes, and were easily applied. On the other hand, they were highly inflammable, and oil from the motor penetrated the film of cellulose nitrate, causing the tension of the fabric to be relaxed.

  11. Wind Tunnel Development of Means to Alleviate Buffeting of the North American XP-82 Airplane at High Speeds

    NASA Technical Reports Server (NTRS)

    Anderson, Joseph L.

    1947-01-01

    This report presents the results of wind-tunnel tests of a 0.22-scale model of the North American XP-82 airplane with several modifications designed to reduce the buffeting of the airplane. The effects of various modifications on the air flow over the model are shown by means of photographs of tufts. The drag, lift, and pitching-moment coefficients of the model with several of the modifications are shown. The result indicate that, by reflexing the trailing edge of the center section of the wing and modifying the radiator air-scoop gutter and the inboard lower-surface wing fillets, the start of buffeting can be delayed from a Mach number of 0.70 to 0.775, and that the diving tendency of the airplane would be eliminated up to a Mach number of 0.80.

  12. Development and flight tests of a gyro-less wing leveler and directional autopilot

    NASA Technical Reports Server (NTRS)

    Garner, H. D.; Poole, H. E.

    1974-01-01

    A gyro-less wing leveler and directional autopilot were developed and flight tested in a single-engine light airplane. The primary purpose of the project was to develop a simple, reliable, low-cost stability augmentation and autopilot system for light aircraft. The wing leveler used a fluidic inertial rate sensor, electronic signal processing circuitry, and vacuum operated servos. A strap-down magnetic heading reference of simple design provided the wing leveler with directional autopilot capability. Flight tests indicated that the performance of the gyro-less wing leveler was equal to that of a commercial wing leveler using a gyroscopic rate sensor. Drift-free, long-term, heading-hold capability of the magnetic heading reference was demonstrated.

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

    NASA Technical Reports Server (NTRS)

    Bihrle, W., Jr.; Hultberg, R. S.

    1979-01-01

    Aerodynamic characteristics obtained in a rotational flow environment utilizing a rotary balance located in the spin tunnel are presented in plotted form for a 1/6.5 scale, single engine, low wing, general aviation airplane model. The configurations tested included the basic airplane, various wing leading-edge devices, tail designs, and rudder control settings as well as airplane components. Data are presented without analysis for an angle-of-attack range of 8 deg to 90 deg and clockwise and counter-clockwise rotations covering an (omega)(b)/2V range from 0 to 0.85.

  14. The influence of lateral stability on disturbed motions of an airplane with special reference to the motions produced by gusts

    NASA Technical Reports Server (NTRS)

    Jones, Robert T

    1938-01-01

    Distributed lateral motions have been calculated for a hypothetical small airplane with various modifications of fin area and dihedral setting. Special combinations of disturbing factors to simulate gusts are considered and the influence of lateral stability on the motions is discussed. Fin area and wing dihedral were found to be of primary importance in side gusts. It was found that the rolling action of the wing with as much as 5 degrees dihedral was distinctly unfavorable, especially when the weathercock stability was small. It is pointed out that the greatest susceptibility to lateral disturbances lies in the inherent damping and coupling moments developed by the wing.

  15. X-Wing Research Vehicle in Hangar

    NASA Technical Reports Server (NTRS)

    1987-01-01

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

  16. Trend of airplane flight characteristics

    NASA Technical Reports Server (NTRS)

    Von Koppen, Joachim

    1933-01-01

    This report describes the development of airplane characteristics since the war and indicates the direction development should take in the immediate future. Some of the major topics include: the behavior of an airplane about its lateral, vertical, and longitudinal axes. Behavior at large angles of attack and landing characteristics are also included.

  17. 77 FR 6023 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-02-07

    ... adequacy of existing regulations, the service history of airplanes subject to those regulations, and... Federal Aviation Administration 14 CFR Part 39 RIN 2120-AA64 Airworthiness Directives; Airbus Airplanes... airplanes and Model A310-203, -204, - 221, and -222 airplanes. This proposed AD was prompted by a report...

  18. 14 CFR 125.93 - Airplane limitations.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane limitations. 125.93 Section...

  19. 14 CFR 23.3 - Airplane categories.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Airplane categories. 23.3 Section 23.3... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES General § 23.3 Airplane categories. (a) The normal category is limited to airplanes that have a seating configuration, excluding...

  20. 14 CFR 23.3 - Airplane categories.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Airplane categories. 23.3 Section 23.3... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES General § 23.3 Airplane categories... airplanes that have a seating configuration, excluding pilot seats, of nine or less, a maximum...

  1. 14 CFR 23.3 - Airplane categories.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Airplane categories. 23.3 Section 23.3... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES General § 23.3 Airplane categories. (a) The normal category is limited to airplanes that have a seating configuration, excluding...

  2. A preliminary design study of a laminar flow control wing of composite materials for long range transport aircraft

    NASA Technical Reports Server (NTRS)

    Swinford, G. R.

    1976-01-01

    The results of an aircraft wing design study are reported. The selected study airplane configuration is defined. The suction surface, ducting, and compressor systems are described. Techniques of manufacturing suction surfaces are identified and discussed. A wing box of graphite/epoxy composite is defined. Leading and trailing edge structures of composite construction are described. Control surfaces, engine installation, and landing gear are illustrated and discussed. The preliminary wing design is appraised from the standpoint of manufacturing, weight, operations, and durability. It is concluded that a practical laminar flow control (LFC) wing of composite material can be built, and that such a wing will be lighter than an equivalent metal wing. As a result, a program of suction surface evaluation and other studies of configuration, aerodynamics, structural design and manufacturing, and suction systems are recommended.

  3. Flight investigation of the effect of control centering springs on the apparent spiral stability of a personal-owner airplane

    NASA Technical Reports Server (NTRS)

    Campbell, John P; Hunter, Paul A; Hewes, Donald E; Whitten, James B

    1952-01-01

    Report presents the results of a flight investigation conducted on a typical high-wing personal-owner airplane to determine the effect of control centering springs on apparent spiral stability. Apparent spiral stability is the term used to describe the spiraling tendencies of an airplane in uncontrolled flight as affected both by the true spiral stability of the perfectly trimmed airplane and by out-of-trim control settings. Centering springs were used in both the aileron and rudder control systems to provide both a positive centering action and a means of trimming the airplane. The springs were preloaded so that when they were moved through neutral they produced a nonlinear force gradient sufficient to overcome the friction in the control surface at the proper setting for trim. The ailerons and rudder control surfaces did not have trim tabs that could be adjusted in flight.

  4. Optimum hovering wing planform.

    PubMed

    Nabawy, Mostafa R A; Crowther, William J

    2016-10-01

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

  5. Propeller/wing interaction

    NASA Technical Reports Server (NTRS)

    Witkowski, David P.; Johnston, Robert T.; Sullivan, John P.

    1989-01-01

    The present experimental investigation of the steady-state and unsteady-state effects due to the interaction between a tractor propeller's wake and a wing employs, in the steady case, wind tunnel measurements at low subsonic speed; results are obtained which demonstrate wing performance response to variations in configuration geometry. Other steady-state results involve the propeller-hub lift and side-force due to the wing's influence on the propeller. The unsteady effects of interaction were studied through flow visualization of propeller-tip vortex distortion over a wing, again using a tractor-propeller configuration.

  6. Feasibility and benefits of laminar flow control on supersonic cruise airplanes

    NASA Technical Reports Server (NTRS)

    Powell, A. G.; Agrawal, S.; Lacey, T. R.

    1989-01-01

    An evaluation was made of the applicability and benefits of laminar flow control (LFC) technology to supersonic cruise airplanes. Ancillary objectives were to identify the technical issues critical to supersonic LFC application, and to determine how those issues can be addressed through flight and wind-tunnel testing. Vehicle types studied include a Mach 2.2 supersonic transport configuration, a Mach 4.0 transport, and two Mach 2-class fighter concepts. Laminar flow control methodologies developed for subsonic and transonic wing laminarization were extended and applied. No intractible aerodynamic problems were found in applying LFC to airplanes of the Mach 2 class, even ones of large size. Improvements of 12 to 17 percent in lift-drag ratios were found. Several key technical issues, such as contamination avoidance and excresence criteria were identified. Recommendations are made for their resolution. A need for an inverse supersonic wing design methodology is indicated.

  7. Preliminary flight-test results of an advanced technology light twin-engine airplane /ATLIT/

    NASA Technical Reports Server (NTRS)

    Holmes, B. J.; Kohlman, D. L.; Crane, H. L.

    1976-01-01

    The present status and flight-test results are presented for the ATLIT airplane. The ATLIT is a Piper PA-34 Seneca I modified by the installation of new wings incorporating the GA(W)-1 (Whitcomb) airfoil, reduced wing area, roll-control spoilers, and full-span Fowler flaps. Flight-test results on stall and spoiler roll characteristics show good agreement with wind-tunnel data. Maximum power-off lift coefficients are greater than 3.0 with flaps deflected 37 deg. With flaps down, spoiler deflections can produce roll helix angles in excess of 0.11 rad. Flight testing is planned to document climb and cruise performance, and supercritical propeller performance and noise characteristics. The airplane is scheduled for testing in the NASA-Langley Research Center Full-Scale Tunnel.

  8. Graphical and Statistical Analysis of Airplane Passenger Cabin RF Coupling Paths to Avionics

    NASA Technical Reports Server (NTRS)

    Jafri, Madiha; Ely, Jay; Vahala, Linda

    2003-01-01

    Portable wireless technology provides many benefits to modern day travelers. Over the years however, numerous reports have cited portable electronic devices (PEDs) as a possible cause of electromagnetic interference (EMI) to aircraft navigation and communication radio systems. PEDs may act as transmitters, both intentional and unintentional, and their signals may be detected by the various radio receiver antennas installed on the aircraft. Measurement of the radiated field coupling between passenger cabin locations and aircraft communication and navigation receivers, via their antennas is defined herein as interference path loss (IPL). IPL data is required for assessing the threat of PEDs to aircraft radios, and is very dependent upon airplane size, the interfering transmitter position within the airplane, and the location of the particular antenna for the aircraft system of concern. NASA Langley Research Center, Eagles Wings Inc., and United Airlines personnel performed extensive IPL measurements on several Boeing 737 airplanes.

  9. Investigation of the Forces Acting on Gliders in Automobile-pulley-winch and Airplane Towed Flight

    NASA Technical Reports Server (NTRS)

    Klemperer, W B

    1942-01-01

    The magnitude, the direction, and the fluctuation of towing forces exerted upon gliders by towing them aloft behind an automobile, by means of a winch, and by airplane were measured under a variety of conditions covering a range from gentle to severe types of operation. For these tests the towing forces did not exceed 92 percent of the gross weight of the glider. The results indicate that in pulley and winch towing the towing forces are of about the same magnitude as in automobile towing. Speed increases in the accelerated phases of the towing jerks encountered in airplane towing can readily become critical as speeds in excess of placard speeds can be attained. Passage through the slipstream of the towing airplane can be equivalent to a severe gust that, at high speed, may impose high wing loads and require large control moments.

  10. 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 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, single engine, and twin-engine design. 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 was used successfully in flight 17 times.

  11. An analytical study of the effect of airplane wake on the lateral dispersion of aerial sprays

    NASA Technical Reports Server (NTRS)

    Reed, Wilmer H , III

    1954-01-01

    Calculations are made to determine the trajectories of liquid droplets introduced into the air disturbances generated by an airplane engaged in aerial spraying. The effects of such factors as the positions at which droplets are ejected into the disturbances, airplane lift coefficient, and altitude are investigated. The distribution of deposit on the ground is computed for several droplet-size spectra, variations in the rate at which mass is ejected along the span, and lateral flight-path spacings. Consideration is then given to the problem of adjusting these factors with the aim of improving the uniformity and increasing the effective width of the deposit. The results indicate that the lateral dispersion of droplets is increased when the spanwise position at which particles are ejected is moved toward the wing tip. Greater dispersion also results when the airplane lift coefficient or altitude is increased.

  12. Linear instability in the wake of an elliptic wing

    NASA Astrophysics Data System (ADS)

    He, Wei; Tendero, Juan Ángel; Paredes, Pedro; Theofilis, Vassilis

    2016-07-01

    Linear global instability analysis has been performed in the wake of a low aspect ratio three-dimensional wing of elliptic cross section, constructed with appropriately scaled Eppler E387 airfoils. The flow field over the airfoil and in its wake has been computed by full three-dimensional direct numerical simulation at a chord Reynolds number of Rec=1750 and two angles of attack, {AoA}=0° and 5°. Point-vortex methods have been employed to predict the inviscid counterpart of this flow. The spatial BiGlobal eigenvalue problem governing linear small-amplitude perturbations superposed upon the viscous three-dimensional wake has been solved at several axial locations, and results were used to initialize linear PSE-3D analyses without any simplifying assumptions regarding the form of the trailing vortex system, other than weak dependence of all flow quantities on the axial spatial direction. Two classes of linearly unstable perturbations were identified, namely stronger-amplified symmetric modes and weaker-amplified antisymmetric disturbances, both peaking at the vortex sheet which connects the trailing vortices. The amplitude functions of both classes of modes were documented, and their characteristics were compared with those delivered by local linear stability analysis in the wake near the symmetry plane and in the vicinity of the vortex core. While all linear instability analysis approaches employed have delivered qualitatively consistent predictions, only PSE-3D is free from assumptions regarding the underlying base flow and should thus be employed to obtain quantitative information on amplification rates and amplitude functions in this class of configurations.

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

    NASA Technical Reports Server (NTRS)

    Hultberg, R. S.; Mulcay, W.

    1980-01-01

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

  14. Aerodynamic Classification of Swept-Wing Ice Accretion

    NASA Technical Reports Server (NTRS)

    Broeren, Andy; Diebold, Jeff; Bragg, Mike

    2013-01-01

    The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current state-of-the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice, and spanwise-ridge ice. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.

  15. Commercial airplanes and seaplanes : thick wings or thin wings - all metal or mixed construction

    NASA Technical Reports Server (NTRS)

    POINT

    1922-01-01

    In this report we will consider, as the dominating characteristic, either the load carried, the speed, the radius of action, the fuel consumption, the activity of transport, or, lastly, the qualities of comfort and safety. The first four factors determine the theoretical efficiency, while the others determine its practical efficiency.

  16. Tests of Wing Machine-Gun and Cannon Installations in the NACA Full-Scale Wind Tunnel, Special Report

    NASA Technical Reports Server (NTRS)

    Czarnecki, K. R.; Guryansky, Eugene R.

    1941-01-01

    At the request of the Bureau of Aeronautics, an investigation was conducted in the full-scale wind tunnel of wing installations of .50-caliber machine guns and 20-millimeter cannons. The tests were made to determine the effect of various gun installations on the maximum lift and the high-speed drag of the airplane.

  17. WINGDES2 - WING DESIGN AND ANALYSIS CODE

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.

    1994-01-01

    This program provides a wing design algorithm based on modified linear theory which takes into account the effects of attainable leading-edge thrust. A primary objective of the WINGDES2 approach is the generation of a camber surface as mild as possible to produce drag levels comparable to those attainable with full theoretical leading-edge thrust. WINGDES2 provides both an analysis and a design capability and is applicable to both subsonic and supersonic flow. The optimization can be carried out for designated wing portions such as leading and trailing edge areas for the design of mission-adaptive surfaces, or for an entire planform such as a supersonic transport wing. This program replaces an earlier wing design code, LAR-13315, designated WINGDES. WINGDES2 incorporates modifications to improve numerical accuracy and provides additional capabilities. A means of accounting for the presence of interference pressure fields from airplane components other than the wing and a direct process for selection of flap surfaces to approach the performance levels of the optimized wing surfaces are included. An increased storage capacity allows better numerical representation of those configurations that have small chord leading-edge or trailing-edge design areas. WINGDES2 determines an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. The objective of the design is the recovery of unrealized theoretical leading-edge thrust of the input flat surface by shaping of the design surface to create a distributed thrust and thus minimize drag. The input consists of airfoil section thickness data, leading and trailing edge planform geometry, and operational parameters such as Mach number, Reynolds number, and design lift coefficient. Output includes optimized camber surface ordinates, pressure coefficient distributions, and theoretical aerodynamic characteristics. WINGDES2 is written in FORTRAN V for batch execution and has been

  18. An integrated study of structures, aerodynamics and controls on the forward swept wing X-29A and the oblique wing research aircraft

    NASA Technical Reports Server (NTRS)

    Dawson, Kenneth S.; Fortin, Paul E.

    1987-01-01

    The results of an integrated study of structures, aerodynamics, and controls using the STARS program on two advanced airplane configurations are presented. Results for the X-29A include finite element modeling, free vibration analyses, unsteady aerodynamic calculations, flutter/divergence analyses, and an aeroservoelastic controls analysis. Good correlation is shown between STARS results and various other verified results. The tasks performed on the Oblique Wing Research Aircraft include finite element modeling and free vibration analyses.

  19. Wind-tunnel studies of the effects of stimulated damage on the aerodynamic characteristics of airplanes and missiles

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.

    1982-01-01

    As an aid in assessing the aerodynamic effects of battle damage that might be sustained by military airplanes or missiles, several wind tunnel investigations were performed at the Langley Research Center in which damage was simulated with models by the removal of all or parts of the wing and tails. Results of the investigations indicate that the loss of a major part of the vertical tail will probably result in the loss of an airplane in any speed range. The loss of major parts of the horizontal tail generally results in catastrophic instability in the subsonic range but, at low supersonic speeds, and for some planform configurations at subsonic speeds, may allow stable flight to the extent that the airplane might return to friendly territory before the pilot must eject. The results further indicate that major damage to the wing, up to the point of the complete removal of one wing panel, and major damage to the horizontal tail may be sustained without necessarily causing the loss of the airplane or pilot.

  20. Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Seamless Flaps

    NASA Technical Reports Server (NTRS)

    Bui, Trong T.

    2016-01-01

    Computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a Gulfstream G-III airplane (Gulfstream Aerospace Corporation, Savannah, Georgia) swept wing modified with an experimental seamless, compliant flap called the Adaptive Compliant Trailing Edge (ACTE) flap. The stall characteristics of the modified ACTE wing were analyzed and compared with the unmodified, clean wing at the flight speed of 120 knots and altitude of 2300 feet above mean sea level, in free air as well as in ground effect. A polyhedral finite-volume unstructured full Navier-Stokes CFD code, STAR-CCM (registered trademark) plus (CD-adapco [Computational Dynamics Limited, United Kingdom, and Analysis & Design Application Co., United States]), was used. Steady Reynolds-averaged Navier-Stokes CFD simulations were conducted for a clean wing and the ACTE wings at various ACTE deflection angles in free air (-2 degrees, 15 degrees, and 30 degrees) as well as in ground effect (15 degrees and 30 degrees). Solution sensitivities to grid densities were examined. In free air, the ACTE wings are predicted to stall at lower angles of attack than the clean wing. In ground effect, all wings are predicted to stall at lower angles of attack than the corresponding wings in free air. Even though the lift curves are higher in ground effect than in free air, the maximum lift coefficients for all wings are lower in ground effect. Finally, the lift increase due to ground effect for the ACTE wing is predicted to be less than the clean wing.

  1. The structure of airplane fabrics

    NASA Technical Reports Server (NTRS)

    Walen, E Dean

    1920-01-01

    This report prepared by the Bureau of Standards for the National Advisory Committee for Aeronautics supplies the necessary information regarding the apparatus and methods of testing and inspecting airplane fabrics.

  2. Gamma rays at airplane altitudes

    SciTech Connect

    Iwai, J.; Koss, T.; Lord, J.; Strausz, S.; Wilkes, J.; Woosley, J. )

    1990-03-20

    An examination of the gamma ray flux above 1 TeV in the atmosphere is needed to better understand the anomalous showers from point sources. Suggestions are made for future experiments on board airplanes.

  3. Wing chord extension on D-558-2

    NASA Technical Reports Server (NTRS)

    1953-01-01

    This 1953 NACA High-Speed Flight Research Station photograph shows the wing chord extensions added to D-558-2 #3. The added width on the outer wing panels was one attempt made to solve the tendence of the D-558-2 aircraft to 'pitch-up' during flight. Other modifications tested included fixed and free-floating slats, one- or two-wing fences, and a slat and wing fence. None of these additions proved truly effective, and the chord extension actually aggravated the pitch-up. The research program nevertheless added greatly to the store of knowledge about how to deal with pitch-up (as well as how not to deal with it, itself valuable information). The best solution was to place the horizontal stabilizer low on the aft fuselage, where it would be below the downwash from the wing and the wing wake. The Douglas D-558-2 'Skyrockets' were among the early transonic research airplanes like the X-1, X-4, X-5, and X-92A. Three of the single-seat, swept-wing aircraft flew from 1948 to 1956 in a joint program involving the National Advisory Committee for Aeronautics (NACA), with its flight research done at the NACA's Muroc Flight Test Unit in Calif., redesignated in 1949 the High-Speed Flight Research Station (HSFRS). Also partners in the flight research were the Navy-Marine Corps and the Douglas Aircraft Co. The HSFRS became the High-Speed Flight Station in 1954 and is now known as the NASA Dryden Flight Research Center. The Skyrocket made aviation history when it became the first airplane to fly twice the speed of sound. The 2 in the aircraft's designation referred to the fact that the Skyrocket was the phase-two version of what had originally been conceived as a three-phase program, with the phase-one aircraft having straight wings. The third phase, which never came to fruition, would have involved constructing a mock-up of a combat-type aircraft embodying the results from the testing of the phase one and two aircraft. Douglas pilot John F. Martin made the first flight at Muroc

  4. Differential equations in airplane mechanics

    NASA Technical Reports Server (NTRS)

    Carleman, M T

    1922-01-01

    In the following report, we will first draw some conclusions of purely theoretical interest, from the general equations of motion. At the end, we will consider the motion of an airplane, with the engine dead and with the assumption that the angle of attack remains constant. Thus we arrive at a simple result, which can be rendered practically utilizable for determining the trajectory of an airplane descending at a constant steering angle.

  5. Wing-kinematics measurement and aerodynamics in a small insect in hovering flight.

    PubMed

    Cheng, Xin; Sun, Mao

    2016-01-01

    Wing-motion of hovering small fly Liriomyza sativae was measured using high-speed video and flows of the wings calculated numerically. The fly used high wingbeat frequency (≈265 Hz) and large stroke amplitude (≈182°); therefore, even if its wing-length (R) was small (R ≈ 1.4 mm), the mean velocity of wing reached ≈1.5 m/s, the same as that of an average-size insect (R ≈ 3 mm). But the Reynolds number (Re) of wing was still low (≈40), owing to the small wing-size. In increasing the stroke amplitude, the outer parts of the wings had a "clap and fling" motion. The mean-lift coefficient was high, ≈1.85, several times larger than that of a cruising airplane. The partial "clap and fling" motion increased the lift by ≈7%, compared with the case of no aerodynamic interaction between the wings. The fly mainly used the delayed stall mechanism to generate the high-lift. The lift-to-drag ratio is only 0.7 (for larger insects, Re being about 100 or higher, the ratio is 1-1.2); that is, although the small fly can produce enough lift to support its weight, it needs to overcome a larger drag to do so. PMID:27168523

  6. Wing-kinematics measurement and aerodynamics in a small insect in hovering flight.

    PubMed

    Cheng, Xin; Sun, Mao

    2016-05-11

    Wing-motion of hovering small fly Liriomyza sativae was measured using high-speed video and flows of the wings calculated numerically. The fly used high wingbeat frequency (≈265 Hz) and large stroke amplitude (≈182°); therefore, even if its wing-length (R) was small (R ≈ 1.4 mm), the mean velocity of wing reached ≈1.5 m/s, the same as that of an average-size insect (R ≈ 3 mm). But the Reynolds number (Re) of wing was still low (≈40), owing to the small wing-size. In increasing the stroke amplitude, the outer parts of the wings had a "clap and fling" motion. The mean-lift coefficient was high, ≈1.85, several times larger than that of a cruising airplane. The partial "clap and fling" motion increased the lift by ≈7%, compared with the case of no aerodynamic interaction between the wings. The fly mainly used the delayed stall mechanism to generate the high-lift. The lift-to-drag ratio is only 0.7 (for larger insects, Re being about 100 or higher, the ratio is 1-1.2); that is, although the small fly can produce enough lift to support its weight, it needs to overcome a larger drag to do so.

  7. Wing-kinematics measurement and aerodynamics in a small insect in hovering flight

    PubMed Central

    Cheng, Xin; Sun, Mao

    2016-01-01

    Wing-motion of hovering small fly Liriomyza sativae was measured using high-speed video and flows of the wings calculated numerically. The fly used high wingbeat frequency (≈265 Hz) and large stroke amplitude (≈182°); therefore, even if its wing-length (R) was small (R ≈ 1.4 mm), the mean velocity of wing reached ≈1.5 m/s, the same as that of an average-size insect (R ≈ 3 mm). But the Reynolds number (Re) of wing was still low (≈40), owing to the small wing-size. In increasing the stroke amplitude, the outer parts of the wings had a “clap and fling” motion. The mean-lift coefficient was high, ≈1.85, several times larger than that of a cruising airplane. The partial “clap and fling” motion increased the lift by ≈7%, compared with the case of no aerodynamic interaction between the wings. The fly mainly used the delayed stall mechanism to generate the high-lift. The lift-to-drag ratio is only 0.7 (for larger insects, Re being about 100 or higher, the ratio is 1–1.2); that is, although the small fly can produce enough lift to support its weight, it needs to overcome a larger drag to do so. PMID:27168523

  8. Slotted Aircraft Wing

    NASA Technical Reports Server (NTRS)

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

    2006-01-01

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

  9. Flying wings / flying fuselages

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  10. A Parametric Study of Accelerations of an Airplane Due to a Wake Vortex System

    NASA Technical Reports Server (NTRS)

    Stewart, Eric C.

    1999-01-01

    A study was conducted using strip theory to systematically investigate the effects of progressively more complete descriptions of the interaction of an airplane with a wake vortex system. The emphasis was in roll-dominant, parallel, vortex encounters. That is, the simulated airplane's longitudinal axis was nearly parallel to the rotation axis of the vortex system for most of the results presented. The study began with a drag-less rectangular wing in the flow field of a single vortex and progressed to a complete airplane with aerodynamic surfaces possessing taper, sweep, dihedral, and stalling and immersed in the flow field of a vortex pair in ground effect. The effects of the pitch, roll, and yaw attitudes of the airplane on the calculated accelerations were also investigated. The airplane had the nominal characteristics of a Boeing 757, and the vortex flow field had the nominal characteristics of the wake of a Boeing 767. The Bumham-Hallock model of a vortex flow field was used throughout the study. The data are presented mainly in terms of contours of equal acceleration in a two-dimensional area centered on the vortex pair and having dimensions of 300 feet by 300 feet.

  11. Prediction of airplane cabin noise due to engine shock cell excitation using statistical energy analysis

    NASA Astrophysics Data System (ADS)

    Marshall, Steven E.

    As part of the effort in the 1980's to design fuel efficient propulsion systems (unducted fan engines) for large commercial airplanes, procedures were developed for predicting interior noise using statistical energy analysis (SEA). Due to stable fuel process and deregulation in the airline industry, the emphasis for propulsion systems on commercial airplanes shifted to higher thrust and lower operating costs. In order to preserve and enhance the knowledge acquired using SEA to predict cabin noise for propeller airplanes, potential noise control applications for more conventional airplane configurations were investigated. The present paper records an effort to extend the experience acquired using statistical energy analysis for unducted fan engines to noise generated by turbofan engine exhaust. The technique is applied to the generic case of a large commercial airplane with twin, wing-mounted engines. Results are presented from an evaluation of the noise source based on an uncommon set of flight test data. Model construction is decribed and prediction results compared to the flight test data. It is then demonstrated how SEA is used to prioritize the transmission paths and judge the merit of the common noise suppression techniques.

  12. Airplane Upset Training Evaluation Report

    NASA Technical Reports Server (NTRS)

    Gawron, Valerie J.; Jones, Patricia M. (Technical Monitor)

    2002-01-01

    Airplane upset accidents are a leading factor in hull losses and fatalities. This study compared five types of airplane-upset training. Each group was composed of eight, non-military pilots flying in their probationary year for airlines operating in the United States. The first group, 'No aero / no upset,' was made up of pilots without any airplane upset training or aerobatic flight experience; the second group, 'Aero/no upset,' of pilots without any airplane-upset training but with aerobatic experience; the third group, 'No aero/upset,' of pilots who had received airplane-upset training in both ground school and in the simulator; the fourth group, 'Aero/upset,' received the same training as Group Three but in addition had aerobatic flight experience; and the fifth group, 'In-flight' received in-flight airplane upset training using an instrumented in-flight simulator. Recovery performance indicated that clearly training works - specifically, all 40 pilots recovered from the windshear upset. However few pilots were trained or understood the use of bank to change the direction of the lift vector to recover from nose high upsets. Further, very few thought of, or used differential thrust to recover from rudder or aileron induced roll upsets. In addition, recovery from icing-induced stalls was inadequate.

  13. Large amplitude flutter of a low aspect ratio panel at low supersonic speeds comparison of theory and experiment

    NASA Technical Reports Server (NTRS)

    Ventres, C. S.; Kang, C. K.

    1973-01-01

    Flutter boundaries, as well as flutter limit cycle amplitudes, frequencies and stresses were computed for a panel of length-width ratio 4.48 exposed to applied in-plane and transverse loads. The Mach number range was 1.1 to 1.4. The method used involved direct numerical integration of modal equations of motion derived from the nonlinear plate equations of von Karman, coupled with linearized potential flow aerodynamic theory. The flutter boundaries agreed reasonably well with experiment, except when the in-plane loading approached the buckling load. Structural damping had to be introduced, to produce frequencies comparable to the experimental values. Attempts to compute panel deflections or stress at a given point met with limited success. There is some evidence, however, that deflection and stress maxima can be estimated with somewhat greater accuracy.

  14. Bluff body wakes with free, fixed, and discontinuous separation at low Reynolds numbers and low aspect ratio

    NASA Astrophysics Data System (ADS)

    Petrusma, M. S.; Gai, S. L.

    1996-01-01

    Some comparative experimental results are presented of bluff body wakes with free, fixed, and discontinuous separation. The particular examples considered are a circular cylinder, a plain blunt trailing edge aerofoil and a segemented blunt trailing edge aerofoil. The evolution of the near wake and vortex shedding modes are compared and discussed.

  15. Modular Coil Design for the Ultra-low Aspect Ratio Quasi-axially Symmetric Stellarator MHH2

    SciTech Connect

    Ku LP, the ARIES-CS Team

    2005-09-27

    A family of two field-period quasi-axisymmetric stellarators generally known as MHH2 with aspect ratios of only {approx}2.5 was found. These configurations have low field ripples and excellent confinement of {alpha} particles. This discovery raises the hope that a compact stellarator reactor may eventually be designed with the property of tokamak transport and stellarator stability. In this paper we demonstrate that smooth modular coils may be designed for this family of configurations that not only yield plasmas with good physics properties but also possess engineering properties desirable for compact power producing reactors. We show designs featuring 16 modular coils with ratios of major radius to minimum coil-plasma separation {approx}5.5, major radius to minimum coil-coil separation {approx}10 and the maximum field in coil bodies to the field on axis {approx}2 for 0.2 m{sup 2} conductors. These coils is expected to allow plasmas operated at 5% {beta} with {alpha} energy loss < 10% for a reactor of major radius <9 m at 5 T.

  16. Formation and sustainment of a very low aspect ratio tokamak using coaxial helicity injection (the Helicity Injected Torus (HIT) experiment)

    SciTech Connect

    Jarboe, T.R.; Nelson, B.A.

    1992-01-01

    In the paper we will detail the progress of the HIT experiment construction, including the following components: preliminary data and interpretation; diagnostic systems; vacuum vessel and pumping system; helicity source and power supplies; toroidal field coil and power supply; data acquisition system; collaboration with general atomics, with a brief summary given on each.

  17. Preparing high- and low-aspect ratio AlB2 flakes from borax or boron oxide

    NASA Astrophysics Data System (ADS)

    Hall, A. C.; Economy, J.

    2000-02-01

    The commercial preparation of aluminum-diboride flakes in aluminum relies on relatively expensive starting materials. A new synthesis has been developed that allows AlB2 to be prepared directly from the reaction of borax (Na2B4O7·10H2O) or boron oxide (B2O3) with aluminum. Aluminum metal at temperatures higher than 900°C has been shown to reduce these boron-containing compounds, producing an Al2O3-containing slag and AlB2. A natural separation occurs, leaving AlB2 in the molten aluminum and Al2O3 as part of a slag that forms at the melt surface. Samples containing up to 10 vol.% AlB2 in an aluminum matrix have been directly prepared using this method.

  18. Airplane design for gusts

    NASA Technical Reports Server (NTRS)

    Houbolt, J. C.

    1977-01-01

    There are two basic approaches used for the structural design of aircraft due to dust encounter. One is a discrete gust approach, the other is based on power spectral techniques. Both of these approaches are explained in this report. Tacit to the above approaches is the assumption that loading on the airplane arises primarily from vertical gusts. A study of atmospheric turbulence was made not only on the vertical component, but on the longitudinal and transverse gust components as well. An analysis was made to establish the loads that develop when explicit consideration is given to both the vertical and head-wind components. The results are reported. Also included in this report are brief comments on gust effects during approach and landing.

  19. Performance Evaluation of Leading Edge Slats on Rigid Wing Sail Catamarans

    NASA Astrophysics Data System (ADS)

    Johnson, Chelsea; O'Neill, Charles

    2015-11-01

    Rigid wing sails have created the fastest catamarans in history, however with the addition of a leading edge slat higher lift and faster speeds may be achieved. Slats are currently used on airplane wings to increase lift, but have not been implemented on a rigid wing sail catamaran. Using 3D modeling and computational fluid dynamics software, this research investigates the effect that slats have on the performance of rigid wing sail catamarans. Aerodynamics and hydrodynamics form the basis of the research. The preliminary results show an increase in the coefficient of lift for sail models with slats over sail models without slats, allowing the catamaran to perform at higher speeds. The ability of the slat to rotate has also been identified as a key factor in increasing the benefit of the slat. This work was supported by NSF site award 1358991.

  20. Flight control system development and flight test experience with the F-111 mission adaptive wing aircraft

    NASA Technical Reports Server (NTRS)

    Larson, R. R.

    1986-01-01

    The wing on the NASA F-111 transonic aircraft technology airplane was modified to provide flexible leading and trailing edge flaps. This wing is known as the mission adaptive wing (MAW) because aerodynamic efficiency can be maintained at all speeds. Unlike a conventional wing, the MAW has no spoilers, external flap hinges, or fairings to break the smooth contour. The leading edge flaps and three-segment trailing edge flaps are controlled by a redundant fly-by-wire control system that features a dual digital primary system architecture providing roll and symmetric commands to the MAW control surfaces. A segregated analog backup system is provided in the event of a primary system failure. This paper discusses the design, development, testing, qualification, and flight test experience of the MAW primary and backup flight control systems.

  1. A Buffet Investigation at High Subsonic Speeds of Wing-Fuselage-Tail Combinations having Sweptback Wings with NACA Four-Digit Thickness Distributions, Fences, and Body Contouring

    NASA Technical Reports Server (NTRS)

    Sutton, Fred B.

    1959-01-01

    An investigation has been made to determine the effect of wing fences, fuselage contouring, varying wing sweepback angle from 40 deg. to 45 deg., mounting the horizontal tail on an outboard boom) and wing thickness distribution upon the buffeting response of typical airplane configurations employing sweptback wings of high aspect ratio. The tests were conducted through an angle-of-attack range at Mach numbers varying from 0.60 to 0.92 at a Reynolds number of 2 million. For the combinations with 40 deg. of sweepback, the addition of multiple wing fences usually decreased the buffeting at moderate and high lift coefficients and reduced the erratic variation of buffet intensities with increasing lift coefficient and Mach number. Fuselage contouring also reduced buffeting but was not as effective as the wing fences. At most Mach numbers, buffeting occurred at higher lift coefficients for the combination with the NACA 64A thickness distributions than for the combination with the NACA four-digit thickness distributions. At high subsonic speeds, heavy buffeting was usually indicated at lift coefficients which were lower than the lift coefficients for static-longitudinal instability. The addition of wing fences improved the pitching-moment characteristics but had little effect on the onset of buffeting. For most test conditions and model configurations, the root-mean- square and the maximum values measured for relative buffeting indicated similar effects and trends; however, the maximum buffeting loads were usually two to three times the root-mean-square intensities.

  2. Ice Formation on Wings

    NASA Technical Reports Server (NTRS)

    Ritz, L

    1939-01-01

    This report makes use of the results obtained in the Gottingen ice tunnel in which the atmospheric conditions are simulated and the process of ice formation photographed. The effect of ice formation is threefold: 1) added weight to the airplane; 2) a change in the lift and drag forces; 3) a change in the stability characteristics.

  3. Langley Full-scale-tunnel Investigation of the Factors Affecting the Static Lateral-stability Characteristics of a Typical Fighter-type Airplane

    NASA Technical Reports Server (NTRS)

    Lange, Roy H

    1947-01-01

    The factors that affect the rate of change of rolling moment with yaw of a typical fighter-type airplane were investigated in the Langley full-scale tunnel on a typical fighter-type airplane.Eight representative flight conditions were investigated in detail. The separate effects of propeller operation, of the wing-fuselage combination, and of the vertical tail to the effective dihedral of the airplane in each condition were determined. The results of the tests showed that for the airplane with the propeller removed, the wing-fuselage combination had positive dihedral effect which increased considerably with increasing angle of attack for all conditions. Flap deflection decreased the dihedral effect of the wing-fuselage combination slightly as compared with that with the flaps retracted. Flap deflection resulted in negative dihedral effect due to the vertical tail. Propeller operation decreased the lateral stability parameter of the airplane for all the conditions investigated with larger decreases being measured for the flaps deflected conditions.

  4. Hypersonic wing test structure design, analysis, and fabrication

    NASA Technical Reports Server (NTRS)

    Plank, P. P.; Penning, F. A.

    1973-01-01

    An investigation to provide the analyses, data, and hardware required to experimentally validate the beaded panel concept and demonstrate its usefulness as a basis for design of a Hypersonic Research Airplane (HRA) wing is reported. Combinations of the beaded panel structure, heat shields, channel caps and corrugated webs for ribs and spars were analyzed for the wing of a specified HRA to operate at Mach 8 with a lifespan of 150 flights. Detailed analyses were conducted in accordance with established design criteria and included aerodynamic heating and load predictions, transient structural thermal calculations, extensive NASTRAN computer modeling, and structural optimization. Optimum beaded panel tests at 922 K (1200 F) were performed to verify panel performance. Close agreement of predicted and actual critical loads permitted use of design procedures and equations for the beaded panel concept without modification.

  5. 14 CFR 36.7 - Acoustical change: Transport category large airplanes and jet airplanes.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... paragraph (b) of this section: (1) Airplanes with high bypass ratio jet engines. For an airplane that has jet engines with a bypass ratio of 2 or more before a change in type design— (i) The airplane, after... do not have high bypass ratio jet engines. For an airplane that does not have jet engines with...

  6. 14 CFR 36.7 - Acoustical change: Transport category large airplanes and jet airplanes.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... paragraph (b) of this section: (1) Airplanes with high bypass ratio jet engines. For an airplane that has jet engines with a bypass ratio of 2 or more before a change in type design— (i) The airplane, after... do not have high bypass ratio jet engines. For an airplane that does not have jet engines with...

  7. 14 CFR 36.7 - Acoustical change: Transport category large airplanes and jet airplanes.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... paragraph (b) of this section: (1) Airplanes with high bypass ratio jet engines. For an airplane that has jet engines with a bypass ratio of 2 or more before a change in type design— (i) The airplane, after... do not have high bypass ratio jet engines. For an airplane that does not have jet engines with...

  8. Theoretical aerodynamic characteristics of a family of slender wing-tail-body combinations

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard; Byrd, Paul F

    1951-01-01

    The aerodynamic characteristics of an airplane configuration composed of a swept-back, nearly constant chord wing and a triangular tail mounted on a cylindrical body are presented. The analysis is based on the assumption that the free-stream Mach number is near unity or that the configuration is slender. The calculations for the tail are made on the assumption that the vortex system trailing back from the wing is either a sheet lying entirely in the plane of the flat tail surface or has completely "rolled up" into two point vortices that lie either in, above, or below the plane of the tail surface.

  9. Analysis of iced wings

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  10. Improved Method for Prediction of Attainable Wing Leading-Edge Thrust

    NASA Technical Reports Server (NTRS)

    Carlson, Harry W.; McElroy, Marcus O.; Lessard, Wendy B.; McCullers, L. Arnold

    1996-01-01

    Prediction of the loss of wing leading-edge thrust and the accompanying increase in drag due to lift, when flow is not completely attached, presents a difficult but commonly encountered problem. A method (called the previous method) for the prediction of attainable leading-edge thrust and the resultant effect on airplane aerodynamic performance has been in use for more than a decade. Recently, the method has been revised to enhance its applicability to current airplane design and evaluation problems. The improved method (called the present method) provides for a greater range of airfoil shapes from very sharp to very blunt leading edges. It is also based on a wider range of Reynolds numbers than was available for the previous method. The present method, when employed in computer codes for aerodynamic analysis, generally results in improved correlation with experimental wing-body axial-force data and provides reasonable estimates of the measured drag.

  11. Static performance and noise tests on a thrust reverser for an augmentor wing aircraft

    NASA Technical Reports Server (NTRS)

    Harkonen, D. L.; Marrs, C. C.; Okeefe, J. V.

    1974-01-01

    A 1/3 scale model static test program was conducted to measure the noise levels and reverse thrust performance characteristics of wing-mounted thrust reverser that could be used on an advanced augmentor wing airplane. The configuration tested represents only the most fundamental designs where installation and packaging restraints are not considered. The thrust reverser performance is presented in terms of horizontal, vertical, and resultant effectiveness ratios and the reverser noise is compared on the basis of peak perceived noise level (PNL) and one-third octave band data (OASPL). From an analysis of the model force and acoustic data, an assessment is made on the stopping distance versus noise for a 90,900 kg (200,000 lb) airplane using this type of thrust reverser.

  12. Preliminary investigation of rolling moments obtained with spoilers on both slotted and plain wings

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Wenzinger, Carl J

    1932-01-01

    A wind-tunnel study has been made to determine the possibility of developing spoilers suitable for providing the lateral control for airplanes in place of the usual ailerons. The first tests were made on a model wing with a fixed tip slot, but when it was found that the effectiveness of the spoilers did not depend to any great extent on the slot, tests were made on a plain wing also. In both cases certain spoiler positions were found which were free from the usual adverse rolling moments with small deflections. Five different forms of spoiler were tested, the best ones being simple plates, either straight or slightly curved to fit the contour of the airfoil when not deflected. Sufficient rolling moment can probably be obtained from spoilers of reasonable size to provide satisfactory lateral control for certain types of airplanes.

  13. The X-15 airplane - Lessons learned

    NASA Technical Reports Server (NTRS)

    Dana, William H.

    1993-01-01

    The X-15 rocket research airplane flew to an altitude of 354,000 ft and reached Mach 6.70. In almost 200 flights, this airplane was used to gather aerodynamic-heating, structural loads, stability and control, and atmospheric-reentry data. This paper describes the origins, design, and operation of the X-15 airplane. In addition, lessons learned from the X-15 airplane that are applicable to designing and testing the National Aero-Space Plane are discussed.

  14. 14 CFR 125.355 - Airplane equipment.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane equipment. 125.355 Section...

  15. 14 CFR 121.605 - Airplane equipment.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane equipment. 121.605 Section 121.605..., FLAG, AND SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.605 Airplane equipment. No person may dispatch or release an airplane unless it is airworthy and is equipped as prescribed in §...

  16. 14 CFR 23.3 - Airplane categories.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Airplane categories. 23.3 Section 23.3 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES General § 23.3 Airplane...

  17. 14 CFR 121.605 - Airplane equipment.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Airplane equipment. 121.605 Section 121.605..., FLAG, AND SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.605 Airplane equipment. No person may dispatch or release an airplane unless it is airworthy and is equipped as prescribed in §...

  18. 76 FR 77934 - Airworthiness Directives; Airbus Airplanes

    Federal Register 2010, 2011, 2012, 2013, 2014

    2011-12-15

    ... series airplanes. Since we issued AD 2005-23-02, Amendment 39-14360 (70 FR 69067, November 14, 2005), The... certain ACT equipped airplanes, produced after AD 2005-23-02, Amendment 39-14360 (70 FR 69067, November 14...-14360 (70 FR 69067, November 14, 2005). Applicability (c) This AD applies to Airbus airplanes listed...

  19. 14 CFR 125.93 - Airplane limitations.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93...

  20. 14 CFR 121.605 - Airplane equipment.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Airplane equipment. 121.605 Section 121.605..., FLAG, AND SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.605 Airplane equipment. No person may dispatch or release an airplane unless it is airworthy and is equipped as prescribed in §...

  1. 14 CFR 121.605 - Airplane equipment.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Airplane equipment. 121.605 Section 121.605..., FLAG, AND SUPPLEMENTAL OPERATIONS Dispatching and Flight Release Rules § 121.605 Airplane equipment. No person may dispatch or release an airplane unless it is airworthy and is equipped as prescribed in §...

  2. 14 CFR 125.93 - Airplane limitations.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93...

  3. 14 CFR 125.93 - Airplane limitations.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93...

  4. Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Sewall, William G.

    1995-01-01

    Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

  5. A piloted evaluation of an oblique-wing research aircraft motion simulation with decoupling control laws

    NASA Technical Reports Server (NTRS)

    Kempel, Robert W.; Mcneill, Walter E.; Gilyard, Glenn B.; Maine, Trindel A.

    1988-01-01

    The NASA Ames Research Center developed an oblique-wing research plane from NASA's digital fly-by-wire airplane. Oblique-wing airplanes show large cross-coupling in control and dynamic behavior which is not present on conventional symmetric airplanes and must be compensated for to obtain acceptable handling qualities. The large vertical motion simulator at NASA Ames-Moffett was used in the piloted evaluation of a proposed flight control system designed to provide decoupled handling qualities. Five discrete flight conditions were evaluated ranging from low altitude subsonic Mach numbers to moderate altitude supersonic Mach numbers. The flight control system was effective in generally decoupling the airplane. However, all participating pilots objected to the high levels of lateral acceleration encountered in pitch maneuvers. In addition, the pilots were more critical of left turns (in the direction of the trailing wingtip when skewed) than they were of right turns due to the tendency to be rolled into the left turns and out of the right turns. Asymmetric side force as a function of angle of attack was the primary cause of lateral acceleration in pitch. Along with the lateral acceleration in pitch, variation of rolling and yawing moments as functions of angle of attack caused the tendency to roll into left turns and out of right turns.

  6. Flutter study of an advanced composite wing with external stores

    NASA Technical Reports Server (NTRS)

    Cole, Stanley R.; Rivera, Jose A., Jr.; Nagaraja, K. S.

    1987-01-01

    A flutter test using a scaled model of an advanced composite wing for a Navy attack aircraft has been conducted in the NASA Langley Research Center Transonic Dynamics Tunnel. The model was a wall-mounted half-span wing with a semi-span of 6.63 ft. The wing had an aspect ratio of 5.31, taper ratio of 0.312, and quarter-chord sweep of 25 degrees. The model was supported in a manner that simulated the load path in the carry-through structure of the aircraft and the symmetric boundary condition at the fuselage centerline. The model was capable of carrying external stores from three pylon locations on the wing. Flutter tests were conducted for the wing with and without external stores. No flutter was encountered for the clean wing at test conditions which simulated the scaled airplane operating envelope. Flutter boundaries were obtained for several external store configurations. The flutter boundaries for the fuel tanks were nearly Mach number independent (occurring at constant dynamic pressure). To study the aerodynamic effect of the fuel tank stores, pencil stores (slender cylindrical rods) which had the same mass and pitch and yaw inertia as the fuel tanks were tested on the model. These pencil store configurations exhibited a transonic dip in the flutter dynamic pressure, indicating that the aerodynamic effect of the actual fuel tanks on flutter was significant. Several flutter analyses methods were used in an attempt to predict the flutter phenomenon exhibited during the wind-tunnel test. The analysis gave satisfactory predictions of flutter for the pencil store configurations, but unsatisfactory correlation for the actual fuel tank configurations.

  7. Design, realization and structural testing of a compliant adaptable wing

    NASA Astrophysics Data System (ADS)

    Molinari, G.; Quack, M.; Arrieta, A. F.; Morari, M.; Ermanni, P.

    2015-10-01

    This paper presents the design, optimization, realization and testing of a novel wing morphing concept, based on distributed compliance structures, and actuated by piezoelectric elements. The adaptive wing features ribs with a selectively compliant inner structure, numerically optimized to achieve aerodynamically efficient shape changes while simultaneously withstanding aeroelastic loads. The static and dynamic aeroelastic behavior of the wing, and the effect of activating the actuators, is assessed by means of coupled 3D aerodynamic and structural simulations. To demonstrate the capabilities of the proposed morphing concept and optimization procedure, the wings of a model airplane are designed and manufactured according to the presented approach. The goal is to replace conventional ailerons, thus to achieve controllability in roll purely by morphing. The mechanical properties of the manufactured components are characterized experimentally, and used to create a refined and correlated finite element model. The overall stiffness, strength, and actuation capabilities are experimentally tested and successfully compared with the numerical prediction. To counteract the nonlinear hysteretic behavior of the piezoelectric actuators, a closed-loop controller is implemented, and its capability of accurately achieving the desired shape adaptation is evaluated experimentally. Using the correlated finite element model, the aeroelastic behavior of the manufactured wing is simulated, showing that the morphing concept can provide sufficient roll authority to allow controllability of the flight. The additional degrees of freedom offered by morphing can be also used to vary the plane lift coefficient, similarly to conventional flaps. The efficiency improvements offered by this technique are evaluated numerically, and compared to the performance of a rigid wing.

  8. Effects of flaps on buffet characteristics and wind-rock onset of an F-8C airplane at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Monaghan, R. C.; Friend, E. L.

    1973-01-01

    Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edge-flap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg.

  9. Effects of airplane characteristics and takeoff noise and field length constraints on engine cycle selection for a Mach 2.32 cruise application

    NASA Technical Reports Server (NTRS)

    Whitlow, J. B., Jr.

    1976-01-01

    Sideline noise and takeoff field length were varied for two types of Mach 2.32 cruise airplane to determine their effect on engine cycle selection. One of these airplanes was the NASA/Langley-LTV arrow wing while the other was a Boeing modified delta-plus-tail derived from the earlier 2707-300 concept. Advanced variable cycle engines were considered. A more conventional advanced low bypass turbofan engine was used as a baseline for comparison. Appropriate exhaust nozzle modifications were assumed, where needed, to allow all engines to receive either an inherent co-annular or annular jet noise suppression benefit. All the VCE's out-performed the baseline engine by substantial margins in a design range comparison, regardless of airplane choice or takeoff restrictions. The choice among the three VCE's considered, however, depends on the field length, noise level, and airplane selected.

  10. Investigation on adaptive wing structure based on shape memory polymer composite hinge

    NASA Astrophysics Data System (ADS)

    Yu, Yuemin; Li, Xinbo; Zhang, Wei; Leng, Jinsong

    2007-07-01

    This paper describes the design and investigation of the SMP composite hinge and the morphing wing structure. The SMP composite hinge was based on SMP and carbon fiber fabric. The twisting recoverability of it was investigated by heating and then cooling repeatedly above and below the Tg. The twisting recoverability characterized by the twisting angle. Results show that the SMP composite hinge have good shape recoverability, Recovery time has a great influence on the twisting recoverability. The twisting recovery ratio became large with the increment of recovery time. The morphing wing can changes shape for different tasks. For the advantages of great recovery force and stable performances, we adopt SMP composite hinge as actuator to apply into the structure of the wing which can realize draw back wings to change sweep angle according to the speed and other requirements of military airplanes. Finally, a series of simulations and experiments are performed to investigate the deformations of morphing wings have been performed successfully. It can be seen that the sweep angle change became large with the increment of initial angle. The area reduction became large with the increment of initial angle, but after 75° the area reduction became smaller and smaller. The deformations of the triangle wing became large with the increment of temperature. The area and the sweep angle of wings can be controlled by adjusting the stimulate temperature and the initial twisting angle of shape memory polymer composite hinge.

  11. F-16XL ship #1 - CAWAP boundary layer hot film, left wing

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This photo shows the boundary layer hot film on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. Hot film is used to measure temperature changes on a surface. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

  12. F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

    NASA Technical Reports Server (NTRS)

    1996-01-01

    This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

  13. Measured and predicted pressure distributions on the AFTI/F-111 mission adaptive wing

    NASA Technical Reports Server (NTRS)

    Webb, Lannie D.; Mccain, William E.; Rose, Lucinda A.

    1988-01-01

    Flight tests have been conducted using an F-111 aircraft modified with a mission adaptive wing (MAW). The MAW has variable-camber leading and trailing edge surfaces that can change the wing camber in flight, while preserving smooth upper surface contours. This paper contains wing surface pressure measurements obtained during flight tests at Dryden Flight Research Facility of NASA Ames Research Center. Upper and lower surface steady pressure distributions were measured along four streamwise rows of static pressure orifices on the right wing for a leading-edge sweep angle of 26 deg. The airplane, wing, instrumentation, and test conditions are discussed. Steady pressure results are presented for selected wing camber deflections flown at subsonic Mach numbers up to 0.90 and an angle-of-attack range of 5 to 12 deg. The Reynolds number was 26 million, based on the mean aerodynamic chord. The MAW flight data are compared to MAW wind tunnel data, transonic aircraft technology (TACT) flight data, and predicted pressure distributions. The results provide a unique database for a smooth, variable-camber, advanced supercritical wing.

  14. Transonic wind tunnel test of a 14 percent thick oblique wing

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

    An experimental investigation was conducted at the ARC 11- by 11-Foot Transonic Wind Tunnel as part of the Oblique Wing Research Aircraft Program to study the aerodynamic performance and stability characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing designed by Rockwell International. The 10.3 aspect ratio, straight-tapered wing of 0.14 thickness/chord ratio was tested at two different mounting heights above the fuselage. Additional tests were conducted to assess low-speed behavior with and without flaps, aileron effectiveness at representative flight conditions, and transonic drag divergence with 0 degree wing sweep. Longitudinal stability data were obtained at sweep angles of 0, 30, 45, 60, and 65 degrees, at Mach numbers ranging from 0.25 to 1.40. Test Reynolds numbers varied from 3.2 to 6.6 x 10 exp 6/ft. and angle of attack ranged from -5 to +18 degrees. Most data were taken at zero sideslip, but a few runs were at sideslip angles of +/- 5 degrees. The raised wing position proved detrimental overall, although side force and yawing moment were reduced at some conditions. Maximum lift coefficient with the flaps deflected was found to fall short of the value predicted in the preliminary design document. The performance and trim characteristics of the present wing are generally inferior to those obtained for a previously tested wing designed at ARC.

  15. Air resistance measurements on actual airplane parts

    NASA Technical Reports Server (NTRS)

    Weiselsberger, C

    1923-01-01

    For the calculation of the parasite resistance of an airplane, a knowledge of the resistance of the individual structural and accessory parts is necessary. The most reliable basis for this is given by tests with actual airplane parts at airspeeds which occur in practice. The data given here relate to the landing gear of a Siemanms-Schuckert DI airplane; the landing gear of a 'Luftfahrzeug-Gesellschaft' airplane (type Roland Dlla); landing gear of a 'Flugzeugbau Friedrichshafen' G airplane; a machine gun, and the exhaust manifold of a 269 HP engine.

  16. Spin-Tunnel Investigation of a 1/30-Scale Model of the North American A-5 Airplane

    NASA Technical Reports Server (NTRS)

    Lee, Henry A.

    1964-01-01

    An investigation has been made to determine the erect and. inverted spin and recovery characteristics of a 1/30-scale dynamic model of the North American A-5A airplane. Tests were made for the basic flight design loading with the center of gravity at 30-percent mean aerodynamic chord and also for a forward position and a rearward position with the center of gravity at 26-percent and 40-percent mean aerodynamic chord, respectively. Tests were also made to determine the effect of full external wing tanks on both wings, and of an asymmetrical condition when only one full tank is carried.

  17. Spin Tunnel Investigation of a 1/30 Scale Model of the North American A-5A Airplane

    NASA Technical Reports Server (NTRS)

    Lee, Henry A.

    1964-01-01

    An investigation has been made in the Langley spin tunnel to determine the erect and inverted spin and recovery characteristics of a 1/30-scale dynamic model of the North American A-5A airplane. Tests were made for the basic flight design loading with the center of gravity at 30-percent mean aerodynamic chord and also for a forward position and a rearward position with the center of gravity at 26-percent and 40-percent mean aerodynamic chord, respectively. Tests were also made to determine the effect of full external wing tanks on both wings, and of an asymmetrical condition when only one full tank is carried.

  18. Summary of Measurements in Langley Full-Scale Tunnel of Maximum Lift Coefficients and Stalling Characteristics of Airplanes

    NASA Technical Reports Server (NTRS)

    Sweberg, Harold H; Dingeldein, Richard C

    1945-01-01

    The results of measurements in the Langley full-scale tunnel of the maximum lift coefficients and stalling characteristics of airplanes have been collected. The data have been analyzed to show the nature of the effects on maximum lift and stall of wing geometry, fuselages and nacelles, propeller slipstream, surface roughness, and wing leading-edge appendages such as ducts, armaments, tip slats, and airspeed heads. Comparisons of full-scale-tunnel and flight measurements of maximum lift and stall are included in some cases, and the effects of the different testing techniques on the maximum-lift measurements are also given.

  19. An application of parameter identification to the oscillatory motion of an airplane at high C(L)

    NASA Technical Reports Server (NTRS)

    Batterson, J. G.; Klein, V.

    1983-01-01

    This paper presents an application of a stepwise regression incorporating polynomial splines to oscillatory flight data from a light research airplane operating at near stall angles of attack. It is shown that data from several experiments can be combined into a large data set for analysis and that hysteresis phenomena can be observed in this large data set. Finally, it is postulated from the analysis of the flight data and theoretical calculations that the observed oscillatory motion is the result of a combination of wing stall and wing wake position at the tail.

  20. Effort of spanwise variation of turbulence on the normal acceleration of airplanes with small span relative to turbulence scale

    NASA Technical Reports Server (NTRS)

    Pratt, K. G.

    1975-01-01

    A rigid airplane with an unswept wing is analyzed. The results show that the power spectrum, relative to that for a one-dimensional turbulence field, is significantly attenuated at the higher frequencies even for airplanes with arbitrarily small ratios of span to scale of turbulence. This attenuation is described by a simple weighting function of frequency that depends only on aspect ratio. The weighting function, together with the attenuation due to the unsteady flow of gust penetration, allows the determination of the average rate of zero crossings for airplanes having very small spans without recourse to an integral truncation which is often required in calculations based on a one-dimensional turbulence field.

  1. The forces and moments acting on parts of the XN2Y-1 airplane during spins

    NASA Technical Reports Server (NTRS)

    Scudder, N F

    1937-01-01

    The magnitudes of the yawing moments produced by various parts of an airplane during spins have previously been found to be of major importance in determining the nature of the spin. Discrepancies in resultant yawing moments determined from model and full-scale tests, however, have indicated the probable importance of scale effect on the model. In order to obtain data for a more detailed comparison between full-scale and model results, flight tests were made to determine the yawing moments contributed by various parts of an airplane in spins. The inertia moment was determined by the usual measurement of the spinning motion, and the aerodynamic yawing moments on the fuselage, fin, and rudder were determined by pressure-distribution measurements over these parts of the airplane. The wing yawing moment was determined by taking the difference between the gyroscopic moment and the fuselage, fin, and rudder moments. The numerical values of the wing yawing moments were found to be of the same order of magnitude as those measured in wind tunnels.

  2. Theory of wing rock

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  3. 14 CFR 125.75 - Airplane flight manual.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane flight manual. 125.75 Section 125... Airplane flight manual. (a) Each certificate holder shall keep a current approved Airplane Flight Manual or... approved Airplane Flight Manual or the approved equivalent aboard each airplane it operates. A...

  4. Longitudinal Trim and Tumble Characteristics of a 0.057-Scale Model of the Chance Vought XF7U-1 Airplane, TED NO. NACA DE311

    NASA Technical Reports Server (NTRS)

    Bryant, Robert L.

    1948-01-01

    Based on results of longitudinal trim and tumble tests of a 0.057-scale model of the Chance Vought XF7U-1 airplane, the following conclusions regarding the trim and tumble characteristics of the airplane have been drawn: 1. The airplane will not trim at any unusual or uncontrolled angles of attack. 2. The airplane will not tumble with the center of gravity located forward of 24 percent of the mean aerodynamic chord. When the center of gravity is located at 24 percent of the mean aerodynamic chord and slats are extended and elevators are deflected full up, the airplane may tumble if given an external positive pitching moment. 3. The tumbling motion obtained will be readily terminated by deflecting the elevators full down so as to oppose the rotation. 4. The accelerations encountered during an established tumble may be dangerous to the pilot and, therefore, action should be taken to terminate a tumble immediately upon its inception. 5. Simultaneous opening of two wing-tip parachutes having diameters of 4 feet or larger and having drag coefficients of approximately 0.7 will effectively terminate the tumble. 6. Model results indicate that the pilot will not be struck by the airplane if it becomes necessary to leave the airplane during a tumble. The pilot may require aid from an ejection-seat arrangement.

  5. The Prevention of the Ice Hazard on Airplanes

    NASA Technical Reports Server (NTRS)

    Geer, William C; Scott, Merit

    1930-01-01

    A review of various methods to prevent ice formation and adhesion to aircraft surfaces is given. It was concluded that the adhesion of ice to a surface may be reduced somewhat by the application of certain waxes and varnishes. In the experiments described, the varnishes containing calcium stearate and calcium oleate gave the best results. In wind tunnel tests, the adhesion was further reduced by the application of these waxes and varnishes to a thin, heat insulating layer of rubber. The adhesion of ice is greatly reduced when the surface consists of a vehicle which carries an oil in sufficient quantity so that the surface of the vehicle is self lubricating. Ice may be removed from wings, struts, wires and other parts of an airplane during flight by the inflation of properly constructed pneumatic rubber members, providing that these members have been previously treated with a suitable low adhesion oil.

  6. Structural response of transport airplanes in crash situations

    NASA Technical Reports Server (NTRS)

    Thomson, R. G.; Caiafa, C.

    1983-01-01

    This report highlights the results of contractural studies of transport accident data undertaken in a joint research program sponsored by the FAA and NASA. From these accident data studies it was concluded that the greatest potential for improved transport crashworthiness is in the reduction of fire related fatalities. Accident data pertaining to fuselage integrity, main landing gear collapse, fuel tank rupture, wing breaks, tearing of tank lower surfaces, and engine pod scrubbing are discussed. In those accidents where the energy absorbing protective capability of the fuselage structure is expended and the airplane experiences major structural damage, trauma caused fatalities are also discussed. The dynamic performance of current seat/restraint systems are examined but it is concluded that the accident data does not adequately define the relationship between occupant response and the dynamic interaction with the seat, floor and fuselage structure.

  7. Supersonic civil airplane study and design: Performance and sonic boom

    NASA Technical Reports Server (NTRS)

    Cheung, Samson

    1995-01-01

    Since aircraft configuration plays an important role in aerodynamic performance and sonic boom shape, the configuration of the next generation supersonic civil transport has to be tailored to meet high aerodynamic performance and low sonic boom requirements. Computational fluid dynamics (CFD) can be used to design airplanes to meet these dual objectives. The work and results in this report are used to support NASA's High Speed Research Program (HSRP). CFD tools and techniques have been developed for general usages of sonic boom propagation study and aerodynamic design. Parallel to the research effort on sonic boom extrapolation, CFD flow solvers have been coupled with a numeric optimization tool to form a design package for aircraft configuration. This CFD optimization package has been applied to configuration design on a low-boom concept and an oblique all-wing concept. A nonlinear unconstrained optimizer for Parallel Virtual Machine has been developed for aerodynamic design and study.

  8. X-Wing RSRA - 80 Knot Taxi Test

    NASA Technical Reports Server (NTRS)

    1987-01-01

    The Rotor Systems Research Aircraft/X-Wing, a vehicle that was used to demonstrate an advanced rotor/fixed wing concept called X-Wing, is shown here during high-speed taxi tests at NASA's Ames-Dryden Flight Research Facility (later redesignated Dryden Flight Research Center), Edwards, California, on 4 November 1987. During these tests, the vehicle made three taxi tests at speeds of up to 138 knots. On the third run, the RSRA/X-Wing lifted off the runway to a 25-foot height for about 16 seconds. This liftoff maneuver was pre-planned as an aid to evaluations for first flight. At the controls were NASA pilot G. Warren Hall and Sikorsky pilot W. Faull. The unusual aircraft that resulted from the Ames Research Center/Army X-Wing Project was flown at the Ames-Dryden Flight Research Facility (now Dryden Flight Research Center), Edwards, California, beginning in the spring of 1984, with a follow-on program beginning in 1986. The program, was conceived to provide an efficient combination of the vertical lift characteristic of conventional helicopters and the high cruise speed of fixed-wing aircraft. It consisted of a hybrid vehicle called the NASA/Army Rotor Systems Research Aircraft (RSRA), which was equipped with advanced X-wing rotor systems. The program began in the early 1970s to investigate ways to increase the speed of rotor aircraft, as well as their performance, reliability, and safety . It also sought to reduce the noise, vibration, and maintenance costs of helicopters. Sikorsky Aircraft Division of United Technologies Laboratories built two RSRA aircraft. NASA's Langley Research Center, Hampton, Virginia, did some initial testing and transferred the program to Ames Research Center, Mountain View, California, for an extensive flight research program conducted by Ames and the Army. The purpose of the 1984 tests was to demonstrate the fixed-wing capability of the helicopter/airplane hybrid research vehicle and explore its flight envelope and flying qualities. These

  9. 3D visualization of unsteady 2D airplane wake vortices

    NASA Technical Reports Server (NTRS)

    Ma, Kwan-Liu; Zheng, Z. C.

    1994-01-01

    Air flowing around the wing tips of an airplane forms horizontal tornado-like vortices that can be dangerous to following aircraft. The dynamics of such vortices, including ground and atmospheric effects, can be predicted by numerical simulation, allowing the safety and capacity of airports to be improved. In this paper, we introduce three-dimensional techniques for visualizing time-dependent, two-dimensional wake vortex computations, and the hazard strength of such vortices near the ground. We describe a vortex core tracing algorithm and a local tiling method to visualize the vortex evolution. The tiling method converts time-dependent, two-dimensional vortex cores into three-dimensional vortex tubes. Finally, a novel approach calculates the induced rolling moment on the following airplane at each grid point within a region near the vortex tubes and thus allows three-dimensional visualization of the hazard strength of the vortices. We also suggest ways of combining multiple visualization methods to present more information simultaneously.

  10. Predicting Tail Buffet Loads of a Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Moses, Robert W.; Pototzky, Anthony S.

    2006-01-01

    Buffet loads on aft aerodynamic surfaces pose a recurring problem on most twin-tailed fighter airplanes: During maneuvers at high angles of attack, vortices emanating from various surfaces on the forward parts of such an airplane (engine inlets, wings, or other fuselage appendages) often burst, immersing the tails in their wakes. Although these vortices increase lift, the frequency contents of the burst vortices become so low as to cause the aft surfaces to vibrate destructively. Now, there exists a new analysis capability for predicting buffet loads during the earliest design phase of a fighter-aircraft program. In effect, buffet pressures are applied to mathematical models in the framework of a finite-element code, complete with aeroelastic properties and working knowledge of the spatiality of the buffet pressures for all flight conditions. The results of analysis performed by use of this capability illustrate those vibratory modes of a tail fin that are most likely to be affected by buffet loads. Hence, the results help in identifying the flight conditions during which to expect problems. Using this capability, an aircraft designer can make adjustments to the airframe and possibly the aerodynamics, leading to a more robust design.

  11. The Testing of Airplane Fabrics

    NASA Technical Reports Server (NTRS)

    Schraivogel, Karl

    1932-01-01

    This report considers the determining factors in the choice of airplane fabrics, describes the customary methods of testing and reports some of the experimental results. To sum up briefly the results obtained with the different fabrics, it may be said that increasing the strength of covering fabrics by using coarser yarns ordinarily offers no difficulty, because the weight increment from doping is relatively smaller.

  12. Glues Used in Airplane Parts

    NASA Technical Reports Server (NTRS)

    Allen, S W; Truax, T R

    1920-01-01

    This report was prepared for the National Advisory Committee for Aeronautics and presents the results of investigations conducted by the Forest Products Laboratory of the United States Forest Service on the manufacture, preparation, application, testing and physical properties of the different types of glues used in wood airplane parts.

  13. Vibration Response of Airplane Structures

    NASA Technical Reports Server (NTRS)

    Theodorsen, Theodore; Gelalles, A G

    1935-01-01

    This report presents test results of experiments on the vibration-response characteristics of airplane structures on the ground and in flight. It also gives details regarding the construction and operation of vibration instruments developed by the National Advisory Committee for Aeronautics.

  14. Paper Airplanes: A Classroom Activity

    ERIC Educational Resources Information Center

    Painter, Richard A.

    1976-01-01

    A learning experience is described for upper elementary or junior high students involving the manufacture, transportation, and marketing of a product for consumers. Steps are given and roles are assigned for students to convert raw material (paper) to a finished product (paper airplanes) and to sell it. (AV)

  15. Qualitative comparison of calculated turbulence responses with wind-tunnel measurements for a DC-10 derivative wing with an active control system

    NASA Technical Reports Server (NTRS)

    Perry, B., III

    1981-01-01

    Comparisons are presented analytically predicted and experimental turbulence responses of a wind tunnel model of a DC-10 derivative wing equipped with an active control system. The active control system was designed for the purpose of flutter suppression, but it had additional benefit of alleviating gust loads (wing bending moment) by about 25%. Comparisions of various wing responses are presented for variations in active control system parameters and tunnel speed. The analytical turbulence responses were obtained using DYLOFLEX, a computer program for dynamic loads analyses of flexible airplanes with active controls. In general, the analytical predictions agreed reasonably well with the experimental data.

  16. Butterfly wing colours: scale beads make white pierid wings brighter.

    PubMed Central

    Stavenga, D. G.; Stowe, S.; Siebke, K.; Zeil, J.; Arikawa, K.

    2004-01-01

    The wing-scale morphologies of the pierid butterflies Pieris rapae (small white) and Delias nigrina (common jezabel), and the heliconine Heliconius melpomene are compared and related to the wing-reflectance spectra. Light scattering at the wing scales determines the wing reflectance, but when the scales contain an absorbing pigment, reflectance is suppressed in the absorption wavelength range of the pigment. The reflectance of the white wing areas of P. rapae, where the scales are studded with beads, is considerably higher than that of the white wing areas of H. melpomene, which has scales lacking beads. The beads presumably cause the distinct matt-white colour of the wings of pierids and function to increase the reflectance amplitude. This will improve the visual discrimination between conspecific males and females. PMID:15306303

  17. Aircraft Wing Preparation

    NASA Technical Reports Server (NTRS)

    1921-01-01

    A LMAL carpenter prepares full scale wings for flight research, 1920. Photograph published in Winds of Change, 75th Anniversary NASA publication (page 36), by James Schultz. Published in Engineer in Charge, NASA SP- 4305 (p. 82), by James R. Hansen.

  18. Pressure Distribution Over the Fuselage of a PW-9 Pursuit Airplane in Flight

    NASA Technical Reports Server (NTRS)

    Rhode, Richard V; Lundquist, Eugene E

    1932-01-01

    This report presents the results obtained from pressure distribution tests on the fuselage of a PW-9 pursuit airplane in a number of conditions of flight. The investigation was made to determine the contribution of the fuselage to the total lift in conditions considered critical for the wing structure, and also to determine whether the fuselage loads acting simultaneously with the maximum tail loads were of such a character as to be of concern with respect to the structural design of other parts of the airplane. The results show that the contribution of the fuselage toward the total lift is small on this airplane. Aerodynamic loads on the fuselage are, in general, unimportant from the structural viewpoint, and in most cases they are of such character that, if neglected, a conservative design results. In spins, aerodynamic forces on the fuselage produce diving moments of appreciable magnitude and yawing moments of small magnitude, but opposing the rotation of the airplane. A table of cowling pressures for various maneuvers is included in the report.

  19. Calculation of tapered monoplane wings

    NASA Technical Reports Server (NTRS)

    Amstutz, E

    1930-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Hultberg, R. S.; Chu, J.

    1980-01-01

    Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin g tunnel are presented in plotted form for a 1/6 scale, single engine, high wing, general aviation model. The configurations tested included the basic airplane and control deflections, wing leading edge devices, tail designs, and airplane components. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg and clockwise and counter clockwise rotations covering a spin coefficient range from 0 to 0.9.