Search NASA⌕ Search

SEARCH · Search NASA

Results for “Wing loading”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

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

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.

Allen, Michael J.↗
NASA NTRS2005

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

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.

Allen, Michael J.↗
NASA NTRS1997

Application of Interface Technology in Nonlinear Analysis of a Stitched/RFI Composite Wing Stub Box

A recently developed interface technology was successfully employed in the geometrically nonlinear analysis of a full-scale stitched/RFI composite wing box loaded in bending. The technology allows mismatched finite element models to be joined in a variationally consistent manner and reduces the modeling complexity by eliminating transition meshing. In the analysis, local finite element models of nonlinearly deformed wide bays of the wing box are refined without the need for transition meshing to the surrounding coarse mesh. The COMET-AR finite element code, which has the interface technology capability, was used to perform the analyses. The COMET-AR analysis is compared to both a NASTRAN analysis and to experimental data. The interface technology solution is shown to be in good agreement with both. The viability of interface technology for coupled global/local analysis of large scale aircraft structures is demonstrated.

Wang, John T.↗
NASA NTRS1996

Design and evaluation of a bolted joint for a discrete carbon-epoxy rod-reinforced hat section

The use of prefabricated pultruded carbon-epoxy rods has reduced the manufacturing complexity and costs of stiffened composite panels while increasing the damage tolerance of the panels. However, repairability of these highly efficient discrete stiffeners has been a concern. Design, analysis, and test results are presented in this paper for a bolted-joint repair for the pultruded rod concept that is capable of efficiently transferring axial loads in a hat-section stiffener on the upper skin segment of a heavily loaded aircraft wing component. A tension and a compression joint design were evaluated. The tension joint design achieved approximately 1.0% strain in the carbon-epoxy rod-reinforced hat-section and failed in a metal fitting at 166% of the design ultimate load. The compression joint design failed in the carbon-epoxy rod-reinforced hat-section test specimen area at approximately 0.7% strain and at 110% of the design ultimate load. This strain level of 0.7% in compression is similar to the failure strain observed in previously reported carbon-epoxy rod-reinforced hat-section column tests.

Rousseau, Carl Q.↗
NASA NTRS1974

A modified Newton-Raphson analysis of flight measurements of the trailing vortices of a heavy jet transport

A modified Newton-Raphson method has been used to determine the parameters in equations describing the vortex flow to obtain a best match with flight measurements of the flow behind the C-5A airplane. The flight measurements were made using a specially instrumented T-33 airplane which passed as closely as possible to the centers of the trailing vortices at several distances behind the C-5A airplane. The flight measurements were transformed to flow velocity relative to an inertial frame of reference. The assumed form of the flow consisted of the superposition of two counterrotating, finite core vortices. The positions of the vortex centers, their total circulation, the effective eddy viscosity and measurement bias were the parameters adjusted. The assumed form of vortex flow fit well the measured velocities for the numerous sets of data, both flaps up and down for the C-5A airplane. The resulting values of total circulation, however, were about two-thirds that expected of a wing with an elliptical loading. A partial explanation of the less than expected circulation is a dip in the spanwise lift distribution at the airplane's center line. The distance between the trailing vortices at the smallest times encountered is somewhat less than that expected for an elliptical wing loading.

Taylor, L. W., Jr.↗
NASA NTRS1996

Design and Evaluation of a Bolted Joint for a Discrete Carbon-Epoxy Rod-Reinforced Hat Section

The use of pre-fabricated pultruded carbon-epoxy rods has reduced the manufacturing complexity and costs of stiffened composite panels while increasing the damage tolerance of the panels. However, repairability of these highly efficient discrete stiffeners has been a concern. Design, analysis, and test results are presented in this paper for a bolted-joint repair for the pultruded rod concept that is capable of efficiently transferring axial loads in a hat-section stiffener on the upper skin segment of a heavily loaded aircraft wing component. A tension and a compression joint design were evaluated. The tension joint design achieved approximately 1.0 percent strain in the carbon-epoxy rod-reinforced hat-section and failed in a metal fitting at 166 percent of the design ultimate load. The compression joint design failed in the carbon-epoxy rod-reinforced hat-section test specimen area at approximately 0.7 percent strain and at 110 percent of the design ultimate load. This strain level of 0.7 percent in compression is similar to the failure strain observed in previously reported carbon-epoxy rod-reinforced hat-section column tests.

Baker, Donald J.↗
NASA NTRS1977

Load distribution on a closed-coupled wing canard at transonic speeds

A wind tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely coupled wing canard configuration is reported. Detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading edge vortex flow conditions encountered are included. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading edge vortex to burst over the wing when the wing was in the presence of the high canard.

Gloss, B. B.↗
NASA NTRS1980

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

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

Tang, M. H.↗
NASA NTRS1955

Theoretical study of the tunnel-boundary lift interference due to slotted walls in the presence of the trailing-vortex system of a lifting model

The equations presented in this report give the interference on the trailing-vortex system of a uniformly loaded finite-span wing in a circular tunnel containing partly open and partly closed walls, with special reference to symmetrical arrangements of the open and closed portions. Methods are given for extending the equations to include tunnel shapes other than circular. The rectangular tunnel is used to demonstrate these methods. The equations are also extended to nonuniformly loaded wings.

Matthews, Clarence W↗
NASA NTRS1937

Full Scale Span Load Distribution on a Tapered Wing with Split Flaps of Various Spans

Pressure-distribution tests were conducted in the full-scale wind tunnel on a 2:1 tapered U.S.A. 45 airfoil equipped with 20 percent chord split trailing-edge flaps of various spans. A special installation was employed in the tests utilizing a half-span airfoil mounted vertically above a reflection plane. The airfoil has a constant chord center section and rounded tips and is tapered in thickness from 18 percent c at the root to 9 percent c at the tip. The aerodynamic characteristics, given by the usual dimension less coefficients, are presented graphically as functions of flap span and angle of attack as well as by semispan load diagrams. The results indicate, in general, that only a relatively small increase in the normal-force coefficient is to be expected by extending the flap span of an airfoil-flap combination, similar to the one tested, beyond 70 percent of the wing span.

Parsons, John F↗
NASA NTRS1977

Recent Loads Calibration Experience With a Delta Wing Airplane

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

Jenkins, Jerald M.↗
NASA NTRS2009

Dynamic Structural Fault Detection and Identification

Aircraft structures are designed to guarantee safety of flight in some required operational envelope. When the aircraft becomes structurally impaired, safety of flight may not be guaranteed within that previously safe operational envelope. In this case the safe operational envelope must be redefined in-flight and a means to prevent excursion from this new envelope must be implemented. A specific structural failure mode that may result in a reduced safe operating envelope, the exceedance of which could lead to catastrophic structural failure of the aircraft, will be addressed. The goal of the DFEAP program is the detection of this failure mode coupled with flight controls adaptation to limit critical loads in the damaged aircraft structure. The DFEAP program is working with an F/A-18 aircraft model. The composite wing skins are bonded to metallic spars in the wing substructure. Over time, it is possible that this bonding can deteriorate due to fatigue. In this case, the ability of the wing spar to transfer loading between the wing skins is reduced. This failure mode can translate to a reduced allowable compressive strain on the wing skin and could lead to catastrophic wing buckling if load limiting of the wing structure is not applied. The DFEAP program will make use of a simplified wing strain model for the healthy aircraft. The outputs of this model will be compared in real-time to onboard strain measurements at several locations on the aircraft wing. A damage condition is declared at a given location when the strain measurements differ sufficiently from the strain model. Parameter identification of the damaged structure wing strain parameters will be employed to provide load limiting control adaptation for the aircraft. This paper will discuss the simplified strain models used in the implementation and their interaction with the strain sensor measurements. Also discussed will be the damage detection and identification schemes employed and the means by which the damaged aircraft parameters will be used to provide load limiting that keeps the aircraft within the safe operational envelope.

Smith, Timothy↗
NASA NTRS1951

Low-speed static longitudinal and lateral stability characteristics of two low-aspect-ratio wings cambered and twisted to provide a uniform load at a supersonic flight condition

A delta wing and a tapered sweptback wing of aspect ratios 1.56 and 2.00, respectively, both of which were cambered and twisted so as to provide a uniform load distribution for a supersonic flight condition, were tested in combination with a fuselage at Reynolds numbers between 384,000 and 1,550,000 in order to determine the low-speed lift-drag and static stability characteristics of such wings.

REYNOLDS NUMBER EFFECTS - COMPLETE WINGS↗
NASA NTRS1951

Lift, Drag, and Pitching Moment of Low-Aspect-Ratio Wings at Subsonic and Supersonic Speeds: Triangular Wing of Aspect Ratio 2 with NACA 0005-63 Thickness Distribution, Cambered and Twisted for a Trapezoidal Span Load Distribution

A wing-body combination having a plane triangular wing of aspect ratio 2 with NACA 0005-63 thickness distribution in streamwise planes, and twisted and cambered for a trapezoidal span load distribution has been investigated at both subsonic and supersonic Mach numbers. The lift, drag, and pitching moment of the model are presented for Mach numbers from 0.60 to 0.90 and 1.30 to 1.70 at a Reynolds number of 3.0 million. The variations of the characteristics with Reynolds number are also shown for several Mach numbers.

Smith, Willard G.↗
NASA NTRS1972

Initial rotation-loading and low speed flutter test results for a straight wing version of the space shuttle vehicle

For three straight semispan model space shuttle wings, the maximum total load during rapid rotation from 66 deg to 0 deg angle of attack, at Mach numbers from 0.28 to 0.60, was essentially no higher than that measured for buffet. During slow rotation over the same angle range, there was no visible flutter. For one of the wings, however, unstable aerodynamic damping was established at two fixed angles of attack.

Warner, R. W.↗