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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.

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At least 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

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

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.

Jenkins, Jerald M.↗
NASA NTRS1952

Effects of wing lift and weight on landing-gear loads

The effects of wing lift and weight on maximum landing gear loads were investigated by drop testing a small landing gear with oleo shock strut in the Langley impact basin. Lift forces were mechanically applied to the dropping mass to produce wing lift factors between 1000 and 2500 pounds. The results show the relationship between lift force, weight, and landing-gear loads for a range of vertical velocities between 0 and 12 feet per second.

Lindquist, Dean C↗
NASA NTRS2005

Deflection-Based Structural Loads Estimation From the Active Aeroelastic Wing F/A-18 Aircraft

Traditional techniques in structural load measurement entail the correlation of a known load with strain-gage output from the individual components of a structure or machine. The use of strain gages has proved successful and is considered the standard approach for load measurement. However, remotely measuring aerodynamic loads using deflection measurement systems to determine aeroelastic deformation as a substitute to strain gages may yield lower testing costs while improving aircraft performance through reduced instrumentation weight. This technique was examined using a reliable strain and structural deformation measurement system. The objective of this study was to explore the utility of a deflection-based load estimation, using the active aeroelastic wing F/A-18 aircraft. Calibration data from ground tests performed on the aircraft were used to derive left wing-root and wing-fold bending-moment and torque load equations based on strain gages, however, for this study, point deflections were used to derive deflection-based load equations. Comparisons between the strain-gage and deflection-based methods are presented. Flight data from the phase-1 active aeroelastic wing flight program were used to validate the deflection-based load estimation method. Flight validation revealed a strong bending-moment correlation and slightly weaker torque correlation. Development of current techniques, and future studies are discussed.

Lizotte, Andrew M.↗
NASA NTRS1925

Calculation of wing spars of variable cross-section and linear load

The calculation of wing spars of constant cross-section and load has been thoroughly treated by a large number of authors. Such is not the case,however, regarding the calculation of wing spars whose section and linear load diminish toward the ends, as in wings of trapezoidal contour and decreasing section.

SPARS↗
NASA NTRS1948

Theoretical symmetric span loading at subsonic speeds for wings having arbitrary plan form

A method is shown by which the symmetric span loading for a certain class of wings can be simply found. The geometry of these wings is limited only to the extent that they must have symmetry about the root chord, must have a straight quarter-chord line over the semispan, and must have no discontinuities in twist. A procedure is shown for finding the lift-curve slope, pitching moment, center of lift, and induced drag from the span load distribution. A method of accounting for the effects of Mach number and for changes in section lift-curve slope is also given. Charts are presented which give directly the characteristics of many wings. Other charts are presented which reduce the problem of finding the symmetric loading on all wings falling within the prescribed limits to the solution of not more than four simultaneous equations. The loadings and wing characteristics predicted by the theory are compared to those given by other theories and by experiment. It is concluded that the results given by the subject theory are satisfactory. The theory is applied to a number of wings to exhibit the effects of such variables as sweep, aspect ratio, taper, and twist. The results are compared and conclusions drawn as to the relative effects of these variables.

Deyoung, John↗
NASA NTRS2016

Full-Field Reconstruction of Structural Deformations and Loads from Measured Strain Data on a Wing Using the Inverse Finite Element Method

A study was undertaken to investigate the measurement of wing deformation and internal loads using measured strain data. Future aerospace vehicle research depends on the ability to accurately measure the deformation and internal loads during ground testing and in flight. The approach uses the inverse Finite Element Method (iFEM). The iFEM is a robust, computationally efficient method that is well suited for real-time measurement of real-time structural deformation and loads. The method has been validated in previous work, but has yet to be applied to a large-scale test article. This work is in preparation for an upcoming loads test of a half-span test wing in the Flight Loads Laboratory at the National Aeronautics and Space Administration Armstrong Flight Research Center (Edwards, California). The method has been implemented into an efficient MATLAB® (The MathWorks, Inc., Natick, Massachusetts) code for testing different sensor configurations. This report discusses formulation and implementation along with the preliminary results from a representative aerospace structure. The end goal is to investigate the modeling and sensor placement approach so that the best practices can be applied to future aerospace projects.

Finite element↗
NASA NTRS1991

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 3: Application of theory for blade loading, wakes, noise, and wing shielding

Results of the program for the generation of a computer prediction code for noise of advanced single rotation, turboprops (prop-fans) such as the SR3 model are presented. The code is based on a linearized theory developed at Hamilton Standard in which aerodynamics and acoustics are treated as a unified process. Both steady and unsteady blade loading are treated. Capabilities include prediction of steady airload distributions and associated aerodynamic performance, unsteady blade pressure response to gust interaction or blade vibration, noise fields associated with thickness and steady and unsteady loading, and wake velocity fields associated with steady loading. The code was developed on the Hamilton Standard IBM computer and has now been installed on the Cray XMP at NASA-Lewis. The work had its genesis in the frequency domain acoustic theory developed at Hamilton Standard in the late 1970s. It was found that the method used for near field noise predictions could be adapted as a lifting surface theory for aerodynamic work via the pressure potential technique that was used for both wings and ducted turbomachinery. In the first realization of the theory for propellers, the blade loading was represented in a quasi-vortex lattice form. This was upgraded to true lifting surface loading. Originally, it was believed that a purely linear approach for both aerodynamics and noise would be adequate. However, two sources of nonlinearity in the steady aerodynamics became apparent and were found to be a significant factor at takeoff conditions. The first is related to the fact that the steady axial induced velocity may be of the same order of magnitude as the flight speed and the second is the formation of leading edge vortices which increases lift and redistribute loading. Discovery and properties of prop-fan leading edge vortices were reported in two papers. The Unified AeroAcoustic Program (UAAP) capabilites are demonstrated and the theory verified by comparison with the predictions with data from tests at NASA-Lewis. Steady aerodyanmic performance, unsteady blade loading, wakes, noise, and wing and boundary layer shielding are examined.

Hanson, D. B.↗
NASA NTRS1979

Fluctuating loads measured on an over-the-wing supersonic jet model

Fluctuating pressure loads on the wing and flap of an over-the-wing supersonic jet model were measured. The model was tested statically and at a Mach number of 0.1 in a small free jet to simulate forward speed. Test parameters were impingement angle, nozzle height, and flap deflection. Load levels as high as 170 db were measured at the center of the impingement region during static tests. Forward speed reduced the loading about 1 db. Load level increased with increasing impingement angle and decreasing nozzle height above the wing. The effect of flap deflection was small. When scaled to full-size aircraft conditions, the maximum amplitude of the one-third-octave fluctuating pressure spectra was about 154 db at about 160 Hz. Maximum load level occurred near the intersection of the nozzle center line with the impinged surface. Downstream of the maximum the fluctuating pressure is inversely proportional to the distance downstream of the nozzle.

Willis, C. M.↗
NASA NTRS1984

Analytical model of the structureborne interior noise induced by a propeller wake

The structure-borne contribution to the interior noise that is induced by the propeller wake acting on the wing was studied. Analytical models were developed to describe each aspect of this path including the excitation loads, the wing and fuselage structures, and the interior acoustic space. The emphasis is on examining a variety of parameters, and as a result different models were developed to examine specific parameters. The excitation loading on the wing by the propeller wake is modeled by a distribution of rotating potential vortices whose strength is related to the thrust per blade. The response of the wing to this loading is examined using beam models. A model of a beam structurally connected to a cylindrical shell with an internal acoustic fluid was developed to examine the coupling of energy from the wing to the interior space. The model of the acoustic space allows for arbitrary end conditions (e.g., rigid or vibrating end caps). Calculations are presented using these models to compare with a laboratory test configuration as well as for parameters of a prop-fan aircraft.

Junger, M. C.↗
NASA NTRS1999

An H-Infinity Approach to Control Synthesis with Load Minimization for the F/A-18 Active Aeroelastic Wing

The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H-infinity synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.

Lind, Rick↗
NASA NTRS1999

An H-infinity Approach to Control Synthesis with Load Minimization for the F/A-18 Active Aeroelastic Wing

The F/A-18 Active Aeroelastic Wing research aircraft will demonstrate technologies related to aeroservoelastic effects such as wing twist and load minimization. This program presents several challenges for control design that are often not considered for traditional aircraft. This paper presents a control design based on H(sub infinity) synthesis that simultaneously considers the multiple objectives associated with handling qualities, actuator limitations, and loads. A point design is presented to demonstrate a controller and the resulting closed-loop properties.

Lind, Rick↗
NASA NTRS2016

Evaluation of the Hinge Moment and Normal Force Aerodynamic Loads from a Seamless Adaptive Compliant Trailing Edge Flap in Flight

A seamless adaptive compliant trailing edge (ACTE) flap was demonstrated in flight on a Gulfstream III aircraft at the NASA Armstrong Flight Research Center. The trailing edge flap was deflected between minus 2 deg up and plus 30 deg down in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The interface hardware instrumentation layout and load calibration are discussed. Twenty-one applied calibration test load cases were developed for each individual fitting. The 2-sigma residual errors for the hinge moment was calculated to be 2.4 percent, and for normal force was calculated to be 7.3 percent. The hinge moment and normal force generated by the ACTE flap with a hinge point located at 26-percent wing chord were measured during steady state and symmetric pitch maneuvers. The loads predicted from analysis were compared to the loads observed in flight. The hinge moment loads showed good agreement with the flight loads while the normal force loads calculated from analysis were over-predicted by approximately 20 percent. Normal force and hinge moment loads calculated from the pressure sensors located on the ACTE showed good agreement with the loads calculated from the installed strain gages.

Hinge Moment Loads↗
NASA NTRS1984

Application of panel method to wake vortex/wing interaction and comparison with experiment

The ability of a low-order panel method to calculate the aerodynamic loads on wings caused by interaction with wake vortices was studied. The loads were calculated for various positions of a downstream following wing relative to an upstream vortex-generating wing. Calculated vortex-induced span loads and rolling-moment coefficients on the following wing were compared with experimental data. A good agreement with experiment was obtained when the following wing was located more than one following-wing chord length from the tip vortex. The predictions deteriorated as the following wing was placed closer to the vortex. At large downstream distances (approximately 10 generating-wing chord lengths), induced rolling-moment coefficients on the following wing were consistently overestimated. Despite the strong interaction between the wake-vortex filaments and surface doublet panels, the accuracy of the calculations was in most cases independent of the panel distribution and density. A good agreement between theoretical and experimental loads was obtained with a minimum of experimentation with panel arrangements.

Smith, B. E.↗