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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

A hybrid formulation of a component mode synthesis method

Component mode synthesis is a substructuring technique frequently employed in structural dynamics. In this method, a given structure is subdivided into components or substructures, each of which is analyzed independently for natural frequencies and for mode shapes. The substructure mode shapes are then assembled to give displacement shapes or load patterns of the original structure. An analytical justification of the basic concept is presented using spectral decompositions, and a variant substructuring approach where intersubstructure continuity is enforced in a weak form is derived. This leads to a hybrid formulation of the basic method which is particularly suitable for assembling heterogeneous substructures and analyzing nonconforming and incompatible finite element substructure models. For problems where both the basic and hybrid methods are applicable, the hybrid variant can be computationally more advantageous.

Farhat, Charbel↗
NASA NTRS2016

Aircraft Design

The present invention is an aircraft wing design that creates a bell shaped span load, which results in a negative induced drag (induced thrust) on the outer portion of the wing; such a design obviates the need for rudder control of an aircraft.

Bowers, Albion H.↗
NASA NTRS2020

Fiber-Optic Strain-Based Deflection and Twist Sensing for a High-Aspect-Ratio Swept Wing

Designs of aircraft structures have been moving toward leaner, lightweight designs for increased fuel efficiency. The Passive Aeroelastic Tailored (PAT) wing developed under the NASA Advanced Air Transport Technology (AATT) project is an example of a swept-wing design with high aspect ratio that incorporates lightweight highly-flexible tailored composite construction. The passive aeroelastic tailored structural design has explored the design space to enable aeroelastically tailored wing structures to increase aspect ratios (from 9 to 14) and ultimately reduce weight by 20 percent to 25 percent without impacting aeroelastic performance. To further study the aeroelastic performance of such a wing, the NASA Armstrong Flight Research Center (AFRC) (Edwards, California) has developed efficient real-time structural algorithms that are used in conjunction with a fiber-optic measurement system for lightweight vehicle applications. The AFRC Fiber Optic Sensing System (FOSS) provides up to 8,000 distributed surface strain measurements at one-half-inch increments and can be used to estimate a variety of structural parameters such as shape and load. This report discusses the implementation of strain-based displacement and twist-sensing techniques applied to the PAT wing test article tested at the NASA AFRC Flight Loads Laboratory. Empirical FOSS strain data are collected under varying loading conditions. Strain data are processed with the displacement and twist-sensing algorithms and independently verified by comparison to conventional ground-based instrumentation.

Pena, Francisco↗
NASA NTRS1983

Specification for a Program for an Interative Aeroelastic Solution (PIAS)

An engineering and software specification which was written for a computer program to calculate aeroelastic structural loads including the effects of nonlinear aerodynamics is presented. The procedure used in the program for an iterative aeroelastic solution (PIAS) is to alternately execute two computer codes: one to calculate aerodynamic loads for a specific wing shape, and another to calculate the deflected shape caused by this loading. A significant advantage to the design of PIAS is that the initial aerodynamic module can be replaced with others. The leading edge vortex (LEV) program is used as the aerodynamic module in PIAS. This provides the capability to calculate aeroelastic loads, including the effects of a separation induced leading edge vortex. The finite element method available in ATLAS Integrated structural analysis and design system is used to determine the deflected wing shape for the applied aerodynamics and inertia loads. The data management capabilities in ATLAS are used by the execution control monitor (ECM) of PIAS to control the solution process.

Manro, M. E.↗
NASA NTRS1992

Effect of crash pulse shape on seat stroke requirements for limiting loads on occupants of aircraft

An analytical study was made to provide comparative information on various crash pulse shapes that potentially could be used to test seats under conditions included in Federal Regulations Part 23 Paragraph 23.562(b)(1) for dynamic testing of general aviation seats, show the effects that crash pulse shape can have on the seat stroke requirements necessary to maintain a specified limit loading on the seat/occupant during crash pulse loadings, compare results from certain analytical model pulses with approximations of actual crash pulses, and compare analytical seat results with experimental airplace crash data. Structural and seat/occupant displacement equations in terms of the maximum deceleration, velocity change, limit seat pan load, and pulse time for five potentially useful pulse shapes were derived; from these, analytical seat stroke data were obtained for conditions as specified in Federal Regulations Part 23 Paragraph 23.562(b)(1) for dynamic testing of general aviation seats.

Carden, Huey D.↗
NASA NTRS1994

The buckling response of symmetrically laminated composite plates having a trapezoidal planform area

The focus of this work is the buckling response of symmetrically laminated composite plates having a planform area in the shape of an isosceles trapezoid. The loading is assumed to be inplane and applied perpendicular to the parallel ends of the plate. The tapered edges of the plate are assumed to have simply supported boundary conditions, while the parallel ends are assumed to have either simply supported or clamped boundary conditions. A semi-analytic closed-form solution based on energy principles and the Trefftz stability criterion is derived and solutions are obtained using the Rayleigh-Ritz method. Intrinsic in this solution is a simplified prebuckling analysis which approximates the inplane force resultant distributions by the forms Nx=P/W(x) and Ny=Nxy=0, where P is the applied load and W(x) is the plate width which, for the trapezoidal planform, varies linearly with the lengthwise coordinate x. The out-of-plane displacement is approximated by a double trigonometric series. This analysis is posed in terms of four nondimensional parameters representing orthotropic and anisotropic material properties, and two nondimensional parameters representing geometric properties. For comparison purposes, a number of specific plate geometry, ply orientation, and stacking sequence combinations are investigated using the general purpose finite element code ABAQUS. Comparison of buckling coefficients calculated using the semi-analytical model and the finite element model show agreement within 5 percent, in general, and within 15 percent for the worst cases. In order to verify both the finite element and semi-analytical analyses, buckling loads are measured for graphite/epoxy plates having a wide range of plate geometries and stacking sequences. Test fixtures, instrumentation system, and experimental technique are described. Experimental results for the buckling load, the buckled mode shape, and the prebuckling plate stiffness are presented and show good agreement with the analytical results regarding the buckling load and the prebuckling plate stiffness. However, the experimental results show that for some cases the analysis underpredicts the number of halfwaves in the buckled mode shape. In the context of the definitions of taper ratio and aspect ratio used in this study, it is concluded that the buckling load always increases as taper ratio increases for a given aspect ratio for plates having simply supported boundary conditions on the parallel ends. There are combinations of plate geometry and ply stackling sequences, however, that reverse this trend for plates having clamped boundary conditions on the parallel ends such that an increase in the taper ratio causes a decrease in the buckling load. The clamped boundary conditions on the parallel ends of the plate are shown to increase the buckling load compared to simply supported boundary conditions. Also, anisotropy (the D16 and D26 terms) is shown to decrease the buckling load and skew the buckled mode shape for both the simply supported and clamped boundary conditions.

Radloff, H. D., II↗
NASA NTRS1990

Mobility power flow analysis of an L-shaped plate structure subjected to distributed loading

An analytical investigation based in the Mobility Power Flow (MPF) method is presented for the determination of the vibrational response and power flow for two coupled flat plate structures in an L-shaped configuration, subjected to distributed excitation. The principle of the MPF method consists of dividing the global structure into a series of subsystems coupled together using mobility functions. Each separate subsystem is analyzed independently to determine the structural mobility functions for the junction and excitation locations. The mobility functions, together with the characteristics of the junction between the subsystems, are then used to determine the response of the global structure and the MPF. In the considered coupled plate structure, MPF expressions are derived for distributed mechanical excitation which is independent of the structure response. However using a similar approach with some modifications excitation by an acoustic plane wave can be considered. Some modifications are required to deal with the latter case are necessary because the forces (acoustic pressure) acting on the structure are dependent on the response of the structure due to the presence of the scattered pressure.

Cuschieri, J. M.↗
NASA NTRS1930

Pressure distribution over a Douglas wing tip on a biplane in flight

This note presents the results obtained in pressure distribution tests on the right upper wing panel and tip of a Douglas M-3 airplane in flight. These tests are a part of a more extensive investigation of the effect of changes in tip shape on the load distribution, the tip reported herein being the first of a series of tip shapes being tested. The results are given in tables and curves in such form that the load distribution for any conditions may be determined easily. Tests were made at Langley Field by the NACA in the spring of 1930.

Rhode, Richard V↗
NASA NTRS2001

Projection Moire Interferometry Measurements of Micro Air Vehicle Wings

Projection Moire Interferometry (PMI) has been used to measure the structural deformation of micro air vehicle (MAV) wings during a series of wind tunnel tests. The MAV wings had a highly flexible wing structure, generically reminiscent of a bat s wing, which resulted in significant changes in wing shape as a function of MAV angle-of-attack and simulated flight speed. This flow-adaptable wing deformation is thought to provide enhanced vehicle stability and wind gust alleviation compared to rigid wing designs. Investigation of the potential aerodynamic benefits of a flexible MAV wing required measurement of the wing shape under aerodynamic loads. PMI was used to quantify the aerodynamically induced changes in wing shape for three MAV wings having different structural designs and stiffness characteristics. This paper describes the PMI technique, its application to MAV testing, and presents a portion of the PMI data acquired for the three different MAV wings tested.

Fleming, Gary A.↗
NASA NTRS2018

Active In-Flight Load Redistribution Utilizing Fiber-Optic Shape Sensing and Multiple Control Surfaces

Morphing wing technologies have gained research interest in recent years as technological advancements pave the way for such innovations. A key benefit of such a morphing wing concept is the ability of the wing to transition into an optimal configuration at multiple flight conditions. Such a morphing wing will have applications not only in drag reduction but also in flutter suppression and gust alleviation. By manipulating the wing geometry to match a given flight profile it is likely that the wing will yield increases in not just aerodynamic efficiency but also structural efficiency. These structurally efficient designs will likely rely on some type of structural sensing system which will ensure the wing maintains positive margins throughout its flight profile.

FOSS↗
NASA NTRS1994

Piezoelectric loudspeaker

A piezoelectric loudspeaker suitable for midrange frequencies uses a dome shaped piezoelectric actuator to drive a speaker membrane directly. The dome shaped actuator is made from a reduced and internally biased oxygen wafer, and generates excursion of the apex of the dome in the order of 0.02 - 0.05 inches when a rated drive voltage of 350 V rms is applied between the convex and the concave surfaces of the dome shaped actuator. The load capacity exceeds 10 lbs. The edge of the rim of the dome shaped actuator must be free to rock when the dome height varies to ensure low distortion in the loudspeaker. This is achieved by mounting the rim of the dome shaped actuator on a support surface by prestress only. An exceptionally simple design uses a planar speaker membrane with the center part of one side pressed against the rim of a dome shaped actuator by prestress from a stretched latex surround member.

Regan, Curtis Randall↗
NASA NTRS1998

High Displacement Solid State Ferroelectric Loudspeaker

A piezoelectric loudspeaker suitable for midrange frequencies uses a dome shaped piezoelectric actuator to a speaker membrane directly is discussed. The dome shaped actuator is made from a reduced and internally biased oxygen wafer, and generates excursion of the apex of the dome in the order of 0.02-0.05 inches when a rated drive voltage of 350 V rms is applied between the convex and the concave surface of the dome shaped actuator. The load capacity exceeds 10 lbs. The edge of the rim of the dome shaped actuator must be free to rock when the dome height varies to ensure low distortion in the loudspeaker. This is achieved by mounting the rim of the dome shaped actuator on a support surface by prestress only. An exceptionally simple design uses a planar speaker membrane with the center part of one side pressed against the rim of a dome shaped actuator by prestress from a stretched latex surround member.

Regan, Curtis R.↗
NASA NTRS1976

The development of large diameter, high pressure, cryogenic radial static seals

Two configurations of radial static seals are developed for high-pressure 773 kg per sq cm cryogenic applications: a U-shaped fluorocarbon seal and a delta-shaped PTFE seal. The U-seal is a common commercial design consisting of a spring-loaded cup-shaped jacket; the basic principle is that the soft plastic jacket provides the interface between the piston and the bore, the jacket being held in place by a metal spring. The delta seal is based on the principle that the soft plastic delta cross section furnishes the interface at the metal face and circumferential face of the gland and cylinder. Test results indicate that both the plastic U-seal and the delta seal designs are sufficiently flexible to accommodate assembly in bore or piston grooves. Of these two configurations, the delta seal is found to be superior as it meets all the design requirements for proof pressure testing the turbopump components of the main engine in the Space Shuttle.

Burr, M. E.↗
NASA NTRS1954

A revised gust-load formula and a re-evaluation of v-g data taken on civil transport airplanes from 1933 to 1950

A revised gust-load formula with a new gust factor is derived to replace the gust-load formula and alleviation factor widely used in gust studies. The revised formula utilizes the same principles and retains the same simple form of the original formula but provides a more appropriate and acceptable basis for gust-load calculations. The gust factor is calculated on the basis of a one-minus-cosine gust shape and is presented as a function of a mass-ratio parameter in contrast to the ramp gust shape and wing loading, respectively, used for the alleviation factor. A summary of gust-velocity data from v-g records taken on civil transport airplanes from 1933 to 1950, reevaluated by the revised formula, is also presented.

Pratt, Kermit G↗
NASA NTRS1946

Numerical procedures for the calculation of the stresses in monocoques III : calculation of the bending moments in fuselage frames

This report deals with the calculation of the bending moments in and the distortions of fuselage rings upon which known concentrated and distributed loads are acting. In the procedure suggested, the ring is divided into a number of beams each having a constant radius of curvature. The forces and moments caused in the end sections of the beams by individual unit displacements of the end sections are listed in a table designated as the operations table in conformity with Southwell's nomenclature. The operations table and the external loads are equivalent to a set of linear equations. For their solution the following three procedures are presented: 1) Southwell's method of systematic relaxations. This is a step-by-step approximation procedure guided by the physical interpretation of the changes in the values of the unknown. 2) The growing unit procedure in which the individual beams are combined successively into beams of increasing length until finally the entire ring becomes a single beam. In each step of the procedure a set of not more than three simultaneous linear equations is solved. 3) Solution of the entire set of simultaneous equations by the methods of the matrix calculus. In order to demonstrate the manner in which the calculations may be carried out, the following numerical examples are worked out: 1) Curved beam with both its end sections rigidly fixed. The load is a concentrated force. 2) Egg-shape ring with symmetric concentrated loads. 3) Circular ring with antisymmetric concentrated loads and shear flow (torsion of the fuselage). 4) Same with V-braces incorporated in the ring. 5) Egg-shape ring with antisymmetric concentrated loads and shear flow (torsion of the fuselage). 6) Same with V-braces incorporated in the ring. The results of these calculations are checked, whenever possible, by calculations carried out according to known methods of analysis. The agreement is found to be good. The amount of work necessary for the solution of ring problems by the methods described in the present report is practically independent of the degree of redundancy of the structure. For this reason the methods are recommended for use particularly in problems of rings having one or more internal bracing elements.

LOADS - FUSELAGES↗
NASA NTRS2015

Structural Testing of a 6m Hypersonic Inflatable Aerodynamic Decelerator System

NASA is developing low ballistic coefficient technologies to support the Nations long-term goal of landing humans on Mars. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current and future launch vehicle fairing limitations. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) are being developed to circumvent this limitation and are now considered a leading technology to enable landing of heavy payloads on Mars. At the beginning of 2014, a 6m diameter HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify its structural performance under flight-relevant loads. The inflatable structure was constructed into a 60 degree sphere-cone configuration using nine inflatable torus segments composed of fiber-reinforced thin films. The inflatable tori were joined together using adhesives and high-strength textile woven structural straps. These straps help distribute the load throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials that are designed to protect the inflatable structure from heat loads that would be seen in flight during atmospheric entry. A custom test fixture was constructed to perform the static load test series. The fixture consisted of a round structural tub with enough height and width to allow for displacement of the HIAD test article as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The rigid centerbody of the HIAD was mounted to a pedestal in the center of the structural tub. Using an impermeable membrane draped over the HIAD test article, an airtight seal was created with the top rim of the static load tub. This seal allowed partial vacuum to be pulled beneath the HIAD resulting in a uniform static pressure load applied to the outer surface. Using this technique, the test article was subjected to loads of up to 50,000lbs. During the test series an extensive amount of instrumentation was used to provide a rich data set, including deflected shape, structural strap loads, torus cord loads, inflation pressures, and applied static load. In this paper the 2014 6m HIAD static load test series will be discussed in detail, including the design of the 6m HIAD test article, the test setup, and test execution. Analysis results will be described supporting the conclusions that were drawn from the test series..

Hypersonic↗