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At least 991 records · Page 55
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Equivalent Plate Structural Modeling for Wing Shape Optimization Including Transverse Shear

A new technique for structural modeling of airplane wings is presented taking transverse shear effects into account. The kinematic assumptions of first-order shear deformation plate theory In combination with numerical analysis, where simple polynomials are used to define geometry, construction, and displacement approximations, lead to analytical expressions for elements of the stiffness and mass matrices and load vector. Contributions from the cover skins, spar and rib caps, and spar and rib webs are included as well as concentrated springs and concentrated masses. Limitations of wing modeling techniques based on classical plate theory are discussed, and the Improved accuracy of the new equivalent plate technique is demonstrated through comparison with finite element analysis and test results. Expressions for analytical derivatives of stiffness, mass, and load terms with respect to wing shape are given. Based on these, it is possible to obtain analytic sensitivities of displacements, stresses, and natural frequencies with respect to planform shape and depth distribution. This makes the new capability an effective structural tool for wing shape optimization.

Livne, Eli↗
NASA NTRS1947

The Load Distribution in Bolted or Riveted Joints in Light-Alloy Structures

This report contains a theoretical discussion of the load distribution in bolted or riveted joints in light-alloy structures which is applicable not only for loads below the limit of proportionality but also for loads above this limit. The theory is developed for double and single shear joints. The methods given are illustrated by numerical examples and the values assumed for the bolt (or rivet) stiffnesses are based partly on theory and partly on known experimental values. It is shown that the load distribution does not vary greatly with the bolt (or rivet) stiffnesses and that for design purposes it is usually sufficient to know their order of magnitude. The theory may also be directly used for spot-welded structures and, with small modifications, for seam-welded structures, The computational work involved in the methods described is simple and may be completed in a reasonable time for most practical problems. A summary of earlier theoretical and experimental investigations on the subject is included in the report.

Vogt, F.↗
NASA NTRS1996

Vibration Characteristics of Partially Covered Double-Sandwich Cantilever Beam

The differential equations of motion together with the boundary conditions for a partially covered, double-sandwich cantilever beam are derived. Bending and extension, rotational and longitudinal inertia of damping layers, and shear deformation and rotational and longitudinal inertia of the constraining layers and the primary beam are included in the equations. The theory is applicable for long as well as short, soft, or stiff damping layer, double-sandwich beams. Also, the effects of different parameters on the system loss factor and resonance frequency are discussed. Differences are found to exist with the previous beam model (called the Euler beam model) when the damping layers are stiff, when the thickness of the damping layer is large compared to the primary-beam thickness, and in the case of higher modes of vibration.

Chen, Qinghua↗
NASA NTRS1998

Recent Advances in the Method of Forces: Integrated Force Method of Structural Analysis

Stress that can be induced in an elastic continuum can be determined directly through the simultaneous application of the equilibrium equations and the compatibility conditions. In the literature, this direct stress formulation is referred to as the integrated force method. This method, which uses forces as the primary unknowns, complements the popular equilibrium-based stiffness method, which considers displacements as the unknowns. The integrated force method produces accurate stress, displacement, and frequency results even for modest finite element models. This version of the force method should be developed as an alternative to the stiffness method because the latter method, which has been researched for the past several decades, may have entered its developmental plateau. Stress plays a primary role in the development of aerospace and other products, and its analysis is difficult. Therefore, it is advisable to use both methods to calculate stress and eliminate errors through comparison. This paper examines the role of the integrated force method in analysis, animation and design.

Patnaik, Surya N.↗
NASA NTRS1999

A Discussion of Zero Spring Rate Mechanisms Used for the Active Isolation Mount Experiment

In the summer of 1995 the Structural Dynamics Branch at NASA Langley Research Center set out to conceive a small, lightweight, low frequency isolation mount that could be used for spaceflight experiments. The Engineering Design Branch undertook the task of developing the isolation mount. This report describes the engineering process that led to three phases of a study entitled "Active Isolation Mounts" (AIM). A zero spring rate mechanism was used to achieve low fundamental frequencies for a payloads in the 1 to 10 pound range. It worked by balancing both a positive and a negative stiffness so that the net result was a small positive stiffness. The study demonstrated devices that could reduce the initial corner frequency by a factor of six for brief periods and a factor of two for extended periods. The designs were relatively simple and minimized weight, volume, and power. They could be scaled down and they were made of spaceflight compatible materials. All designs offered the ability to continuously vary the fundamental frequency. Yet, the goal of reducing the frequency by an order of magnitude was not achieved because the systems were too unstable at low frequencies. There was a trade between performance and stability.

Teter, John E., Jr.↗
NASA NTRS2000

Equivalent Skin Analysis of Wing Structures Using Neural Networks

An efficient method of modeling trapezoidal built-up wing structures is developed by coupling. in an indirect way, an Equivalent Plate Analysis (EPA) with Neural Networks (NN). Being assumed to behave like a Mindlin-plate, the wing is solved using the Ritz method with Legendre polynomials employed as the trial functions. This analysis method can be made more efficient by avoiding most of the computational effort spent on calculating contributions to the stiffness and mass matrices from each spar and rib. This is accomplished by replacing the wing inner-structure with an "equivalent" material that combines to the skin and whose properties are simulated by neural networks. The constitutive matrix, which relates the stress vector to the strain vector, and the density of the equivalent material are obtained by enforcing mass and stiffness matrix equities with rec,ard to the EPA in a least-square sense. Neural networks for the material properties are trained in terms of the design variables of the wing structure. Examples show that the present method, which can be called an Equivalent Skin Analysis (ESA) of the wing structure, is more efficient than the EPA and still fairly good results can be obtained. The present ESA is very promising to be used at the early stages of wing structure design.

Liu, Youhua↗
NASA NTRS2000

Adjoint Techniques for Topology Optimization of Structures Under Damage Conditions

The objective of this cooperative agreement was to seek computationally efficient ways to optimize aerospace structures subject to damage tolerance criteria. Optimization was to involve sizing as well as topology optimization. The work was done in collaboration with Steve Scotti, Chauncey Wu and Joanne Walsh at the NASA Langley Research Center. Computation of constraint sensitivity is normally the most time-consuming step of an optimization procedure. The cooperative work first focused on this issue and implemented the adjoint method of sensitivity computation (Haftka and Gurdal, 1992) in an optimization code (runstream) written in Engineering Analysis Language (EAL). The method was implemented both for bar and plate elements including buckling sensitivity for the latter. Lumping of constraints was investigated as a means to reduce the computational cost. Adjoint sensitivity computation was developed and implemented for lumped stress and buckling constraints. Cost of the direct method and the adjoint method was compared for various structures with and without lumping. The results were reported in two papers (Akgun et al., 1998a and 1999). It is desirable to optimize topology of an aerospace structure subject to a large number of damage scenarios so that a damage tolerant structure is obtained. Including damage scenarios in the design procedure is critical in order to avoid large mass penalties at later stages (Haftka et al., 1983). A common method for topology optimization is that of compliance minimization (Bendsoe, 1995) which has not been used for damage tolerant design. In the present work, topology optimization is treated as a conventional problem aiming to minimize the weight subject to stress constraints. Multiple damage configurations (scenarios) are considered. Each configuration has its own structural stiffness matrix and, normally, requires factoring of the matrix and solution of the system of equations. Damage that is expected to be tolerated is local and represents a small change in the stiffness matrix compared to the baseline (undamaged) structure. The exact solution to a slightly modified set of equations can be obtained from the baseline solution economically without actually solving the modified system.. Shennan-Morrison-Woodbury (SMW) formulas are matrix update formulas that allow this (Akgun et al., 1998b). SMW formulas were therefore used here to compute adjoint displacements for sensitivity computation and structural displacements in damaged configurations.

Akgun, Mehmet A.↗
NASA NTRS1999

Optimization of Aerospace Structure Subject to Damage Tolerance Criteria

The objective of this cooperative agreement was to seek computationally efficient ways to optimize aerospace structures subject to damage tolerance criteria. Optimization was to involve sizing as well as topology optimization. The work was done in collaboration with Steve Scotti, Chauncey Wu and Joanne Walsh at the NASA Langley Research Center. Computation of constraint sensitivity is normally the most time-consuming step of an optimization procedure. The cooperative work first focused on this issue and implemented the adjoint method of sensitivity computation in an optimization code (runstream) written in Engineering Analysis Language (EAL). The method was implemented both for bar and plate elements including buckling sensitivity for the latter. Lumping of constraints was investigated as a means to reduce the computational cost. Adjoint sensitivity computation was developed and implemented for lumped stress and buckling constraints. Cost of the direct method and the adjoint method was compared for various structures with and without lumping. The results were reported in two papers. It is desirable to optimize topology of an aerospace structure subject to a large number of damage scenarios so that a damage tolerant structure is obtained. Including damage scenarios in the design procedure is critical in order to avoid large mass penalties at later stages. A common method for topology optimization is that of compliance minimization which has not been used for damage tolerant design. In the present work, topology optimization is treated as a conventional problem aiming to minimize the weight subject to stress constraints. Multiple damage configurations (scenarios) are considered. Each configuration has its own structural stiffness matrix and, normally, requires factoring of the matrix and solution of the system of equations. Damage that is expected to be tolerated is local and represents a small change in the stiffness matrix compared to the baseline (undamaged) structure. The exact solution to a slightly modified set of equations can be obtained from the baseline solution economically without actually solving the modified system. Sherrnan-Morrison-Woodbury (SMW) formulas are matrix update formulas that allow this. SMW formulas were therefore used here to compute adjoint displacements for sensitivity computation and structural displacements in damaged configurations.

Akgun, Mehmet A.↗
NASA NTRS1998

Recent Advances in the Method of Forces: Integrated Force Method of Structural Analysis

Stress that can be induced in an elastic continuum can be determined directly through the simultaneous application of the equilibrium equations and the compatibility conditions. In the literature, this direct stress formulation is referred to as the integrated force method. This method, which uses forces as the primary unknowns, complements the popular equilibrium-based stiffness method, which considers displacements as the unknowns. The integrated force method produces accurate stress, displacement, and frequency results even for modest finite element models. This version of the force method should be developed as an alternative to the stiffness method because the latter method, which has been researched for the past several decades, may have entered its developmental plateau. Stress plays a primary role in the development of aerospace and other products, and its analysis is difficult. Therefore, it is advisable to use both methods to calculate stress and eliminate errors through comparison. This paper examines the role of the integrated force method in analysis, animation and design.

Patnaik, Surya N.↗
NASA NTRS2000

Preventing Damaging Pressure Gradients at the Walls of an Inflatable Space System

An inflatable structural system to deploy a space system such as a solar shield, an antenna or another similar instrument, requires a stiffening element after it is extended by the inflated gas pressure. The stiffening element has to be packaged in a folded configuration before the deployment. It must be relatively small, lightweight, non-damaging to the inflated system, and be able to become stiff in a short time. One stiffening method is to use a flexible material inserted in the deployable system, which, upon a temperature curing, can become stiff and is capable to support the entire structure. There are two conditions during the space operations when the inflated volume could be damaged: during the transonic region of the launch phase and when the curing of the rigidizing element occurs. In both cases, an excess of pressure within the volume containing the rigid element could burst the walls of the low-pressure gas inflated portion of the system. This paper investigates those two conditions and indicates the vents, which will prevent those damaging overpressures. Vent openings at the non-inflated volumes have been calculated for the conditions existing during the launch. Those vents allow the initially folded volume to exhaust the trapped atmospheric gas at approximately the same rate as the ambient pressure drops. That will prevent pressure gradients across the container walls which otherwise could be as high as 14.7 psi. The other condition occurring during the curing of the stiffening element has been investigated. This has required the testing of the element to obtain the gas generation during the curing and the transformation from a pliable material to a rigid one. The tested material is a composite graphite/epoxy weave. The outgassing of the uncured sample at 121C was carried with the Cahn Microbalance and with other outgassing facilities including the micro-CVCM ASTM E-595 facility. The tests provided the mass of gas evolved during the test. That data, including the chemical nature of the evolved gas, provided the data for the calculation of the pressure produced within the volume. The evaluation of the areas of the vents that would prevent excessive pressures and provide a rapid release of the gas away from contamination sensitive surfaces has been carried out. The pressure decay with time has been indicated.

Scialdone, John J.↗
NASA NTRS2000

Preventing Damaging Pressure Gradients at the Walls of an Inflatable Space System

An inflatable structural system to deploy a space system such as a solar shield, an antenna or another similar instrument requires a stiffening element after it is extended by the inflated gas pressure. The stiffening element has to be packaged in folded configuration before the deployment. It must be relatively small, lightweight, non-damaging to the inflated system and be able to become stiff in a short time. One stiffening method is to use a flexible material inserted in the deployable system, which, upon a temperature curing, can become stiff and is capable of supporting the entire structure. There are two conditions during the space operations when the inflated volume could be damaged: during the transonic region of the launch phase and when the curing of the rigidizing element occurs. In both cases, an excess of pressure within the volume containing the rigid element could burst the walls of the low-pressure gas inflated portion of the system. This paper investigates those two conditions and indicates the vents, which will prevent those damaging overpressures. Vent openings at the non-inflated volumes have been calculated for the conditions existing during the launch. Those vents allow the initially folded volume to exhaust the trapped atmospheric gas at approximately the same rate as the ambient pressure drops. That will prevent pressure gradients across the container walls which otherwise could be as high as 14.7 psi. The other condition occurring during the curing of the stiffening element has been investigated. This has required the testing of the element to obtain the gas generation during the curing and the transformation from a pliable material to a rigid on The tested material is a composite graphite/epoxy weave. The outgassing of the uncured sample at 121 deg Celcius was carried with the Cahn Microbalance and with other outgassing facilities including the micro-CVCM ASTM E-595 facility. The test provided the mass of gas evolved during the test. That data, including the chemical nature of the evolved gas, provided the data for the calculation of the pressure produced within the volume. The evaluation of the areas of the vents that would prevent excessive pressures and provide a rapid release of the gas away from contamination sensitive surfaces has be carried out. The pressure decay with time has been indicated.

Scialdone, John J.↗
NASA NTRS2000

Precision of Four Acoustic Bone Measurement Devices

Though many studies have quantified the precision of various acoustic bone measurement devices, it is difficult to directly compare the results among the studies, because they used disparate subject pools, did not specify the estimation methodology, or did not use consistent definitions for various precision characteristics. In this study, we used a repeated measures design protocol to directly determine the precision characteristics of four acoustic bone measurement devices: the Mechanical Response Tissue Analyzer (MRTA), the UBA-575+, the SoundScan 2000 (S2000), and the Sahara Ultrasound Done Analyzer. Ten men and ten women were scanned on all four devices by two different operators at five discrete time points: Week 1, Week 2, Week 3, Month 3 and Month 6. The percent coefficient of variation (%CV) and standardized coefficient of variation were computed for the following precision characteristics: interoperator effect, operator-subject interaction, short-term error variance, and long-term drift, The MRTA had high interoperator errors for its ulnar and tibial stiffness measures and a large long-term drift in its tibial stiffness measurement. The UBA-575+ exhibited large short-term error variances and long-term drift for all three of its measurements. The S2000's tibial speed of sound measurement showed a high short-term error variance and a significant operator-subject interaction but very good values ( < 1%) for the other precision characteristics. The Sahara seemed to have the best overall performance, but was hampered by a large %CV for short-term error variance in its broadband ultrasound attenuation measure.

Miller, Christopher↗
NASA NTRS2000

Precision of Four Acoustic Bone Measurement Devices

Though many studies have quantified the precision of various acoustic bone measurement devices, it is difficult to directly compare the results among the studies, because they used disparate subject pools, did not specify the estimation methodology, or did not use consistent definitions for various precision characteristics. In this study, we used a repeated measures design protocol to directly determine the precision characteristics of four acoustic bone measurement devices: the Mechanical Response Tissue Analyzer (MRTA), the UBA-575+, the SoundScan 2000 (S2000), and the Sahara Ultrasound Bone Analyzer. Ten men and ten women were scanned on all four devices by two different operators at five discrete time points: Week 1, Week 2, Week 3, Month 3 and Month 6. The percent coefficient of variation (%CV) and standardized coefficient of variation were computed for the following precision characteristics: interoperator effect, operator-subject interaction, short-term error variance, and long-term drift. The MRTA had high interoperator errors for its ulnar and tibial stiffness measures and a large long-term drift in its tibial stiffness measurement. The UBA-575+ exhibited large short-term error variances and long-term drift for all three of its measurements. The S2000's tibial speed of sound measurement showed a high short-term error variance and a significant operator-subject interaction but very good values (less than 1%) for the other precision characteristics. The Sahara seemed to have the best overall performance, but was hampered by a large %CV for short-term error variance in its broadband ultrasound attenuation measure.

Miller, Christopher↗
NASA NTRS2000

Effective Widths of Compression-Loaded Plates With a Cutout

A study of the effects of cutouts and laminate construction on the prebuckling and initial postbuckling stiffnesses, and the effective widths of compression-loaded, laminated-composite and aluminum square plates is presented. The effective-width concept is extended to plates with cutouts, and experimental and nonlinear finite-element analysis results are presented. Behavioral trends are compared for seven plate families and for cutout-diameter-to-plate-width ratios up to 0.66. A general compact design curve that can be used to present and compare the effective widths for a wide range of laminate constructions is also presented. A discussion of how the results can be used and extended to include certain types of damage, cracks, and other structural discontinuities or details is given. Several behavioral trends are described that initially appear to be nonintuitive. The results demonstrate a complex interaction between cutout size and plate orthotropy that affects the axial stiffness and effective width of a plate subjected to compression loads.

Hilburger, Mark W.↗
NASA NTRS2000

Dynamic Investigation of Static Divergence: Analysis and Testing

The phenomenon known as aeroelastic divergence is the focus of this work. The analyses and experiment presented here show that divergence can occur without a structural dynamic mode losing its oscillatory nature. Aeroelastic divergence occurs when the structural restorative capability or stiffness of a structure is overwhelmed by the static aerodynamic moment. This static aeroelastic coupling does not require the structural dynamic system behavior to cease, however. Aeroelastic changes in the dynamic mode behavior are governed not only by the stiffness, but by damping and inertial properties. The work presented here supports these fundamental assertions by examining a simple system: a typical section airfoil with only a rotational structural degree of freedom. Analytical results identified configurations that exhibit different types of dynamic mode behavior as the system encounters divergence. A wind tunnel model was designed and tested to examine divergence experimentally. The experimental results validate the analytical calculations and explicitly examine the divergence phenomenon where the dynamic mode persists. Three configurations of the wind tunnel model were tested. The experimental results agree very well with the analytical predictions of subcritical characteristics, divergence velocity, and behavior of the noncritical dynamic mode at divergence.

Heeg, Jennifer↗
NASA NTRS2001

Structural Modeling of a Five-Meter Thin Film Inflatable Antenna/Concentrator With Rigidized Support Struts

Dynamic characterization of a non-rigidized thin film inflatable antenna/solar concentrator structure with rigidized composite support struts is described in detail. A two-step finite element modeling approach in MSC/NASTRAN is utilized, consisting of: (1) a nonlinear static pressurization procedure used to obtain the updated stiffness matrix, and (2) a modal "restart" eigen solution that uses the modified stiffness matrix. Unique problems encountered in modeling of this large 5-m lightweight inflatable are identified, including considerable difficulty in obtaining convergence in the nonlinear pressurization solution. It was found that the extremely thin polyimide film material (.001 in or I mil) presents tremendous problems in obtaining a converged solution when internal pressure loading is applied. It was concluded that the ratios of film thickness to other geometric dimensions such as torus cross-sectional and ring diameter and lenticular diameter are the critical parameters for convergence of the pressurization procedure. Comparison of finite element predictions for frequency and mode shapes with experimental results indicated reasonable agreement considering the complexity of the structure, the film-to-air interaction, and the nonlinear material properties of the film. It was also concluded that analysis should be done using different finite element to codes to determine if a more robust and stable solution can be obtained.

Smalley, Kurt B.↗
NASA NTRS2001

Vibroacoustic Model Validation for a Curved Honeycomb Composite Panel

Finite element and boundary element models are developed to investigate the vibroacoustic response of a curved honeycomb composite sidewall panel. Results from vibroacoustic tests conducted in the NASA Langley Structural Acoustic Loads and Transmission facility are used to validate the numerical predictions. The sidewall panel is constructed from a flexible honeycomb core sandwiched between carbon fiber reinforced composite laminate face sheets. This type of construction is being used in the development of an all-composite aircraft fuselage. In contrast to conventional rib-stiffened aircraft fuselage structures, the composite panel has nominally uniform thickness resulting in a uniform distribution of mass and stiffness. Due to differences in the mass and stiffness distribution, the noise transmission mechanisms for the composite panel are expected to be substantially different from those of a conventional rib-stiffened structure. The development of accurate vibroacoustic models will aide in the understanding of the dominant noise transmission mechanisms and enable optimization studies to be performed that will determine the most beneficial noise control treatments. Finite element and boundary element models of the sidewall panel are described. Vibroacoustic response predictions are presented for forced vibration input and the results are compared with experimental data.

Buehrle, Ralph D.↗
NASA NTRS2001

A Wind-Tunnel Parametric Investigation of Tiltrotor Whirl-Flutter Stability Boundaries

A wind-tunnel investigation of tiltrotor whirl-flutter stability boundaries has been conducted on a 1/5-size semispan tiltrotor model known as the Wing and Rotor Aeroelastic Test System (WRATS) in the NASA-Langley Transonic Dynamics Tunnel as part of a joint NASA/Army/Bell Helicopter Textron, Inc. (BHTI) research program. The model was first developed by BHTI as part of the JVX (V-22) research and development program in the 1980's and was recently modified to incorporate a hydraulically-actuated swashplate control system for use in active controls research. The modifications have changed the model's pylon mass properties sufficiently to warrant testing to re-establish its baseline stability boundaries. A parametric investigation of the effect of rotor design variables on stability was also conducted. The model was tested in both the on-downstop and off-downstop configurations, at cruise flight and hover rotor rotational speeds, and in both air and heavy gas (R-134a) test mediums. Heavy gas testing was conducted to quantify Mach number compressibility effects on tiltrotor stability. Experimental baseline stability boundaries in air are presented with comparisons to results from parametric variations of rotor pitch-flap coupling and control system stiffness. Increasing the rotor pitch-flap coupling (delta(sub 3) more negative) was found to have a destabilizing effect on stability, while a reduction in control system stiffness was found to have little effect on whirl-flutter stability. Results indicate that testing in R-134a, and thus matching full-scale tip Mach number, has a destabilizing effect, which demonstrates that whirl-flutter stability boundaries in air are unconservative.

Piatak, David J.↗