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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Application of formal optimization techniques in thermal/structural design of a heat-pipe-cooled panel for a hypersonic vehicle

Nonlinear mathematical programming methods are used to design a radiantly cooled and heat-pipe-cooled panel for a Mach 6.7 transport. The cooled portion of the panel is a hybrid heat-pipe/actively cooled design which uses heat pipes to transport the absorbed heat to the ends of the panel where it is removed by active cooling. The panels are optimized for minimum mass and to satisfy a set of heat-pipe, structural, geometric, and minimum-gage constraints. Two panel concepts are investigated: cylindrical heat pipes embedded in a honeycomb core and an integrated design which uses a web-core heat-pipe sandwich concept. The latter was lighter and resulted in a design which was less than 10 percent heavier than an all actively cooled concept. The heat-pipe concept, however, is redundant and can sustain a single-point failure, whereas the actively cooled concept cannot. An additional study was performed to determine the optimum number of coolant manifolds per panel for a minimum-mass design.

Camarda, Charles J.↗
NASA NTRS1988

Composite panel development at JPL

Parametric computer studies can be use in a cost effective manner to determine optimized composite mirror panel designs. An InterDisciplinary computer Model (IDM) was created to aid in the development of high precision reflector panels for LDR. The materials properties, thermal responses, structural geometries, and radio/optical precision are synergistically analyzed for specific panel designs. Promising panels designs are fabricated and tested so that comparison with panel test results can be used to verify performance prediction models and accommodate design refinement. The iterative approach of computer design and model refinement with performance testing and materials optimization has shown good results for LDR panels.

Mcelroy, Paul↗
NASA NTRS1988

Lightweight composite reflector panels

The Hexel Corp. has produced additional composite panels, based on JPL designs, that: (1) have increased the panel size from 0.15 to 0.40 meters, (2) have improved the as-manufactured surface precision 3.0 to approx. 1.0 micron RMS, (3) have utilized different numbers of face sheet plys, (4) have improved face sheet fiber orientation, (5) have variations of aluminum honeycomb core cell size, (6) have combined graphite/epoxy (Gr/Ep) face sheets with E-glass honeycomb cores, and (7) have used standard aluminum core with face sheets composed of combinations of glass, Kevlar, and carbon fibers. Additionally, JPL has identified candidate alternate materials for the facesheets and core, modified the baseline polymer panel matrix material, and developed new concepts for panel composite cores. Dornier designed and fabricated three 0.6 meter Gr/Ep panels, that were evaluated by JPL. Results of both the Hexel and Dornier panel work were used to characterize the state-of-the-art for Gr/Ep mirrors.

Freeland, R. E.↗
NASA NTRS1990

JANNAF liquid rocket combustion instability panel research recommendations

The Joint Army, Navy, NASA, Air Force (JANNAF) Liquid Rocket Combustion Instability Panel was formed in 1988, drawing its members from industry, academia, and government experts. The panel was charted to address the needs of near-term engine development programs and to make recommendations whose implementation would provide not only sufficient data but also the analysis capabilities to design stable and efficient engines. The panel was also chartered to make long-term recommendations toward developing mechanistic analysis models that would not be limited by design geometry or operating regime. These models would accurately predict stability and thereby minimize the amount of subscale testing for anchoring. The panel has held workshops on acoustic absorbing devices, combustion instability mechanisms, instability test hardware, and combustion instability computational methods. At these workshops, research projects that would meet the panel's charter were suggested. The JANNAF Liquid Rocket Combustion Instability Panel's conclusions about the work that needs to be done and recommendations on how to approach it, based on evaluation of the suggested research projects, are presented.

Klem, Mark D.↗
NASA NTRS1990

JANNAF Liquid Rocket Combustion Instability Panel Research Recommendations

The Joint Army, Navy, NASA, Air Force (JANNAF) Liquid Rocket Combustion Instability Panel was formed in 1988, drawing its members from industry, academia, and government experts. The panel was chartered to address the needs of near-term engine development programs and to make recommendations whose implementation would provide not only sufficient data but also the analysis capabilities to design stable and efficient engines. The panel was also chartered to make long-term recommendations toward developing mechanistic analysis models that would not be limited by design geometry or operating regime. These models would accurately predict stability and thereby minimize the amount of subscale testing for anchoring. The panel has held workshops on acoustic absorbing devices and combustion instability computational methods. At these workshops, research projects that would meet the panel's charter were suggested. The panel's conclusions about the work that needs to be done and recommendations on how to approach it, based on evaluation of the suggested research projects, are presented.

Klem, Mark D.↗
NASA NTRS1991

Optimization of composite sandwich cover panels subjected to compressive loadings

An analysis and design method is presented for the design of composite sandwich cover panels that include the transverse shear effects and damage tolerance considerations. This method is incorporated into a sandwich optimization computer program entitled SANDOP. As a demonstration of its capabilities, SANDOP is used in the present study to design optimized composite sandwich cover panels for for transport aircraft wing applications. The results of this design study indicate that optimized composite sandwich cover panels have approximately the same structural efficiency as stiffened composite cover panels designed to satisfy individual constraints. The results also indicate that inplane stiffness requirements have a large effect on the weight of these composite sandwich cover panels at higher load levels. Increasing the maximum allowable strain and the upper percentage limit of the 0 degree and +/- 45 degree plies can yield significant weight savings. The results show that the structural efficiency of these optimized composite sandwich cover panels is relatively insensitive to changes in core density. Thus, core density should be chosen by criteria other than minimum weight (e.g., damage tolerance, ease of manufacture, etc.).

Cruz, Juan R.↗
NASA NTRS1991

Panel installation development in Automated Structures Assembly Lab

The design requirements for panel installation process are listed and photographs are presented of the following: panel latches to truss via adaptors; panel end effector; arm retrieving panel from canister; arm installing panel; and facility view with panel installed.

Cuong Quach↗
NASA NTRS1992

Finite element nonlinear panel flutter with arbitrary temperatures in supersonic flow

A finite element frequency domain method for predicting nonlinear flutter response of panels with temperature effects is presented. By using the principle of virtual work, the element nonlinear stiffness formulation for a panel under a combined thermal and aerodynamic loads is derived on the bases of von Karman's large deflection plate theory, the first-order piston theory aerodynamics and the quasi-steady thermal stress theory. The system equations of motion can be mathematically separated into two sets of equations and then solved in sequence. The first set of equations yields the panel thermal-aerodynamic equilibrium and the second set of equations of motion leads to the flutter limit-cycle oscillations. Stability and flutter boundaries can also be obtained from the two sets of system equations. Finite element large amplitude limit-cycle flutter results at different uniform temperatures are obtained for a simply supported square panel and are compared with existing Galerkin/time integration and other finite element solutions. Effects of nonuniform temperature distributions, panel length-to-width ratios, and boundary conditions on flutter responses of rectangular and triangular panels are presented.

Xue, David Y.↗
NASA NTRS1991

Titanium honeycomb panel testing

The paper describes the procedures of thermal mechanical tests carried out at the NASA Dryden Flight Research Facility on two tianium honeycomb wing panels bonded using liquid interface diffusion (LID) technique, and presents the results of these tests. The 58.4 cm square panels consisted of two 0.152-cm-thick Ti 6-2-4-2 face sheets LID-bonded to a 1.9-cm-thick honeycomb core, with bearing plates fastened to the perimeter of the upper and the lower panel surfaces. The panels were instrumented with sensors for measuring surface temperature, strain, and deflections to 315 C and 482 C. Thermal stress levels representative of those encountered during aerodynamic heating were produced by heating the upper panel surface and restraining all four edges. After more than 100 thermal cycles from room temperature to 315 C and 50 cycles from room temperature to 482 C, no significant structural degradation was detected in the panels.

Richards, W. L.↗
NASA NTRS1991

Optimization of composite sandwich cover panels subjected to compressive loadings

An analysis and design method is presented for the design of composite sandwich cover panels that includes transverse shear effects and damage tolerance considerations. This method is incorporated into an optimization program called SANDOP (SANDwich OPtimization). SANDOP is used in the present study to design optimized composite sandwich cover panels for transport aircraft wing applications as a demonstration of its capabilities. The results of this design study indicate that optimized composite sandwich cover panels have approximately the same structural efficiency as stiffened composite cover panels designed to identical constraints. Results indicate that inplane stiffness requirements have a large effect on the weight of these composite sandwich cover panels at higher load levels. Increasing the maximum allowable strain and the upper percentage limit of the 0 degree and plus or minus 45 degree plies can yield significant weight savings. The results show that the structural efficiency of these optimized composite sandwich cover panels is relatively insensitive to changes in core density.

Cruz, Juan R.↗
NASA NTRS1992

Nonlinear vibration and radiation from a panel with transition to chaos

The dynamic response of an aircraft panel forced at resonance and off-resonance by plane acoustic waves at normal incidence is investigated experimentally and numerically. Linear, nonlinear (period doubling), and chaotic responses are obtained by increasing the sound pressure level of the excitation. The response time history is sensitive to the input level and to the frequency of excitation. The change in response behavior is due to a change in input conditions, triggered either naturally or by modulation of the bandwidth of the incident waves. Off-resonance bifurcation is diffused and difficult to maintain; thus the panel response drifts into a linear behavior. The acoustic pressure emanated by the panel is either linear or nonlinear as is the vibration response. The nonlinear effects accumulate during the propagation with distance. Results are also obtained on the control of the panel response using damping tape on an aluminum panel and a graphite epoxy panel having the same size and weight. Good agreement is obtained betwen the experimental and numerical results.

Maestrello, Lucio↗
NASA NTRS1992

Nonlinear vibration and radiation from a panel with transition to chaos induced by acoustic waves

The dynamic response of an aircraft panel forced at resonance and off-resonance by plane acoustic waves at normal incidence is investigated experimentally and numerically. Linear, nonlinear (period doubling) and chaotic responses are obtained by increasing the sound pressure level of the excitation. The response time history is sensitive to the input level and to the frequency of excitation. The change in response behavior is due to a change in input conditions, triggered either naturally or by modulation of the bandwidth of the incident waves. Off-resonance, bifurcation is diffused and difficult to maintain, thus the panel response drifts into a linear behavior. The acoustic pressure emanated by the panel is either linear or nonlinear as is the vibration response. The nonlinear effects accumulate during the propagation with distance. Results are also obtained on the control of the panel response using damping tape on aluminum panel and using a graphite epoxy panel having the same size and weight. Good agreement is obtained between the experimental and numerical results.

Maestrello, Lucio↗
NASA NTRS1993

Stiffness, thermal expansion, and thermal bending formulation of stiffened, fiber-reinforced composite panels

A method is presented for formulating stiffness terms and thermal coefficients of stiffened, fiber-reinforced composite panels. The method is robust enough to handle panels with general cross sectional shapes, including those which are unsymmetric and/or unbalanced. Nonlinear, temperature and load dependent constitutive material data of each laminate are used to 'build-up' the stiffened panel membrane, bending, and membrane-bending coupling stiffness terms and thermal coefficients. New thermal coefficients are introduced to quantify panel response from through-the-thickness temperature gradients. A technique of implementing this capability with a single plane of shell finite elements using the MSC/NASTRAN analysis program (FEA) is revealed that provides accurate solutions of entire airframes or engines with coarsely meshed models. An example of a composite, hat-stiffened panel is included to demonstrate errors that occur when an unsymmetric panel is symmetrically formulated as traditionally done. The erroneous results and the correct ones produced from this method are compared to analysis from discretely meshed three-dimensional FEA.

Collier, Craig S.↗
NASA NTRS1991

Development and testing of lightweight composite reflector panels

Several years ago the Jet Propulsion Laboratory embarked upon a program to develop advanced polymer composite mirror elements as enabling technology for orbiting far IR/submillimeter telescopes. Structural composite mirrors have the advantages of high specific stiffness, good thermal stability and low cost that beneficially affect the entire telescope design. The goal of this panel development program is to design and fabricate prototype mirror panels: up to one meter in size with a surface precision and orbital thermal performance of a few microns, to achieve real densities close to 5 kg/sq m and to demonstrate the thermally stable performance of these panels experimentally. Studies leading to current mirror design are summarized. The precision and thermal performance of the mirror panels were tested in a dedicated test facility. Data showing panel performance are presented. Finally, the test results are compared to the analytically predicted panel performance.

Helms, R. G.↗
NASA NTRS1956

Theoretical investigation of flutter of two-dimensional flat panels with one surface exposed to supersonic potential flow

A Rayleigh type analysis involving chosen modes of the panel as degrees of freedom is used to treat the flutter of a two-dimensional flat panel supported at its leading and trailing edges and subjected to a middle-plane tensile force. The panel has a supersonic stream passing over its upper surface and still air below. The aerodynamic forces due to the supersonic stream are obtained from the theory for linearized two-dimensional unsteady flow and the forces due to the still air are obtained from acoustical theory. In order to study the effect of increasing the number of modes in the analysis, two and then four modes are employed. The modes used are the first four natural modes of the panel in a vacuum with no tensile force acting. The analysis includes these variables: Mach number, structural damping, tensile force, density of the still air, and edge fixity (clamped and pinned). For certain combinations of these variables, stability boundaries are obtained which can be used to determine the panel thickness required to prevent flutter for any panel material and altitude.

Nelson, Herbert C↗
NASA NTRS1994

Benchmark panels

Results are presented for an analytical and experimental study of curved, stiffened composite panels for a commercial transport aircraft. Five panels representative of fuselage crown design alternatives were fabricated for concept evaluation. These panels are scheduled to be tested in the Pressure Box Test Machine at NASA Langley Research Center. The panels are characterized by a 122 in. radius, tow-placed skins, hot-drape-formed hat-section stringers, and braided/RTM J-section frames. A geometrically nonlinear finite element analysis was conducted to support the development of test fixturing and to provide predictions of undamaged panel response. A strain softening model for damaged structures is described and used to predict the response of damaged panels. Results for a study of the stringer-frame intersection ('mousehole') design detail are also presented. Bonded frame pull-off tests and supporting analyses are described.

T H Walker↗
NASA NTRS1995

Active control of panel vibrations induced by a boundary layer flow

The problems of active and passive control of sound and vibration has been investigated by many researchers for a number of years. However, few of the articles are concerned with the sound and vibration with flow-structure interaction. Experimental and numerical studies on the coupling between panel vibration and acoustic radiation due to flow excitation have been done by Maestrello and his associates at NASA/Langley Research Center. Since the coupled system of nonlinear partial differential equations is formidable, an analytical solution to the full problem seems impossible. For this reason, we have to simplify the problem to that of the nonlinear panel vibration induced by a uniform flow or a boundary-layer flow with a given wall pressure distribution. Based on this simplified model, we have been able to consider the control and stabilization of the nonlinear panel vibration, which have not been treated satisfactorily by other authors. Although the sound radiation has not been included, the vibration suppression will clearly reduce the sound radiation power from the panel. The major research findings are presented in three sections. In section two we describe results on the boundary control of nonlinear panel vibration, with or without flow excitation. Sections three and four are concerned with some analytical and numerical results in the optimal control of the linear and nonlinear panel vibrations, respectively, excited by the flow pressure fluctuations. Finally, in section five, we draw some conclusions from research findings.

Chow, Pao-Liu↗
NASA NTRS1995

Finite element buckling and postbuckling solutions for multilayered composite panels

A study is made of the buckling and postbuckling responses of flat, unstiffened composite panels subjected to various combinations of mechanical and thermal loads. The analysis is based on a first-order shear deformation von Karman-type plate theory. A mixed formulation is used with the fundamental unknowns consisting of the strain components, stress resultants and the generalized displacements of the plate. The stability boundary, postbuckling response and the sensitivity coefficients are evaluated. The sensitivity coefficients measure the sensitivity of the buckling and postbuckling responses to variations in the different lamination and material parameters of the panel. Numerical results are presented for both solid panels and panels with central circular cutouts. The results show the effects of the variations in the fiber orientation angels, aspect ratio of the panel, and the hole diameter (for panels with cutouts) on the stability boundary, postbuckling response and sensitivity coefficients.

Noor, Ahmed K.↗