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At least 217 records · Page 12

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NASA NTRS1983

Study of noise reduction characteristics of double-wall panels

The noise reduction characteristics of general aviation type, flat, double-wall structures were investigated. The experimental study was carried out on 20-by-20 inch panels with an exposed area of 18 by 18 inches. A frequency range from 20 to 5000 Hz was covered. The experimental results, in general, follow the expected trends. At low frequencies the double-wall structures are no better than the single-wall structures. However, for depths normally used in the general aviation industry, the double-wall panels are very attractive. The graphite-spoxy skin panels have higher noise reduction at very low frequencies ( 100 Hz) than the Kevlar skin panels. But the aluminum panels have higher noise reduction in the high frequency region, due to their greater mass. Use of fiberglass insulation is not effective in the low frequency region, and at times it is even negative. But the insulation is effective in the high-frequency region. The theoretical model for predicting the transmission loss of these multilayered panels is also discussed.

Navaneethan, R.↗

NASA NTRS1980

Fabrication and evaluation of brazed titanium-clad borsic/aluminum skin-stringer panels

A successful brazing process was developed and evaluated for fabricating full-scale titanium-clad Borsic/aluminum skin-stringer panels. A panel design was developed consisting of a hybrid composite skin reinforced with capped honeycomb-core stringers. Six panels were fabricated for inclusion in the program which included laboratory testing of panels at ambient temperatures and 533 K (500 F) and flight service evaluation on the NASA Mach 3 YF-12 airplane. All panels tested met or exceeded stringent design requirements and no deleterious effects on panel properties were detected followng flight service evaluation on the YF-12 airplane.

Bales, T. T.↗

NASA NTRS1985

Glass/Epoxy Door Panel for Automobiles

Lightweight panel cost-effective. Integrally-molded intrusion strap key feature of composite outer door panel. Strap replaces bulky and heavy steel instrusion beam of conventional door. Standard steel inner panel used for demonstration purposes. Door redesigned to exploit advantages of composite outer panel thinner. Outer panel for automobilie door, made of glass/epoxy composite material, lighter than conventional steel door panel, meets same strength requirements, and less expensive.

Bauer, J. L. JR.↗

NASA NTRS1984

POSTOP: Postbuckled open-stiffener optimum panels-theory and capability

The computer program POSTOP was developed to serve as an aid in the analysis and sizing of stiffened composite panels that are loaded in the postbuckling regime. A comprehensive set of analysis routines was coupled to a widely used optimization program to produce this sizing code. POSTOP is intended for the preliminary design of metal or composite panels with open-section stiffeners, subjected to multiple combined biaxial compression (or tension), shear and normal pressure load cases. Longitudinal compression, however, is assumed to be the dominant loading. Temperature, initial bow eccentricity and load eccentricity effects are included. The panel geometry is assumed to be repetitive over several bays in the longitudinal (stiffener) direction as well as in the transverse direction. Analytical routines are included to compute panel stiffnesses, strains, local and panel buckling loads, and skin/stiffener interface stresses. The resulting program is applicable to stiffened panels as commonly used in fuselage, wing, or empennage structures. The analysis procedures and rationale for the assumptions used therein are described in detail.

Dickson, J. N.↗

NASA NTRS1984

POSTOP: Postbuckled open-stiffener optimum panels, user's manual

The computer program POSTOP developed to serve as an aid in the analysis and sizing of stiffened composite panels that may be loaded in the postbuckling regime, is intended for the preliminary design of metal or composite panels with open-section stiffeners, subjected to multiple combined biaxial compression (or tension), shear and normal pressure load cases. Longitudinal compression, however, is assumed to be the dominant loading. Temperature, initial bow eccentricity and load eccentricity effects are included. The panel geometry is assumed to be repetitive over several bays in the longitudinal (stiffener) direction as well as in the transverse direction. Analytical routines are included to compute panel stiffnesses, strains, local and panel buckling loads, and skin/stiffener interface stresses. The resulting program is applicable to stiffened panels as commonly used in fuselage, wing, or empennage structures. The capabilities and limitations of the code are described. Instructions required to use the program and several example problems are included.

Biggers, S. B.↗

NASA NTRS1984

Postbuckling behavior of axially compressed graphite-epoxy cylindrical panels with circular holes

The results of an experimental and analytical study of the effects of circular holes on the postbuckling behavior of graphite-epoxy cylindrical panels loaded in axial compression are presented. The STAGSC-1 general shell analysis computer code is used to determine the buckling and postbuckling response of the panels. The loaded, curved ends of the specimens were clamped by fixtures and the unloaded, straight edges were simply supported by knife-edge restraints. The panels are loaded by uniform end shortening to several times the end shortening at buckling. The unstable equilibrium path of the postbuckling response is obtained analytically by using a method based on controlling an equilibrium-path-arc-length parameter instead of the traditional load parameter. The effects of hole diameter, panel radius, and panel thickness on postbuckling response are considered in the study. Experimental results are compared with the analytical results and the failure characteristics of the graphite-epoxy panels are described.

Knight, N. F., Jr.↗

NASA NTRS1985

Space shuttle elevon seal panel mechanism

The orbiter elevon seal panel mechanism controls the position of fairing panels between the orbiter wing and elevon. Early mechanism designs used linkages which approximately matched the panel motion to elevon position, depending on panel deflections to maintain sealing. These linkages were refined during orbiter development to match panel motion to elevon motion more exactly, thus reducing panel deflections, loads, and weight. Changes to the adjacent cove seal resulted in the use of curved tension compression links. Mechanism temperatures up to 750 F (locally) posed difficulties in bearing lubrication. Despite the adverse effect of the many fabrication tolerances, the system successfully prevented the entry of 1200 F hot gases into the wing/elevon joint.

Ripley, J. G.↗

NASA NTRS1985

Postbuckling behavior of selected curved stiffened graphite-epoxy panels loaded in axial compression

Results of an experimental and analytical study of the postbuckling behavior of selected curved stiffened graphite-epoxy panels loaded in axial compression are presented. The postbuckling response and failure characteristics of the panels are described. Each panel had four equally-spaced I-shaped stiffeners and 16-ply quasi-isotropic skins. Panels with three different stiffener spacings were tested. Failure of all panels initiated in a skin-stiffener interface region. Analytical results from a nonlinear general shell finite element analysis computer code correlate well with typical postbuckling test results up to failure. The analytical modeling detail necessary to predict accurately the response of the panel is described. Measured initial geometric imperfections were included in the postbuckling analysis.

Knight, N. F., Jr.↗

NASA NTRS1985

Application of the finite element method in the calculation of transmission loss of flat and curved panels

This investigation represents an extension of a study of Roussos (1985) who considered the noise transmission loss of a rectangular plate in an infinite baffle. Roussos, who employed an analytical formulation, considered an unstiffened plate. While it is difficult to consider stiffeners by means of analytical methods, the difficulties can be avoided by employing a finite element procedure. For this reason, the present study is concerned with the implementation of a finite element method. The representation of the panel transmission loss is discussed, and the determination of the panel motion by means of the finite element technique is described, taking into account an isotropic flat panel, the exciting force, an eigenvalue problem, the radiation pressure, a plate element, and a cylindrical shell element. Numerical results are considered for a flat panel, a curved panel, and a stiffened flat panel.

Koval, L. R.↗

NASA NTRS1987

Analysis and test of superplastically formed titanium hat-stiffened panels under compression

Four hat-stiffened titanium panels with two different stiffener configurations were fabricated by superplastic forming/weld brazing and tested under a moderately heavy compressive load. The panels had the same overall dimensions but differed in the shape of the hat-stiffener webs; three panels had stiffeners with flat webs and the other panel had stiffeners with beaded webs. Analysis indicated that the local buckling strain of the flat stiffener web was considerably lower than the general panel buckling strain or cap buckling strain. The analysis also showed that beading the webs of the hat stiffeners removed them as the critical element for local buckling and improved the buckling strain of the panels. The analytical extensional stiffness and failure loads compared very well with experimental results.

Davis, Randall C.↗

NASA NTRS1985

Damage tolerant composite wing panels for transport aircraft

Commercial aircraft advanced composite wing surface panels were tested for durability and damage tolerance. The wing of a fuel-efficient, 200-passenger airplane for 1990 delivery was sized using grahite-epoxy materials. The damage tolerance program was structured to allow a systematic progression from material evaluations to the optimized large panel verification tests. The program included coupon testing to evaluate toughened material systems, static and fatigue tests of compression coupons with varying amounts of impact damage, element tests of three-stiffener panels to evaluate upper wing panel design concepts, and the wing structure damage environment was studied. A series of technology demonstration tests of large compression panels is performed. A repair investigation is included in the final large panel test.

Smith, Peter J.↗

NASA NTRS1989

Design and testing of thermal-expansion-molded graphite-epoxy hat-stiffened sandwich panels

Minimum weight configurations for two types of graphite-epoxy hat-stiffened compression-loaded panels fabricated by the thermal-expansion-molding (TEM) manufacturing process were evaluated analytically and experimentally for designs with load index Nx/L values ranging from 100 to 800. The two types of panels contain graphite-epoxy face sheets with a foam core and hat stiffeners which are either open or filled with foam. Constraints on the extensional and shear stiffnesses are imposed on the design so that the panels will satisfy typical constraints for aircraft wing structures. Optimal structurally efficient TEM panels are compared to commercially available aluminum aircraft structures. Predicted load-strain relationships agree well with experimental results. Significant impact damage to the unstiffened face sheet and foam core does not noticeably reduce the load carrying ability of the panels, but damage to the stiffened face sheet reduces the failure load by 20 percent compared to unimpacted panels.

Jegley, Dawn C.↗

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.↗