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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 811 records · Page 45
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Advances in Structural Integrity Analysis Methods for Aging Metallic Airframe Structures with Local Damage

Analysis methodologies for predicting fatigue-crack growth from rivet holes in panels subjected to cyclic loads and for predicting the residual strength of aluminum fuselage structures with cracks and subjected to combined internal pressure and mechanical loads are described. The fatigue-crack growth analysis methodology is based on small-crack theory and a plasticity induced crack-closure model, and the effect of a corrosive environment on crack-growth rate is included. The residual strength analysis methodology is based on the critical crack-tip-opening-angle fracture criterion that characterizes the fracture behavior of a material of interest, and a geometric and material nonlinear finite element shell analysis code that performs the structural analysis of the fuselage structure of interest. The methodologies have been verified experimentally for structures ranging from laboratory coupons to full-scale structural components. Analytical and experimental results based on these methodologies are described and compared for laboratory coupons and flat panels, small-scale pressurized shells, and full-scale curved stiffened panels. The residual strength analysis methodology is sufficiently general to include the effects of multiple-site damage on structural behavior.

fatigue-crack growth↗
NASA NTRS1998

Fracture Analysis of the FAA/NASA Wide Stiffened Panels

This paper presents the fracture analyses conducted on the FAA/NASA stiffened and unstiffened panels using the STAGS (STructural Analysis of General Shells) code with the critical crack-tip-opening angle (CTOA) fracture criterion. The STAGS code with the "plane-strain" core option was used in all analyses. Previous analyses of wide, flat panels have shown that the high-constraint conditions around a crack front, like plane strain, has to be modeled in order for the critical CTOA fracture criterion to predict wide panel failures from small laboratory tests. In the present study, the critical CTOA value was determined from a wide (unstiffened) panel with anti-buckling guides. The plane-strain core size was estimated from previous fracture analyses and was equal to about the sheet thickness. Rivet flexibility and stiffener failure was based on methods and criteria, like that currently used in industry. STAGS and the CTOA criterion were used to predict load-against-crack extension for the wide panels with a single crack and multiple-site damage cracking at many adjacent rivet holes. Analyses were able to predict stable crack growth and residual strength with a few percent (5%) of stiffened panel tests results but over predicted the buckling failure load on a unstiffened panel with a single crack by 10%.

Seshadri, B. R.↗
NASA NTRS1999

Fracture Analysis of the FAA/NASA Wide Stiffened Panels

This paper presents the fracture analyses conducted on the FAA/NASA stiffened and unstiffened panels using the STAGS (STructural Analysis of General Shells) code with the critical crack-tip-opening angle (CTOA) fracture criterion. The STAGS code with the "plane-strain" core option was used in all analyses. Previous analyses of wide, flat panels have shown that the high-constraint conditions around a crack front, like plane strain, has to be modeled in order for the critical CTOA fracture criterion to predict wide panel failures from small laboratory tests. In the present study, the critical CTOA value was determined from a wide (unstiffened) panel with anti-buckling guides. The plane-strain core size was estimated from previous fracture analyses and was equal to about the sheet thickness. Rivet flexibility and stiffener failure was based on methods and criteria, like that currently used in industry. STAGS and the CTOA criterion were used to predict load-against-crack extension for the wide panels with a single crack and multiple-site damage cracking at many adjacent rivet holes. Analyses were able to predict stable crack growth and residual strength within a few percent (5%) of stiffened panel tests results but over predicted the buckling failure load on an unstiffened panel with a single crack by 10%.

Seshadri, B. R.↗
NASA NTRS1975

Advanced beaded and tubular structural panels

A program to develop lightweight beaded and tubular structural panels is described. Applications include external surfaces, where aerodynamically acceptable, and primary structure protected by heat shields. The design configurations were optimized and selected with a computer code which iterates geometric parameters to satisfy strength, stability, and weight constraints. Methods of fabricating these configurations are discussed. Nondestructive testing produced extensive combined compression, shear, and bending test data on local buckling specimens and large panels. The optimized design concepts offer 25 to 30% weight savings compared to conventional stiffened sheet construction.

Musgrove, M. D.↗
NASA NTRS1991

A new fabrication method for precision antenna reflectors for space flight and ground test

Communications satellites are using increasingly higher frequencies that require increasingly precise antenna reflectors for use in space. Traditional industry fabrication methods for space antenna reflectors employ successive modeling techniques using high- and low-temperature molds for reflector face sheets and then a final fit-up of the completed honeycomb sandwich panel antenna reflector to a master pattern. However, as new missions are planned at much higher frequencies, greater accuracies will be necessary than are achievable using these present methods. A new approach for the fabrication of ground-test solid-surface antenna reflectors is to build a rigid support structure with an easy-to-machine surface. This surface is subsequently machined to the desired reflector contour and coated with a radio-frequency-reflective surface. This method was used to fabricate a 2.7-m-diameter ground-test antenna reflector to an accuracy of better than 0.013 mm (0.0005 in.) rms. A similar reflector for use on spacecraft would be constructed in a similar manner but with space-qualified materials. The design, analysis, and fabrication of the 2.7-m-diameter precision antenna reflector for antenna ground tests and the extension of this technology to precision, space-based antenna reflectors are described.

Sharp, G. Richard↗
NASA NTRS1976

Theoretical and experimental pressure distributions for a 71.2 degree swept arrow-wing configuration at subsonic, transonic, and supersonic speeds

A wind-tunnel test of an arrow-wing body configuration consisting of flat and twisted wings, as well as a variety of leading- and trailing-edge control-surface deflections, was conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. Theory-to-experiment comparisons of detailed pressure distributions were made using current state-of-the-art and newly developed attached- and separated-flow methods. Conditions were delineated under which these theories provide accurate basic and incremental aeroelastic loads predictions. Current state-of-the-art linear and nonlinear attached-flow methods were adequate only at small-angle-of-attack cruise conditions. Of the several separated-vortex methods evaluated, only the one utilizing a combination of linear source and quadratically varying doublet panels showed promise of yielding accurate loads distributions at moderate to large angles of attack.

Bobbitt, P. J.↗
NASA NTRS1978

Three-dimensional canard-wing shape optimization in aircraft cruise and maneuver environments

This paper demonstrates a numerical technique for canard-wing shape optimization at two operating conditions. For purposes of simplicity, a mean surface wing paneling code is employed for the aerodynamic calculations. The optimization procedures are based on the method of feasible directions. The shape functions for describing the thickness, camber, and twist are based on polynomial representations. The primary design requirements imposed restrictions on the canard and wing volumes and on the lift coefficients at the operating conditions. Results indicate that significant improvements in minimum drag and lift-to-drag ratio are possible with reasonable aircraft geometries. Calculations were done for supersonic speeds with Mach numbers ranging from 1 to 6. Planforms were mainly of a delta shape with aspect ratio of 1.

De Silva, B. M. E.↗
NASA NTRS1986

Optimization of outgassing bake-out temperatures and duration of space systems

Satellite components, e.g., solar panels, were subjected to bake-out tests to model outgassing characteristics and to identify methods for minimizing the bake-out time. Bake-out removes gases and particulates that are trapped at the surfaces of components during manufacture, storage, transit and handling. The tests covered temperatures from 45-85 C for up to 200 hr. Measurements of the mass loss rates were used to model the process as a function of the temperature, time and mass. Data sampling with a quartz crystal microbalance was found to be an effective means for ascertaining the mass loss rate and the activation energy of the release. The tests showed that the duration of the bake-out depends on the amount of material which must be removed and the type of outgassing which occurs. Materials which are resistant to high temperatures can be baked at the highest feasible temperatures to accelerate the outgassing process.

John Scialdone↗
NASA NTRS1989

Investigation of a free-tip rotor configuration for research on spanwise life distributions and wake velocity surveys of a semi-span wing with a discontinuous twist

A wind tunnel test was conducted in the NASA Ames 7 x 10 Foot Wind Tunnel to investigate the lift distribution on a semi-span wing with a discontinuous change in spanwise twist. The semi-span wing had a tip with an adjustable pitch angle independent on the inboard section pitch angle simulating the free-tip rotor blade when its free-tip is at a deflected position. The spanwise lift distribution over the wing and the tip were measured and three component velocity surveys behind the wing were obtained with a three dimensional laser Doppler velocimeter (LV) with the wing at one angle of attack and the tip deflected at different pitch angles. A six component internal strain gage balance was also used to measure total forces and moments on the tip. The three dimensional lift was computed from the two dimensional life distributions obtained from the LV and from the strain gage balance. The results from both experimental methods are shown to be in agreement with predictions made by a steady, three dimensional panel code, VSAERO.

Fortin, Paul↗
NASA NTRS1989

Spanwise lift distributions and wake velocity surveys of a semi-span wing with a discontinuous twist

A wind tunnel test was conducted in the NASA-Ames 7 x 10 ft wind tunnel to investigate the lift distribution on a semispan wing with a discontinuous change in spanwise twist. The semispan wing had a tip with an adjustable pitch angle independent on the inboard section pitch angle simulating the free tip rotor blade when its free tip is at a deflected position. The spanwise lift distribution over the wing and the tip were measured and three component velocity surveys behind the wing were obtained with a 3-D laser Doppler velocimeter (LV) with the wing at one angle of attack and the tip deflected at different pitch angles. A six-component internal strain gage balance was also used to measure total forces and moments on the tip. The 3-D lift was computed from the 2-D lift distributions obtained from the LV and from the strain gage balance. The results from both experimental methods are shown to be in agreement with predictions made by a steady, 3-D panel code, VSAERO.

Kumagai, Hiroyuki↗
NASA NTRS1989

A shear deformable theory of laminated composite shallow shell-type panels and their response analysis. I - Free vibration and buckling

This paper deals with the substantiation of a shear deformable theory of cross-ply laminated composite shallow shells. While the developed theory preserves all the advantages of the first order transverse shear deformation theory it succeeds in eliminating some of its basic shortcomings. The theory is further employed in the analysis of the eigenvibration and static buckling problems of doubly curved shallow panels. In this context, the state space concept is used in conjunction with the Levy method, allowing one to analyze these problems in a unified manner, for a variety of boundary conditions. Numerical results are presented and some pertinent conclusions are formulated.

Librescu, L.↗
NASA NTRS1990

Buckling analysis and test correlation of hat stiffened panels for hypersonic vehicles

The paper discusses the design, analysis, and test of hat stiffened panels subjected to a variety of thermal and mechanical load conditions. The panels were designed using data from structural optimization computer codes and finite element analysis. Test methods included the grid shadow moire method and a single gage force stiffness method. The agreement between the test data and analysis provides confidence in the methods that are currently being used to design structures for hypersonic vehicles. The agreement also indicates that post buckled strength may potentially be used to reduce the vehicle weight.

Percy, Wendy C.↗
NASA NTRS2000

Flexible Wing Model for Structural Sizing and Multidisciplinary Design Optimization of a Strut-Braced Wing

This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.

Gern, Frank H.↗
NASA NTRS2004

Subscale Testing of a Ceramic Composite Cooled Panel Led to Its Design and Fabrication for Scramjet Engine Testing

In a partnership between the NASA Glenn Research Center and Pratt & Whitney, a ceramic heat exchanger panel intended for use along the hot-flow-path walls of future reusable launch vehicles was designed, fabricated, and tested. These regeneratively cooled ceramic matrix composite (CMC) panels offer lighter weight, higher operating temperatures, and reduced coolant requirements in comparison to their more traditional metallic counterparts. A maintainable approach to the design was adopted which allowed the panel components to be assembled with high-temperature fasteners rather than by permanent bonding methods. With this approach, the CMC hot face sheet, the coolant containment system, and backside structure were all fabricated separately and could be replaced individually as the need occurred during use. This maintainable design leads to both ease of fabrication and reduced cost.

Jaskowiak, Martha H.↗
NASA NTRS2012

4D Light Field Imaging System Using Programmable Aperture

Complete depth information can be extracted from analyzing all angles of light rays emanated from a source. However, this angular information is lost in a typical 2D imaging system. In order to record this information, a standard stereo imaging system uses two cameras to obtain information from two view angles. Sometimes, more cameras are used to obtain information from more angles. However, a 4D light field imaging technique can achieve this multiple-camera effect through a single-lens camera. Two methods are available for this: one using a microlens array, and the other using a moving aperture. The moving-aperture method can obtain more complete stereo information. The existing literature suggests a modified liquid crystal panel [LC (liquid crystal) panel, similar to ones commonly used in the display industry] to achieve a moving aperture. However, LC panels cannot withstand harsh environments and are not qualified for spaceflight. In this regard, different hardware is proposed for the moving aperture. A digital micromirror device (DMD) will replace the liquid crystal. This will be qualified for harsh environments for the 4D light field imaging. This will enable an imager to record near-complete stereo information. The approach to building a proof-ofconcept is using existing, or slightly modified, off-the-shelf components. An SLR (single-lens reflex) lens system, which typically has a large aperture for fast imaging, will be modified. The lens system will be arranged so that DMD can be integrated. The shape of aperture will be programmed for single-viewpoint imaging, multiple-viewpoint imaging, and coded aperture imaging. The novelty lies in using a DMD instead of a LC panel to move the apertures for 4D light field imaging. The DMD uses reflecting mirrors, so any light transmission lost (which would be expected from the LC panel) will be minimal. Also, the MEMS-based DMD can withstand higher temperature and pressure fluctuation than a LC panel can. Robotics need near complete stereo images for their autonomous navigation, manipulation, and depth approximation. The imaging system can provide visual feedback

Bae, Youngsam↗
NASA NTRS1997

Design Criteria for X-CRV Honeycomb Panels: A Preliminary Study

The objective of this project is to perform the first step in developing structural design criteria for composite sandwich panels that are to be used in the aeroshell of the crew return vehicle (X-CRV). The preliminary concept includes a simplified method for assessing the allowable strength in the laminate material. Ultimately, it is intended that the design criteria be extended to address the global response of the vehicle. This task will require execution of a test program as outlined in the recommendation section of this report. The aeroshell of the X-CRV is comprised of composite sandwich panels consisting of fiberite face sheets and a phenolic honeycomb core. The function of the crew return vehicle is to enable the safe return of injured or ill crewpersons from space station, the evacuation of crew in case of emergency or the return of crew if an orbiter is not available. A significant objective of the X-CRV project is to demonstrate that this vehicle can be designed, built and operated at lower cost and at a significantly faster development time. Development time can be reduced by driving out issues in both structural design and manufacturing concurrently. This means that structural design and analysis progresses in conjunction with manufacturing and testing. Preliminary tests results on laminate coupons are presented in the report. Based on these results a method for detection material failure in the material is presented. In the long term, extrapolation of coupon data to large scale structures may be inadequate. Test coupons used to develop failure criteria at the material scale are typically small when compared to the overall structure. Their inherent small size indicates that the material failure criteria can be used to predict localized failure of the structure, however, it can not be used to predict failure for all failure modes. Some failure modes occur only when the structure or one of its sub-components are studied as a whole. Conversely, localized failure may not indicate failure of the structure as a whole and the amount of reserve capacity, if any, should be assessed. To develop a complete design criteria experimental studies of the sandwich panel are needed. Only then can a conservative and accurate design criteria be developed. This criteria should include effects of flaws and defects, and environmental factors such as temperature and moisture. Preliminary results presented in this report suggest that a simplified analysis can be used to predict the strength of a laminate. Testing for environmental effects have yet to be included in this work. The so called 'rogue flaw test' appears to be a promising method for assessing the effect of a defect in a laminate. This method fits in quite well with the philosophy of achieving a damage tolerant design.

Caccese, Vincent↗
NASA NTRS1992

A comparative evaluation of in-plane shear test methods for laminated graphite-epoxy composites

The objectives were to evaluate popular shear test methods for various forms of graphite-epoxy composite materials and to determine the shear response of graphite-epoxy composites with various forms of fiber architecture. Numerical and full-field experimental stress analyses were performed on four shear test configurations for unidirectional and bidirectional graphite-epoxy laminates to assess the uniformity and purity of the shear stress (strain) fields produced in the specimen test section and to determine the material in-plane shear modulus and shear response. The test methods were the 10 deg off-axis, the +/- 45 deg tension, the Iosipescu V-notch, and a compact U-notch specimen. Specimens were prepared from AS4/3501-6 graphite-epoxy panels, instrumented with conventional strain gage rosettes and with a cross-line moire grating, and loaded in a convenient testing machine. The shear responses obtained for each test method and the two methods of specimen instrumentation were compared. In a second phase of the program the shear responses obtained from Iosipescu V-notch beam specimens were determined for woven fabric geometries of different weave and fiber architectures. Again the responses of specimens obtained from strain gage rosettes and moire interferometry were compared. Additional experiments were performed on a bidirectional cruciform specimen which was also instrumented with strain gages and a moire grating.

Morton, John↗
NASA NTRS1980

Loading tests of a wing structure for a hypersonic aircraft

Room-temperature loading tests were conducted on a wing structure designed with a beaded panel concept for a Mach 8 hypersonic research airplane. Strain, stress, and deflection data were compared with the results of three finite-element structural analysis computer programs and with design data. The test program data were used to evaluate the structural concept and the methods of analysis used in the design. A force stiffness technique was utilized in conjunction with load conditions which produced various combinations of panel shear and compression loading to determine the failure envelope of the buckling critical beaded panels The force-stiffness data did not result in any predictions of buckling failure. It was, therefore, concluded that the panels were conservatively designed as a result of design constraints and assumptions of panel eccentricities. The analysis programs calculated strains and stresses competently. Comparisons between calculated and measured structural deflections showed good agreement. The test program offered a positive demonstration of the beaded panel concept subjected to room-temperature load conditions.

Fields, R. A.↗