SEARCH · Engineering Papers

Results for “ultimate loads”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 91 records · Page 5

All sources NASA NTRS DOE OSTI

NASA NTRS1990

Damage states in laminated composite three-point bend specimens: An experimental-analytical correlation study

Damage states in laminated composites were studied by considering the model problem of a laminated beam subjected to three-point bending. A combination of experimental and theoretical research techniques was used to correlate the experimental results with the analytical stress distributions. The analytical solution procedure was based on the stress formulation approach of the mathematical theory of elasticity. The solution procedure is capable of calculating the ply-level stresses and beam displacements for any laminated beam of finite length using the generalized plane deformation or plane stress state assumption. Prior to conducting the experimental phase, the results from preliminary analyses were examined. Significant effects in the ply-level stress distributions were seen depending on the fiber orientation, aspect ratio, and whether or not a grouped or interspersed stacking sequence was used. The experimental investigation was conducted to determine the different damage modes in laminated three-point bend specimens. The test matrix consisted of three-point bend specimens of 0 deg unidirectional, cross-ply, and quasi-isotropic stacking sequences. The dependence of the damage initiation loads and ultimate failure loads were studied, and their relation to damage susceptibility and damage tolerance of the mean configuration was discussed. Damage modes were identified by visual inspection of the damaged specimens using an optical microscope. The four fundamental damage mechanisms identified were delaminations, matrix cracking, fiber breakage, and crushing. The correlation study between the experimental results and the analytical results were performed for the midspan deflection, indentation, damage modes, and damage susceptibility.

Starbuck, J. Michael↗

NASA NTRS1989

Assessment of structural analysis technology for elastic shell collapse problems

The prediction of the ultimate load carrying capability for compressively loaded shell structures is a challenging nonlinear analysis problem. Selected areas of finite element technology research and nonlinear solution technology are assessed. Herein, a finite element analysis procedure is applied to four shell collapse problems which have been used by computational structural mechanics researchers in the past. This assessment will focus on a number of different shell element formulations and on different approaches used to account for geometric nonlinearities. The results presented confirm that these aspects of nonlinear shell analysis can have a significant effect on the predicted nonlinear structural response. All analyses were performed using the CSM Testbed software system which allowed a convenient assessment of different element formulations with a consistent approach to solving the discretized nonlinear equations.

Knight, N. F., Jr.↗

NASA NTRS1991

Detailed analysis and test correlation of a stiffened composite wing panel

Nonlinear finite element analysis techniques are evaluated by applying them to a realistic aircraft structural component. A wing panel from the V-22 tiltrotor aircraft is chosen because it is a typical modern aircraft structural component for which there is experimental data for comparison of results. From blueprints and drawings supplied by the Bell Helicopter Textron Corporation, a very detailed finite element model containing 2284 9-node Assumed Natural-Coordinate Strain (ANS) elements was generated. A novel solution strategy which accounts for geometric nonlinearity through the use of corotating element reference frames and nonlinear strain displacements relations is used to analyze this detailed model. Results from linear analyses using the same finite element model are presented in order to illustrate the advantages and costs of the nonlinear analysis as compared with the more traditional linear analysis. Strain predictions from both the linear and nonlinear stress analyses are shown to compare well with experimental data up through the Design Ultimate Load (DUL) of the panel. However, due to the extreme nonlinear response of the panel, the linear analysis was not accurate at loads above the DUL. The nonlinear analysis more accurately predicted the strain at high values of applied load, and even predicted complicated nonlinear response characteristics, such as load reversals, at the observed failure load of the test panel. In order to understand the failure mechanism of the panel, buckling and first ply failure analyses were performed. The buckling load was 17 percent above the observed failure load while first ply failure analyses indicated significant material damage at and below the observed failure load.

Davis, D. Dale, Jr.↗

NASA NTRS1943

Tests of Aluminum-alloy Stiffened-sheet Specimens Cut from an Airplane Wing

The specimens used in the present tests were cut from an actual airplane wing of the stressed-skin type. The specimens thus obtained were not representative of the usual type of laboratory specimens because the stiffeners were not exactly parallel nor evenly spaced and, in one case, the skin consisted of pieces of sheet of different thicknesses. The test data obtained indicate that the buckling strain of stiffened curved sheet can be computed with reasonable accuracy by the equation given by Wenzek. The ultimate loads of the specimens when tested as flat sheet were within +/-11 percent of the product of the compressive yield strength and the cross-sectional area of the stiffeners. A rivet spacing equal to 98 times the sheet thickness was a source of weakness, and rivet spacings up to 36 times the sheet thickness appeared satisfactory.

Holt, Marshall↗

NASA NTRS1993

Analysis and test of low profile aluminum aerospace tank dome

In order to increase the structural performance of cryogenic tanks, the aerospace industry is beginning to employ low-profile bulkheads in new generation launch vehicle designs. This report details the analysis and test of one such dome made from 2219 aluminum. Such domes have two potential failure modes under internal pressure, general tensile failure and hoop compression buckling (in regions near the equator). The test determined the buckling load and ultimate tensile load of the hardware and showed that both compared well with the analysis predictions. This effort was conducted under the auspices of NASA and the General Dynamics Cryogenic Tank Technology Program (CTTP).

Ahmed, R.↗

NASA NTRS1994

Impact force as a scaling parameter

The Federal Aviation Administration (FAR PART 25) requires that a structure carry ultimate load with nonvisible impact damage and carry 70 percent of limit flight loads with discrete damage. The Air Force has similar criteria (MIL-STD-1530A). Both civilian and military structures are designed by a building block approach. First, critical areas of the structure are determined, and potential failure modes are identified. Then, a series of representative specimens are tested that will fail in those modes. The series begins with tests of simple coupons, progresses through larger and more complex subcomponents, and ends with a test on a full-scale component, hence the term 'building block.' In order to minimize testing, analytical models are needed to scale impact damage and residual strength from the simple coupons to the full-scale component. Using experiments and analysis, the present paper illustrates that impact damage can be better understood and scaled using impact force than just kinetic energy. The plate parameters considered are size and thickness, boundary conditions, and material, and the impact parameters are mass, shape, and velocity.

Poe, Clarence C., Jr.↗

NASA NTRS1997

Advanced Technology Composite Fuselage - Repair and Damage Assessment Supporting Maintenance

Under the NASA-sponsored contracts for Advanced Technology Composite Aircraft Structures (ATCAS) and Materials Development Omnibus Contract (MDOC), Boeing is studying the technologies associated with the application of composite materials to commercial transport fuselage structure. Included in the study is the incorporation of maintainability and repairability requirements of composite primary structure into the design. This contractor report describes activities performed to address maintenance issues in composite fuselage applications. A key aspect of the study was the development of a maintenance philosophy which included consideration of maintenance issues early in the design cycle, multiple repair options, and airline participation in design trades. Fuselage design evaluations considered trade-offs between structural weight, damage resistance/tolerance (repair frequency), and inspection burdens. Analysis methods were developed to assess structural residual strength in the presence of damage, and to evaluate repair design concepts. Repair designs were created with a focus on mechanically fastened concepts for skin/stringer structure and bonded concepts for sandwich structure. Both a large crown (skintstringer) and keel (sandwich) panel were repaired. A compression test of the keel panel indicated the demonstrated repairs recovered ultimate load capability. In conjunction with the design and manufacturing developments, inspection methods were investigated for their potential to evaluate damaged structure and verify the integrity of completed repairs.

Flynn, B. W.↗

NASA NTRS1997

Biochemical Assessment of Stress in Cardiac Tissue in Response to Weightless Space Travel

The absence of unit gravity may cause physiological changes in the cardiovascular system. For instance, in the absence of Earth's gravity, venous return to the heart may increase due, in pan, to decreased pooling of the blood in the extremities. We hypothesize that this would produce an increase in the heart's work load ultimately resulting in hypertrophy.

Brunton, Laurence L.↗

NASA NTRS2002

Site- and compartment-specific changes in bone with hindlimb unloading in mature adult rats

The purpose of this study was to examine site- and compartment-specific changes in bone induced by hindlimb unloading (HU) in the mature adult male rat (6 months old). Tibiae, femora, and humeri were removed after 14, 21, and 28 days of HU for determination of bone mineral density (BMD) and geometry by peripheral quantitative computed tomography (pQCT), mechanical properties, and bone formation rate (BFR), and compared with baseline (0 day) and aging (28 day) controls. HU resulted in 20%-21% declines in cancellous BMD at the proximal tibia and femoral neck after 28 day HU vs. 0 day controls (CON). Cortical shell BMD at these sites was greater (by 4%-6%) in both 28 day HU and 28 day CON vs. 0 day CON animals, and nearly identical to that gain seen in the weight-bearing humerus. Mechanical properties at the proximal tibia exhibited a nonsignificant decline after HU vs. those of 0 day CON rats. At the femoral neck, a 10% decrement was noted in ultimate load in 28 day HU rats vs. 28 day CON animals. Middiaphyseal tibial bone increased slightly in density and area during HU; no differences in structural and material properties between 28 day HU and 28 day CON rats were noted. BFR at the tibial midshaft was significantly lower (by 90%) after 21 day HU vs. 0 day CON; this decline was maintained throughout 28 day HU. These results suggest there are compartment-specific differences in the mature adult skeletal response to hindlimb unloading, and that the major impact over 28 days of unloading is on cancellous bone sites. Given the sharp decline in BFR for midshaft cortical bone, it appears likely that deficits in BMD, area, or mechanical properties would develop with longer duration unloading.

Non-NASA Center↗

NASA NTRS1998

Advanced Electric Distribution, Switching, and Conversion Technology for Power Control

The Electrical Power Control Unit currently under development by Sundstrand Aerospace for use on the Fluids Combustion Facility of the International Space Station is the precursor of modular power distribution and conversion concepts for future spacecraft and aircraft applications. This unit combines modular current-limiting flexible remote power controllers and paralleled power converters into one package. Each unit includes three 1-kW, current-limiting power converter modules designed for a variable-ratio load sharing capability. The flexible remote power controllers can be used in parallel to match load requirements and can be programmed for an initial ON or OFF state on powerup. The unit contains an integral cold plate. The modularity and hybridization of the Electrical Power Control Unit sets the course for future spacecraft electrical power systems, both large and small. In such systems, the basic hybridized converter and flexible remote power controller building blocks could be configured to match power distribution and conversion capabilities to load requirements. In addition, the flexible remote power controllers could be configured in assemblies to feed multiple individual loads and could be used in parallel to meet the specific current requirements of each of those loads. Ultimately, the Electrical Power Control Unit design concept could evolve to a common switch module hybrid, or family of hybrids, for both converter and switchgear applications. By assembling hybrids of a common current rating and voltage class in parallel, researchers could readily adapt these units for multiple applications. The Electrical Power Control Unit concept has the potential to be scaled to larger and smaller ratings for both small and large spacecraft and for aircraft where high-power density, remote power controllers or power converters are required and a common replacement part is desired for multiples of a base current rating.

Soltis, James V.↗

NASA NTRS2011

Structural Efficiency and Behavior of Pristine and Notched Stitched Structure

Two driving factors in aircraft panel design are structural efficiency and response to in-service damage. Stitching through the thickness can improve both of these considerations. Combining stitching with a post-buckling design approach can provide additional improvements. The buckling behavior of stitched structure is considered since lighter structures can be achieved if local skin buckling is allowed to occur at less than design ultimate load. Through-the-thickness stitching can suppress delamination between skin and flange, thereby allowing the structure to reliably carry load into the postbuckling range. Hat-stiffened and rod-stiffened panels in which the skin and flanges were stitched together through-the-thickness prior to curing are considered through experiment and analysis. In both types of panels no mechanical fasteners were used for the assembly. Specimens were loaded to failure in axial compression. In this study all specimens buckled in the skin between the stiffeners and continued to carry load. In addition, the behavior of panels with a severed stringer or notch are considered. Failure loads and strain distributions in the notched panel are compared to those in the unnotched panel.

Jegley, Dawn C.↗

NASA NTRS2015

Vibroacoustic Tailoring of a Rod-Stiffened Composite Fuselage Panel with Multidisciplinary Considerations

An efficient multi-objective design tailoring procedure seeking to improve the vibroacoustic performance of a fuselage panel while maintaining or reducing weight is presented. The structure considered is the pultruded rod stitched efficient unitized structure, a highly integrated composite structure concept designed for a noncylindrical, next-generation flight vehicle fuselage. Modifications to a baseline design are evaluated within a six-parameter design space including spacing, flange width, and web height for both frame and stringer substructure components. The change in sound power radiation attributed to a design change is predicted using finite-element models sized and meshed for analyses in the 500 Hz, 1 kHz, and 2 kHz octave bands. Three design studies are carried out in parallel while considering a diffuse acoustic field excitation and two types of turbulent boundary-layer excitation. Kriging surrogate models are used to reduce the computational costs of resolving the vibroacoustic and weight objective Pareto fronts. The resulting Pareto optimal designs are then evaluated under a static pressurization ultimate load to assess structural strength and stability. Results suggest that choosing alternative configurations within the considered design space can reduce weight and improve vibroacoustic performance without compromising strength and stability of the structure under the static load condition considered, but the tradeoffs are significantly influenced by the spatial characteristics of the assumed excitation field.

Allen, Albert R.↗

NASA NTRS2015

Nonlinear Analysis and Preliminary Testing Results of a Hybrid Wing Body Center Section Test Article

A large test article was recently designed, analyzed, fabricated, and successfully tested up to the representative design ultimate loads to demonstrate that stiffened composite panels with through-the-thickness reinforcement are a viable option for the next generation large transport category aircraft, including non-conventional configurations such as the hybrid wing body. This paper focuses on finite element analysis and test data correlation of the hybrid wing body center section test article under mechanical, pressure and combined load conditions. Good agreement between predictive nonlinear finite element analysis and test data is found. Results indicate that a geometrically nonlinear analysis is needed to accurately capture the behavior of the non-circular pressurized and highly-stressed structure when the design approach permits local buckling.

Przekop, Adam↗

NASA NTRSDate not supplied

FEA-Based Failure Analysis of Post-Buckled Composite Multi-Hat Stiffened Panel Under Static Compression

The aerospace industry uses composite multi-hat stiffened panels for skins because of their high strength-to-weight and stiffness-to-weight ratios. However, laminated composites with initial flaws are susceptible to delamination under buckling loads. The goal is to ensure damage initiation occurs past the buckled state, only stable damage progression at loads below the ultimate compressive load, and debonding failure past the ultimate compressive load. This approach reduces weight and addresses a potential design review for future aircraft repairs. The need exists to better understand damage initiation and growth mechanisms in composite structures while assisting in the design, analysis, and sustainment methods. Teflon inserts simulate damage defects during manufacturing, while barely visible impact helps capture human error during manufacturing. This work shows the interaction between the skin-stringer and ply-ply separations at the post-buckled state for both defect events using Abaqus Virtual Crack Closure Technique (VCCT). The initial configuration for the Teflon case was to leave the Teflon region unbonded, and the initial impact damage region approximation came from the experimental data and was considered the initial configuration. When compared against the experimental data produced through the NASA Advanced Composites Project (ACP), the present model captured damage transition from skin-hat to ply1-ply2. The model validations were within ~5% of the experimental data.

composite↗

DOE OSTI2024

Critical design load case fatigue and ultimate failure simulation for a 10-m H-type vertical-axis wind turbine

While previous studies investigating critical VAWT design load cases have focused on large and relatively flexible Darrieus designs, the bulk of current commercial products seeking certification fall in the relatively small, stiff, and H-type configuration, such as the XFlow Energy Corporation turbine that this study compares against. Understanding the critical design load case impacts for both fatigue and ultimate failure for this size and type of VAWT are imperative for certification. The abil

Brownstein, Ian↗

NASA NTRS2001

Effect of Load Rate on Ultimate Tensile Strength of Ceramic Matrix Composites at Elevated Temperatures

The strengths of three continuous fiber-reinforced ceramic composites, including SiC/CAS-II, SiC/MAS-5 and SiC/SiC, were determined as a function of test rate in air at 1100 to 1200 C. All three composite materials exhibited a strong dependency of strength on test rate, similar to the behavior observed in many advanced monolithic ceramics at elevated temperatures. The application of the preloading technique as well as the prediction of life from one loading configuration (constant stress-rate) to another (constant stress loading) suggested that the overall macroscopic failure mechanism of the composites would be the one governed by a power-law type of damage evolution/accumulation, analogous to slow crack growth commonly observed in advanced monolithic ceramics. It was further found that constant stress-rate testing could be used as an alternative to life prediction test methodology even for composite materials, at least for short range of lifetimes and when ultimate strength is used as the failure criterion.

Choi, Sung R.↗

NASA NTRS2014

Long Term Displacement Data of Woven Fabric Webbings Under Constant Load for Inflatable Structures

Inflatable modules for space applications offer weight and launch volume savings relative to current metallic modules. Limited data exist on the creep behavior of the restraint layer of inflatable modules. Long-term displacement and strain data of two high strength woven fabric webbings, Kevlar and Vectran, under constant load is presented. The creep behavior of webbings is required by designers to help determine service life parameters of inflatable modules. Four groups of different webbings with different loads were defined for this study. Group 1 consisted of 4K Kevlar webbings loaded to 33% ultimate tensile strength and 6K Vectran webbings loaded to 27% ultimate tensile strength, group 2 consisted of 6K Kevlar webbings loaded to 40% and 43% ultimate tensile strength, and 6K Vectran webbings loaded to 50% ultimate tensile strength, group 3 consisted of 6K Kevlar webbings loaded to 52% ultimate tensile strength and 6K Vectran webbings loaded to 60% ultimate tensile strength, and group 4 consisted of 12.5K Kevlar webbings loaded to 22% ultimate tensile strength, and 12.5K Vectran webbings loaded to 22% ultimate tensile strength. The uniquely designed test facility, hardware, displacement measuring devices, and test data are presented. Test data indicate that immediately after loading all webbings stretch an inch or more, however as time increases displacement values significantly decrease to fall within a range of several hundredth of an inch over the remainder of test period. Webbings in group 1 exhibit near constant displacements and strains over a 17-month period. Data acquisition was suspended after the 17th month, however webbings continue to sustain load without any local webbing damage as of the 21st month of testing. Webbings in group 2 exhibit a combination of initial constant displacement and subsequent increases in displacement rates over a 16-month period. Webbings in group 3 exhibit steady increases in displacement rates leading to webbing failure over a 3-month period. Five of six webbings experienced local damage and subsequent failure in group 3. Data from group 4 indicates increasing webbing displacements over a 7-month period. All webbings in groups 1, 2, and 4 remain suspended without any local damage as of the writing of this paper. Variations in facility temperatures over test period seem to have had limited effect on long-term webbing displacement data.

Kenner, Winfred S.↗

NASA NTRS1990

Structural behavior of composites with progressive fracture

Structural characteristics such as natural frequencies and buckling loads with corresponding mode shapes were investigated during progressive fracture of multilayer, angle-plied polymer matrix composites. A computer program was used to generate the numerical results for overall mechanical response of damaged composites. Variations in structural characteristics as a function of the previously applied loading were studied. Results indicate that most of the overall structural properties were preserved throughout a significant proportion of the ultimate fracture load. For the cases studied, changes in structural behavior began to occur after 70 percent of the ultimate fracture load was applied. However, the individual nature of the structural change was rather varied depending on the laminate configuration, fiber orientation, and the boundary conditions.

Minnetyan, Levon↗