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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

Unconservatism of Linear-Elastic Fracture Mechanics (LEFM) Analysis Post Autofrettage

The scope of the NASA Engineering and Safety Center (NESC) Unconservatism of Linear-Elastic Fracture Mechanics (LEFM) Analysis Post Autofrettage assessment was to evaluate the influence of the elastic-plastic autofrettage cycle on elastic cycles post autofrettage and generate data to support damage tolerance life analysis verification approach. The decision to include an elastic-plastic autofrettage cycle is informed by other factors beyond damage tolerance, including manufacturing and fatigue crack initiation. The results of the NESC assessment are contained in this report.

Linear-Elastic Fracture Mechanics↗
NASA NTRS1988

Review of fatigue and fracture research at NASA Langley Research Center

Most dynamic components in helicopters are designed with a safe-life constant-amplitude testing approach that has not changed in many years. In contrast, the fatigue methodology in other industries has advanced significantly in the last two decades. Recent research at the NASA Langley Research Center and the U.S. Army Aerostructures Directorate at Langley are reviewed relative to fatigue and fracture design methodology for metallic components. Most of the Langley research was directed towards the damage tolerance design approach, but some work was done that is applicable to the safe-life approach. In the areas of testing, damage tolerance concepts are concentrating on the small-crack effect in crack growth and measurement of crack opening stresses. Tests were conducted to determine the effects of a machining scratch on the fatigue life of a high strength steel. In the area of analysis, work was concentrated on developing a crack closure model that will predict fatigue life under spectrum loading for several different metal alloys including a high strength steel that is often used in the dynamic components of helicopters. Work is also continuing in developing a three-dimensional, finite-element stress analysis for cracked and uncracked isotropic and anisotropic structures. A numerical technique for solving simultaneous equations called the multigrid method is being pursued to enhance the solution schemes in both the finite-element analysis and the boundary element analysis. Finally, a fracture mechanics project involving an elastic-plastic finite element analysis of J-resistance curve is also being pursued.

Everett, Richard A., Jr.↗
NASA NTRS1991

Impact response of composite materials

Composite materials composed of carbon fibers and resin matrices offer great promise in reducing the weight of aerospace structures. However they remain extremely vulnerable to out of plane impact loads, which lead to severe losses in strength and stiffness. The results of an experimental program, undertaken to investigate the low velocity impact damage tolerance of composite materials is presented. The objectives were to identify key neat resin/composite properties that lead to enhancement of composite impact damage tolerance and to find a small scale test that predicts compression after impact properties of panels. Five materials were selected for evaluation. These systems represented different classes of material behavior such as brittle epoxy, modified epoxies, and amorphous and semicrystalling thermoplastics. The influence of fiber properties on the impact performance was also studied in one material, i.e., in polyether ether ketone (PEEK). Several 24 and 48 ply quasi-isotropic and 24 ply orthotropic laminates were examined using an instrumented drop weight impactor. Correlations with post impact compression behavior were made.

Tiwari, S. N.↗
NASA NTRS1978

Conference on Helicopter Structures Technology, Moffett Field, Calif., November 16-18, 1977, Proceedings

Work on advanced concepts for helicopter designs is reported. Emphasis is on use of advanced composites, damage-tolerant design, and load calculations. Topics covered include structural design flight maneuver loads using PDP-10 flight dynamics model, use of 3-D finite element analysis in design of helicopter mechanical components, damage-tolerant design of the YUH-61A main rotor system, survivability of helicopters to rotor blade ballistic damage, development of a multitubular spar composite main rotor blade, and a bearingless main rotor structural design approach using advanced composites.

Source record↗
NASA NTRS1994

An Analytical Tool for Fracture Control

The structural integrity of critical NASA space flight hardware is ensured through a damage tolerance program which is dictated in the fracture control requirements. The goal of a fracture control program is to ensure proper material selection and control, control of working stress levels, use of fracture resistant design concepts, manufacturing process control, and the use of qualified inspection procedures. The damage tolerance philosophy assumes that all components contain initial defects and the damage accumulation process is thus entirely crack propagation. The service capability of fracture critical components is determined by analysis and/or test. A large percentage of NASA fracture critical hardware is certified by analysis only and it is therefore imperative that the analysis code utilized contain the latest state of the art technology. Not all space contractors and vendors are able to develop and maintain an analysis code which meets this requirement. A response to this need has been fulfilled by the development of the NASA/FLAGRO (NASGRO) fracture mechanics code here at the Johnson Space Center. The NASGRO code is based primarily in Linear Elastic Fracture Mechanics (LEFM) and is continually updated to meet to needs of the user community. Development assistance comes from other NASA centers as well as the European Space Agency (ESA). The code currently has a vast materials data file (over 350 different entries) which defines the material strength, fracture toughness, crack growth rate, and crack growth threshold. The code also allows for user defined inputs of data from other sources. The code also provides a large library of stress intensity factor solutions for numerous geometries and loading conditions.

Patin, Raymond M.↗
NASA NTRS2014

The Use of the STAGS Finite Element Code in Stitched Structures Development

In the last 30 years NASA has worked in collaboration with industry to develop enabling technologies needed to make aircraft more fuel-efficient and more affordable. The focus on the airframe has been to reduce weight, improve damage tolerance and better understand structural behavior under realistic flight and ground loading conditions. Stitched structure is a technology that can address the weight savings, cost reduction, and damage tolerance goals, but only if it is supported by accurate analytical techniques. Development of stitched technology began in the 1990's as a partnership between NASA and Boeing (McDonnell Douglas at the time) under the Advanced Composites Technology Program and has continued under various titles and programs and into the Environmentally Responsible Aviation Project today. These programs contained development efforts involving manufacturing development, design, detailed analysis, and testing. Each phase of development, from coupons to large aircraft components was supported by detailed analysis to prove that the behavior of these structures was well-understood and predictable. The Structural Analysis of General Shells (STAGS) computer code was a critical tool used in the development of many stitched structures. As a key developer of STAGS, Charles Rankin's contribution to the programs was quite significant. Key features of STAGS used in these analyses and discussed in this paper include its accurate nonlinear and post-buckling capabilities, its ability to predict damage growth, and the use of Lagrange constraints and follower forces.

Jegley, Dawn C.↗
NASA NTRS2003

Probabilistic Evaluation of Blade Impact Damage

The response to high velocity impact of a composite blade is probabilistically evaluated. The evaluation is focused on quantifying probabilistically the effects of uncertainties (scatter) in the variables that describe the impact, the blade make-up (geometry and material), the blade response (displacements, strains, stresses, frequencies), the blade residual strength after impact, and the blade damage tolerance. The results of probabilistic evaluations results are in terms of probability cumulative distribution functions and probabilistic sensitivities. Results show that the blade has relatively low damage tolerance at 0.999 probability of structural failure and substantial at 0.01 probability.

Chamis, C. C.↗
NASA NTRS2017

Damage Simulation in Composite Materials: Why It Matters and What Is Happening Currently at NASA in This Area

Use of lightweight composite materials in space and aircraft structure designs is often challenging due to high costs associated with structural certification. Of primary concern in the use of composite structures is durability and damage tolerance. This concern is due to the inherent susceptibility of composite materials to both fabrication and service induced flaws. Due to a lack of general industry accepted analysis tools applicable to composites damage simulation, a certification procedure relies almost entirely on testing. It is this reliance on testing, especially compared to structures comprised of legacy metallic materials where damage simulation tools are available, that can drive costs for using composite materials in aerospace structures. The observation that use of composites can be expensive due to testing requirements is not new and as such, research on analysis tools for simulating damage in composite structures has been occurring for several decades. A convenient approach many researchers/model-developers in this area have taken is to select a specific problem relevant to aerospace structural certification and develop a model that is accurate within that scope. Some examples are open hole tension tests, compression after impact tests, low-velocity impact, damage tolerance of an embedded flaw, and fatigue crack growth to name a few. Based on the premise that running analyses is cheaper than running tests, one motivation that many researchers in this area have is that if generally applicable and reliable damage simulation tools were available the dependence on certification testing could be lessened thereby reducing overall design cost. It is generally accepted that simulation tools if applied in this manner would still need to be thoroughly validated and that composite testing will never be completely replaced by analysis. Research and development is currently occurring at NASA to create numerical damage simulation tools applicable to damage in composites. The Advanced Composites Project (ACP) at NASA Langley has supported the development of composites damage simulation tools in a consortium of aerospace companies with a goal of reducing the certification time of a commercial aircraft by 30%. And while the scope of ACP does not include spacecraft, much of the methodology and simulation capabilities can apply to spacecraft certification in the Space Launch System and Orion programs as well. Some specific applications of composite damage simulation models in a certification program are (1) evaluation of damage during service when maintenance may be difficult or impossible, (2) a tool for early design iterations, (3) gaining insight into a particular damage process and applying this insight towards a test coupon or structural design, and (4) analysis of damage scenarios that are difficult or impossible to recreate in a test. As analysis capabilities improve, these applications and more will become realized resulting in a reduction in cost for use of composites in aerospace vehicles. NASA is engaged in this process from both research and application perspectives. In addition to the background information discussed previously, this presentation covers a look at recent research at NASA in this area and some current/potential applications in the Orion program.

McElroy, Mack↗
NASA NTRS1995

Thermomechanical Fatigue Behavior of Three Cfcc's

The thermomechanical fatigue (TMF) behavior of three continuous-fiber ceramic composites (CFCC's) was examined. The three matrices consisted of two different glass-ceramics and silicon carbide, respectively. The matrices of some of the specimens with the silicon carbide matrix were enhanced to improve oxidation resistance. All three were reinforced with Nicalon fibers with various fiber architectures. The thermoelastic properties of the matrix relative to the fibers, the elastic moduli and coefficients of thermal expansion (CTE), were different in each composite, providing a comparison of the effects of these critical properties. The specimens were tested under in-phase (IP) and out-of-phase (OP) cyclic loadings with respect to thermal cycling between 600 C and 1100 C. One of the glass-ceramic composites had greatly increased TMF lives compared to the other glass-ceramic matrix composites. This was probably caused by the reversal of the CTE mismatch between the fibers and the matrices and caused by the different oxidation resistances of the composites. For the same TMF lives a cross-ply reinforced composite had a maximum cyclic stress half as great, as that of unidirectionally reinforced composite. This indicated that the (0 deg) ply fibers had a strong influence of TMF life. Both the glass ceramic matrix composites had shallow-slope stress-life plots that indicated sensitivity to damage and therefore low damage tolerance. The composites with an unenhanced silicon carbide matrix experienced the shortest TMF lives of all the CFCC's tested. However, the enhanced composite had the longest. Also, the enhanced composite demonstrated the best damage tolerance as evidenced by a stress-life curve that curves sharply upward to the left. In all the composites, out-of-phase cyclic loading was worse than in-phase cyclic loadings despite the CTE mismatch between fiber and matrix. A damage investigation and a preliminary analysis of the stresses in the fibers and matrices taking into account their respective thermoelastic properties provided insight into the relative behavior of the glass-ceramic matrix composites under IP and OP TMF.

Worthem, Dennis W.↗
NASA NTRS1992

Resin transfer molding of textile preforms for aircraft structural applications

The NASA LaRC is conducting and supporting research to develop cost-effective fabrication methods that are applicable to primary composite aircraft structures. One of the most promising fabrication methods that has evolved is resin transfer molding (RTM) of dry textile material forms. RTM has been used for many years for secondary structures, but has received increased emphasis because it is an excellent method for applying resin to damage-tolerant textile preforms at low cost. Textile preforms based on processes such as weaving, braiding, knitting, stitching, and combinations of these have been shown to offer significant improvements in damage tolerance compared to laminated tape composites. The use of low-cost resins combined with textile preforms could provide a major breakthrough in achieving cost-effective composite aircraft structures. RTM uses resin in its lowest cost form, and storage and spoilage costs are minimal. Near net shape textile preforms are expected to be cost-effective because automated machines can be used to produce the preforms, post-cure operations such as machining and fastening are minimized, and material scrap rate may be reduced in comparison with traditional prepreg molding. The purpose of this paper is to discuss experimental and analytical techniques that are under development at NASA Langley to aid the engineer in developing RTM processes for airframe structural elements. Included are experimental techniques to characterize preform and resin behavior and analytical methods that were developed to predict resin flow and cure kinetics.

Hasko, Gregory H.↗
NASA NTRS1993

Performance of resin transfer molded multiaxial warp knit composites

Composite materials that are subjected to complex loads have traditionally been fabricated with multidirectionally oriented prepreg tape materials. Some of the problems associated with this type of construction include low delamination resistance, poor out-of-plane strength, and labor intensive fabrication processes. Textile reinforced composites with through-the-thickness reinforcement have the potential to solve some of these problems. Recently, a relatively new class of noncrimp fabrics designated as multiaxial warp knits have been developed to minimize some of the high cost and damage tolerance concerns. Multiple stacks of warp knit fabrics can be knitted or stitched together to reduce layup labor cost. The through-the-thickness reinforcement can provide significant improvements in damage tolerance and out-of-plane strength. Multilayer knitted/stitched preforms, in conjunction with resin transfer molding (RTM), offer potential for significant cost savings in fabrication of primary aircraft structures. The objectives of this investigation were to conduct RTM processing studies and to characterize the mechanical behavior of composites reinforced with three multiaxial warp knit fabrics. The three fabrics investigated were produced by Hexcel and Milliken in the United States, and Saerbeck in Germany. Two resin systems, British Petroleum E9O5L and 3M PR 500, were characterized for RTM processing. The performance of Hexcel and Milliken quasi-isotropic knitted fabrics are compared to conventional prepreg tape laminates. The performance of the Saerbeck fabric is compared to uniweave wing skin layups being investigated by Douglas Aircraft Company in the NASA Advanced Composites Technology (ACT) program. Tests conducted include tension, open hole tension, compression, open hole compression, and compression after impact. The effects of fabric defects, such as misaligned fibers and gaps between tows, on material performance are also discussed. Estimated material and labor cost savings are projected for the Saerbeck fabric as compared to uniweave fabric currently being used by Douglas in the NASA ACT wing development program.

Dexter, H. Benson↗
NASA NTRS1992

Comparison of resin film infusion, resin transfer molding, and consolidation of textile preforms for primary aircraft structure

Innovative design concepts and cost effective fabrication processes were developed for damage tolerant primary structures that can perform at a design ultimate strain level of 6000 micro inch/inch. Attention focused on the use of textile high performance fiber reinforcement concepts that provide improved damage tolerance and out-of-plane load capability, low cost resin film infusion (RFI) and resin transfer molding (RTM) processes, and thermoplastic forming concepts. The fabrication of wing 'Y' spars by four different materials and/or processes methods is described: fabricated using IM7 angle interlock 0 to 90 deg woven preforms with + or - 45 deg plies stitched with Toray high strength graphite thread and processed using RFI and 3501-6 epoxy; fabricated using G40-800 knitted/stitched preforms and processed using RFI and 3501-6 epoxy; fabricated using G40-800 knitted/stitched preforms using RTM and Tactix 123/H41 epoxy; and fabricated preforms using AS4(6K)/PEEK 150 g commingled angle interlock 0 to 90 deg woven preforms with + or - 45 deg commingled plies stitched using high strength graphite thread and processed by consolidation. Structural efficiency, processability, and acquisition cost are compared.

Suarez, J.↗
NASA NTRS1994

Micromechanics of fatigue in woven and stitched composites

The goals of this research program were to: (1) determine how microstructural factors, especially the architecture of reinforcing fibers, control stiffness, strength, and fatigue life in 3D woven composites; (2) identify mechanisms of failure; (3) model composite stiffness; (4) model notched and unnotched strength; and (5) model fatigue life. We have examined a total of eleven different angle and orthogonal interlock woven composites. Extensive testing has revealed that these 3D woven composites possess an extraordinary combination of strength, damage tolerance, and notch insensitivity in compression and tension and in monotonic and cyclic loading. In many important regards, 3D woven composites far outstrip conventional 2D laminates or stitched laminates. Detailed microscopic analysis of damage has led to a comprehensive picture of the essential mechanisms of failure and how they are related to the reinforcement geometry. The critical characteristics of the weave architecture that promote favorable properties have been identified. Key parameters are tow size and the distributions in space and strength of geometrical flaws. The geometrical flaws should be regarded as controllable characteristics of the weave in design and manufacture. In addressing our goals, the simplest possible models of properties were always sought, in a blend of old and new modeling concepts. Nevertheless, certain properties, especially regarding damage tolerance, ultimate failure, and the detailed effects of weave architecture, require computationally intensive stochastic modeling. We have developed a new model, the 'binary model,' to carry out such tasks in the most efficient manner and with faithful representation of crucial mechanisms. This is the final report for contract NAS1-18840. It covers all work from April 1989 up to the conclusion of the program in January 1993.

Cox, B. N.↗
NASA NTRS2003

Case Studies of Fatigue Life Improvement Using Low Plasticity Burnishing in Gas Turbine Engine Applications

Surface enhancement technologies such as shot peening, laser shock peening (LSP), and low plasticity burnishing (LPB) can provide substantial fatigue life improvement. However, to be effective, the compressive residual stresses that increase fatigue strength must be retained in service. For successful integration into turbine design, the process must be affordable and compatible with the manufacturing environment. LPB provides thermally stable compression of comparable magnitude and even greater depth than other methods, and can be performed in conventional machine shop environments on CNC machine tools. LPB provides a means to extend the fatigue lives of both new and legacy aircraft engines and ground-based turbines. Improving fatigue performance by introducing deep stable layers of compressive residual stress avoids the generally cost prohibitive alternative of modifying either material or design. The X-ray diffraction based background studies of thermal and mechanical stability of surface enhancement techniques are briefly reviewed, demonstrating the importance of minimizing cold work. The LPB process, tooling, and control systems are described. An overview of current research programs conducted for engine OEMs and the military to apply LPB to a variety of engine and aging aircraft components are presented. Fatigue performance and residual stress data developed to date for several case studies are presented including: * The effect of LPB on the fatigue performance of the nickel based super alloy IN718, showing fatigue benefit of thermal stability at engine temperatures. * An order of magnitude improvement in damage tolerance of LPB processed Ti-6-4 fan blade leading edges. * Elimination of the fretting fatigue debit for Ti-6-4 with prior LPB. * Corrosion fatigue mitigation with LPB in Carpenter 450 steel. *Damage tolerance improvement in 17-4PH steel. Where appropriate, the performance of LPB is compared to conventional shot peening after exposure to engine operating temperatures.

Prevey, Paul S.↗
NASA NTRS2023

Reduction in Computational Cost of Progressive Failure Analysis of Composite Structures

Designing aircraft structures requires efficient modeling approaches to iterate on multiple structural configurations to achieve an optimal design. Typically, damage tolerance is not considered at the design stage because of the high computational cost in its implementation within a finite element modeling approach. Therefore, analytical or empirical approaches are often used to size critical damage-tolerant structures once an optimal design is determined. In this study, the Progressive Release eXplicit Virtual Crack-Closure Technique (PRX-VCCT) is assessed for its capability to cost-effectively evaluate skin stringer separation of a blade-stiffened panel that is subjected to seven point bend loads. An initial verification study was performed to evaluate PRX VCCT to accurately simulate skin-stringer separation with respect to existing cohesive element approaches. Furthermore, the influence of element size, ranging from 0.10 in. to 0.40 in., on the total computational time using the PRX-VCCT is investigated. The results indicate that the PRX-VCCT can be used to accurately simulate skin stringer separation using large element lengths (0.40 in.). Additionally, a significant reduction in the computational time to simulate skin stringer separation is observed using the PRX-VCCT. Large-scale progressive damage analysis using PRX-VCCT can be implemented early in the design cycle of composite structures without requiring a global-to-local modeling approach.

Composites↗
NASA NTRS1998

Damage Progression in Bolted Composites

Structural durability,damage tolerance,and progressive fracture characteristics of bolted graphite/epoxy composite laminates are evaluated via computational simulation. Constituent material properties and stress and strain limits are scaled up to the structure level to evaluate the overall damage and fracture propagation for bolted composites. Single and double bolted composite specimens with various widths and bolt spacings are evaluated. The effect of bolt spacing is investigated with regard to the structural durability of a bolted joint. Damage initiation, growth, accumulation, and propagation to fracture are included in the simulations. Results show the damage progression sequence and structural fracture resistance during different degradation stages. A procedure is outlined for the use of computational simulation data in the assessment of damage tolerance, determination of sensitive parameters affecting fracture, and interpretation of experimental results with insight for design decisions.

Minnetyan, Levon↗
NASA NTRS1998

Damage Progression in Bolted Composites

Structural durability, damage tolerance, and progressive fracture characteristics of bolted graphite/epoxy composite laminates are evaluated via computational simulation. Constituent material properties and stress and strain limits are scaled up to the structure level to evaluate the overall damage and fracture propagation for bolted composites. Single and double bolted composite specimens with various widths and bolt spacings are evaluated. The effect of bolt spacing is investigated with regard to the structural durability of a bolted joint. Damage initiation, growth, accumulation, and propagation to fracture are included in the simulations. Results show the damage progression sequence and structural fracture resistance during different degradation stages. A procedure is outlined for the use of computational simulation data in the assessment of damage tolerance, determination of sensitive parameters affecting fracture, and interpretation of experimental results with insight for design decisions.

Minnetyan, Levon↗
NASA NTRS1983

DC-10 composite vertical stabilizer ground test program

A review of the structural configuration and ground test program is presented. Particular emphasis is placed on the testing of a full-scale stub box test subcomponent and full span ground test unit. The stub box subcomponent was tested in an environmental chamber under ambient, cold/wet, and hot/wet conditions. The test program included design limit static loads, fatigue spectrum loading to approximately two service lifetimes (with and without damage), design limit damage tolerance tests, and a final residual strength test to a structural failure. The first full-scale ground test unit was tested under ambient conditions. The test unit was to have undergone static, fatigue, and damage tolerance tests but a premature structural failure occurred at design limit load during the third limit load test. A failure theory was developed which explains the similarity in types of failure and the large load discrepancy at failure between the two test articles. The theory attributes both failures to high stress concentrations at the edge of the lower rear spar access opening. A second full-scale ground test unit has been modified to incorporate the various changes resulting from the premature failure. The article has been assembled and is active in the test program.

Palmer, J. M., Jr.↗