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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Testing and Analysis Validation of a Metallic Repair Applied to a PRSEUS Tension Panel

A design and analysis of a repair concept applicable to a stiffened composite panel based on the Pultruded Rod Stitched Efficient Unitized Structure was recently completed. The damage scenario considered was a midbay-to-midbay saw-cut with a severed stiffener, flange and skin. Advanced modeling techniques such as mesh-independent definition of compliant fasteners and elastic-plastic material properties for metal parts were utilized in the finite element analysis supporting the design effort. A bolted metallic repair was selected so that it could be easily applied in the operational environment. The present work describes results obtained from a tension panel test conducted to validate both the repair concept and finite element analysis techniques used in the design effort. The test proved that the proposed repair concept is capable of sustaining load levels that are higher than those resulting from the current working stress allowables. This conclusion enables upward revision of the stress allowables that had been kept at an overly-conservative level due to concerns associated with repairability of the panels. Correlation of test data with finite element analysis results is also presented and assessed.

Przekop, Adam↗
NASA NTRS1977

Effect of bow-type initial imperfection on the buckling load and mass of graphite-epoxy blade-stiffened panels

A structural synthesis computer code which accounts for first order effects of an initial bow and which can be used for sizing stiffened composite panels having an arbitrary cross section is used to study graphite blade-stiffened panels. The effect of a small initial bow on both the load carrying ability of panels and on the mass of panels designed to carry a specified load is examined. Large reductions in the buckling load caused by a small initial bow emphasize the need for considering a bow when a panel is designed.

Stroud, W. J.↗
NASA NTRS1978

Graphite/epoxy orthogrid panel fabrication

The structural concept considered for a spacecraft body structure is a grid stiffened skin with a skin laminate configuration and the stiffener grid geometry selected to best suit the design requirements. The orthogrid panel developed weighs 0.55 lb/sq ft and resisted an ultimate in-plane shear load of 545 lbf/in. The basic concept of a grid stiffener composite panel is that a relatively thin skin is reinforced with a gridwork of stiffeners so that the overall panel can resist design loads without becoming structurally unstable or being overstressed. The main feature of the orthogrid panel design is that it provides the potential for low cost structural panels when advanced to the production phase. The most innovative part of the fabrication method is the foam/fiberglass stiffener web grid billet fabrication and machining to size.

Lager, J. R.↗
NASA NTRS1992

Static and fatigue testing of full-scale fuselage panels fabricated using a Therm-X(R) process

Large, curved, integrally stiffened composite panels representative of an aircraft fuselage structure were fabricated using a Therm-X process, an alternative concept to conventional two-sided hard tooling and contour vacuum bagging. Panels subsequently were tested under pure shear loading in both static and fatigue regimes to assess the adequacy of the manufacturing process, the effectiveness of damage tolerant design features co-cured with the structure, and the accuracy of finite element and closed-form predictions of postbuckling capability and failure load. Test results indicated the process yielded panels of high quality and increased damage tolerance through suppression of common failure modes such as skin-stiffener separation and frame-stiffener corner failure. Finite element analyses generally produced good predictions of postbuckled shape, and a global-local modelling technique yielded failure load predictions that were within 7% of the experimental mean.

Dinicola, Albert J.↗
NASA NTRS1990

EZDESIT: A Computer Program for Structural Element Sizing and Vehicle Weight Prediction

A method is presented for formulating stiffness terms and thermal coefficients of fiber reinforced composite, stiffened panels for input to finite element analysis (FEA). The method is robust enough to handle panels with general cross sectional shapes, including those which are unsymmetric and/or unbalanced. New thermal coefficients are introduced to quantify panel response from through-the-thickness temperature gradients. Equations are defined for stiffness, thermal expansion, and thermal bending that consider the full complement of membrane, bending, membrane-bending coupling, and transverse shear behavior. A technique of implementing this capability with a single plane of shell finite elements using the MSC/NASTRAN™ FEA program is revealed. Finally, an example of a composite, hat-stiffened panel is included to demonstrate errors that occur when an unsymmetric panel is symmetrically formulated as traditionally done. These erroneous values and the correct ones produced from the presented method are listed for the panel's stiffness matrices, thermal expansion and bending vectors, and the forces and moments that are thermally induced from in-plane and through-the-thickness temperature gradients.

Jeffrey A Cerro↗
NASA NTRS1992

Structural analysis and sizing of stiffened, metal matrix composite panels for hypersonic vehicles

The present method for strength and stability analyses of stiffened, fiber-reinforced composite panels to be used in hypersonic vehicle structures is of great generality, and can be linked with planar finite-element analysis (FEA). Nonlinear temperature and load-dependent material data for each laminate are used to 'build-up' the stiffened panel's membrane, bending, and membrane-bending coupling stiffness terms, as well as thermal coefficients. The resulting, FEA-solved thermomechanical forces and moments are used to calculate strain at any location in the panel; this allows an effective ply-by-ply orthotropic strength analysis to be conducted, together with orthotropic instability checks for each laminated segment of the cross-section.

Collier, Craig S.↗
NASA NTRS1994

Benchmark panels

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

T H Walker↗
NASA NTRS1985

Failure mechanisms of laminates transversely loaded by bolt push-through

Stiffened composite panels proposed for fuselage and wing design utilize a variety of stiffener-to-skin attachment concepts including mechanical fasteners. The attachment concept is an important factor influencing the panel's strength and can govern its performance following local damage. Mechanical fasteners can be an effective method for preventing stiffener-skin separation. One potential failure mode for bolted panels occurs when the bolts pull through the stiffener attachment flange or skin. The resulting loss of support by the skin to the stiffener and by the stiffener to the skin can result in local buckling and subsequent panel collapse. The characteristic failure modes associated with bolt push-through failure are described and the results of a parametric study of the effects that different material systems, boundary conditions, and laminates have on the forces and displacements required to cause damage and bolt pushthrough failure are presented.

Waters, W. A., Jr.↗
NASA NTRS1996

Design and evaluation of a bolted joint for a discrete carbon-epoxy rod-reinforced hat section

The use of prefabricated pultruded carbon-epoxy rods has reduced the manufacturing complexity and costs of stiffened composite panels while increasing the damage tolerance of the panels. However, repairability of these highly efficient discrete stiffeners has been a concern. Design, analysis, and test results are presented in this paper for a bolted-joint repair for the pultruded rod concept that is capable of efficiently transferring axial loads in a hat-section stiffener on the upper skin segment of a heavily loaded aircraft wing component. A tension and a compression joint design were evaluated. The tension joint design achieved approximately 1.0% strain in the carbon-epoxy rod-reinforced hat-section and failed in a metal fitting at 166% of the design ultimate load. The compression joint design failed in the carbon-epoxy rod-reinforced hat-section test specimen area at approximately 0.7% strain and at 110% of the design ultimate load. This strain level of 0.7% in compression is similar to the failure strain observed in previously reported carbon-epoxy rod-reinforced hat-section column tests.

Rousseau, Carl Q.↗
NASA NTRS1996

Design and Evaluation of a Bolted Joint for a Discrete Carbon-Epoxy Rod-Reinforced Hat Section

The use of pre-fabricated pultruded carbon-epoxy rods has reduced the manufacturing complexity and costs of stiffened composite panels while increasing the damage tolerance of the panels. However, repairability of these highly efficient discrete stiffeners has been a concern. Design, analysis, and test results are presented in this paper for a bolted-joint repair for the pultruded rod concept that is capable of efficiently transferring axial loads in a hat-section stiffener on the upper skin segment of a heavily loaded aircraft wing component. A tension and a compression joint design were evaluated. The tension joint design achieved approximately 1.0 percent strain in the carbon-epoxy rod-reinforced hat-section and failed in a metal fitting at 166 percent of the design ultimate load. The compression joint design failed in the carbon-epoxy rod-reinforced hat-section test specimen area at approximately 0.7 percent strain and at 110 percent of the design ultimate load. This strain level of 0.7 percent in compression is similar to the failure strain observed in previously reported carbon-epoxy rod-reinforced hat-section column tests.

Baker, Donald J.↗
NASA NTRS1990

Postbuckling behavior of curved panels under combined compression and shear loads

A validated semiempirical design procedure and fatigue data were developed for curved, stiffened composite panels operating in the postbuckled regime under the action of combined compression and shear loading. A previously developed design methodology for composite panels under pure shear or pure compression loading was used as the starting point for the program. Initially, the well established interaction rules for metal panels were adopted to predict buckling under combined loading. Test data were then developed to verify these rules and suggest modifications where necessary. Postbuckling failure envelopes were developed by accounting for the failure modes possible under shear loading only, and under pure compression loading. Static failure predictions under combined loading were based on test verified interaction criteria. Fatigue tests were conducted under combined loading to determine strength degradation and the possible failure modes.

Deo, Ravi B.↗
NASA NTRS1981

PASCO: Structural panel analysis and sizing code: Users manual - Revised

A computer code denoted PASCO is described for analyzing and sizing uniaxially stiffened composite panels. Buckling and vibration analyses are carried out with a linked plate analysis computer code denoted VIPASA, which is included in PASCO. Sizing is based on nonlinear mathematical programming techniques and employs a computer code denoted CONMIN, also included in PASCO. Design requirements considered are initial buckling, material strength, stiffness and vibration frequency. A user's manual for PASCO is presented.

Anderson, M. S.↗
NASA NTRS2010

Assessment of the Damage Tolerance of Postbuckled Hat-Stiffened Panels Using Single-Stringer Specimens

A procedure is proposed for the assessment of the damage tolerance and collapse of stiffened composite panels using a single-stringer compression specimen. The dimensions of the specimen are determined such that the specimen s nonlinear response and collapse are representative of an equivalent multi-stringer panel in compression. Experimental tests are conducted on specimens with and without an embedded delamination. A shell-based finite element model with intralaminar and interlaminar damage capabilities is developed to predict the postbuckling response as well as the damage evolution from initiation to collapse.

Bisagni, Chiara↗
NASA NTRS1992

User's manual for MacPASCO

A user's manual is presented for MacPASCO, which is an interactive, graphic, preprocessor for panel design. MacPASCO creates input for PASCO, an existing computer code for structural analysis and sizing of longitudinally stiffened composite panels. MacPASCO provides a graphical user interface which simplifies the specification of panel geometry and reduces user input errors. The user draws the initial structural geometry and reduces user input errors. The user draws the initial structural geometry on the computer screen, then uses a combination of graphic and text inputs to: refine the structural geometry; specify information required for analysis such as panel load and boundary conditions; and define design variables and constraints for minimum mass optimization. Only the use of MacPASCO is described, since the use of PASCO has been documented elsewhere.

Lucas, S. H.↗
NASA NTRS1980

PASCO: Structural panel analysis and sizing code, capability and analytical foundations

A computer code denoted PASCO which can be used for analyzing and sizing uniaxially-stiffened composite panels is described. Buckling and vibration analyses are carried out with a linked-plate analysis computer code denoted VIPASA, which is incorporated in PASCO. Sizing is based on nonlinear mathematical programming techniques and employs a computer code denoted CONMIN, also incorporated in PASCO. Design requirements considered are initial buckling, material strength, stiffness, and vibration frequency. The capability of the PASCO computer code and the approach used in the structural analysis and sizing are described.

Stroud, W. J.↗
NASA NTRS1991

MacPASCO - A Macintosh-based, interactive graphic preprocessor for structural analysis and sizing

MacPASCO, an interactive, graphic preprocessor for panel design is described. MacPASCO creates input for PASCO, an existing computer code for structural analysis and optimization of longitudinal stiffened composite panels. By using a graphical user interface, MacPASCO simplifies the specification of panel geometry and reduces user input errors, thus making the modeling and analysis of panel designs more efficient. The user draws the initial structural geometry on the computer screen, then uses a combination of graphic and text inputs to: refine the structural geometry, specify information required for analysis such as panel load conditions, and define design variables and constraints for minimum-mass optimization. Composite panel design is an ideal application because the graphical user interface can: serve as a visual aid, eliminate the tedious aspects of text-based input, and eliminate many sources of input errors.

Lucas, S. H.↗
NASA NTRS1993

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 34th and AIAA/ASME Adaptive Structures Forum, La Jolla, CA, Apr. 19-22, 1993, Technical Papers. Pts. 1-6

Topics addressed include the prediction of helicopter component loads using neural networks, spacecraft on-orbit coupled loads analysis, hypersonic flutter of a curved shallow panel with aerodynamic heating, thermal-acoustic fatigue of ceramic matrix composite materials, transition elements based on transfinite interpolation, damage progression in stiffened composite panels, a direct treatment of min-max dynamic response optimization problems, and sources of helicopter rotor hub inplane shears. Also discussed are dynamics of a layered elastic system, confidence bounds on structural reliability, mixed triangular space-time finite elements, advanced transparency development for USAF aircraft, a low-velocity impact on a graphite/PEEK, an automated mode-tracking strategy, transonic flutter suppression by a passive flap, a nonlinear response of composite panels to random excitation, an optimal placement of elastic supports on a simply supported plate, a probabilistic assessment of composite structures, a model for mode I failure of laminated composites, a residual flexibility approach to multibody dynamics,and multilayer piezoelectric actuators.

Source record↗
NASA NTRS1994

Out of plane gradients in laminated composites

The research under this grant focused on three different, but related, research projects. All three projects were concerned with predicting the interlaminar stresses that develop in laminated composites in the presence of free edges. The first project was dedicated to the development of a new semi-analytical approximate solution for predicting interlaminar stresses. The second project was devoted to a study of the influence of boundary conditions, geometric parameters, and material modelling on interlaminar stresses in stiffened composite panels. The third project was dedicated to the development of a new finite element for efficient analysis of interlaminar stresses in curved, stiffened panels. A summary of each of the projects is given.

Herakovich, Carl T.↗