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At least 271 records · Page 15

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NASA NTRS1975

Repair of boron/aluminum composites

A program was performed to determine the repairability of boron/aluminum aerospace structural components. During the program, a list of repair guidelines was prepared to aid in determining the proper repair techniques for a given structure. These guidelines included specifying types of repair material and their applicability, corrosion prevention procedures, design criteria, and inspection criteria. Six sets of boron/aluminum structural components were repaired and tested to compare as-fabricated and repaired performance. The specimens included a honeycomb-stiffened panel, elastically buckled tubes, a skin-stringer panel, a tube combining bending and tension, a splice joint specimen, and a tension field panel. All but one set of specimens, when repaired, exceeded the strength of the original specimens. The repairs resulted in an average weight increase per structure of 9% and an average performance increase of 27%. It is concluded that metal-matrix composite material, damaged in service, can be repaired by techniques not very different from those currently in use for conventional materials.

Christian, J. L.↗

NASA NTRS1993

A weight-efficient design strategy for cutouts in composite transport structures

Two design procedures for composite panels with cutouts are described and illustrated by example applications. One of these procedures uses a specialized cutout analysis code to obtain preliminary sizing information for the panel laminate, cutout padup, and cutout stiffener reinforcements. The other procedure uses a finite element based structural optimization code to develop a minimum weight panel design. The best features of both procedures form the basis of a design strategy for weight-efficient cutout panels.

Russell, S. G.↗

NASA NTRS2018

Detection of Damage During Quasi-Static Loading of a Single Stringer Panel Using Passive Thermography and Acoustic Emission

The purpose of this paper is to describe the methodology of measuring damage onset and growth in a composite structure during quasi-static loading using passive thermography and acoustic emission. The early detection and measurement of damage progression is important to understand failure modes. A single stringer panel was subjected to quasi-static loading to induce deformation which resulted in the formation of damage between the stiffener flange and skin. The loading was stopped when damage growth was detected. Passive thermography and acoustic emission were used to detect damage in real-time as a function of the applied load. Of particular interest are the small transient thermographic signals resulting from damage formation which can be challenging to detect, as compared to the persistent passive thermography indications of cyclic fatigue loading. We describe a custom developed thermal inspection system for detection of composite damage during quasi-static loading. The thermal results are compared to a two-dimensional multi-layered thermal simulation based on the quadrupole method. Acoustic emission is used to further characterize the damage by comparing the acoustic emission events with the thermal imagery. Results are compared to ultrasonic measurements to document the damage through-the-thickness.

Zalameda, Joseph N.↗

NASA NTRS1997

Global Cost and Weight Evaluation of Fuselage Side Panel Design Concepts

This report documents preliminary design trades conducted under NASA contracts NAS1 18889 (Advanced Technology Composite Aircraft Structures, ATCAS) and NAS1-19349 (Task 3, Pathfinder Shell Design) for a subsonic wide body commercial aircraft fuselage side panel section utilizing composite materials. Included in this effort were (1) development of two complete design concepts, (2) generation of cost and weight estimates, (3) identification of technical issues and potential design enhancements, and (4) selection of a single design to be further developed. The first design concept featured an open-section stringer stiffened skin configuration while the second was based on honeycomb core sandwich construction. The trade study cost and weight results were generated from comprehensive assessment of each structural component comprising the fuselage side panel section from detail fabrication through airplane final assembly. Results were obtained in three phases: (1) for the baseline designs, (2) after global optimization of the designs, and (3) the results anticipated after detailed design optimization. A critical assessment of both designs was performed to determine the risk associated with each concept, that is the relative probability of achieving the cost and weight projections. Seven critical technical issues were identified as the first step towards side panel detailed design optimization.

Polland, D. R.↗

NASA NTRS1988

Composite transport wing technology development: Design development tests and advanced structural concepts

Numerous design concepts, materials, and manufacturing methods were investigated for the covers and spars of a transport box wing. Cover panels and spar segments were fabricated and tested to verify the structural integrity of design concepts and fabrication techniques. Compression tests on stiffened panels demonstrated the ability of graphite/epoxy wing upper cover designs to achieve a 35 percent weight savings compared to the aluminum baseline. The impact damage tolerance of the designs and materials used for these panels limits the allowable compression strain and therefore the maximum achievable weight savings. Bending and shear tests on various spar designs verified an average weight savings of 37 percent compared to the aluminum baseline. Impact damage to spar webs did not significantly degrade structural performance. Predictions of spar web shear instability correlated well with measured performance. The structural integrity of spars manufactured by filament winding equalled or exceeded those fabricated by hand lay-up. The information obtained will be applied to the design, fabrication, and test of a full-scale section of a wing box. When completed, the tests on the technology integration box beam will demonstrate the structural integrity of an advanced composite wing design which is 25 percent lighter than the metal baseline.

Griffin, Charles F.↗

NASA NTRS2013

Applying a Stitched, Rod-Stiffened Concept to Heavily Loaded Structure

NASA and the Boeing Company have worked to develop new low-cost, light-weight composite structures for aircraft. A stitched carbon-epoxy material system was developed to reduce the weight and cost of transport aircraft wing structure, first in the NASA Advanced Composites Technology (ACT) program in the 1990's and now in the Environmentally Responsible Aviation (ERA) Project. By stitching through the thickness of a dry carbon fiber material prior to cure, the labor associated with panel fabrication and assembly can be significantly reduced and the need for mechanical fasteners is almost eliminated. Stitching provides the benefit of reducing or eliminating delaminations, including those between stiffener flanges and skin. Stitching also reduces part count, and therefore, cost of the structure. The stitched panel concept used in the ACT program in the 1990's used simple blade-stiffeners as stringers, caps and clips. Today, the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept is being developed for application to advanced vehicle configurations. PRSEUS provides additional weight savings through the use of a stiffener with a thin web and a unidirectional carbon rod at the top of the web which provides structurally efficient stiffening. A comparison between the blade-stiffened structure and PRSEUS is presented focusing on highly loaded structure and demonstrating improved weight reduction.

Jegley, Dawn C.↗

NASA NTRS1979

Fabrication research for supersonic cruise aircraft

Advanced fabrication and joining processes for titanium and composite materials are being investigated by NASA to develop technology for the Supersonic Cruise Research (SCR) Program. Full-scale structural panels are being designed and fabricated to meet the criteria of an existing integrally stiffened shear panel on the upper wing surface of the NASA YF-12 aircraft. The program consists of laboratory testing and Mach 3 flight service of full-scale structural panels and laboratory testing of representative structural element specimens. Borsic/aluminum honeycomb-core, titanium clad Borsic/aluminum skin-stringer, graphite/PMR-15 polyimide honeycomb-core, and titanium superplastically formed/diffusion bonded panels have been designed, fabricated, and tested. Graphite/LARC-160 polyimide skin-stringer panels have been designed, and fabrication methods are being developed.

Hoffman, E. L.↗

NASA NTRS1992

Test and analysis results for composite transport fuselage and wing structures

Automated tow placement (ATP) and stitching of dry textile composite preforms followed by resin transfer molding (RTM) are being investigated by researchers at NASA LaRC and Douglas Aircraft Company as cost-effective manufacturing processes for obtaining damage tolerant fuselage and wing structures for transport aircraft. The Douglas work is being performed under a NASA contract entitled 'Innovative Composites Aircraft Primary Structures (ICAPS)'. Data are presented in this paper to assess the damage tolerance of ATP and RTM fuselage elements with stitched-on stiffeners from compression tests of impacted three-J-stiffened panels and from stiffener pull-off tests. Data are also presented to assess the damage tolerance of RTM wing elements which had stitched skin and stiffeners from impacted single stiffener and three blade-stiffened compression tests and stiffener pull-off tests.

Deaton, Jerry W.↗

NASA NTRS2005

On Structural Design of a Mobile Lunar Habitat With Multi- Layered Environmental Shielding

This report presents an overview of a Mobile Lunar Habitat (MLH) structural design consisting of advanced composite materials. The habitat design is derived from the cylindrical-shaped U.S. Lab module aboard the International Space Station (ISS) and includes two lateral ports and a hatch at each end that geometrically match those of the ISS Nodes. Thus, several MLH units can be connected together to form a larger lunar outpost of various architectures. For enhanced mobility over the lunar terrain, the MLH uses six articulated insect-like robotic, retractable legs enabling the habitat to .t aboard a launch vehicle. The carbon-composite shell is sandwiched between two layers of hydrogen-rich polyethylene for enhanced radiation shielding. The pressure vessel is covered by modular double-wall panels for meteoroid impact shielding supported by externally mounted stiffeners. The habitat s structure is an assembly of multiple parts manufactured separately and bonded together. Based on the geometric complexity of a part and its material system, an appropriate fabrication process is proposed.

Pruitt, J. R.↗

NASA NTRS2012

Failure at Frame-Stringer Intersections in PRSEUS Panels

NASA, the Air Force Research Laboratory and The Boeing Company have worked to develop new low-cost, light-weight composite structures for aircraft. A Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept has been developed which offers advantages over traditional metallic structures. In this concept a stitched carbon-epoxy material system has been developed with the potential for reducing the weight and cost of transport aircraft structure by eliminating fasteners, thereby reducing part count and labor. By adding unidirectional carbon rods to the top of stiffeners, the panel becomes more structurally efficient. This combination produces a more damage tolerant design. This study focuses on the intersection between the rod-stiffener and the foam-filled frame in a PRSEUS specimen. Compression loading is considered, which induces stress concentrations at the intersection point that can lead to failures. An experiment with accompanying analysis for a single-frame specimen is described, followed by a parametric study of simple reinforcements to reduce strains in the intersection region.

Jegley, Dawn C.↗

NASA NTRS1989

Computational procedures for postbuckling of composite shells

A recently developed finite-element capability for general nonlinear shell analysis, featuring the use of three-dimensional constitutive equations within an efficient resultant-oriented framework, is employed to simulate the postbuckling response of an axially compressed composite cylindrical panel with a circular cutout. The problem is a generic example of modern composite aircraft components for which postbuckling strength (i.e., fail-safety) is desired in the presence of local discontinuities such as holes and cracked stiffeners. While the computational software does a reasonable job of predicting both the buckling load and the qualitative aspects of postbuckling (compared both with experiment and another code) there are some discrepancies due to: (1) uncertainties in the nominal layer material properties, (2) structural sensitivity to initial imperfections, and (3) the neglect of dynamic and local material delamination effects in the numerical model. Corresponding refinements are suggested for the realistic continuation of this type of analysis.

Stanley, G. M.↗

NASA NTRS1991

Advanced fiber placement of composite fuselage structures

The Hercules/NASA Advanced Composite Technology (ACT) program will demonstrate the low cost potential of the automated fiber placement process. The Hercules fiber placement machine was developed for cost effective production of composite aircraft structures. The process uses a low cost prepreg tow material form and achieves equivalent laminate properties to structures fabricated with prepreg tape layup. Fiber placement demonstrations planned for the Hercules/NASA program include fabrication of stiffened test panels which represent crown, keel, and window belt segments of a typical transport aircraft fuselage.

Anderson, Robert L.↗

NASA NTRS1998

Development of Textile Reinforced Composites for Aircraft Structures

NASA has been a leader in development of composite materials for aircraft applications during the past 25 years. In the early 1980's NASA and others conducted research to improve damage tolerance of composite structures through the use of toughened resins but these resins were not cost-effective. The aircraft industry wanted affordable, robust structures that could withstand the rigors of flight service with minimal damage. The cost and damage tolerance barriers of conventional laminated composites led NASA to focus on new concepts in composites which would incorporate the automated manufacturing methods of the textiles industry and which would incorporate through-the-thickness reinforcements. The NASA Advanced Composites Technology (ACT) Program provided the resources to extensively investigate the application of textile processes to next generation aircraft wing and fuselage structures. This paper discusses advanced textile material forms that have been developed, innovative machine concepts and key technology advancements required for future application of textile reinforced composites in commercial transport aircraft. Multiaxial warp knitting, triaxial braiding and through-the-thickness stitching are the three textile processes that have surfaced as the most promising for further development. Textile reinforced composite structural elements that have been developed in the NASA ACT Program are discussed. Included are braided fuselage frames and window-belt reinforcements, woven/stitched lower fuselage side panels, stitched multiaxial warp knit wing skins, and braided wing stiffeners. In addition, low-cost processing concepts such as resin transfer molding (RTM), resin film infusion (RFI), and vacuum-assisted resin transfer molding (VARTM) are discussed. Process modeling concepts to predict resin flow and cure in textile preforms are also discussed.

Dexter, H. Benson↗

NASA NTRSDate not supplied

Progressive Damage Failure Analysis of Post-Buckled Composite Single-Stringer Panel with Teflon Inserts

The high strength-to-weight and stiffness-to-weight ratio materials, such as laminated composites, are advantageous for modern aircraft. Laminated composites with initial flaws are susceptible to delamination under buckling loads. PDA tools help enhance the industry’s understanding of the mechanisms for damage initiation and growth in composite structures while assisting in the design, analysis, and sustainment methods of these composite structures. The global-local modeling approach for the single-stringer post-buckled panel was evaluated through this effort, using Teflon inserts to simulate the defect of damage during manufacturing. This understanding is essential for designing the post-buckled structure, reducing weight while predicting damage initiation location, and addressing a potential design review for future aircraft repairs. In this work, the initial damage was captured with Teflon inserts as the starting configuration; and any reference to the damage initiation refers to any damage beyond the “initial unbonded region.” The effort aims to develop, evaluate, and enhance methods to predict damage initiation and progression and the failure of post-buckled hat-stiffened panels using multiple Abaqus FEA Virtual Crack Closure Technique (VCCT) definitions. Validation of the PDA using the VCCT material model was performed on a large single-stringer panel subjected to compressive loading. The compressive loading of the panel caused the skin to buckle before any damage began to occur locally. In addition, comparisons are made for critical aspects of the damage morphology, such as a growth pattern that included delamination from the skin-stiffener interface to the skin and ply interfaces. When compared against the experimental data produced through the NASA Advanced Composites Project (ACP), the present model captured damage migration from one surface to another, and model validations were within 4% of the experimental data.

PDA↗

NASA NTRS1983

Structures and subsystems

New and improved materials combined with efficient structural design concepts have made an essential contribution towards the shaping of the present transportation system (STS), and now, for the replacement of the STS in the year 2005, new materials and novel designs are being studied to identify the technologies which should be developed for a low-cost future space transportation system (FSTS). Three basic structural arrangements were considered for the FSTS orbiter. They include a nonintegral tank arrangement, an integral-tank arrangement, and a hybrid of the first two. Three representative arrangements regarding wall constructions are considered. Each employs a blade-stiffened aluminum tank with reinforced closed-cell-foam cryogenic insulation. Attention is given to an aluminum-alloy structure, a graphite-epoxy structure, a graphite-polyimide structure, a carbon-carbon surface panel structure, a graphite-composite fuselage structure, serviceability and all-weather considerations, and structural concept ratings.

Taylor, A. H.↗

NASA NTRS1975

Fabrication methods for YF-12 wing panels for the Supersonic Cruise Aircraft Research Program

Advanced fabrication and joining processes for titanium and composite materials are being investigated by NASA to develop technology for the Supersonic Cruise Aircraft Research (SCAR) Program. With Lockheed-ADP as the prime contractor, full-scale structural panels are being designed and fabricated to replace an existing integrally stiffened shear panel on the upper wing surface of the NASA YF-12 aircraft. The program involves ground testing and Mach 3 flight testing of full-scale structural panels and laboratory testing of representative structural element specimens. Fabrication methods and test results for weldbrazed and Rohrbond titanium panels are discussed. The fabrication methods being developed for boron/aluminum, Borsic/aluminum, and graphite/polyimide panels are also presented.

Hoffman, E. L.↗

NASA NTRS1997

Advanced Technology Composite Fuselage - Manufacturing

The goal of Boeing's Advanced Technology Composite Aircraft Structures (ATCAS) program is to develop the technology required for cost-and weight-efficient use of composite materials in transport fuselage structure. Carbon fiber reinforced epoxy was chosen for fuselage skins and stiffening elements, and for passenger and cargo floor structures. The automated fiber placement (AFP) process was selected for fabrication of stringer-stiffened and sandwich skin panels. Circumferential and window frames were braided and resin transfer molded (RTM'd). Pultrusion was selected for fabrication of floor beams and constant-section stiffening elements. Drape forming was chosen for stringers and other stiffening elements cocured to skin structures. Significant process development efforts included AFP, braiding, RTM, autoclave cure, and core blanket fabrication for both sandwich and stiffened-skin structure. Outer-mold-line and inner-mold-line tooling was developed for sandwich structures and stiffened-skin structure. The effect of design details, process control and tool design on repeatable, dimensionally stable, structure for low cost barrel assembly was assessed. Subcomponent panels representative of crown, keel, and side quadrant panels were fabricated to assess scale-up effects and manufacturing anomalies for full-scale structures. Manufacturing database including time studies, part quality, and manufacturing plans were generated to support the development of designs and analytical models to access cost, structural performance, and dimensional tolerance.

Wilden, K. S.↗

NASA NTRS2018

Real Time Detection of Damage During Quasi-Static Loading of a Single Stringer Panel Using Passive Thermography

Real time nondestructive evaluation is required for composites load testing. The early detection and measurement of damage progression is important to understand failure modes. A single stringer panel was subjected to quasi-static loading to induce deformation which can result in the formation of delamination damage between the stiffener flange and skin. Passive thermography was used to detect damage in real time as a function of the applied load. The loading was stopped when damage growth was detected. Of particular interest was the early detection of damage formation which can be challenging, as compared to cyclic fatigue loading. Passive thermography data were acquired and processed in real time and revealed damaged areas due to heating from fiber breakage and delamination formation. The processed thermal imagery was also compared to acoustic emission and ultrasound data.

Zalameda, Joseph N.↗