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Composite materials

2nd International Conference on Composites Testing and Model Identification : Comptest 2004

The site provides access to the proceedings of the Comptest 2004 conference, held on the 21st - 23rd September 2004, hosted by the Department of Aerospace Engineering, University of Bristol, U.K. In all 106 papers were presented both aurally and on posters by authors from 25 different countries. The presentations are arranged by author and by session and there is a book of abstracts and a delegate list both of which are in PDF format.

A Comparison of CEN and ASTM Test Methods for Composite Materials

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA AR-04/24 by Daniel O. Adams dated June 2004. A detailed test method comparison was performed to assess the equivalence of test methods for composite materials, with emphasis on the Committee for European Standardization and ASTM International test methods referred to in the SAE International Aerospace Material Specification (AMS) 2980 and 3970 specifications. This comparison included both the parameters associated with two comparable test methods and the additions and changes listed in the AMS specifications. For the types of tests where only one test method is referred to in the SAE specifications, a second comparable ASTM or Suppliers of Advanced Composite Materials Association test method was selected for comparison purposes. In total, two test methods were reviewed and compared for a total of 16 different types of tests. For each type of test, three comparison tables are presented, focusing on geometric features of the specimen and test fixture, parameters associated with the test procedure, and procedures for data reduction and reporting. Each table contains a list of individual parameters specified in the test methods as well as any additions or changes provided in the AMS 2980 and 3970 specifications. For every parameter listed in the comparison tables, an assessment of the equivalence was made using a 0-4 rating scale. A brief summary of each test method comparison is provided, which emphasizes the most significant differences between the test methods. Based on the comparative assessments performed, 4 of the 16 types of tests were recommended for follow-on testing to further assess test method equivalency. Note that the selection of these four tests for follow-on testing only reflects the need for additional test data to assess equivalency and is not a reflection of their degree of equivalency relative to the other tests. The four types of tests recommended for follow-on testing are lamina compression testing to assess the effects of gage length, laminate compression testing to assess the effects of loading method, in-plane shear testing to investigate the effects of specimen thickness, and constituent content determinations to investigate the effects of specimen size and weighing accuracy. [Taken from abstract]. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

A Contribution to the Finite Element Formulation for the Analysis of Composite Sandwich Shells

This web site provides access to a University of Cincinnati, Department of Aerospace Engineering and Engineering Mechanics PhD dissertation, by Romil R. Tanov, dated 2000. The objective of the study was to devise an accurate and efficient analysis approach for composite and sandwich shells, which would be simple enough to be capable of implementing into a FE code without significantly affecting its computational efficiency, and at the same time would give good accuracy in predicting the behavior of layered shells. Bibliographic and abstract details are available in HTML format. The title page, contents and the full text of the document are accessible online in PDF format (3.14 MB). This title is part of the OhioLink Electronic Theses and Dissertations Project.

A Decision Support System for Advanced Composites Manufacturing Cost Estimation

This web site provides access to a Virginia Polytechnic Institute and State University Department of Industrial and Systems Engineering PhD dissertation, by Mark Alan Eaglesham, dated 10 April 1998. The dissertation describes the development of a methodology for the improvement of cost estimation at the conceptual design phase. This methodology uses intelligent searching and storage of existing accounting data in order to enhance access and retrieval. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format [951.15 Kb]. This title is part of Virginia Tech's Electronic Thesis and Dissertation Collection (VT ETD).

A New Scheme for the Optimum Design of Stiffened Composite Panels with Geometric Imperfections

This web page provides access to a Virginia Polytechnic Institute and State University Department of Aerospace and Ocean Engineering PhD dissertation, by Mohamed Elseifi, dated 9 November 1998. The thesis desribes the development of new methodology in which a manufacturing model and a convex model for uncertainties are used in conjunction with a nonlinear design tool in order to improve the design of thin walled stiffened composite panels. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in a series of PDF format files. This title is part of Virginia Tech’s Electronic Thesis and Dissertation Collection (VT ETD).

A Novel Method for Characterizing the Impact Response of Functionally Graded Plates

This is the full text of a thesis written by Reid Larson which was presented to the Air Force Institute of Technology, Wright Patterson Air Force Base, Ohio, in September 2008. Functionally graded material (FGM) plates are advanced composites with properties that vary continuously through the thickness of the plate. Metal-ceramic FGM plates have been proposed for use in thermal protection systems where a metal-rich interior surface of the plate gradually transitions to a ceramic-rich exterior surface of the plate. The ability of FGMs to resist impact loads must be demonstrated before using them in high-temperature environments in service. This dissertation presents a novel technique by which the impact response of FGM plates is characterized for low-velocity, low- to medium-energy impact loads. An experiment was designed where strain histories in FGM plates were collected during impact events. These strain histories were used to validate a finite element simulation of the test. An optimization technique was applied to estimate local material properties in the FGM plate to enhance the finite element simulation. The optimized simulation captured the physics of the impact events. The method allows research & design engineers to make informed decisions necessary to implement FGM plates in aerospace platforms. [Taken from Abstract]. This is in PDF format so Adobe Acrobat software is required in order to read it.

A Predictive Methodology for Delamination Growth in Laminated Composites : Part 1 : Theoretical Development and Preliminary Experimental Results

This final report (DOT/FAA/AR-97/87) was published by the Federal Aviation Administration in April 1998, and was written by Barry D. Davidson. A methodology is presented for the prediction of delamination growth in laminated structures. The report also addresses the issue that many laminated composites exhibit large-scale crack-tip damage zones and, as such, a singular field-based mode mix decomposition may not accurately account for the dependence of toughness on the loading [extracted from author abstract]. This is a PDF file so Adobe Acrobat software will be required in order to read it.

A Predictive Methodology for Delamination Growth in Laminated Composites Part II : Analysis, Applications, and Accuracy Assessment : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/56, by Barry D. Davidson, dated October 2001. A nonclassical, energy release rate-based approach to predict delamination growth is described that overcomes the limitations of current, state-of-the art methodologies. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

A Three-Phase Constitutive Model for Macrobrittle Fatigue Damage of Composites

This web site provides access to a West Virginia University Department of Mechanical and Aerospace Engineering PhD dissertation, by Gasser F. Abdelal, dated 15 June 2000. The thesis presents an examination of damage in composite materials. A micromechanical macrobrittle damage (damage under monotonic load) model is developed and compared with available experimental data. An anisotropic fatigue damage (damage under cyclic load) model is derived and integrated with the micromechanical macrobrittle damage model to develop a micromechanical fatigue damage model. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format [1.18 Mb]. This title is part of West Virginia University's Electronic Thesis and Dissertation Project.

Advanced Aircraft Materials, Engine Debris Penetration Testing

This technical report (DOT/FAA/AR-03/37) was published by the Federal Aviation Administration (FAA) Office of Aviation Research in December 2005 and was written by Steven J. Lundin and Richard B. Mueller. This report documents the results of testing conducted in July and August 2001 at the Naval Air Warfare Center-Weapons Division, China Lake, CA, as part of the continued effort to characterize uncontained engine events. This effort was performed in support of the Federal Aviation Administration Aircraft Catastrophic Failure Prevention Program. Data generated from this test will support the penetration equation development for the Uncontained Engine Debris Damage Analysis Model (UEDDAM), a developmental design tool for conducting aircraft safety analysis for engine rotor burst events. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Advanced Manufacturing Research Centre (AMRC)

The Advanced Manufacturing Research Centre (AMRC) is a partnership between the University of Sheffield and Boeing. Its aim is to become a world-class global research facility developing innovative and advanced technology solutions for advanced materials forming. The site provides detailed information about AMRC's core research areas (structural integrity, dynamic analysis, damping, surface integrity, virtual reality, machining, complexity, manufacturing management, additive processes and thixo-forming), and a list of publications. Activities within schools and FE colleagues are described as are services that can be offered to industry (e.g. consultancy, rapid prototyping). There is also an FAQ, a news archive, and a career opportunities section. Part of the site is for members only,

Advanced Polymeric & Metallic Composite Materials for Space and Aerospace Vehicle Structures and Strength Optimization of Composite Structures and their Certification

This is Research and Technology Organization (RTO) AGARD Lecture Series report, AGARD-LS-204, dated December 1995. This lecture series presents and discusses the sci entific problem of advanced polymer and metallic c omposite materials for aerospace structures, stren gth optimization of composite structures, and thei r certification. Some challenges of using composit e structures, including airframe concept definitio n, are studied. Fiber orientation optimization pri nciples for composite panels and shells are outlin ed. Procedures for certification of assemblies mad e out of composites are dealt with. Certification requirements, including requirements to estimate s tatic and fatigue strengths, are formulated. Desig n conditions for composite structures are analyzed , including development. For individual titles, see N96-23937 through N96-23946. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (33 kB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

AEA Technology

AEA Technology focuses on five key areas: technology-based products, specialised science, environmental management, improving the efficiency of industrial plant, and risk assessment and safety management. The site describes the capabilities, products and services of the company. There is a site search facility and a products and services catalogue which can be searched and browsed. The 'investor relations' area provides annual reports and account information from 1997 in PDF format. Up-to-date company news is available.

AeroInfo

AeroInfo is a searchable and browsable aerospace information directory, produced by the Information Center for Aerospace Science and Technology (ICAST) at the National Aerospace Laboratories in Bangalore. It covers most fields of aerospace including aeronautics, aviation, aerodynamics, CFD, avionics, composites, mechanical engineering, materials science, and computer science. There is also a section on worldwide aerospace agencies and a searchable database of papers, journals, technical reports, standards, patents and regulatory information.

Air Force Research Laboratory : Space Vehicles Directorate

Space Vehicles is one of nine Air Force Research Laboratories directorates, and its main purpose is to develop technologies that support evolving warfighter requirements to control and exploit space. The keys areas of research include the battlespace environment, protection of space assets, space vehicle control, space-based sensing, space vehicle technologies (structures, power, thermal management) wargaming, and performing a variety of integrated space technology demonstrations. The resources available from the site include news releases, a newsletter, fact sheets, an overview of the directorate's research interests and activities, and a bibliographic database of directorate technical reports, which can be searched and/or browsed.

Aircraft Wiring Degradation Study

This technical report (DOT/FAA/AR-08/2) was published by the Federal Aviation Administration (FAA) Office of Aviation Research and Development in January 2008 and was written by Robert Bernstein, Mike Etheridge, Gary LaSalle, Roy McMahon, Jim Meiner, Noel Turner, Michael Walz and Cesar Gomez. The purpose of this initial research program was to evaluate the aging characteristics of three types of aircraft electrical wire: polyimide, poly trafluoroethylene/polyimide composite, and polyvinyl chloride/nylon. In addition, predictive models forthe aging of these wire types were developed. These wire types were chosen because of their widespread use in commercial aircraft and the amount of reported incidents concerning them. The factors that cause the wire insulation to degrade were examined and techniques to determine when a wire will no longer be capable of transfer of electrical current were evaluated. The results in this study provided a platform to evaluate existing and new test methods that could be used to monitor the aging of wire in aircraft. The results found were similar to the aging samples found from the Aging Transport Systems Rulemaking Advisory Committee Intrusive Inspection Report. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Alliant Techsystems (ATK)

Alliant Techsystems is a leading designer, developer, and manufacturer of space and strategic propulsion systems, spacecraft and aircraft structures, munitions propellants, tactical missile propulsion systems and warheads. The ATK web site provides corporate information, including financial and shareholder details, a searchable index of recent press releases and a description of the Company's main business activities.

Analytical Modeling of ASTM Lap Shear Adhesive Specimens : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/130, by Charles Yang, John S Tomblin, and Guan, Zhidong, dated February 2003. An analytical model was developed to predict the stress distribution within the specimen specified in ASTM D 3165 “Strength Properties of Adhesives in Shear by Tension Loading of Single-Lap-Joint Laminated Assemblies.” The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Application of Damage Tolerance Principles for Improved Airworthiness of Rotorcraft

This web site provides access to a NATO Research and Technology Organization (RTO) report titled: Application of Damage Tolerance Principles for Improved Airworthiness of Rotorcraft, RTO-Mp-024, February 2000. The report contains papers presented at the Applied Vehicle Technology Panel (AVT) Specialists Meeting, held in Corfu, Greece, 21-22 April 1999. The metting addressed issues associated with aging systems, and in particular with the application of damage tolerance principles for improved airworthiness of rotorcraft. The papers were grouped into three sessions covering the following: materials data and crack growth models for damage tolerance approaches to helicopter structures; design application of DT principles; and operator experience and certification issues. The citation and abstract information is in HTML format, and the full text is available online in PDF format (23 Mbytes).

Applications of Fracture Mechanics to the Durability of Bonded Composite Joints

This final report (DOT/FAA/AR-97/56) was published by the Federal Aviation Administration (FAA) in May 1998, and was written by W. Steven Johnson et al. This report covers an effort which focused on using fracture mechanics to evaluate the Mode I fracture and fatigue properties of several adhesively bonded aerospace material systems. Results are discussed with respect to their relevance and applicability to bonded joint design. Key results include the identification of significant degradation in some varieties of bonded joints subjected to long-term isothermal exposure under hot/wet conditions [extracted from FAA abstract]. This is a PDF file [60 pages, 1.03Mb] so Adobe Acrobat software will be required in order to read it.

Applied Aerospace Structures Corporation

This is the web site of the Applied Aerospace Structures Corporation (AASC), a US company involved in the manufacture of lightweight composites and metallic structures and components for space and aircraft applications. The focus of the company is on creating high strength lightweight structures for aircraft and spacecraft. The site gives some background information about the company and also gives details about its products and services. These are divided into five main product areas - space, aircraft, precision structures, ground systems, engineering and testing. It is possible to view the full text of the company's press releases.

ASME Digital Library

This service allows you to search ASME (American Society of Mechanical Engineers) journal articles back to 1985 and conference proceedings back to 2002. Search results provide bibliographic information and an abstract, with the option to purchase the full text, if you are not a subscriber.

Assessment of Probabilistic Certification Methodology for Composite Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/74, by Han-Pin Kan, dated January 2001. A sensitivity study has been conducted to assess the currently available probabilistic structural analysis methods. The influence of the distribution parameters on the probability of failure was investigated analytically. The significant parameters that have an impact on development of probabilistic certification procedures were identified. The technical gaps which need to be filled for probabilistic certification of composite structures were discussed. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Aurora Flight Sciences

Aurora specialized in the design and production of Unmanned Aerial Vehicles (UAVs). Aurora's business groups focus on the design, development, manufacturing and operaion of a high altitude long endurance UAVs for a wode range of tactical and science / research applications. The web site contains company information, including a brief history, and provides descriptions of various Aurora products and programmes. These include HALE designs such as Perseus A and Perseus B, Theseus A and Theseus B; the development of UAVs to fly in the Martian environment, MarsFlyer, ARES, and two High Altitude Deployment Demonstrators (HADD1 and HADD2); as well as tactical systems including GoldenEye-50, GoldenEye-100, Excalibur, Hunter II, and Orion. The site also provides access to a media collection for downloading including press releases and fact sheets. The site also gives employment and contact details.

Best Practice in Adhesive Bonded Structures and Repairs

This technical report (DOT/FAA/AR-TN06/57) was produced by the Federal Aviation Administration (FAA) in April 2007 and was written by Max Davis and John Tomblin. The opinions expressed in this technical note were presented at the Federal Aviation Administration (FAA) Bonded Structures workshop in 2004. The FAA, realizing their value, commissioned a written record of these observations and recommendations. These observations and recommendations represent the experiences, some anecdotal, in the application and maintenance of bonded structures on one group. This document does not represent a comprehensive survey and analysis of the failures or best corrective actions for bonded structures, but data that resulted from real-world applications and experience with disbands and other adhesive failures in structural applications. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Bolted/Bonded Joints in Polymeric Composites

This is Research and Technology Organization(RTO) AGARD Conference Proceedings, AGARD-CP-590, dated January 1997. This proceedings was sponsored by the Advisory Group for Aerospace Research and Development. The objective of this Meeting was to examine the state of the art in joining polymeric composites, to consider the relative merits of the various methods and to highlight gaps in the technology which should be addressed. The papers presented cover a number of aspects concerning the application of adhesively bonded and mechanically fastened joints in the analysis, design, manufacturing, and repair of fiber-polymer composites. The focus is on aerospace rather than commercial products. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text of the document (88 Mb) can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Calculating Polymer Flammability From Molar Group Contributions : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/31, by Richard Walters and Richard E. Lyon, dated September 2001. Hundreds of polymers of known chemical composition have been tested to date, providing over 40 different empirical molar group contributions to the heat release capacity. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Characterisation of Fibre Reinforced Titanium Matrix Composites

This is Research and Technology Organization (RTO) AGARD-R-796, dated February 1994. The combination of stiffness, strength, and high temperature resistance provided by fiber reinforced titanium matrix composites offers major benefits for aircraft engine and airframe applications, where these materials could be used to reduce weight or improve performance. This workshop on the subject of characterization of titanium composites was intended to provide a forum for the exchange of information in this important area. Characterization in this case refers to the understanding of the behavior of the composites as it relates to the ability to predict their performance in real-life applications. It covers various topics that include mechanical test techniques, NDE methods, life prediction models, and other factors that will affect the level of confidence with which these relatively new materials will be accepted for application. For individual titles, see N94-36650 through N94-36670. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (79.74MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Characterization and Structural Behavior of Braided Composites

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA/AR-08/52 written by Ajit Kelkar and John Whitcomb dated January 2009. The growing interest in small business jets in the general aviation industry is the motivation of this present research. The major objective in the small business jet industry is to reduce costs while keeping takeoff weights below 12,500 lb (5670 kg), which is a requirement of the Federal Aviation Administration. The overall objective of this research is the performance evaluation and modeling of biaxial braided composites manufactured using vacuum-assisted resin transfer molding. Biaxial braided composites with different braid angles were manufactured using carbon braids and two different resin systems (vinyl ester and epoxy). Static tension and tension-tension fatigue tests were performed. It was concluded that the Sigmoidal function accurately represents the stress-fatigue life curve of braided composites. It was observed that the braid angle has great influence on mechanical properties but has less effect on endurance limits. The variation in ultimate tensile strength between the specimens caused large scatter in the fatigue data. A new approach based on statistical analysis techniques for conducting fatigue tests is recommended. A new approach was developed to model stiffness degradation curves. This model was proven to be very efficient in all the three major stages of stiffness degradation. A computational micromechanics strategy was developed to model 2x2 braids. Since the tow cross-section along the towpath is not uniform, direct finite element mesh generation for the model is difficult. A mapping technique was developed to generate the finite element mesh for various 2x2 biaxial braids from the previously developed mesh for the twill weave. This resulted in substantial time savings. By exploiting symmetry operations such as mirroring, rotation, or a combination of the two, the analysis region was greatly reduced. The analyses for different braids have shown that the peak stresses in the tow mainly occur at the undulating region and along the edges of the tow. Stress distribution in braids was also compared with those in equivalent laminates. A considerable volume of the tow (10%-45% for the range of parameters studied) had stresses larger than an equivalent lamina. The severity of stresses in a braidas compared to those in an equivalent lamina depends upon braid geometric parameters. Braid angle changes the stress distribution in the tow considerably. The severity of peak stresses seems to be increasing linearly with an increase in waviness ratio. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Characterization of Compressive Creep Behavior of Oxide/Oxide Composite with Monazite Coating at Elevated Temperature

This is the full text of a Master's thesis by Second Lieutenant Patrick R. Jackson, USAF, AFIT/GAE/ENY/06-M17, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2006. The compressive creep behavior of a N610/monazite/alumina composite was investigated in this work. The composite consists of a porous alumina matrix reinforced with NextelTM610 fibers coated with monazite in a symmetric cross-ply (0 deg/90 deg/0 deg/90 deg)s orientation. Compressive stress-strain behavior was investigated as well. The addition of monazite coating resulted in ~ 35% loss in compressive strength at 900 deg C and in ~45% loss in compressive strength at 1100 deg C. Compressive creep behavior was examined at 900 and 1100 deg C for creep stresses ranging from 50 to 95 MPa. Primary and secondary creep regimes were observed at both temperatures. Minimum creep rate was reached in all tests. At 900 ?C both monazite containing and control specimens produced creep strains ≤ 0.05. Conversely, at 1100 deg C creep strains were significant, approaching 9%, with monazite containing specimens accumulating larger creep strains at a given stress than the control samples. Creep strain rates were on the order of 10-7 s-1. Creep run-out, defined as 100 h at creep stress, was achieved in all tests. The residual strength and modulus of specimens that achieved run-out at 1100 deg C were characterized. Composite microstructure, as well as damage and failure mechanisms were investigated. Furthermore, effects of variation in microstructure on mechanical response were examined. While differences in processing and consequently the composite microstructure did not have a significant effect on tensile response of the CMC, effects on the compressive properties were dramatic. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Characterization of In-Plane, Shear-Loaded Adhesive Lap Joints - Experiments and Analysis : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-03/21, by John Tomblin, Waruna Seneviratne, Hyonny Kim and Jungmin Lee, dated May 2003. In this experimental investigation, failure strengths of in-plane, shear-loaded bonded joints were compared with analytical predictions of the Shear-Loaded Bonded Joint (SLBJ) theory. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Chris Heintz Design College

This web site brings together a number of full text articles that deal with various aspects of light aircraft design. Fourteen of the articles were published in the Experimental Aircraft Association's (EAA) Light Plane World or Experimenter publications. The topics covered include: light aircraft materials and their properties; flight and performance testing; riveted joints; airfoils; pitch, stability and control; control surfaces; determining weight and balance; and STOL aircraft design. The author, Chris Heintz, is resonsible for the design of a range of kit aircraft for the Zenith Aircraft Company.

Comparative Testing to Assess the Equivalence of CEN and ASTM Test Methods for Composite Materials.

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA/AR-04/50 by Daniel Adams dated February 2005. An initial assessment of equivalence between CEN and ASTM test methods was performed and documented in DOT/FAA/AR-04/24. That study recommended additional test data to assess test method equivalency for four types of tests. The results of the four types of tests are contained in this report. First, lamina compression testing was performed to investigate the effects of different specimen gage lengths between either the prEN 2850–B and SACMA SRM 1 test methods (5 and 4.75 mm, respectively) and the AMS 2980 and 3970 specifications (12.5 mm). The results suggest that the gage length difference investigated (4.75 mm versus 12.5 mm) can produce significant differences in the apparent lamina compression strength. However, this difference is believed to be due to specimen buckling that occurred for specimens with longer gage lengths, which is an unacceptable failure mode. Second, laminate compression tests were performed to investigate the method used to load the test fixture. The results suggest that the end-loaded test fixture shown in the SACMA SRM 3-94 test method is not suitable without additional clamping force applied near the specimen ends to prevent end-brooming failures. However, it is expected that through the modification of the existing test fixture to include more clamping bolts, an end-loaded test procedure can be produced that is capable of producing equivalent test results to the shear-loaded ASTM D 6484 test method. Third, in-plane shear testing was performed using ±45° type composite laminates loaded in tension to investigate laminate thickness effects. Test results showed that the 0.2% offset shear strength, the 5% shear strength measures, and the shear moduli calculated following the prEN 6031/AMS and ASTM D 3518 test methods are in good agreement and independent of laminate thickness over the thickness ranges specified in the two test methods. Thus, these two shear test methods are believed to produce equivalent results when testing either tape or fabric laminates. Finally, constituent content was determined to compare procedures EN 2564 Method A and ASTM D 3171 Method I procedure B. The results showed that EN 2564 Method A and ASTM D 3171 Method I procedure B produced similar values for both the fiber volume percent and the matrix volume percent. However, the increased accuracy of the ASTM D 3171 test method due to the larger specimen size and the greater weighing accuracy produced more realistic values of void volume fraction than the EN 2564 Method A test method. [Taken from abstract]. The full text of the report is available in PDF format from the online catalogue of the FAA William J Hughes Technical Center Library.

Comparative Testing to Assess the Equivalence of CEN and ASTM Test Methods for Composite Materials

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA AR-04/50 by Daniel Adams dated February 2005. An initial assessment of equivalence between CEN and ASTM test methods was performed and documented in DOT/FAA/AR-04/24. That study recommended additional test data to assess test method equivalency for four types of tests. The results of the four types of tests are contained in this report. First, lamina compression testing was performed to investigate the effects of different specimen gage lengths between either the prEN 2850–B and SACMA SRM 1 test methods (5 and 4.75 mm, respectively) and the AMS 2980 and 3970 specifications (12.5 mm). The results suggest that the gage length difference investigated (4.75 mm versus 12.5 mm) can produce significant differences in the apparent lamina compression strength. However, this difference is believed to be due to specimen buckling that occurred for specimens with longer gage lengths, which is an unacceptable failure mode. Second, laminate compression tests were performed to investigate the method used to load the test fixture. The results suggest that the end-loaded test fixture shown in the SACMA SRM 3-94 test method is not suitable without additional clamping force applied near the specimen ends to prevent end-brooming failures. However, it is expected that through the modification of the existing test fixture to include more clamping bolts, an end-loaded test procedure can be produced that is capable of producing equivalent test results to the shear-loaded ASTM D 6484 test method. Third, in-plane shear testing was performed using ±45° type composite laminates loaded in tension to investigate laminate thickness effects. Test results showed that the 0.2% offset shear strength, the 5% shear strength measures, and the shear moduli calculated following the prEN 6031/AMS and ASTM D 3518 test methods are in good agreement and independent of laminate thickness over the thickness ranges specified in the two test methods. Thus, these two shear test methods are believed to produce equivalent results when testing either tape or fabric laminates. Finally, constituent content was determined to compare procedures EN 2564 Method A and ASTM D 3171 Method I procedure B. The results showed that EN 2564 Method A and ASTM D 3171 Method I procedure B produced similar values for both the fiber volume percent and the matrix volume percent. However, the increased accuracy of the ASTM D 3171 test method due to the larger specimen size and the greater weighing accuracy produced more realistic values of void volume fraction than the EN 2564 Method A test method. [Taken from abstract]. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

Composite Landing Gear Components made with Resin Transfer Moulding (RTM)

This technical report (NLR-TP-2004-122) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2004 and was written by H. G. S. J. Thuis, J. F. M. Wiggenraad and H. P. J. de Vries. Composite materials are used in aircraft structures because of weight and cost benefits. These components are mostly shell structures. For helicopters, weight savings are even more important than for fixed wing aircraft. Hence, components other than shell structures, which can be made of composite materials, are being considered. A project is described, focused on the development of the fabrication technology for selected composite landing gear components. The project is being carried out by a Dutch consortium: landing gear supplier SPa&vs, aerospace research centre NLR, software company MSC.Software and preform supplier Eurocarbon. The Composites Group of Twente University is involved with software developments. The components considered are a pair of torque links and a trailing arm, baselined on the NH-90 helicopter. The torque links were fabricated and tested successfully. A trailing arm has been fabricated with the over-braiding technique, and will be tested later in 2004. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Composite Repair of Military Aircraft Structures

This is Research and Technology Organization (RTO) AGARD-CP-550 , dated January 1995. The AGARD Structures and Materials Panel held a specialists' Meeting to address composite repair of military aircraft. The meeting focused on two main areas, repair of metal structures using composite patches and repair of composite structures using composite or metal patches. The work presented had direct application to the maintenance and support of military aircraft. Repair of military aircraft provides both a means to extend the useful life of the airframe beyond the original design life and a method to maintain military readiness by returning damaged aircraft to service. For individual titles, see N95-27505 through N95-27528. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (58.6MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Crashworthiness Research at NLR (1990 - 2003)

This technical report (NLR-TP-2003-317) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2004 and was written by J.F.M. Wiggenraad. In the past decade, NLR has carried out research and development activities in several areas to improve the crashworthiness of future helicopters and fixed-wing aircraft. The crashworthiness of composite helicopters has been studied, and components for the NH90 helicopter have been developed. In several international collaborations, NLR has been involved in the development of crashworthy composite and metal aircraft structures. The potential for further improvements is expected to incite continuing collaborative research and development. [Taken from abstract]. The full text is available as a PDF file.

Creep Behavior of an Oxide/Oxide Composite with Monazite Coating at Elevated Temperatures

This is the full text of a Master's thesis by First Lieutenant Sean S. Musil, USAF, AFIT/GAE/ENY/05-M14, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2005. This study focuses on experimental investigation of stress-rupture behavior (creep response) of an oxide/oxide composite in a cross-ply (0/90) lay-up at elevated temperature. The test material, Nextel 610/monazite/alumina composite, employs monazite, an oxidation-resistant interfacial coating designed to improve performance at elevated temperatures. The experimental program included monotonic tensile tests to failure and creep-rupture tests at elevated temperatures. Tensile tests served to establish an ultimate tensile strength (UTS) for the material. The ensuing creep-rupture tests involved stress levels at varying percentages of the UTS. Stress-rupture curves at 900 and 1100 degrees C were established. A family of creep curves for various constant stress levels at 900 and 1100 degrees C was produced. Composite microstructure, as well as damage and failure mechanisms were also investigated. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Creep-Rupture and Fatigue Behaviors of Notched Oxide/Oxide Ceramic Matrix Composite at Elevated Temperature

This is the full text of a Master's thesis by Captain Mark A. Sullivan, USAF, AFIT/GAE/ENY/06-M30, which was presented to the Faculty Department of Aeronautical and Astronautical Engineering of Air University's Air Force Institute of Technology (AFIT), in March 2006. Oxide/oxide composites are being considered for use in high temperature aerospace applications where their inherent resistance to oxidation provides for better long life properties at high temperature than most other ceramic matrix composites (CMCs). One promising oxide/oxide CMC is Nextel 720/A (N720/A) which uses an 8-harness satin weave (8HSW) of Nextel 720 fibers embedded in a porous alumina matrix. Possible aerospace applications for N720/A will likely require inserting holes into the material for mounting and cooling purposes. The notch characteristics must be understood to ensure designs using the material are sufficient for the desired application. This research effort examined the fatigue and creep-rupture characteristics of N720/A with a 0 /90 fiber orientation and notch to width ratio (2a/w) of 0.33. Specifically, 12.0 mm wide rectangular specimens with a 4.0 mm center hole were subjected to axial fatigue and creep-rupture loads in 1200 C laboratory air. Monotonic tensile tests at 1200 C were performed on unnotched specimens to provide a baseline for comparison with previous research. Fracture surfaces were examined under microscope to observe microstructure and damage mechanisms. Comparisons to previous unnotched research results at 1200 C show N720/A to be primarily insensitive to 0.33 notch ratios. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Creep-Rupture Behavior of a Woven Ceramic Matrix Composite at Elevated Temperatures in a Humid Environment

This is the full text of a Master's thesis by Second Lieutenant Jennifer L. Ryba, USAF, AFIT/GMS/ENY/06-M02, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2006. This study focused on moisture and temperature effects on the embrittlement and stress-rupture life of the SiC/SiC CMC Syl-iBN/BN/SiC. The Syl-iBN/BN/SiC is composed of Sylramic fibers with an in-situ layer of boron nitride (Syl-iBN), boron nitride interphase (BN), and SiC matrix. Stress rupture tests and monotonic tests were performed on the specimens. Tests were conducted under 100% humidity and laboratory air environments at three temperatures, 450 C, 750 C, and 950 C. These temperatures were chosen because they fall below the intermediate range, within the range, and above the range, respectively. This study found that while this CMC does experience embrittlement at intermediate temperatures, it also occurs at temperatures above the intermediate range. Scanning Electron Microscopy (SEM) analysis showed the embrittlement and pesting in the specimens increased with time, temperature, and moisture exposure, leading to premature failure. An analysis of the data confirmed that with increase in temperature and exposure to moisture, the stress-rupture life of the Syl-iBN/BN/SiC was considerably shortened. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Cyclic Creep and Recovery Behavior of Nextel(Trademark) 720/Alumina Ceramic Matrix Composite at 1200deg C in Air and in Steam Environments

This is the full text of a thesis written by Bridgett Whiting which was presented to the Air Force Institute iof Technology, Wright Patterson Air Force Base, Ohio in September 2007. The cyclic creep and recovery behaviors of the N720/Al2O3 composite were investigated in this research. The ceramic matrix composite (CMC) contains a porous alumina matrix with laminated, woven mullite/alumina (NextelTM 720) fibers. The composite does not have an interface between the fiber and matrix. The CMC relies on the porous nature for flaw tolerance. The objective of this study the influences of monotonic creep and cyclic creep loading histories on the creep lifetime, creep strain rate, accumulated creep strain as well as on the recovery of creep strain at near zero stress. The cyclic creep and recovery tests were performed at 1200 ?C with maximum creep stress levels of 100 and 125 MPa in air and in steam. The creep and recovery periods were ranged from 3 min to 30 h. The laboratory air tests significantly exceeded the life of the monotonic creep tests. Introduction of intermittent periods of unloading and recovery at near zero tress into the monotonic creep history resulted in one to two orders of magnitude improvement in the creep life and rate. The presence of steam greatly reduced the performance of the material. The results in steam were similar to those of the monotonic creep. The omposite microstructure, damage and failure mechanisms were also explored. [Taken from abstract]. This is in PDF format so Adobe Acrobat software is required in order to read it.

Damage Resistance and Tolerance of Composite Sandwich Panels—Scaling Effects

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA AR-03/75 by John S. Tomblin, The impact responses and the damage states in flat sandwich panels with thin facesheets are known to be dependent on the diameter of the spherical steel impactor. The residual strength of impact-damaged sandwich panels under static in-plane compressive loads was dependent on the nature of the damage state. The coupon sizes used in these investigations were relatively small, and the finite size effects may be embedded in the observed trends. The effects of scaling the planar dimensions of sandwich specimens on the damage resistance were studied by conducting experiments and finite element analysis. The impact force and the damage size were observed to decrease as both the planar dimensions were increased. These effects were negligible when only a single dimension was scaled. The off-center impacts indicated that for a given energy level, the impacts occurring closer to the boundary supports were more severe compared to those farther away from the boundaries. The impacts on sandwich specimens supported by a rigid base proved to be the most severe case in terms of the impact behavior and the resulting damage metrics. The parametric study conducted using the finite element model confirmed the observed experimental trends and further indicated that the damage formation is the dominant energy dissipation mechanism when the ratio of the impactor mass to that of the target is greater than 2 and vibrational energy transfer is dominant for ratios less than 2. The effects of the ratio of specimen width to planar damage size on the compressive residual strength and failure modes were investigated for two sandwich configurations. A subsurface damage state was considered for the study and inflicted using a 3�� diameter impactor. The scaling effects were characterized in terms of the residual strength and strain distributions in the vicinity of the damage region. The latter was measured using a photogrammetry method. The residual strength was found to increase by 12% when the ratio of the specimen size to damage size was increased from 4.6 to 12.4 for sandwich specimens with two-ply facesheets. No trends were, however, observed for sandwich specimens with four-ply facesheets. The strain and displacement distributions indicated bending of the facesheet within the damage region leading to a strain concentration-driven failure mode resembling an open hole configuration for the two-ply facesheet sandwich panels. The 6.5�� wide sandwich specimens with four-ply facesheets failed by global buckling initiated by an unstable dimple propagation. For wider specimens, there was a dimple growth-arrest mechanism that lead to eventual facesheet fracture. The increase of the specimen height resulted in a slight decrease in residual strength. This study showed that the results obtained from small specimens are valid as far as the compressive residual strength obtained experimentally. However, larger specimen sizes will sustain less impact damage for equivalent impact energy levels. Thus, it is valid to test smaller panels if the damage is simulated correctly from the larger specimens. The effects of specimen height were not investigated but could effect the buckling response of the panels. The impact responses and the damage states in flat sandwich panels with thin facesheets are known to be dependent on the diameter of the spherical steel impactor. The residual strength of impact-damaged sandwich panels under static in-plane compressive loads was dependent on the nature of the damage state. The coupon sizes used in these investigations were relatively small, and the finite size effects may be embedded in the observed trends. The effects of scaling the planar dimensions of sandwich specimens on the damage resistance were studied by conducting experiments and finite element analysis. The impact force and the damage size were observed to decrease as both the planar dimensions were increased. These effects were negligible when only a single dimension was scaled. The off-center impacts indicated that for a given energy level, the impacts occurring closer to the boundary supports were more severe compared to those farther away from the boundaries. The impacts on sandwich specimens supported by a rigid base proved to be the most severe case in terms of the impact behavior and the resulting damage metrics. The parametric study conducted using the finite element model confirmed the observed experimental trends and further indicated that the damage formation is the dominant energy dissipation mechanism when the ratio of the impactor mass to that of the target is greater than 2 and vibrational energy transfer is dominant for ratios less than 2. The effects of the ratio of specimen width to planar damage size on the compressive residual strength and failure modes were investigated for two sandwich configurations. A subsurface damage state was considered for the study and inflicted using a 3�� diameter impactor. The scaling effects were characterized in terms of the residual strength and strain distributions in the vicinity of the damage region. The latter was measured using a photogrammetry method. The residual strength was found to increase by 12% when the ratio of the specimen size to damage size was increased from 4.6 to 12.4 for sandwich specimens with two-ply facesheets. No trends were, however, observed for sandwich specimens with four-ply facesheets. The strain and displacement distributions indicated bending of the facesheet within the damage region leading to a strain concentration-driven failure mode resembling an open hole configuration for the two-ply facesheet sandwich panels. The 6.5�� wide sandwich specimens with four-ply facesheets failed by global buckling initiated by an unstable dimple propagation. For wider specimens, there was a dimple growth-arrest mechanism that lead to eventual facesheet fracture. The increase of the specimen height resulted in a slight decrease in residual strength. This study showed that the results obtained from small specimens are valid as far as the compressive residual strength obtained experimentally. However, larger specimen sizes will sustain less impact damage for equivalent impact energy levels. Thus, it is valid to test smaller panels if the damage is simulated correctly from the larger specimens. The effects of specimen height were not investigated but could effect the buckling response of the panels. [Taken from abstract]. The full text of the report is available in PDF format from the online catalogue of the FAA William J Hughes Technical Center Library.

Damage Resistance Characterization of Sandwich Composites Using Response Surfaces : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/71, by T. E. Lacy et al, dated March 2002. The influence of material configuration and impact parameters on the damage resistance characteristics of sandwich composites comprised of carbon-epoxy woven fabric face sheets and Nomex honeycomb cores were investigated using empirically based response surfaces. The effects of impact energy and velocity on damage formation were also examined. The full text of the report is available in PDF format

Damage Tolerance and Durability of Selectively Stitched, Stiffened Panels : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-03/46, by H. Thomas Hahn, Jenn Ming Yang, Sung Suh, Tan Yi, and Guocai Wu, dated June 2003. The goal of this project was to investigate the effectiveness of selective stitching on the damage tolerance and durability of a stiffened structural element applicable to air transportation systems. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Damage Tolerance Characterization of Sandwich Composites Using Response Surfaces : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/101, by T.E.Lacy, I.K.Samarah, and J.S.Tomblin, dated November 2002. The influence of material configuration and impact parameters on the damage tolerance characteristics of sandwich composites comprised of carbon-epoxy woven fabric facesheets and Nomex honeycomb cores was investigated using empirically based response surfaces. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Damage Tolerance of Composite Sandwich Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/91, by R. Clifton Moody and Anthony J. Vizzini, dated January 2000. The report focuses on the modeling aspects of damage tolerances of thin-gage composite sandwich structures. full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Delphion Patent Search Form

This site allows you to search for United States patents, European patents and patent applications, Patent Cooperation Treaty (PCT) application data from the World Intellectual Property Office, the Patent Abstracts of Japan and INPADOC data. The service can be searched in several different ways, including patent number, US classification and Boolean keyword search. It is possible to view to the bibliographic information of granted US patents free of charge, all other services are payable. You will need to register to use this service, which is free of charge.

Design and Analysis of Stiffened Composite Panels for Damage Resistance and Tolerance

This technical report (NLR-TP-2002-193) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2002 and was written by J. F. M. Wiggenraad, E. S. Greenhalg and R. Olsson. This paper describes a joint Dutch-Swedish-British programme for reducing the effect of impact on skin-stringer composite panels. An optimisation code for the design of such panels was extended with the capability to optimise panels for "damage resistance". Hereto, a damage initiation constraint was implemented. Three different panels were designed, fabricated and tested: one baseline configuration with previous design methodology, and two "damage resistant" configurations, according to the new design capability. The new criterion was found accurate for impacts at stiffener foot locations, and conservative for impacts at midbay locations between stiffeners. Although the panels were not fully damage resistant for 35 J impacts as intended, the results were close, and clearly superior to the baseline configuration. The penalty for increased damage resistance is increased panel weight, while the advantages are savings of fabrication and maintenance costs. Delamination growth after impact was simulated with a moving mesh FE-model. The simulations demonstrated a strong interaction between global skin buckling and the location and growth of delaminations. Delamination growth is promoted by location at a buckle crest and retarded by location at a node line. [Taken from abstract]. The full text is available as a PDF file.

Design and Testing of a Composite Bird Strike Resistant Leading Edge

This technical report (NLR-TP-2003-054) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2003 and was written by L.C. Ubels, A.F. Johnson, J.P. Gallard and M. Sunaric. The paper describes several innovative designs for a bird strike resistant, composite leading edge for a Horizontal Tail Plane of a transport aircraft. These designs are based on a novel application of composite materials with high energy-absorbing characteristics: the tensor-skin concept. This paper describes the development of this energy-absorbing concept and its application to an impact resistant aircraft structure. The design philosophy, the fabrication and test of the first prototypes are discussed. Three improved leading edge structures with different energy-absorbing tensor concepts were manufactured. Bird-strike tests on these leading edges with a 4 lb synthetic bird at impact velocities around 100 m/s were performed. Finite element models were developed to simulate the unfolding of the tensor ply. Before each test was carried out, pre-test bird impact simulations were used to determine the impact test parameters and to predict the dynamic behaviour and failure mode of the structure. [Taken from abstract]. The full text is available as a PDF file.

Design, Manufacturing, and Performance of Stitched Stiffened Composite Panels With and Without Impact Damage : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/111, by H. Thomas Hahn, Jenn Ming Yang, Sung S. Suh, and Nanlin Han, dated October 2002. The goal of this project was to develop the knowledge base required for certification of composite structures in air transportation systems in the form of a design-manufacturing-performance relationship. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Determination of Temperature / Moisture Sensitive Composite Properties : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/40, by John S.Tomblin, Lamia Salah and Yeow C. Ng, dated September 2001. An investigation of the temperature and moisture sensitive composite properties was conducted. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Development and Evaluation of a Hard Patch Repair Method for Composite Stiffened Wing Panels

This technical report (NLR-TP-99033) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 1999 and was written by A. L. P. J. Michielsen, L. C. Ubels and F. Baas. Within the framework of European program Euclid RTP 3.1 'Repair methodology on-aircraft', repair methods were developed for composite structures in fighter aircraft. The National Aerospace Laboratory (NLR), in co-operation with Hogeschool Haarlem, developed a hard patch repair method. In this approach, the damaged zone is completely removed and substituted with a pre-cured patch, flush mounted on a flange that is bolted (and bonded) to the interior skin surface. Two types of repair were studied, which involve either a skin bay between two stiffeners or a stiffener zone. For each case, the method was developed on small specimens. After a feasible solution was established in terms of structural performance, its applicability in structures with one-sided access only was validated on a full scale panel structure. It was concluded that the repair method is suited to be performed "on aircraft", because it is possible to perform the repair with access from the outside only, by field and maintenance personnel and with relatively simple tools. The method used results in an aerodynamic smooth surface: the patches are flush at the outer skin side apart from a number of rivets. The fatigue and residual strength tests showed that the requirements were fulfilled with respect to repair performance for all configurations studied in the development phase and for the configuration tested in the validation phase. The method is thought to be promising for the future but is still not developed to the extent, that the application in service is possible without further improvements in repair skills, adaptation of repair materials and further validation tests. Additional research has to be performed if the method is selected to become a standard repair procedure for a specific aircraft. [Taken from abstract]. The full text is available as a PDF file.

Development and Evaluation of the V-Notched Rail Shear Test for Composite Laminates

This technical report (DOT/FAA/AR-03/63) was produced by the Aviation Research office of the Federal Aviation Administration (FAA) in September 2003 and was written by Daniel O. Adams, Joseph M. Moriarty, Adam M. Gallegos and Donald F. Adams. The V-notched rail shear test developed in this investigation appears to be well-suited for measuring the in-plane shear modulus and shear strength of unidirectional and multidirectional composite laminates. This test method incorporates attractive features from both the Iosipescu V-notched shear test and the standard two-rail shear test. The proposed V-notched specimen provides a larger gage section than the standard Iosipescu shear specimen and enhanced loading capability compared to either existing test method. Finite element analysis was used to evaluate several notched and tabbed rail shear specimen configurations. A 90° notch-angle V-notched specimen configuration, with notch depths that were 22.7 percent of the gage section height, was found to produce a desirable stress state in the gage section. Extensive rail shear testing was performed on a series of 16-ply carbon/epoxy laminates ranging from 0% ±45 plies to 100% ±45 plies consisting of [0]16, [0/90]4S, [(0/90)2/±45/0/90]S, [0/±45/90]2S, and [±45/90/±45/0/±45]S. Both a tabbed rectangular specimen and the V-notched specimen produced significantly higher shear strengths than the baseline rectangular specimen. The V-notched specimen configuration was selected over the tabbed rectangular configuration based on the higher shear strengths obtained, acceptable gage section failures produced, ease and economy of specimen preparation, and accuracy of shear modulus measurements. The proposed 90° notch-angle V-notched specimen was shown to produce accurate measures of shear modulus as predicted by finite element analysis. A new rail shear test fixture was developed to accommodate the V-notched rail shear specimen. [Taken from abstract] This is a PDF file, so Adobe Acrobat will be required in order to read it.

Development of a Composite Torque Link for Helicopter Landing Gear Applications

This technical report (NLR-TP-1999-026) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 1999 and was written by H.G.S.J. Thuis. In the framework of a composite landing gear technology programme, a composite torque link for helicopter landing gear applications was developed. The torque link was designed by finite element analysis and optimised for minimal weight. The torque link was fabricated by Resin Transfer Moulding (RTM) for which a tooling concept was developed. Static tests demonstrated the load carrying capabilities in undamaged and damaged condition of the torque link since all specimens failed beyond their Design Ultimate Load level. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Development of a Crashworthy Composite Fuselage Concept for a Commuter Aircraft

This technical report (NLR-TP-2001-108) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2001 and was written by J. F. M. Wiggenraad, D. Santoro, F. Lepage, C. Kindervater and H. Climent Manez. Within the framework of Brite-Euram programme CRASURV "Commercial Aircraft - Design for Crash Survivability", technology was developed for the design of composite air frames with respect to crashworthiness. The ultimate goal of the project was to develop computer codes for the simulation of the crash behaviour of composite fuselage structures. A significant part of the project consisted of the design, fabrication and droptesting of two representative composite fuselage sections, to generate the experimental data needed for the validation of the new code developments. The present paper gives an overview of the development, test and numerical analysis of one of the fuselage sections, a one-bay section representative of a commuter aircraft like the ATR-42/72. The fuselage section consists of the sub-floor structure, which is the major area that will be crushed during a potentially survivable crash. The structure failed in a mode which was not predicted. The deficiencies of the model were repaired and a post-test analysis gave satisfactory results. The project has resulted in improved simulation capabilities. However, it cannot be concluded that the state-of-the-art is such, that the behaviour of new composite structures can be predicted accurately in the near future. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Development of Generic Composite Box Structures with Prepeg Preforms and RTM

This technical report (NLR-TP-2002-019) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2002 and was written by H. P. J. de Vries. In the framework of a national technology program an one shot manufacturing process for closed composite box structures with accurate dimensions was developed. The goals of the program were the development of a manufacturing process which leads to net shaped box structures with a dimensional tolerance of ñ °®05 mm and the development of a reliable stringer/skin attachment without rivets. An RTM manufacturing concept with matched Invar tooling and a prepreg preform with stitched spars and stringers was developed. Three generic box structures were manufactured successfully. The dimensional tolerances of the boxes were within ñ °®05 mm and within ñ °®3?. [Taken from abstract]. The full text is available as a PDF file.

Effect of Environment on Creep Behavior of an Oxide/Oxide CFCC with 45 deg. Fiber Orientation

This is the full text of a Master's thesis by Ensign Gregory T. Siegert, USN, AFIT/GAE/ENY/06-J15, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in June 2006. Aerospace applications require materials capable of maintaining superior mechanical properties while operating at high temperatures and oxidizing environments. Nextel(trademark) 720/A (N720/A), an oxide/oxide ceramic matrix composite (CMC) with a porous alumina matrix was developed specifically to provide improved long-term properties and performance at 1200 deg C. This research evaluated the creep behavior of N720/A with a plus or minus 45 deg fiber orientation at 1200 deg C in: laboratory air, 100% steam, and 100% argon environments. Creep-rupture tests at the creep stress levels of: 45, 40, 35, and 15 MPa were conducted in each environment. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Effect of Hold Times on Fatigue Behavior of NEXTEL (trademark) 720/Alumina Ceramic Matrix Composite at 1200 deg C in Air and in Steam Environment

This is the full text of a Master's thesis by Captain John M. Mehrman, USAF, AFIT/GAE/ENY/06-M23, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2005. The aerospace field requires structural materials that can maintain superior mechanical properties while subjected to high temperatures and oxidizing environments. This research investigated the effect of hold times at maximum load on fatigue performance of a Nextel 720/Alumina ceramic matrix composite at 1200 C, explored the influence of environment on material response to cyclic loading with hold times at maximum load, and assessed the effects of loading history on material behavior and environmental durability. The N720/A composite relies on an oxide/oxide composition for inherent oxidation resistance and a porous matrix with no interphase between the fiber and matrix for damage tolerance. Mechanical testing results showed a significant decrease in material life and performance in a steam environment when compared to tests conducted in a laboratory air environment. Prior fatigue of specimens tested in the air environment resulted in an order of magnitude increase in creep life. Fracture surface observations with a Scanning Electron Microscope showed a correlation between an increase in fiber pull-out and increased time to failure. A qualitative spectral analysis indicated evidence of silicon species migration from the fiber to the matrix, especially in the steam environment. This may be the cause of the decreased creep performance of the material in the steam environment. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Effect of Temperature and Steam Environment on Fatigue Behavior of an Oxide-Oxide Continuous Fiber Ceramic Composite

This is the full text of a Master's thesis by First Lieutenant Chalene A. Eber, USAF, AFIT/GA/ENY/05-M09, which was presented to the Faculty Department of Aeronautical and Astronautical Engineering of Air University's Air Force Institute of Technology (AFIT), in March 2005. There is an ever-increasing need for materials that maintain high strength and fracture toughness at elevated temperatures and in complex environments. Advanced aerospace applications are motivating the development of composite materials that can meet demanding requirements. This research effort investigates mechanical behavior of an oxide-oxide continuous fiber ceramic composite (CFCC) consisting of a porous alumina matrix reinforced with mullite/alumina Nextel 720 fibers developed specifically for advanced aerospace applications. Tension-tension fatigue behavior of this CFCC was studied at 1200 and 1330 C in laboratory air and 100% steam environments. Fatigue resistance and retained strength properties were determined. Effects of environmental degradation was addressed in detail. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Effects of Environment on Creep Behavior of Two Oxide-Oxide Ceramic Matrix Composites at 1200 degs C

This is the full text of a Master's thesis by Captain Pavlos A. Koutsoukos, Hellenic Air Force, AFIT/GAE/ENY/06-S05, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in September 2005. Previous studies by the advisor and graduate students examined creep behavior of the Nextel720/Alumina CMC in air and in 100% steam environments at 1200 and 1330 deg C. Results showed that while this oxide/oxide system exhibits an exceptionally high fatigue limit at 1200 deg C it also experiences substantial strain accumulation under sustained loading conditions. Furthermore, these earlier investigations revealed a significant degrading effect of 100% steam environment on material performance under both static and cyclic loadings. The present effort will investigate creep rupture behavior of Nextel720/Alumina composite in the inert gas environment. In addition, creep rupture behavior of Nextel720/Aluminosilicate CMC will be investigated in both inert gas and in 100% steam environments. Combined with existing data, results of this research will fully reveal effects of progressively more oxidizing environment on creep resistance of these CMCs. In addition, degradation of Nextel720 fibers under load in oxidizing environments will be assessed. The study followed a systematic plan. Baseline tensile tests were performed to verify the at-temperature basic properties and to guide selection of the creep stress levels. Creep-rupture tests were carried out at different stress levels at 1200 deg C. In order to examine combined effects of temperature and exposure to oxidizing environment on the creep response, creep-rupture tests were performed in the inert gas and in 100% steam environments. For selected creep stress levels, creep tests in laboratory air were performed as well. As a result of this effort creep-rupture curves, as well as families of creep curves were established. Degradation of creep resistance due to increasing moisture exposure were assessed. Composite microstructure, as well as damage and failure mechanisms were examined. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Effects of Frequency and Environment on Fatigue Behavior of an Oxide-Oxide Ceramic Matrix Composite at 1200 Deg. C

This is the full text of a Master's thesis by Ensign Griffin Hetrick, USN, AFIT/GAE/ENY/06-J05, which was presented to the Faculty Department of Aeronautical and Astronautical Engineering of Air University's Air Force Institute of Technology (AFIT), in June 2006. Advances in aeronautical engineering in the 21st century depend upon materials that can perform well in extreme environments such as high temperatures and oxidizing conditions. Nextel(Trademark)720/Alumina (N720/A) is an oxide/oxide ceramic matrix composite with a porous alumina matrix that has been identified as a candidate material for such applications. This research investigated the effects of frequency on fatigue response of N720/A at 1200C in both air and steam environment. Prior investigation of this material by Eber [8] in 2005 studied fatigue behavior at 1200C in air and in steam environments at the frequency of 1.0 Hz. The current research focused on fatigue response at the frequencies of 0.1 Hz and 10 Hz. Results of mechanical testing showed a significant decrease in fatigue performance in steam versus air. Specimens tested at 0.1 Hz exhibited shorter fatigue lives and smaller strains at failure than those tested at 10 Hz. Scanning Electron Micrographs of specimen fracture surfaces revealed higher degrees of fiber pull-out and greater variation in fiber failure locations in specimens tested at 10 Hz, indicating a weakening of the fiber/matrix interface. Qualitative assessment using Energy Dispersive Spectroscopy showed correlations between frequency and amount of silicon species migration between fiber and matrix. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Effects of Surface Preparation on Long-Term Durability of Composite Adhesive Bonds : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/8, by Jason Bardis and Keith Kedward, dated April 2001. The report examines the long-term affects of surface preparation techniques for composite bonded joints. It focuses on the effects of peel plies and grit blasting on the fracture toughness and failure mode of adhesively bonded composites. The full text of the report is available in PDF format.

Effects of Surface Preparation on the Long-Term Durability of Adhesively Bonded Composite Joints

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA/AR-03/53 by Jason Bardis and Keith Kedward dated January 2004. The long-term durability of adhesively bonded composite joints is critical to modern aircraft structures, which are increasingly using bonding as an alternative to mechanical fastening. The effects of surface preparation for the adherends are critical to initial strength, long-term durability, fracture toughness, and failure modes of bonded joints. In this study, several potential factors are evaluated, with focus on the following: 1. Effects of possible chemical contamination from release fabrics, release films, and peel plies during adherend cure. 2. Chemical and mechanical effects of abrasion on the fracture toughness and failure mode. 3. Characterization of paste and film adhesives using mechanical test methods. There are several standard test methods to evaluate specimen fracture, but the majority concentrate on bonded metals and interlaminar composite fracture. Testing is concentrated on mode I tests. A custom double cantilever beam specimen was devised and used, and two forms of a wedge crack test (traveling and static) were also used. Additionally, mode II single lap shear tests were run to compare to the mode I tests. Nondestructive testing included X-ray photography of crack fronts, energy dispersive spectroscopy and X-ray photoelectron spectroscopy surface chemistry analyses, and scanning electron microscope imaging of prepared surfaces. All mode I test methods tended to be in agreement in the ranking of different surface preparation methods. Test results showed that release agents deposited on adherend surfaces during their cure cycle prevented proper adhesion. While mechanical abrasion did improve their fracture toughness and lower their contamination greatly, the test values did not reach the levels of samples that were not contaminated before bonding; therefore, the interfacial modes of failure did not always change to desirable modes. [Taken from abstract]. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

Embedded Piezoelectric Sensors and Actuators for Control of Active Composite Structures

This paper was given at the 6th Dynamics and Control of Systems and Structures in Space (DCSSS) conference by G. Sala, M. Olivier, P. Bettini and D. Sciacovelli. The conference took place in July 2004 at Riomaggiore, Italy. This paper illustrates the results obtained at Composite and Smart Materials Laboratory of Politecnico di Milano by an industrial-academic team, that designed, developed and produced a technological demonstrator for space structures, embedding piezoelectric sensors/actuators and optical sensors. The selection of optical sensors, the use of PZT material, and their embedding techniques into composite laminates are described in the following, as well as embedding a sensing/actuation pack, i.e. robust and easy to handle sensor\actuator unit, in the demonstrator structure. Finally, actuation/sensing capability have been tested to validate both the embedded sensors/actuators and the technologies adopted for the manufacturing process. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Enhanced Reliability Prediction Methodology for Impact Damaged Composite Structures

This final report (DOT/FAA/AR-97/79) was published by the Federal Aviation Administration (FAA) in October 1998, and was written by H P. Kan. A thorough review of the existing impact test data and analysis methods was conducted and the results were used to identify a reliability prediction methodology for further development. The integrated residual strength/reliability method developed by Northrop Grumman under a Navy/Federal Aviation Administration sponsored program was selected and modified. A structural damage tolerance evaluation was conducted using the modified model and the results compared to those obtained from the existing model [extracted from FAA abstract]. This is a PDF file [90 pages, 437Kb] so Adobe Acrobat software will be required in order to read it.

ESDU International

ESDU create and maintain over 22 series of validated engineering design data covering structural, mechanical, aeronautical and chemical process engineering. They are produced by committees of independent experts who ensure that each Data Item is a sound technical document which presents a clear explanation of the recommended approach. The bulletins that each committee produce are available for viewing using Adobe Acrobat. Subscribers to the service can access the full text of all data items. Non subscribers may view abstracts of each Data Item. An FAQ, glossary and technical notes are also available.

European Aeronautics Science Network (EASN)

This is a three year funded project which aims to bring the European universities with aeronautics activities into an integrated network, operating in parallel with industry and the national research establishments. EASN has a Steering Committee representing partners to oversee the activities of the network, as well as a number of Interest Groups (IGs) addressing various thematic issues. There are 10 interest areas covering Flight Physics, Aerostructures, Propulsion, Aircraft Avionics Systems and Equipment, Flight Mechanics, Integrated Design and Validation, Air Traffic Management, Airports, Human Factors, and Innovative Concepts and Scenarios. Within these areas several Interest Groups have been established for Advanced Combustion Chambers, Ageing Aircraft, Crashworthiness and Structural Impact, Emission Minimizing Flight Operations, Fault Tolerant Systems, Increased Exploitation of Composites, Manufacturing Processes and Technologies of Aero-Engines, Risk Analysis Based LCE in Aeronautics, Surface Engineering Treatments, Vortical Structures and IG Innovative Contacts and Scenarios. The central element is an open, Internet based network that will enable communication between groups and will provide access to a database. The Network Database contains Information on the university institutes with their aerospace competence profiles, companies and organisations in the aerospace supply chain, research establishments and information on national aeronautics research programmes. The web site describes the network members and provides details of each of the regional contact points. It identifies R&T areas and Interest Groups and the Universities who are engaged in research activities in these areas. The site also provides news and a list of related links.

Evaluation of the Probabilistic Design Methodology and Computer Code for Composite Structures

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/12, by Michael Shiao, dated June 2001. The report presents the results of an independent evaluation on the numerical accuracy and computational efficiency of a probabilistic design methodology for composite aircraft structures. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Experimental and Computational Failure Analysis of Graphite/Bismaleimide Laminated Composite and Carbon Foam in Sandwich Construction

This is the full text of a thesis by Troy C. Welker which was presented to the Air Force Institute of Technology (AFIT) in 2003. Sandwich beams consisting of a carbon foam core and graphite/bismaleimide face sheets were constructed and tested. Nine specimens were fabricated, using three distinct cross-ply symmetric face sheet layups with a constant core thickness. Four-point bend testing, controlled by a constant rate of midspan vertical displacement, was used to load the specimens to failure. Strains were recorded at midspan on the top and bottom faces, and vertical displacement was measured at midspan. Failure modes were observed for the beams and compared to results from a layerwise finite element method. The finite element code was originally written to dynamically calculate stresses in a sandwich plate dropped on an elastic foundation, and was enhanced to handle concentrated static forces. In addition to in-plane strains and stresses, the finite element method takes into account both normal and shear strains and stresses through the thickness of the beam. The stress results were used to evaluate failure criteria for both the composite face sheets and foam core. Displacements and strains from the experiment were compared to analytical sandwich beam theory, and displacements and failure loads were compared to the finite element solution. A phenomenological failure criterion was developed that compares favorably with experimental failure data. The finite element solution gives failure within an average of 7.58% of experiment, and a stiffness within an average of 11.16%. The analytical sandwich beam theory predicts stiffness within 4.83% of experiment, and strain within 5.27% of experiment. A primary goal of the research was to evaluate the ability of the layerwise finite element method, a method having potential use in a structural health monitoring system, to predict failure onset and location in a structure. Another goal was to determine if the structures responded in such a way that core failure in shear was reinforced by the face sheets. This research shows that the finite element theory has the ability to predict failure onset and location in a sandwich structure, and that the face sheet layups with higher stiffness delay the onset of shear failure in the core. [taken from abstract]. The full text is available in PDF format and is provided by the Air University ResearchWeb site.

Experimental Aspects and Mechanical Modeling Paradigms for the Prediction of Degradation and Failure in Nanocomposite Materials Subjected to Fatigue Loading Conditions

This is the full text of a Georgia Institute of Technology, Department of Polymer, Textile and Fibre Engineering, dated August 2008. The objective of the current research was to contribute to the area of mechanics of composite polymeric materials. This objective was reached by establishing a quantitative assessment of the fatigue strength and evolution of mechanical property changes during fatigue loading of nanocomposite fibers and films. Both experimental testing and mathematical modeling were used to gain a fundamental understanding of the fatigue behavior and material changes that occurred during fatigue loading. In addition, the objective of the study was to gain a qualitative and fundamental understanding of the failure mechanisms that occurred between the nanoagent and matrix in nanocomposite fibers. This objective was accomplished by examining scanning electron microscopy (SEM) fractographs. The results of this research can be used to better understand the behavior of nanocomposite materials in applications where degradation due to fatigue and instability of the composite under loading conditions may be a concern. These applications are typically encountered in automotive, aerospace, and civil engineering applications where fatigue and/or fracture are primary factors that contribute to failure.[taken from abstract]. The full text of the thesis is available in PDF format via the Georgia Institute of technology

Experimentation and Analysis of Composite Scarf Joint

This is the full text of a Master's thesis by Captain Benjamin M. Cook, USAF, AFIT/GA/ENY/05-M03, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2005. Composite bonded scarf repairs were examined by experimentally measuring and analytically predicting the residual curing strains and strains due to mechanical loading. To accomplish this a three prong approach was used: a full strain field through a repaired laminate's thickness was measured for both a loaded specimen and a specimen with the residual strain released, models were developed for comparison to both states, and data was collected for large tensile test specimens at various stages of being scarf repaired. A ^14:1 straight scarfed one-inch wide specimen was used to collect Moire interferometry data to calculate a full strain field due to mechanical loading and strain release. A three-dimensional thermo mechanical linear elastic analysis using an Air Force Research Laboratory in-house stress analysis program B-Spline Analysis Method (BSAM) results were correlated to the Moire interferometry test results. Three large tensile test specimens were tested as manufactured, three were tested with a scarfed hole in the center, and the remaining were tested with a scarf repair centered on a hole in the center. The strain gage results from the panels are presented. An additional feature of this work was to document each of the difficulties present in the given methods incorporated in this research. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Fatigue and Stress Relaxation of Adhesives in Bonded Joints.

This provides access to a Federal Aviation Administration (FAA), DOT/FAA/AR-03/56 by John Tomblin dated October 2003. Increased confidence in composite materials have increased the use of adhesively bonded components in primary aircraft structures. Applications on primary structures require rigorous characterization of the material, unlike applications on secondary structures. Previously funded Federal Aviation Administration programs at the National Institute for Aviation Research at Wichita State University in Wichita, Kansas, have extensively researched thick bondline adhesive joint behavior, as well as adhesive joint behavior for a wide range of bondline thickness, with respect to several aircraft operating environmental conditions. The primary goal of this investigation was to characterize the long-term durability of adhesives. Two simultaneous investigations were performed to characterize fatigue and stress relaxation behavior. To gain more information about high-stress behavior (low-cycle fatigue), stress amplitudes resulting in adhesive failure at low levels of 1,000 to 100,000 cycles were used. Other subsequent tests were conducted at those levels, including temperature and frequency dependence of adhesives. In addition, stress relaxation behavior of various cases, such as during fuselage pressurization, was studied under three different environmental conditions and at three different stress levels. Both fatigue and stress relaxation tests provide vital design data for long-term durability of adhesive-bonded structures. Fatigue tests showed that moisture absorption shortened the fatigue life of adhesives from the lives at ambient room temperature or cold dry conditions. For most cases, stress relaxation tests indicated that the higher the stress level and test temperature, the higher the stress decay during relaxation. These results also revealed the danger of designing an adhesive structure to operate at temperatures near the glass transition temperature. [Taken from abstract]. The full text of the report is avialable in PDF format from the online catalogue of the FAA William J Hughes Technical Center Library.

Fatigue Behavior of a Functionally-Graded Titanium Matrix Composite

This is the full text of a Master's thesis by Captain Scott R. Cunningham, USAF, AFIT/GAE/ENY/05-M06, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in March 2005. Functionally-graded Titanium Matrix Composites are an attempt to utilize the high-strength properties of a titanium matrix composite with a monolithic alloy having the more practical machining qualities. This work studied the mechanical characteristics of the joint region as a first step toward future evaluation of this material. The scope of this effort involved testing under monotonic tension and fatigue loading conditions. Mechanical properties and cyclic behavior were evaluated for the joint area and then compared to those of the parent materials. The results of this study found that the strength of the transition region was slightly higher than the unreinforced alloy. However, the presence of fiber ends in the transition region proved to be the source of failure under fatigue loading conditions. Failure in the transition region did not occur at the tip of the taper joint as anticipated. Instead, failure occurred at the transition to the next ply in the taper. This indicates that fiber volume, in conjunction with the presence of fiber ends, plays a key role in the fatigue life of the entire material. These findings encourage and provide the basic scientific knowledge for further evaluation and development of functionally-graded titanium matrix composites. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Fiber Composite Analysis and Design : Composite Materials and Laminates : Vol. 1

This final report (DOT/FAA/AR-95/29) was published by the Federal Aviation Administration (FAA) in October 1997, and was written by Z. Hashin et al. The report provides extensive background information on the characteristics and mechanics of fibre reinforced composites which permits engineers experienced in the evaluation of structures involving conventional materials - especially metals - to extend their competence to an assessment of fibrous composite structures. The emphasis is on the definition of the nature and magnitude of the differences associated with the use of composites rather than conventional materials and especially towards the elucidation of the reasons for the differences and their implications for design evaluation. Accordingly, a broad spectrum of technologies is involved, ranging from detailed methods of analysis to more qualitative discussions on methods of analysis and design. This is a PDF file [339 pages, 16.7 Mb] so Adobe Acrobat software will be required in order to read it.

Fibre Composite Aircraft : Capability and Safety

This provides access to an Australian Transport Safety Bureau Research and Analysis Report No.AR2007021 dated June 2008. For many decades, fibre composites have been replacing traditional aluminium structures in a wide variety of aircraft types. From the first all-composite kit plane released in 1957, composites are widespread today in commercial aircraft and many other aircraft types. This is due to the cost and weight savings that materials such as glass/phenolic and carbon/epoxy offer aircraft manufacturers over aluminium, while maintaining or surpassing its strength and durability. This study provides an overview of fibre composite use in aircraft and the issues associated with its use, with a focus on aircraft operating in Australia that contain these materials. There are almost 2,000 aircraft on the Australian civil register made of, or containing, fibre composite materials. This includes most of the mainline jet fleet, effectively all sailplanes and gliders, many popular general aviation aircraft, and a third of the growing amateur-built aircraft category. There is a lot of conflicting or incorrect information in the aviation community about the safety and capability of fibre composite materials. Composite structures behave very differently under normal loads than equivalent metal structures. Fatigue and corrosion have been proven through trials of composite repair patches to be much less prevalent in composites compared with metals. Subsurface damage such as delamination however can go undetected for long periods and result in sudden catastrophic failure. It is important that operators of fibre composite aircraft be aware of correct detection and repair procedures for the unique types of damage that occur to composites. First responders involved in post-crash cleanup operations have expressed concerns about the long-term effects from exposure to carbon fibres released from burning composites. Fibre dust can pose an inhalation risk similar to asbestos. Released fibres or splinters are needle-sharp, and can cause skin and eye irritation. In the event of a post-crash fire, smoke and toxic gases are also released from decomposing composites, presenting further health risks. [Taken from abstract]. The full text is available in PDF so Adobe Acrobat software is required in order to read it.

Finite Element Analysis of a Composite Cylindrical Shell with a Cutout Under Fatigue Loading

The site provides access to an Air University Air Force Institute of Technology MSc Thesis, by Captain Joshua T. Boatwright, USAF, AFIT/GA/ENY/00M-03, dated March 2000. The thesis describes the analysis of compressive and tensile loads on a graphite/epoxy laminated cylinder containing a square cutout. Citation details and an abstract are available in HTML format. The full text can be accessed in PDF format (2,650,489 bytes). The document is part of the Air University Research Database.

Fire Safety of Advanced Composites for Aircraft

This web site provides access to an Australian Bureau of Air Safety Investigation (BASI) document (Grant 20040046) written by Professor A.P. Mouritz and dated April 2006. Fire contributes to aircraft accidents and many fatalities. The growing use of polymer composite materials in aircraft has the potential to increase the fire hazard due to the flammable nature of the organic matrix. This report assesses the fire hazard of current and next-generation polymer composites for aircraft, and identifies those materials with improved flammability resistance. A comprehensive review of the scientific literature was performed to develop a database on the fire properties of a large number of polymer composite materials. For both aircraft cabin materials and aircraft structural materials the following fire properties were considered in the determination of fire safety: time-to-ignition, limiting oxygen index, peak heat release rate, average heat release rate, total heat release, flame spread rate, smoke, and combustion gases. The data is presented as performance tables which rank the composite materials in order from best to worst. The composite most often used in pressurised aircraft cabins is glass/phenolic, and the database shows that this material has excellent fire reaction performance and that very few next-generation composites display superior properties. The most used structural composite is carbon/epoxy, and this material has poor fire resistance and can pose a serious fire hazard. A number of advanced structural composites with superior fire properties are identified, including materials with high temperature thermoset polymer, thermoplastic or inorganic polymer matrices. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software will be required in order to read it.

Fire-Resistant Cyanate Ester-Epoxy Blends : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/53, by Richard N. Walters, dated May 2002. The cure chemistry, thermal stability, and fire behavior in a series of fire-resistant cyanate ester-epoxy blends were examined. The dicyanate and diepoxide of 1,1-dichloro-2,2-bis(4-hydroxyphenyl)ethylene (bisphenol-C, BPC) were combined in various molar ratios, and the reaction chemistry was monitored using Fourier Transform Infrared Spectroscopy (FTIR) and differential scanning calorimetry (DSC). Fire behavior of the BPC cyanate-epoxy blends was studied in flaming and nonflaming combustion, using fire calorimetry and pyrolysis-combustion flow calorimetry, respectively. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Fire-Resistant Elastomers

This is a Federal Aviation Administration (FAA) report, DOT/FAA/AR-TN01/104, by Richard E. Lyon, dated May 2002. Molecular design of semi-inorganic rubbers has yielded flexible polysilphenylene-siloxane and polyphosphazene elastomers having the fire resistance of rigid, high-temperature engineering plastics (e.g., polyaramids, polyetherketones, and polyarylsulfones) based on the results of microscale combustibility data. In flaming combustion, a commercially viable polyphosphazene exhibited a 75% reduction in heat release rate compared to the polyurethane rubber currently used in fire-blocked foam aircraft seat cushions. A comparable reduction in heat release rate is obtained at lower cost by adding expandable graphite flakes to the polyurethane formulation. The graphite flakes exfoliate during heating to produce a 2000% volumetric expansion of the burned rubber which shields and insulates the underlying material from the heat source and lowers the heat release rate significantly. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Fire-Safe Polymers and Polymer Composites

This technical report (DOT/FAA/AR-04/11) was produced by the Fire Safety Branch of the Federal Aviation Administration (FAA) in September 2004 and was written by Huiqing Zhang. The intrinsic relationships between polymer structure, composition, and fire behavior have been explored to develop new fire-safe polymeric materials. Three milligram-scale methods (pyrolysis-combustion flow calorimetry (PCFC), simultaneous thermal analysis, and pyrolysis gas chromatography/mass spectrometry (GC/MS)) have been combined to fully characterize the thermal decomposition and flammability of polymers and polymer composites. Thermal stability, mass loss rate, char yield, and properties of decomposition volatiles were found to be the most important parameters in determining polymer flammability. Most polymers decompose by either an unzipping or a random chain scission mechanism with an endothermic decomposition of 100-900 J/g. Aromatic or heteroaromatic rings, conjugated double or triple bonds, and heteroatoms such as halogens, N, O, S, P, and Si, are the basic structural units for fire-resistant polymers. The flammability of polymers can also be successfully estimated by combining the pyrolysis GC/MS results or chemical structures with the thermogravimetric analysis. The thermal decomposition and flammability of two groups of inherently fire-resistant polymers—poly(hydroxyamide) (PHA) and its derivatives and bisphenol C (BPC II) polyarylates—have been systematically studied. PHA and most of its derivatives have extremely low heat release rates and very high char yields upon combustion. PHA and its halogen derivatives can completely cyclize into quasi-polybenzoxazole structures at low temperatures. However, the methoxy and phosphate derivatives show a very different behavior during decomposition and combustion. Molecular modeling shows that the formation of an enol intermediate is the rate-determining step in the thermal cyclization of PHA. BPC II-polyarylate is another extremely flame-resistant polymer. It can be used as an efficient flame-retardant agent in copolymers and blends. From the PCFC results, the total heat of combustion of these copolymers or blends changes linearly with composition, but the change of maximum heat release rates also depends on the chemical structure of the components. The flammability of various polymers and polymer composites measured by PCFC; cone calorimeter, ASTM E1354; and the Ohio State University (OSU) calorimeter, ASTM E906, were also compared. For pure polymers, there was a relatively good correlation between different methods. However, for polymer composites with inert fillers or flame-retardant additives, the OSU and cone calorimetries are more suitable evaluation methods. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Flammability of Polymer Composites

This provides access to a US Federal Aviation Administration (FAA) technical report DOT/FAA/AR-08/18 written by Richard Walters and Richard Lyon dated May 2008. The flammability and mechanical properties of fiber-reinforced thermoset resin structural composites were evaluated. Processing characteristics, thermal stability, and flammability of the neat resins were measured using rheology, thermogravimetry, and pyrolysis-combustion flow calorimetry, respectively. Structural laminates were fabricated from liquid resins and woven glass fabric by vacuum-assisted resin transfer molding. Single-layer specimens (lamina) were prepared for fire testing using a hand lay-up technique. Mechanical properties of the laminates were measured in three-point bending. Fire behavior of the lamina and laminates was measured according to Title 14 Code of Federal Regulations 25.853(a-1) and Military Standard MIL-STD-2031. Results for flammability, fire performance, and mechanical properties of these composites are presented in this report. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat softwareis required in order to read it.

Flammability Properties of Aircraft Carbon-Fiber Structural Composite

This technical report (DOT/FAA/AR-07/57) was produced by the Fire Safety Branch of the Federal Aviation Administration in October 2007 and was written by James G. Quintiere, Richard N. Walters, and Sean Crowley. This study investigated the flammability of a carbon-fiber composite material for use in aircraft structures. In particular, it considered a composite material manufactured by Toray Composites (America) to Boeing Material Specification 8-276. The objective was to establish a complete set of properties pertaining to the heating and burning characteristics of these materials in fires. Several apparatuses were used, including the cone calorimeter, microscale combustion calorimeter, thermogravimetric analyzer, differential scanning calorimeter, and a flame spread rig to promote spread with preheating by radiation. An attempt was made to measure the thermal conductivity of the composite over a range of temperatures through its decomposition, but the heat losses from the apparatus likely caused an overestimate in the measurement. Data from standard tests were also reported for the Ohio State University calorimeter and the smoke density chamber. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Fuel Tank Flammability Assessment Method User's Manual

This provides access to a US Federal Aviation Administration (FAA) technical report DOT/FAA/AR-05/8 written by Steven Summer and dated May 2008. The Fuel Tank Flammability Assessment Method (FTFAM) is a Federal Aviation Administration-developed computer model designed as a comparative analysis tool to determine airplane fuel tank flammability as a requirement of Title 14 Code of Federal Regulations 25.981. The model uses Monte Carlo statistical methods to generate flammability data for certain unknown variables over known distributions for a large number of flights. The FTFAM iterates through each flight, calculating the flammability exposure time of each flight given the data input provided by the user. Calculating this flammability exposure time for a sufficiently large number of flights results in statistically reliable flammability exposure data. These calculations can be performed by the user for virtually any type of airplane fuel tank (body tank, wing tank, auxiliary tank, etc.) both with and without a flammability reduction method being employed. This report serves as a user’s manual for this computer model to assist the user in its operation and to discuss the permissible changes that may be made to this model specific to a particular fleet of aircraft. It is updated through version 10 of the FTFAM. The user should reference Advisory Circular 25.981-2A for additional guidance on when to use this model and for a discussion of interpretation of results. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Future Materials Science Research on the International Space Station

This web site provides access to a report prepared by the National Research Council's Committee on Materials Science Research on the International Space Station, Washington, D.C., National Academy Press, 1997. The report presents the results of the Committee's evaluation of NASA's high-temperature, microgravity materials research plan. In particular it focuses on the Space Station Furnace Facility (SSFF) Core. This will provide the mechanical, power and control infrastructure for an array of experiment modules designed for crystal growth and solidification research in the fields of electronic and photonic materials, metals and alloys, glasses and ceramics. The report also reviews NASA's selection procedures for identifying and supporting research within the developing field of materials science and engineering. An executive summary is available in HTML format, and the full text of the report can be accessed online in Open Book format.

Going to Extremes : Meeting the Emerging Demand for Durable Polymer Matrix Composites

This is a full text book made available by National Academies Press and produced by the National Research Council's Committee on Durability and Life Prediction of Polymer Matrix Composites in Extreme Environments, August 2005. Advanced polymer matrix composites (PMCs) have many advantages such as light weight and high specific strength that make them useful for many aerospace applications. Enormous uncertainty exists, however, in predicting long-term changes in properties of PMCs under extreme environmental conditions, which has limited their use. To help address this issue, the Department of Defense requested a study from the NRC to identify the barriers and limitations to the use of PMCs in extreme environments. The study was to focus on issues surrounding methodologies for predicting long-term performance. This report provides a review of the challenges facing application of PMCs in extreme environments, the current understanding of PMC properties and behavior, an analysis of the importance of data in developing effective models, and recommendations for improving long-term predictive methodologies. Bibliographic and abstract information is available in HTML format, access to the full text is provided online in Open Book format with printable PDF files, and an abridged reports version is also available in PDF format.

Graphitized Carbon Foam with Phase Change Material

This is the full text of a Master's thesis by Captain Angelinda D. Fedden, USAF, AFIT/GA/ENY/06-M02, which was presented to the Faculty Department of Aeronautical and Astronautical Engineering of Air University's Air Force Institute of Technology (AFIT), in March 2006. This thesis examines the transient heating and cooling responses of graphitized carbon foam infiltrated with phase change material (PCM). The carbon foam provides rapid heat transfer throughout the PCM volume, while the PCM stores the heat for later removal. The foam/PCM system was heated with a copper heating block, and then cooled with a liquid-cooled heat removal block. Infiltrating the foam with PCM significantly increased the length of time before the system reached maximum temperature. The temperature response of the foam/PCM system was consistent over multiple cycles of heating and cooling. A high density foam had a faster heating and cooling response than a low density foam. A comparison of the temperature profile at various locations within a sample shows that it can be modeled as a lump block. The effects of contact resistance were shown by using different substances between the heater and the test article. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Guide to the measurement of the transient performance of aircraft turbine engines and components

This is Research and Technology Organization (RTO) AGARD-AR-320, dated March 1994. This report provides a guide for the measurement of transient aerothermodynamic performance parameters of aircraft gas turbine engines or components for engine developers, test agencies, certifying authorities, and operators of overhaul facilities and aircraft. It may be treated as an extension of AGARD Advisory Reports AR-245 and AR-248. It includes discussion of recommended procedures for the transient measurement of pressures, temperatures, flows, component geometry including rotational speed and clearances, thrust, torque, and the use of the engine control system for transient parameter measures. Typical examples are presented. A section on data acquisition and processing is included. Higher frequency dynamic measurements are excluded. Two examples, the measurement of compressor ratio and air flow at surge and a measurement of engine acceleration time are discussed in detail. This Advisory Report was prepared at the request of the Propulsion and Energetics Panel of AGARD. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (19.39MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Guidelines and Recommended Criteria for the Development of a Material Specification for Carbon Fiber/Epoxy Fabric Prepregs

This technical report (DOT/FAA/AR-06/10) was produced by the Federal Aviation Administration (FAA) Office of Aviation Research and Development in May 2007 and was written by Stephen Ward, William McCarvill and John Tomblin. The building block approach is used within the composite aircraft industry for the substantiation of composite structure. A key element supporting the building block approach is material and process specifications. Material and process specificationsare interwoven throughout the qualification and validation process. They are used to define the material’s attributes, define the qualification characterization tests, purchase the material, and define and control the processes used for the fabrication. It is critical that the qualification specimens fabricated through the various levels of the building block approach use the same process, which is representative of the one that will be used in the fabrication of production aircraft and rotorcraft. This document recommends guidance and criteria for the development of material specifications for carbon fiber/epoxy prepreg fabric materials to be used on aircraft structures. This report is intended to be a companion to previous Federal Aviation Administration reports which established methodology for developing material allowable data, control of the data, and sharing the resulting database, DOT/FAA/AR-02/109 and DOT/FAA/AR-02/110. These reports provide recommendations for prepreg tape materials and processing. The guidelines and recommendations are meant to be a documentation of current knowledge and practices, and application of sound engineering principles to the development and implementation of composite material procurement specifications. A list of material control areas needing improvement and enhancement is provided for discussion. This document can also be used to develop common industry specifications. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Guidelines and Recommended Criteria for the Development of a Material Specification for Carbon Fiber/Epoxy Unidirectional Prepregs : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/109, by William McCarvill, Stephen Ward, Gregg Bogucki, and John Tomblin, dated March 2003. This report establishes recommendations to guide the development of composite prepreg material specifications to be used on aircraft structures.The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Guidelines and Recommended Criteria for the Development of a Material Specification for Carbon Fiber/Epoxy Unidirectional Prepregs Update

This technical report (DOT/FAA/AR-07/3) was produced by the Federal Aviation Administration (FAA) Office of Aviation Research and Development in May 2007 and was written by William McCarvill, Stephen Ward and John Tomblin. This report establishes recommendations to guide the development of composite prepreg material specifications. This is intended to advance the work that has been done through previous Federal Aviation Administration and National Aeronautics and Space Administration programs such as the Advanced General Aviation Transport Experiment. These programs have established methodologies for developing design allowable data, control of the data, and sharing the resulting database. In the current work, a generalized approach to the development of a shared composite material database is proposed. It is intended to remove the restrictions placed on those general aviation methods to allow a broader market to use the shared database. This document recommends guidance and criteria for the development of material specifications for carbon fiber/epoxy unidirectional prepreg tape materials to be used on aircraft structures. These recommendations were prepared by a team of industry experts. The guidelines and recommendations are meant to be a documentation of current knowledge and application of sound engineering principles to the development and implementation of composite material procurement specifications. A list of material control areas needing improvement and enhancement is given in appendix A. This document can also be used to develop common industry specifications. This document is limited to recommendations and guidance on the development of the material specification. Additional guidance on the development of process specifications, instructions, and controls for making high-quality laminates can be found in the companion report Guidelines for the Development of Process Specifications, Instructions, and Controls for the Fabrication of Fiber-Reinforced Polymer Composites, DOT/FAA/AR-02/110. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Guidelines for Analysis, Testing, and Nondestructive Inspection of Impact-Damaged Composite Sandwich Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/121, by Peter Shyprykevich, J.Tomblin, and L.Ilcewicz, dated March 2003, In this report, past work is summarized and synthesized to provide guidance for analysis, testing, and nondestructive inspection of impact-damaged composite structures. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Guidelines for the Development of a Critical Composite Maintenance and Repair Issues Awareness Course

This provides access to a US Federal Aviation Administration (FAA)technical report DOT/FAA/AR-08/54 written by Larry LLcewiz ...[et al] and dated February 2009. This report documents the results of a Federal Aviation Administration Cooperative Agreement with Edmonds Community College to develop a standard for teaching an introductory course on critical composite maintenance and repair issues. The course will serve as an introduction and provide an awareness of safety issues regarding the maintenance and repair of composite materials used in aerospace. The course will also provide a background for further study for those interested in becoming qualified practitioners. This course is intended for engineers, technicians, inspectors, and other personnel involved with the maintenance and repair of composite structures. The framework for the awareness course is defined by Terminal Course Objectives, which are summarized further into terminal course modules. The four main areas at the highest outline level include base knowledge, teamwork and disposition, damage detection and characterization, and repair processes. Content includes text, laboratory instructions, and videos in support of the course objectives. Materials for evaluating the effectiveness of the course content in meeting the standards represented in this report are provided with reference to industry documents, especially CACRC AIR report 5719, which provides a checklist of detailed teaching points. An instructor’s guide is also included to assist in developing the course. Additionally, several collaborative workshops and other forums that were held during the development process, involving industry, academia, and government regulatory agencies are described. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Guidelines for the Development of Process Specifications, Instructions, and Controls for the Fabrication of Fiber-Reinforced Polymer Composites : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/110, by Gregg Bogucki, William McCarvill, Stephen Ward, and John Tomblin, dated March 2003. This document provides (1) a set of guidelines for the development of process specifications for the fabrication of continuous fiber-reinforced polymer composite laminate test panels used in the generation of mechanical properties and (2) an approach for the validation of composite fabrication processes used during the certification of composite aircraft structure. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center library.

Handbook of brittle material design technology

This is Research and Technology Organization (RTO) AGARD-AG-152, dated February 1971. Interest in the structural use of brittle non-metallic refractory materials in aerospace vehicles arises as a result of interest in re-entry vehicles and the continuing need for propulsion systems of increased performance. In both of these situations performance is dependent on the temperature capability of the structural materials. The presence of temperatures beyond the capability of most metallic materials has encouraged designers to study “ceramics� but, with the possible exception of inserts in the throats of some solid rocket motors, significant applications have not developed. Characteristically these materials show no plastic deformation before failure at temperatures of interest, and they have little toughness to arrest crack growth. These characteristics have resulted in a lack of confidence by designers, in such materials, and unwillingness to use them. The materials to which this handbook is intended to apply include oxides, carbides, borides and similar compounds. Graphite, in its many forms, is also included. The important characteristic of these materials is refractoriness, which permits them to be used in applications where the more structurally efficient metals are useless. Such materials have been used extensively in the past for high temperature applications, such as furnace linings, but these have involved ground installations where weight was generally not important. Currently the interest involves the high temperature applications generated by re-entering space vehicles and rocket engine compounds, but since these are extremely weight critical applications, substantial improvements in structural efficiency and reliability over those typical of the furnace type application must be obtained. Among the applications which are of current interest for this class of material are numerous components for winged re-entry vehicles or hypersonic atmospheric vehicles, and these include leading edge elements, nose caps, surface panels, which may or may not include elements for insulation, control surface structural parts, and engine intake structural parts. Rocket engine nozzles and chambers are the obvious propulsion applications. Bibliographic and abstract details are available in HTML format. A table of contents and the full text (76.36MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Health Hazards of Combustion Products from Aircraft Composite Materials

This final report (DOT/FAA/AR-98/34) was published by the Federal Aviation Administration in September 1998, and was written by Sanjeev Gandhi and Richard E. Lyon. Concerns about the potential health hazards of burning fiber-reinforced polymer composites in aircraft fires parallel the rising usage of these materials for commercial aircraft primary and secondary structures. An overview of the nature and the potential hazards associated with airborne carbon fibers released during flaming combustion of aircraft composites is presented [extracted from FAA abstract]. This is a PDF file [29 pages, 223Kb] so Adobe Acrobat software will be required in order to read it.

High-Performance Structural Fibers for Advanced Polymer Matrix Composites

This is a full text book made available by National Academies Press and produced by Committee on High-Performance Structural Fibers for Advanced Polymer Matrix Composites, National Research Council (NRC), in 2005. Military use of advanced polymer matrix composites (PMC) consisting of a resin matrix reinforced by high-performance carbon or organic fibers while extensive, accounts for less that 10 percent of the domestic market. Nevertheless, advanced composites are expected to play an even greater role in future military systems, and the U.S. Department of Defense (DOD) will continue to require access to reliable sources of affordable, high-performance fibers including commercial materials and manufacturing processes. As a result of these forecasts, DOD requested the NRC to assess the challenges and opportunities associated with advanced PMCs with emphasis on high-performance fibers. This report provides an assessment of fiber technology and industries, a discussion of R&D opportunities for DOD, and recommendations about accelerating technology transition, reducing costs, and improving understanding of design methodology and promising technologies. The full text of the report can be read online in open book format, and a summary is available for downloading in PDF format.

Impact and Delamination Failure Characterization of BMS 8-212 Composite Aircraft Material

This provides access to a US Federal Aviation Administration (FAA) technical report DOT/FAA/AR-08/48 written by David Powell, Tarek Zohdi and George Johnson dated October 2008. This study was part of a program, sponsored by the Federal Aviation Administration that focused on understanding the behavior of different aircraft materials under impact in the speed range generated from engine uncontained failures. This test program investigated the material response of 8-, 16-, and 32-ply BMS8-212 composite panels provided by The Boeing Company. All tests were performed at the University of California, Berkeley ballistics laboratory using a pneumatic gas gun and half-inch-diameter spherical and flat-ended cylindrical projectiles. The ballistic impact tests indicated that the amount of energy absorbed by a similar composite target panel during impacts above the ballistic limit was nearly constant, showing only a slight increase with increasing initial energy. The amount of energy absorbed per ply increased only slightly for thicker panels. The tests also showed that the cylindrical projectiles required more energy to penetrate the composite panels than the spherical projectiles. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Impact Damage Characterization and Damage Tolerance of Composite Sandwich Airframe Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/44, by John S. Tomblin et al, dated January 2001. The report describes the results of damage tolerance experiments on sandwich panels conducted at Wichita State University. The effect of impactor size on impact resistance and residual strength was investigated. The effectiveness of traditional nondestructive inspection (NDI) methods for detecting and quantifying damage distribution in sandwich panels was also studied and the salient results are presented. The full text of the report is available in PDF format.

Impact Damage Characterization and Damage Tolerance of Composite Sandwich Airframe Structures – Phase II : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-02/80, by Tomblin, John S., Raju, K.S., Acosta, J.F., Smith, B.L., and Romine, N.A. dated October 2002. The impact responses and the damage states in flat composite sandwich panels with thin facesheets were investigated in Phase I and were found to be dependent on the diameter of the spherical steel impactor. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Impact of Acoustic Loads on Aircraft Structures

This is Research and Technology Organization (RTO) AGARD-CP-549, dated September 1994. A broad band of different activities was addressed in the Specialists' Meeting held by the Structures and Materials Panel of AGARD in May 1. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (152.88MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Intelligent Processing of High Performance Materials

This site provides access to a Research and Technology Organisation Meeting Proceedings, RTO-MP-009, Paris, November 1998. The document contains papers presented at a Workshop on Intelligent Processing of High Performance Materials organised by the Applied Vehicle Technology Panel (AVT) of RTO, in Brussels, Belgium, 13-14 May 1998. It describes various aspects of intelligent processing, a methodology for simulating and controlling the processing and manufacture of materials, which is finding widespread application during the manufacture of functional electronic, photonic and composite materials as well as primary metals such as steel and aluminium. The papers are presented under the following headings: •Overview and analytical techniques •Metallic materials applications •Non-metallic materials applications Bibliographic and abstract details are available in HTML format. A table of contents, and the full text of the document (18.5 Mbytes) can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Introduction of Ceramics into Aerospace Structural Composites

This is Research and Technology Organization (RTO) AGARD-R-795 , dated November 1993. Ceramics have been considered over the last two decades as a possible alternative to refractory metals and alloys to be used as structural materials for aeronautical use. The main disadvantage of these materials is their brittleness and the very low value of the critical size of defects leading to fracture. The concept of ceramic matrix composites has been recognized as one of the ways to escape this difficulty. Extensive work has been performed to identify the mechanisms of crack propagation and general fracture for unidirectional composites, laminates or other fabrics, including the understanding of their long term response: creep and fatigue effects or environmental degradation. The Workshop which has been held by AGARD SMP at Antalya (Turkey), April 1993, aimed at reviewing the present knowledge on all these aspects. For individual titles, see N94-24229 through N94-24240. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (47.65 MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Investigation of Adhesive Behavior in Aircraft Applications : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/57, by Charles Yang and John S. Tomblin, dated September 2001. There are two parts included in this report. The first part evaluates the most commonly used test method for adhesive properties, ADTM D 5656. The second part of this report provides (1) an analytical model for predicting stress distributions within an adhesive-bonded composite joint using ASTM D 3165 test specimen dimensions and (2) a method for predicting joint strength under the adherend failure mode. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Investigation of Thick Bondline Adhesive Joints : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/33, by John S. Tomblin et al, dated June 2001. The report examines bondline thicknesses from 0.010 - 0.160 inch for three paste adhesive systems using different test methods and substrate materials. The full text of the report is available in PDF format.

Italian Aerospace Research Centre

This is the homepage of the Italian Aerospace Research Centre (CIRA), a non-profit making research consortium which carries out the Italian National Aerospace Research Programme (PRORA). The site gives an introduction to the organisation and the scientific and the educational areas covered. It describes CIRA's aerospace activities (which include fluid dynamics, air structures, flight systems and computer science) and testing facilities (which include a plasma wind tunnel, an icing wind tunnel, a transonic research wind tunnel and an aerospace structures impact facility. The projects area of the site describes work being carried out on unmanned aerial vehicles and unmanned space vehicles. The educational area covers topics such as flight safety, pssengers comfort, air transportation and access to space. The general support area provides access to the Library Catalogue. The CIRA Newsletter is also available.

Laminate Statistical Allowable Generation for Fiber-Reinforced Composite Materials : Lamina Variability Method

This provides access to a Federal Aviation Administration technical report DOT/FAA/AR-06/53 written by John Tomblin and Waruna Seneviratne dated January 2009. Substantiation of composite structures often requires analysis as well as tests conducted at multiple building-block levels. Typically, as one proceeds up the building block and becomes more confident with the analytical design methods, the number of samples or replicates is reduced since the structural complexity increases. As the complexity of higher-level tests increases, more variables are introduced; i.e., fastener holes and multidirectional laminates. Currently, there is no standardized statistical methodology that allows progression up the pyramid of tests with a concurrent reduction of the number of tests while maintaining statistical reliability established at the lower replicate level. Therefore, conservative approaches are exercised at the higher level of building-block tests to be able to use the material variability of lower levels to represent the higher levels. Because a reduction in the number of samples results in excessively low or overly conservative material allowables, this research program explores the use of lamina-level variability to generate B-basis allowables for fiber-reinforced laminated composites. A multibatch laminate database was created to form baseline B-basis values. Then, B-basis allowables were obtained from single-batch data sets and compared with multibatch data to investigate the statistical reliability of the proposed small-sample B-basis method—the lamina variability method (LVM). This investigation included six material systems from three prepreg manufacturers: Toray Composites, FiberCote Industries, and Advanced Composites Group. In addition to the proposed methodology, two additional small-sample, B-basis methods proposed by Steve Ward were compared. Analysis of lamina and laminate test data for six different prepreg materials showed that LVM works well and gives conservative B-basis allowables that are about 90% of the three-batch laminate allowables. Caution should be exercised if the laminate data exhibits unusual variability due to test methods. If the variability is due to the material, the LVM method should be used with caution. If test variability is too high, for example, bearing strength, then three batches of laminate tests should be performed for that property to obtain B-basis allowables. It is imperative to realize the advantages as well as the risks associated with small-sample methods, and the data may need to be validated with additional tests based on specific application. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Low Cost Composite Structures and Cost Effective Application of Titanium Alloys in Military Platforms

This site provides access to a Research and Technology Organization Meeting Proceedings, RTO-MP-069 (II), Paris, March 2003. The document contains papers presented at the Symposium of the RTO Applied Vehicle Technology Panel (AVT) Specialists’ Meeting held in Loen, Norway, 7-11 May 2001. Bibliographic and abstract details are available in HTML format. The meeting provided a forum for information exchange and discussions by specialists on the cost effective application of titanium alloys to air, land and sea platforms. Twenty presentations from nine countries discussed titanium extraction, casting technology, component fabrication, use for ballistic protection, and applications in platform subsystems. A table of contents, and the full text (98.6 Mbytes) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Material Qualification and Equivalency for Polymer Matrix Composite Material Systems

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/47, by John S. Tomblin et al, dated April 2001. The report documents a qualification plan that will provide the detailed background information and engineering practices to help ensure the control of repeatable base material properties and processes, which are applied to both primary and secondary structures for aircraft products using composite materials. The full text of the report is available in PDF format.

Mechanical Behavior of Cracked Panels Repaired with Bonded Composite Patch

This is the full text of a Master's thesis by Captain Michael A. Hansen, USAF, AFIT/GA/ENY/05-J01, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in June 2005. This research focuses on investigating the mechanical behavior of cracked aluminum panels repaired with bonded boron/epoxy composite patches. The effects of crack initiation and growth on the residual strength of the repaired panels are characterized. This research establishes a correlation between damage modes, residual strength and evolution of strain within as well as outside the patch. Monotonic tensile tests on specimens with a perfectly bonded patch were used to determine the base line strength. Likewise, fatigue tests on specimens with a perfectly bonded patch served to establish baseline fatigue life. In addition, several specimens with a perfectly bonded patch were subjected to different fractions of the expected fatigue life, introducing damage, which were quantified by NDE techniques. These specimens were then subjected to a monotonic tensile test to failure in order to characterize the residual strength and the evolution of strain within and outside the patch, and the correlation between the disbonds and strain measurements at various locations on the specimen. This research looks to help in extending the service life of military and commercial aging aircraft, by using bonded composite patches on developing cracks in the structure. Bonded composite patches may be able to replace the crack patching technique of using bolted joints, which have the disadvantage of requiring holes to be machined in the metallic structure, which decreases its load-carrying capacity, creating stress concentrations and sites for crack initiation. In this study it was learned how the strain values increase as the crack grows. And despite differing crack growth rates, the strain values followed the growth of the crack closely throughout all the tests. The effects of overload situations were seen, and how this produces a retardation effect in the rate of growth of the crack. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Modeling Fracture in Z-Pinned Composite Co-Cured Laminates Using Smeared Properties and Cohesive Elements in DYNA3D

This is the full text of a Master's thesis by Captain Jason K. Freels, USAF, AFIT/GMS/ENY/06-S01, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in September 2006. The purpose of the present research was three-fold: 1) gain a more sophisticated understanding of the response of co-cured composite joints with and without through-thickness reinforcement (TTR), 2) compare the behavior of specimens reinforced with various sizes and densities of reinforcement, and 3) use experimental data to verify the existing DYNA3D smeared property model. Double cantilever beam, end-notch flexure and T-section specimens reinforced with 0.011" diameter z-pins at 2% and 4% volume densities were tested to determine the mode I, mode II and mixed mode (I and II) behavior. Results were added to preliminary research in which tests were conducted on previously mentioned specimen geometries reinforced with 0.022" diameter z-pins at similar densities. Experiments were modeled in DYNA3D using shell and cohesive elements. The energy release rate, G, determined through a curve fit developed from beam theory, was smeared across the region of reinforcement treating it as a separate material. The research validated Z-pinning as an effective means of improving the fracture toughness of polymer matrix laminated composites in mode I and mixed mode loading conditions and determined that the existing model works well in simulating the behavior in mode I tests. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Molar Group Contributions to the Heat of Combustion

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-TN01/75, by Richard N. Walters, dated September 2001. Experimental results for the gross heat of combustion of over 140 commercial and developmental polymers and small molecules of known chemical structure were used to derive additive molar group contributions. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Molecular Nanotechnology in Aerospace : 1999

This web page provides online access to a NASA Ames Research Center report, prepared by Al Globus, published January 2000. The report presents a high-level discussion of molecular nanotechnology, and addresses potential aerospace applications including, computers, materials and sensors. A bibliographic description and abstract is available in HTML format. The full text can be accessed online in either HTML format (2 Mbytes) or PDF format (3.4 Mbytes). The document is contained in the Papers and Reports section of the NASA NAS Systems Division's web site.

NASA Langley Research Center

This is the home page of the NASA Langley Research Center. Its primary concerns are airframe systems, atmospheric sciences and structures and materials research. The site describes the research that the Center carries out, information on doing business with it, and provides access to LISAR, the Langley Image Scanning Archival and Retrieval system which contains a database of photographs of NASA and NACA research undertaken at Langley. The site also features the Langley Factsheet Server which provides full text details about many of Langley's research programmes, the Langley Technical Report Server (LTRS) for searching and browsing technical reports, some of which are available in full in PDF format, and latest news.

NASA Tech Briefs Online : Engineering Solutions for Design and Manufacturing

This newsletter site includes articles and news items on the latest spin-off developments from NASA including a TechSearch of over 6000 technologies available for license. Free subscription is available and an archive is available to search as well as a Tech Brief library to browse. A new NanoTech Brief is available as well as news items and articles on the latest developments in the bio-medical, materials, mechanics, motion control, manufacturing, machinery, photonics and test and measurement fields.

Nondestructive Evaluation of Aircraft Composites Using Terahertz Time Domain Spectroscopy

This is the full text of a thesis written by Christopher Stoik which was presented to the Air Force Institute of Technology, Wright Patterson Air Force Base, Ohio, in December 2008. Terahertz (THz) time domain spectroscopy (TDS) was assessed as a nondestructive evaluation technique for aircraft composites. Material properties of glass fiber composite were measured using both transmission and reflection configuration. The interaction of THz with a glass fiber composite was then analyzed, including the effects of scattering, absorption, and the index of refraction, as well as effective medium approximations. THz TDS, in both transmission and reflection configuration, was used to study composite damage, including voids, delaminations, mechanical damage, and heat damage. Measurement of the material properties on samples with localized heat damage showed that burning did not change the refractive index or absorption coefficient noticeably; however, material blistering was detected. Voids were located by THz TDS transmission and reflection imaging using amplitude and phase techniques. The depth of delaminations was measured via the timing of Fabry-Perot reflections after the mail pulse. Evidence of bending stress damage and simulated hidden cracks was also detected with terahertz imaging.[Taken from Abstract]. This is in PDF format so Adobe Acrobat software is required in order to read it.

Office National d'Etudes et de Recherches Aerospatiales (ONERA)

Office National d'Etudes et de Recherches A�rospatiales (ONERA) is the French national aerospace research establishment. It is a scientific and technical public establishment managed according to industrial and commercial practice, placed under the supervision of the Minister of Defense. Its missions are to develop and guide aerospace research; design, develop and implement the facilities it requires to conduct its research and testing; publish and promote the results of its research; and contribute to the education of engineers and scientists. In addition to an overview of current research projects and interests, facilities and expertise, the site does provide a searchable database of scientific publications. This includes articles published in journals, papers delivered before conferences, theses, Technical Notes (NT), other publications, but not unpublished research reports. Some recent documents are available in full text format. An French language version of the site is also available.

ONERA Publications Database

ONERA (Office National d'Etudes et de Recherches Aerospatiale) is the French national aerospace establishment which has expertise in all the disciplines involved in aircraft, spacecraft and missile design. The ONERA database allows for searching of journal articles, papers delivered at conferences, theses, Technical Notes (NT), and other publications. Unpublished ONERA research reports are not included. The full text of selected recent documents is available in PDF format. It is possible to search by author name, information contained within the bibliographic record or the full text (if it is available). Searching in the full text field will return full text reports only.

Outgassing Data for Selecting Spacecraft Materials Online

This site features a database of outgassing data of materials intended for spacecraft use, obtained at the Goddard Space Flight Center (GSFC) Materials Engineering Branch, and utilising equipment developed at Stanford Research Institue (SRI) under contract to the Jet Propulsion Laboratory (JPL). The site provides a system description, and the possibility to download the entire Outgassing database. The Outgassing Materials Search function provides the capabilility to search all across materials tested at GSFC for determining the mass loss in a vacuum and for collecting the outgassed products. A Manufacturer Lookup allows to search on manufacturer name or code, or to browse the manufacturer list. There are also category and alphabetical listings and a report documentation page.

Performance Materials Net

This is a service that provides news and features about advanced and high-performance materials. Coverage includes advanced composites, engineering ceramics, performance plastics, and biomedical materials. All the items on the website are indexed so that information on specific applications, companies, or brands can be located using the menu links. The search engine can also be used to locate all relevant documents. Subscribers have free unlimited access, non-subscribers may view short descriptions of the content. There is a free email alert service and a forthcoming events area.

Preliminary Guidelines and Recommendations for the Development of Material and Process Specifications for Carbon Fiber-Reinforced Liquid Resin Moulded Materials

This provides access to a Federal Aviation Administration (FAA) report DOT/FAA/AR-06/25 by Gregg Bogucki ...[et al] dated June 2007. This document recommends guidance and criteria for the development of material and process specifications and material acceptance documents for liquid resins and continuous carbon fiber reinforcement materials used in liquid molding processes to manufacture structures for aircraft and space structures. The guidelines and recommendations are meant to be a documentation of current knowledge and application of sound engineering principles to the development and implementation of composite material procurement and process specifications. This document can also be used to develop common industry specifications. This report is limited to recommendations and guidance on the development of material and process specifications. The guidelines and recommendations contained in this document should not be viewed as Federal Aviation Administration policy or as the only acceptable method for composite material specifications and qualification procedures. They are meant to be a documentation of current knowledge and application of sound engineering principles to the development and implementation of composite material procurement specifications specific to the liquid resins and continuous fiber reinforcements used in liquid molding processes. The goal of any material procurement document is to provide the necessary controls to ensure the material used to establish the qualification data and certification data has not changed beyond its normal variability. The goal of any process specification is to ensure that the process remains in control and produces material consistent with specified requirements. [Taken from abstracts]. The full text is available in PDF format from the online catalogue of the William J. Hughes Technical Center Library.

PrePRINT Network : Department of Energy

This service is a searchable gateway to preprint servers provided by the US Department of Energy (DOE) Office of Scientific and Technical Information (OSTI). It is possible to search one site, a group of sites, or all the sites, or to browse an alphabetical listing of all preprint sites included on the service. An additional feature notifies individuals of new preprints that match a personally defined profile of subject interests. Subject areas covered include physics, mathematics, computer science, nonlinear sciences, engineering, and materials science.

Probabilistic Design Methodology for Composite Aircraft Structures

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/2, by M.W. Long and J.D. Narciso, dated June 1999. The report describes the evolution of probabilistic analysis, and the basic theory is discussed and explained via examples. Aerospace industry applications are discussed. It concludes with an assessment of potential benefits and concerns. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Probabilistic Design of Damage Tolerant Composite Aircraft Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/55, by A. Ushakov, A. Stewart, I. Mishulin, and A. Pankov, dated January 2002. The document describes the efforts under Memorandum of Cooperation between the Federal Aviation Administration (FAA), USA, and the Central Aero−Hydrodynamic Institute (TsAGI), Russian Federation. Under this effort, a methodology for calculating reliability of composite aircraft structures was developed and is contained in software called Probabilistic Design of Damage Tolerant Composite Structures (ProDeCompoS). The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Proceedings of the Institution of Mechanical Engineers, Part G : Journal of Aerospace Engineering

Published six times a year, the Journal of Aerospace Engineering is a forum for the communication of ideas and methods presently in use at the forefront of technology in the field of aerospace engineering. It contains papers on both theoretical and practical aspects of all types of civil and military aircraft and spacecraft and their support systems. The scope is wide, covering research, design, development, production, operation, servicing and repair, components and auxiliary equipment, safety and reliability. The site provides contents information for the journal. If you wish to view full text check with your library to see if they have a subscription.

Proceedings of the Institution of Mechanical Engineers, Part L: Journal of Materials: Design and Applications

Published four times a year, the Journal of Materials: Design and Application, is a forum for the communication of ideas and methods presently in use at the forefront of technology in the field of materials science. This journal is addressing a whole range of materials engineering and technology, which includes metallic materials, polymers, composites and ceramics. In addition, metal matrix composites and ceramic matrix composites are part of the portfolio. The site provides contents information for the journal. If you wish to view full text check with your library to see if they have a subscription.

Process Induced Residual Stresses and Dimensional Distortions in Advanced Laminated Composites

This web site provides access to a University of Florida, PhD dissertation, by Xiaokai Niu, dated 1999. The thesis describes the development of a technique called cure referencing method (CRM) for determining the residual stresses in flat laminated composites. The technique was validated using the shadow moiré method to measure the curvature of asymmetrical laminates while lamination theory is used to calculate the curvature from lamina strain information measured with CRM. Process induced strains on multi-directional composites were also measured with the CRM. Bibliographic and abstract details are available in HTML format. The title page, contents and the full text of the document are accessible online in PDF format. This title is part of the University of Florida's Electronic Theses and Dissertations Project

Rapra Polymer Bulletin : Aerospace Applications

This Polymer Bulletin is a current awareness service from the Polymer Library, the world's largest database dedicated to polymer literature. Each time the abstracts database is updated with new records (approx. every two weeks) you will be sent a bulletin alerting you to any items that relate to aerospace applications. This is a service that requires an annual subscription of £250.

Reduction of Thermal Residual Strains in Adhesively Bonded Composite Repairs

This is the full text of a thesis by Heather R. Crooks which was presented to the Air Force Institute of Technology (AFIT) in 2003. Many military and commercial aircraft are being called upon to fly well beyond their original intended service lives. This has forced the United States Air Force (USAF) to increasingly rely on structural repairs to address fatigue induced damage and to extend aircraft useful life. The focus of this research is the use of a high-strength composite patch technique to repair a fatigue crack on an aluminum aircraft structure. This study investigates the thermal residual strains that occur as a direct result of the coefficient of thermal expansion (CTE) mismatch between the repair patch and the underlying cracked metallic structure to which the patch is bonded. This research examines the response of a precracked, 24 inches x 6 inches x 0.125 inch, 1015-T6 aluminum panel repaired with a 15-ply graphite/epoxy patch. Two adhesives: EA 9696 and FM 13M were used with varying cure cycles. The hypothesis is that by reducing cure temperatures, the CTE mismatch will be less dramatic, thus yielding a more robust repair with a comparable fatigue crack growth rate. The research concluded that reducing the cure cycle temperature could decrease the thermal residual strains by as much as 26.5% between the graphite/epoxy composite patch and the aluminum structure when FM 13M adhesive is used to bond them together and 1.4% when EA 9696 is used. The research also concluded that a lower cure cycle temperature did not detrimentally affect the panels' fatigue crack growth rates. [Taken from abstract]. The full text of the thesis is available in PDF format and is provided by the Air University ResearchWeb site.

Repair of Composite Laminates : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/46, by Sung-Hoon Ahn and George S. Springer, dated December 2000. The report examines the repair effectiveness of damaged fibre reinforced composite laminates. The effectiveness of the repair was assessed by the tensile failure load of the repaired laminate. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

Residual Stress Characterization for Laminated Composites

This web site provides access to a University of Florida, PhD dissertation, by Shao-Chun Liu, dated 1999. The thesis describes the development of a technique called cure referencing method (CRM) for measuring the residual stresses in symmetric laminated composite plates. The chemical curing shrinkage of the polymer matrix was also investigated. A technique was developed to measure the post-gel chemical curing shrinkage. Bibliographic and abstract details are available in HTML format. The title page, contents and the full text of the document are accessible online in PDF format. This title is part of the University of Florida's Electronic Theses and Dissertations Project

Response and Failure of Composite Plates With a Bolt-Filled Hole : Interim Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-97/85, by U.M. Yan and others, dated June 1998. The report focuses on net-tension failure of composites containing a circular cutout with or without a mechanically tightened bolt. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

Review of Damage Tolerance for Composite Sandwich Airframe Structures : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/49, by J. Tomblin and others, dated August 1999. The report presents a review of previous damage tolerance investigations. It includes a compilation of damage tolerance certification procedures; a survey of past and current airframe industry sandwich constructions; and recommendations for future research. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

Scaled Composites LLC

Scaled Composites is an aerospace and specialty composites development company located in Mojave, California. Founded in 1982 by Burt Rutan, Scaled has a broad experience in air vehicle design, tooling, and manufacturing, specialty composite structure design, analysis and fabrication, and developmental flight test. The site provides brief descriptions about the company's products, expertise, and activities, including tilt-rotor and tilt-body UAV designs. The site also includes notices of career opportunities. There is also a link to short description on the areas of expertise namely: Conceptual Design; Aerodynamic Design; Structural Analysis and Design; Tooling; Fabrication; Structural Testing; and Flight Testing.

Selected Current Aerospace Notices

Provided by NASA, this service is an electronic current awareness journal which is published twice a month. The aim of the service is to alert users to recently published report and journal literature about aeronautics and aerospace research. It is divided into broad topic areas (aeronautics, astronautics, chemistry and materials, engineering, geosciences, life sciences, mathematical and computer sciences, physics, social sciences and space sciences) which are further subdivided into 191 specific subject topics to aid browsing and identification of relevant publications. The service is also searchable.

Smart Structures and Materials: Implications for Military Aircraft of New Generation

This is Research and Technology Organization(RTO) AGARD Lecture Series, AGARD-LS-205, dated October 1996. It is sponsored by the Advisory Group for Aerospace Research and Development of AGARD. Smart materials and structures technology is the integration of sensing and actuation elements into a structure or even more ambitiously into a material, with sensor and actuator being linked by a controller. Materials actually favored for integration include optical fibers and piezoelectric materials with respect to sensors, piezoelectric and electrostrictive materials, shape memory alloys or electro-rheological fluids with respect to actuators and microprocessors, neural networks, fuzzy logic and various types of signal processing with respect to control. The first part of the lecture series is mainly focussed on understanding the fundamentals of smart materials and structures technology and achieving the capability to judge the use of that technology with respect to individual applications. Presentations related to sensor and actuator materials, mechanics of smart structures, control and data processing, as well as structural integration of sensors, actuators, and generally electronics are therefore the focus of this part. In a second part, applications of smart structures technology are considered with respect to aircraft. Topics to be covered include monitoring the health/damage of aircraft structures or components, conceptual design of an adaptive wing, and electromagnetic antennae and their structural integration. For individual titles, see N97-11476 through N97-11487. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Smart Structures for Aircraft and Spacecraft

This is Research and Technology Organization (RTO) AGARD-CP-531, dated April 1993. An overview of the state-of-the-art of 'Smart Structures' technology as well as detailed descriptions of specific applications is presented. This technology offers extremely attractive advantages in the design, development, and operation of aerospace structures. For individual titles, see N94-11318 through N94-11347. Bibliographic and abstract details are available in HTML format. A table of contents, and the full text (117.93MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Some Aspects of the Mechanical Response of PMR-15 Neat Resin at 288 deg. C: Experiment and Modeling

This is the full text of a Master's thesis by 2nd Lieutenant Christina M. Falcone, USAF, AFIT/GAE/ENY/06-S03, which was presented to the Faculty Department of Aeronautics and Astronautics of Air University's Air Force Institute of Technology (AFIT), in September 2006. The mechanical response of PMR-15 neat resin was investigated at 288?C. Monotonic loading/unloading tests performed at several constant stress rates revealed considerable rate dependence, especially on the unloading path. Effect of prior stress rate on creep behavior was evaluated in creep tests preceded by uninterrupted loading to a target stress. Creep response was dependent on the prior stress rate. Effect of loading history was studied in stepwise creep tests, where specimens were subjected to a constant stress rate loading followed by unloading to zero stress with intermittent creep periods during both loading and unloading. Comparison of creep strains accumulated during a stepwise creep test to those accumulated during creep preceded by uninterrupted loading indicate that the prior stress history affects the creep behavior. A nonlinear viscoelastic model (Schapery?s formulation) was characterized using creep and recovery tests. The model was verified by comparing the predictions with experimental results obtained in monotonic loading/unloading, single-step, and multi-step creep tests. The model qualitatively predicted creep response to single- and multi-step creep tests, including negative creep and creep rate reversal during unloading. However, predictions were not quantitatively accurate. The model was unable to accurately predict the recovery behavior and could not account for rate effects. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

Structural Integrity of Discontinuous Stiffened Integrally Braided and Woven Composite Panels : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/24, by K.N. Shivakumar, M.J. Sundaresan, and V.S. Avva, dated March 1999. The report describes an investigation of the strength and failure modes of discontinuous blade stiffened panels made of textile preform composites using 3D finite element analysis and testing. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

Structural Testing and Analysis of Honeycomb Sandwich Composite Fuselage Panels

This provides access to a Federal Aviation Administration (FAA) technical report DOT/FAA/AR-08/51 written by Frank Leone ...[et al] and dated January 2009. This study investigated the damage tolerance characteristics and failure mechanisms of six honeycomb sandwich composite fuselage panels subjected to quasi-static pressurization and axial loading using the Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) fixture located at Federal Aviation Administration William J. Hughes Technical Center, Atlantic City International Airport, NJ. The FASTER fixture is capable of testing full-scale fuselage panel specimens under conditions representative of those seen by an aircraft in actual operation. The faster fixture is capable of applying pressurization, axial, hoop, and shear loads to a fuselage panel. The damage tolerance of composite sandwich panels with impact damage, holes, and notches under in-plane tensile and compressive loading was previously investigated at coupon and element-scale levels. A typical airframe predominantly experiences in-plane loads, although damaged regions may experience localized out-of-plane bending and bulging due to internal pressurization. The objective of these tests was to study the effects of holes and notches on the damage tolerance of full-scale, curved composite panels that reflect a typical sandwich fuselage structure subjected to combined loading. All six panels were loaded quasi-statically up to failure, studying damage growth and strain redistribution behavior with increasing load and recording the residual strength. The test articles were instrumented with strain gages near the damage and in the far-field regions for strain surveys. A digital image correlation method was used to obtain full-field displacement and strain measurements at equal load intervals and after any visible surface damage was observed. The acoustic emission method was used to monitor for damage growth in real time and served as an early warning for imminent failure. Several nondestructive inspection methods, including flash thermography and computer-aided tap testing, were used to scan for nonvisual damage. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Study of the of the Flammability of Commercial Transport Airplane Wing Fuel Tanks

This provides access to Federal Aviation Administration (FAA) technical report DOT/FAA/AR-08/08 written by William Cavage and Steven Summer dated February 2008. The Fire Safety Team of the Airport and Aircraft Safety Research and Development Division performed tests at the Federal Aviation Administration (FAA) William J. Hughes Technical Center using the environmental chamber and the air induction facility (wind tunnel) to examine individual effects that contribute to commercial transport wing fuel tank flammability. Additionally, previously acquired wing tank flammability measurements taken during flight tests were compared with the results from the FAA Fuel Air Ratio Calculator in an effort to see if the calculations agreed with existing flight test data. The results of the scale fuel tank testing in the environmental chamber showed that (1) fuel height in the tank had little or no effect on the flammability, (2) increasing the amount of heat on the top surface and a higher ambient temperature caused increased flammability, and (3) lower fuel flash point increased flammability greatly. Wind tunnel tests conducted with a section of a Boeing 727 wing tank showed that, under dynamic airflow conditions, change in ullage temperature was the primary mechanism affecting ullage flammability, not fuel temperature, as observed in environmental chamber tests. Other wind tunnel tests showed that the angle of attack of the fuel tank played little role in reducing fuel tank flammability, but that a cross-venting condition of the fuel tank would lead to a very rapid decrease in hydrocarbon concentration. An input temperature algorithm could be used with the FAA Fuel Air Ratio Calculator to significantly improve predictions of wing tank ullage flammability, based on tests that showed in-flight changes of ullage flammability in a wing tank are driven largely by the ullage temperature. This is very different from what had been shown with a center wing fuel tank, in which fuel temperature continues to be the main driver of flammability even during flight. [Taken from abstract]. The full text is available in PDF format so Adobe Acrobat software is required in order to read it.

Tabbing Guide for Composite Test Specimens

This technical report (DOT/FAA/AR-02/106) was produced by the Aviation Research Office of the Federal Aviation Administration (FAA) in October 2002 and was written by Daniel O. Adams and Donald F. Adams. This document provides guidelines for selecting suitable tabbing configurations for composite material text specimens. Additionally, a practical methodology is detailed for preparing and applying tabs to composite test specimens. This document is based on research performed within the Mechanical Engineering Department at the University of Utah and previously by the Composite Materials Research Group (CMRG) at the University of Wyoming, both sponsored by the Federal Aviation Administration. [Taken from abstract] This is a PDF file, so Adobe Acrobat software will be required in order to read it.

Tensile Stress Rupture Behavior of a Woven Ceramic Matrix Composite in Humid Environments at Intermediate Temperature

This is the full text of a Doctoral thesis by Major Kevin J. LaRochelle, USAF, AFIT/DS/ENY/05-01, which was presented to the Faculty Graduate School of Engineering and Management of Air University's Air Force Institute of Technology (AFIT), in March 2005. Stress rupture tests on the Sylramic(TM) fiber with an in-situ layer of boron nitride, boron nitride interphase, and SiC matrix ceramic matrix composite were performed at 550 degrees C and 750 degrees C with 0.0, 0.2, or 0.6 atm partial pressure of water vapor, pH(sub 2)O. The 550 degrees C, 100-hr strengths were 75%, 65% and 51% of the monotonic room temperature tensile strength, respectively. At 750 degrees C, the strengths were 67%, 51%, and 49%, respectively. Field Emission Scanning Electron Microscopy analysis estimated the total embrittlement times for 550 degrees C with 0.0, 0.2, and 0.6 atm pH(sub 2)O were >63 hrs, >38 hrs, and between 8 and 71 hrs, respectively. Corresponding estimated embrittlement times for the 750 degrees C were >83 hrs, between 13 and 71 hrs, and between 1 and 6 hrs. A time-dependent, phenomenological, Monte Carlo-type failure model was developed and simulated total embrittlement times that were within the experimentally determined range for all cases. Variation in the room temperature ultimate strength, the elevated temperature ultimate strength, and the fiber reference strength affected the model the most. Variation in the modulus of elasticity of the matrix and fiber affected it the least. Stress rupture strength degradation increases with temperature, moisture content level, and exposure time. [Taken from abstract]. The full text is available in PDF format on the Scientific and Technical Information Network (STINET) which is provided by the Defense Technical Information Center (DTIC).

The characterization and application of materials

This is Research and Technology Organization (RTO) AGARD-LS-51, dated May 1971. This is a lecture series edited by the Structures and Materials Panel and the Consultant and Exchange Programme of AGARD. The Lecture Series will begin with a discussion of the systems approach to the selection and application of materials, to be given by Dr Robert Maddin. The second in the series will be given by Dr Walter S.Owen and will be primarily concerned with the characterization, selection and use of high strength steels. The third lecture will be given by Dr Joseph Pask and will be concerned with the characterization, selection and uses of ceramic materials. The fourth in the series will be given by Professor Wippler and will cover the characterization, selection and use of polymeric materials. Dr Kelly will present the fifth lecture which will deal with characterization, selection and use of composite materials. The last lecture will cover two fields of special interest to aerospace: aluminium alloys and titanium, their characterization and selection aspects. This lecture will be divided into two parts and will be presented by Mr Syre for the titanium, and by Mr Tigeot for the aluminium alloys part of the paper. Bibliographic and abstract details are available in HTML format. A table of contents and the full text (25.52 MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

The Composite Materials Handbook

The site is maintained by the MIL-17 organisation in order to disseminate and share information on composite materials. The site is primarily concerned with supporting the use of MIL-HDBK-17. This handbook is jointly produced by the Department of Defence and the Federal Aviation Administration. It documents engineering methodologies for the development of standardised, statistically-based material property data for polymer matrix composites. It also provides selected guidance on materials selection, materials specification, material processing, design, analysis, quality control, and repair of typical polymer matrix composite materials. The site describes the content of the volumes, a user's forum, links to related documents, and instructions on how to obtain copies of the handbook. Please note that the text of MIL-HDBK-17/1-5 are available online from the Assist-Quick Search web site.

The Development of a Composite Landing Gear Component for a Fighter Aircraft

This technical report (NLR-TP-2002-020) was published by NLR (the National Aerospace Laboratory of the Netherlands) in 2002 and was written by H. G. S. J. Thuis. The use of Resin Transfer Moulding (RTM) as fabrication technique for structural components for the aerospace industry is increasing gradually. Although RTM moulds often are complex and expensive, RTM has several advantages compared to the autoclave fabrication method, which at this moment is the standard method used in the aerospace industry. One of these advantages is that complex shaped components can be made that would be very cumbersome or even impossible to make by autoclave processing. This means that designers now can design composite components as replacement of components made with metal forging. In the framework of a technology programme, a composite landing gear component for a fighter aircraft was developed as replacement of a steel component. The targets of the programme were to achieve a weight reduction of 20 % and a cost reduction of 15 %, which both were met. Several components were fabricated by RTM and tested successfully. In the paper a brief presentation is given of the design of the composite component, the RTM tooling concept and RTM set-up, and a brief overview of the test results. [Taken from abstract]. The full text is available as a PDF file.

The Development of Composite Landing Gear Components for Aerospace Applications

This technical report (NLR-TP-2004-141) was produced by NLR (the National Aerospace Laboratory of the Netherlands) in 2004 and was written by G. S. J. Bert. Composites are being used increasingly for structural components for aircraft and space applications because of their superior specific strength and stiffness properties in comparison aluminum and steel. The weight savings that were realized by applying composites used to be one of the main drivers to apply these materials. However, nowadays a reduction in fabrication cost is becoming important as well. The objective therefore is to combine new cost effective fabrication methods with lightweight structural concepts in order bring the exploitation of composite materials to a higher level. Up to now, the autoclave process is the standard fabrication technique to produce composite components for the aerospace industry. Recent developments show the evolution of new cost efficient fabrication techniques and of composite materials for these new techniques. One of these (for the aerospace) new fabrication methods is Resin Transfer Moulding (RTM). The RTM fabrication concept is based on the injection of resin into a mould cavity containing dry fibers (preform). During the injection process, air in the mould is being replaced by resin and the fibers are impregnated. Although RTM tooling can be complex and expensive, RTM has several advantages compared to autoclave processing. One of these advantages is that thick complex shaped components can be made that would be very cumbersome or even impossible to make by autoclave processing. This means that designers now can design composite components as replacement of components made with metal forging. In the framework of several technology programs the Structures Technology Department as part of the Aerospace Vehicles division of NLR developed several composite landing gear components for a large military helicopter and a fighter aircraft. These programs were carried out in close collaboration with the landing gear manufacturer SP aerospace and vehicle systems. The targets of the programs were to achieve not only weight reductions of 20% but also to reduce the manufacturing costs by 15% and to achieve a reduction in lead-time. Several different landing gear components were fabricated successfully by RTM and tested. All tested landing gear components failed beyond their required failure load levels. All program targets were met. The present paper will present an overview of the design concepts of these composite landing gear components. The RTM tooling concepts and the RTM manufacturing set-ups will be described and a brief overview of the test results will be given. Based on the results achieved, composite landing gear components are now considered to be feasible for application in next generation civil and military aircraft. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

The Effect of Loading Parameters on Fatigue of Composite Laminates : Part III : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-99/22, by H. Thomas Han, Milan Mitrovic, and Ozgur Turkgenc, dated June 1999. The report describes a study of the long-term mechanical fatigue of quasi-isotropic graphite / epoxy laminates to determine the influence of loading parameters on impact induced delamination growth during constant amplitude and spectrum fatigue loading. The full text of the report is available in PDF format from the online catalogue of the FAA William J. Hughes Technical Center Library.

The Effect of Loading Parameters on Fatigue of Composite Laminates : Part IV Information Systems : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/48, by H. Thomas Hahn and Ozgur Turkgenc, dated December 2000. This is the fourth in a continuing series of reports that provide a comprehensive study of the damage induced by spectrum fatigue loading in composite laminates and its influence on residual mechanical properties. This report focuses on the development of an information system to organise and retrieve damage tolerance and durability data. The full text of the report is available in PDF format

The Effect of Loading Parameters on Fatigue of Composite Laminates : Part V : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/24, by H. Thomas Han and Sung Won Choi, dated June 2001. The report examines the effects of load type, load level, load sequence, and spectrum modification on the damage growth of notched and visibly impact-damaged graphite/epoxy quasi-isotropic laminates. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library.

The Effect of Peel-Ply Surface Preparation Variables on Bond Quality

This technical report (DOT/FAA/AR-06/28) was published by the Federal Aviation Administration (FAA) Office of Aviation Research in August 2006 and was written by Brian Flinn and Molly Phariss. To further understand the effect of the variables present in surface preparation on the durability of cobonded and secondary-bonded composite joints, surface analysis coupled with mechanical testing and fractography were used to analyze samples prepared using peel-ply removal as the sole surface preparation technique. Laminates were made from aerospace carbon fiber prepreg and bonded with two different film adhesives. Nylon and polyester peel plies, and siloxane-coated release fabrics, were investigated to determine the effect of peel-ply material on bond quality and to examine why some peel-ply/adhesive systems are incompatible. Varying weaves of nylon and polyester peel plies were used to investigate the effect of peel-ply texture (weave) on bond quality. The moisture content of polyester peel plies was varied to examine its effect on bond quality. Laminate surfaces and peel plies were analyzed after peel-ply removal through scanning electron microscopy (SEM) and x-ray photoelectron spectroscopy (XPS); the results of which were correlated to Mode I double cantilever beam strain energy release rates to determine bond quality. Uncoated polyester peel plies were easily removed from laminate surfaces after curing and produced good bonds with both adhesives for all textures (GIC> 850 J/m2, cohesive failure). Super Release Blue-coated polyester peel ply created surfaces that bonded very poorly in both cases (GIC< 94 J/m2, adhesion failure), the result of the transfer of the peel-ply siloxane coating to the composite surface. Nylon peel plies were more difficult to remove from the laminate and, in the coarser weaves, could not be removed without damaging the laminates. Laminate surfaces prepared with nylon peel plies bonded well with AF555 (GIC> 750J/m2, cohesive failure). Laminate surfaces prepared with nylon peel plies bonded poorly with MB1515-3 (GIC<150 J/m2, adhesion failure). This may be explained by the transfer of nylon to the prepared surface, which was found during SEM and XPS analysis. Peel-ply texture or peel-ply moisture content had no significant effect on fracture energy or mode of failure. [Taken from abstract]. This is a PDF file, so Adobe Acrobat software will be required in order to read it.

The Effects of Advanced Materials on Airframe Operating and Support Costs

This provides access to RAND report DB-398-AF, prepared by Raj Raman, John C. Graser, and Obaid Younossi, dated 2003. The report has been published as part of the Project Air Force programme. The report examine whether airframe parts made of advanced materials, such as polymer composites and titanium, cost more to maintain than parts made of aluminium. A description of the report is available in HTML, while the full text of the document is available for browsing online and downloading in PDF (2.6 MB) format.

The Manufacture, Characterization and Aging of Novel High Temperature Carbon Fibre Composites

This web site provides access to an Australian National University PhD thesis, by Bronwyn Louise Fox, dated February 2001. This thesis has examined the effect of isothermal aging on two high temperature composite materials, a novel CSIRO composite and a commercial composite, both based on bismaleimides. Changes in mechanical properties and resin chemistry at two different temperatures were measured in order to assess the validity of accelerated aging tests. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format. This title is part of the Australian Digital Theses Program.

Thermal Decomposition of Cyanate Ester Resins : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-01/32, by Michael L. Ramirez, Richard Walters, Edward P. Savitski, and Richard E. Lyon, dated September 2001. The thermal decomposition chemistry of nine different polycyanurates was studied by thermogravimetry and infrared analysis of solid films and analysis of the gases evolved during pyrolysis using infrared spectroscopyand gas chromatography-mass spectrometry. The full text of the report is available in PDF format, from the online catalogue of the FAA William J. Hughes Technical Center Library

Thermophysical properties of solid materials - Cooperative thermal expansion measurements up to 1000 C. Project section 1A

This is Research and Technology Organization (RTO) AGARD-AR-31, dated March 1971. As a part of a major project concerning thermophysical properties of solid materials at high temperatures this is a report on the special section of this project dealing with thermal expansion up to temperatures of 1000°C. The complete project TX 44 is aimed at determining absolute accuracy of thermophysical property data of engineering materials at high temperatures by means of a cooperative measurement programme. The results are to indicate the best way of obtaining accurate data. This involves clarifying the following: (1) whether it is best to experimentally determine the properties of a material at the temperatures for which the data are needed, (2) whether known low temperature values on the material considered should be extrapolated to high temperatures, (3) whether it is better to derive estimated high temperature data of the material considered by a comparison with a similar material whose data are available from data compilations. Additionally the aim of the Project Section "Thermal Expansion up to 1OOO"C" is to determine the thermal expansion of solid materials up to 1000°C, which is considered conventionally as the most simple measurement within the realm of thermophysical property measurements. Customarily such measurements, especially in industrial laboratories, are performed by means of pushrod dilatometers, using both quartz glass and alumina reference systems. In as much as this type of apparatus does not yield absolute values and the accuracy of data obtained thereby is frequently considered in doubt, it is a special objective of this project section to determine the accuracy of this type of measurements by using the noble metals gold and platinum. A further objective is to investigate the thermal expansion behaviour of technical materials up to 1OOO"C, to find out if such materials can be used as calibration standards and to establish if an extension of the study on these materials to higher temperatures is justified. Bibliographic and abstract details are available in HTML format. A table of contents and the full text (6.02MB) of the document can be accessed online in PDF format. The document is contained in the RTO's Full Text Publication Library.

Two-Level Optimization of Composite Wing Structures Based on Panel Genetic Optimization

This web site provides access to a University of Florida, PhD dissertation, by Boyang Liu, dated 2001. Bibliographic and abstract details are available in HTML format. The title page, contents and the full text of the document are accessible online in PDF format. This title is part of the University of Florida's Electronic Theses and Dissertations Project.

Vacuum Assisted Resin Transfer Molding (VARTM) : Model Development and Verification

This is a Virginia Polytechnic Institute and State University Department of Engineering Science and Mechanics PhD dissertation, by Xiaolan Song, dated April 14, 2003. It describes the development and verification of a comprehensive Vacuum Assisted Resin Transfer Molding (VARTM) process simulation model. The model incorporates resin flow through the preform, compaction and relaxation of the preform, and viscosity and cure kinetics of the resin. The computer model can be used to analyze the resin flow details, track the thickness change of the preform, predict the total infiltration time and final fiber volume fraction of the parts, and determine whether the resin could completely infiltrate and uniformly wet out the preform. VARTM of two flat composite panels was conducted to verify the simulation model. The composite panels were fabricated using the SAERTEX multi-axial warp knit carbon fiber fabric and SI-ZG-5A epoxy resin. The simulation code was also used to investigate the VARTM of a new form of sandwich structure with through-the-thickness reinforcements, which is being considered for use in primary aircraft structure. The infiltration of three foam core sandwich preforms with different stitch densities was studied. The objective of the study was to determine whether the preforms could be completely infiltrated and how the stitch density affects the infiltration process. The visualization experiments were conducted to verify the simulation. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format [3.54 Mb]. This title is part of Virginia Tech’s Electronic Thesis and Dissertation Collection (VT ETD)

Verification of the Combined Load Compression (CLC) Test Method : Final Report

This provides access to a Federal Aviation Administration (FAA) report, DOT/FAA/AR-00/26, by Peter M. Wegner and Donald F. Adams, dated August 2000. The reports examines a Combined Loading Compression (CLC) test method for the determination of lamina compressive strength and modulus, which has been developed by the University of Wyoming. The full text of the report is available in PDF format, from the FAA's Office of Aviation Research web site.

Vibration Analysis of Cracked Composite Bending-torsion Beams for Damage Diagnosis

This is a Virginia Polytechnic Institute and State University Department of Mechanical Engineering PhD dissertation, by Kaihong Wang, dated , November 29, 2004. It describes the development of an analytical model of cracked composite beams vibrating in coupled bending-torsion. Based on the crack model, the aeroelastic characteristics of an unswept composite wing with an edge crack are investigated. The cracked composite wing is modelled by a cracked composite cantilever and the inertia coupling terms are included in the model. An approximate solution on critical flutter and divergence speeds is obtained by Galerkin’s method in which the fundamental mode shapes of the cracked wing model in free vibration are used. model-based crack detection (size and location) by changes in natural frequencies is addressed. The Cawley-Adams criterion is implemented and a new strategy in grouping frequencies is proposed to reduce the probability of measurement errors. Finally, sensitivity of natural frequencies to model parameter uncertainties is investigated. Uncertainties are modeled by information-gap theory and represented with a collection of nested sets. Five model parameters that may have larger uncertainties are selected in the analysis, and the frequency sensitivities to uncertainties in the five model parameters are compared in terms of two immunity functions. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format [1.88 Mb]. This title is part of Virginia Tech’s Electronic Thesis and Dissertation Collection (VT ETD)

Vibroacoustic Behavior and Noise Control Studies of Advanced Composite Structures

This is a University of Pittsburgh, School of Engineering PhD dissertation, by Deyu Li, defended July 16, 2003. The research presented in this thesis is devoted to the problems of sound transmission and noise transmission control for advanced composite payload fairings. There are two advanced composite fairings under study. The first is a tapered, cylindrical advanced grid-stiffened composite fairing, and the second is a cylindrical ChamberCore composite fairing. A fully coupled mathematical model for characterizing noise transmission into a finite elastic cylindrical structure with application to the ChamberCore fairing is developed. Structural-acoustic dynamic parameters of the two fairings are obtained using a combination of numerical, analytical, and experimental approaches. An in-situ method for experimentally characterizing sound transmission into the fairings called noise reduction spectrum (NRS) is developed based on noise reduction. The regions of interest in the NRS curves are identified and verified during a passive control investigation, where various fill materials are added into wall-chambers of the ChamberCore fairing. Both Helmholtz resonators (HRs) and long T-shaped acoustic resonators (ARs) are also used to successfully control noise transmission into the ChamberCore fairing. In the process, an accurate model for the resonant frequency calculation and design of cylindrical HRs is derived. Further, a novel and more general model for the design of multi-modal, long, T-shaped ARs is developed, including three new end-correction equations that are validated experimentally. Bibliographic and abstract details are available in HTML format. The full text of the document is accessible online in PDF format [2.26 Mb]. This title is part of the University of Pittsburgh's Electronic Thesis and Dissertation (ETD) Collection.

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