DOCSLIB.ORG

Crash Simulation of Fibre Metal Laminate Fuselage 2014

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Crash Simulation of Fibre Metal Laminate Fuselage 2014 CRASH SIMULATION OF FIBRE METAL LAMINATE FUSELAGE A THESIS SUBMITTED TO THE UNIVERSITY OF MANCHESTER FOR THE DEGREE OF DOCTOR OF PHILOSOPHY (PhD) IN THE FACULTY OF ENGINEERING AND PHYSICAL SCIENCES 2014 AHMAD SUFIAN ABDULLAH SCHOOL OF MECHANICAL, AEROSPACE AND CIVIL ENGINEERING Table of Contents Contents 2 List of Figures 6 List of Tables 9 Abstract 10 Declaration 11 Copyright Statements 12 Acknowledgement 13 Chapter 1 INTRODUCTION ................................................................................................. 14 1.1. Background and Motivation ............................................................................... 14 1.2. Aim of Research ................................................................................................. 16 1.3. Outline of Structure ........................................................................................... 16 Chapter 2 LITERATURE REVIEW ......................................................................................... 18 2.1 Crash Simulation of Aircraft ............................................................................... 18 2.1.1 Introduction of Aircraft Crash Simulation ................................................... 18 2.1.2 Methodology of Aircraft Crash simulation .................................................. 18 2.1.3 Crash Simulation of Composite Aircraft Fuselage ........................................ 21 2.2 Failure and Impact Response of Fibre Metal Laminate and its Constituents ...... 24 2.2.1 Introduction to Failure and Impact Response of Fibre Metal Laminate ....... 24 2.2.2 General Review on Mechanical Properties of Fibre Metal Laminate ........... 24 2.2.3 Bending and Buckling Behaviour of Fibre Metal Laminate .......................... 26 2.2.4 Impact Response and Damage of Fibre Metal Laminate under Low Velocity Impact …………………………………………………………………………………………………………………..27 2.2.5 Review on Failure and Impact Response of Fibre Reinforced Composite Laminate under Low Velocity Impact ......................................................................... 30 2.2.6 Review on Failure and Impact Response of Metal under Low Velocity Impact …………………………………………………………………………………………………………………..34 2.2.7 Finite Element Modelling of Impact and Damage on Fibre Metal Laminate and Its Constituents................................................................................................... 35 2.3 Summary of the Literature Review ..................................................................... 38 Chapter 3 BACKGROUND THEORY OF FIBRE METAL LAMINATE ......................................... 39 3.1 Introduction on Fibre Metal Laminate ................................................................ 39 3.1.1 GLARE: A Glass Fibre Reinforced Based Fibre Metal Laminate .................... 39 3.1.2 Impact Behaviour of GLARE ........................................................................ 40 3.2 Aluminium Alloys ............................................................................................... 42 3.2.1 Stress-strain Relationship of Isotropic and Homogeneous Materials ........... 42 3.2.2 Impact Behaviour of Aluminium Alloys Structure ........................................ 45 3.3 Fibre Reinforced Composite Laminate................................................................ 48 3.3.1 Stress-strain Relationships of Fibre Reinforced Composite Laminate .......... 48 3.3.2 Analysis of a Composite Lamina ................................................................. 49 3.3.3 Failure of Fibre-reinforced Composite Laminate ......................................... 52 3.3.4 Impact Behaviour of Fibre-Reinforced Composite Laminate under Low Velocity Impact ......................................................................................................... 56 Chapter 4 FINITE ELEMENT METHOD................................................................................. 59 4.1. Introduction....................................................................................................... 59 4.1.1. Introduction of Finite Element Method in Aircraft Crash Analysis ............... 59 4.1.2. General Description of Finite Element Method ........................................... 60 4.1.3. Abaqus Finite Element (FE) Software .......................................................... 61 4.2. Nonlinear Dynamic Analysis ............................................................................... 62 4.2.1. Nonlinear Analysis of Aircraft Structure ...................................................... 62 4.2.2. Dynamic Analysis of Aircraft Structure........................................................ 64 4.3. Selection of Elements for Discretisation ............................................................. 66 4.3.1. Shell element ............................................................................................. 66 4.3.2. Incompatible Mode Solid Element .............................................................. 69 4.3.3. Reduced Integration Element ..................................................................... 69 4.3.4. Hourglass Control ....................................................................................... 70 4.3.5. Cohesive Element ....................................................................................... 71 4.4. Material and Damage Model of Aluminium Alloy ............................................... 76 4.4.1. Material Model of Aluminium Alloy ............................................................ 76 4.4.2. Damage model of Aluminium Alloy ............................................................ 77 4.4.3. Onset of damage in Aluminium Alloy .......................................................... 77 4.4.4. Damage Evolution of Aluminium Alloy........................................................ 77 4.5. Material and Damage Model of Fibre-Reinforced Composite Laminate .............. 79 3 4.5.1. Material Model of Fibre-Reinforced Composite Laminate ........................... 79 4.5.2. Onset of damage in Fibre-Reinforce Composite Lamina .............................. 80 4.5.3. Damage Evolution of Fibre-Reinforced Composite Lamina .......................... 82 4.6. Interaction and Contact Modelling ..................................................................... 85 4.7. Constraint and Connection Modelling ................................................................ 88 4.7.1. Mesh Tie Constraints .................................................................................. 88 4.7.2. Mesh Independent Fasteners ..................................................................... 88 4.8. Computational Facilities in The University of Manchester .................................. 89 Chapter 5 DEVELOPMENT OF FIBRE METAL LAMINATE FUSELAGE CRASH MODEL ............. 90 5.1. Introduction of Aircraft Crash Methodology ....................................................... 90 5.2. Methodology of Crash Modelling of Fibre Metal Laminate Fuselage................... 91 5.3. Validation of Material and Damage Model Subjected to Impact Loading ............ 93 5.3.1. Validation of Aluminium Alloy Material and Damage Model ....................... 93 5.3.2. Validation of Composite Laminate Material and Damage Model .............. 102 5.4. Validation of General Impact Modelling ........................................................... 112 5.5. Verification of Fuselage Frame Impact Modelling ............................................. 119 5.5.1. Finite Element Modelling of Fuselage Frame ............................................ 119 5.5.2. Verification Results of Fuselage Frame Impact Model ............................... 122 5.6. Development of Crash Impact FE Model of Aluminium Alloy Fuselage Section . 126 5.6.1. Geometric Information and Assumptions ................................................. 126 5.6.2. Discretisation of the Fuselage Section ...................................................... 127 5.6.3. Material Assignment ................................................................................ 127 5.6.4. Impact and Contact Modelling ................................................................. 127 5.6.5. Location of Mass ...................................................................................... 128 5.7. Development of Crash Impact FE Model of GLARE Fuselage Section ................ 129 5.8. Evaluation of Acceleration Response at Floor-Level.......................................... 130 5.8.1. Data collection and processing of the acceleration response during crash event …………………………………………………………………………………………………………………130 5.8.2. Human tolerance towards acceleration .................................................... 130 Chapter 6 RESULTS AND DISCUSSIONS ............................................................................ 133 6.1. Introduction..................................................................................................... 133 6.2. Energy Dissipation during Crash ....................................................................... 133 6.3. Structural Deformation of Fuselage Structure .................................................. 141 6.4. Acceleration at Floor Level ............................................................................... 150 4 Chapter 7 Conclusions and Future Work.......................................................................... 153 7.1. Conclusions.....................................................................................................
Load more

Recommended publications
  • Testing Several Composite Materials in a Material Science Course Under the Engineering Technology Curriculum
    AC 2010-133: TESTING SEVERAL COMPOSITE MATERIALS IN A MATERIAL SCIENCE COURSE UNDER THE ENGINEERING TECHNOLOGY CURRICULUM N.M. Hossain, Eastern Washington University Dr. Hossain is an assistant professor in the Department of Engineering and Design at Eastern Washington University, Cheney. His research interests involve the computational and experimental analysis of lightweight space structures and composite materials. Dr. Hossain received M.S. and Ph.D. degrees in Materials Engineering and Science from South Dakota School of Mines and Technology, Rapid City, South Dakota. Jason Durfee, Eastern Washington University Professor DURFEE received his BS and MS degrees in Mechanical Engineering from Brigham Young University. He holds a Professional Engineer certification. Prior to teaching at Eastern Washington University he was a military pilot, an engineering instructor at West Point and an airline pilot. His interests include aerospace, aviation, professional ethics and piano technology. Page 15.1201.1 Page © American Society for Engineering Education, 2010 Testing Several Composite Materials in a Material Science Course under the Engineering Technology Curriculum Abstract The primary objective of a material science course is to provide the fundamental knowledge necessary to understand important concepts in engineering materials, and how these concepts relate to engineering design. In our institution, this course involves different laboratory performances to obtain various material properties and to reinforce students’ understanding to grasp the course objectives. As we are on a quarter system, this course becomes very aggressive and challenging to complete the intended course syllabus in a satisfactory manner within the limited time. It leaves very little time for students and instructor to incorporate thorough study any additional items such as composite materials.
    [Show full text]
  • The Residual Stress Characteristics and Mechanical Behavior of Shot Peened Fiber Metal Laminates Based on the Aluminium-Lithium Alloy
    Delft University of Technology The residual stress characteristics and mechanical behavior of shot peened fiber metal laminates based on the aluminium-lithium alloy Li, Huaguan; Wang, Hao; Alderliesten, René; Xiang, Junxian; Lin, Yanyan; Xu, Yingmei; Zhao, Haidan; Tao, Jie DOI 10.1016/j.compstruct.2020.112858 Publication date 2020 Document Version Final published version Published in Composite Structures Citation (APA) Li, H., Wang, H., Alderliesten, R., Xiang, J., Lin, Y., Xu, Y., Zhao, H., & Tao, J. (2020). The residual stress characteristics and mechanical behavior of shot peened fiber metal laminates based on the aluminium- lithium alloy. Composite Structures, 254, [112858]. https://doi.org/10.1016/j.compstruct.2020.112858 Important note To cite this publication, please use the final published version (if applicable). Please check the document version above. Copyright Other than for strictly personal use, it is not permitted to download, forward or distribute the text or part of it, without the consent of the author(s) and/or copyright holder(s), unless the work is under an open content license such as Creative Commons. Takedown policy Please contact us and provide details if you believe this document breaches copyrights. We will remove access to the work immediately and investigate your claim. This work is downloaded from Delft University of Technology. For technical reasons the number of authors shown on this cover page is limited to a maximum of 10. Green Open Access added to TU Delft Institutional Repository 'You share, we take care!' - Taverne project https://www.openaccess.nl/en/you-share-we-take-care Otherwise as indicated in the copyright section: the publisher is the copyright holder of this work and the author uses the Dutch legislation to make this work public.
    [Show full text]
  • Compressive Strength of the Glare Composite Laminates After Impact Load
    IJISET - International Journal of Innovative Science, Engineering & Technology, Vol. 2 Issue 3, March 2015. www.ijiset.com ISSN 2348 – 7968 Compressive Strength of the Glare Composite Laminates after Impact Load 1 2 P SanthoshP Kumar D, M.Tech , P Dr.R.AsokanP ,Professor, [email protected],Hindustan university. [email protected], Hindustan university ABSTRACT composites simultaneously. GLARE is widely used in aircraft skins. GLARE got some serious attentions in recent days due GLARE is a hybrid composite consisting of thin to its large amount usage in the skin of AIRBUS A-380. aluminium and glass/epoxy layers. GLARE has good impact resistance and has high resistance to fatigue loading.This paper In the past investigation of GLARE , some scholars like SeyedYaghoubi et al (2012) presented the experimental and presents the experimental behaviour of the GLARE and numerical investigations on ballistic impact behaviours of investigate its impact damage. The analysis involved for various GLARE 5 fibre-metal laminated (FML) beams of various thickness of the aluminium fibre. The thickness of the aluminum thicknesses and also analysed using 3D finite element (FE) sheet is 0.2.After investigated the laminate in impact testing it is code, LS-DYNA and compared the results.Alderliesten et al tested in the compression testing. (2012) calculated the impact loading using the combination of suitable properties of metals and fibre reinforced composites, Key words -GLARE,Impact test,Compression test used superior impact properties as well as considerable improvement in fatigue performance.The results for different 1.INTRODUCTION material is studied .Hamedahmadi et al( 2011) studied the 2/1 GLARE laminates that are manufactured and impacted by The aircraft industry is very conservative in the adoption of cylindrical projectiles at energies up to that required to new designs and technologies.
    [Show full text]
  • NI\SI\ ! F\1V\"VJN, VII(GINIA National Aeronautics and Space Administration Langley Research Center Hampton
    lII/JSlJ TIlf- glS7~ NASA Technical Memorandum 87572 .. NASA-TM-87572 19850023836 THERMAL EXPANSION OF SELECTED GRAPHITE REINFORCED POLYIMIDE-, EPOXY-, AND GLASS-MATRIX COMPOSITE Stephen S. Tompkins JULY 1985 \ '~-·"·'·1 LI~HARY COpy AUG 1 91005 I i,l ,:.,1 II I,Lsc"nCH CUJ I FII lIHr~ARY, NlISA NI\SI\ ! f\1V\"VJN, VII(GINIA National Aeronautics and Space Administration Langley Research center Hampton. Virginia 23665 '1 Lli':'lIT.22./84-:S5· 1 RN/NASA-TM-87572 [1 ISPLA\i 251,,/·~:! ../'l 85N321"49*# ISSUE 21 PAGE 3561 CATEGORY 24 RPT#: NASA-TM-87572 NAS 1.15:87572 85/07/00 20 PAGES UNCLASSIFIED DOCUMENT t~)fTL~ Tt-"ierrfi,31 e~:<p,3rr3il)(l {)f ~3elec·-ted ~9r"cH:it-'tit2 (e·irlr(;;t"c:eci l::tol'ilrni(ie-, EtJ{)){'l-, ar\c! glass-matrix COffiPostte . I~UTH: A.·/·TCit:,tlPK If\!:=;, ~; •.s. LvRP: National. Aeronautics and Space Administration. LangleY Rese~rch Center, Hampton,V~. AVAIL. NTIS SAP: He A02/MF AOl p\R·._.IS ~ /"::'::8()F~CI:31 LI C~ATE (;LASS ....... ~·::C:AR.8(Jr\1 FI BEF~:3 .../:,~~!~F:AF!H ITE -PCiL\tI t~'l ILiE COf'·'lPf):; ITES.·./::-:; T~ERMAL EXPANSION (ftiNS: ! 'OIMENSIONAL STABILITY; LAMINATES/ RESIDUAL STRESS HbA: At!thCj( j J J J J J J J J J J J J J J J J J J J J J J J J J J J SUMMARY .The thermal expansion of three epoxy-matrix composites, a polymide-matrix composite and a borosilicate glass-matrix composite, each reinforced with con­ tinuous carbon fibers, has been measured and compared.
    [Show full text]
  • Comparative Study of Tensile and Flexural Behaviour for Glass-Fiber- Reinforced Aluminium (Glare) Laminates and Aluminum
    Asian Journal of Computer Science and Technology (AJCST) Vol1.No.1 2013 pp 5-13. available at: www.goniv.com Paper Received :19-02-2013 Paper Published:20-03-2013 Paper Reviewed by: 1. John Arhter 2. Hendry Goyal Editor : Prof. P.Muthukumar COMPARATIVE STUDY OF TENSILE AND FLEXURAL BEHAVIOUR FOR GLASS-FIBER- REINFORCED ALUMINIUM (GLARE) LAMINATES AND ALUMINUM Asha Melba.V, Senthil Kumar.A, Vino.A # Department of CAD /CAM Engineering Sethu Institute of Technology, pulloor-626 115. [email protected] [email protected] [email protected] ABSTRACT The objective of this research article was investigated to evaluate tensile and flexural properties of Glass-fiber-Reinforced Aluminium (Glare) laminates as well as Aluminium sheets of same thickness. In addition to that the tensile and flexural strength of Glare laminates were compared with values of aluminium sheets. In this article, the aluminium based FMLs were fabricated using hand layup technique and cut as per ASTM standards. Three types of layers such as 2/1 Glare, 3/2 Glare and 5/4 Glare laminates were prepared. Computer controlled UTM machine used to determine the tensile, flexural properties and failure mode of the Glare laminates and Plain aluminium sheets with same thickness. From the test results the graphs were plotted for load vs displacement, tensile and flexural strength vs layers thickness. It shows that tensile and flexural strength of Glass-fiber-Reinforced Aluminium (Glare) laminates purely depend on the volume percentage of fibre and it exhibits improvement over the properties of aluminium sheets. Keywords: Glass-fiber-Reinforced Aluminium laminates (Glare), aluminium sheet, hand lay-up technique, tensile property and flexural property.
    [Show full text]
  • Tensile Fracture Strength of Boron (Sae- 1042)/Epoxy/Aluminum (5083 H112) Laminates
    International Journal of Mechanical And Production Engineering, ISSN: 2320-2092, Volume- 5, Issue-2, Feb.-2017 http://iraj.in TENSILE FRACTURE STRENGTH OF BORON (SAE- 1042)/EPOXY/ALUMINUM (5083 H112) LAMINATES 1RAVI SHANKER V CHOUDRI, 2S C SONI, 3A.N. MATHUR 1,2Mewar University, Chittorgarh, Rajasthan, India. 3Former Dean, College of Technology and Engineering, MPUAT, Udaipur, Rajasthan, India Abstract- Recognizing the mechanical property of a component is tremendously important, before it is used as an purpose to keep away from any failure. Composite structures put into practice can be subjected to various types of loads. One of the main loads among such is tensile load. This paper shows experimental study and results of the flat rectangular dog bone tensile specimen of Boron (SAE-1042)-Epoxy/Aluminum (5083 H112) (B/Ep/Al) laminated metal composite (LMC). LMCs are a single form of composite material in which alternating metal or metal containing layers are bonded mutually with separate interfaces. Boron metal is among the hardest materials from the earth’s crest coupled with the aluminum metal, it is known that one side of boron and aluminum plates are appropriately knurled and later adhesively bonded with an epoxy which acts as binder and consolidated by using heavy upsetting press. Processes of tensile tests have been carried out with a hydraulic machine where three specimens with the same geometry were tested which all of them showing similar stress- strain curves. These results of tensile tests were evidently indicated the nature of ductility. Keywords- FML, GLARE, LMC, Boron (SAE-1042), Epoxy, Aluminum (5083 H112), Aluminum (6061-t6), Aluminum (6083-t6), LMC, BEA, tensile behavior, Stress- strain curve.
    [Show full text]
  • Toray TC275-1E
    Toray TC275-1E PRODUCT DATA SHEET DESCRIPTION TC275-1E is an extended out time version of the TC275-1 resin system. TC275-1E is a flexible cure toughened epoxy prepreg designed to enable composite part construction with low pressure or vacuum bag only (VBO) cures. The resin system provides a 28-day total out time to facilitate construction of thick or larger composite structures. TC275-1E may be cured at temperatures from 135°C (275°F) to 177°C (350°F) for higher temperature service. FEATURES f Robust OOA/VBO system f Capable of freestanding post cure for higher Tg f Flexible cure prepreg system f Excellent resistance to hot/wet exposure f Long out time and tack life for shop floor handling f High toughness PRODUCT TYPE TYPICAL NEAT RESIN PROPERTIES 135–177°C (275–350°F) Cure Toughened Epoxy Resin System Density 1.17 g/cc TYPICAL APPLICATIONS Gel Time at 177°C (350°F) 9–25 min f Aircraft structures f Launch and space structures Tg by DMA after 6 hours at 135°C (275°F) 168°C (334°F) f Thick parts cured under low pressure f Honeycomb stiffened parts UD tape Compression and Open-Hole Compression Strengths of IM7 12K, 150gsm/TC275-1E, 35% RC* SHELF LIFE MPa Out Life: Up to 28 days at ambient 1600 1400 Frozen Storage Life: 12 months at -18°C (< 0°F) 1200 1000 800 600 Out life is the maximum time allowed at 21°C (70°F) or below 400 200 and 60% or less RH before cure, after a single frozen storage 0 cycle in the original unopened packaging at -18°C (0°F) or CS OHC below for a period not exceeding the frozen storage life noted RTD ETD ETW above.
    [Show full text]
  • Composite Laminate Modeling
    Composite Laminate Modeling White Paper for Femap and NX Nastran Users Venkata Bheemreddy, Ph.D., Senior Staff Mechanical Engineer Adrian Jensen, PE, Senior Staff Mechanical Engineer Predictive Engineering Femap 11.1.2 White Paper 2014 WHAT THIS WHITE PAPER COVERS This note is intended for new engineers interested in modeling composites and experienced engineers who would like to get acquainted with the Femap interface. This note is intended to accompany a technical seminar and will provide you a starting background on composites. The following topics are covered: o A little background on the mechanics of composites and how micromechanics can be leveraged to obtain composite material properties o 2D composite laminate modeling Defining a material model, layup, property card and material angles Symmetric vs. unsymmetric laminate and why this is important Results post processing o 3D composite laminate modeling Defining a material model, layup, property card and ply/stack orientation When is a 3D model preferred over a 2D model o Modeling a sandwich composite Methods of modeling a sandwich composite 3D vs. 2D sandwich composite models and their pros and cons o Failure modeling of a 2D composite laminate Defining a laminate failure model Post processing laminate and lamina failure indices Predictive Engineering Document, Feel Free to Share With Your Colleagues Page 2 of 66 Predictive Engineering Femap 11.1.2 White Paper 2014 TABLE OF CONTENTS 1. INTRODUCTION ...........................................................................................................................................................
    [Show full text]
  • HPC Sept 02 Airbus A380
    Commercial Aircraft Extra large by any measure Advanced composites play a big role in the success of the world’s largest commercial aircraft. t 555 passengers and a maximum take-off According to Airbus’ current plans, the A380 weight of 1,235,000 lb (560 metric tons), will carry 30 metric tons/66,000 lb of structural A“huge” is the most appropriate word to composites, primarily of carbon-fiber/epoxy, or 16 describe the Airbus A380-800, the world’s largest percent of its airframe weight (approx. 170 metric jet airliner, set to begin regular service in 2006. tons), making it the most composite-intensive The plane itself is an enormous achievement, but commercial aircraft to date. Due to the higher so is its impact on the worlds of commercial avia- stiffness-to-weight performance of carbon fiber tion, advanced materials and composite manufac- composites, this is equivalent to the replacement turing technologies. Over a decade in the planning of 20 percent of conventional aluminum struc- and design phase, the first flying aircraft is cur- ture. This figure could rise to 35 metric tons (77,000 lb) as the final component designs near completion. Wing leading edges will take advan- tage of economies realized in the use of glass-rein- forced thermoplastics. And 4 percent of the air- Source: Airbus Industrie frame will be GLARE (GLAss fiber-REinforced aluminum), a multi-layer laminate of fiberglass/epoxy and aluminum to be used in the upper fuselage panels (discussed below and in HPC May/June 1996, p. 28). Beyond the structural composites under discussion here, as many as 30 metric tons of composites, mainly fiberglass/phe- nolic, may be used in each plane’s interior.
    [Show full text]
  • Ballistic Impact Response of Ceramic-Faced Aramid Laminated Composites Against 7.62 Mm Armour Piercing Projectiles
    Defence Science Journal, Vol. 63, No. 4, July 2013, pp. 369-375, DOI : 10.14429/dsj.63.2616 2013, DESIDOC Ballistic Impact Response of Ceramic-faced Aramid Laminated Composites Against 7.62 mm Armour Piercing Projectiles N. Nayak*, A. Banerjee, and P. Sivaraman# Proof and Experimental Establishment, Chandipur-756 025, India #Naval Materials Research Laboratory, Ambernath-421 506, India *E-mail: [email protected] ABSTRACT Ballistic impact response of ceramic- composite armor, consisting of zirconia toughened alumina (ZTA) ceramic front and aramid laminated composite as backing, against 7.62 mm armor piercing (AP) projectiles has been studied. Two types of backing composite laminates i.e. Twaron-epoxy and Twaron-polypropylene (PP) of 10 mm and 15 mm thickness were used with a ceramic face of 4mm thick ZTA. The ceramic- faced and the stand alone composite laminates were subjected to ballistic impact of steel core 7.62 mm AP projectiles with varying impact velocities and their V50 ballistic limit (BL) was determined. A sharp rise in BL was observed due to addition of ceramic front layer as compared to stand alone ones. The impact energy was absorbed during penetration primarily by fracture of ceramic, deformation and fracture of projectile and elastic-plastic deformation of flexible backing composite layer. The breaking of ceramic tiles were only limited to impact area and did not spread to whole surface and projectile shattering above BL and blunting on impact below BL was observed. The ceramic- faced composites showed higher BL with Twaron-PP as backing than Twaron-epoxy laminate of same thickness. This combination of ceramic-composite laminates exhibited better multi-hit resistance capability; ideal for light weight armor.
    [Show full text]
  • Probabilistic Modeling of Ceramic Matrix Composite Strength
    NASA / TM--1998-208492 Probabilistic Modeling of Ceramic Matrix Composite Strength Ashwin R. Shah Sest, Inc., North Royalton, Ohio Pappu L.N. Murthy Lewis Research Center, Cleveland, Ohio Subodh K. Mital University of Toledo, Toledo, Ohio Ramakrishna T. Bhatt U.S. Army Research Laboratory, Lewis Research Center, Cleveland, Ohio Prepared for the 39th Structures, Structural Dynamics, and Materials Conference and Exhibit cosponsored by AIAA, ASME, ASCE, AHS, and ASC Long Beach, California, April 20-23, 1998 National Aeronautics and Space Administration Lewis Research Center September 1998 Available from NASA Center for Aerospace Information National Technical Information Service 7121 Standard Drive 5287 Port Royal Road Hanover, MD 21076 Springfield, VA 22100 Price Code: A03 Price Code: A03 PROBABILISTIC MODELING OF CERAMIC MATRIX COMPOSITE STRENGTH Ashwin R. Shah Sest, Inc. North Royalton, Ohio 44133 Pappu L.N. Murthy National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135 Subodh K. Mital 1 The University of Toledo Toledo, Ohio 43606 Ramakrishna T. Bhatt U.S. Army Research Laboratory Lewis Research Center Cleveland, Ohio 44135 ABSTRACT Uncertainties associated with the primitive random variables such as manufacturing process (processing temperature, fiber volume ratio, void volume ratio), constituent properties (fiber, matrix and interface), and geometric parameters (ply thickness, interphase thickness) have been simulated to quantify the scatter in the first matrix cracking strength (FMCS) and the ultimate tensile strength of SCS-6/RBSN (SiC fiber (SCS-6) reinforced reaction-bonded silicon nitride composite) ceramic matrix composite laminate at room temperature. Cumulative probability distribution function for the FMCS and ultimate tensile strength at room temperature (RT) of [0] 8, 102/ 902] s, and [+452] s laminates have been simulated and the sensitivity of primitive variables to the respective strengths have been quantified.
    [Show full text]
  • Effect of Boric Acid Content in Aluminosilicate Matrix On
    materials Article Effect of Boric Acid Content in Aluminosilicate Matrix on Mechanical Properties of Carbon Prepreg Composites Eliška Haincová * and Pavlína Hájková Unipetrol Centre for Research and Education, Revoluˇcní 84, 40001 Ústí nad Labem, Czech Republic; [email protected] * Correspondence: [email protected]; Tel.: +420-475-309-233 Received: 30 October 2020; Accepted: 25 November 2020; Published: 27 November 2020 Abstract: This work presents carbon fabric reinforced aluminosilicate matrix composites with content of boric acid, where boron replaces aluminum ions in the matrix and can increase the mechanical properties of composites. Five different amounts of boric acid were added to the alkaline activator for preparing six types (including alkaline activator without boric acid) of composites by the prepreg method. The influence of boric acid content in the matrix on the tensile strength, Young’s modulus and interlaminar strength of composites was studied. Attention was also paid to the influence of boron content on the behavior of the matrix and on the internal structure of composites, which was monitored using a scanning electron microscope. The advantage of the aluminosilicate matrix is its resistance to high temperatures; therefore, tests were also performed on samples affected by temperatures of 400–800 ◦C. The interlaminar strength obtained by short-beam test were measured on samples exposed to 500 ◦C either hot (i.e. measured at 500 ◦C) or cooled down to room temperature. The results showed that the addition of boron to the aluminosilicate matrix of the prepared composites did not have any significant effect on their mechanical properties. The presence of boron affected the brittleness and swelling of the matrix and the differences in mechanical properties were evident in samples exposed to temperatures above 500 ◦C.
    [Show full text]