advances repair metallic aircraft structure volume2 docx

ADVANCES IN THE BONDED COMPOSITE REPAIR OF METALLIC AIRCRAFT STRUCTURE Volume 2 Editors A.A.. More recently, he is recognised for pioneering work on bonded composite repair of metallic

Bonded Comp Repair o f Metallic Aircraft Structure

VOLUME 2

A

7

ADVANCES IN THE BONDED COMPOSITE

STRUCTURE Volume 2

Elsevier Science Internet Homepage - http://www.elsevier.com

Consult the Elsevier homepage for full catalogue information on all books, journals and electronic products and

services

Elsevier Titles of Related Interest

VALERY V VASILEV & EVGENY V MOROZOV

Mechanics and Analysis of Composite Materials

ISBN: 0 08 042702 2

JANG-KYO KIM & YIU WING MA1

Engineered Interfaces in Fiber Reinforced Composites

ISBN: 0 08 042695 6

J.G WILLIAMS & A PAVAN

Fracture of Polymers, Composites and Adhesives

ISBN: 0 08 043710 9

D.R MOORE, A PAVAN & J.G WILLIAMS

Fracture Mechanics Testing Methods for Polymers Adhesives and Composites

ISBN: 0 08 043689 7

Related Journals:

Composite Structures - www.elsevier.com/locate/compstruct

Composites Part A: Applied Science and Manufacturing - www.elsevier.com/locate/compositesa

Composites Part B: Engineering - www.elsevier.com/locate.compositesb

Composites Science and Technology - www.elsevier.com/locate.compscitech

Major Reference Work

Comprehensive Composite Materials - www.elsevier.com/locate/isbn/0080429939

To contact the Publisher

Elsevier Science welcomes enquiries concerning publishing proposals: books, journal special issues, conference proceedings, etc All formats and media can be considered Should you have a publishing proposal you wish to

discuss, please contact, without obligation, the publisher responsible for Elsevier's Composites and Ceramics programme:

Emma Hurst

Assistant Publishing Editor

Elsevier Science Ltd

OX5 IGB, UK E.mail: e.hurst@elsevier.com

General enquiries, including placing orders, should be directed to Elsevier's Regional Sales Offices -please access the Elsevier homepage for full contact details (homepage details at the top of this page)

Book Butler logo to search for more Elsevier books, visit the Books Butler at http://www.elsevier.com/homepage/

booksbutler/

ADVANCES IN THE BONDED COMPOSITE REPAIR OF METALLIC AIRCRAFT

STRUCTURE Volume 2

Editors

A.A Baker

Defence Science and Technology Organisation,

Air Vehicles Division, Victoria, Australia

L.R.F Rose

Department of Defence, Defence Science and Technology Organisation,

Air Vehicles Division, Victoria, Australia

Amsterdam - Boston - London - New York - Oxford - Paris

San Diego San Francisco - Singapore Sydney Tokyo

The Boulevard, Langford Lane

Kidlington, Oxford OX5 IGB, U K

0 2002 Elsevier Science Ltd All rights reserved

This work is protected under copyright by Elsevier Science, and the following terms and conditions apply to its use:

Photocopying

Single photocopies of single chapters may be made for personal use as allowed by national copyright laws Permission of the Publisher and payment of a fee is required for all other photocopying, including multiple or systematic copying, copying for advertising or promotional purposes, resale, and all forms of document delivery Special rates are available for educational institutions that wish to make photocopies for non-profit educational classroom use

Permissions may be sought directly from Elsevier Science via their home page (httD.iiwww.elsevler.com),

by selecting ‘Customer support’ and the ‘Permissions’ Alternatively you can send an e-mail to:

permissions~eelsevier.co.uk, or fax to: (+44) 1865 853333

In the USA, users may clear permissions and make payments through the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, M A 01923, U S A ; phone (+I) 978 750 8400, fax: (+I) 978 750 4744, and in the U K through the Copyright Licensing Agency Rapid Clearance Service (CLARCS), 90 Tottenham Court Road, London

W I P OLP; phone (+44) 207 63 1 5555; fax: (+44) 207 63 1 5500 Other countries may have a local reprographic rights agency for payments

Derivative Works

Tables of contents may be reproduced for internal circulation, but permission of Elsevier Science is required for external resale or distribution of such material

Permission of the Publisher is required for all other derivative works, including compilations and translations

Electronic Storage or Usage

Permission of the Publisher is required to store or use electronically any material contained in this work, including any chapter or part of a chapter

Except as outlined above, no part of this work may be reproduced, stored in a retrieval system or transmitted in any

form or by any means, electronic, mechanical, photocopying, recording or otherwise, without written permission of the Publisher

Address permission requests to: Elsevier Science Global Rights Department, a t the mail, fax and email addresses note above

Notice

N o responsibility is assumed by the Publisher for any injury and/or damage to persons or property as a matter of products liability, negligence or otherwise, or from any use or operation of any methods, products, instructions or ideas contained in the material herein Because of rapid advances in the medical sciences, in particular, independent verification of diagnosis and drug dosages should be made

First Edition 2002

Library of Congress Cataloging in Publication Data

A catalog of record from the Library of Congress has been applied for

British Library Cataloguing in Publication Data

A catalogue record from the British Library has been applied for

ISBN: 0-08-042699-9

@ The paper used for this publication meets the requirements of ANSIiNISO 239.48-1992 (Permanence of Paper) Printed in The Netherlands

BIOGRAPHIES

Dr Alan Baker

Dr Alan Baker is Research Leader Aerospace Composite Structures, in Airframes and Engines Division, Defence Science and Technology (DSTO), Aeronautical and Maritime Research Laboratory and Technical Adviser to the Cooperative Research Centre-Advanced Composite Structures, Melbourne Australia He is a Fellow of the Australian Academy of Technological Sciences and Engineering and an Adjunct Professor in Department of Aerospace Engineering, Royal Melbourne Institute of Technology Dr Baker is a member of the International Editorial Boards of the Journals Composites Part A Applied Science and Manufacturing, Applied Composites and International Journal of Adhesion and Adhesives

He is recognised for pioneering research work on metal-matrix fibre composites while at the Rolls Royce Advanced Research Laboratory More recently, he is recognised for pioneering work on bonded composite repair of metallic aircraft components for which he has received several awards, including the 1990 Ministers Award for Achievement in Defence Science

Dr Francis Rose

Dr Francis Rose is the Research Leader for Fracture Mechanics in Airframes and Engines Division, Defence Science and Technology (DSTO), Aeronautical and Maritime Research Laboratory He has made important research contributions in fracture mechanics, non-destructive evaluation and applied mathematics In particular, his comprehensive design study of bonded repairs and related crack- bridging models, and his contributions to the theory of transformation toughening

in partially stabilised zirconia, have received international acclaim His analysis of laser-generated ultrasound has become a standard reference in the emerging field of laser ultrasonics, and he has made seminal contributions to the theory of eddy- current detection of cracks, and early detection of multiple cracking

He is the Regional Editor for the International Journal of Fracture and a member

of the editorial board of Mechanics of Materials He was made a Fellow of the Institute of Mathematics and its Applications, UK, in 1987, and a Fellow of the Institution of Engineers, Australia, in 1994 He is currently President of the Australian Fracture Group, and has been involved in organising several local and international conferences in the areas of fracture mechanics and engineering mathematics He currently serves on the Engineering Selection Panel of the Australian Research Council and of several other committees and advisory bodies

vi Biographies

Professor Rhys Jones

Professor Rhys Jones joined Monash University in early 1993 and is currently Professor of Mechanical Engineering, and Head of the Defence Science and Technology Organisation Centre of Expertise on Structural Mechanics Professor Jones’ is best known for his in the fields of finite element analysis, composite repairs and structural integrity assessment Professor Jones is the Founding Professor of both the BHP-Monash Railway Technology Institute and the BHP-Monash Maintenance Technology Institute He is heavily involved with both Australian and overseas industry In this context he ran the mechanical aspects of the Australian Governments Royal Commission into the failure at the ESSO plant in

Victoria, and the Tubemakers-BHP investigation into the failure of the McArthur River gas pipe line in the Northern Territory

He is the recipient of numerous awards including the 1982 (Australian) Engineering Excellence Award, for composite repairs to Mirage 111, the Institution

of Engineers Australia George Julius Medal, for contributions to failure analysis, a TTCP Award, for contributions to Australian, US, UK, Canada and NZ Defence Science in the field of composite structures, and a Rolls-Royce-Qantas Special Commendation, for his work on F-111 aircraft Since 1999 Professor Jones has been Co-Chair of the International Conference (Series) on Composite Structures

Acknowledgement

The editors are very pleased to acknowledge their appreciation of the great assistance provided by Drs Stephen Galea and Chun Wang of the Defence Science and Technology Organisation, Aeronautical and Maritime Research Laboratory, who made important contributions, in the collation and editing of this book

FOREWORD

The introduction of the technology for bonded composite repairs of metallic airframe structures could not have come at a more opportune time Today, many countries are facing the challenge of aging aircraft in their inventories These airframes are degrading due to damage from fatigue cracking and corrosion Repair with dependable techniques to restore their structural integrity is mandatory The concept of using bonded composite materials as a means to

maintain aging metallic aircraft was instituted in Australia approximately thirty years ago Since that time it has been successfully applied in many situations requiring repair These applications have not been limited to Australia Canada, the United Kingdom, and the United States have also benefited from the use of this technology The application for the solution of the problem of cracking in the fuel drain holes in wing of the C-141 is credited with maintaining the viability of this fleet

The concept for composite repair of metallic aircraft is simple The bonded repair reduces stresses in the cracked region and keeps the crack from opening and therefore from growing This is easy to demonstrate in a laboratory environment It

is another thing to do this in the operational environment where many factors exist that could adversely affect the repair reliability The researchers at the Aeronautical and Maritime Research Laboratory in Australian realized there were many obstacles to overcome to achieve the desired reliability of the process They also realized that failures of the repair on operational aircraft would mean loss of confidence and consequently enthusiasm for the process They proceeded slowly Their deliberate approach paid off in that they developed a process that could be transitioned to aircraft use by engineers and technicians The essential ingredient for application of this technology is discipline When the applicator of this process maintains the discipline required for the process and stays within the bounds of appropriate applications, then the repair will be successful

This book, edited by Drs A.A Baker, L.R.F Rose and R Jones, includes the essential aspects of the technology for composite repairs The editors have chosen some of the most knowledgeable researchers in the field of bonded repairs to discuss the issues with the many aspects of this technology Included are discussions

on materials and processes, design of repairs, certification, and application considerations These discussions are sufficiently in-depth to acquaint the reader with an adequate understanding of the essential ingredients of the procedure The application case histories are especially useful in showing the breadth of the possible uses of the technology

vii

It is easy to be excited about the future of composite repairs to metallic airframes It has all the ingredients for success Today’s applications have shown that it is reliable, there is typically a significant return on the investment, and it can

be transitioned to potential users Additional research will open up possible new applications

This book is intended to provide the reader with a good understanding of the basic elements of this important technology It fulfills that purpose

John W Lincoln

Technical Adviser for Aircraft Structural Integrity

United States Air Force

It is rare to find in science and engineering, such a giant in the field who was so

modest, approachable and friendly Jack was regarded both as a supportive father figure and the expert to be convinced on all airworthiness issues, particularly as related to the USAF

ix

DEFAULT NOMENCLATURE

Boron/epoxy

Shear modulus (also used for

Characteristic crack length

Strain Displacement Thickness Applied load Force per unit width Stiffness ratio Thermal expansion coefficient Temperature range

0

I

*

xi

The case for adhesively bonded repairs

Composite versus metallic patches

Design and certification of airframe structures

Problems with ageing metallic airframe components

Chapter 2 Materials Selection and Engineering

2.1.1 Factors affecting adhesion

Materials for patches and reinforcements

Primers and coupling agents

Adhesive and composite test procedures

Materials engineering considerations

3.4.4 Bond durability model

Requirements of surface preparation

3.6.1 Factors controlling bondline thickness

3.6.2 Void formation and minimisation

Surface treatment quality control

3.7.1 Waterbreak Test

3.7.2 Surface work function methods

3.7.3 Fourier transform infrared spectroscopy

Bondline pressurisation and adhesive cure

Standards and environments for adhesive bonding

Qualification of bonding procedures and performance

Fracture mechanics and the cleavage specimen

Surface roughness and bond durability

Surface hydration and bond durability

Surface contamination and bond durability

Creation of a high energy surface oxide

Process control coupons (traveller or witness specimens) Practitioner education, skill and standards

On-aircraft acid anodisation and acid etch processes

Sol-Gel technology for adhesive bonding Hot solution treatment for adhesive bonding

Contents

Chapter 4 Adhesives Characterisation and Data Base

P Chalkley and A.A Baker

Mode I1 and mixed mode

In situ shear stress-strain allowables

Fickean diffusion coefficients for moisture absorption

Chapter 5 Fatigue Testing of Generic Bonded Joints

P.D Chalkley, C.H Wang and A.A Baker

5.3.2 Experimental method and results

5.3.3 Fracture mechanics approach

5.4 Discussion

References

Damage-tolerance regions in a bonded repair

The generic design and certification process

Stress state in the DOFS

5.2 The DOFS

5.3 The skin doubler specimen

Stress state in the skin doubler specimen

Chapter 6 Evaluating Environmental Effects on Bonded Repair Systems

Using Fracture Mechanics

L.M Butkus, R.V Valentin and W.S Johnson

xvi Contents

Chapter 7 Analytical Methods for Designing Composite Repairs

L.R.F Rose and C.H Wang

Formulation and notation

Load transfer of bonded reinforcement

Symmetric repairs

7.4.1 Stage I: Inclusion analogy

7.4.2

7.4.3 Plastic adhesive

7.4.4 Finite crack size

7.4.5 Finite element validation

Shear mode

One-sided repairs

7.6.1 Geometrically linear analysis

7.6.2 Crack bridging model

7.6.3 Geometrically non-linear analysis

Residual thermal stress due to adhesive curing

Stage 11: Stress intensity factor

Residual stress due localised heating Residual stresses after cooling from cure Thermal stress due to uniform temperature change

Chapter 8 Recent Expansions in the Capabilities of Rose’s Closed-form

Analyses for Bonded Crack-patching

Universal efficiency charts for isotropic patches

Equivalence between octagonal and elliptical patch shapes

Effects of patch tapering on the adhesive stresses

Universal charts for the effects of corrosion

Design of patches to compensate for corrosion damage

Analysis of patches over cracks in stiffened panels

Designing to avoid crack initiation

Universal efficiency charts for orthotropic patches

Effects of residual thermal stresses on bonded repairs

Effects of adhesive non-linearity and disbonds on crack-tip stress-intensity factors

Out-of-plane bending effects with one-sided patches

Remaining challenges involving closed-form analyses

The 2D finite element formulation

9.2.1 Element stiffness matrix

Initial design guidelines

Comparison with experimental results for non rib stiffened panels

Repair of thick sections

9.5.1 Experimental results

Repair of cracked holes in primary structures

Repair of cracked lugs

Repairs to cracked holes under bi-axial loading

Findings relevant to thick section repair

9.12.1

References

Repair of cracks in aircraft wing skin

Governing differential equations for bonded joints/repairs

The effect of variable adhesive thickness and material non-linearity

Comparison of commercial finite element programs for the 3 0

10.3.2 Results for no-fillet case

10.3.3 Results for fillet case

10.3.4 Discussion of results

Gradientless FE method for optimal through-thickness shaping

10.4.1 Optimal adherend taper profile at the end of a bonded joint

Sensitivity FE method for optimal joint through-thickness shaping

10.5.1 Initial geometry, materials and loading arrangement

10.5.2 Optimisation method

10.5.3 Analysis for symmetric crack repair with aluminium patch

10.5.4 Analysis for non-symmetric crack repair with boron/epoxy patch

Optimal through-thickness shaping for F/A-I 8 bulkhead reinforcement

10.6.1 Initial geometry, materials and loading arrangement

10.6.2 Parameters for reinforcement optimisation analyses

10.6.3 Stress results for optimal reinforcement designs

Shape Optimisation for Bonded Repairs

M Heller and R Kaye

Context for finite element based shape optimisation

10.2

General configuration for symmetric stepped patches

Analysis for single step case

Analysis for patch with multiple steps

Estimate for optimal first step length

Minimum bound for peak shear strain due to patch length

Minimum bound for peak shear strain due to stiffness of first

xviii Contents

10.6.4 Discussion

Optimisation for F/A- 18 aileron hinge reinforcement

10.7.1

10.7.2 Shape optimisation before reinforcement

10.7.3 Iterative reinforcement design

Initial geometry, materials and loading arrangement

10.8 In-plane shaping effects

Geometry, loading and modelling considerations

Determination of Kt from FEA output

Uniaxial loading and patches with aspect ratios of 2.1

Uniaxial loading and other patch aspect ratios Stress reduction at the centre of the patch for uniaxially loaded plate

Summary of results and discussion 10.9 Conclusions

Chapter 11 Thermal Stress Analvsis

Finite element thermal stress analysis

11.3.1 Two-dimensional strip joints

11.3.2 Three-dimensional strip joints

Application of analysis to large repairs of aircraft wings

Chapter 12 Fatigue Crack Growth Analysis a

Overload effect and validation using finite element method

Thermal residual stresses and comparison with experimental results

12.4.1 Thermal residual stresses

12.4.2

References

Large-scale yielding solution for a stationary crack Plasticity induced crack closure under large-scale yielding solutions

13.4.2 Influence of stress range

13.4.3 Influence of patch thickness

Model for estimating stress intensity

Use of model to estimate crack growth

Extension of the model for growth of disbond damage

Disbond damage in the patch system

Influence of panel thickness variation

Residual strength of patched panels

Glare Patching Efficiency Studies

R Fredell and C Guijt

Overview and background of fibre metal laminates

Chapter 15 Graphite/epoxy Patching Efficiency Studies

Repair of thin skin components

Repair of thick sections

Graphite/epoxy versus boron/epoxy

Effect of bondline defects

Effect of impact damage

Effect of service temperature

Effect of exposure to hot-wet environments

Repair of battle damage

16.2 Specimen and loading

Repair of Multi-site Damage

R Jones and L Molent

Contents

xx

16.2.1 Boeing lap joints

16.2.2 Airbus lap joints

16.5 Specimen fatigue test results

Unreinforced baseline fuselage lap joint specimens Reinforced baseline fuselage lap joint specimens

16.6

Airbus A330/A340 fatigue test article Boeing 727, 747 and 767 in-flight demonstrators

16.8 Conclusions

Chapter 17 Damage Tolerance Assessment of Bonded Composite Doubler

Repairs for Commercial Aircraft Applications

D Roach

17.1 Introduction

17.1.1

17.1.2

Composite doubler damage tolerance tests

Conformity inspection and FAA oversight

Validation of Stress Intensity Estimations in Patched Panels

B Aktepe and A.A Baker

18.2.1 K-gauge equations

Theory of KI measurement using strain gauges

18.3.1 Westergaard equations

18.3.2

18.3.3 Wang’s crack-bridging model

Rose’s inclusion model for stress intensity 18.4 Experimental procedure

Chapter 19 Bonded Repair of Acoustic Fatigue Cracking

R.J Callinan and S.C Galea

19.2.2 Aft fuselage cracking

Sound pressure levels

19.3.1 Inlet nacelle

19.3.2 Aft fuselage

19.3.3 Power spectral density

Random response analysis

Stress intensity factors

FEA of cracked nacelle inlet

19.6.1 Crack growth study

19.7.4 Results and discussion

Aft fuselage finite element model

19.8.1 Modes and frequencies

19.8.2

19.8.3 Residual thermal stresses

19.8.4 Damping data

19.8.5 Adhesive data

Thermal environment for F/A- 18

Summary of repair failure investigation

Design of highly damped patch

Damping of highly damped patch

Analysis of repaired cracked plate

Acoustic fatigue crack growth data

19 IO Analytical results

20.2 Smart patch approach

20.3 Damage detection studies

S.C Galea

20.3.1 Load transfer (strain) technique

20.3.2 Residual strain technique

20.3.3

20.3.4

Laboratory smart patch conceptional demonstrators

Electro-mechanical impedance, transfer function and stress wave

Chapter 21 Adhesively Bonded Repairs: Meeting the Safety Requirements

Implied within Existing Aviation Industry Certification Regulations

The need to certify a repair

Fatigue and damage tolerance analysis

Current limitations of crack patching

Justifying credit for patching efficiency - fatigue concerns

22.2.1

22.2.2

22.2.3 Validation of patching analysis

Justifying credit for patching efficiency - environmental durability concerns 22.3.1

22.3.2

Justifying credit for patching efficiency ~ the Smart Patch approach

Influence of fatigue on patching efficiency Obtaining patch system fatigue allowables 22.3

Assurance of patch system environmental durability Australian experience on service durability

22.4

22.5 Discussion

22.6 Conclusions

References

Chapter 23 Nondestructive Evaluation and Quality Control for Bonded Composite

Repair of Metallic Aircraft Structures

D.P Roach and C.M Scala

23.3.3 Challenges in crack monitoring

Quality control issues in service

Use of realistic calibration standards

Chapter 24 Practical Application Technology for Adhesive Bonded Repairs

24.1.1 Management of repair technology

Repair application technology

Training and certification

Deficient repair concepts

25.2 Aircraft battle damage repair

Rapid Application Technology: Aircraft Battle Damage Repairs

R Bartholomeusz, P Pearce and R Vodicka

765

766

766

767

xxiv Contents

25.3.4 Simplified design methods for ABDR

25.3.5 Surface treatment

25.3.6 Forming the bonded composite patch

25.3.7 Mechanically fastened, metallic repair

25.3.8 Fatigue and static testing of specimens

25.3.9 Comparison of test results

Development of a bonded composite ABDR system

25.5.2 Pre-bonding surface treatment procedures

25.5.3 Repair consolidation and application

Current approaches to training and certification

26.4.1 The purpose of a trade structure

26.4.2 A four-tiered trade structure - the ARTI model

The ARTI model for training of bonded repair specialists

Certification of bonded repair specialists

26.6.1

26.6.2 Administration of certification tests

26.7 Conclusion

References

Benefits of improved training and process control - an example

Building a database of reliable repairs - “We’re all in this together”

27.2 Crack location and residual strength

27.3 Repair substantiation requirements

27.3.1 Design load cases

27.3.2 Fatigue loading

Design validation (finite element analysis)

27.5.1 Uncracked, unpatched wing model

Cracked, patched model including thermal effects

Repair substantiation (representative specimen testing)

27.7.1 Representative bonded joints

Case History: F-111 Lower Wing Skin Repair Substantiation

K.F Walker and L.R.F Rose

28.2 Fuselage door surround structure tests

28.2.1 Full-scale structural testing philosophy

28.2.2 L-1011 fuselage structure

28.2.3

28.2.4 Biaxial test facility description

Fuselage door surround structure test results

28.3.1

28.3.2

28.3.4

28.3.5 Nondestructive inspection

Component level tests: door corner specimen

28.4.1 Door corner test overview

28.4.2 Subsize door corner test results

28.5.1 Composite doubler repair of L-1011 aircraft passenger door

28.5.2 Non-destructive inspection of door surround structure and

composite doubler

28.5.3 Inspection intervals for L-1011 aircraft

28.5.4 Quality assurance measures

Repair of fuselage test article with a composite doubler

28.3

Structural tests before composite doubler installation

Structural tests after composite doubler installation

Validation of finite element model analytical results

28.4

28.5 L-1011 composite doubler installation

28.6 FAA and industry approvals

Selection and evaluation of materials

Selection and evaluation of the reinforcement

29.4.1 Mechanical test evaluation

29.4.2

29.4.3

29.4.4

29.4.5 Modifications to doubler system

29.4.6 Residual stress minimisation

Doubler application technology

Cure characterisation and formability studies

Selection and evaluation of candidate adhesives

Selection and evaluation of surface treatment procedures

xxvi Contents

Chapter 30 Case History: Bonded Composite Reinforcement of

the F/A-18 Y470.5 Centre Fuselage Bulkhead

R.A Bartholomeusz and A Sear1

FE analysis of bulkhead and reinforcement

30.2 I Results of the bulkhead FE analysis

30.2.2 Measurement of adhesive through-thickness stresses

F E design of representative specimen (curved beam specimen)

Experimental test program

30.4.1 Static testing of curved beam specimen

30.4.2 Durability testing of the curved beam specimen

30.4.3 Residual strength after fatigue

Trial installation of reinforcement to full-scale fatigue test article

Discussion

30.6.1 Pre ECP reinforcement

30.6.2 Post ECP reinforcement

C-5A Fuselage Crown Cracking

C Guijt and S Verhoeven

Design of the bonded repair

FEM model of the patched crack

Case History: F-16 Fuel Vent-hole Repairs

C Guijt and J Mazza

32.3.1 Mechanically fastened aluminum patch

Design of the bonded repair

Fatigue analysis of the aluminum doubler

Chapter 33 Reinforcement of the F/A-18 Inboard Aileron Hinge

Static testing and repair validation

Certification and implementation to aircraft

34.3 Repairs to RAF aircraft

34.3.1 Secondary structure repairs

34.3.2 Primary structure repairs

35.6 Current status of DC-lO/MD-l 1 commercial aircraft repairs

Chapter 36 Case History: CF-116 Upper Wing Skin Fatigue Enhancement Boron

Bonded composite doublers

Doubler design and analysis

Doubler manufacturing and installation procedures

36.6.1 Doubler qualification testing

Doubler fractographic analysis

37.2.1 QANTAS demonstrator program

37.2.2 Ansett keel beam reinforcement

In-service environment and repair location

37.3.1 Temperature

37.3.2 Foreign object damage

37.3.3 Airflow and erosion

37.3.4

37.3.5 Miscellaneous

Bond durability and surface treatment

37.5.1 QANTAS program

37.5.2 Ansett keel beam demonstrator reinforcement

37.6 Discussion and lessons learnt

37.6.1 Erosion protection by the use of shields

37.6.2 Repair location and design

37.6.3 Applicability of demonstrator programs

39.2.2 Large scale wing reinforcement

39.2.3 Large scale fuselage reinforcement

Current state of the technology

39.3.1

39.3.2 Grit blast/silane process steps

Full length rotor blade doublers

Contents

39.4 Process areas requiring adaptation

39.4.1 Solvent scrubbing step

39.4.2 Grit blast step

40.2.3 Strain gage installation

40.2.4 Test spectrum and equipment

Test results

40.3.1 Strain gage results

40.3.2 Crack growth results

Comparison between test results and analytical predictions

Application of composite reinforcement to a full scale wing test

Conclusions

References

Composite reinforcement fabrication and bonding

Chapter 41 Case History: Advanced Composite Repairs of USAF C-141 and

Industrialization and repair

Success and failures

Materials development and characterisation

Installation of composite reinforcement

Reinforcement efficiency assessment

Chapter 19

BONDED REPAIR OF ACOUSTIC FATIGUE

CRACKING

R.J CALLINAN and S.C GALEA

Defence Science and Technology Organisation, Air Vehicles Division, Fishermans Bend, Victoria, 3207, Australia

19.1 Introduction

Acoustic fatigue is due to a very high intensity excitation as a result of pressure waves caused by either engine or aerodynamic effects Acoustically-induced cracking has occurred on the external surface of the lower nacelle and aft fuselage skins on the F/A-18, as illustrated in Figure 19.1 In the former region, overall sound pressure levels greater than 170dB have been measured in flight [l] These high sound pressure levels appear to be a result of an aerodynamic disturbance at the inlet lip [l] Typical cracks occur along a line of rivets or run parallel to the rivet line and may turn into the centre of the panel, as shown in Figure 19.1 Cracking generally occurs along the longer side of the panel where the bending stresses due to out-of-plane vibrations are a maximum Up to a third of the F/A-18’s in the RAAF fleet are affected by these cracks

The standard repair for such cracking is to remove and replace the panel The standard long term fix is to incorporate additional stiffeners on the inside to stiffen the panel This has two effects; firstly to reduce the panels response, i.e lower stress for a given load and secondly it increases the resonant frequencies of the panel to

frequencies well outside the recorded excitation frequencies

In order to reduce the cost of repairing such cracked structures a bonded composite repair would be preferred The benefits of such a repair are reflected in the time required to carry out the repair For example, the inlet nacelle repair typically requires a repair time of 60 h for the mechanical repair and approximately 15-25 h for the bonded repair In the case of the repair to the aft fuselage the time for the mechanical repair is 15-30 h spread over three or four days This repair also requires engine removal and installation which takes a three man team approximately 8 h followed by engine ground runs

532 Advances in the bonded composite repair of metallic aircraft structure

Y557.500 y566.m Y574.500 Y580500 Y598.000

Fig 19.1 Location of the cracking in the lower nacelle inlet and the aft fuselage skins (Chemically

milled fillets are indicated by dotted line.)

A bonded repair was designed for the inlet nacelle area, based on a standard repair design procedure, and implemented on an existing cracked aircraft While in the past boron fibre/epoxy resin patches have been extremely successful, in repairing cracked metallic secondary and primary structures [2], (see also Chapters

13 and 16) in this case significant crack growth occurred after the application of the repair This unsuccessful application of a bonded repair to a cracked metallic structure was due to the fact that the standard design procedure [3] is based on in- plane low cycle (quasi-static) loading condition (See Chapter 7) However for the acoustic environment a new design approach for bonded repairs is required, i.e a design procedure based on dynamic out-of-plane high-cycle loading Work carried out by [4,5] showed that the addition of damping to a bonded repair can result in significant reductions of crack growth These authors showed that a low-damping patch covering the entire panel resulted in a reduction of the mode Z stress intensity,

KI, from 18 MPa ml/*, for a cracked unrepaired panel, to 6 MPa m1'2 Although this

is not a high value in comparison with the fracture toughness value, K I C , in an environment of high-cycle fatigue it leads to a high crack growth rate They found that by further increasing the damping of the repaired structure, from a loss factor

of 0.032 to 0.128, reduced the stress intensity to 4MPam1l2 Work carried out by

[6,7] also indicates that the addition of damping can significantly reduce the crack growth rate

Chapter 19 Bonded repair of acoustic fatigue cracking

This chapter firstly investigates the failure of the low damped boron/epoxy repair Secondly the report details the design of a highly-damped bonded repair, incorporating constrained layered damping (CLD), using finite element analysis (FEA) of the structure subject to simulated acoustic loading conditions CLD material incorporates a layer of visco-elastic material (VEM) with a constraining layer attached to the top of the VEM The study involves the estimation of the root mean square (rms) stress intensity factor ( K ) in the cracked and cracked/repaired

cases Various patch design parameters, such as reinforcing patch dimensions, thicknesses of the (viscoelastic) damping material and thickness of the constraining layer are considered in order to maximise the effectiveness of the patch and thus lower the stress intensity Finally it is theoretically verified that the highly-damped patch can significantly reduce crack growth and therefore yield acceptable crack lengths for a 6000 h lifetime condition

19.2 Cracking history

19.2.1 Inlet nacelle

As a result of the failure of the boron/epoxy repair to prevent crack growth, the particular panel was removed from the aircraft for replacement and assessment, [8],

of the fracture surface The inside surface of a section of the cracked panel is shown

in Figure 19.2, and the missing section was that used for the examination It was found that the crack had started on the inside of the thicker section, and developed

as a semi-elliptical surface crack, as indicated in Figure 19.3, until it became deep enough to penetrate the thickness of the panel The crack proceeded as a through crack, with crack fronts being inclined as shown in Figures 19.4 and 19.5 corresponding to a final crack length of 135mm The central region of the crack just before the boron/epoxy repair was applied is shown in Figure 19.6 and indicates rubbing of the crack faces due to transverse movement This would

Advances in the bonded composite repair of metallic aircraft structure

Fig 19.3 The origin of the crack in the panel (arrowed) The remains of the semi-elliptical surface crack can be seen just below the dotted line This crack was located on the inside surface of the thicker section

of panel Note the smooth rubbed character of the surface, indicating large relative motions between the

two halves of the crack [8]

Fig 19.4 View of the tip of the crack in the thick end of the panel The fatigue crack surface is the bright faceted region to the left and the deliberately broken surface is the mottled grey area to the right Note the shape of the crack, much longer along the inside surface of the panel, indicating that a degree of

bending was involved in the development of the crack Also note that the fatigue crack surface is bright

and highly detailed, and does not appear to have been rubbing [8]

Chapter 19 Bonded repair of acoustic fatigue cracking 535

Fig 19.5 View of the tip of the crack in the thinner part of the panel The same comments as above for

Fig 19.4 apply here [8]

Fig 19.6 View of the central area of the crack, showing the presence of pronounced rubbing [8]

correspond to a mode I11 crack tip driving force However after the patch was

applied a change in crack growth mode occurred This is seen in Figures 19.4 and 19.5 where the crack is bright and highly detailed in comparison to that shown in Figure 19.6 After the repair the shape of the crack front indicates that bending was involved This is consistent with a neutral axis offset caused by the bending of both the boron/epoxy and panel in which the maximum stress occurred in the inner surface of the panel

Both the length of the crack tip after the repair and flight hours are known for the right hand crack tip shown in Figure 19.2, and is 2mm/flight hour While the crack growth rate before the repair is unknown it is thought to be considerably less than that after the repair There is a possibility that before the repair, the rubbing

of the crack faces provided some friction damping to the panel, this in turn led to a lower crack growth rate

536 Advances in the bonded composite repair of metallic aircraft structure

, , , , , , \ ‘ ,

, ,

The chemically milled steps for panels in the aft fuselage are shown dotted in Figure 19.1 Cracks occur in the chemically milled step region, and are generally parrallel to the longest side and midway along the side The aft fuselage section shown in Figure 19.1 contains a composite of all known cracking locations which occur between fuselage stations Y557.500 and Y598.000 The conventional repair involves removal of the crack and use of a metallic doubler attached by mechanical fasteners To carry out this type of repair the engine needs to be removed

19.3 Sound pressure levels

19.3.1 Inlet nacelle

One third octave sound pressure measurements have been made in flight [ 11, using microphones located at the nacelle inlet area, and are plotted in Figure 19.7 The spectrum level, relative to the overall sound pressure level (OASPL), is derived from this data and is also shown in Figure 19.7 In this case the higher than expected

sound pressure levels (SPL) were caused by an aerodynamic disturbance at the inlet

lip At present no data exists for the calculation of SPL’s in the aft fuselage This spectrum is now used as the excitation pressure on the inlet nacelle F.E model

19.3.2 Aft fuselage

Frequency (Hz)

Fig 19.7 Spectrum and one-third octave band levels of sound pressure over the external nacelle inlet,

(where OASPL= 172.2dB), [l]

Chapter 19 Bonded repair of acoustic fatigue cracking 537 Table 19.1

Power spectral density for inlet nacelle

Three points Pressure Frequency for curve spectrum level

31.5 - 32.2 140 4.0 x IO-' 1000.0 - 35.2 137 2.005 x IO-*

disturbance is responsible for the higher than expected SPL The possibility exists that the flap may be responsible for this In the case of the B52, [9], it was found that cracking in the aft fuselage skin was a result of the aerodynamic disturbance caused by the deflected flap position during take-off However aerodynamic disturbances can also be caused by high angle of attack manoeuvres resulting in separation of airflow over the flap

19.3.3 Power spectral density

The relationship between the spectrum SPL and the r.m.s fluctuating pressure

b) is given in [lo] as:

19.4 Random response analysis

The random response analysis capability of the NASTRAN program has been used to solve this problem [Ill This involves a solution in the frequency domain

after the transfer function, H ( w ) , is generated Together with the PSD of the

excitation, Sr(o), the PSD of the response, S J ( W ) , is determined:

(19.3) This analysis allows the statistical properties of the system to be evaluated Random vibrations considered here involve all frequencies at any one instant in time After calculating the PSD, the root mean square (r.m.s.) of the response can