97199844-Composites-Manual-NAVAIR-01-1A-21.pdf

NAVAIR 01-1A-21

TECHNICAL MANUAL

ORGANIZATIONAL AND INTERMEDIATEMAINTENANCE

GENERAL COMPOSITEREPAIR

This publication supersedes NAVAIR 01-1A-21, dated 01 November 2001.

DISTRIBUTION STATEMENT C. Distribution authorized to U.S. Government agencies and their contractors toprotect publications required for official use of for administrative or operational purposes, determined on31 January 1994. Other requests for this document shall be referred to: Commanding Officer, Naval Air TechnicalData and Engineering Service Command, Naval Air Station North Island P.O. Box 357031, Building 90 Distribu-tion, San Diego, CA 92135-7031.

DESTRUCTION NOTICE - For unclassified, limited documents, destroy by any method that will preventdisclosure of contents or reconstruction of the document.

PUBLISHED BY DIRECTION OF COMMANDER, NAVAL AIR SYSTEMS COMMAND

01 SEPTEMBER 20050801LP1046361

NATEC ELECTRONIC MANUAL

NAVAIR 01-1A-21

LIST OF EFFECTIVE PAGESDates of issue for original and changed pages are:

Original ....................... 0 ........................ 01 Sep 2005Change ....................... x ....................... xx XXX 199X

Insert latest changed pages; dispose of superseded pages in accordance with applicable regulations.

NOTE: On a changed page, the portion of the text affected by the latest change is indicated be a vertical line,or other change symbol in the outer margin of the page. Change in illustrations are indicated by miniaturepointing hands. Changes to wiring diagrams are indicated by shaded areas.

Total number of pages in this manual is 320, consisting of the following:

Page *Change Page *Change Page *ChangeNo. No. No. No. No. No.

Change ....................... 0 ......................... 15 Sep 1993Change ....................... x ........................ xx XXX 199X

A Change X

*Zero in this column indicates an original page.

Title ........................................ 0A ............................................. 0i-xi .......................................... 0xii Blank ................................. 01-1 - 1-3 ................................. 01-4 Blank................................ 02-1 - 2-2 ................................. 03-1 - 3-13 ............................... 03-14 Blank ............................. 04-1 - 4-9 ................................. 04-10 Blank ............................. 05-1 - 5-30 ............................... 06-1 - 6-78 ............................... 07-1 - 7-113 ............................. 07-114 Blank ........................... 08-1 - 8-32 ............................... 09-1 - 9-3 ................................. 09-4 Blank................................ 010-1 - 10-9 ............................. 010-10 Blank ........................... 0Glossary-1 - Glossary-7 ........ 0Glossary-8 Blank ................... 0

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LIST OF ILLUSTRATIONS ........................................ iii

LIST OF TABLES ...................................................... vii

WARNINGS APPLICABLE TO HAZARDOUSMATERIALS .................................................. ix

I INTRODUCTION

1-1. Purpose ........................................... 1-11-2. Contents and Limitations................ 1-11-3. Requisitioning and Automatic

Distribution ................................... 1-11-4. Abbreviations and Symbols ........... 1-11-5. Reference Material ......................... 1-1

II INTRODUCTION TO COMPOSITES

2-1. Advanced Composite Materials ..... 2-12-2. Prepreg ........................................... 2-22-3. Laminate ......................................... 2-2

III TYPICAL DAMAGE

3-1. General ........................................... 3-13-2. Damage to Advanced Composite

Materials ....................................... 3-13-3. Damage Assessment ..................... 3-6

IV REPAIR PROCEDURE SELECTION CRITERIA

4-1. Design Criteria ................................ 4-14-2. Additional Criteria ........................... 4-14-3. Basic Repair Joints ......................... 4-24-4. Damage Disposition ....................... 4-74-5. Repair Versus Replace .................. 4-9

V REPAIR MATERIALS

5-1. Incorporated Materials ................... 5-15-2. Incorporated Repair Material

Selection Criteria ........................ 5-145-3. Unincorporated Materials (Ancillary) 5-155-4. Repair Material Shipping,

Receiving and Storage .............. 5-185-5. Repair Material Preparation ......... 5-205-6. Material Evaluation Testing .......... 5-275-7. Disposal of Materials

Used for Repair .......................... 5-31

VI REPAIR PROCESSES

6-1. Description ...................................... 6-16-2. Cleaning .......................................... 6-1

a. General .................................... 6-1b. Procedure ................................ 6-1

6-3. Damage Removal ........................... 6-1a. Outlining the Damage ............. 6-1b. Penetration Damage ............... 6-2c. Partial Thickness Damage...... 6-2d. Core Damage .......................... 6-2

6-4. Machining, Drilling, Reamingand CountersinkingAdvanced Composites ................. 6-4a. Background ............................. 6-4b. General .................................... 6-5c. General Air Tool Safety .......... 6-5d. Machining Boron/Epoxy

Composites ............................. 6-6e. Drilling Boron/Epoxy

Composites ............................. 6-6f. Machining Carbon/Epoxy,

Carbon/Bismaleimide andCarbon/PolyimideComposites ............................. 6-6

g. Drilling, Reaming andCountersinking Carbon/Epoxy,Carbon/Bismaleimide andCarbon/PolyimideComposites ............................. 6-6

h. Machining Kevlar/EpoxyComposites ............................. 6-7

i. Drilling and CountersinkingKevlar/Epoxy Composites ...... 6-8

j. Align-A-Drill Setup .................. 6-96-5. Paint Removal .............................. 6-11

a. General .................................. 6-11b. Procedure .............................. 6-11

6-6. Joint Machining ............................. 6-11a. General .................................. 6-11b. Blunt Cut Joint Machining ..... 6-11c. Scarf Joint Machining ........... 6-12d. Step Joint Machining ............ 6-13

6-7. Bonded Repair Processes .......... 6-14a. Drying .................................... 6-14b. Core Replacement Methods . 6-17c. Core Machining ..................... 6-19d. Ply Orientation and Layup of

Advanced Composite RepairPlies and Patches ................. 6-20

TABLE OF CONTENTS

Section Page Section Page

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e. Standard Wet Layup Process ..6-22f. Double Vacuum Debulk

(DVD) Wet Layup Process ... 6-27g. Precured Patch Material

Cure Processes..................... 6-36h. Surface Preparation for

Bonding ................................. 6-37i. Patch Installation .................. 6-39j. Methods For Applying

Pressure and Heat toCure Bonded Repairs ........... 6-42

k. Adhesive Cure Processes .... 6-54l. Repair Verification ................ 6-59m. Heat Survey .......................... 6-61

6-8. Injection Repair Processes .......... 6-61a. General .................................. 6-62b. Adhesive Characteristics ...... 6-62c. Damage Classification .......... 6-62d. Positive Pressure Injection

Repair .................................... 6-62e. Vacuum Injection Repair ...... 6-63

6-9. Bolted Repair Processes .............. 6-63a. Patch Preparation ................. 6-63b. Blind Side Drilling .................. 6-64c. Drilling/Reaming Patch and

Skin ........................................ 6-68d. Patch and Fastener

Installation ............................. 6-69e. Sealing Repairs ..................... 6-76

VII REPAIR PROCEDURES

7-1. Description ...................................... 7-17-2. Partial Thickness Skin Repair ........ 7-1

a. Procedure 1. Surface Repair .. 7-1b. Procedure 2. Partial Thickness

Damage: Bonded Repair ........ 7-2c. Procedure 3. Partial Thickness

Damage: Bolted Repair .......... 7-67-3. Delamination Repair ..................... 7-10

a. Procedure 4. DelaminationOpen to An Edge .................. 7-10

b. Procedure 5. Delamination NotOpen to An Edge (Blister) ..... 7-13

7-4. Disbond Repair ............................. 7-16a. Procedure 6. Skin to

Core Disbond ........................ 7-16b. Procedure 7. Skin to Closure

Member Disbond andDelaminations ....................... 7-19

7-5. Edge Damage Repair ................... 7-21a. Procedure 8. Edge Damage

Repair .................................... 7-21b. Procedure 9. Edge Damage

Rebuild .................................. 7-22c. Procedure 10. Flush Corner

Repair .................................... 7-31d. Procdure 11. Flush Trailing

Edge Repair .......................... 7-387-6. Fastener Hole Repair ................... 7-41

a. Procedure 12. CountersinkRepair .................................... 7-41

b. Procedure 13. Fill and DrillRepair .................................... 7-42

c. Procedure 14. Fastener HoleDelamination Repair ............. 7-44

d. Procedure 15. Fastener HoleRepair: Swagged Grommet .. 7-47

e. Procedure 16. Fastener HoleRepair: Captive Bushing ....... 7-50

7-7. Penetration Damage Repair ........ 7-53a. Procedure 17. Penetration

Damage Bonded Repair ....... 7-53b. Procedure 18. Backside

Sealing for Installation ofExternally Bonded Patches .. 7-56

c. Procedure 19. PenetrationDamage Bolted Repair,External Patch ....................... 7-60

d. Procedure 20. PenetrationDamage Bolted Repair,External/Internal Patch ......... 7-65

e. Procedure 21. PenetrationDamage Bolted Repair,Internal Patch ........................ 7-73

7-8. Substructure Repairs .................... 7-78a. Procedure 22. Honeycomb

Core Repair: Core Fill Method ... 7-78b. Procedure 23. Honeycomb

Core Repair: Paste AdhesiveMethod................................... 7-81

c. Procedure 24. HoneycombCore Repair: Film/FoamMethod................................... 7-85

d. Procedure 25. Closure RibBonded Repair ...................... 7-91

e. Procedure 26. SubstructureBolted Repair ........................ 7-96

f. Procdure 27. Skin and PartialRib Repair ........................... 7-105

TABLE OF CONTENTS (Cont.)Section Page Section Page

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VIII REPAIR EQUIPMENT/TOOLS8-1. General ........................................... 8-18-2. Composite Repair Tool Sets .......... 8-18-3. Equipment/Tools to Perform

Cutting/Machining,Drilling/Countersinking andReaming Operations ofAdvanced Composite Materials ... 8-1

8-4. Fastener Installation andRemoval Tools ........................... 8-13

8-5. Equipment and Tools to PerformSpecialized Operations .............. 8-23

8-6. Temperature/Vacuum ControlRepair Sets................................. 8-28

IX FACILITY REQUIREMENTS9-1. Background ..................................... 9-19-2. Requirements ................................. 9-1

TABLE OF CONTENTS (Cont.)Section Page Section Page

9-3. General Ventilation ......................... 9-29-4. Equipment/Utility Requirements .... 9-2

X HEALTH AND SAFETY

10-1. Background ................................... 10-110-2. Exposure Routes .......................... 10-110-3. Exposure Limits ............................ 10-210-4. Toxicity and Hazards of

Advanced Composite MaterialsUsed for Repair .......................... 10-3

10-5. Personal Protective Equipment,Equipment/Facilities andPersonal Hygiene ....................... 10-7

10-6. Emergency and First Aid Procedures 10-9

GLOSSARY ................................................ Glossary-1

Figure Title Page Figure Title Page

LIST OF ILLUSTRATIONS

2-1. Laminate Cross Section Cut 90 Degreesto the 0 Degree Fiber Direction ....................2-2

3-1. Typical Impact Damage on 0.1 Inch Thick Carbon/Epoxy Laminate,3 Ft-Lbs Impact Energy................................3-1

3-2. Dented Honeycomb Panel ...............................3-23-3. Penetration Damage .......................................3-23-4. Penetration Damage on 0.75 Inch Thick

Carbon/Epoxy Laminate ..............................3-33-5. Airstream Stripping Damage............................3-33-6. Edge Damage .................................................3-43-7. Fire Damaged Advanced Composites .............3-63-8. Pulse-Echo Ultrasonic Inspection with

A-Scan Presentation ....................................3-83-9. Through-Transmission Ultrasonic

Inspection, A-Scan Presentation ................3-113-10. Defect Mapping of Damage Extent

from NDI ....................................................3-13

4-1. Basic Repair Joints (General) ..........................4-24-2. Basic Repair Joints (Bonded) ..........................4-34-3. External Bonded Patch Joint

Eccentricity Effects ......................................4-4

4-4. External Bonded Patch Shear StressConcentration Effects ..................................4-5

4-5. Scarf Bonded Patch Joint Eccentricity andShear Stress Concentration Effects .............4-6

4-6. Basic Repair Joints (Bolted) ............................4-74-7. Externally Bolted Patch Joint

Eccentricity Effects ......................................4-74-8. External/Internal Bolted Joint

Eccentricity Elimination ................................4-84-9. Effects of Close Tolerances on

Displacement Required to LoadFasteners ....................................................4-8

4-10. Increased Load Sharing of FastenersCaused by Tapering Patch ..........................4-8

5-1. Dry Woven Carbon Cloth Weave Patterns .......5-95-2. Carbon/Epoxy Honeycomb Sandwich

Assembly ...................................................5-105-3. Honeycomb Core Designation .......................5-115-4. Close Tolerance Structural Screw

Installation .................................................5-115-5. Blind Fastener Installation .............................5-125-6. Examples for Preparing Two Part

Adhesives and Filler Compounds...............5-22

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LIST OF ILLUSTRATIONS (Cont.)Figure Title Page Figure Title Page

6-35. Heat Blanket Layup - Partial Vacuum BagCross Section ............................................6-48

6-36. Heat Blanket Layup - Partial Vacuum Bag .....6-506-37. Heat/Vacuum Blanket Installation ..................6-526-38. Typical Envelope Bag Installation ..................6-546-39. Typical Two Part Adhesive

Heat Cure Cycle ........................................6-556-40. The Effect of Undercuring on

Adhesive Strength .....................................6-566-41. Typical Film Adhesive Cure Cycle .................6-586-42. Inspection of Adhesive Squeeze Out

Following Cure...........................................6-616-43. Types of Delaminations .................................6-626-44. Injection Repair .............................................6-626-45. Impact Damage Injection Repair ....................6-626-46. EA956 Isothermal Rheological Response ......6-636-47. Hole Finder Method .......................................6-656-48. Blind Hole Transfer Punch Method ................6-666-49. Measuring and Scaling Method .....................6-666-50. Hydrocal Drill Blanket Method........................6-686-51. Hi-Lok Installation: Pneumatic Tooling ...........6-716-52. Blind Fastener Inspection ..............................6-736-53. Fastener Removal Methods ..........................6-736-54. Depth Gauge Adjustment ..............................6-746-55. Vacuum Pad Indexing ...................................6-756-56. Blind Fastener Drilling ...................................6-756-57. Blind Bolt Knockout .......................................6-766-58. Removal of Tightly Clamped

Blind Fasteners ..........................................6-776-59. Removal of Loose Blind Fasteners ................6-776-60. Sealing of Bolted Repairs ..............................6-78

7-1. Surface Damage .............................................7-17-2. Process Flow Diagram for Surface Repair,

Procedure 1 .................................................7-17-3. Process Flow Diagram for Partial Thickness

Bonded Repair, Procedure 2 ........................7-37-4. Partial Thickness Bonded Repair .....................7-47-5. Process Flow Diagram for Partial Thickness

Bolted Repair, Procedure 3 ..........................7-77-6. Partial Thickness Bolted Repair,

Generic Patch Layout ..................................7-87-7. Fabrication/Installation of Countersink Filler .....7-97-8. Composite Blind Fastener

Inspection and Acceptability Limits .............7-117-9. Process Flow Diagram for Delamination

Open to An Edge Repair, Procedure 4 .......7-127-10. Delamination Repair Open to An Edge ..........7-137-11. Process Flow Diagram for Delamination

Not Open to an Edge Repair, Procedure 5 ....7-14

5-7. Use of Triple Beam Balance withTwo Part Adhesives ...................................5-23

5-8. Film/Foaming Adhesive Out-Time LogExample ....................................................5-24

5-9. Film/Foaming Adhesive Out-Time Log ...........5-255-10. Vertical Flow Test Fixture Assembly ..............5-265-11. Temperature and Humidity Operating

Environment for Adhesive Preparation .......5-27

6-1. Damage Outlining ...........................................6-26-2. Penetration Damage Removal -

Template Method.........................................6-36-3. Partial Thickness Damage Removal ................6-36-4. In-Plane Versus Out-Of-Plane

Cutting Forces .............................................6-46-5. Align-A-Drill Setup .........................................6-106-6. Basic Repair Joints ........................................6-116-7. Scarf Joint Outline Layout ..............................6-126-8. Scarf Joint Machining ....................................6-126-9. Scarf Joint Inspection Requirements .............6-136-10. Step Joint Outline Layout ..............................6-146-11. Step Joint Inspection Requirements ..............6-146-12. Core Replacement Methods ..........................6-176-13. Replacement Core Fit ...................................6-186-14. Weight Versus Hole Diameter for Two

Core Replacement Methods ......................6-196-15. Lamina and Lamina Fiber Direction ...............6-206-16. Fiber Orientations ..........................................6-216-17. Typical Laminate ...........................................6-216-18. Stacking Sequence Effects ............................6-216-19. Cutting Template Alignment ..........................6-236-20. Ply Layup Log ...............................................6-256-21. Three Ply (45,0,45)w Repair Patch .................6-276-22. DVD Wooden Box Tool .................................6-286-23. V-22 DVD Tool ..............................................6-296-24. Minimum Vacuum Level for DVD Process .....6-326-25. Cure Stacking Sequences .............................6-356-26. The Effect of Cure Pressure on

Interlaminar Shear Strength (ILSS) andVoid Content ..............................................6-37

6-27. Typical Carbon/Epoxy Laminate Cure Cycle ..6-386-28. Patch Edge Taper Dimensions ......................6-396-29. Layup of Stacked Patches and Adhesive .......6-416-30. Methods for Applying Positive Pressure .........6-436-31. C-Clamp Sequence and Placement ...............6-446-32. Temperature Variations Underneath a

Typical 12 Inch x 12 Inch Heat Blanket ......6-456-33. Heat Blanket Selection/Thermocouple

Placement (Typical) ...................................6-456-34. Heat Lamp Temperature Effects ....................6-46

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7-12. Verification of Leak Path:Delamination Repair .................................7-14

7-13. Delamination Repair Not Open to an Edge(Blister) ......................................................7-15

7-14. Process Flow Diagram for Disbond Repair,Procedure 6 ...............................................7-17

7-15. Verification of Leak Path: Disbond Repair ......7-177-16. Disbond Repair .............................................7-187-17. Process Flow Diagram for Skin to

Closure Member Disbond andDelamination Repair, Procedure 7 .............7-20

7-18. Skin to Closure Member Disbond Repair .......7-207-19. Edge Damage and Repair .............................7-217-20. Process Flow Diagram for

Edge Damage Repair, Procedure 8 ...........7-227-21. Typical Edge Damage ...................................7-237-22. Process Flow Diagram for

Edge Damage Rebuild, Procedure 9 ..........7-247-23. Edge Damage Rebuild ..................................7-267-24. Machine Scarf ...............................................7-287-25. Aluminum Support Plate ................................7-287-26. OML Patch Adhesive.....................................7-297-27. Machining Repair Core ..................................7-297-28. Machining Ramp in Repair Core ....................7-297-29. Cutting Impregnated Carbon Cloth ................7-297-30. Impregnated Carbon Cloth Repair Plies.........7-307-31. Ply Layup ......................................................7-307-32. Vacuum Bag Layup .......................................7-317-33. Process Flow Diagram for Flush Corner

Repair, Procedure 10 .................................7-327-34. Flush Corner Repair Sequence .....................7-337-35. Repair Rib Layup and Tool ............................7-357-36. Marking Splice Plate for Alignment ................7-357-37. Adhesive Applied to Mating Surface

of Patch .....................................................7-377-38. Repair Patch Taped to Repair Rib and

Splice Plate ...............................................7-377-39. Flush Trailing Edge Repair Sequence ...........7-397-40. Process Flow Diagram for

Countersink Repair, Procedure 12 .............7-417-41. Process Flow Diagram for

Fill and Drill Fastener Hole Repair,Procedure 13 .............................................7-43

7-42. Template Fabrication .....................................7-437-43. Fastener Hole Sealing ...................................7-437-44. Process Flow Diagram for

Fastener Hole Delamination Repair(Vacuum Injection), Procedure 14 ..............7-45

7-45. Vacuum Cup Installation ................................7-46

7-46. Application of Clamp-Up Pressure .................7-477-47. Process Flow Diagram for

Fastener Hole Repair: Swagged Grommet,Procedure 15 .............................................7-48

7-48. Swagged Grommet Installation ......................7-497-49. Captive Bushing Repair Components ............7-507-50. Process Flow Diagram for

Captive Bushing Repair, Procedure 16 ......7-517-51. Captive Bushing Repair Flange and

Countersink Bushing Installation ................7-527-52. Installation View of Captive Bushing Repair ...7-537-53. Process Flow Diagram for Penetration

Damage Bonded Repair, Procedure 17 .....7-547-54. Process Flow Diagram for

Backside Sealing for Installation ofExternally Bonded Patches,Procedure 18 .............................................7-57

7-55. Safety Wire and Slotted Backside Patch ........7-587-56. Installation of Backside Patch ........................7-587-57. Application of Adhesive .................................7-597-58. Backside Patch Pulled Into Position ...............7-597-59. Application of Pressure ..................................7-607-60. Process Flow Diagram for Penetration

Damage Bolted Repair, External Patch, Proce-dure 19 ......................................................7-61

7-61. Repair Arrangement, Bolted Repair,External Patch ...........................................7-62

7-62. Sump Removal and Installation .....................7-627-63. Process Flow Diagram for

Penetration Damage Bolted RepairExternal/Internal Patch, Procedure 20 ........7-66

7-64. Damage Definition and Cleanup ....................7-667-65. Repair Kit Components .................................7-677-66. Patch and Center Plug Aligned on Part ..........7-677-67. Align-A-Drill Setup .........................................7-687-68. Locating Drill Guide .......................................7-687-69. Transferring Holes From Patch to Skin ..........7-687-70. Insertion of Backup Plates Into Cavity ............7-697-71. Backup Plates Held in Place With

Temporary Fasteners ................................7-707-72. Measuring Gap Between

Center Plug and OML ................................7-707-73. Reaming Operation .......................................7-717-74. Installation of Backing Plates .........................7-717-75. Threaded Assembly Pin ................................7-717-76. Backing Plates Pulled Into Position ................7-717-77. Center Plug and Backing Plates

Correctly Installed ......................................7-727-78. Finished Repair .............................................7-72

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7-79. Process Flow Diagram for PenetrationDamage Bolted Repair, Internal Patch,Procedure 21 .............................................7-74

7-80. Repair Arrangement for PenetrationDamage Bolted Repair, Internal Patch .......7-75

7-81. Fastener Pattern Layout ................................7-757-82. Internal Patch and Splice Plate Assembly ......7-767-83. Internal Patch Aligned and Secured ...............7-767-84. Fay Surface Sealing ......................................7-777-85. Process Flow Diagram for

Honeycomb Core Repair,Core Fill Method, Procedure 22 .................7-79

7-86. Procedure for Core ReplacementUsing the Core Fill Method.........................7-80

7-87. Estimating Filler Material forCore Fill Method ........................................7-81

7-88. Process Flow Diagram forHoneycomb Core Repair,Paste Adhesive Method, Procedure 23 ......7-82

7-89. Procedure for Core ReplacementUsing the Paste Adhesive Method .............7-84

7-90. Process Flow Diagram forHoneycomb Core Repair,Film/Foam Method, Procedure 24 ..............7-86

7-91. Procedure for Core ReplacementUsing the Film/Foam Method .....................7-88

7-92. Process Flow Diagram for Closure RibBonded Repair, Procedure 25 ....................7-91

7-93. Repair Details .............................................7-927-94. Damaged Closure Rib .................................7-927-95. Damage Removed ......................................7-937-96. Replacement Closure Rib Fabrication .........7-937-97. Positioning Replacement Core ....................7-937-98. Positioning Replacement Rib ......................7-947-99. Applying Pressure to Bondline

With C-Clamps ........................................7-947-100. Machine Replacement Core Flush

With OML ................................................7-957-101. Application of Adhesive to Repair Area ........7-957-102. Repair Patch Applied to Repair Area ...........7-967-103. NDI Performed on Final Patch Bond............7-967-104. Repair Complete .........................................7-967-105. Process Flow Diagram for Substructure

Bolted Repair, Procedure 26 ....................7-977-106. Repair Arrangement for C-Channel

Substructure Bolted Repair ......................7-987-107. Process Flow Diagram for Skin and

Partial Rib Repair, Procedure 27 .............. 7-106

7-108. Skin and Partial Rib Repair Sequence .......7-1077-109. Skin and Partial Rib Repair,

Repair Angle Bond ..................................7-1097-110. Repair Angle Layup and Tool ....................7-1107-111. Repair Angle Clamp Arrangement .............7-1117-112. Pressure Application ................................. 7-112

8-1. Router Motors and Accessories .......................8-28-2. Cutters/Sanding Equipment .............................8-68-3. Honeycomb Core Cutters ................................8-78-4. Drilling/Reaming/Countersinking/

Counterboring Tools ....................................8-88-5. Drilling Equipment ...........................................8-98-6. Temporary Fasteners ....................................8-138-7. Blind Fastener Grip Length Gauges...............8-158-8. Blind Fastener Installation:

Close Quarter Pneumatic Tooling ..............8-188-9. Blind Fastener Installation:

Pneumatic Tooling .....................................8-188-10. Blind Fastener Installation:

Close Quarter Hand Tooling ......................8-198-11. Blind Fastener Installation: Hand Tooling .......8-198-12. Pistol Extension Assemblies ..........................8-208-13. Installation Tool Conversion ..........................8-208-14. Fastener Removal Kit: Vacuum System ........8-228-15. Hi-Lok Installation: Hand Tooling ...................8-228-16. Hi-Lok Fastener Tools ...................................8-238-17. HEPA Filter Vacuum Cleaner ........................8-248-18. Industrial Hypodermic Syringe and Needles ..8-248-19. SEMCO Model 250 Sealant Gun ...................8-258-20. Vacuum Cup .................................................8-258-21. Moisture Indicator ..........................................8-298-22. SK340-00192 Adhesive Comb ......................8-298-23. F-14 Composite Structure Repair

Console and Blanket Assembly .................8-308-24. 4230-211 Two Zone Heat Blanket

Temperature Sensor Placement ................8-308-25. F-18 Temperature/Vacuum Control

Repair Set, P/N 74D110165-1001..............8-318-26. Generic Temperature/Vacuum Control

Repair Set, P/N 1935AS100-1 ...................8-32

10-1. Carbon Fiber and Human Hair DiametersCompared to Filtration Level of HEPA Filter10-2

10-2. Removing Disposable Gloves ........................10-8

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LIST OF TABLESTable Title Page Table Title Page

1-1. Abbreviations and Symbols .............................1-21-2. Reference Material ..........................................1-3

5-1. Incorporated Repair Materials:Adhesives/Sealants/Fillers ...........................5-2

5-2. Incorporated Repair Materials:Patch Materials ............................................5-5

5-3. Incorporated Repair Materials:Honeycomb Core Materials .......................5-11

5-4. Incorporated Repair Materials:Mechanical Fasteners ................................5-13

5-5. Unincorporated Repair Materials ...................5-165-6. Vacuum Bag Repair Materials Kit,

P/N 135040-1 ............................................5-175-7. Two Part Adhesive Shelf-Life ........................5-185-8. Two Part Adhesives: Pot Life and

Maximum Amount of Material ....................5-205-9. Vertical Flow Test Limits ................................5-27

6-1. Two Part Adhesive Cure Cycles ....................6-266-2. DVD Material and Process Differences

(8 Harness Vs. Plain Weave Fabric) ..........6-316-3. Film Adhesive Cure Cycles............................6-576-4. Common Bonded Repair Errors ....................6-606-5. Blind Fastener Inspection Requirements .......6-72

6-6. Tool Selection to Remove Blind Bolts UsingRK3042B Fastener Removal Kit ................6-74

6-7. Pilot and Shank Drill Sizes forBlind Fastener Removal .............................6-77

7-1. Clamp-Up Bolts and Torque Values ..............7-45

8-1. Router Motors and Accessories .......................8-28-2. Cutters/Sanding Equipment .............................8-58-3. Honeycomb Core Cutters ................................8-78-4. Drilling/Reaming/Countersinking/

Counterboring Tools ....................................8-108-5. Drilling Equipment .........................................8-128-6. Temporary Fasteners ....................................8-148-7. Composi-Lok Installation Tooling ...................8-168-8. Composi-Lok II and IIa Installation Tooling ....8-168-9. Visu-Lok Installation Tooling ..........................8-178-10. Visu-Lok II Installation Tooling .......................8-178-11. Blind Fastener Removal Kit, RK3042B ..........8-218-12. Miscellaneous Equipment ..............................8-26

10-1. Permissible Exposure Limits forComposite Materials ..................................10-3

10-2. Personal Protective Equipment, Equipment/Facilities and Personal Hygiene for Workingwith Advanced Composite Materials ..........10-4

NAVAIR 01-1A-21

viii

Report Control Number (RCN) Location

alskjalkj0000/00000 Pg x-xx

0000/00000 Pg x-xx








Report Control Number (RCN) Location

HMH-466 QA/TPL53998-2005-0028 Pg iii








LIST OF TECHNICAL PUBLICATIONS DEFICIENCY REPORTS INCORPORATED

ix

NAVAIR 01-1A-21

WARNINGS APPLICABLE TO HAZARDOUS MATERIALS

Cryogenic

The symbol of a hand in a block of iceshows that the material is extremelycold and can injure human skin or tis-sue.

Explosion

This rapidly expanding symbol showsthat the material may explode if sub-jected to high temperature, sources ofignition or high pressure.

Eye Protection

The symbol of a person wearing gogglesshows that the material will injure theeyes.

Fire

The symbol of a fire shows that thematerial may ignite or overheat and causeburns.

Poison

The symbol of a skull and crossbonesshows that the material is poisonous oris a danger to life.

Vapor

The symbol of a human figure in a cloudshows that material vapors present adanger to life or health.

1. Warnings and cautions for hazardousmaterials listed are designed to apprise personnel ofhazards associated with such items when they comein contact with them by actual use. Additionalinformation related to hazardous materials is providedin Section X of this manual, Navy Hazardous MaterialControl Program NAVSUPPINST 5100.27, NavyOccupational Safety and Health (NAVOSH) ProgramManuals OPNAVINST 5100.23 (Ashore) andOPNAVINST 5100.19 (Afloat) and the DOD 6050.5Hazardous Materials Information System (HMIS)series publications. For each hazardous materialused within the Navy, a Material Safety Data Sheet(MSDS) must be provided and available for reviewby users. Consult your local safety and health staffconcerning any questions regarding hazardousmaterials, MSDS, personal protective equipmentrequirements, appropriate handling and emergencyprocedures and disposal guidance.

2. Under the heading HAZARDOUS MATERIALSWARNINGS, complete warnings, including relatedicon(s) and a numeric identifier, are provided forhazardous materials used in this manual. The numericidentifiers have been assigned to the hazardousmater ia l in alphabet ical order by mater ia lnomenclature. Each hazardous material is assignedonly one numerical identifier. Repeat use of a specifichazardous material references the numeric identifierassigned at its initial appearance. The approvedicons and their application are shown below.

3. In the text of the manual, the caption WARNINGis not used for hazardous material warnings. Hazardsare cited with appropriate icon(s), the nomenclatureof the hazardous material and the numeric identifierthat relates to the complete warning. Users ofhazardous materials shall refer to the completewarnings, as necessary.

4. EXPLANATION OF HAZARDOUS MATERI-ALS ICONS.

Chemical

The symbol of a liquid dripping onto ahand shows that the material will causeburns or irritation to human skin or tissue.

xNAVAIR 01-1A-21

HAZARDOUS MATERIALS WARNINGSINDEX MATERIAL WARNING

1 Two Part Adhesive Adhesives are toxic. DO NOT breathe vapors.Avoid contact with eyes, skin and clothing. Mix

and use only in well ventilated areas. Wear faceshield, gloves and apron to prevent eye and skincontact. If eye contact occurs, flush immediatelywith large amounts of water. If skin contactoccurs, wash immediately with soap and water.

2 Two Part Residual Adhesive To prevent excessive exotherm, mix no morethan the "Maximum Amount to Mix" grams

specified in Table 5-8 in any one container.If more resin is required, mix in separate mixingcups in the ""Maximum Amount to Mix" gramsspecified in Table 5-8. Do not mix resins whenambient temperatures exceed 90F.

3 Two Part Liquid Adhesive Mixed liquid adhesive may generate largeamounts of heat. Liquid adhesive mixed in

excess of 30 grams may melt the disposableinjection cartridges. For ambient temperatures above80F, do not use injection cartridges to inject liquidadhesive if more than 20 minutes have elapsedafter mixing. For ambient temperatures below 80F,do not use injection cartridges if more than 40 minuteshave elapsed after mixing. Pressurized cartridgesmay spray hot adhesive after the safe operatingtimes have elapsed, potentially injuring artisansand bystanders.

4 Solvent Solvents are toxic and flammable. DO NOTbreathe vapors. Avoid contact with eyes, skin

and clothing. DO NOT use near open flame,sparks or heat. Use only in well ventilated areas.Wear goggles and gloves to prevent eye and skincontact. If eye contact occurs, flush immediatelywith large amounts of water. If skin contactoccurs, wash with soap and water.

5 Sealing Compound Sealing compounds are toxic and flammable.DO NOT breathe vapors. Avoid contact with

eyes, skin and clothing. DO NOT use near openflame, sparks or heat. Use only in well ventilatedareas. Wear goggles and gloves to prevent eyeand skin contact. If eye contact occurs, flushimmediately with large amounts of water. If skincontact occurs, wash with soap and water.

xi

NAVAIR 01-1A-21

HAZARDOUS MATERIALS WARNINGS (Cont.)INDEX MATERIAL WARNING

6 Rubber Primer Rubber primer is toxic and flammable. DO NOTbreathe vapors. Avoid contact with eyes, skin

and clothing. DO NOT use near open flame,sparks or heat. Use only in well ventilated areas.Wear goggles and gloves to prevent eye andskin contact. If eye contact occurs, flushimmediately with large amounts of water. If skincontact occurs, wash with soap and water.

7 Composite Materials Sanding, cutting or drilling composite materialsproduces a fine dust that may cause eye, skin

and lung irritation. Breathing this dust may beinjurious to health. When sanding, cutting ordrilling composite materials, the following pro-tective equipment shall be worn: a respiratorcontaining a HEPA filter, gloves, goggles and longsleeve coveralls. Tape coverall sleeves closed atthe wrist. Use a vacuum cleaner equipped with aHEPA filter to control dust during and aftersanding, cutting or drilling.

8 Dry Ice Dry ice (solid C02) is extremely cold (-110F).DO NOT handle with bare hands. Use gloves with

adequate insulation when handling. Dry ice passesdirectly from the solid state to the gaseous statewhen exposed to ambient temperatures. Whenmaterial is stored in confined spaces, gaseous C02can displace oxygen. Personnel entering oxygendeficient areas may become unconscious.

9 Film/Foaming Adhesive Adhesives are toxic. Avoid prolonged or repeatedcontact with skin. Wear gloves and long sleeve

coveralls to prevent skin contact. If contact occurs,immediately wash with soap and water.

xi/(xii Blank)

xii

NAVAIR 01-1A-21

THIS PAGE LEFT INTENTIONALLY BLANK

1-1

NAVAIR 01-1A-21

SECTION IINTRODUCTION

1-1. PURPOSE.

a. The purpose of this technical manual is to providerepair methods for structures manufactured from ad-vanced composite materials (ACM). In addition, the repairprocess rationale is described (where applicable) to pro-vide the repair technician with an understanding of theprocess sensitive nature of advanced composite repair.This manual also lists the approved equipment andmaterials required for performing the repairs. Theserepair methods are for use at organizational and inter-mediate levels of maintenance.

b. The repairs described in this manual are perma-nent and will restore the part being repaired to itsrequired strength, stiffness and service life.

1-2. CONTENTS AND LIMITATIONS.

a. This manual provides a description of the equip-ment, materials and processes used to repair navalaircraft parts manufactured from ACM. Repair methodsinclude both bonded and bolted techniques. The meth-ods are applicable to monolithic laminates, bondedhoneycomb sandwich assemblies and thin, stiffenedskin assemblies. (The majority of this manual deals withthe repair of carbon/epoxy composites as they com-prise over 90% of the advanced composite assembliescurrently in use on naval aircraft).

b. This manual is a supplement, not a replacementfor a part specific structural repair manual (SRM). Theindividual part specific SRM must be consulted as thelimitations, procedures and materials listed in it takeprecedence over this manual. Information such asoperating environment, damage size limits, weight andbalance limits and repair moldline protrusion limits areestablished by the aircraft manufacturer based uponthe criticality of specific parts. Violation of SRM limitsmay result in excessive part deflection, dynamic instabilityor structural failure. Deviation or substitutions from partspecific SRM materials and processes can only be autho-rized by the Fleet Support Team (FST) for the specificpart in question.

1-3. REQUISITIONING AND AUTOMATIC DISTRI-BUTION OF NAVAIR TECHNICAL MANUALS. Proce-dures to be used by Naval activities and other Depart-ment of Defense activities requiring NAVAIR technicalmanuals are defined in NAVAIR 00-25-100.

1-4. ABBREVIATIONS AND SYMBOLS. Table 1-1lists abbreviations and symbols that do not appear inMIL-STD-12.

1-5. REFERENCE MATERIAL. All references appli-cable to this manual are listed in Table 1-2.

1-2

NAVAIR 01-1A-21

Abbreviations/Symbol Definition

ACM Advanced Composite Materials

ACS American Chemical Society

AIMD Aircraft IntermediateMaintenance Department

BCM beyond the capability of maintenance

BMI bismaleimide

CRT cathode ray tube

CTE coefficient of thermal expansion

CRES corrosion resistant steel

DED Damage Engineering Disposition

DVD double vacuum debulk

D scarf outline dimension

d diameter

EMI electromagnetic interference

E heat blanket edge distance

e joint eccentricity

FST Fleet Support Team

HEPA high efficiency particulate air

HSS high speed steel

h damage depth

IML inner moldline

Abbreviations/Symbol Definition

L patch overlap

LML lower moldline

mrA mix ratio of part A

mrB mix ratio of part B

mrF mix ratio of filler material

NDI nondestructive inspection

NHMA next higher maintenance activity

OML outer moldline

P load

PCF pounds per cubic foot

PPE personal protective equipment

psi pounds per square inch

r damage layout radius

SRM Structural Repair Manual

Tc thermocouple

T length of taper for partialthickness damage

t skin thickness

w woven

UML upper moldline

Table 1-1. Abbreviations and Symbols

1-3

NAVAIR 01-1A-21

Table 1-2. Reference Material

1-3/(1-4 Blank)

Title Number

Technical Manual, Aircraft and Missile Repair Structural Hardware NAVAIR 01-1A-8

Aerospace Metals - General Data and Usage Factors NAVAIR 01-1A-9

General Use of Cements, Sealants and Coatings NAVAIR 01-1A-507

Aircraft Weapons Systems Cleaning and Corrosion Control NAVAIR 01-1A-509

Nondestructive Inspection Methods NAVAIR 01-1A-16

Organizational, Intermediate and Depot Maintenance, Structure, Typical Repair A1-F18AC-SRM-250

Organizational, and Intermediate Maintenance Structure Repair A1-AV-8B-SRM-250

Organizational, Intermediate and Depot Maintenance Repair Instructions, Horizontal Stabilizer - Boron/Epoxy Structure NAVAIR 01-F14AAA-3-2.4

Temperature/Vacuum Control, Advanced Composite Structural Repair Test Set Operation and Maintenance Instructions NAVAIR 17-1-131

Navy Occupational Safety and Health (NAVOSH) Program Manual (Ashore) OPNAVINST 5100.23B

Navy Occupational Safety and Health (NAVOSH) Program Manual (Afloat) OPNAVINST 5100.19B

Hazardous Material Information System DOD 6050.5

Navy Environmental Health Center Technical Manual NEHC-TM-91-6

Adhesive Bonded Aerospace Structure Repair MIL-HDBK-337

Naval Air Systems Command Technical Manual Program NAVAIR 00-25-100

Distribution of NAVAIR Technical Publications NAVAIRINST 5605.5

Military Standard Abbreviations for Use On Drawings, and in Specifications,Standards and Technical Documents MIL-STD-12

Interchangeability and Replaceability of Component Parts for Aerospace Vehicles MIL-I-8500

1-4

NAVAIR 01-1A-21


2-1

NAVAIR 01-1A-21

SECTION IIINTRODUCTION TO COMPOSITES

(3) Aramid (Kevlar). Aramid material is a syntheticpolymer. Kevlar fibers are made from the polymer using adry-jet wet spinning process. DuPonts Kevlar fiber is0.0005 inch in diameter and is currently the onlycommercially available aramid fiber. The Kevlar fiber, likethe carbon fiber, has a low bending stiffness allowing it tobe woven in various patterns. Although the fiber hasexcellent properties in tension, the compression propertiesare poor limiting its use in structural applications to internalducts, non-structural access covers, fairings and lightlyloaded helicopter skins.

b. Matrix. The material that holds, or supportsthe fibers in the laminate is termed the matrix. Inaddition to providing support for the fibers, the matrixprovides fiber to fiber bonding and bonds the plies orlaminae together forming a laminate. The three matrixmaterials currently in use on naval aircraft (which areall thermosets) are epoxies, bismaleimides andpolyimides. The governing criteria for selection of amatrix material is generally based upon the operatingtemperature of the part in question and the servicetemperature of the material.

(1) Epoxies. Epoxy resins used for matrixmaterials on naval aircraft cure at 350F and have aservice temperature of 250F. The majority ofadvanced composite parts on naval aircraft and incommercial applications use epoxy resins as thematrix material.

(2) Bismaleimides. Bismaleimide resins areintermediate service temperature polyimide resins.They cure at 450F and have a service temperatureof 350F. Generally, they are used on areas of theaircraft that experience operating temperatures inexcess of 250F but less than 350F.

(3) PMR Polyimides. PMR polyimide resinsused as matrix materials are high service temperatureresins. They cure at 650F and have a servicetemperature of 550F. PMR polyimide materials haveseen very limited use as matrix materials and havebeen used exclusively on engine components.

2-1. ADVANCED COMPOSITE MATERIALS (ACM).ACM consist of high strength, high extensional stiffnessfibers imbedded in a matrix or binder material. Thiscomposite of two separate and distinct materials forms asingle new material with properties different from eitherconstituent material. It is the high extensional stiffness ofthe fibers (high resistance to applied loads) that allowadvanced composite materials to replace aluminum orsteel as a structural material. One of the unique featuresof ACM which makes them so appealing to designers isthe ability to tailor laminates by putting the fibers wherethey are needed to carry loads. This results in a structuralmaterial with higher strength and lower weight than currentlyis available using metallic materials.

a. Fibers. The primary function of the fibers is to carryload and to provide the required part stiffness. Carbon,boron and aramid (Kevlar) are the three advanced fibersin use on naval aircraft.

(1) Carbon. The carbon (or graphite) fiber is 0.0003inch in diameter. It is made from a synthetic material similarto rug yarn. The fibers are carbonized in an inert environmentat temperatures around 3000F. The carbon fiber has lowbending stiffness allowing it to be woven into variouspatterns. It can be bent around an object about thediameter of a pencil point and can be readily used tomanufacture complex contoured parts. The majority ofadvanced composite parts used on naval aircraft aremade from carbon fibers.

(2) Boron. The boron fiber is 0.005 inch in diameter.It is made by chemical vapor deposition of elemental borononto a cleaned 0.0005 inch diameter tungsten wire. Theboron fiber has a high bending stiffness and cannot bebent around an object any smaller than the diameter of adime. Due to this relatively high bending stiffness, boronfibers cannot be woven into cloth or used for complexcontoured parts. Their use on naval aircraft has beenlimited to the F-14 horizontal stabilator.

2-2

NAVAIR 01-1A-21

c. Adhesives. Film adhesives are used in theconstruction of parts manufactured from ACM tobond honeycomb core and/or substructure membersto laminate skins. These adhesives may be cocuredduring the laminate curing process or they may besecondarily cured after the laminate curing processis complete. Epoxies are the predominant materialsused for adhesives on naval aircraft.

(1) Epoxy film adhesives that cure at 350Fgenerally have a service temperature of 250F. Anoted exception is the adhesive system used to bondF-14 honeycomb sandwich parts. This system has aservice temperature of 350F.

(2) Epoxy film adhesives that cure at 250Fgenerally have a service temperature of 180F.

2-2. PREPREG. Prepreg material is the basic buildingblock of advanced composites. It consists of fiberspreimpregnated with a partially cured (B-staged) matrixmaterial. It is supplied by a prepreg manufacturer in thin

sheets in two different forms, unidirectional prepreg(fibers all in one direction) and woven prepreg (fiberswoven into a specified weave). The fibers in both of theseforms are continuous.

2-3. LAMINATE. Skins and substructure details (ribs,spars, etc.) are manufactured by laminating plies of prepreg.Plies of prepreg are cut to the required orientation andshape and stacked together in specified directions toobtain the required stiffness and strength. This stackup isthen cured in an autoclave using heat and pressure(100-200 psi), to form a solid laminate. Excess resin bledduring the cure process bonds the plies together forminga laminate. Typical ply orientations used for aircraft partsare 0, 90, +45, and -45 degrees (see paragraph 6-7d fora description of ply orientation). A cross section of alaminate made from unidirectional carbon/epoxy prepregis shown in Figure 2-1. Typical thickness per ply for a curedlaminate made from unidirectional carbon/epoxy prepregis 0.005 inch. Note the thin resin bands in between eachply. Interply bonding occurs in these resin band areas.

0.005INCH

INTERPLYRESINBAND

MATRIX

FIBER

Figure 2-1. Laminate Cross Section Cut 90 Degrees to the 0 Degree Fiber Direction

0

+45

-45

90

NAVAIR 01-1A-21

3-1

SECTION IIITYPICAL DAMAGE

3-1. GENERAL. Most damage to naval aircraft occurson the ground during aircraft servicing, maintenance andhandling. Impacts from dropped tools, forklifts, maintenancestands, dropped panels and in-flight foreign objects aremajor causes of damage.

3-2. DAMAGE TO ADVANCED COMPOSITEMATERIALS (ACM). Damage incurred by ACM is quitedifferent than that experienced by metallic materials. ACMdo not deform like metals. ACM either resist an impact forceand spring back or rupture. Due to the brittle nature of mostACM, these ruptures can occur at rather low impact energies.The damage resulting from a rupture produces cracks inthe matrix, delaminations between plies and broken fibers.

a. Impact Related Delaminations. Impact relateddelaminations tend to occur at multiple depths throughoutthe thickness of the laminate. They consist of inter and intraply matrix cracks which are not always interconnected.Subsurface delaminations and matrix cracks can existwithout any indication on the part surface. (See Figure 3-1).Laminates subjected to or suspected of having beensubjected to impacts from foreign objects (such asmaintenance stands, tool boxes, dropped tools, etc.) musthave the suspected area inspected by an appropriatemethod (usually ultrasonics). Inspect for the presence ofsubsurface damage per paragraphs 3-3b and 3-3c as wellas the part specific structural repair manual (SRM). Impactrelated delamination damage is the most common type ofdamage experienced with ACM.

B. Cross Section of Impacted AreaC. Magnified Cross Sections of Impacted Area

INTERPLYDELAMINATION

INTRAPLYMATRIXCRACKS

INTERPLYDELAMINATION

INTRAPLYMATRIXCRACKS

Figure 3-1. Typical Impact Damage on 0.1 Inch Thick Carbon/Epoxy Laminate, 3 Ft-Lbs Impact Energy

A. Impacted Laminate: No Indication of SubsurfaceDamage

EXTENT OFNDI INDICATION

NAVAIR 01-1A-21

3-2

DENT/SKINPENETRATION

BROKENFIBERS

b. Dents. The presence of dents in composite skinscan indicate delaminations and/or matrix cracks. Ifhoneycomb core is present, crushing of the core will existbeneath the dented composite skin (see Figure 3-2). Thecrushed core can be hidden by the brittle ACM separatingin the laminate, springing back and masking the buckledcore beneath (see Figure 3-3).

c. Penetration Damage. Composite skin penetrationdamage is characterized by broken fibers, matrix cracksand delaminations. This type of damage usually results insubsurface delaminations and matrix cracks larger than

that apparent visually (see Figure 3-3). Mapping of theseareas is required as described in paragraph 3-3c(3)(c).Visually apparent penetration damage on the exit side istypically 3-5 times larger than entrance side damage size(see Figure 3-4).

d. Airstream Stripping Damage. Penetrations thatoccur in-flight result in damage much larger than the actualpenetration. The air flow over the part lifts the outer ply ofthe composite at the edge of the penetration and strips itback off the part (see Figure 3-5).

Figure 3-3. Penetration Damage

MATRIXCRACK

DELAMINATION

DENT

BUCKLED CORE

Figure 3-2. Dented Honeycomb Panel

DELAMINATION

BUCKLED CORE

NAVAIR 01-1A-21

3-3

Figure 3-5. Airstream Stripping Damage

Figure 3-4. Penetration Damage on 0.75 Inch Thick Carbon/Epoxy Laminate

A. Entrance Side B. Exit Side

OUTER PLYSTRIPPING

CAUSED BYTHE AIRSTREAM

SKINPENETRATIONS

NAVAIR 01-1A-21

3-4

e. Edge and Corner Damage. Edge and cornerdamage to panels result in edge delaminations and/orbroken off pieces sometimes requiring a rebuilding effort(see Figure 3-6).

f. Partial Thickness Damage. Partial thicknessdamage due to gouging of the part surface results in outerply splintering, broken/removed fibers and delaminations.

g. Resin Damage. Damage to the resin can occur dueto the effects of heat and chemical attack.

(1) Temperature exposures in excess of the partcure temperature can degrade matrix strength. Epoxymatrix materials that cure at 350F and that are exposed to

temperatures above 400F but less than 600F canexperience a marked reduction in strength. Little or novisual indication of damage to the laminate is apparent.However, if the laminate is painted, discoloration of thepaint system provides an indicator that laminate damagemay have occurred. For 350F curing epoxy matrix materials,exposures beyond 600F may result in visual blistering andpyrolyzation of the outer plies of the laminate. For bondedcomposite assemblies, epoxy adhesives degrade at lowertemperatures (typically 50F lower) than laminate matrixmaterials. ACM and bonded composite assembliesexhibiting discolored paint or that are suspected of beingexposed to excessive temperatures (above 400F forepoxy matrix composites or 500F for bismaleimide matrixcomposites) may have experienced heat damage. TheseACM and bonded composite assemblies are suspect. The

A. Honeycomb Sandwich Panel Edge

Figure 3-6. Edge Damage (Sheet 1 of 2)

NAVAIR 01-1A-21

3-5

matrix material beneath any blistered, delaminated orpyrolyzed plies are suspect as well. The suspect areasrequire evaluation and disposition by Fleet Support Team(FST) engineering to evaluate laminates, laminate to corebonds and laminate to substructure bonds.

(2) ACM exposed to aircraft fires may experiencepyrolyzation of the matrix material on the outer plies, thuseliminating support for the fibers (see Figure 3-7), while theunderlying plies experience blistering and interplydelamination. Although appearing intact, the matrix materialbeneath blistered, delaminated or pyrolyzed pliesexperience a strength reduction and requires evaluationand disposition as discussed in paragraph 3-2g(1).

(3) ACM exposed to chemical paint strippersexperience a long term degradation of matrix strength withno visual indication on the part surface. They requireevaluation and disposition by FST engineering.

(4) ACM exposed to most other chemicals found inthe maintenance environment (cleaners, solvents, jet fuel,hydraulic fluid, engine oil, etc.) show no effect due toexposure.

h. Disbonds. Skin to core disbonds and skin tosubstructure disbonds occur as a result of core corrosion,or part exposure to temperatures at or above the part curetemperature. They can also be caused by impact forces ifthe adhesive used to bond the core or substructure is morebrittle than the composite matrix material. Disbondindications between composite skins and composite closuremembers bonded with FM300 adhesive should be checkedclosely. In most cases (contamination being the exception),the indications are either skin delaminations or closuremember delaminations as the FM300 adhesive is tougherthan the matrix material.

B. Monolithic Panel Edge

Figure 3-6. Edge Damage (Sheet 2)

NAVAIR 01-1A-21

3-6

i. Fluid Intrusion. This type of damage occurs withhoneycomb sandwich assemblies when a leak pathdevelops which allows fluid to enter the honeycomb corecells. This can be detrimental to weight critical flight controlsurfaces, as well as causing material degradation to bothmetallic and non-metallic honeycomb core. Fluid intrusionis of major concern in performing elevated temperaturecures during bonded repairs.

j. Fastener Hole Damage. Gouging of countersinksand areas surrounding fastener holes caused by fastenerdrivers can occur during fastener removal and installation.Repeated fastener removal and installation can result inexcessive wear of fastener holes. Delamination of fastenerholes can be caused by over torquing fasteners duringinstallation and by generation of an excessive amount ofheat when drilling out damaged fasteners.

3-3. DAMAGE ASSESSMENT.

a. Damage Categories. Prior to actually beginning arepair, the damage should be assessed and thencategorized to determine if the repair is required/feasible.Three types of damage are categorized below:

(1) Negligible Damage. Damage which,because of its size, nature and location that does notadversely affect the structural integrity of the part isdefined as negligible. It may be allowed to existwithout repair, or may only require a cosmetic repairto be performed to prevent further damage fromoccurring (such as further stripping of outer plymaterial). Refer to the part specific SRM for furtherguidance on what constitutes negligible damage.

Figure 3-7. Fire Damaged Advanced Composites

NAVAIR 01-1A-21

3-7

(2) Repairable Damage. This is defined not onlyas damage requiring repair, but also damage that is withinthe repair capability of the activity at which the repair is tobe performed. The location of damage, complexity of therepair procedure, repair weight limitations, availability ofrepair equipment and materials, repair time/cost, sparepart availability, etc., are all factors in deciding whether apart is beyond the capability of maintenance (BCM) at thatactivity. Parts that are BCM must be forwarded to the nexthigher maintenance activity (NHMA). Refer to the partspecific SRM for guidance on repairability due to damagelocation and to provide specific repair weight limits.

(3) Non-Repairable Damage. Parts determined to benon-repairable must be forwarded to depot level for disposition.

b. Damage Assessment Methodology. Four stepsare involved in assessing damage. Locating damage,characterizing the damage and determining its extent,zoning the damage on the part being repaired and re-evaluation of the damaged area after damage removal.

(1) First, locate the damage. This is usuallyperformed by visual inspection. However, caution must beexercised as non-visible subsurface damage may existbeneath impact areas and areas suspected of having beenimpacted. Areas impacted (with or without visual indicationon the part surface) or suspected of having been impactedmust be further evaluated for delaminations and matrixcracks. Use the nondestructive inspection (NDI) methodslisted in paragraph 3-3c as well as the part specific SRM.

(2) Once the damage has been located, the extentof the damage must be determined and the damagecharacterized. The depth of delamination and the presenceof skin to core or skin to substructure disbonds (if applicable)should be determined to characterize the detected damage.Damage to honeycomb core should be characterizedusing radiographic techniques. Determining the extent ofdamage and characterizing the damage is an importantpart of the damage assessment process, as it will have adirect bearing on the repair procedures to be employed.

(3) After the damage has been characterized andthe extent determined, the repair zone in which the damageis located is determined using the part specific SRM.Overlap of damage from one repair zone to another requiresthe damage limits for the worst case zone be used. If thedamage limits for the repair zone in which the damage islocated are exceeded, the part must be forwarded to theNHMA for repair. If the damage lies in a non-repairablezone, the part must be forwarded to depot for disposition.

(4) Following damage removal, reinspect thedamage area to ensure all the damage was in fact removed.Current NDI methods used to detect subsurfacedelaminations are capable of only finding the firstdelamination nearest the surface on which the probe wasapplied. Deeper delaminations can be masked by the firstdelamination (see Figure 3-8, View D). After removingwhat initially appears to be all the damage present, it isnecessary to reinspect the area to ensure no delaminationsremain below the originally defined damage.

c. Damage Assessment Techniques. The followingnondestructive methods are used to inspect ACM to evaluatedamage. The ultrasonic inspection techniques may also beused following a bonded repair to evaluate the adequacy ofthe performed repair.

(1) Visual Inspection. As discussed above, visualinspection is used to initially locate damage. Penetrationdamage is readily apparent. The presence of dents requiresa closer look and can be aided by using a straight edge overa suspected dent area and comparing the suspected dentarea with the surrounding part. The visual method can beenhanced by using a flashlight and magnifying glass. Thepresence of edge delaminations may sometimes bedetected by wiping the edge of the part with a solvent. If theedge is delaminated, the solvent will wick into thedelaminated area. The solvent will evaporate on theundelaminated edge area leaving a wet mark along thedelaminated edge. NDI penetrants should not be used asthey may contaminate surface cracks or edge delaminations,foiling subsequent repair attempts. Internal flaws such asdelaminations not open to an edge and skin disbondscannot be detected using visual methods and require theuse of ultrasonic inspection techniques.

(2) Coin Tap. This method can be used to detectthe presence of disbonds and/or delaminations in bondedhoneycomb sandwich assemblies with thin compositefacesheets. It is not effective for thick laminates and cannotdetermine defect depth or distinguish a disbond from adelamination. The technique involves lightly tapping thesurface of a composite in the area of a suspected defectand comparing the acoustic response due to tapping in thesuspect area with the acoustic response from a good area.Good areas have a sharp glassy ring to them when tapped.Areas containing disbonds or delaminations have a dull orflat sound when tapped. Caution must be exercised whenusing coin tap. Experience has shown that areas thatprovide a defect indication do in fact contain defects.However, areas that sound good by coin tap may stillcontain disbonds and delaminations and must be inspectedusing ultrasonic techniques.

NAVAIR 01-1A-21

3-8

Figure 3-8. Pulse-Echo Ultrasonic Inspection with A-Scan Presentation (Sheet 1 of 3)

A. Composite Laminate/No Defects

B. Composite Laminate/Delamination at 0.060 Inch

C. Composite Laminate/Multi-Level Delamination

TRANSMITTER/RECEIVERTRANSDUCER

0.160INCH 0 1 2 3 4 5 6 7 8 9 10

0.160INCH

0102030405060708090

100DELAY LINE/PART SURFACEINTERFACE RESPONSE

BACK SURFACERESPONSE FROMCOMPOSITELAMINATE

CRT DISPLAY

0 1 2 3 4 5 6 7 8 9 100

102030405060708090

100

RESPONSE FROM0.040 INCH DEPTH

CRT DISPLAYMULTI-LEVEL DELAMINATION

AT 0.020,0.040, AND 0.060 INCH




0.160INCH0.160INCH

0.160INCH 0 1 2 3 4 5 6 7 8 9 10

0.060INCH

0102030405060708090

100 RESPONSE FROM0.060 INCH DEPTH

CRT DISPLAY

0.060INCH


DELAMINATIONAT 0.060 INCH

NAVAIR 01-1A-21

3-9

Figure 3-8. Pulse-Echo Ultrasonic Inspection with A-Scan Presentation (Sheet 2of 3)

D. Composite Laminate/Masked Delaminations

E. Repair Patch to Laminate Bond/No Defects

F. Repair Patch to Laminate Bond/Bondline Void or Disbond


0 1 2 3 4 5 6 7 8 9 100

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100 RESPONSE FROM0.060 INCH DEPTH

CRT DISPLAY

0.060INCH

DELAMINATIONAT 0.060 INCH

DELAMINATIONS MASKEDBY DELAMINATION AT 0.060 INCH

0.160INCH


0 1 2 3 4 5 6 7 8 9 100

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100

CRT DISPLAY

ABSENCE OFLAMINATEBACK SURFACERESPONSE

PATCH TO LAMINATE BONDLINEVOID OR DISBOND

PATCHSURFACERESPONSE

PATCH/ADHESIVERESPONSE

PATCHADHESIVE

LAMINATE


0 1 2 3 4 5 6 7 8 9 100

102030405060708090

100 PATCHSURFACERESPONSE

PATCH/ADHESIVERESPONSE

CRT DISPLAY

PATCHADHESIVE

LAMINATE

LAMINATEBACK SURFACERESPONSE

NAVAIR 01-1A-21

3-10

between the transducer and the part as the air gapsinterrupt sound being induced into the part. An internaldefect, such as a delamination, interrupts the sound travelingthrough the laminate and an indication is received on theultrasonic unit indicating the presence of a defect.Interpretation of ultrasonic inspection indications requiresa reference standard made from the same material as thepart being tested and containing flaws of known size andlocation for comparison purposes. Ultrasonic techniquesrequire a certified NDI technician to perform the inspectionsand are performed using procedures in part specific SRMs.General procedures are provided in NAVAIR 01-1A-16.Two techniques are in use in the field, pulse-echo ultrasonicsand through-transmission ultrasonics.

NOTE

The following summary provides an overviewonly and is not sufficiently detailed forperforming ultrasonic or radiographic inspectionof ACM. These inspections require a trainedand certified NDI technician.

(3) Ultrasonics. Currently, the contact ultrasonicinspection techniques are the most widely used methodsfor assessing damage to ACM in the field. This methoduses high frequency sound waves, called ultrasound,transmitted into the part by a transducer placed in contactwith the part. A liquid couplant is used to eliminate air gaps

Figure 3-8. Pulse-Echo Ultrasonic Inspection with A-Scan Presentation (Sheet 3)

H. Composite Laminate Honeycomb Sandwich Assembly/Skin to Core Disbond


0 1 2 3 4 5 6 7 8 9 100

102030405060708090

100 OVER SKIN AREAWHERE ADHESIVE ISDISBONDED FROMBACK SURFACE OFSKIN, AMPLITUDE MAYINCREASE &ADHESIVE/AIRINTERFACE RESPONSEWILL BE ABSENT

CRT DISPLAY

SKIN TO COREDISBOND

G. Composite Laminate Honeycomb Sandwich Assembly/No Defects


0 1 2 3 4 5 6 7 8 9 100

102030405060708090

100 COMPOSITE SKINBACK SURFACERESPONSE

CRT DISPLAY

ADHESIVE/AIRINTERFACERESPONSE

SKIN TO COREBOND ADHESIVE ADHESIVE/AIR

INTERFACE

NAVAIR 01-1A-21

3-11

(a) Pulse-Echo Ultrasonics. This is the mostcommon technique used in the field. It makes use of asingle transducer that sends and receives sound energy.The sound energy is reflected back to the transducer by theinitial surface of the part, by the backside surface of the part,by interfaces between different materials (such as compositeand adhesive interfaces) and by locations of internal defects.It can be used to determine defect area, and defect depth.Field inspection test results are displayed using A-scanpresentation on the cathode ray tube (CRT) of an ultrasonicflaw detector. When this method is used by a skilledtechnician, a skin delamination can usually be distinguishedfrom a skin to core or skin to substructure disbond. Figure 3-8shows typical pulse-echo A-scan presentation CRT displaysfor the defects indicated.

(b) Through-Transmission Ultrasonics. Thistechnique uses two transducers, one to transmit soundenergy and one to receive. Sound energy is sent from one

side of the part through the part to the second transduceron the opposite side. A defect encountered in the partsignificantly reduces the intensity of the sound energy. Thisreduction in sound energy intensity is used to detect thepresence of defects. It can be used to determine the areaof a defect. Unlike pulse-echo ultrasonics, the sound doesnot have to traverse the thickness of the part twice. Through-transmission ultrasonics is usually more sensitive for flawdetection in bonded assemblies. This technique is not ableto determine defect depth or type. In addition, it requiresaccess to both sides of the part and alignment of the twotransducers during inspection to ensure the receivingtransducer picks up the sound energy sent by the transmittingtransducer (see Figure 3-9, View B). Like pulse-echoultrasonics, field inspection test results are displayed usingA-scan presentation. See Figure 3-9, View C for typicalthrough-transmission A-scan presentation CRT responsesfor the defects indicated.

A. Composite Laminate Honeycomb Sandwich Assembly with No Defects

Figure 3-9. Through-Transmission Ultrasonic Inspection, A-Scan Presentation (Sheet 1 of 2)

TRANSDUCER MISALIGNMENT CAN CAUSEAN ERRONEOUS DEFECT INDICATION

SEARCH UNITALIGNMENT

FIXTURE USEDTO MAINTAINTRANSDUCERALIGNMENT

B. Transducer Alignment

0 1 2 3 4 5 6 7 8 9 100

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100

OPPOSITESIDE SURFACERESPONSE

CRT DISPLAY

TRANSMITTERTRANSDUCER

RECEIVERTRANSDUCER

PART SURFACERESPONSE

T

T

NAVAIR 01-1A-21

3-12

(c) Defect Mapping. Using the visual indicationof damage as a guide, mark a grid of 0.5 inch squares outto a point at least 1 inch away from the edge of the damageon the part surface using a marking pen. Using one of theultrasonic techniques listed above, inspect each 0.5 inchsquare of the marked grid. Mark the location of defectindications (and depths if using the pulse-echo ultrasonictechnique) as indicated on the CRT on the part surfaceusing a marking pen (see Figure 3-10). Tape a piece ofmylar over the damage area and transfer the defect indicationmap to the mylar using a permanent marking pen. Marklocating lines on the part away from the damage area.Transfer these lines to the mylar to aid in positioning themylar after the paint has been removed from the part duringthe repair process. The defect map will be used to determine

the damage layout and extent of material removed asdescribed in paragraph 6-3. The mylar map provides an aidwhen performing subsequent NDI of the area after damageremoval and after repair as well as providing a permanentrecord of the defect indications.

(4) Radiographic Techniques. This techniquemakes use of x-rays to detect defects in materials andassemblies. X-rays penetrate a material and are absorbeddifferently based upon the material's density. Defects lessdense than surrounding material (such as voids in acomposite laminate) absorb less radiation and are shownon x-ray film as darker areas as compared to nearbyimages. Defects more dense than surrounding areas absorbmore radiation (such as water in honeycomb sandwich

Figure 3-9. Through-Transmission Ultrasonic Inspection, A-Scan Presentation (Sheet 2)

C. Composite Laminate Honeycomb Sandwich Assembly/Skin to Core Disbond and Skin Delamination

0 1 2 3 4 5 6 7 8 9 100

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100

ABSENCE OFOPPOSITE SIDESURFACERESPONSE

CRT DISPLAY

CTR Response for Both Skin to CoreDisbond and Skin Delamination

DELAMINATION

SKIN TO CORE UNBOND

Through-Transmission Ultrasonics Is Unableto Determine Depth of Defect or Defect Type

NAVAIR 01-1A-21

3-13

assemblies) and are shown as lighter areas on x-ray filmwhen compared to nearby images. The recording mediumin the field is film. The need to precisely orient the part toobtain the required sensitivity to detect defects may precludethe use of radiographic techniques for inspection of complexcontoured parts. In addition, the requirement for backsideaccess to position x-ray film limits its usefulness. Thefollowing types of defects are detectable using radiographictechniques:

(a) Voids in Patch to Part Bondlines. Voidscontain less material than the surrounding adhesive and showup as darker areas when compared to areas lacking voids.

Figure 3-10. Defect Mapping of Damage Extent from NDI

(b) Water Entrapment in HoneycombSandwich Assemblies. The water present is excess oradded material for the x-rays to penetrate and appearslighter when compared to images of adjacent cells notcontaining water.

(c) Honeycomb Core Damage. Damage tometallic core material, such as blown, crushed, corroded,fatigued, or distorted core material, is best detected whenthe cell walls appear to be laid over on the x-ray film. (Thearea of interest should be offset from the central ray of thex-ray beam to allow viewing of the cell walls and determiningif damage is present).

A. Layout of NDI Grid

B. Damage Extent from Ultrasonic Inspection

ULTRASONICINDICATIONOF DAMAGE

VISUALINDICATIONOF DAMAGE

3-13/(3-14 Blank)

NAVAIR 01-1A-21

3-14


NAVAIR 01-1A-21

4-1

SECTION IVREPAIR PROCEDURE SELECTION CRITERIA

4-1. DESIGN CRITERIA. The design of a repairprocedure is dependent upon many criteria. The followingcriteria are used when designing repairs defined in partspecific structural repair manuals (SRMs) and by FleetSupport Team (FST) engineers.

a. Strength. An important aspect of any repair isrestoring strength to the part. Some composite parts aredesigned such that the full laminate strength is needed tocarry the part load. These are referred to as strengthcritical parts. However, other composite parts are designedas stiffness critical. A damaged stiffness critical part maynot require restoration to full strength as the part was notdesigned to carry high loads. Examples of structures forwhich full strength restoration is not required are asfollows: lightly loaded secondary structure (non-structuralaccess covers, doors, and fairings), some areas of flightcontrol surfaces designed for resisting deflections ratherthan carrying high loads and some areas of fuselage skinsor wing skins (skin thickness sized for resistance tohandling damage rather than carrying loads). Part specificSRMs zone the part to identify the type of repair needed.This permits the use of simpler repairs in areas where fullstrength restoration is not required. Zoning can also beused to define other requirements such as weight andbalance or moldline protrusion limitations.

b. Stiffness. It is always necessary to restore astructure to its required stiffness.

(1) Fixed structures such as wings and verticaltails have required bending and torsional stiffnesses toprevent excessive deflections when loaded. Improperstiffness restoration can result in excessive deflection orpossible structural failure.

(2) Actuated doors (such as landing gear doors)have stiffness requirements to prevent excessive deflectionduring actuation or during application of aerodynamicloads. Lack of stiffness restoration can result in improperdoor function, an increase in aerodynamic drag or possiblestructural failure.

(3) Flight Control Surfaces. These surfaces aresensitive to aerodynamic flutter and their stiffnesses aredesigned to prevent flutter from occurring. Any significantchange in part stiffness can result in improper function ofthe control surface or possible structural failure.

(4) Load Path Changes. As a general rule, thelocal stiffness of the patch is designed to correspond to thatof the surrounding material in order to avoid load pathchanges. This is especially important when restoration ofstrength is required. Attention must be given to the effectof the stiffness of the repair patch on the load distributionin the structure. If the patch has less stiffness than thesurrounding structure (such as use of a fiberglass epoxypatch on a carbon epoxy structure), the patch may notcarry its share of the load and overload of the surroundingstructure may result. Conversely, an excessively stiffpatch may attract more than its share of the load causingadjacent areas to which it is attached to be overloaded.

c. Weight and Balance. The weight addition of mostrepairs is small in comparison to overall aircraft weight.However, the mass balance of most flight control surfacesis such that very little weight addition can adversely affectthe balance of the part. Caution must be exercised whenrepairing flutter sensitive flight control surfaces (such asailerons, rudders and stabilators). Both the weight additionand the distance of the repair from the hinge axis of the partmust be considered to prevent premature failure of thepart. The further aft of the hinge axis the repair is located,the less additional repair weight is allowed. Consult thepart specific SRM or FST engineering for further guidance.

4-2. ADDITIONAL CRITERIA. In addition to designcriteria, other criteria which must be considered includethe following:

a. Moldline Protrusion. High-performance aircraftdepend upon smooth external surfaces to minimize drag.Some areas are more critical than others depending uponaerodynamic considerations. In general, the most criticalareas from an aerodynamic moldline protrusion standpointare leading edges of aerodynamic surfaces (wings, flightcontrol surfaces, flaps, vertical tails) and engine inletareas. In addition to aerodynamic considerations, someaccess covers and doors have moldline fit up constraintswith other parts of the structure that preclude repairs fromextending above the surface of the part. Requirements forlow observable structures also provide moldline protrusionconstraints. The use of a flush repair may be required tomeet some of these requirements.

NAVAIR 01-1A-21

4-2

b. Aircraft Systems. The effects of the repair onaircraft systems must be considered.

(1) Fuel Systems. Repairs performed in areasused to contain fuel, must seal adequately to prevent fuelleaks. The repair design must take into account fuelpressure loads as well as interference with fuel systemcomponents due to geometric constraints. Protection offuel system components against high repair cure cycletemperatures must be considered during repair selection.

(2) Mechanical Systems. Mechanically actuatedparts, such as landing gear doors and control surfaces,must function correctly after repair. In addition to fit problemsmentioned above, the parts may required rerigging orrebalancing after repair.

(3) Protection Systems. Composite structures thathave fire suppression, survivability, noise suppression,electromagnetic interference (EMI) shielding, lowobservable technology or lightning protection must havethose systems restored to their original function in additionto restoring part stiffness and strength.

c. Part Manufacturing Methods. Advanced compositeparts used on naval aircraft consist of laminates or skinsand some form of substructure members. Joining of theseskins to substructure members (spars, ribs, honeycombcore, integral stiffeners, etc.) involves either a bolting orbonding process. In general, bolted assemblies use boltedrepair concepts and bonded assemblies use bondedrepair concepts.

4-3. BASIC REPAIR JOINTS. Restoration of strengthand/or stiffness to damaged parts requires damagedareas to be joined for load path continuity. This restorationinvolves a repair joint and the joining of the patch materialto an undamaged area of the part through this joint (seeFigure 4-1). Load travels along the undamaged skinsneutral axis through a joining material (adhesive ormechanical fasteners) into the patch. The patch providesa bridge for the load across the damaged area. Loadtravels along the neutral axis of the patch, through thejoining material and back into the part skin. As can be seenin Detail A of Figure 4-1, the load in the skin and the loadin the patch are horizontally opposed to one another. Thissets up a shear force in the joining material. Some basicjoints used for repair and their load transfer mechanismsare described below.

a. Bonded Joints. Bonded joints make use of a repairpatch to carry load across the damaged region. A commonmisconception is that the adhesive is applied to hold thepatch in place on the part. While that is one function of the

adhesive, the primary function is to transfer load from theundamaged part to the patch. Load transfer is accomplishedvia shearing action in the adhesive. Three commonly usedbonded repair joints are shown in Figure 4-2. Only thosebonded repair joints and patch/adhesive materials specifiedin either a part specific SRM or by FST engineering shallbe used for repair.

(1) External Bonded Patch. The external bondedpatch is the easiest bonded repair joint to fabricate.Adhesive bondline area for load transfer is provided byoverlapping the repair patch and the part skin. One of the

AA

COMPOSITESKIN

P P

BASIC REPAIR JOINT

PATCHDAMAGEHOLE

PATCH BRIDGESDAMAGE HOLE

SKIN

HOLEEDGE

CENTERLINE OFDAMAGE CLEANUP

HOLEDETAIL A

SECTION A-ACROSS SECTION OF REPAIR JOINT

PP

DETAIL A

DIRECTIONOF TRAVELOF LOAD P

PATCHNEUTRAL

AXISSKIN

NEUTRALAXIS

HOLEEDGE

Figure 4-1. Basic Repair Joints (General)

NAVAIR 01-1A-21

4-3

limiting factors for this type of joint is the eccentric load pathcaused by the offset of patch and part neutral axes due tothe overlap. (See Figure 4-3). This eccentricity can resultin interlaminar failure of the patch or part skin, or prematurefailure of the adhesive. Another limiting factor in the use ofan external bonded patch is the stress concentration thatexists at the edge of the patch and the edge of the damagecleanup hole in the part skin. (See Figure 4-4). This stressconcentration can be reduced at the edge of the patch bygradually increasing the thickness at its edge (by steppingor tapering the edge of the patch) to ease the load into therepair joint, but the stress concentration at the hole edgeremains and limits the load carrying capability of theexternal bonded patch.

(2) Internal Bonded Patch. The internal bondedpatch is similar to the external bonded patch and has thesame limiting factors. However, the internal bonded patchis very difficult to incorporate without backside access andvirtually impossible on honeycomb sandwich parts due tointerference with the core. For these reasons, the use ofinternal bonded patches has been limited.

(3) Scarf or Step Bonded Joints. A scarf joint, (orstep joint) machined in the part skin reduces the stressconcentration and the adhesive shear stress at the edgeof the damage cleanup hole. In addition, the scarf or stepjoint almost eliminates joint eccentricity as the patch andpart skin neutral axes are nearly coincident. Adhesivebondline area is provided along the scarf or step surfaceswithin the thickness of the part skin. (See Figure 4-5). Ascarf or step joint can result in joints as strong as theoriginal part skin. The machining of the scarf or step jointin the part skin is time consuming, must be done withaccuracy and removes a large quantity of sound material.A major disadvantage to using this type of joint is the needto very accurately layup and position replacement plymaterial in the repair joint. In addition, curing of replacementply material can result in significantly reduced strength ifnot cured in an autoclave. For these reasons, this type ofjoint is usually performed only at depot level unless the partis lightly loaded. The scarf joint is preferred over the stepjoint as it is more efficient from a load transfer standpoint,and is easier to fabricate. Step joints have been used forrepair of parts fabricated from woven materials utilizingKevlar and glass fibers.

b. Bolted Joints. Bolted joints also make use of arepair patch to carry loads across the damage region.Fasteners attaching the repair patch to the damagedstructure complete the joint. Three commonly used boltedrepair joints are shown in Figure 4-6. A commonmisconception is that the purpose of the fasteners are tohold the patch in place. The fasteners do provide clamp-up between the plate and the part surface, but their primarypurpose is to allow the load to be transferred from theoriginal part surface through the patch. This load istransferred through the fasteners and patch by shearforces as the fasteners contact the loaded structure andthe plate at the edge of the fastener hole. This load transferis illustrated in Figures 4-7 and 4-8. These forces aretransferred more efficiently with tighter fastener holetolerances. For fastener holes other than interference fitfasteners, a deflection in the structure and plate is requiredfor the fasteners to contact the loaded structure and theplate. Figure 4-9 illustrates how a smaller deflection isrequired to load the adjoining structure for fasteners withsmaller initial clearances than another hole with a largerfastener clearance. However, interference fit fastenersshould not be used in composite structures. The compositematrix may crack or delaminate resulting in strengthdegradation as the interference fit fastener is pressed intoplace. A Class II fit (clearance of +0.0025/-0.000 inch) isnormally specified for structural fasteners in compositemater