ADVANCE COMPOSITE REPAIR

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Aircraft Composite Repair Technology

CHAPTER 1: INTRODUCTION.................................................................................5 REFERENCES.....................................................................................................8

CHAPTER 2: FITTER & FINISH PROCESS .............................................................9

2.1 Introduction ....................................................................................................9 2.2 Trimming Part.................................................................................................9 2.3 Smoothing And Edge Filling............................................................................9 2.4 Application of Resin......................................................................................10 2.5 General Rules and Procedures of Smoothing and Edge Filling......................11

CHAPTER 3: INSERT INSTALLATION...................................................................12

3.1 Insert ............................................................................................................12 3.2 Installing the Inserts......................................................................................13

3.2.1 Mechanically Installed Inserts (Grommet Type)......................................13 3.2.2 Molded in (Potted Inserts Bonded) .........................................................14

CHAPTER 4: COMPOSITES DAMAGE REPAIR ....................................................16

4.1 Introduction ..................................................................................................16 4.2 Types of Composite Repair...........................................................................17

4.2.1 Hot Bond Repair....................................................................................17 4.2.2 Cold Bond Repair ..................................................................................18

CHAPTER 5: DAMAGE ASSESMENT....................................................................21

5.1 General ........................................................................................................21 5.2 Defect & Damage Classifications ..................................................................21

CHAPTER 6: TYPES OF DEFECT & DAMAGE......................................................30

6.1 Introduction ..................................................................................................30 6.1.1 Blisters ..................................................................................................30 6.1.2 Tedlar Wrinkles .....................................................................................31 6.1.3 Fabric Wrinkles......................................................................................31 6.1.4 Rich Resin Areas...................................................................................32 6.1.5 Resin Ridge...........................................................................................32 6.1.6 Resin Starved Areas..............................................................................32 6.1.7 Tacky Areas ..........................................................................................33 6.1.8 Scratches/ Nicks or Gorges ...................................................................33 6.1.9 Cracks...................................................................................................34 6.1.10 Fractures/ Punctures ...........................................................................34 6.1.11 Delamination .......................................................................................34 6.1.12 Incorrect Ply Orientation ......................................................................35 6.1.13 Skin-Core delamination/ Disbond.........................................................35 6.1.14 Core depression ..................................................................................35 6.1.15 Core Crushing .....................................................................................36 6.1.16 Core Displacement ..............................................................................36 6.1.17 Core Nodal Delamination.....................................................................36 6.1.18 Bridging...............................................................................................37 6.1.19 Pitting located in Center of Cells ..........................................................37 6.1.20 Blisters in the center of cells ................................................................37 6.1.21 Telegraphing .......................................................................................37 6.1.22 Porosity ...............................................................................................38 6.1.23 Foreign Object Inclusions ....................................................................38 6.1.24 Geometrical Deviations........................................................................38

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Aircraft Composite Repair Technology 6.2 Damage Terminology in Maintenance Activity ...............................................38

6.2.1 Cosmetic Defects ..................................................................................39 6.2.2 Impact Damage.....................................................................................39 6.2.3 Delamination and Disbond.....................................................................40 6.2.4 Cracks...................................................................................................40 6.2.5 Hole Damage ........................................................................................41 6.2.6 Water Ingression Damage .....................................................................41 6.2.7 Lightning Strike Damage........................................................................42 6.2.8 Abrasion................................................................................................42 6.2.9 Burn Marks............................................................................................43 6.2.10 Chemical Attack Abrasion....................................................................43

CHAPTER 7: INSPECTION METHODOLOGY........................................................45

7.1 Non Destructive Inspection ...........................................................................45 7.2 Visual Inspection ..........................................................................................45 7.3 Tap Test.......................................................................................................46 7.4 Ultrasonic Inspection ....................................................................................47

7.4.1 Ultrasonic Pulse-Echo Inspection...........................................................47 7.4.2 Through Transmission Ultrasonic Bond Inspection.................................47

7.5 Radiography or X-ray....................................................................................48 7.6 Infrared/ Thermography ................................................................................49 7.7 Laser Shearography .....................................................................................49 7.9 Dye Penetrant ..............................................................................................50 7.8 Hardness Testing .........................................................................................51

CHAPTER 8: PREPARATION AND GENERAL REPAIR PROCEDURES ...............53

8.1 Determine Damage ......................................................................................53 8.2 Determine Repair Area Configurations..........................................................53 8.3 General Preparation .....................................................................................54

8.3.1 Material Preparation ..............................................................................54 8.3.2. Facilities, Equipment and Tools Preparation .........................................56 8.3.3 Personnel Safety ...................................................................................56 8.3.4 Freezer..................................................................................................58 8.3.4 Oven, Autoclaves and Heating Blankets ................................................59 8.3.5 Vacuum.................................................................................................60

8.4 Tooling .........................................................................................................61 8.5 Tools ............................................................................................................61

8.5.2. Air Driven Motors..................................................................................62 8.5.3 Other Equipment ...................................................................................62

8.6 Surface Preparation......................................................................................63 8.7 Damage Removal.........................................................................................63

8.7.1 Removing Paint & Tedlar/Moisture Barrier Film......................................63 8.7.2 Removing Plies .....................................................................................64 8.7.3 Removing Core .....................................................................................66

8.8 Cleaning After Damage Removal..................................................................68 8.9 Water/Moisture Removal After Damage Removal .........................................69 8.10 Replacing Core...........................................................................................70

8.10.1 Replacing Honeycomb Core ................................................................70 8.10.2 Replace with Replacement Honeycomb Core ......................................70

8.11 Replacing Plies...........................................................................................73 8.11.1 Cut plies ..............................................................................................73 8.11.2 Cutting Extra Patches ..........................................................................75 8.11.3 Ply Lay-up ...........................................................................................76

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Aircraft Composite Repair Technology 8.11.4 Ply Compaction ...................................................................................79 8.11.5 Vacuum Debulking ..............................................................................79

8.12 Curing Process ...........................................................................................79 8.12.1 Application of Pressure ........................................................................80 8.12.2 Application of Heat ..............................................................................80 8.12.3 Curing Damaged Honeycomb Core .....................................................82 8.12.4 Curing Damaged Laminates and Damaged Sandwich Structure...........82 8.12.5 Refinishing ..........................................................................................84

CHAPTER 9: GENERAL REPAIR METHODS & PROCEDURES TO COMPOSITE STRUCTURE BY USING WET LAY UP AND COLD BOND METHOD....................86

9.1 Repair In General .........................................................................................86 9.2 Repair ..........................................................................................................88 9.3 Repair Procedures of Various types of defects..............................................90

9.3.1 Resin Injection.......................................................................................90 9.3.2 Sanding and Resin Filling ......................................................................91 9.3.3 Ply Replacement ...................................................................................92 9.3.4 Core Replacement.................................................................................94 9.3.5 Filler or potting compound .....................................................................95

CHAPTER 10: GENERAL REPAIR METHODS & PROCEDURES TO COMPOSITE STRUCTURE BY USING DRY LAY UP AND HOT BOND METHOD ......................99

10.1 Introduction.................................................................................................99 10.2. STRUCTURAL REPAIR MANUAL (SRM) ................................................100 10.3. Determine Damage..................................................................................100 10.4. Water Removal from Damaged Area........................................................101 10.5. Damage Removal....................................................................................102 10.6. Honeycomb Core Plug Fabrication and Cleaning .....................................104 10.7. Core Plug Installation...............................................................................104 10.8 Curing ......................................................................................................105 10.9 Type of Damages .....................................................................................107

10.9.1 Damage on Laminate ........................................................................107 10.9.2 Damage on the Core .........................................................................108

CHAPTER 11: GENERAL REPAIR METHODS & PROCEDURES TO METALLIC BONDING STRUCTURE BY USING DRY LAY UP AND HOT BOND METHOD ...112

11.1. Introduction..............................................................................................112 11.2. Repair Procedure ....................................................................................113 11.3 Repair of a Delamination at the Edge of Aluminum Honeycomb Structure.115 11.4 External Doubler Repair of a Dent.............................................................116 11.5 External Doubler Repair of a Skin Crack ...................................................118 11.6 External Doubler Repair on the Square Edge of a Panel with Corrosion Damage or Delamination ..................................................................................120 11.7 External Doubler Repair of One Skin and the Aluminum Honeycomb Core123

CHAPTER 12: FUTURE REPAIR AND TECHNOLOGY........................................139

12.1 Introduction...............................................................................................139 12.2 Aircraft Structural Health Monitoring (SHM)...............................................139

12.2.1 Comparative Vacuum Sensors (CVM)................................................141 12.2.2 Acoustics Emissions (AE) ..................................................................141 12.2.3 Eddy Current Foil Sensors (ETFS).....................................................141 12.2.4 Fiber Bragg Grating (FBG).................................................................142 12.2.5 Acoustic Ultrasonic (AU) ....................................................................142

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Aircraft Composite Repair Technology 12.3 Thick Laminate Repair..........................................................................143 12.3.1 Bonded repair....................................................................................144 12.3.2 Bolted repair......................................................................................145 12.3.3 Bonded Versus Bolted Repair............................................................145 12.3.4 On-Aircraft Curing Options.................................................................146

12.3 Glass-Aluminum Reinforcement (GLARE) Laminate Repair ......................147 CHAPTER 13: SAFETY IN REPAIR .....................................................................150

13.1 General ....................................................................................................150 13.2 Precautions ..............................................................................................150 13.3 Emergency Action Procedures..................................................................153

CHAPTER 14: QUALITY CONTROL IN REPAIR ..................................................155

13.1 General ....................................................................................................155 13.2 Control Procedure ....................................................................................155 13.3 Use Of Master Parts With Typical Defects.................................................155

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CHAPTER 1: INTRODUCTION

In the past decades there are many achievement have been made in the

composite technology fields either in aerospace or non aerospace application. In the aerospace/aviation industry, the use of composite material have been extensively replaced the traditional aluminum alloy construction starting from the non-structural part to primary structural application.

Composite by definition is a combination of two or more materials that retain

their own identify after they are bonded or cured. These combinations produce new material that has unique and superior characteristics than traditional aluminum alloy material. This unique characteristics is due to it’s anisotropic characteristic in which the strength carrying fibers can be tailored to the optimum direction where the stresses are concerned. Modern composites that are using fiber-resin composites are the most popular application today either in dry lay up or wet lay up application in order fully utilize their capability.

The composite material application in the aircraft have been known in the

late 70-s used in the military aircraft such as in the F-14 fighter aircraft at the horizontal stabilizer. Boron was utilized for it’s superior strength compared to the metallic counterparts. However, due to it’s toxicity and flammability, the material was discontinued and other type of material was introduced such as Kevlar and carbon graphite. Current aircraft such as the Eurofighter, F/A-18, F-22 Raptor have used much of their structural material from those advanced composite material.

In the early eighties, the civilian aircraft started to use advanced composite

with DC-10 and continue by other models. Firstly used on fairings and control surface, now they have been used as a fuselage construction. By using composite material, weight reduction is achieved without sacrificing the strength of the material. In this case, it is translated to better fuel economy, increase of payload, reduce maintenance time etc.

Figure 1: Boeing 777 and Boeing 787 composite construction

(Courtesy of www.boeing.com)

Composite materials have been increasingly used in the aircraft constructions nowadays. Here are few examples of how the transition of commercial

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Aircraft Composite Repair Technology aircraft towards composite material. Boeing 777 uses only 25% of its construction and Boeing 787 uses 80% of it’s structure made from composite as in Figure 1. Airbus Industries have also utilized composite structure on all its fleet. The Airbus A380 is using GLARE for it’s fuselage construction. GLARE is glass aluminum reinforcement in which fiberglass is layered between aluminum skin in which has a fatigue resistant characteristic. This can reduce the problem of crack propagation extending from one layer to another layer and traditional aluminum skin still can be used to replace it when damage occurred.

Since the use of composite has generated several benefit, the actual

problems lies during the maintenance while operating the aircrafts. Composite by nature is not the same as metal. Metal has the isotropic properties in which the material has the same identities in any locations. The stresses are distributed evenly at all directions. On the other hand, the directions of the load carrying members are tailored only to the desired strength carrying capability; therefore, the stresses are not evenly distributed at any location. Furthermore the laminates itself are connected with resins and there are layers of lamiae that made up this laminates.

Delamination is the greatest issue in composite parts or component. The

disbond sometimes is hard to detect and seen. It may occur internally or externally. Stringent inspection requirement need to be done to ensure the hidden damage is detected and repair accordingly. The use of NDI techniques have also been improved to do this task. Issues, decision and solution need to be solved in order to ensure the safety of the aircraft. Various types of NDI techniques are discussed at another chapter for further understanding.

The repair procedures of metallic structures are easy and understand since

the characteristic is isotropic. It means that the forces are equally distributed to all directions. In addition, there are a lot testing and data that shows consistency in terms of the behavior of the material. The data can be retrieved and obtained from America Society of Testing Material (ASTM), Boeing Material Specification (BMS) and others. But bare in mind that some of data is not freely given. Either it can be purchase or kept as a trade secret especially if it is a new material innovated by different process, metallurgical content and competitive. Acceptance test need to be done in order to further check the specifications.

When dealing with composite parts, there are many considerations need to

be explore such as the location, thickness, type of material, fiber orientation and quality. Composite parts have different orientation of fibers that are embedded and hold by resin. The repair procedures of repairing a composite structure can be found in the Structural Repair Manual (SRM) for large aircraft and Service Manual (SM) for small aircraft. The repair procedures stated inside the manuals are the repair that is approved to be used and done by the operators. Since the repair procedure is not standardized between the aircraft manufacturers, therefore the level of allowable or authority of repair is different between the aircraft manufacturers. For instances most typical minor and major repair are allowed to be done by a competent operator unless if the damage or defect is out of specification. Comparatively, the Airbus manufacturer does not allow any major repair don by the operator. They have to get consent of the repair by describing the location, type of defect and other report so that they will examine the damage and further advice on the repair either by giving work instruction or send their repair team to fix the damage.

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Figure 2: Major repair of a helicopter blade.

(Courtesy of www.amtonline.com)

The general idea when talking about minor and major repair is that minor repair does not involve with removing and replacing the fabric plies or honeycomb when compared to major repair as in Figure 2. This is because when the fiber or honeycomb is removed, the stress distribution path is taken away and new path is introduced. The part is also weaken and layers of doubler need to be added. Weight is also added because of addition of weight is introduced.

Therefore this course will hopefully guide the students for gaining the basic

understanding how to deal with composite repair methods and technology. The curriculum will cope the student with the basic understanding of where to get the repair procedure, how to read the manuals, and practice on some of the basic repair.

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Aircraft Composite Repair Technology REFERENCES 1. Mazumdar S.K, Composite Manufacturing: Materials, Product and Process

Engineering, CRC Press, 2002. 2. Dawson D.K., Aerospace Composite: A Design & Manufacturing Guide First

Edition, Gardner Publications, 2008 3. Armstrong K.B and Barret R.T, Care and Repair of Advanced Composites,

Society of Automotive Engineers, 1998. 4. Foreman C., Advanced Composites, Jeppensen Sanderson, 2000. 5. www.amtonline.com

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CHAPTER 2: FITTER & FINISH PROCESS 2.1 Introduction

Fitter and finish is a process to achieve the proper exact shape, size and configuration of the part according to the engineering drawing. It is normally done after the part has been cured and prior to any final processes such as painting (finishes) and assembly. 2.2 Trimming Part

When part has been cured, parts produced will have larger parts in terms of

shape and size. The excess material will be trimmed in order to get the proper shape and dimension of the part as per required by the drawing. Part will be trimmed up to the part-line that is pre-built during manufacturing process as in Figure 3.

Figure 3: The parting line is used to guide the fitting process.

2.3 Smoothing And Edge Filling

The process is done by sealing the edge with an appropriate type of resin. The resin may be in liquid or paste form, and in some cases, it is a mixture with microballon. The correct mixture will be known when the viscosity is high like a paste. For thick laminates, the sealing process is just an application of resin whether in liquid or paste form. For sandwich construction, these processes are done first prior to application of resin.

The honeycomb core is removed using a special drill bits. The amount of

honeycomb to be removed is normally between 3 – 10 mm from the edge. The special tool can be brought from the supplier which cost so must or it can easily fabricated by using an ordinary hardware. A 2 - 3 inches long nail can be used by knocking the tip until flat and bend it 90 from the shank. The desired depth of removal tip can be set from the bend radius to the end of tip. Then the hardware is inserted to a drill gun to cut the honeycomb segment. It is advice to use high speed rotation instead of slow in order to have a clean and nice cut.

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Figure 4: Special modified drill bit or undercut tool.

Any crushed or collapsed cell walls must be removed. Any material of

honeycomb that is left in the cavity should be removed using special tools. The special tool can be made in the shop by modify a nail and flatten the tip and bend it to 900. The bent tip will have approximately 3 to 10 mm or 1/16” of length, see Figure 4. This modified nail will be inserted to the chuck of the drill and act like a drill bit. Be careful when using this bit as the tip is swinging in a rotation. The excess honeycomb on the skin side can be cleared by using scrapper but be careful not to damage the fiber. The edge of honeycomb segment is also trimmed by the same process in order to fill the edges with moisture barrier filler as in Figure 5.

Figure 5: The removed core by undercut tool. 2.4 Application of Resin

After the core is removed, the exposed core needs to be filled in by using either resin mix with microballons, paste, adhesive foam etc. This is to prevent the moisture and water entering the core section as the core is sensitive to water. This will wicked or corrode the core material. Make sure the resin applied protruded from edge of the part due to shrinkage. Mask off the surrounding area of the part such as in Figure 6. Please ensure during the application of the resin mix with microballons, paste adhesive etc., applied an adequate pressure to these materials, so that the void can be pushed out and this will ensure a tight air barrier.

Figure 6: Filling resin at the edge of honeycomb sandwich structure.

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Once the resin, adhesive foam or paste are cured, the part is inspection for any damages and a repair is required when the defects is found. The area masked by the tape will be open and any access of the filler will be trimmed by sanding or mechanical means to retain the called dimensions. 2.5 General Rules and Procedures of Smoothing and Edge Filling When trimming and removing the edges for edge filling a few rules and regulations need to be adhered. Dry composite when sanding will produce a very thin dust that is very dangerous to health and environment. Serious attention need to be followed by the person who is working in composite field in terms of attire, disposable area and environment. Appropriate methods and procedures must be observed as follow:-

• The surrounding area must be free from any contamination and have proper conditions.

• Wear personal protection equipment • Use approved resin. • Make sure the resin is in good condition and not expired. • Clean any trimming process before the application of sealing the edge is

done. • Mask off surrounding area. • Cure the resin according to manufacturer instructions. • Trim and clean part after resin has cured.

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CHAPTER 3: INSERT INSTALLATION 3.1 Insert

The main purpose of using inserts is to giving honeycomb and sandwich panels a method of attachment where it can pick up secondary moderately concentrated loads on sandwich panel where the load must be applied away from the edges. The load can be in the form of shear, tension, compression and torsional stress that can damage the sandwich panel if the inserts is not utilized as in Figure 7. It is necessary to stiffen the core locally to aid in distribution of the load from the insert. This is normally accomplished by filling the area around the insert with potting compound.

Figure 7: Direction of force on the sandwich structure.

(Courtesy of Fairchild Fastener) Potting compound is a low density material which has the following uses: a) Material inserts b) Edge filling The purpose of the using the insert are to a) To prevent crushing by local concentrated loads in honeycomb core. b) To join material in honeycomb constructions c) To secure inserts installed in either honeycomb foam core.

Figure 8: The position of the insert after installation and typical inserts type.

The typical configuration of an inserts comes with a flat surface on the top

and bottom. The difference between the top and bottom is the top usually has potting holes for resin transfer. The top part also is temporarily attached with a hat that is used to prevent the inserts from dropping into the sandwich structure. The hat has two potted holes inline with the top-face inserts such as in Figure 8. The inserts shanks are usually but sometimes is threaded or combined of these two types. The length of the inserts are also varies where by some of them are allowed to protrude out on the opposite skin and some of the inserts are sink partially into the sandwich structure.

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Aircraft Composite Repair Technology Since the honeycomb structure consists of 90% to 99% open space per unit areas there are a few types of inserts that are available in the market nowadays. The different of these inserts are depended to the method of attachment of the intended component to the honeycomb structure. Some of the examples of typical sandwich applications in aircraft industry are at the floor panel, interior walls, galley assemblies, lavatory assemblies, stowage bin and exterior pieces.

There are two commonly methods of attachment of the inserts are in the sandwich structure which are mechanically installed inserts and molded-in (potted) inserts bonded type. 3.2 Installing the Inserts

There are a few types of inserts in the market depends of the product offered by a company. The application of installing the inserts is divided into two types. Mechanically or grommet types uses tools for installation and molded type by bonding.

3.2.1 Mechanically Installed Inserts (Grommet Type)

Permanently installed at subassembly, the fasteners are self-retained through a telescopic press fit that is a function of the sleeve and plug sections. The use of threaded or threaded self locking type permits the attachment of the components without the use of additional lock nuts. Panel preparation requires the following:- - A single diameter thru-hole - Standard drill sizes (comparable to the body diameter) - Access to both sides of the panel.

The most common method of applying the necessary pressure is the use of hand arbor press, a hydraulic squeezer or any pneumatically operated press. To assure proper alignment and to direct the pressure to the head of the fastener, the use of a piloted anvil type tool as illustrated in Figure 9 is suggested.

Figure 9: Mechanically installed inserts

(Courtesy of Fairchild Fastener)

Alignment tool such as these can be manufactured by the tooling facility if exist in your place. An average of 1800 lbs for installation pressure is recommended. Extensive pressure may force the telescopic section to over expand and become loose. Panel facing sheet up to 0.032” will dimple automatically to obtain a flush head condition. Thicker sheets may require the use of the non flush head type fastener. If flushness is required in these thicker facings, pre-dimpling or spot-facing is common

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Aircraft Composite Repair Technology practice in the industry. Fastener cannot be installed by conventional methods (such as field installations), may be installed by hand operated pull up tools. 3.2.2 Molded in (Potted Inserts Bonded)

Permanently installed at subassembly, the fasteners are secured after the resin is cured and prevent it from shearing. There are a few type of the inserts with no threads, all the way threaded inserts, half way threaded inserts etc. Panel preparation Requires the following:- - A single diameter thru-hole - Standard drill sizes (comparable to the body diameter) - Either access to both sides of the panel or one face only - Resin and hardener - Easily manufactured on our own.. Most blind applications for potted in fasteners can use the “pre-pot technique. This involves filling the cavity near full, giving consideration to the displacement factor of an installed fastener. Sufficient potting material must be used to bond securely yet to avoid overflow. Other head types use head tab to position and hold the fastener in a flush, perpendicular position. Slots or holes in the tabs and insert head, allow additional potting material to be injected into the panel cavity such as in Figure 10. Usually there are two holes and if the resin is adequate, one of the holes will be over-flown by the resin mixture injected on the other hole.

Figure 10: Bonding inserts installation

(Courtesy of Fairchild Fastener)

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Aircraft Composite Repair Technology REFERENCES 1. Delron Inserts for Honeycomb & Sandwich Panels, Fairchild Fasteners, 1994 2. Armstrong K.B. and Barret R.T., Original Design Criteria, Care and repair of

Advanced Composites, Society of Automotive Engineers Inc., 1998.

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CHAPTER 4: COMPOSITES DAMAGE REPAIR 4.1 Introduction Repair is basically a process to remove the damage or defect from the part and to make the part serviceable up to the standards that are required. In manufacturing and maintenance point of view, the purpose of repair is as follows: § To reduce the number of unserviceable § To reduce the damaged parts due to the mishandling or improper

manufacturing process. § To reduce cost in manufacturing new parts or buying new parts. § To maintain the parts in good condition. Damage must be present first in order repair can be made to the structure. In

manufacturing, damage may be due to the: § improper manufacturing process § mishandling of the parts, § misassembling of parts and components of aircraft. In maintenance, aircraft composite structure may experience damage due to the: § aircraft operating conditions, § environmental conditions and also, § from mishandling of the parts.

Damage and defect can exist in both laminates and sandwich type composite

structure. To be exact, damage can exist either on (1) fiber, (2) matrix-resin, (3) the core, or (4) combination of those three materials.

The types of defects in manufacturing and in maintenance operation are

somewhat similar; however they do have difference from each other. Despite of that, it is very important to evaluate the damage to determine its type, depth and location. Some defect may be more serious to the performance of the part. Then, this information is used to determine the method of repair.

The flow of sequence for a composite repair is depicted in below chart. The

method used to accomplish the repair depends on the manufacturers or SRM limitations, the extent of the damage and the availability of time and material. Refer Figure 11.

In any repair situation, the person must document the repair done to the aircraft parts or component. This is to ensure the traceability of the job done to that particular parts. Either it comes from manufacturing or maintenance side. Usually in maintenance, the parts of the aircraft will be placed on the part itself. When it is installed in the aircraft and some defects are found, it can be trace back by using this number. So the other parts which come from the same batches will eventually inspected when defect or damage due to manufacturing is detected. This is why an Airworthiness Directives (AD) received by the operator are only affected to the aircraft with certain serial number.

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Figure 11: Repair flow sequence. (Courtesy of CN 235 SRM)

4.2 Types of Composite Repair

It is understandable that any damage found after the aircraft is sold to the

customer can be referred to the SRM. Sometimes the defects cannot be found neither in the Maintenance Manual (MM) or SRM. In this case, the operator of the aircraft has to consult the aircraft manufacturer regarding the repair. Usually the manufacturer will send a team to inspect the damage and the recommended procedures to do the repair. Usually at this stage the aircraft need to be grounded until the repair is finish and it safe to fly. In general, there are two methods of repair curing system which are: 4.2.1 Hot Bond Repair

A repair process similar to cure the dry lamination process, which used additional prepreg materials, utilized vacuum process and polymerized at high temperature using heating blanket, oven or autoclave. The typical repair involved either/and replacement of the laminate or core. The flow of the repair process is depicted as in Figure 12.

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Figure 12: A typical hot bond repair process flow

Hot bond repairing of advanced composite structures is accomplished at 93-

121OC (200-230 OF) – Wet Lay-up, 121 OC (250 OF), or 177 OC (350 OF) as applicable. The SRM provides repair data for each component and specifies where these repairs may be used and the maximum size permitted. Heating blankets are normally used to accomplish hot bond repairs. The use of ovens or autoclaves is optional, provided adequate tooling is available to support the assembly during the curing process. For hot bond repairs, pre-impregnated material, adhesive films and foams, procured patches or wet lay-up materials are used. 4.2.2 Cold Bond Repair

A repair process similar to wet lamination process, which used dry fiber fabric either mat or woven roving and resin matrix material, and cured at room temperature. The resin then is infused into the fabric. Refer Figure 13 for the typical cold bond repair flow. Some time, the manufacturer may call for cure at higher temperature to this type of repair.

Cold-bond repairs using wet lay-up materials are made at room temperature

and cured at room temperature to 66 OC (150 OF). The use of heating blankets, heat lamps or warm air ovens is permitted, in order to accelerate the curing of the resin, up to 66 OC (150 OF). Cold-bond repair may not be all the time be permanent repair, it may also be temporary or definitive repair.

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Figure 13: A typical cold bond repair process flow.

Again, it is very important to follow manufacturer recommendations and/or structural repair manual when it comes to determine the type, extent or depth and location of damage, and also its repair procedures and methods. It is wise to note that old repair method to an advanced composite method will result an unapproved repair. Any method of repair must be specific to that type of damage. Failure to comply could result in an unacceptable and unauthorized repair.

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Aircraft Composite Repair Technology REFERENCES 1. 51-50 Composite Parts Repair, CN 235 Structural Repair Manual, IPTN, 2001 2. 51-70 Repair Typical, Boeing 737-400 Structural Repair Manual, Boeing Inc.

2005 3. Armstrong K.B. and Barret R.T., Structural Repair Manual (SRM) Repair

Method Selection, Society of Automotive Engineers Inc, 1998

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CHAPTER 5: DAMAGE ASSESMENT 5.1 General

The repair methods and damage classification are not standardized yet in aviation industry. Each manufacturer has developed a method of classifying damage with an appropriate repair procedure. Moreover, damage classifications differ in terms of manufacturing and maintenance. However, the repair methods are similar to one and another.

Damage will be assessed to determine the (1) type, (2) extent – size & depth

and (3) the location of damage. Once the type and the location of damage are determined, the extent of damage can be classified in terms of:

§ Absolute

- Not more than 100 cm2. - Minimum spacing of 2 times from the maximum size of defects. - Damage on the first layer. - Not penetrating more than 3 layers.

§ Relative

- Total damage does not exceed 25% from the total area of the part. - Damage does not exceed 25% of the part’s thickness.

Only then, proper evaluation can be made whether the damage can be repair or not. Furthermore, proper damage assessment will determine the proper type, configuration and method for that specific damage.

5.2 Defect & Damage Classifications

Damage classification is very important in determine the proper method of repair. In general, depending on the manufacturer of the aircraft, classification of damage is usually placed in one of three categories: (1) negligible, (2) repairable and (3) non-repairable.

Negligible repair is damage that may be corrected by simple procedures with

no restrictions on flight operations. Repairable damage is damage to skin, bond or core that cannot exist without placing restrictions on the aircraft or part. All permanent repairs must be structural, load carrying repairs (normally restore a minimum of 90% of original strength unless specified) that meet the aerodynamic smoothness requirements. Temporary repair is normally involved in maintenance operations.

A non-repairable composite part is one that is damaged beyond established

repair limits. A composite part that is damaged beyond limits must be rejected and/or replaced, unless a structurally sound repair can be designed by a structural engineer – normally repair by remanufactured method. In manufacturing industry, negligible damage is referred to as acceptable, repairable as correctable, and non-repairable as rejectable.

In order that any damage can be classified within these categories,

manufacturer will specify damage limits that can exist on the composite structure.

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Aircraft Composite Repair Technology These will be provided in damage classification chart by the manufacturer. In maintenance industry, the chart can be found in a structural repair manual.

It is important to note that the damage classification will vary in terms of type

of structure and its applications, the type of reinforcing material and its matrix, and also its cure temperature cycles. Next is some example for a typical damage classification chart for aramid & Hybrid structure completed assemblies in manufacturing industry.

Discrepancy Acceptable Correctable A. Surface scratches Scratches in surface resin

only. Scratches must not cut or damage fibers

None

B. Surface depressions on Bag-side and Tool-side Facings

1. Depressions less than or equal to 0.023 cm are acceptable provided the maximum dimension does not exceed 2.54 cm.

2. Depressions greater than

0.023 cm but less than 0.051 cm are acceptable provided the minimum dimensions is 0.508 cm and the maximum is 2.54 cm.

3. Depressions must not

occur more than once in any 30.48 X 30.48 cm area.

4. The edge of each

depression must be at least 15.24cm from any hole or panel edge. Note 1

5. No fiber damage is

allowed.

None

C. Tool-side resin surface impressions or surface imperfections other than surface resin, starvation, or tape edge

1.Impressions or imperfections (e.g. peel ply splice, TFE tape splice or edge mark-off or tool machining mark) less than 0.0124 cm deep which are transferred from tool surface and eliminated by

Not Applicable

pg 23

Aircraft Composite Repair Technology splice gaps. finishing.

2.No fiber distortion or damage is allowed.

D. Bridging,

Delamination and Voids. Note 2

1. Core edge bridging 2.54/10.16 cm wide x 2.54/5.08 cm long and not more than in each 12 linear inches.

2. Radii bridging – 2.54/20.32

cm wide x 2.54/5.08 cm long and not more in each 12 linear inches.

3. All others – 2.54 cm or less

in any dimensions and not more than one in any 30.48 cm x 30.48 cm area.

1. Core edge bridging 2.54/10.16 cm wide x 2 inches long and only once in 12 linear inches.

2. Radii bridging –

2.54/20.32 cm wide x 5.08 cm long and only once in 12 linear inches.

3. All others – 3.81 x 5.08

cm and not only once in any 30.48 cm x 30.48 cm section.

E. Surface Resin Ridges/ Tedlar Wrinkles

1. Tool Side Facings

2. Bagside

Facings

3. Faying Surfaces

None Resin/Tedlar wrinkles on bagside and non-faying surface with next assembly and/or installation to a maximum height of 0.0508 cm. None

All ridges not containing fibers None All resin ridges/wrinkles not containing fibers.

F. Visual Material Inclusions

None None

G. Resin-Rich Areas None To a maximum depth of 0.0508 cm (0.02 inches)

H. Fabric Wrinkles

None None

I. Honeycomb Discrepancies in Completed Assemblies

1. Splice Gaps

None

None

pg 24

Aircraft Composite Repair Technology 2. Node Bond

Separation § Partial

§ Complete

3. Core Chamfer Flatness (Waviness)

10 % of total nodes ad 34 % of the nodes in any 5.08 in x 5.08 cm area. 1 % of total nodes and two nodes maximum in any 5.08 cm x 5.08 cm area (2 in X 2 in) Core chamfer face must be flat to within +/- 0.127 cm (+/- 0.050 in.)

None 1 percent of total nodes or 7 continuous cells in any 15.24 x 15.24 cm2 (6 x 6 in2 ) None

J. Depressions at the edge of honeycomb core or edge of core members (applies to bag-side only) Note 3

0.0508 cm deep and 10 % of the length of the edge-band or beam member

None

K. Assembly Warpage

Maximum gap of 0.076cm (0.03 in) between assemble attachment points and checking fixture when 4.53 kg (10 lbs.) maximum localized forces are applied at 30.48 cm (12 in.) minimum intervals or equivalent smaller intervals [ e.g. 2.26 kg (5 lbs) for every 15.24 cm (6 in.) interval]

None

L. Faying Surface Flatness

Flat within 0.0254cm (0.010 in.) except for step changes caused by splices and ply terminations as allowed by drawings.

None

M. Ultrasonic Indications Note 4

One indication in 30.48 cm x 30.48 cm (12 in x 12 in) section and not greater than ½ in any direction.

None

N. Frayed, Burred, Scorched or Delaminated Area

0.150 inch from machined edges

None

pg 25

Aircraft Composite Repair Technology near Machining Edges

O. Fiber Breakout from Drilling Operations (Drilling, Countersunk) Note 5

See Section XXX. See Section XXX

P. Scorched and/or charred Tedlar

None 100 % of surface provided there has been no fiber damage

Q. Surface Resin Starvation - On adhesive surfaced parts, non-woven surface taped, laminates, sandwich parts, and procured skins

All, provided non indications of porosity, voids, bridging, delamination and other discrepancy criteria.

Not applicable

R. Edgeband Thickness below drawing tolerance thickness

None For nominal drawing thickness 0.203 cm – NONE For nominal drawing thickness 0.203 maximum correction of 0.0254 cm (0.01 inch)

Note:

1. Not related to honeycomb cell mark off 2. Not applicable to delamination related to machined and drilled edges. 3. This condition refers to sharp depression at the edge of the honeycomb core.

This condition could be caused by ply drop-off or improperly placed filler plies and does not refer to a gradual taper caused by resin bleed into the core.

4. Through transmission (TTU) indications having attenuation level 18 dB over a sound are of the applicable standard.

5. Aramid fuzz is acceptable except where it interferes with fit-up to adjacent structure or surface finish.

Below is the typical damage classification chart for laminated Carbon/Epoxy Structural Skins or Bonded Carbon/Epoxy Skins to Honeycomb Sandwich Structure in Structural Repair Manual of Maintenance Industry.

TABLE 4-1: Typical damage classification for aramid & hybrid structure in manufacturing industry.

pg 26


Repair Area A – Laminated Carbon/Epoxy Structural Skins (Cure at 250 OF)

Type of Damage

Negligible Damage (note 8)

Repairable Damage

Scratches Glass Ply Damage (Note 5)

0.010 to 0.030 inch in depth and less than 3.25 inches in diameter or length (Note 1)

Dents (Note 1) Less than 0.010 inch in depth (Note 6)

Not applicable

Panel Edge Damage Less than 0.125 inch wide by 6.0 inches in length and less than depth of skin. (Note 1)

None

Surface Damage Not defined 1. Less than 1.0 inch diameter and less than 0.085 inch deep.

2. Greater than 1.0 inch

diameter and less than 3.25 inch diameter, and less than 0.035 inch deep.

3. Greater than 1.0 inch

diameter and 0.035 inch deep but less than 3.25 inch diameter and 0.085 inch deep.

Surface Damage and Holes

Not Defined Greater than hole limits set in Fig. XX (XX-XX-XX) but less than 6.0 inches in diameter.

TABLE 4-2: Typical damage classification for carbon/epoxy structural/skin damage classification

pg 27


Repair Area B – Bonded Carbon/Epoxy Skins to Honeycomb Sandwich Structure (Cure at 250 OF)

Type of Damage

Negligible Damage (note 8)

Repairable Damage

Scratches Does not penetrate beyond protective glass ply into Carbon Composite

1. Penetrates one or more carbon plies but not through more than 0.01 inches.

2. Penetrates one or more

carbon plies but through skin and not longer than 3.25 inches.

Dents (Note 1 & 3) Less than 0.01 inch in

depth 0.01 to 0.03 inch in depth and less than 1.0 inch in diameter. (Note 1)

Panel Edge Member Damage

Less than 0.125 inch wide and 6.0 inches in length and less than depth of skin (Note 1)

None

Holes and/or Cracks through One Skins

Not Defined 1. 1.0 inch diameter hole or less (Note 1)

2. 1.0 inch to 3.0 inch diameter

hole (Note 1)

Holes and/or Cracks through Both Skins

Not Defined 1. 1.0 inch diameter hole or less on either side (Note 1)

2. 1.0 inch to 3.0 inch diameter

hole on either side (Note 1)

Skin to Core Voids Less than 0.5 inch diameter in area.

Greater than 0.5 inch but less than 2.50 inches in diameter or no greater than 0.70 inch wide by 4.0 inches long.

Leading or Trailing Edge Damage

Less than 0.25 inch deep (Note 3)

Greater than 0.25 inch deep but less than 0.380 inch beyond 0.008 inch stainless steel leading edge and 3.0 inches in length. (Note 1)

TABLE 4-3: Typical damage classification for carbon/epoxy skin to core damage classification.

pg 28


Repair Area A & B

Laminate to Laminate Voids

Skin Thickness (in.) 0.016 0.016-0.020 0.021-0.032 0.033-0.051 0.052-0.064 0.065

Length (in.) 0 0.25 0.50 0.60 0.84 0.93

Greater than allowable but less than 10 times allowable in length.

NOTE:

1. Any dents causing delamination, breaking and/or creasing of the skin must be considered fracture and must be repaired accordingly.

2. The repair adhesives do not adhere well if are not properly bonded initially. 3. It is permissible to straightened out dents in the 0.008 gauge edge that are

confined to within 0.25 inch of edge. Use the three-ounce hammer and back-up bar. Care must be taken to avoid debonding.

4. Surface damage is defined as cuts, deep scratches, abrasions and dents with broken fibers that do not penetrate the skin.

5. Surface damages such as scratches and abrasions that damage paint and/or protective fiberglass outer ply but do not scratch or abrade the carbon laminate fibers underneath are classified as negligible damage.

6. Dents in skin that are stable and are not accompanied with delamination or broken fibers are classified as negligible damage.

7. Sum of void dimensions in any direction shall not exceed 20 % of maximum dimensions in that direction.

8. There are no restrictions on size, locations or number of negligible repairs 9. Repairable holes in vertical stabilizer box skin are limited to holes that do not

extend into the internal structure before or after clean-up. 10. The table is taken from structural repair manual. It is not to be used while

making a repair

TABLE 4-4: Typical damage classification for laminate to laminate voids.

pg 29

Aircraft Composite Repair Technology REFERENCES

1. 51-50 Composite Parts Repair, CN 235 Structural Repair Manual, IPTN, 2001 2. Armstrong K.B. and Barret R.T., Structural Repair Manual (SRM) Repair

Method Selection, Society of Automotive Engineers Inc, 1998.

pg 30

Aircraft Composite Repair Technology CHAPTER 6: TYPES OF DEFECT & DAMAGE 6.1 Introduction

The types of damage discussed in this section are the typical damage that can

be found especially in the manufacturing and fabrication process. However similar defects also can be found in the maintenance and operation of the aircraft. Any damage of composite materials can be divided into three sections. There are:-

1) Laminate The damage occurs on within the fabric or the matrix failure. The failure is

usually due to bonding strength between the plies. This bonding strength is referred to the matrix. However, a good preparation and care will reduce the tendency of damage on this area.

2) Core Some of the core material is very sensitive to moisture or chemical. The top

and bottom plies are preventing the entry of any FOD to the cavities. However the edge filling will prevent the intrusion of any potentially hazardous materials from entering the internal core area. Fiber breakage and matrix cracking may also cause any liquid to sipping through the core area thus weakening the structure.

3) Interface between laminate and core. Any joining between the core material to the laminate must be joined by

adhesive film. Most type of failure in this region is due to either adhesive failure or cohesive failure shown in Figure 14. Adhesive failure is due to the adhesive strength between the ply near to the core and the core itself. However the latter is due to matrix strength. a This area is joint by adhesive film.

Figure 14: Failure on interface of honeycomb structure

6.1.1 Blisters

Hollows or air pockets caused by gas occlusion. They may appear either in solid laminates or sandwich construction shown in Figure 15. Most of the time is due to the instantaneous rate of heat-up in cure cycles.

pg 31


Figure 15: Blister

6.1.2 Tedlar Wrinkles

Wrinkles appear on the surface of any part protected with (Tedlar) PVF waterproofing films or similar material, but not affecting prepreg material or plies. This is caused by uneven layup and improperly stretched of the bagging materials as shown in Figure 16.

Figure 16: Tedlar wrinkles

6.1.3 Fabric Wrinkles

The result of the defect is protrusions and depressions on the part surface that is caused by bending or overlapping of one or more layers of fabric. The way of measuring this defect is by specifying the maximum height (h) in the case of protruding wrinkles or maximum depth (p) in case of intruding wrinkles such as in Figure 17

. The defect may also be caused by improper preparation of vacuum bag of polymerization (wrinkles on mylar, thermocouple marks, ect) , with no overlapping or bending of fabric itself)

Figure 17: Fabric wrinkles

pg 32

Aircraft Composite Repair Technology 6.1.4 Rich Resin Areas

Areas with excess resin which occur on chamfered or stepped edges, radii and etc. They are visually noticed since the affected area shows a muddier shade than the remainder portion of the part with same thickness such as in Figure 18. The cause of the defect is due to excess resin accumulated on the said location. During vacuum this resin accumulated and cannot be dispersed out, therefore, cured at these location.

Figure 18: Resin rich area

6.1.5 Resin Ridge

Sudden accumulation of resin on the part surface after cured. The defect can also be felt when running on the surface as shown in Figure 19. Sharp peaks or protrusions are another indication of these defects. It is due to uneven tooling surface and too much resin that is accumulated at that path.

Figure 19: Resin ridge

6.1.6 Resin Starved Areas

This can be identified as areas where the reinforcing materials are not uniformly covered with resins such as in Figure 20. They can be located due to their high dryness, roughness and porosity. Defect may be caused by lack of pressure on the affected area during polymerization. Furthermore, improperly spreading up the resin is also the cause due to some resin does not wet the fabric.

pg 33


Figure 20: Resin starved area

6.1.7 Tacky Areas

The defect spots on part where resin has not entirely polymerized as in Figure 21. It may caused either by improper mixing of resin with catalyst, or insufficient time or temperature during polymerization. Tacky area by any means is not acceptable and rejected. The uncured resin has damaged the structural integrity of the part.

Figure 21: Tacky areas

6.1.8 Scratches/ Nicks or Gorges

It is a surface damage appearing on the laminate surface and only affecting the resin of the first coat of plies, or the finishing on the composite structure as seen in Figure 22. Scratches affecting one or more layers of fabric shall be considered as cracks or fractures. There are many causes that contribute to this damage but usually it happen during transportation because of rubbing. Therefore, certain covering is applied during this movement of the particular component.

Figure 22: Scratches

pg 34


6.1.9 Cracks

Cutouts, cracks, or deep scratches going into one or more layers of fabric without reaching opposite surface are the same type of defect as shown in Figure 23. The area close to the cracks usually has a shade clearer than adjacent material. Main cause is impact due to tool drop, drop, etc. If it is small, the repair may not require the removal of the laminate and becoming minor repair. However, if laminate replacement is required, then it is becoming the major repair.

Figure 23: Cracks

6.1.10 Fractures/ Punctures

Similar to cracks where cuts or cracks extending the whole fabric layers making up laminate, which may be caused during laminates de-molding, improper handling, blows and etc. In case of sandwich structure, any other damage involving the core will be considered as a fracture as shown in Figure 24. Any misplaced holes in the laminate area, inserts improperly located in sandwich construction shall be considered as fractures.

Figure 24: Fractures

6.1.11 Delamination

It is the improper bonding between two or more layers of fabrics such as in Figure 25. It is detectable by percussion, particularly when present on the last layers of laminate or when it is a thin one. Can be situated at the edge of part, edge of hole or edge of ply delaminated. Delamination is caused by improper resin/catalyst, improper mixing of weighing of the components, inadequate amount of pressure or heat during the cure cycle and improper cleaning of dirt, grease or foreign materials

pg 35


Figure 25: Delamination

6.1.12 Incorrect Ply Orientation

The defect is found mostly in the manufacturing process after the plies are cured. This is due to imbalance lay-up caused by improper design especially for a very long part such as the wings, control surfaces etc. It is notified by the curling or twisting of the part during the curing process. 6.1.13 Skin-Core delamination/ Disbond

It is the un-bonded area between core and laminate layer shown in Figure 26. This is caused by improper pressure application, contamination or adhesive failure during curing. The defect is detectable with the bare eyes on the laminates made with a few layers.

Figure 26: Skin to core void

6.1.14 Core depression

It is due to local distortion of core without full smashing in the direction of the

cell axis as shown in Figure 27. This happens when a when heavy load is applied perpendicular to the nodes and due to an impact of an object.

pg 36


Figure 27: Core depression

6.1.15 Core Crushing

It is a large distortion or smashing of core cells such as in Figure 28. The crushing may appear in three different manners which are crushing in the direction of cell axis, distortion of face hexagons in the direction normal to center line, and “card castle” type crushing which may occur at chamfered edges, radii, etc.

Figure 28: Core depression

6.1.16 Core Displacement

A defect that is due to movement of the core area either to the front of the actual dimensional line or rear. This is due to the pressure applied during blanket is too high or the material does not secured properly that the part displaced from the actual position. See Figure 29

Figure 29: Core displacement

6.1.17 Core Nodal Delamination

This defect usually appears on sharp bends, where stress concentration is

higher or where core shows failures prior to the lay up as shown in Figure 30. It is

pg 37

Aircraft Composite Repair Technology shown at the faying surface is sheared off that separation of the attachment is seen clearly.

Figure 30: Nodal delamination of core material

6.1.18 Bridging

It is void or voids between two or more plies of the glass cloth reinforcement in a recessed corner due to improper layup techniques or lack of machining between one or several layers of pre-impregnated webbing and core in core chamfered areas. Another factor is due to the minimum radius in which the fabric bends. Please see Figure 31. It is detectable by percussion when bridging is located on the side of laminate or where laminate is made of few layers.

Figure 31: Bridging effect at between core and laminates

6.1.19 Pitting located in Center of Cells

Cavities located in the center of cells appearing on the sandwich panels, tool face, and usually affecting the first coat of pre-preg material as seen in Figure 33. The reason of its appearing is usually associated with excessive depression or sinking of fabric into cells, and insufficient resin flow.

Figure 32: Pitting of resin affecting first layer

6.1.20 Blisters in the center of cells

The interval cavities appearing in the intermediate areas of laminate covering the core in center of cell, these activities being originated by the same causes as pitting. In some instances, they are detectable by transparency. 6.1.21 Telegraphing

The sinking of the fabric layers into the core cells as shown in Figure 33. This defect may appear both and/or either in vacuum bag face or in the tool side face. The

pg 38

Aircraft Composite Repair Technology design of sandwich structure is the most important in order to eliminate this defect that a too big cavities will sink the fabric.

Figure 33: Telegraphing defects

6.1.22 Porosity

A small diameter external porosity which usually appears on insufficient pressure areas on the tool side face. (Do not confuse with resin starved areas). The part shows on the affected areas, a lusterless/ unpolished look as compared with surrounding material. 6.1.23 Foreign Object Inclusions

Incorporation of foreign object matters into laminate, such as fillings, shaving, pieces of cured resin, dust, during the lay-up process. The use of brush during the wet layup is the most contributing factor pertaining to the existence of the brush trapped underneath the fabric. For manufacturing that is using pre-preg material, the isolation and quarantine between the production and another department is the way to control the existence of the defect. See Figure 34

Figure 34: Foreign object inclusion defect

6.1.24 Geometrical Deviations

This type of damage usually occurred due to failure to follow proper procedure or miscommunication in the process sheet form. Detail step need to be written and understand by the responsible personnel. The defect may originated by either pieces of material missing during production, thickness difference or improper fitter and finish.

6.2 Damage Terminology in Maintenance Activity Mainly this area has the most type of defect found. Majority of the defect is caused by the usage of the aircraft itself. Some of the defects found in the manufacturing side can also be found in this area. During operation the aircraft is subjected to various types of load, environment and human mistakes that lead to defect. As the construction of the aircrafts are changing from the traditional metallic structures to advanced composite structures, more type of defects found are different and unique. Within the different reinforcing materials will show various types of defect that must be handled differently.

pg 39

Aircraft Composite Repair Technology 6.2.1 Cosmetic Defects

A cosmetic defect is a defect on the outer surface skin that does not involve any damage on the reinforcing fibers. It may be caused by chipping, scratching or abrasion during handling; does not usually affecting the strength of the part, and usually repaired for esthetic reasons.

On some structural components made of either aramid or carbon/graphite,

their top layer may be of fiberglass. If the damage occurred to the fiberglass, it may be considered negligible or cosmetic defects due to fiberglass is considered as the stack up ply or protective layers.

Figure 35: Abrasion/ cosmetic defect

6.2.2 Impact Damage

Impact damage may occur if the part is struck by foreign object. Either it is due to environment conditions or human error, this type of defect must be taken seriously because it can lead to a very shocking findings. Most of the aircraft primary structure utilized carbon graphite. The color is black and it is hard to see by naked eyes. Assistance from modern non destructive inspection (NDI) techniques is used to check, measure and see the degree of defect caused by this damage. The degree of damage may range from slight to quite severe depends on the velocity of the impact.

Probably the most common cause of impact damage results from careless

handling during transportation, storage or by standing parts on their edge without adequate protection, and also FOD during take-off and landing. Because of the thin face sheets on a sandwich panel, they are susceptible to impact damage. An area which has been subjected to impact damage should be inspected for delamination around the impacted area. Delamination, denting, nicking, chipping, cracking or fracture of the edge or corner can also be caused from improper handling.

Figure 36: High, medium and low energy impact from left to right.

pg 40

Aircraft Composite Repair Technology There are types of impact damage as stated in Figure 36, The most dangerous type of defect is the third type in which a low impact damage. The initial two damages shown physical topography changes on the surfaces. Unlike the former, low impact damage cannot be seen and the residual defect if its not detected can lead to catastrophic effect. Figure 37 shows the actual damage on an aircraft due to impact.

Figure 37: Impact damage. (Courtesy of www.netcomposite.com)

6.2.3 Delamination and Disbond

Delamination is the separation of layers of material in a laminate where as the separation between the laminate and core is called disbond. This defect can occur with no visible indications. To compound the problem, delamination often accompanies other types of damage, such as impact damage, moisture in the fabric or lightning strike. The delamination may also occur during manufacturing, or more often during a repair operation. 6.2.4 Cracks

Cracks can occur in advanced composite structures, just as in metallic ones. Sometimes they can be detected visually, other times they may require more advanced methods of NDI as seen in Figure 38. A crack may just be in the top of paint or matrix layer, and not penetrate into fiber material at all. A crack may also extend into the fiber material and into the core, but appear to be just in the top surface. A thorough inspection should be made to determine the extent of each crack.

pg 41


Figure 38: Crack under radiography inspection.

6.2.5 Hole Damage

Hole damage may occur from impact damage, over-torque fasteners or as a result fastener pull-through. The holes drilled in the wrong location, wrong size, or wrong number of holes can also be classified as hole damage. Such as in Figure 39. Different type of fibers especially Kevlar need to be drilled by using a special type of drill bit known as herringbone drill . The drill is designed so that when it is used on any Kevlar material, it will minimize the fuzziness on the edges.

Unlike carbon-graphite, the use of general drill will make the drill blunt faster and create minutes delamination inside the drilled hole. Spade drill or dagger drill is designed to overcome this problem

Figure 39: Hole damage

6.2.6 Water Ingression Damage

Moisture absorption or trapped water may cause corrosion to metallic

composite sandwich structure. In nonmetallic composite sandwich structure, water can wick the cell and causing the part to be weaker. Apart from that, trapped water can cause delamination due to the temperature variation. For metallic structure, white flakes can be seen as a result of aluminum corrosion and the sandwich structure is weak and compressible. Such as in Figure 40.

pg 42


Figure 40: Moisture absorption damage

6.2.7 Lightning Strike Damage

The damage is due to the lightning strike that causes burn off resins, leaving bare cloth. Other type damage that accompanies this damage is crack, blister, chip, scorch and/or discolored paint such as in Figure 41. This is due to non-conductive path existed in the non-metallic composite. To overcome this problem, a wire mesh is embedded within the laminated plies or a metallic pigment is sprayed on the coating to provide the conductive path for the lightning back to the atmosphere.

Figure 41: Lightning strike damage on composite structure.

6.2.8 Abrasion

Damage that is caused by wearing away of a portion of the surface by either natural (rain, wind etc.), mechanical (misfit, etc.) or manmade (oversanding etc) and it only penetrate the surface finish only such as in Figure 42.

pg 43


Figure 42: Erosion due to hail storm.

6.2.9 Burn Marks

A type of damage that is showing evidence of thermal decomposition or charring through some discoloration, distortion, destruction or conversion of the surface of the plastic, sometimes to a carbonaceous char such as in Figure 43. This damage can be caused by lightning strike, heat shock on part or fire.

Figure 43: Burn marks caused by lightning

(Courtesy of NASA)

6.2.10 Chemical Attack Abrasion The damage is caused by resin matrix by accidental contact with unauthorized use of chemicals. Typical type of chemical spillage is including fuel, hydraulic oil or leak by chemical during the transportation. The spill may caused structural damage to the structure that can lead to a costly repair or non-economical repair that lead the aircraft to be written off from the service.

pg 44

Aircraft Composite Repair Technology REFERENCES 1. http://www.netcomposites.com/education.asp?sequence=67 2. Radtke T.C, Charon A & Vodicka, Hot/ Wet Environmental Degradation of

Honeycomb Sandwich Structure Representative of F/A 18: Fatwise Tension Strength, DSTO Aeronautical and Maritime Research Library, 1999.

3. Armstrong K.B. and Barret R.T., Structural Repair Manual (SRM) Repair Method Selection, Society of Automotive Engineers Inc, 1998.

4. ATA 51-50 Composite Parts Repair, CN 235 Structural Repair Manual, IPTN, 2001

5. http://oea.larc.nasa.gov/PAIS/Concept2Reality/lightning.html

pg 45

Aircraft Composite Repair Technology CHAPTER 7: INSPECTION METHODOLOGY 7.1 Non Destructive Inspection

In manufacturing industry, there are two types of inspection which are nondestructive inspection (NDI) and destructive testing. In maintenance, the inspection is only done through nondestructive testing. In this section, the discussion will be primarily on the NDI subject. Most of time, visual inspection is carried out first since it is the principal method of damage detection. Careful observation during walk-around and servicing inspections will help to assure early detection of manufactured parts or service-incurred defects.

TABLE 7-1. Service inspection of composites

Other types of NDI are normally carried out to determine the extent of the defect that has been visually detected or detect the damage that is not visible to the naked eye. It is also used to detect damage that has no visual indication and for post-repair inspection. TABLE 7-1 shows the type of NDT methods used in maintenance industry to detect service-incurred defects in composite parts.

7.2 Visual Inspection Visual inspection is the principal method of damaged detection. It is often thought as initial inspection which accuracy is unreliable. However, can be base for further inspection which will determine the extent of damage such as location, depth and size of damage. A magnifying glass or microscope, mirror, borescope and flash lights are very useful in this inspection. Refer Figure 44. Magnifying glass or microscope is used to detect a small defect that hard to be seen by naked eyes. Only the surface defect can be detected and seen by using this technique. Flash light provides lighting that the shadow created behind the defects are exposed. Borescope is used to detect damage expecially in an enclose surface such as in the engine or hollow contained area. Fiberoptic inspection is similar to borescope but the shaft can be bend around corners. It is used to view areas deep inside an assembly. Some manufacturer can inspect up to a maximum length of 4 feet.

pg 46


Figure 44: Typical visual inspection tools

Videoscope is an extended ranged of fiberoptic scope which it can provide high

quality images. It can give live feedback and record for inspection records and review. Most modern techniques is using this type of inspection method. All of this method of traces of defect will only spot surface flaw such as crack, fractures, wrinkles, resin rich, starve resin and depression.

A visual inspection checks for surface flaws such as scratches, cracks, fractures, blisters, wrinkles, rich or starve resin, and depressions. Damage that occurs inside the structure such as delamination, bridging and entrapped water, will not be able to be visually inspected and must utilize more sophisticated equipment in which high acquisition cost will acquire. 7.3 Tap Test It is the simplest methods to detect delamination or other internal flaw. However, it is not accurate. Although anybody seems can perform this job but confirmation is done by a person who has approval to this. This can be achieved through certain non destructive inspection program.

Figure 45: Coin tap test and the tap hammer machine

In this method, coin is tapped lightly along a bond line or area suspected of having delamination. Listen for variations in the tapping sound. A sharp solid sound indicates good bond. A dull thud indicates bond separation. However, changes in the thickness of the part, reinforcements, fasteners, and previous repairs may give false indications. Tap hammer is an electronic equipment which utilize sound wave during the tapping. The variation of the sound wave is detected by the hammer. Any delamination or disbond will be indicated at higher value on the lcd screen.

pg 47

Aircraft Composite Repair Technology 7.4 Ultrasonic Inspection The most common and useful inspection in detecting internal damage and delamination in composite parts is ultrasonic. Ultrasonic testing uses a high frequency sound wave as a means of detecting flaws in a part. This is done by beaming a high frequency wave (through a medium such as water and gel) through the part and viewing the echo pattern on an oscilloscope. By examining the variations of a given response, delamination, flaws or other conditions are detected. There are two types of method used in ultrasonic inspection: 7.4.1 Ultrasonic Pulse-Echo Inspection

It is a single search unit containing both transmitting and receiving

transducer such as in Figure 46. An initial pulse activates the transmitting transducer element, which creates sonic energy travels through a Teflon contact tip (and a medium such gel – to reduce signal lost) into the test part.

Figure 46: Ultrasonic pulse-echo inspection equipment

(Courtesy of www.jetsinc.net) A waveform is generated in the test part and is picked up by the receiving

transducer element. Observing any changes in phase or amplitude of the received signal, or time required for the echo to return to the transducer, indicates defects or delamination present. 7.4.2 Through Transmission Ultrasonic Bond Inspection

This method uses two impinging water columns to transmit sound between two yoked-mounted transducers that are positioned on opposite sides of the part. Water is the medium for transmitting the sonic beam as shown in Figure 47.

Ultrasonic waves produced by the sending transducer are transmitted along

the water column, through the inspection part, and continue along the water column on the opposite side to the receiving transducer. Any delamination will cause a reduction in the transmitted sound, producing a greatly reduced signal response on the CRT screen. The output may be plotted on a recording system.

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Figure 47: Ultrasonic through transmission inspection diagram.

7.5 Radiography or X-ray Radiographic inspection is used to detect cracks in the surface of a component and internal cracks that cannot be detected visually as seen in Figure 48. It is done exposing the suspected area with an X-ray in which the film is developed on the other side of the structure. Only a highly qualified personal with approval tin this method are allow to do this inspection with precaution of nobody are allow to enter the site while the job is in progress. One the film is developed, defects is shown contrasted with the un-defected area. Cracks and moisture embedded inside the honeycomb are indicated by the darkened areas on the radiographic film.

Figure 48: A schematic diagram radiogram or X-ray inspection

The extend of the damages are visibly seen in this inspection such as in Figure 49. One of the current development in this method is the output can be seen in a colorful form that can make interpretation much more easy and user-friendly. Other development is the x-ray unit become more portable and easily moving from one site to another.

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Figure 49: Radiography is one of the best method in detecting entrapped water.

7.6 Infrared/ Thermography This method is also known as Thermography. It locates the flaws by temperature/heat variations at the surface of damage part as shown in Figure 50. Heat is applied to the part, and then the temperature gradients are measured using an infrared camera or film.

Figure 50: A schematic diagram of how thermography/infrared inspection works

Thermography requires knowledge of the thermal conductivity of the test specimen and a reference standard for comparison purposes. Thermography techniques use infra-red sensitive camera detectors and film or video display and recording methods. The most common type of Thermography is known as pulse-video Thermography in which a fully integrated systems including flash tube, thermal camera and data processing hardware. It can be applied to a variety of materials systems including composites, sandwich structures, ceramics and metallic systems. 7.7 Laser Shearography

It is a method that using the laser to take an image differences (pre-stress & stress) as shown in Figure 51 and 52. It is design for strain analysis where very sensitive to slight changes in surface strain due to subsurface discontinuities The type of defect it can detect are metal skin to metal doubler disbonds, moisture in composites. It also can find discontinuities in honeycombs and foam laminar composites. The system is differed from the conventional methods such as

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Aircraft Composite Repair Technology ultrasonic, X-ray or eddy current, whereby, the system presentation allows a fast and large area survey inspection. The major advantages are the high lateral resolution and the relative large penetration depth. Laser shearography is also known as optical method. A typical area of application of optical methods is the field of aerospace with components of composite or other lightweight structures.

Figure 51: A schematic diagram of laser shearography

Figure 52: Laser shearography application

7.9 Dye Penetrant Dye penetrant has been used successfully for detecting cracks in metallic surfaces, however, with the advanced composites, their use is still questioned. The reason is that if a dye penetrant is used on the nonmetallic composite structure and is allowed to sit on the surface, the wicking action of the fibers may take in the dye penetrant and hen they would no longer bond to the new material. The entire area which was affected by the dye penetrant would have to be removed before new patches could be applied. This in effect could extend the damage to the size which would make the part non-repairable. Figure 53 is the process of dye penetrant inspection from the application of penetrant until the crack is seen.

Figure 53: Application of petrant, remover and developer upon a cracked surface.

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Aircraft Composite Repair Technology 7.8 Hardness Testing After a repair has cured, a hardness tester, such as Barcol tester in Figure 54, could be used to determine whether the resins have reached their proper strength. A special chart is used to interpret the results for different types of resins and prepregs. Hardness testing does not test the strength of the composite but only the matrix strength.

Figure 54: Hardness tester.

(Courtesy of Qualitest)

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REFERENCE 1. Roach D., Assessing Conventional and Advanced NDI for Composite Aircraft,

High Performance Composites, Gardner Publications, Volume 6, Number 4, July 2008.

2. Armstrong K.B. and Barret R.T., Damage and Repair Assessment, Society of Automotive Engineers Inc, 1998.

3. Foreman C., Assessment and Repair, Advanced Composites, Jeppensen Sanderson Inc., 1990

4. Dawson D.K., Maintenance & Repair, Aerospace Composites A Design & Manufacturing Guide, Gardner Publication, 2008

5. Niu MCY, Quality Assurance, Composite Airframe Structures Practical Design Information and Data, Hong Kong Conmilit Press Ltd, 1992.

6. www.laserndt.com/technology

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Aircraft Composite Repair Technology CHAPTER 8: PREPARATION AND GENERAL REP AIR PROCEDURES The repair procedures and preparation are basically the same as for maintenance and as for manufacturing. It is important to follow all the recommendations and guidelines set forth by the manufacturer in making a repair. Failure to comply will result unapproved and un-airworthy repair. One must know that each type of damage will have its own specific type of repair procedures. All topics below will show the repairs and procedures which are common and typical to many types of advanced composite repair. In actual situations, manufacturer’s repair manual and other approved technical documents such as Engineering drawings, must be consulted regarding such information as operating environment, damage size limits, materials and cure cycle, repair proximity limits and other information pertaining to a specific repair. 8.1 Determine Damage Before repair can be done, a rectification procedure must be follow in order to determine how far is the damage looks. Before checking for the defect the aircraft need to be cleaned and once the aircraft is ready for inspection a visual examination is taken over to know the type, extend and seriousness of damage. Again adherence to the approve documentation is a must in applying the procedures. If for maintenance operation, check in the vicinity of damaged area for entry of water, oil, fuel, dirt or other foreign matter. Once a the location of the defect is detected visually , further check to find the extent of suspected damage with other NDI methods such as coin tapping, ultrasonic or X-ray. Once the damage has been classified, the area is marked off for further action 8.2 Determine Repair Area Configurations

From the Engineering drawings or structural repair manual, identify the parts materials and configuration of the damaged area, such as:

§ Material type, class, style or grade and fabrication process such hot

or wet lay-up, cure cycles and etc. § Number of plies, orientation and stacking sequence § Adhesive and matrix system § Type of core, ribbon direction, core splicing adhesive and potting

compound. § When making 350 OF cure repairs, check for certain condition such

as sealant or fittings which may limit cure temperature to 250 OF.

Identify the proper repair configuration in making a repair by referring to the structural repair manual or manufacturer recommendations. In general, remove damage in circular or oval shapes, and do not use sharp corners. If an irregular shape must be used, then round off each corner to as large a radius as practical. This is to avoid and eliminate any stress concentration in repair and other complication during repair. Refer Figure 55 below.

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Figure 55: Round shape is mostly desired damage removal

configuration compared to sharp edges. Apart from shapes, you should determine the repair area that need to be

covered, how the skins and core should be removed and patched, and also, the cure cycles. However, always follow manufacturer recommendations. It is wise to sketch the damage configuration on a paper before actual repair are performed. 8.3 General Preparation Below are the steps taken and compiled by various types of resources and reference. However when all information is examined carefully, a generic procedure can be concluded. 8.3.1 Material Preparation The materials needed for the repair are identified and gathered together. The manual are the best reference to determine the items needed. The structural repair manual or some other manufacturer publications such as engineering drawings will list the materials you needed for a repair. Be sure all materials such as reinforcing fibers, resins, adhesives, prepregs, core materials and potting compound that are to be prepared and incorporated in the repair must be the same nature as the base material of the composite parts. They must be of the same type, class, style or grade. 8.3.1.1 Pre-impregnated Material Prepreg material is used to build up the plies of damaged area being repaired. Use only the type of material specified on the engineering drawing or SRM. If carbon fiber is specified, do not replace with Kevlar, fiberglass or other materials. Do not use prepreg material that has gone beyond its storage or “out-of-freezer” work life. 8.3.1.2 Pre-cured Patch Pre-cured patch repairs are repairs where the plies are cured and then bonded to the component similar to Figure 55. The bonding temperature will depend on the adhesive used, usually 120 OC (250 OF) or 175 OC (350 OF).

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Figure 56: Pre-cured patch used in composite repair.

8.3.1.3. Dry Fiber & Resin Wet lay-up repairs are made with dry fiber, matt or woven roving, which are manually impregnated with resin by the user. Resin content of the impregnated fabric should be 40 to 50 percent by weight. The ratio is hard to achieve and it depends on the personnel experience. Tape material is usually replaced with two layers of woven roving fabric material. 8.3.1.4 Film Adhesive Adhesives are used in hot-bond repair to assure bonding of the replacement core and of the prepreg repair plies or the procured patches to the replacement core or existing structure. Film adhesive is cut to the proper size, placed in the repair, and cured at the same time as the repair material. Adhesive have a limited shelf life and must be stored in a freezer. 8.3.1.5. Honeycomb Core When damage extends into the honeycomb core, the damaged core must be removed and replaced with new honeycomb core. Honeycomb material is specified in the drawings. The orientation of the honeycomb core is also specified. If you do not have the same core (material, cell shape, size and density), you can substitute with different core. Generally you can use a denser core of the same cell size and shape. Core with larger cell size and the same density will not be as strong as the original core. Check the SRM for the substitution material. Honeycomb core processing requires that core be cut to size and trimmed to fit the repair area. Where honeycomb tapers, the correct angle must be achieved and edges must be chamfered. Cutting and trimming tools must be sharp and provide clean-cut edges without tearing, shredding or crushing the honeycomb. When deep curvature parts are to be repaired, heat forming of the honeycomb is required. Core crushing must be avoided. In most cases, any repair or changes to the structure will require authorization from Material Review Board (MRB) especially if there is a required substitution of materials with cure schedule or fibers different from that incorporated in the original design. Identify and understand all deviations from the original manufacturing materials.

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Aircraft Composite Repair Technology Be sure all materials are in good conditions especially resins, adhesives and prepregs are within their usable life. They also must be store in a proper manner. Use proper resins. Weigh and mix resins properly. Make a test lay-up. In cases where mold is required to accomplish the repair, the mold surface shall be prepared first by cleaning with MEK, acetone or aliphatic naphtha. Parting films or mold releases are then applied to the mold surface. Finally cover the surface until they are ready to be used. Remember that some release agents require baking prior to lay-up and also, do not wipe a mold surface on which a mold release has been applied. 8.3.2. Facilities, Equipment and Tools Preparation 8.3.2.1 Lay-up area condition Every repair step shall be performed in shop areas properly conditioned as to moisture, temperature, and absence of contaminating agents. Proper housekeeping need to be done to ensure cleanliness and safety while in the area, refer to Figure 57

Figure 57: Typical lay-up area condition.

Lay-up areas must be at a temperature between 18 OC to 27 OC, and relative humidity of not more than 70%. The lay-up area should be reasonably free from dust, moisture, oil mist, exhaust fumes, and any other materials that may be detrimental to adhesion. As a suggestion, the lay-up area should not be located in the same room where cutting, sanding and drilling are accomplished. 8.3.3 Personnel Safety This is the most important safety aspect that easy to say but hard to do. It covers the personnel and place the person works. Government agencies such as the Occupational Safety and Health Administration (OSHA) of any countries has safety codes and regulations that the industries have to follow in order to ensure the safety and health to any person working in this environment. These safety issues are also connected to environmental issues due to the fact that many materials used in composite are chemically hazardous as in Figure 58.

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Figure 58: Personnel safety during machining composite material

8.3.3.1 Ventilation Attention should be paid to personnel safety. Good ventilation of the work area is a requirement. Epoxies and cleaning agents may cause discomfort or illness when used in confined area not having adequate ventilation. 8.3.2.2 Skin Protection Most resins are harmful to the skin. Some people are especially sensitive to these materials. For this reason, gloves are recommended for protection. Gloves are required when handling adhesives or prepreg materials since hand creams can have an adverse effect on bonding operations. For wet lay-up repairs, rubber gloves must be worn to provide protection from the resins. 8.3.2.3 Masks and Safety Glasses Dust masks and safety glasses should be provided to personnel performing sanding operations and to others in the immediate vicinity. Vacuum dust extraction equipment is also recommended. See Figure 59.

Figure 59: Personnel safety

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Aircraft Composite Repair Technology 8.3.2.4 Material Safety and Data Sheet (MSDS) The MSDS or CSDS list all the harmful chemicals and their effects. It will also tell you of any potential hazard such as fire or explosion. Read the MSDS before using a material for the first time to make sure you know what protection you need. Normally copy of MSDS will be supplied by the vendor or aircraft manufacturers. 8.3.4 Freezer The equipment is used to store materials that requires low freezing temperature in order to retard the curing of a material such as pre-preg material, adhesive film , adhesive foam etc. These materials have very limited and short shelf life. Any removal of these materials to the room temperature will hasten the curing process and the shelf life of the material becomes shortened as in Figure 60. Filling up a record is crucial in order to monitor the shelf life. Information on the rate of exposure with respect to the balance of the shelf life can be retrieved from the data sheet attached with the shipment.

Figure 60: A typical storage freezer

8.3.3.1 Storage Requirements Various materials, including resins, adhesives and prepreg materials, used for composite repairs have limited shelf life which required refrigerated storing condition. Refrigerated storage capability below -180C (O0F) shall be provided for these materials. Specific information is available from the material suppliers. In general, these materials are stored in freezer at -120C (100F) or below. Allowable storage life will vary from 3 months to 12 months depending on their storage temperature. 8.3.3.2 Prepreg Work Life The work life, or accumulated time out of the freezer, is a critical factor since materials which have been kept at room temperature beyond a specific of time or beyond specified temperature will not cure properly. These exposure limits are also defined by the material manufacturer. They vary considerably, depending on the material and actual exposure temperature. A typical example is at 240 exposure

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Aircraft Composite Repair Technology hours (10 days) at 27 OC (80 OF) for a commonly used graphite epoxy prepreg material. 8.3.3.3 Time Record Since these exposure hours represent the total cumulative time out of the freezer, it is essential that record cards be used to log all out and in times as in Figure 61. Each time the pre-preg material is take out, the life time become shorten and eventually it become expire before the due. Once the material is due, it cannot be used although the material is still look perfect and perfect.

Figure 61: Freezer for time-temperature sensitive materials. 8.3.4 Oven, Autoclaves and Heating Blankets Apart of safety and material storage requirement, the curing system is also a crucial part. This is because the curing system use will determine the cost, quality and production efficiency of a composite repair facility. A big and expensive equipment is not well to justified if it use to cure a small composite panel. Therefore, the part can be send to other repair station that have the facility to cost. 8.3.4.1 Size Oven and autoclaves must have the physical capacity to accept the part to be repaired. Heating blankets must be sufficient size to extend a minimum of 2 inches beyond the edges of repair. 8.3.4.2 Temperature Oven, autoclaves and heating blankets with their power sources must be capable of providing controlled heat-up rates of 1 to 80F/minute and cure temperatures to 355 ± 100F (higher temperature capability may be required for newer type of matrix). The recommended heating blanket capacity is 5 watts/sq. inch.

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Aircraft Composite Repair Technology 8.3.4.3 Pressure The pressure created by a minimum of 22 inches of mercury must be maintained during the repairs. If an autoclave is used, the pressure is 45 ± 5 psi for honeycomb sandwich structure or 80 ± 15 psi for solid laminates. Follow the cure cycle shown in the process specification. See in Figure 62

Figure 62: Heat Blanket, Hot Bonder Machine and Autoclave.

8.3.5 Vacuum Ply lay-up compaction in composite is achieved by exerting external pressure on the lay-up through the use of vacuum bag and vacuum. During production, assemblies are then placed in autoclaves where additional compaction is gained by using pressure of 45 ± 5 psi for honeycomb sandwich structure or 80 ± 15 psi for solid laminates. Repair must also be cured under vacuum pressure in order to compact the repair plies and to bleed off the entrapped air and gasses. A minimum vacuum of 22 inches of mercury must be maintained during the cure cycle in order to effectively compact the repair. This will only produce about 10 psi pressures on the repair, considerably less than the autoclave pressure. See Figure 63

Figure 63: Vacuum gage placed adjacent to the repair area.

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The vacuum can be obtained by using plant vacuum if available, or by converting plant air pressure by means of a venturi (usually provided as part of the control unit). Figure below shows that a gage installed adjacent to the repair provides more reliable reading data than the built-in gage on the control unit which would not note any pressure drop due to bag leakage or hose obstruction. 8.4 Tooling Tooling tools are often needed to support the repair plies or hold the contour of a part during the repair. The tools can be a plaster or plastic casting, wet lay-up fiberglass or carbon fiber epoxy, metal plate, or full part prepreg tool. The materials used will depend on repair requirements, cure temperature, and time required to make the tool. Casting can be made from the plastic resin or a combination of both. When a small area of a part needs to be supported during a repair (such as a hole thru both skins) you can make a splash of an area that has the same contour to support the repair such as in Figure 64. For a tool that will be used more than once, you can make a plastic faced plaster casting, solid plastic casting, wet lay-up fiberglass or carbon fiber epoxy mold or a stiffened prepreg tool. To support edge band repair, you can tape a sheet of metal to the part. For large repairs cure at high temperature, you will have to build a tool that will support the entire part.

Figure 64: Tooling used for repair 8.5 Tools Ensure all equipment and tools are properly identified. They must be capable and in good condition to be used in the repair. Examples of typical tools & equipment are as follows: 8.5.1 Cutting Tools For repair of advanced components, various air powered hand tools are used to cut damaged area. These include routers, router guide and templates, rotary sanders with sanding disks and abrasive paper. See Figure 65.

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Figure 65: Cutting tools used in repairs. 8.5.2. Air Driven Motors Only air driven motors should be used with power tools. Electric motors are not recommended when working on Carbon/Graphite components because carbon is a conductive material. Blowing and drifting dust or loose filaments will be attracted to nearby electrical equipment and connections, shorting them and making them inoperative. 8.5.3 Other Equipment Other equipment includes masking tape, clamps, knives, spatula, and squeegee. These equipment will ensure a proper repair job is done and also to achieve higher quality task. 8.5.3.1 Dust Collector Shop vacuum should always be used to capture dust, loose fibers and other scrap during repair operations. The materials collected are sometimes irritating, toxic and flammable. A good understanding of these materials will ensure a proper disposal during cleaning the collection system. 8.5.3.2 Thermocouple Thermocouple is used to detect and measure high temperature by using a wire that is twisted together. Inspection need to be done to check for damaged at ends of thermocouple bare wires. The ends are either welded or twisted. A twisted thermocouple does not form as reliable junction as a welded thermocouple and the reliability of a twisted thermocouple decreases significantly with repeated use. Thermocouple plugs must be of the same type designated for the thermocouple wire being used. See Figure 66

NOTE: Twisted wire junctions are susceptible to splitting open. Corrosion at non-welded junctions can cause erratic readings or failure. Silver solder type welds often leave unwanted marks on the finished product

NOTE: MRB will also approved the use of facilities, equipment and tools that are needed to make a specific repair.

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Figure 66: Thermocouples

8.6 Surface Preparation Proper surface preparation is the key element to a good composite repair. It is one area that can cause a repair to fail. Therefore, be sure to remove all surface contaminates such as exhaust residue, hydraulic fluids, and other dirt by using appropriate medium such as soap & water, MEK, acetone or aliphatic naphtha. For in-service parts, clean with a mild soap & water. Careful consideration must be taken to avoid any water seeping through the inner part of composite structure. Then allow it to dry before any further cleaning. For specific cleaning at that repair area, wipe surfaces with a clean dry cloth moistened with (Metyl Isobutyl Ketone) MIBK, (Metyl Ethyl Ketone) MEK, acetone or aliphatic naphtha. Allow solvent to evaporate before proceeding with the repair. Never use shop rag or contaminated rags to clean the damaged surface. Never touch cleaned parts with bare hands or hands that had been applied with protective creams. This applies to handling prepreg and adhesive materials. After cleaning the surface, moisture barrier film such as Tedlar and paint must be removed from the damaged area in order to allow adhesion. Leftover paint in the repair area does not allow the resins to properly adhere, which compromises the repair. Once the surface finish has been removed from the damaged area, the repair zone should be masked off to protect from surrounding areas. 8.7 Damage Removal 8.7.1 Removing Paint & Tedlar/Moisture Barrier Film Nonmetallic composite paint and tedlar material can only be removed mechanically. Any composite material is protected from moisture by these material. On the outside the composite material is protected by the finishes and on the internal it is protected by the tedlar film. A manual sanding is easy to control to avoid excessive material removal but a very tedious process and cost more time. Mechanical sanding is much faster but care need to be taken to ensure that undamaged area is not removed accidently. Paint stripper cannot be used to remove the finishing as it can damage the resin chemically thus more damage is made.

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Figure 67: Tedlar removal in composite repair

Paint and Tedlar moisture barrier may be removed using either of the two following methods: 8.7.1.1 Sanding

Remove paint or Tedlar film using no. 240 or finer Scothbrite or no. 150 or finer sandpaper in the masked off area. Sanding can be done by manual or mechanical operation. Sanding operation must not expose fabric fibers in surface ply or loss of structural strength will occur. 8.7.1.2 Blasting-Medium

Vacuum-blast with No. 230 cast steel shot, or sandblast no. 60 soda lime glass beads with air pressure of 60 to 80 psi or silica sand with air pressure of 30 to 40 psi. Apply blasting medium over paint or over repair area until Tedlar is loosened. Then, remove Tedlar by peeling it from the surface.

Do not penetrate Tedlar with blasting media and use blasting method when repair is being done on the aircraft, contamination of aircraft equipment may result. Use of silica sand should be limited wherever possible because it is a potential health hazards. Observe safety requirements. Do not use higher pressure from the recommended pressure. It may cause further delamination on the part. It is very important to wear appropriate PPE during sanding operation. Vacuum cleaner is normally used while performing sanding operation. 8.7.2 Removing Plies Determine the number of plies that have been damaged, and the number of extra repair plies, if any, required in the repair. If none are specified then one extra ply is automatically required. To remove damage plies, sanding is the method that is usually used. Sanding can be done by either manually or mechanically. Manual sanding offers a great control but is tedious and time consuming. Mechanical sanding is faster and easier, but it is also more likely to cause additional damage by sanding away too much material or accidentally sanding off undamaged area. The grit number for the sandpaper and abrasive paper or disk to remove plies is no. 80.

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Figure 68: Sanding composite plies using pneumatic right-angle sander.

One of the best tools for mechanical sanding composites is a small pneumatic right-angle sander as seen in Figure 68. If this is not available, a drill motor equipped with sanding disk can be used. The sanding of composite structures gives off a fine dust that may cause skin irritations, and breathing of this dust may be injurious to your health. There are two ways of removing plies: 8.7.2.1 Scarf sanding

Scarf cutting is used to remove damaged material with a tapered cutout. Dimensions of the scarf are based on the ratio of the total height of the plies to a given length. The scarf should be evenly tapered down to the center of the repair. By looking the glossiness or plies orientation, you can identify the layer transition. A 1:30 ratio means that the thickness of the skin is 1, the taper is 30. See Figure 69Cores may be either straight-cut or tapered at different ratio than the laminates.

Figure 69: Scarf cut

8.7.2.2 Step sanding

Step cutting is done by removing damage in step manner whereby, each successive layer of the fiber and matrix are removed without damaging the underlying layer. Great care must be exercised during this portion of the repair procedure to avoid damaging the fiber surrounding the area being removed.

This type of joint is usually done to radome repair whereby the need to have

radar transmission by having the minimizing the number of extra plies. Usually if extra plies are use, the radar transmission will bouncing back and indicated as error. The step effect is accomplished by sanding away approximately one-half inch of each layer as you taper down to the center of the repair. The sanding is started at the

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Aircraft Composite Repair Technology outermost mark and work down, toward the center, removing one layer at a time. See Figure 70.

Figure 70: Step cut

Initially, aramid will fuzz, and fiberglass & carbon/graphite will produce a fine powder as each layer is being sanded through. Eventually, the materials will show a gloss area for each ply removed. As the fiber/matrix is being sanded, watch for a slight glossing of the work area. The glossing indicates that one layer of material has been removed and the top of the next layer has been exposed.

When glossing effect is seen, be cautious to stop sanding in that area or the next layer of material may be damaged. The layers are very thin and an inexperienced technician may sand through into the next layer. However, with practice, this portion of the repair technique can be mastered.

Another way to detect if one layer has been sanded is to look for a change in

major fiber orientation. This is only possible when the warp direction of each layer has been manufactured in alternating positions. As the top ply is sanded, the next layer will produce weave in a different orientation, signaling that one layer has been removed, exposing the top of the next layer.

Probably the most difficult layer to sand is second to the last layer, especially if over a core structure. Be careful not to sand through the last layer and expose the core during this sanding operation. Do not use excessive pressure on the sander, or excessive speed. 8.7.3 Removing Core

If damage has occurred to the core material of a sandwich structure, the damaged core material must be removed first, prior to step cutting the laminate. The discussion will be based that the top fiber layers may show some damage that penetrates to the core. The bottom sandwich layers are undamaged.

There are two types of core usually used in the sandwich structures which

are the light density core and higher density core. The different between those cores are the size of the honeycomb cell size, the smaller is the denser.

The removing of the light density core can be achieved by using a sharp

pocket knife or a sharp putty knife of a chisel. A vertical cut is made through the core proceeding along the edge of the routed skin. The knife is then slid along the inner skin surface to make a horizontal cut. The core is then removed down to the adhesive bond line. Any traces of honeycomb excess can be cleaned by using

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Aircraft Composite Repair Technology sanding disk or sandpaper. A router can also be used to clean the core material. Figure 71 to Figure 73 shows a typical process of core removal.

Figure 71: Removal of core with a sharpened putty knife for edge removal

It is highly recommended that for high density core, this type of techniques is

used. Outline the area where the core must be removed using a suitable marking device. A router is usually used to remove the damaged core material. The first cut should be done with a pointed router bit to outline the plug in the damaged core. This top laminate layer plug should then be removed with a pair of pliers. It is recommended to have the router be supported on a flat template so that unnecessary damage can be avoided and the core removal will be flat.

Figure 72: Core removal by using templates.

If the surface to be routed is flat, simply set and lock the depth of the cut on

the router base. The depth can be determined by referring to the structural repair manual. It is recommended to make a trial and error first prior to actual routing. The second cut is made by a flush bit to route out the entire honeycomb area. The first may be skipped if the core damage diameter area is small such as less than one inch.

The damaged core is usually trimmed back or routed flush with the skin trim

line. Be careful not to route the surface of inner ply. As precaution, leave about 1/16 inch of the honeycomb trace so that this can sanded down to the opposite layer. The existing core trace should be sanded flush with the surface of the inner without damaging the fibers in the surface of the ply. Lightly sand the surface of the undamaged inner ply to which replacement core will be bonded. If the opposing laminate is damaged by the router or sander, it will too have to be sanded and repair plies added to repair.

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Aircraft Composite Repair Technology If the repair is sloped (such as on a trailing edge) or curved (such as on a

leading edge), it may be necessary to use a wedged template to ensure that the router is held at correct relative depth of cut.

Figure 73: Core removal using wedge template

Sometimes, it is better to route to a partial depth rather than to route through

the opposite layer of laminate. This can normally be done if the core is greater than one inch deep. However, honeycomb core needs stabilization process first before it can be routed. In manufacturing,

Septums are used for honeycomb’s stabilization. They are used in repairs to

hold the adhesive in place to bond the replacement material. Note that for replacement, if repair stability is a concern, use a higher density of honeycomb material. For foam and wood core, the process of removal is basically similar to the removal process of honeycomb core except for equipment of tool usage in removing those cores. 8.8 Cleaning After Damage Removal All repairs must be cleaned after sanding or drilling operations in order to remove all dust, dirt, oils and other residue. The strength of the bond is directly related to the condition of adhesion surface. If the surface is not properly cleaned, the repair may not bond. A vacuum cleaner is routinely used to remove the dust from sanding. Do not used compressed air where it may further delaminate the damage area. Some compressed air may contain oil residue. A solvent wash dampened with dry clean cloth of MIBK, MEK, acetone or other approved cleaners is then used to further clean dirt and oils at the damage area. Always let the part completely dry before beginning the repair. Once the part is cleaned it is important not to touch the surface or any repair materials with bare hands, or the entire cleaning process must be repeated. Wear gloves during the repair process. Prior to the bonding process, some manufacturer may require a water-break test to make sure all oil and grease has been removed from the surface. The water-break test is accomplished by flushing the repair with room temperature water. If the water beads on the surface, the cleaning process must be repeated. Before continuing with bonding or lay-up process, make sure the surface is completely dry.

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Aircraft Composite Repair Technology 8.9 Water/Moisture Removal After Damage Removal Moisture or water trapped within a composites structure can be very dangerous. It may expand when heated which would build up pressure and cause delamination. Entrapped water that freezes may also cause delamination. The water may also act as a plasticizer, reducing the composite structure strength characteristics. If water is not removed before bonding the patches, blisters may form, or the patches will not bond at all. Water or moisture can enter around improperly sealed edges and fastener holes, and any other area that is not properly sealed. The moisture may be just the humidity in the air, or from leaving the part out where rain and snow can be soaked in the edges. See Figure 74. When the part has been damaged, it is very important to bring the part into a dry area to avoid excessive moisture from wicking into the part. If the part has been properly sealed with paint or sealant after manufacturing or repairing, moisture should not be a problem. In general, there are two options in order to remove standing water. The first option is

1. Remove any standing water with wet/dry vacuum. 2. Vacuum bag the surface to pull the water out. The bleeder, a barrier

normally used to absorb excess resin, in the vacuum bag system helps soak up additional water and the heat blankets evaporates any remaining water or moisture. Most popular method and typical procedures are as follows:

Figure 74: Moisture removal by using vacuum bagging method.

pg 70


Procedures:

a) Apply rigid metal screen over exposed area or core. b) Apply a thermocouple to the center of the screen. c) Apply bleeder cloth – may be glass fabric or breather material, and

heating blanket d) When opposite face is accessible, apply screen, thermocouple,

bleeder cloth and heating blanket to both sides. However, it is acceptable to used heat blanket at one side only – remember that additional heating blanket accelerates water removal.

e) When opposite ace is inaccessible, perform the following steps: ~ Apply a layer of glass fabric bleeder cloth over the screen and hold in place with masking tapes. ~ Apply thermocouple to the center of the bleeder cloth. Place a vacuum line on the edge of the bleeder cloth and hold in place with masking tape. ~ Place the heating blanket over the bleeder cloth.

f) Place sealant tapes are on the entire area and seal the area with vacuum bag material.

g) Evacuate the lay-up to a minimum 20 inches of mercury. (Water trap may be necessary to prevent water from entering the vacuum system)

h) Heat the area normally between 1 hour minimum at 150 OF to 170 OF with the rate of temperature rise must not exceed 5 OF per minute.

3. Use heat lamp to dry out composite structures. 4. If entrapped water cannot be removed, remove and repair the affected

area. 8.10 Replacing Core Once the damage has been removed and cleaned, the new core will replace the one that is taken out. Several steps and procedures need to be followed to ensure a good bonding between the new core to the repaired structure. 8.10.1 Replacing Honeycomb Core

First determine the repair options available from the structural repair manual. There are several options in replacing honeycomb core depending on type and size of the damage and also the type of the composite structure.

8.10.2 Replace with Replacement Honeycomb Core 8.10.2.1 Fabrication of Core Plug

For butt splicing & partial splicing, the honeycomb core plug should match

the original core. It should have the same type of material, type, and cell density. As fabric has a warp direction, honeycomb also has a specific way it should be inserted into a repair. The ribbon direction of replacement core must be the same as the core being replaced. It must be pre-fit to allow a fine gap all around the replacement core. There will be a tight fit against the cell walls of surrounding core material with the

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Aircraft Composite Repair Technology foaming adhesive or a mix of resin and microballoons wrapped around it. See Figure 75

Figure 75: Honeycomb core plug fabrication

For crush splicing, the honeycomb core plug should be made one to three

cells (0.38 maximum) larger than the repair cavity. For pre-preg repair material, an excess minimum of 1/16 inch in core thickness should be provided to allow for core settling during cure and for sanding after cure to obtain a smooth, uniform surface. An excess of minimum 1/8 inch above the core plug thickness in foaming adhesive should be allowed to prevent voids due to the settling during the cure cycle.

For wet lamination repair material, an excess minimum of 1/8 inch in core

thickness should be provided to allow for core settling during the bond operation.

8.10.2.2 Cleaning Core Plug

Clean contaminated core plug by dipping (a maximum of four times) in an MIBK, MEK or acetone bath for 60 seconds; or by vapor degreasing, limiting immersion to 30 seconds per cycle for a maximum of 4 cycles. Locally contaminated areas can be washed with acetone or MEK. The core must be completely dry, clean and free of evidence of solvents before installations. 8.10.2.3 Pre-preg Core Replacement

After pre-fitting, remove replacement core and clean bonding surface with

appropriate solvents such as MIBK, MEK or acetone. Then wipe surfaces clean of all solvent and let the repair area completely dry before proceeding to the next process.

Apply one layer of structural adhesive film to the exposed undamaged skin or

septum of the core cavity. For partial splicing, one layer of adhesive film followed by one ply of pre-preg material and another layer of adhesive film is placed on top of the original core.

Then add more core stability to the repair, a layer of structural adhesive foam is wrapped around the replacement core edge. Finally the core plug is inserted into the core cavity and oriented to the proper ribbon direction. The actual visual presentation is shown in Figure 76.

pg 72


Figure 76: Dry lay-up core replacement (Butt & Partial Slice).

8.10.2.3. Wet Lay-Up Core Replacement

After pre-fitting, remove replacement core and clean bonding surface with

appropriate solvents such as MIBK, MEK or acetone. Then wipe surfaces clean of all solvent and let the repair area completely dry before proceeding to the next process. Depending on manufacturer, two layers of fabric impregnated with resin are placed on the bonding surface (exposed core – butt splicing or undamaged core – partial splicing) of the core cavity).

Then apply a uniform coating of resin with microballoons added around the

core plug edges. Adhesive may be used especially if the curing is a hot cure. Insert core plug into the core plug cavity. Orient ribbon direction in the direction of the original core. See in Figure 77

Figure 77: Wet lay-up core replacement (Butt & Partial Slice).

8.10.2.4. Replace with Potting Compound Potting repair is normally done for damage area less than one inch in diameter or nonstructural parts. Potted repair do not give as much as strength to the composite structure as refitting the hole with a new core. It is normally a mixture of resin with microballoon, or any other filler such as flox or chopped fibers. If this method is used, a light coat of the mixture is all that is needed. Remember that

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Aircraft Composite Repair Technology excessive resin will make the part heavier and more susceptible to cracking. See Figure 78.

Figure 78: Potting compound repair

8.10.2.5 Replacing Foam or Wood Core For foam core, the replacement core must match with the original core. The method is similar to the honeycomb core replacement. Potting compound may also be used to repair damaged foam sandwich composite structure. For wood core, the replacement should also match with the original core. It is very important that the grain of the replacement wood core be perpendicular to the skin of composite structure. Again, potting compound may be used to repair the damaged core composite structure. 8.11 Replacing Plies 8.11.1 Cut plies The patches made from the reinforcing material for the repair must carry the stress loads which were originally carried by the fibers that were manufactured into the part. There should be one fabric bonding patch ply of the same thickness and ply orientation for each damaged ply removed. For example, a bonding patch should be cut and laid into a repair at +45, -45, 0 and 90 degrees to replace those plies which have been removed which were +45, -45, 0 and 90 degrees. However, each replacement plies must be 30 mm (1 inch larger than the each damaged plies. See Figure 79.

pg 74


Figure 79: A typical configuration of step cut replacement plies dimension. When there is a fracture or puncture damage on the composite laminate, all the plies are removed and an additional plies or doubler are placed on both sides. This is to ensure the part have the required strength same as the original structure. The figure 80 shows a typical dry layup replacement plies.

Figure 80: Plies replacement on tapered (above) and step (below) cut on laminates structure.

(Courtesy of www.hexcel.com)

This is another typical damage in which it cracks until the core but not puncturing to the other side. The total number of replacement plies must be the same

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Aircraft Composite Repair Technology as the removed plies and with extra ply used as a doubler. The replacement of the core is discussed at later chapter.

Figure 81: Plies replacement on tapered (above) and step (below) cut on sandwich structure.

(Courtesy of www.hexcel.com)

Whenever repairing a composite component, the fiber orientation of the new patches must be in the same direction as the original structure. This is so that any stress which is imposed on the structure will be able to flex through the repair as well as it does throughout the entire structure. If the fiber orientation is not correctly cut and applied during repair operation, the strength of the part is dramatically reduced and may cause failure of the component. The structural repair manual or engineering drawing will supply information on the ply direction of each layer of the part. Therefore, the repair plies, including any extra plies if required, should be cut from the correct type of fabrics or prepregs. The amount and the orientation of the plies must be exact as the original structure. See Figure 81. 8.11.2 Cutting Extra Patches For hot bond repairs, the required top overlap patch ply must be 30 mm (one inch) larger in radius than the largest replacement plies. In most cases when replacing plies over core, an additional ply the size of the damaged area, called a filler ply, is required to minimize surface depression on flush patches. This extra patch is placed on top of the core plug at either 0 or 90 degrees. In some instance, a manufacturer may recommend to have another extra ply on top of the extra ply that was required for over-patch. This will be the nonstructural ply normally will be glass fabric. Another type of extra plies is normally required for a repair that does not to be flush to the skin surface or for a stronger repair requirement.

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Aircraft Composite Repair Technology 8.11.3 Ply Lay-up It is very important that the surface area that will be laminated has been properly clean, completely dry and free from any solvents and contamination before the lamination process. 8.11.3.1 Wet Lay-up

This process is basically similar to the wet lay-up process in making the composite part except for the impregnation of resin into dry fabrics, rather than pouring resin on the part and applies the patch. First, the amount, shape and direction of plies must be determined by using a sheet of plastic and a warp compass. An easy way to cut the replacement patches to the correct size and shape is to proceed as follows – :

1. Lay up the warp compass onto the repair. 2. Orient the zero degree reference marks on the compass with the

zero degree reference of the part. See Figure 82 3. Lay up a compatible clear piece of plastic material on top of the

repair area. 4. Trace the shape of the repair cutout onto the plastic (starting with

the bottom cut). 5. Note the warp orientation on the plastic for the particular layer

being traced.

Figure 82: Warp compass or clock is a tool which can be used to reference the orientation of the warp to the fiber.

6. Remove the plastic from the repair and leave it for a while and free

from any contamination. 7. Place replacement fabric onto another clear piece of plastic

material. Weigh and mix resin properly and work the resin onto the fabric. This is the impregnation process. See Figure 83.

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Figure 83: Manually impregnated the fabric with catalyzed resin.

8. Take the first damaged-traced plastic material and place it over the

replacement fabric, being careful to orient the warp in the correct direction.

9. Then cut the fabric to the correct shape using sharp razor knife to cut along the shape outlined on the plastic. Refer Figure 84.

Figure 84: Once the fabric is impregnated with resin, the marked plastic sheet placed over it, the patches

can be cut out with utility knife.

10. As the patches are cutout, keep plastic stuck onto the patch so that the fiber direction can be easily recognized and so that the patches can be laid up into the repair in the proper sequence.

It is very important to have plastic material is compatible with the resin and

also free from any contamination. Be sure to weigh and mix the resins properly. Follow all manufacturers’ instructions. Exercise the correct safety procedures. The liquid resin should be poured onto the fabric and worked in using a squeegee. Remember that the resins must fully permeate the fabric so that after curing, the resin and the fabric will form a single solid structure. Caution must be exercised when working with the squeegee not to damage the fiber orientation of fray the fabric.

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The fabric/resin mixture should be about 60:40 depending on manufacturer. A rich resin repair is more susceptible to cracking due to the lack of adequate fiber support – brittle. A starved resin repair is weak in those area where sufficient resin does not provide stiffness or because the fibers are not held together and supported as in a completely impregnated repair. Mix only enough resin.

8.11.3.2 Dry Lay-up

In most cases, some repairs may require cure temperatures that are the same as those used during the manufacturing process. In those cases, prepreg materials are often used. They are remarkably easy to use in repair process.

Prepregs can simply be cut from the roll using the pre-traced plastic sheet

over the material. Remember to cut an amount off the roll of the fabric first, and then store the roll back into the freezer as quickly as possible. The fabric that was cut off should be sufficient size to cut out all patches.

In most cases, manufacturers will call for an adhesive to be used over the sanded surface before the patches are applied. A thin layer of adhesive film, or adhesive tape, is used and extends the outermost damage mark with a minimum of 15 mm or one-half inch in radius. Refer Figure 85.

Figure 85: Adhesive Film usage in dry lay-up.

In some cases, the manufacturer will call for multiple adhesive to be placed

on the repair. For example, every two to three plies, adhesive film has to be placed. The adhesive film with the plastic backing is applied to the sanded area with the adhesive side down. After the film has been worked properly, remove the plastic backings. Heat gun may be used to soften the adhesive from film. With light strokes from a squeegee, the plastic should soon separate from the adhesive.

Another form of adhesive used could be catalyzed resins. These are either in

liquid or paste form, which is applied to the sanded step-cut or scarf-cut area using brush or roller. The pre-cut pre-preg patches can be laid into the repair area at this time. Be sure to remove any plastic backing from the patches. If the plastic is allowed to remain, the patches will not bond to the surface, causing un-airworthy repair.

Pre-preg may be squeegee after each patch is applied. There is no danger of

wet resin dripping into the core area and the process will not be messy. A perforated release film, or peel ply may be added at this time and will prevent any shifting of the patches during the vacuum bagging operation.

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Aircraft Composite Repair Technology 8.11.4 Ply Compaction This process is basically done in a manner where the each patches are worked down really well using appropriate squeegee. This action will compact the fiber layers, eliminate any entrapped bubble, relatively shaping the contour of repair area and in wet lay-up, removes any excess resin. Furthermore, proper compaction will hold the repair securely to prevent shifting during curing. Some manufacturers may have certain procedures which require proper compaction. For example, Boeing required that each ply must be compacted with a temporary bay for at least three minutes. This is required in repair procedure for the Boeing model 777. The compacting process on each ply in repair process will reduce repair porosity. 8.11.5 Vacuum Debulking For some repairs, it may be necessary to debulk the repair plies. When you have a thick repair or when you try to form the plies in or around small radius, you may need to debulk the repair. Debulking will soften the repair plies into or over a small radius. Vacuum debulking is not used for each ply, but after 6 to 10 plies had been laid up. To debulk the repair, first work the plies into or over the radius. Next, the vacuum bag was set at 1200F for no more than 30 minutes for the repair to take place. Leave the plies under vacuum until they are back at room temperature. You can use a heat blanket, heat lamp or a heat gun. 8.12 Curing Process The curing process of repair is basically dependent upon the type of matrix or prepreg materials being used in the repair. The choices of matrices or pre-preg used in repair are dictated by the original materials when the structures were manufactured. In another words, in most cases, curing process depends on the previous curing process done on the structures when they were originally manufactured. In some instances, it is permissible to have cold curing when the part was originally undergone hot curing process. This can be seen especially when holes on parts that has been cured under hot curing method are miss-drilled. In certain circumstances, it is allowed to have the repair part to be cured at the temperature of 2500F when originally; the part was cure at 3500F. When this happen, the repaired structure will not gain or restore to its full strength condition. Therefore, most of the time extra plies and/or overlap will be necessary. However, to the best of our interest; in any repair conditions, the method of curing must be followed based on the manufacturer recommendation or instructions. In general, the curing methods in composite repairs, for the fact, are not different from the curing process in composite manufacturing. The exact principles are applied which are application of pressure and also heat. The equipment and facilities are basically similar.

pg 80

Aircraft Composite Repair Technology 8.12.1 Application of Pressure When performing composite repair, pressure should be applied to the surface during the curing operation. Applying pressure assists in the removal of the excess resin that is squeezed out. In addition, sustained pressure compacts the fiber layers together – reduce porosity, removes entrapped air, and maintains the contour of the repair relative to the original part. It also holds the repair securely to prevent shifting. Several methods can be used to apply pressure on the repair such as sand bags, springs and etc. However, the most common and acceptable to the industry is the use of vacuum bagging and autoclave. However, autoclave is normally used in manufacturing activities. 8.12.2 Application of Heat Some repair does not need heat since composite matrix system cure by chemical reaction. Some matrix can cure at room temperature while others require heat to achieve maximum strength. Failure to follow proper curing requirements and improper use of curing equipment can cause unapproved repair. 8.12.2.1 Room Temperature Curing (Cold-Bond Repair) Some type of composite repairs may be cured at room temperature (65-800F) over a time period of 8 to 24 hours depending on the type of resin used. In some cases, room temperature curing can be accelerated by applying low heat (140-1600F). Full cure strength is usually achieved for five to seven days. However, to have a reliable repair, room temperature curing must be done in a controlled environment where humidity and temperature are carefully controlled and monitored. Room temperatures cures are normally used on nonstructural or lightly loaded parts. 8.12.2.2 Heat Curing (Hot Bond Repair) Most advanced composite utilize resins that require high temperatures during curing process to develop full strength. Consequently, the repair parts that use these types of resins must also cure at high heat settings (250-7500F) to restore the original strength. Typical hot curing temperatures for composite materials are 2500F and 3500F. The amount of heat applied must be controlled by monitoring the surface temperature of the repair. Although heat curing may produce a stronger repair, overheating can cause a severe damage. Do not exceed recommended temperature, so as to avoid disintegration or further delamination of the part. It is very important to note that rise of heating and cooling is very important in heat curing to avoid other damage such as blisters as well as trying to get a homogenous composite part. Heat curing can be accomplished using several different methods, described below:

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Aircraft Composite Repair Technology 8.12.2.3. Heat Lamp Most of time, the use of heat lamp is not recommended unless call for. Heat lamp may localize the heat in one spot causing uneven curing. To prevent this problem, we must consult the SRM for the correct distance between the lamp and the part. This is due to the temperature applied is depended on the distance between the lamp to the part. Below is the graph that has to be used in order to determine the temperature by given the distance between the heat source to the part. See Figure 86. It is taken from the SRM and it need to be noted that this is not a replacement of hot bond curing.

Figure 86 : The heat lamp curing graph (Courtesy of Boeing 737-400 Manual)

8.12.2.4 Heat Guns Heat gun without temperature controller or monitor is highly not recommended. Excessive heat can evaporates the resin, leaving dry areas in the part. 8.12.2.5 Oven Curing It offers controlled and uniform heating of all repair surfaces. Some ovens incorporate vacuum ports to provide pressure while curing. Disadvantages to oven curing are that the part must be removed from the aircraft and must be small enough to fit into the oven. 8.12.2.6 Autoclaves It is customarily used in manufacturing of composites, rather than in repair. However, the advantage of autoclave is that it can repair a large structure that cannot be repaired by any other equipment. On the other hand an autoclave is so expensive in terms of initial cost and operation cost. Therefore, only a repair station and component manufacturers have the capability to acquire and offer the repairing service.

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Aircraft Composite Repair Technology 8.12.2.7 Hot Patch Bonding It is the most popular equipment to be used in repair especially in maintenance industry. It utilizes a flexible silicon heating blanket that incorporates a temperature control. Most hot patch bonding machines also incorporate a vacuum pump to apply pressure during the curing process. 8.12.3 Curing Damaged Honeycomb Core This method is only applicable in curing the replacement honeycomb core plug only that is to be bonded into the core cavity without any plies laminated on top of it. In another words, the core plug is cured separately from the repair plies. For hot-bond repair, after the core has been properly inserted into the core cavity, place thermocouples in contact with all adhesive bond lines to monitor cure temperatures. Failure to meet required cure temperatures is cause for rejection. Vacuum bag as IAW manufacturer recommendation and evacuate to a minimum of 22 inches of mercury. The cure cycle is based on resin & repair. 8.12.4 Curing Damaged Laminates and Damaged Sandwich Structure Depending on the requirement of repair, some manufacturer will call for bondable Tedlar material to be placed over the lay-up as required for bagside, using typically 0.5 to 1 inch overlap of the last prepreg ply. Tedlar film may be held in place with small pieces of Mylar tape – pressure tape.

Figure 87: Honeycomb core cure by using vacuum bag. (Courtesy of Boeing 737-400 Manual)

pg 83

Aircraft Composite Repair Technology However, Tedlar material is rarely used if the repair is a wet lay-up repair. If Tedlar material is not replaced, then a proper moisture-proof type of resin must be used in the finishing process. In some cases, peel ply are used instead of bondable Tedlar material. Peel plies may be used as necessary to prepare surface for subsequent bonding or finishing operations. Vacuum bagging the repair accordingly and for hot bond repair, place thermocouples in contact with the edge of the outermost patch, and on all sides of the repair. Additional cure temperature monitoring is necessary in areas of possible heat sinks. Place control thermocouples over and/or under heat blanket to assist in reaching and maintaining steady state cure temperature. Refer Figure 88. The heating blanket must extend a minimum of 2 inches beyond the repair patch edges. In some cases, when using a heating blanket larger than 12 inches on one side, a light gage aluminum caul - pressure plate (0.040 inch maximum) can be used under the heating blanket to minimize localized heating.

Figure 88: Typical thermocouples placing and vacuum bagging if heat blanket is used.

(Courtesy of AC 43.13-1B)

Vacuum bag the repair and evacuate to a minimum of 22 inches of mercury. The number of perforated film, non-perforated parting film, bleeder/breather cloth depends on the repair requirements. Pressure plate must be used when two or more heat blankets are applied to the same repair. If the is a potential heat sink in the area of repair, heat, when allowed, shall be applied to the heat sink area to prevent heat from being drawn away from the repair materials. The cure cycle for curing will be based on the type of resin used and the type of repair. See Figure 89

pg 84


Figure 89: Typical Thermocouple placing with heat blankets.

(Courtesy of Boeing 737-400 Manual)

8.12.5 Refinishing Remove all bagging material and peel ply if present. Check for the condition of the repair. If flush repair and aerodynamic characteristic of the repair is required, then sand the surface appropriately. Then lightly sand the surface and/or edge of the topmost ply around the repair area with 240-grit or finer abrasive to produce a feather edge. Finish surface of repair with one of the following applicable methods:

1. Inside surface if Tedlar was not replaced:

a) Seal repaired surface where Tedlar film was removed with moisture-proof type of resin. Excess resin must be scrapped off before resin gels and cure in accordance with manufacturer recommendation.

b) Make proper surface preparation. c) Touch up the repaired area with proper type of primer and enamel in

accordance with manufacturer instructions. Match color! 2. Outside surface flame-sprayed with aluminum: Prepare surface & apply

aluminum coating by over-spraying in accordance with manufacturers’ instructions.

3. Painted surface: Prepare surface and finishing in accordance with manufacturers’ instructions and/or drawing requirement

REFERENCES 1. www.hexcel.com 2. Fiberglass and Plastic, AC 43.13.1B Acceptable Methods, Techniques and

Practices- Aircraft Inspection, Repair & Alteration, Aviation Supplies & Academics, Inc., 1998

3. Crane D., Non Metallic Aircraft Structures, Airframe, Aviation Supplies & Academics Inc., 1996

pg 85

Aircraft Composite Repair Technology 4. Wood, Composites, and Transparent Plastic Structures, A & P Technician

Airframe Testbook, Jeppensen Sanderson, 2000 5. 737-400 Structural Repair Manual 6. Foreman C, Composite Safety, Advanced Composite, Jeppensen Sanderson,

1990 7. Armstrong K.B. and Barrett R.T., Safety & Environment, Care and Repair of

Advanced CompositeSAE International, 1998

pg 86


CHAPTER 9: GENERAL REPAIR METHODS & PROCEDURES T O COMPOSITE STRUCTURE BY USING WET LAY UP AND COLD BOND METHOD 9.1 Repair In General In manufacturing industry, there is always one type of repair which is permanent repair; however, in maintenance industry, two types of repair exist which are permanent and temporary. Bear in mind that the discussion in this section is mostly covers permanent type repair. In this chapter, the repair steps are mainly for wet lay up method by using cold bond curing system. The use of heat is only limited to heat lamp and oven. However vacuum bagging without the use of heat blanket is also considered in this method. Nevertheless, in any cases, the best way to make a permanent repair on damaged composite structure is to laminate sets of new plies and/or core after damage has been removed, observing the choice of material, overlap dimensions, ply orientation and curing process. The configuration of repair is normally depending on the location of damage occurred on the composite structure. Therefore, basically damage in terms of location can be classified as related to the type of structure. Examples are as follows:

§ Cosmetic Defect: surface defect not affecting plies on laminates and sandwich

§ Laminates: Damage to one side or through the part § Sandwich: Damage to plies, damage to one side or through the part.

In any of these damaged mentioned above, some may be considered as minor repair and some are major repair, except for cosmetic repair which is definitely will be considered as minor repair. In reality, there is no exact definition of major or minor repair, and they can be the same type of damage. However, the interpretation of minor and major can be classified by looking at variety of aspects such as:

§ The severity of damage: affecting certain desired performance such as

strength, aerodynamic, stiffness, dielectric and etc. § Type of structure and its application: primary, secondary or tertiary;

nonstructural or structural § Complexity of repair procedures.

In the event of conflict between any approved technical repair manual or

documents and Engineering drawings, the Engineering drawings shall have precedence. In general, repair shall restore the original strength or must at least 90% of the original strength of the composite part. This section will only show the general procedures and methods in repairing damage on composite structure based on manufacturing point of view. Do remember that the examples are just to expose the idea of how the repair is performed. It

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Aircraft Composite Repair Technology cannot be used as a basis to make an actual repair or certain judgment. For actual repair, it is best to follow manufacturer’s instruction and recommendation. The typical and/or more advanced repair procedures on definite type of structure will be discussed in advanced composite damage repair course. Those types of repair are concentrated more on the maintenance activity. However, the repair methods and procedures for manufacturing and maintenance are very similar except for the configuration of repair. A typical example of the repair flow is shown in Figure 90.

Figure 90: Repair flow

(Courtesy of Boeing 737-400 SRM)

pg 88

Aircraft Composite Repair Technology 9.2 Repair There are a lot of defects that have been discussed in the previous Chapter 5. Some of the defects found in the manufacturing can also be found in the maintenance. Those types of defects can be group as in Figure 91.

DEFECT

Interface SandwichLaminate

Not EffectingFiber

Effecting Fiber

OneSided

ThroughHole

Not Allow

Sanding and Resin Apply

Resin Injection

Scaft Sanding

Step Sanding

Small Large

Under Cut Remove PlyAnd Core

Remove PlyAnd Core

Bonding

Figure 91: Category of defects

From the above diagram, most of the defect on composite parts is occurred on the said laminate. The degree of defects with it corrective actions must be referred to what the manuals or procedural document says. Below are the typical type of defects occurred on the composite section as in Figure 92.

pg 89


Figure 92: Typical type of defects in a group.

Inside the aircraft Structural Repair Manual there are two classifications of repairs inside the aircraft Structural Repair Manual. They are:-

1) Repairs that have been evaluated and analyzed for damage tolerance capability are classified as Category A, B, or C repairs.

2) Repairs that have not been evaluated and analyzed for damage tolerance capability and are classified as Permanent, Interim or Time-Limited Repairs.

In any occasion when doing the inspection and the repair can not be

identified as an interim or time-limited repair, it will become as a permanent repair. Below is the characterization of the different categories of damage tolerant repairs. Damage tolerance is the ability of structure to sustain anticipated loads in the presence of damage, such as fatigue cracks until it is detected through inspection or malfunction, and repaired. The repair categories that can be analyzed by using the damage tolerance are as follows:

1) Category A Repair is a permanent repair for which the inspections

given in the Baseline Zonal Inspection (BZI) are sufficient and no other actions are necessary.

2) Category B Repair is a permanent repair for which supplemental inspections are necessary at the specified threshold and repeat intervals.

pg 90

Aircraft Composite Repair Technology 3) Category C Repair is a time-limited repair where supplemental

inspections are necessary at the specified threshold and repeat intervals followed by a replaced or reworked repair at the specified time limit. This is also known as definitive repair.

The definitions of the different types of repairs that have not been evaluated

and analyzed for damage tolerance are as follows: 1) Permanent repair: A repair where no action is necessary, except the

operator's normal maintenance. 2) Interim repair: A repair that has the necessary structural strength and

could stay on the airplane indefinitely. The repair must be inspected at specified intervals and replaced if deterioration is detected or damage is found.

3) Time limited repair : A repair that has the necessary structural strength but does not have sufficient durability. This repair must be replaced after a specified time, usually given as a number of flight cycles, flight hours or a calendar time.

9.3 Repair Procedures of Various types of defects From Figure 92, the type of repair on a composite structure can be divided into 5 types from 3 categories. Some of the repair can be done by combining one and another. They are:-

1) Resin Injection 2) Sanding and Resin Filling 3) Ply replacement 4) Core replacement 5) Filler or potting compound

9.3.1 Resin Injection There are two types of repair that is required for resin injection method. Location is the most important things to be considered. The method of repair on the edges and within the laminate area is different. These repairs are only applicable when it involved with disbonding between layers. Another type of repair will be considered if the disbonding is followed by a suspected disconnected main fiber The first type of repair is for the edge delamination or disbonding repair. The steps of doing the repair as in Figure 93:-

Figure 93: Delamination on edge


This type of repair is only for small diameter damage. Some only allow if the damage area is within 1 inch radius only. Typical type of defects that can be repaired

pg 91

Aircraft Composite Repair Technology by using this method are blister, delamination, Small size nodal delamination and bridging that does not interfere with any adjacent component.

a) Drill one 2 mm or 0.06 inches diameter hole in each end of defect and check intercommunication between both holes, using clean dry air. Mask off the surrounding holes area.

b) If possible, place the part in vertical position; inject resin through lower hole, using s syringe, or injection gun or other suitable equipment, allowing air to flow out through the upper hole, until the defective area is completely filled with resin. Be careful to prevent air bubble from being caused in the repaired area. As in Figure 94.

c) Cover holes with tape, if possible. d) Cure according to the characteristics of resin used.

Figure 94: Resin injection


9.3.2 Sanding and Resin Filling

The method is only applied to the surface of a part that does not interfere with the stress carrying fibers. The defects only appear on the outer surface of the composite part. Tedlar wrinkles, resin ridge, resin rich, scratches, Pitting on center of cells, porosity and geometric deviation are the type of defects that can be repair by using this technique.

a) Remove the defect by sanding them off using no. 240 or finer Scothbrite

or no. 150 or finer sandpaper in the masked off area or Vacu-blast with No. 230 cast steel shot, or sandblast no. 60 soda lime glass beads with air pressure of 60 to 80 psi or silica sand with air pressure of 30 to 40 psi, and then peel it off. Be careful not to damage the top plies.

b) Clean with filtered low pressure air, free from oil and water, and/or vacuum the sanded area.

c) Make proper cleaning. Dry if required d) Coat the sanded area with water-proofing resin, and cure. e) See Figure 95

pg 92


Figure 95: Repair on dent


9.3.3 Ply Replacement

This type of repair requires the removal of plies or the laminates. Any repair includes the removal of plies are considered major repair since the stress carrying load is disturbed. See Figure 96. A new load path has to be introduced in order to ensure the stress is transferred smoothly when it go to the repair area. Doublers are place to support the repair structure to ensure the load path smoothly transferred the stress. There are two repair situations that requires ply replacement. For small laminate damage it only requires to do sanding without damaging the first ply and resin filling to smoothen the surface. The type of damages involved are tedlar wrinkles, resin rich, resin ridge, scratches, pitting on center of cells and geometrical deviation.

Figure 96: Typical ply replacement of one (top) and double sided (bottom) repairs.

(Courtesy of www.comtekadvanced.com)

The other type of repair is the replacement of the laminated ply. The type of damage that involves are large area blister, fabric wrinkles with overlapping fabrics, crack, fracture, delamination, blister in the center of the cells, porosity, foreign object damage for structural part and geometrical deviation. The steps are as follow

pg 93

Aircraft Composite Repair Technology a) Using a masking tape, encircle a total diameter area to be determined by

the following formula:

9.3.3.1 Step Method Diameter to be rounded = D + 30N (mm) = D + 1N (inches) where: D – Approximate diameter of the damaged area. N – Number of involved layer

Start sanding on A area limit in a stepped pattern, (15 mm or 0.5 inch offset) and following an oval pattern, if possible as in Figure 97. Should the mentioned step pattern not possible to be made because of the geometrical requirements, the indicated offset may be reduced in accordance with manufacturer instructions. In certain case, scarf method is utilized due to flexibility during sanding.

Figure 97: Sanding and overlap requirement


9.3.3.2 Scarf method Damage is scarfed or tapered based on a ratio, such as 15:1. 1 indicates the thickness of damaged to be scarfed and 15 is the diameter of damaged area to be scarfed. However, each tapered layer must have exposed sanded area a minimum of 15 mm or ½ inch.

a) Smoothly sand the top surface that to be patched until it is feather. b) Blow off sanded areas with low pressure filtered air free from oil and

water and/or vacuum the sanded area. c) Clean the sanded surfaces. Dry the part if required.

pg 94

Aircraft Composite Repair Technology d) Trim as many patches as delaminated layers plus any extra plies

that are required, from prepreg or raw fabric material, as the same type and direction as those removed. As in Figure 98.

Figure 98: Taper sanding


d) If material from raw fabric, impregnated the raw fabric. e) Place adhesive film and followed by patches on the proper position.

Additional adhesive film may be placed if required between the patch. Adhesive film may not be required if for wet lay-up.

f) Prepare vacuum bag as per required by manufacturer. g) Cure according to the type of material used. h) Remove vacuum bag i) Rework the excess of resin, and flush the repair if required and if

applicable, finish with water-proofing resin and cure. 9.3.4 Core Replacement

Core placement is also considered as a major repair. Both plies and

core are removed in this process. The removal of the plies is similar to either by stepping or scarfing methods. The best practice is to sand through the core

Figure 99: Core displacement repair

and start trimming from the center to the outer layer. See Figure 99. This is a major repair and it is permanent on the aircraft structure. There are two types of repair which are the plies and core removal with resin filling and laminate replacement. The type of damage involves is crack at mislocated holes. The

pg 95

Aircraft Composite Repair Technology other typical repair is laminate and core removal and replacement. The type of repair involve are fracture with mislocated holes, fracture, crushed core, core displacement, large nodal delamination and bridging. In core replacement, the core is removed and replaced by a new core which is the same size or denser than the original.

a) For sandwich structure, the repair is started by removing the plies as the steps stated in 8.3.3. core

Figure 100: One sided repair on sandwich structure

(Courtesy of Boeing 737-400)

Figure 101: Both sided repair on a sandwich structure


9.3.5 Filler or potting compound

This type of repair is usually done to small damage area and for non structural part. Refer to Figure 102.

pg 96

Aircraft Composite Repair Technology a) Open puncture with a drill or router to remove the ragged edges

and broken fibers. b) Clean out crushed core. If undercut potted repair is required,

undercut the core approximate a minimum of 0.125”. c) If overlap plies is necessary, mark the outline. The plies are

remove if it is required d) Prepare & clean the surface properly. e) Inject potting compound or foam filler or mixture of resin &

microballoons into the cavity. Allow them to cure. f) Trim excess potting compound. If no ply replacement required,

fitter & finish. g) If plies replacement is required, cut repair patches, prepare

surface and patch. h) Apply pressure and cure. Refinish.

Figure 102: Potting compound repair for crack less than 1 inch on sandwich structure.


The summary of typical repair techniques correspond to damage can be seen from the table 9-1. The details of the repair procedure can be retrieved from the SRM.

pg 97


Type of defect

Res

in In

ject

ion

Sand

and

app

ly fi

ller

Sand

with

out

dam

age

1st p

ly a

nd

appl

y w

ater

proo

f re

sin

Step

cut

/ sc

arf c

ut o

n la

min

ate

and

ply

repl

acem

ent

Step

cut

/ sca

rf cu

t on

lam

inat

e an

d co

re

resi

n fil

ling

Step

cut

/ sca

rf cu

t on

lam

inat

e an

d co

re

repl

acem

ent

Not

acc

epta

ble

Acce

ptab

le

Blister √ √ Tedlar Wrinkles √ Resin rich area √ Resin ridge area √ Resin starve area √ Tacky area √ Fabric Wrinkles (no fabric overlapping)

√

Fabric wrinkle with overlapping fabric)

√

Scratches √ Crack √ Crack Mislocated hole √ Fracture √ Fracture mislocated hole √ Delamination √ √ Core depression √ Crushed Core √ Core displacement √ Nodal delamination (small) √ Nodal delamination (large) √ Bridging (not interfere) √ √ Bridging (interfere) √ Pitting on center of cells √ Blister in the center of cells √ Thread Telegraphing (bagside) √ Thread Telegraphing (toolside) √ Porosity √ Foreign Object Damage (structural Part)

√

Foreign Object Damage (non-structural part)

√

Geometric Deviation (short laminates)

√


√


√

Table 9-1: Typical repair of general damage.

pg 98


1. AC 43.13.1B Acceptable Methods, Techniques and Practices- Aircraft

Inspection, Repair & Alteration, Aviation Supplies & Academics, Inc., 1998

2. Crane D., Non Metallic Aircraft Structures, Airframe, Aviation Supplies & Academics Inc., 1996

3. Wood, Composites, and Transparent Plastic Structures, A & P Technician Airframe Textbook, Jeppensen Sanderson, 2000

4. 737-400 Structural Repair Manual 5. Foreman C, Composite Safety, Advanced Composite, Jeppensen

Sanderson, 1990 6. Armstrong K.B. and Barrett R.T., Safety & Environment, Care and Repair

of Advanced Composite, SAE International, 1998

pg 99


CHAPTER 10: GENERAL REPAIR METHODS & PROCEDURES T O COMPOSITE STRUCTURE BY USING DRY LAY UP AND HOT BOND METHOD

10.1 Introduction

In this chapter, it represents the typical repair techniques and procedure for dry lay up material. In this case, the application will involve the use of pre-preg material, adhesive film, adhesive foam and those materials that constitute in this repair. The application of heat is assisted by various type of equipment. In the case of aircraft maintenance, the most widely use equipment is the heat blanket system. It can cover the temperature for high temperature (HT) to the very high temperature( VHT).

Although the dry lay up must undergo this high temperature curing, but some of the wet lay up resin system can undergo this temperature as well. Care must be taken to read the instruction for the curing resin. A low temperature curing resin when place at high temperature will burn it out and if high temperature resin is cured at low curing temperature it might not be cure at all thus does not achieve the strength it requires. Aircraft resin system is mostly epoxy based especially for the structural application and used at the external surface. For internal application the requirement is to use a phenolic type resin in which it has the fire retardant characteristic. Current fiber applications are consisted of fiberglass, Kevlar and carbon graphite. Most of the aircraft structure are mostly use carbon graphite due to its stiffness and strength to weight ratio when compare to aluminum.

Figure 103: Location of graphite/aramid/fiberglass panel areas


pg 100

Aircraft Composite Repair Technology The application of these fibers does not only limit to one type laminate. It can in the form of hybrid which is the fibers are combination of different fabric types to make up the component as seen in Figure 103. SRM is the reference document in order to know the typical material specification used in the aircraft fabrication. This note will only discuss a typical curing system at high temperature (up to 2500F) and the application can be extended to the very high temperature application (up to 3500F). It also goes well with all material either hybrid or single material laminates. All of these materials will undergo the same curing temperature and curing time as instructed by the manufacturer. But when refer to SRM, each material and curing temperature will be explained in different chapter. 10.2. STRUCTURAL REPAIR MANUAL (SRM)

This subject contains repairs to components made from epoxy resin reinforced with several layers of graphite tape or fabric, aramid fabric, or a combination of graphite tape or fabric, aramid fabric, or glass fabric (hybrids). The most common construction is a sandwich of two laminated skins separated by a nonmetallic honeycomb core. Solid laminate is used for small components, honeycomb panel edgebands, some primary and secondary structure, and at fitting locations. Glass fabric is also known as fiberglass cloth. This subject describes repairs made using 2500F (1210C) or 3500F (1770C) cure materials (prepreg layup).

To get an acceptable quality repair, user need to refer to the applicable component which you are going to repair for the repair limits and material of the component before you use these repair instructions. Use the correct materials for the type of component and repair that is made. The materials need to be agreed with the specifications in the SRM. A 2500F (1210C) cure repairs are restricted to specific components as shown in the specific component structure repair subject. 2500F (1210C) cure repairs will not restore the strength or durability of components originally made using 3500F (1770C) cure materials. Their use is restricted as shown in the specific component structure repair subject. Failure to comply will result in an unacceptable and unauthorized repair. 10.3. Determine Damage

When damage is detected, the area need to be clean. Soap and water are used to remove any oil traces, dirt etc. to ensure the defect can be seen clearer. Once the area has been cleaned, the part need to be inspected to determine the degree of damage. Once inspection is finished, the area is masked off for removal and sanding . Chemical paint strippers cannot be used to remove paint before making damage evaluations. It can cause the damage to the adhesive resin system.

Once the coating is removed, then the area is examined visually for extent of

damage. The part was checked in vicinity of damage for entry of water, oil, fuel, dirt or other foreign matter. Water can be detected by radiographic or thermographic methods. A part from this, any indication of delamination was also inspected. Delamination can be detected by instrumental nondestructive inspection (NDI) methods or by tap test. For the tap test, a solid metal disk is used and repair area lightly tapped, but firmly. Void areas will give a dull sound. Solidly bonded areas will give a sharp ring.

pg 101

Aircraft Composite Repair Technology 10.4. Water Removal from Damaged Area

Once the damaged skin plies is removed, the damage area of the honeycomb is exposed. Any indication of standing water is removed by using vacuum or oil-free compressed air. The top surface of the core is gently sanded to remove the traces of adhesive but the removal of adhesive fillets on core is not required.

Then the part is prepared as per Figure 104 by using aluminum screen, thermocouple, bleeder cloth, heat blanket, vacuum bag and sealed. The aluminum screen is placed over exposed core. The thermocouple is placed at the center of the screen. Another thermocouple is placed on the honeycomb core if the opposite face is accessible. When opposite face is accessible, apply thermocouple to the opposite side of honeycomb sandwich panel and also the heat blanket. An additional heat blanket may be used on the near face at the backside. The additional heat blanket accelerates water removal. When opposite face is inaccessible, use of a heat blanket on the near side is required.

When opposite face is inaccessible or when using an additional heat blanket,

perform the following steps: 1) Apply a layer of glass fabric bleeder cloth over the screen and hold in

place with masking tape. 2) Apply a thermocouple to the center of the bleeder cloth. Place a vacuum

line on the edge of the bleeder cloth and hold in place with masking tape. 3) Place the heat blanket over the bleeder cloth.

Figure 104: Water Removal from Honeycomb Sandwich


pg 102

Aircraft Composite Repair Technology Once the arrangement of the plies is satisfied, then, place the extruded

sealing compound around the entire area and seal the area with vacuum bag material. The pressure is evacuated from the layup to a minimum 22 inches of mercury vacuum is required. On the heat blanket machine, the heat is set for 1 hour minimum at 1500F (660C) to 1700F (770C). The rate of temperature rise must not exceed 50F (30C) per minute. Once completed the layup materials are removed and proceeded with repair procedure.

Special care must be taken during this process to the aluminum structure to prevent loss of strength in the aluminum parts. All aluminum alloys, except 2219, must be kept below 2000F (930C). Aluminum 2219 can be kept up to 4000F (2040C). Sealants, paints and primers need to be removed in those areas where the temperature could reach above the maximum allowable temperature. 10.5. Damage Removal

Trim out the damaged lamination to a smooth shape with rounded corners, or a circular or oval shape. Take care not to damage the undamaged plies, core, or surrounding material. Remove only damaged plies, damaged doublers, and damaged fillers. When the core is also damaged, remove the core by trimming to the same outline as the skin. The core area removed should extend at least 0.50 inch further than the visible core damage limits. Take care to avoid cutting into an undamaged skin on the opposite side. In cores greater than 1.0 inch thick, partially remove core (at least 0.5 inch deep) sufficient to clean up damage.

In areas where contamination cannot be removed by cleaning or drying refer to SRM to remove the contaminated structure along with the other damage. When opposite inner skin is also damaged, trim out the damage to a smooth shape as described in SRM. When the core is removed from the inner surface of opposite skin, the core needs to be smoothened down to the adhesive film. The cut area is inspected to ensure that all damage has been removed once it is finish.

The sanding for adhesion or finish removal must not expose or damage filaments in the untapered surface repair area. The sanding must not expose or damage filaments in each ply when step sanding or in the ply bonded to the core. Refer to Figure 105 and 106.

Figure 105: Sanding and Overlap Requirement


pg 103


Figure 106: Sanding and Overlap Requirement

(Courtesy of Boeing 737-400 SRM) To achieve either tapered or stepped sanding removal, firstly paint finish or

tedlar film is removed by sanding in the masked off area using No. 240 or finer Scotch-Brite abrasive, or No. 150 or finer sandpaper. If the damage area has a layer of aluminum foil, cut back the foil enough to allow a new foil layer to have a 1.0 inch minimum overlap beyond the outermost repair

Taper sand or step sand each ply around the cleaned up damage using No. 80 sandpaper. Taper or step sand each ply a minimum of 0.50 inch per ply. An optional procedure for two or three ply laminate face sheets is to fill the cleaned up damage area flush with the original surface using filler plies during the repair layup. The repair plies are then installed directly on the resulting smooth surface of the repair area. Abrade surfaces around repair using No. 240 or finer, Scotch-Brite abrasive.

Taper sand a uniform taper around the cleaned up damage using No. 80 sandpaper for preparation. The taper is to be over a minimum distance of 0.50 inch for each existing ply of the laminate. As an option on sandwich structure, step sanding per par. 2.C.(2)(b)2) may be used instead of taper sanding. Refer to 51-10-01 for locations of areas of critical aerodynamic smoothness. Taper sanding must always be used on solid laminate structures (Fig. 4). For sanding, use a flexible disk sander, a belt sander, a rotating pad sander, or sand by hand. Remove exterior finishes, including enamel finish and conductive coating, from the surfaces around repair using No. 150 or finer sandpaper.

Clean the repair area once the sanding and removal process is finished.

Remove all sanding dust using oil-free compressed air. Keep solvents away from sources of heat, fire, or sparks. Heat, fire, or sparks can cause an explosion. Avoid contact of solvent with skin, eyes, and clothing. Wear eye protection and use mechanical ventilation or respiratory protection when working in a confined space or area. Breathing vapors or allowing solvent to contact skin or eyes is hazardous. Do not immerse parts in trichloroethane solvent or allow standing solvent on part. Damage to part will occur.

pg 104


10.6. Honeycomb Core Plug Fabrication and Cleaning

Refer to specific component structural identification subject to determine type of core called out on engineering drawing. The honeycomb core plug should fit flush with original core and with ribbon direction the same as in original core. Trim the replacement core up to 0.05 of an inch smaller than the cutout. The core is to fit snugly in the cutout after it is wrapped with the foaming adhesive. Trim core plug to full or partial depth of original core. When applicable, depth of core plug should allow for shrinkage during cure and for thickness of extra plies of fabric cloth and adhesive film between core plug and undamaged core or skin. Clean core plug once the cutting process is finished. The core must be completely dry, clean and free of evidence of solvents before installation.€ 10.7. Core Plug Installation.

Most of the materials used in this procedure have limited life and require controlled storage conditions. Refer to the applicable material specifications for the maximum time out of the controlled storage and for the uncontrolled storage conditions. Before opening the adhesive film, foaming adhesive, preimpregnated material wrappers, condition them to room temperature until moisture no longer condenses on the wrapper.

During use, suspend the rolls of adhesive or preimpregnated material horizontally through their axes free from other rolls or objects. The following procedure is based on the core plug installation being cured separately from the repair plies. As an option, the core plug installation and the repair plies may be cured at the same time if the far side skin is accessible to place thermocouples and a heat blanket on it. For partial core replacement, perform the following steps, Cut two plies of adhesive film to fit the core cavity. Cut one piece of, Type 120 glass fabric to fit the core cavity. Remove the separator film from one ply of adhesive film and place the adhesive film into the core cavity. Remove the separator film from the glass fabric and place the glass fabric into the core cavity. Remove the separator film from the second ply of adhesive film and place the adhesive film into the core cavity. For full depth core replacement where damage does not extend through both skins, perform the following steps. Cut one ply of adhesive film to fit the core cavity.

Remove the separator film from the adhesive film and place the adhesive film into the core cavity. If both skins are damaged, apply a caul plate against the exterior surface of the far side skin and tape in place. If the replacement core plug is being cured separately from the repair plies, place thermocouples at all adhesive bond lines. Cut a piece of foaming adhesive to wrap around the replacement core plug.

pg 105


Figure 107: Application of Pressure During Cure


Remove the separator film from the foaming adhesive, wrap the adhesive around the replacement core plug, and install into the core cavity. Orient the ribbon direction to align with the original core ribbon. Put the layup materials and equipment in place per Fig 107. If damage extends through both skins, put layup materials and equipment on both sides of part 10.8 Curing

Evacuate the repair area to a vacuum of 22 inches of mercury minimum. Maintain a vacuum of 22 inches of mercury minimum during the entire cure cycle. Check the vacuum bag for leaks. Cure per Fig. 8. Determination of the temperature must be made using the thermocouples placed at the adhesive bond line. Determine cure temperature from the thermocouple with the lowest reading. Allow the repair area to cool under vacuum until the temperature of the repair area is 1250F (520C) or less. Then remove layup materials and equipment.

Sand repair core plug approximately flush with surrounding material, making

allowance for adhesive film and slight core crush during cure of repair plies. Remove all sanding dust using oil-free compressed air.

pg 106


Figure 108: Repair Cure Cycle

(Courtesy of Boeing 737-400 SRM) A part of using heat blanket system or autoclave for curing, the using of heat

lamp is also needed depend on the requirement from the SRM. The parameter of temperature, pressure and time is shown in Figure 108. When using heat lamp is used, the temperature is only applied up to 1500F (770C). The application of heat is depended on the distance of the heat source to the repaired surface. The SRM requires a 250W heat lamp is used for these application. The relation between the surface temperature on the repaired material and the distance from the heat source to the surface is given in the Figure 109.

Figure 109: Sealer Resin Cure Temperature and Heat Lamp Temperature Curve.


pg 107

Aircraft Composite Repair Technology 10.9 Type of Damages As been discussed in the early chapter, the type of damage on composite structures are divided to laminate, core and interface layer. Any removal of the reinforcing fibers is considered as major repair. The rest of the figures shows the type of repair suitable for this chapter 10.9.1 Damage on Laminate

Figure 110: Repair of damage to one skin of honeycomb panel.


Figure 111: Repair of damage to plies on edge of honeycomb panel


pg 108


Figure 112: Repair of damage greater than 0.50” diameter to solid laminate


Figure 113: Repair of delamination between plies of panel edgeband.


10.9.2 Damage on the Core

Figure 114: Repair of damage to honeycomb panel where access is limited to one side.


pg 109


Figure 115: Repair of damage to both skins and honeycomb core.


Figure 116: Repair of damage to honeycomb panel where access is limited to one side.


pg 110


Figure 117: Repair of damage to honeycomb core on edge of honeycomb panel.


Figure 118: Repair of damage, 0.50 inch diameter or less, to solid laminate.


pg 111

Aircraft Composite Repair Technology REFERENCES 1. AC 43.13.1B Acceptable Methods, TYechniques and Practices- Aircraft

Inspection, Repair & Alteration, Aviation Supplies & Academics, Inc., 1998 2. Crane D., Non Metallic Aircraft Structures, Airframe, Aviation Supplies &

Academics Inc., 1996 3. Wood, Composites, and Transparent Plastic Structures, A & P Technician

Airframe Testbook, Jeppensen Sanderson, 2000 4. 737-400 Structural Repair Manual 5. Foreman C, Composite Safety, Advanced Composite, Jeppensen Sanderson,


Advanced Composite, SAE International, 1998

pg 112

Aircraft Composite Repair Technology CHAPTER 11: GENERAL REPAIR METHODS & PROCEDURES T O METALLIC BONDING STRUCTURE BY USING DRY LAY UP AND HOT BOND METHOD

11.1. Introduction. Metallic bonding is another type of composite structure which compose of metal and metal bonded together to make up a component. This structure has more stiffness and lighter compared to traditional metallic plate. When metals are bonded, an increment of fatigue life is much more attainable thus can make the structure live longer. Latest application of metallic bonding have been introduced in the fuselage construction by combining metallic and fiberglass or known as GLARE or glass-aluminum reinforcement. GLARE has been introduced and used by modern Airbus fleet such as the A340 and A380 aircrafts see Figure 119. This chapter will only discuss a small portion of the procedures found in the actual Structural Repair Manual or SRM. A more detail and thorough procedures of each type of defects can be found in the said documents. Some of the repair technique may be generic to a typical aircraft. As the technology of the aircraft structure and advanced composite becoming more advance, there will be new techniques and procedures implemented in the future.

Figure 119: Airbus A 380 under construction (Courtesy of Airbus Industries)

Different manufacturers have different recommendation and restriction in giving the aircraft operator to do the repair. Airbus aircrafts have been extensively used composite structure on their aircraft and they are more conservative if other than them making the repair especially on the primary and secondary repair. Only a non-structural repairs are allowed to be done and the information on the procedures can be found inside the SRM. However, for the major structure, they have to contact the manufacturer and get consent on the repair procedure.

pg 113

Aircraft Composite Repair Technology Ironically, Boeing have tabulated a complete repair techniques inside the manual with it’s specification. Other than what is specify inside the manual, then the operator will contact them and get advice on the repair.

The metal bond procedures stated are applicable for repairs that are

necessary to get back to the full strength of the initial aircraft structure provided if all recommendation from the manual is followed. The discussed repair procedure are for limited size repairs only. Larger repair sizes may be approved by specific component repair sections in this SRM, other official aircraft manufacturer approvals, or local regulatory authorization.

In the case of Boeing, for these larger repairs, they recommends to use the

specified Boeing manufacturing processes that are discussed in this note, in which, was derived from the SRM. Approved documented processes that have been demonstrated to give equivalent results as the aircraft manufacturing processes are also acceptable. Failure to follow or adhere to the procedures, can result a decrease in the strength and durability of large size repairs.

The curing temperature is applicable only to a subject with curing

temperature of 2500F (1210C) and the parts are made from aluminum sheets separated by aluminum honeycomb core. The sheets are bonded on the two sides of the honeycomb core as face skins.

Detail information pertaining to the damage structure must consult the criteria

stated below: (1) The identification of the materials. (2) The allowable damage data. (3) The references to this subject. (4) The types of repairs that are permitted. (5) The repair size limits and other information that is not given in this

subject. (6) The engineering drawing references.

Damage is specified as any visible change to the surface of a part that is

caused by deterioration, delamination, erosion, dents, gouges, scratches, punctures, and holes. Some damage found in the non metallic composite structure can also be found in the metallic composite structure. The actual repair procedures have to be followed from the SRM at the specific topics. 11.2. Repair Procedure

SRM is needed to proceed the repair and SRM 51-00-06 is used to seek information on the definitions of terms. Make sure that you do the repairs in a clean location. It is best to a controlled temperature and humidity so that the effect of moisture in the atmosphere can be reduced. Make sure that the air in the area has no oil, mist, exhaust fumes, gasses, soot, rain, dust, or other unwanted materials.

If the place do not have the facility in order to controlled the outcome of the

repair part, a tent can be used to seal the area to prevent outside contamination. The resin systems and potting compounds need to be kept in sealed containers at a temperature of 40 to 800F (4 to 270C). The adhesive film, adhesive foam and pre-preg materials need to be keep at -120C to retard the curing time. While storing the component some information need to be identified by the label on the container.

pg 114

Aircraft Composite Repair Technology The label should have the info on:- (a) BMS Specification (b) Type (c) Class (d) Supplier name (e) Batch number (f) Date of preparation (g) Due date Figure 120 shows a typical flow of repairing non metallic composite structure.

Some of the steps are eliminated depends on the type of damage. Most type of repair procedures are basically the same as per discussed in the earlier chapter. There are three types of damage location on a composite part. It is either on the laminate, core and interface. The discussion on the repair procedure will cover only for a small type of damage repair. Further information on other type of repair procedure may need to refer the SRM

Figure 120: Flow Chart for the Repair Steps


pg 115


11.3 Repair of a Delamination at the Edge of Aluminum Honeycomb Structure A delamination of a face-sheet can be repaired if the damage is not be deeper than 0.5 inch (12.5 mm) into the panel edge or longer than 30percent or the length of the edge. The delamination must also not go into the honeycomb core. If the type of damage is more than what is permitted as stated above, then you must use the damaged face-sheet procedures to repair the panel. The limits of the damage as retrieved from the general repair section.

Isolate the damage area and remove the finish in the damaged area only. The

damage area needed to be cleaned with a soft cloth moist with cleaning solvent. All of the contaminant requires to be removed, in and around the damaged area. Water is removed from the damage area and these steps need to be followed.

a. The area must be fully dry before you continue on with this repair procedure.

b. If necessary, you can dry the area at a faster rate with an external heat source, but the heat source must be limited to a maximum temperature of 1500F (660C)

Once nicks, scratches, gouges, burrs and sharp edges from the repair area

are removed, a two part paste adhesive system is used to repair the delaminated area. The adhesive components are mixed together by following the recommended instructions. Then, the adhesive material is placed between the delaminated plies of the aluminum. Make sure that there is no air caught in-between the adhesive material and the repair area.

A clamp is used to hold the skins together as shown in Figure 121. Use a

light clamping pressure to hold the skins together. Wood, metal, or phenolic wedges are used to apply the clamping pressure equally, use As an alternative, you can use wood, metal, or phenolic blocks between the skin and the clamps. Any unwanted adhesive material that can appear from the edge of the delaminated area must be removed.

Figure 121: Figure Repair of a Delamination at the Edge of Aluminum Honeycomb Structure


pg 116

Aircraft Composite Repair Technology Depends on the type resin, the curing temperature can be varies and care

must be taken to ensure the cured resin at the particular condition produced resin that is structurally stronger. Once the parts are cured then post-repair inspection is continue to ensure the repaired part are following the specification and compliant to the regulation. Finally the repair area is seal and apply the surface finish. 11.4 External Doubler Repair of a Dent

The limit of damage found on the component must be cross check with the General Repair stated in the SRM. An inspection on the damaged area need to be done by determining the skin cracks or delamination. This type of repair cannot be done for skin crack r delamination. Another repair method will be used for the said damage.

Isolate the damage area and remove the finish in the damaged area only. The damage area needs to be cleaned with a soft cloth moist with cleaning solvent. All of the contaminant requires to be removed, in and around the damaged area. The area must be fully dry before you continue on with this repair procedure.

In doing this repair, several measures need to be taken such as the

application of clad aluminum 7075 as a doubler is not permitted in the bonded metal repairs. It can causes corrosion at the bond interface of the repair. On the other hand, clad aluminum 7075 repair parts that are clad on one side can only be used. Make sure that the clad side of the repair part is not bonded to the bond interface of the repair area.

Figure 122: External Doubler Repair of a Dent

(Courtesy of Boeing 737-400 SRM) By referring to Figure 122, proceed with the external repair doubler. The

doubler size is referred to determine the necessary dimension and the thickness of the external doubler.


are removed, the damage area need to be cleaned. Water break test is used to to examine the repair area for free surface. Anodize, or as an alternative, apply a chemical conversion coating to the non- primed surface of the repair area and the aluminum repair parts as given in the General Repair section.

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The step is used only for bare metal surfaces. Remove all of the primer and paint before anodize, or apply a chemical conversion coating to the repair surfaces. The Corrosion Inhibiting Adhesive Primer (CIAP) is then applied to the repair doubler(s), and the repair area.

Figure 123: Installation of the Doubler (Courtesy of Boeing 737-400 SRM)

The dented area is filled up with approved filler material and cured according

to the instruction manuals. Once the filler is cured, then the surface need to be sanded to flush the material with the outer surface of the initial skin. The surface is cleaned before installing the external doubler (s) on the repair area. Adhesive film is applied to the external repair doubler(s) as shown in the Figure 123.

A polyester tape is use to prevent movement of the doubler during curing. Do

not seal the edges of the panel as it can prevent the flow of the adhesive material and the removal of air during the curing cycle. The polyester tape need to be applied around the edges of the doubler to prevent adhesive flash during the curing cycle.

Seal the part by using vacuum bag with the heat blanket application as per

Figure . Once everything in place, the vacuum is applied and examines the vacuum bag for leaks. A vacuum bag that has a leak can cause porosity in the adhesive and a bond failure of the repair. A vacuum to a minimum of 22 inches (560 mm) of Hg (mercury) is applied Then, the vacuum source is removed and the vacuum gage is monitored. After 5 minutes, the total difference in the vacuum must be less than 5 inches (125mm) of mercury. Apply and keep a vacuum to a minimum of 22 inches (560 mm) of Hg (mercury) in the vacuum bag during the cure cycle. Refer to Figure 124.

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Figure 124: Application of Pressure During the Cure for Flat Panels


Use clamps, weight, sand bags, or hydraulic jacks to apply pressure. The pressure must be applied equally to all the areas. This pressure must not be more than 35 to 40 psi (241 to 276 kPa). Use also cast ceramic, plaster, or plastic tools when you apply pressure to the outer skin on a contoured or rounded panel.

Apply the temperature with the heat blankets, heat lamps, or hot air, as necessary for curing. Monitor the temperature with the thermocouples. Make sure that the temperature does not increase at a rate more than 50F (30C) for each minute. Do not increase the temperature of the aluminum part above 2600F (1270C). If you do not obey, damage to the heat treat of the aluminum can occur.

Refer to instruction manual for the correct cure time and cure temperature of the adhesive material that you use on the repair. Do not increase the temperature in areas that are non-pressurized, or not under vacuum, to more than 2000F (930C). This will help to prevent delamination in an area that is not part of the repair.

The heat equipment especially the heat blankets must have a minimum heat capacity of 5 watts per square inch (600 sq. mm) of area. Heat blankets must be a minimum of 1 inch (25 mm) larger than the repair area, but they are not recommended to be more than 2 inches (50 mm) larger. Warm air or heat lamp may be used to help the heat blankets to cure the repair area.

Once the parts are cured then post-repair inspection is continue to ensure

the repaired part are following the specification and compliant to the regulation. Finally the repair area is sealed and the surface finish is applied. 11.5 External Doubler Repair of a Skin Crack

This repair is merely for non non-structural repair. Find the limit of damage

from the General Repair stated in the SRM. The finishes need to be removed before repair can be carried on. The damaged area needs to be cleaned with a soft cloth moist with cleaning solvent. All of the contaminant requires to be removed, in and

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Aircraft Composite Repair Technology around the damaged area. The area must be fully dry before you continue on with this repair procedure.

Drill a hole in the aluminum skin that is 0.25 inch (6mm) in diameter at each

of the outer most boundaries of the crack. It acts as a crack stopper to prevent the crack from elongate. Dye penetrant is not allowed to find the end of the crack. This is because the dye penetrant can sip within the core and damage the core internally.

Figure 125: External Doubler Repair of a Skin Crack

(Courtesy of Boeing 737-400 SRM) In doing this repair several measures need to be taken such as the


By referring to Figure 125, continue with the external repair doubler. The

doubler size is referred to determine the necessary dimension and the thickness of the external doubler.


are removed, the damage areas need to be cleaned. Water break test is used to to examine the repair area for free surface. Anodize, or as an alternative, apply a chemical conversion coating to the non- primed surface of the repair area and the aluminum repair parts as given in the General Repair section.

The step is used only for bare metal surfaces. Remove all of the primer and paint before anodize, or apply a chemical conversion coating to the repair surfaces. The Corrosion Inhibiting Adhesive Primer (CIAP) is then applied to the repair doubler(s), and the repair area.

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Aircraft Composite Repair Technology The external doubler is installed on the repair area by applying the adhesive

material to it. Place the external doubler as shown in the Figure 124. A polyester tape is use to prevent movement of the doubler during curing similar to Figure 123. Do not seal the edges of the panel as it can prevent the flow of the adhesive material and the removal of air during the curing cycle. The polyester tape need to be applied around the edges of the doubler to prevent adhesive flash during the curing cycle.


Figure 11.6. . Once everything in place, the vacuum is applied and examines the vacuum bag for leaks. A vacuum bag that has a leak can cause porosity in the adhesive and a bond failure of the repair. A vacuum to a minimum of 22 inches (560 mm) of Hg (mercury) is applied Then, the vacuum source is removed and the vacuum gage is monitored. After 5 minutes, the total difference in the vacuum must be less than 5 inches (125mm) of mercury. Apply and keep a vacuum to a minimum of 22 inches (560 mm) of Hg (mercury) in the vacuum bag during the cure cycle.

Use clamps, weight, sand bags, or hydraulic jacks to apply pressure. The pressure must be applied equally to all the areas. This pressure must not be more than 35 to 40 psi (241 to 276 kPa). (d) Use cast ceramic, plaster, or plastic tools when you apply pressure to the outer skin on a contoured or rounded panel.

Apply the temperature with the heat blankets, heat lamps, or hot air, as necessary for curing. Monitor the temperature with the thermocouples. Make sure that the temperature does not increase at a rate more than 50F (30C) for each minute. Do not increase the temperature of the aluminum part above 2600f (1270c). If you do not obey, damage to the heat treat of the aluminum can occur.




the repaired part are following the specification and compliant to the regulation. Finally the repair area is sealed and the surface finish is applied. 11.6 External Doubler Repair on the Square Edge of a Panel with Corrosion Damage or Delamination This repair is similar to the repair stated in Chapter 11.5. The difference between the repair Figure 11.7 and Figure 11.8 is the removal of the aluminum laminate/layer with certain dimension according to the recommendation from the SRM. Upon the removal of the material, filler is replaced according to the dimension

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Aircraft Composite Repair Technology removed. A layer of adhesive film is placed between the undamaged part and the filler in order to strengthen the adherence force. A doubler is which is bigger than the filler is place on top to ensure a complete transition of load moving across the structure. Adhesive film is also placed in between the doubler and filler layer to add more strength. The curing time and curing temperature are similar to the repair stated above.

Figure 126: External Doubler Repair on the Square Edge of a Panel with Corrosion Damage or

Delamination (Courtesy of Boeing 737-400 SRM)

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Figure 127: External Doubler Repair of the Honeycomb Edge with Damage Through the Skin and the

Doublers (Courtesy of Boeing 737-400 SRM)

Figure 128: External Doubler Repair on the Square Edge of the Panel with Puncture Damage


Both Figure 127 and Figure 128 are similar in repair procedure except the type and laminate removal is different. Since both damages only occurred at the face

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Aircraft Composite Repair Technology sheet level, care need to be taken during the removal process. Ensure, the removal process does not protrude the core or the backside of the face sheet. Once the removal process is completed refer to Figure 11.5 for doubler installation and Figure 11.6 for curing by using heat blanket. The curing temperature and time are similar. 11.7 External Doubler Repair of One Skin and the Aluminum Honeycomb Core

The repair involved with the removal of the facesheet and the core material. For only one sheet and core removal it is known as one sided and if it is a fracture, then the repair is called two sided repair. In order to achieve proper at the end of the repair, the cure will be a step repair with the initial repair will be done on one face sheet. Upon curing, the next face sheet will be cured with the same settings. There many type of repair that can be found in this category. More types of repair are shown at the end of this chapter

This repair is applicable to damage that is 64 square inches (0.041 square

meters), or less, in total area. On one side only, the maximum length of damage permitted is 12 inches (305 mm). Find the limit of damage from the General Repair stated in the SRM. The finishes need to be removed before repair can be preceded.

All nicks, gouges, burrs, sharp edges, and other unwanted material nee to be removed to a smooth finish. The damaged skin need to trim to remove any delaminated areas. No sharp edges can present and it needs to be cornered to a minimum of 0.50 inch (12.5 mm) in radius. Air-powered router with a valve stem cutter can be used to remove the damaged skins and the honeycomb core from the repair area. See Figure 129 for a cutting position when making tapered cutting for the core.

Figure 129: Machining Aluminum Core with a Valve Stem Cutter


Make sure you do not allow the router to cut into the opposite aluminum skin if there is no damage on the other side. The bond area cannot be heat up to prevent damage to the adhesive system. For flush repairs, remove the honeycomb core down to the adhesive on the opposite aluminum skin as shown in Fig. 11.12a and Figure 11.12b

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Figure 130: Square Repair Doubler Specifications for External Patch Repair.


Figure 131: Repair Doubler specification for flush repairs.

(Courtesy of Boeing 737-400 SRM) The usage of this figure is to determine the doubler dimension only. The figure above only shows a rectangular type of removal area. However, if the shape is round, then the details can be found inside the SRM. The material for the repair parts must be equivalent to the same BMS and heat treat as the skin to be repaired. The bonding surfaces of 20-24 aluminum doublers can be clad or non-clad (bare). On the other hand, the bonding surfaces of 7075 aluminum doublers must be non-clad (Bare). For the initial skin thickness that is from 0.012 to 0.020 inch, the Part 3 external doubler is not necessary. Refer to Figure 11.13 for the size of doubler needed. Part 2 is already sufficient to use. Part 2 must be 1” larger than Part 3 if the

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Aircraft Composite Repair Technology dimension is bigger that stated above. Part 2 external doubler size can be retrieved from the chart below. Further information need to consult the SRM

Figure 132: Doubler specification chart


The core must be cut parallel to the direction of the bonded cell walls as in Figure 133. The damage area needs to be cleaned with a soft cloth moist with cleaning solvent. All of the contaminant requires to be removed, in and around the damaged area. The area must be fully dry before you continue on with this repair procedure. If the core thickness is partially remove, then see Figure 133.

Figure 133: External Doubler Repair of One Skin and the Aluminum Honeycomb Core


When making the repair, BMS need to be referred for the correct type. The repair core has to be the same shape and ribbon direction as the core that was removed. The core must be height thick enough so that it can be sanded flush with the outer surface of the skin.

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Figure 134: Bonded Doubler Repair That Uses Two Cores of Equal Depth and an Internal Septum for Full-

Depth Core Replacement (Courtesy of Boeing 737-400 SRM)

In doing this repair, several measures need to be taken such as the



are removed, the damage area needs to be cleaned. Water break test is used to examine the repair area for free surface. Anodize, or as an alternative, apply a chemical conversion coating to the non- primed surface of the repair area and the aluminum repair parts as given in the General Repair section.

The step is used only for bare metal surfaces not the core section. Remove all of the primer and paint before anodize, or apply a chemical conversion coating to the repair surfaces. The Corrosion Inhibiting Adhesive Primer (CIAP) is then applied to the repair doubler(s), and the repair area.

Before the repair core is install, a fit-up check between the repair core and

the core to be repaired need to be done. If necessary, sand the repair core flush with the outer surface of the initial skin make sure the gap between the repair core and the core to be repaired is a maximum of 0.10 inch (2.5 mm). Use only light finger pressure to make sure that the gap is constant over the surfaces to be bonded together. Then, Sand the repair core flush with the outer surface of the skin. Next clean the repair area by using a vacuum to remove the loose debris and other types of contaminants from the honeycomb core cells. If necessary, the repair surface may be clean by solvent.

Once everything is ready, then the core splice adhesive or adhesive foam is applied to the repair core as in Figure 135. After the adhesive material has been removed from a storage area, you must use the adhesive material before the manufacturer's time limit. If you do not obey, the bonding strength of the adhesive may be unsatisfactory.

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Figure 135: Partial and full depth core replacement.


Foaming adhesive, two-part adhesive or potting compound can be used to bond the repair core mating surfaces. Before you open the adhesive film or foaming adhesive containers, let them stay at a temperature of 60 to 800F (16 to 260C) until no more condensation is seen on the container.

Make sure that the adhesive materials do not touch other objects before they

are used to prevent them from contamination. As an alternative, the repair core and the repair doublers can be cured at the same time.

Cut one ply of the adhesive film to the same dimension as the hole in the core. Use two pieces of Grade 05 adhesive as an alternative for Grade 10. When using Grade 05 adhesive, make sure that you remove the separator between the pieces of adhesive materials. Remove the separator sheet from one ply of the adhesive film and put the film into the hole in the core. The foaming adhesive was cut to the correct dimension and winded around the repair core. The foaming adhesive was trimmed to make each end fit in a butt-joint, or you can allow a 0.125-inch (3 mm) overlap. A sufficient amount of the adhesive material was applied so that the adhesive thickness will fill a minimum of 2/3 the distance of the gap dimension. Multiple layers of adhesive material can be applied to the repair core depending on the need.

A non-stick flexible strip was winded around the repair core and the foaming adhesive during the application of the adhesive around the honeycomb core. The separator sheet was leave on the outside surface for protection before repair is assembled. Make sure the ribbon direction of the repair core is aligned in the same direction as the ribbon direction of the core to be repaired. Then put the repair core into the core hole. Make sure that you remove the separator sheet from the adhesive film before you put it into the core hole.

Make sure that the repair core has a tight interference fit in the core hold after it is wound with the adhesive film. Remove the non-stick flexible strip from the foaming adhesive. Push the adhesive smoothly and tightly in place. Use care not to trap any air under the adhesive. Make sure that the cell walls of the repair core are installed parallel to the initial core cell walls with the maximum error permitted is 3

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Aircraft Composite Repair Technology degrees. The maximum gap permitted between the edge of the repair core and the core to be repaired is 0.10 inch (2.5 mm).

To bond doublers and fillers, use one of the adhesives and apply the

adhesive film to the mating surface of the doubler. Do not remove the separator sheet from the adhesive film. This will help prevent the contamination of the adhesive film until you are ready to assemble the repair parts. Push the adhesive film smoothly and tightly in place. Use care not to trap any air under the adhesive film.

Remove the separator sheet from the adhesive film and push the repair doubler on the mating surface of the repair area. Hold the doublers in place with tape to prevent the movement of the doubler during cure in Figure 11.16 and Figure 11.17. Apply tape around the edges of the repair doublers for flash control. Apply a layer of adhesive film to the mating surface of the repair part. Put the repair part on to the repair mating surface and apply the mechanical pressure.

Figure 136: Installation of the doubler (Courtesy of Boeing 737-400 SRM)

Figure 11.16 shows the application of flash control tapes are in the areas that

require a close tolerance fit up but do not seal the edges of the panel. If the edges of panel are sealed, it can prevent the flow of the adhesive material and the removal of air during the final stage of the cure. Apply less than 25% of the edge length of the repair doubler with flash control tape. Make sure that there is sufficient perforated FEP parting film and breather material to conform to the contour of the part.

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Figure 137: Installation of the Doubler


Hold the doublers in place with tape to prevent the movement of the doubler during cure. Apply tape around the edges of the repair doublers to prevent leakage of the adhesive material. Apply the pressure for the adhesive cure. Do the vacuum bag procedures for curing as per Figure 137

The cure temperature for the adhesive film must not be more than 2600F (1270C). The result can be delamination in the part. To prevent delamination, you can apply 3 to 5 psi (20.7 to 34.5 kpa) pressure around this repair area by using shotbags or something similar. Make sure that you use thermocouples and a measurement system that can measure the cure temperature up to an accuracy of at least 50F (30C). If you do not, the result can be an unsatisfactory repair.

Place four thermocouples at the edge of the repair parts as shown in Fig. 11.8 and Figure 11.9.

1) Four thermocouples near the edge of the caul plate. 2) Four thermocouples near the edge of the doubler. 3) One thermocouple on each skin surface that will be heated and

pressurized. 4) One thermocouple on the nonpressured skin surface that is located on

the opposite side of panel from the repair. Lay up the applicable release film, heat blankets, support plate, and bleeder

padding as shown in Fig. The heat blankets must be a minimum of 2.0 inches (50 mm) larger all around than the repair doubler. Hot air and heat lamps can be used with the heat blankets to increase the temperature at the bondline.


Figure 138. Once everything in place, the vacuum is applied and examines the vacuum bag for leaks. A vacuum bag that has a leak can cause porosity in the adhesive and a bond failure of the repair. A vacuum to a minimum of 22 inches (560

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Aircraft Composite Repair Technology mm) of Hg (mercury) is applied Then, the vacuum source is removed and the vacuum gage is monitored. After 5 minutes, the total difference in the vacuum must be less than 5 inches (125mm) of mercury. Apply and keep a vacuum to a minimum of 22 inches (560 mm) of Hg (mercury) in the vacuum bag during the cure cycle.

Figure 138: Application of pressure during the cure for flat panel.


Figure 139: Application of pressure during the cure for rounded or contoured panels.


Use clamps, weight, sand bags, or hydraulic jacks to apply pressure. The pressure must be applied equally to all the areas. This pressure must not be more than 35 to 40 psi (241 to 276 kPa). (d) Use cast ceramic, plaster, or plastic tools when you apply pressure to the outer skin on a contoured or rounded panel.

Apply the temperature with the heat blankets, heat lamps, or hot air, as necessary for curing. Monitor the temperature with the thermocouples. Make sure that the temperature does not increase at a rate more than 50F (30C) for each minute.

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Aircraft Composite Repair Technology Do not increase the temperature of the aluminum part above 2600f (1270c). If you do not obey, damage to the heat treat of the aluminum can occur.




the repaired part are following the specification and compliant to the regulation. Finally the repair area is sealed and the surface finish is applied.

11.8 Other Typical Repair Below are the other typical repair that can use similar procedure as in the Chapter 11.7. Care must be taken in order to determine the type of repair used by looking at the SRM.

Figure 140: External Doubler Repair of Damage to One Aluminum Skin and the Aluminum Honeycomb

Core at the Edge of a Panel (Courtesy of Boeing 737-400 SRM)

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Figure 141: External Doubler Repair of Damage to Two Aluminum Skins and the Aluminum Honeycomb

Core at the Edge of a Panel (Courtesy of Boeing 737-400 SRM)

Figure 142: External Doubler Repair of Damage to One Aluminum Skin, the Internal Doubler, and the

Aluminum Honeycomb Core (Courtesy of Boeing 737-400 SRM)

pg 133


Figure 143: Repair of the Honeycomb Core at the Trailing Edge


Figure 144: Repair of the Honeycomb Core With Damage to a Single Lower Skin


pg 134


Figure 145: Repair of the Honeycomb Core With Damage to a Single Lower Skin and the Internal Doubler

Figure 11.25: (Courtesy of Boeing 737-400 SRM)

Figure 146: Repair of the Honeycomb Core With Damage to Two Skins by the Use of an External Doubler

and an Internal Doubler (Courtesy of Boeing 737-400 SRM)

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Figure 147: Repair of the Honeycomb Core With Damage to Two Skins by the Use of Two External

Doublers (Courtesy of Boeing 737-400 SRM)

Figure 148: Repair of the Honeycomb Core With Damage to Two Skins by the Use of External Doublers

on Each Skin Figure 11.29 (Courtesy of Boeing 737-400 SRM)

pg 136


Figure 149: Repair of the Honeycomb Panel from Damage or Delamination to the Outer Surface


Figure 150: Repair of the Honeycomb Panel from Major Skin Damage or Delamination


pg 137


Figure 151: Repair of the Honeycomb Core and the Skin from Damage


pg 138


REFERENCES 1. AC 43.13.1B Acceptable Methods, Techniques and Practices- Aircraft

Inspection, Repair & Alteration, Aviation Supplies & Academics, Inc., 1998 2. Crane D., Non Metallic Aircraft Structures, Airframe, Aviation Supplies &

Academics Inc., 1996 3. Wood, Composites, and Transparent Plastic Structures, A & P Technician

Airframe Textbook, Jeppensen Sanderson, 2000 4. 737-400 Structural Repair Manual 5. Foreman C, Composite Safety, Advanced Composite, Jeppensen Sanderson,


Advanced Composite, SAE International, 1998

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CHAPTER 12: FUTURE REPAIR AND TECHNOLOGY 12.1 Introduction

Carbon-fiber reinforced plastic (CFRP) has been used as a prime material for

aerospace structures because of its high specific strength and modulus in fiber direction. The use of CFRP dramatically reduces the weight of those structures, which is necessary in the aerospace field. However, the specific strength and modulus of CFRP in the fiber transverse direction is comparatively low. Therefore, CFRP is generally used as laminates or woven-type composites, which results in various complicated damage characteristics. Consequently, safety factors remain high compared with metal materials where weight reduction has not been fully achieved.

12.2 Aircraft Structural Health Monitoring (SHM)

When the aircraft DeHavilland Comet was introduced and operated, nobody knew that fatigue damage can lead to catastrophic event in a blink of an eye. From that on fatigue was one of the key study in the aircraft development phase. The science of mastering the prediction of fatigue failure has determined the duration of an aircraft needs to be inspected structurally. The newly development aircraft have placed millions of dollars in order to study this phenomenon and find a way to combat this problems. In order to acknowledge the problem, one must know when the structural failure can occur. A sensor that is embedded in the structural may be the solution of this problem. This sensors can trigger any traces of structural compromise by sending signals to the main computer in the aircraft to locate, access and warn any potential of structural failure.

Figure 152: The SHM will imitate the nerve system of a human

(Courtesy of www.airbus.com) Structural Health Monitoring System (SHMS) was develop to sense this potential structural failure thus to overcome the problem while it is in the infant stage. It is like the human body nerve system that can diagnose and interpret if any unhealthy symptom is sensed. See Figure 152. It is cheaper to counter and rectify the problem at this stage rather than it becoming major and the aircraft need to be grounded for long time. Before we further the discussion we need to understand the concept of SHM.

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Aircraft Composite Repair Technology The process of implementing a damage detection strategy for aerospace,

civil, and mechanical engineering infrastructure is referred to as Structural Health Monitoring (SHM). Apart from SHM, usage monitoring is also part of monitoring process. UM attempts to measure the inputs to and responses of a structure before damage so that the analysis can be used to predict the onset of damage and deterioration in structural condition. Prognosis is the coupling of information from SHM, UM, current environmental and operational conditions, previous component and system level testing, and numerical modelling to estimate the remaining useful life of the system.

The SHM process involves the observation of a system over time using

periodically sampled dynamic response measurements from an array of sensors, the extraction of damage-sensitive features from these measurements, and the statistical analysis of these features to determine the current state of the system’s health. For long-term SHM, the output of this process is periodically updated information regarding the ability of the structure to perform its intended function in light of the inevitable aging and degradation resulting from operational environments. After extreme events, such as impact or lightning strikes, is used for rapid condition screening and aims to provide, in near real time, reliable information regarding the integrity of the structure.

The advantage of placing the SHM in the aircraft has tremendous effect on

cost, availability and effectiveness in managing the aircraft fleet. Being an aircraft operator, the more the aircraft is flying, the more revenue it can generate. Any damage found during the operation of the aircraft will lead to un-schedule visit for rectifying the damage. By having the sensor, the occurrence of the defect can be monitored progressively thus avoiding unscheduled down-time for repairs due to the continuous damage monitoring. Furthermore it can avoid complex detail of visual inspection and NDT special inspections thus reduce penalizing inspection in small intervals. The sensors placement will also avoid the complex inspection with difficult access and/or with limited reparability.

There are many types of sensor used for the SHM application. These

sensors will we placed at a strategic place within the aircraft structure. Various types of sensors available and currently under research are used due to its unique function to detect certain type of damage and suitability. The sensors can detect any type of impact, delamination, debonding, water ingression and load/strain.

Figure 153: Sensor placement on A380 experimental aircraft

(Courtesy of www.airbus.com)

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Aircraft Composite Repair Technology Figure 153 some of the sensors placed on an actual aircraft for testing and

data collecting purposes.

12.2.1 Comparative Vacuum Sensors (CVM)

It uses adhesive-backed patch at 1.5mm thick contains tiny tubes or “galleries” that are open to the surface of the part being inspected. When the patch is applied to part using hand pressure, the galleries are closing with the part surface. Any changes of the pressure will leak the vacuum and detection can be easily done. CVM technology is very small and low weight. The system is passive and only be use during periodic interrogation. The patch materials is durable, fuel resistance and very useful at hard to reach target such as at wing box or rear pressure bulkhead. See Figure 154.

Figure 154: CVM placement on structure

(Courtesy of www.airbus.com) 12.2.2 Acoustics Emissions (AE)

AE is an elastic radiation generated by the rapid release of energy from sources within a material. These elastic waves are detected and converted to voltage signals by small piezoelectric sensors mounted to a convenient surface of the material. See Figure 155. Sources of this AE is fracture, plastic deformation, delamination, impacts etc. in which the sensors can detect with noise at the ambient environment. It has the sensitivity to detect newly formed crack surface down to a few hundred square micrometer and less.

Figure 155: AE in application

12.2.3 Eddy Current Foil Sensors (ETFS)

ETPS is used on metallic surfaces only. Cracks and corrosion patches alter the electromagnetic field induced by the eddy current generated by the sensors. The

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Aircraft Composite Repair Technology system works by monitoring of crack growth and corrosion by interaction of the defects with eddy current field. See Figure 156.

Figure 156: ETPS sensor film 12.2.4 Fiber Bragg Grating (FBG)

FBG is a type of Bragg sensor reflector constructed in a short segment of an optical fiber that reflects particular wavelengths of light and transmits all others. This sensors are used for static strain measurement. See Figure 157 The use of this sensor is by embedded the sensors within the laminate. This application does not degrade the strength of the material and the diagnosing method is by monitoring all rib strains in their longitudinal directions. A suitable arrangement or pattern of FBG needs to be placed to collect the data.

Figure 157: FBG embedded inside a laminate

12.2.5 Acoustic Ultrasonic (AU)

AU uses ultrasonic methods in a frequency range typical of acoustic emission application. The technique is able to detect and characterize difference in the structure of single and multilayer metallic, ceramic and composite sheet materials. AU uses pulser and receiver transducer with resonant frequencies in the low ultrasonic range, couple with wave propagation dynamics prediction to detect damage. When damage occurs, changes in the signal indicate the type of damage. The damage can be evaluated from AU measurement by calculating the changes of the signals from a given types and degrees of damage. See Figure 158.

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Figure 158: AU application on a structure

12.3 Thick Laminate Repair

Boeing has spent considerable effort attempting to make the 787’s solid laminates less susceptible to the dents and dings associated with hail damage, tool drops and jetway impacts that plague thinner composite structures. Since most of the structures are made from advanced composite material which is carbon-graphite, they are very prone to impact damage. See Figure 159. The company has performed extensive research in damage probability, calculating energies associated with everything from the corner of a toolbox striking an upward facing structure to the impact of a jetway on a cargo or passenger entry door. They have tried to correlate the damage that we see today and the repairs that are required and translate it into impact energy. The result has been new sizing for upward facing structures and structures around doors to make the airplane more resistant to damage.

Figure 159: Application of material on 787 structures

(Courtesy of www.boeing.com)

Damage will occur, however, and when it does, the first step is to assess if and how much internal damage there is within the laminate, which will require specialized NDI such as ultrasonic inspection. Yet, when a small dent is noticed during visual inspection at the gate, performing ultrasonic testing can be time consuming, and will require a technician with specialized training. When small impact damage is noticed, the SRM will provide allowable damage limits, based on the dent size. A thorough inspection on the delamination inside that structure will show that the limits are more generous in allowing you to dispatch the airplane. To aid this process, Boeing developed the Ramp Damage Checker — a simple “go/no go” handheld device (similar to a stud finder) that identifies the presence of delamination.

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Aircraft Composite Repair Technology It’s not intended to provide detailed information, such as depth, but it can confirm the presence of delamination. Boeing en-vision it will be used in the field by a maintenance lead or inspector and not necessarily an ultrasonic specialist. The first commercial Ramp Damage Checker is now available from Olympus NDT (Waltham, Mass.). Other brands will follow.

For damage repair, Boeing has specified a variety of options in the 787 SRM: traditional bonded repairs; the company’s patented quick composite repair technique; and conventional bolted repair. 12.3.1 Bonded repair

It is the typical repair that is applicable to both laminate and sandwich structure. It includes the 250°F/121°C to 350°F/177°C prepreg repairs as well as 200°F/93°C wet layup repairs. Repair times may increase when multiple cure cycles are needed for thicker structures or if extensive repair is necessary. The general rule of thumb for scarfing is to remove 0.5 inch/13 mm per ply or 1:40 ratio of thickness to the surface laminate. Ply thickness varies depending on the materials used in the original structure, but for carbon fiber unidirectional tape, it can be as thin as 0.13 mm/0.005 inch. A 0.375-inch/9.525-mm thick structure could be approximately 75 plies thick. By the standard rule of thumb, this would translate to a scarf of ~37.5 inches/~952.5 mm.

Scarf distance is measured not from the center of the damage but from the edge of the cleaned-up damaged area. So if the damaged structure has a 6-inch/152-mm diameter hole in it after the damage is removed, then the outer diameter of the scarfed area would be 37.5 inches/952.5 mm plus 6 inches/152 mm plus 37.5 inches/952.5 mm, for a total diameter of 81 inches/2m. A scaft area of this size is very time consuming to work with.

A bolted doubler repair is much quicker and generally easier to perform at this size. However, there may be aerodynamic reasons, concerns about damaging the underlying structure during drilling, or radar signature reasons in military aircraft, which would require a scarfed repair.

Other engineering point of view supports a different scarf ratio, such as 0.25 inch/6.35 mm per ply. If so, then the total diameter of the repair would be 43.5 inches/1.1m, which is still large but an easier repair to perform. With a smaller scarf, less undamaged material is removed but the area of the adhesive bond, which transfers load through the repair plies, also is reduced.

The quick composite repair is not permanent, but rather a “Band-Aid” designed to get a damaged airplane back into service quickly. The precured composite patch is epoxy bonded onto the outside of the airplane over the damaged area. It restores enough residual strength into the damaged area to provide revenue service on a temporary basis. The process allows small-area damage to be repaired in less than an hour. The adhesive is cured by relatively low temperatures provided by a chemical heat pack, which eliminates the need to dry the part, and is designed to be applied at the gate if necessary.

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Aircraft Composite Repair Technology 12.3.2 Bolted repair

It is the same types of mechanically fastened repairs — doublers, scab patch, flush, etc. — that are performed on metal airplanes have been included in the SRM for the 787. In a bolted repair, a cover plate is mechanically fastened around the damaged area. Although fasteners create stress concentrations that can degrade the performance of the parent structure, bolted repair of composites has been service proven on Boeing’s 777.

Typically these repairs would be carried out using a titanium sheet, although Boeing has also successfully tested carbon fiber patch materials and is specifying aluminum as an additional option. A special care must be taken to protect the galvanic coupling between an aluminum patch and the carbon fiber skin if titanium or carbon fiber is not used. Because an aluminum patch also would require periodic inspection for corrosion, its use would most likely be temporary, in which case, paint scheme and faying surface sealant protection on the aluminum part would be adequate to protect against galvanic corrosion. Titanium fasteners, however, are required in all cases.

At the highest level, the bolted repair for a composite aircraft is the same as that on a metal aircraft. It is the same process, same skills, and, in general, the same tools except when you drill into a composite structure, you use the same drill motor but a different drill bit, and you have to adjust speed and pressure. Fiber breakout on the backside need to be aware when performing the repair.

There is certainly a need for training to understand a few minor differences, but, in general, a line technician or mechanic who performs bolted repairs on a metal aircraft today will very easily transition into bolted repairs on a composite structure. Generally, the thickest laminates would be in areas of heavier loads, such as cargo door surround structure. When these areas are damaged on a metal aircraft, it generally requires a stacked doubler repair and significant teardown and re-assembly. The composite repair to the same areas will have a different work scope, but the anticipation of the downtime is significantly different

12.3.3 Bonded Versus Bolted Repair

The choice between a bonded and a bolted repair may come down to how much time you have available to do the repair. The huge advantage of the bolted repair is there is no heat required. Although a bolted repair can impact the aerodynamics and radar signature of the aircraft, in Boeing’s view, flush bolted repairs are nearly equivalent to bonded repairs in terms of aerodynamics and cosmetics. Further, when a metal/composite stack-up is drilled, residual metal chips can damage composite holes. Also, dull bits or incorrect drill speeds can burn composites, and proper steps must be followed to avoid hole misalignment.

Generally, bonded patches provide more efficient load transfer than bolted repairs and are more attractive from an aerodynamic and cosmetic standpoint. And, as with bolted repairs, the quality of a bonded repair depends on many variables: age and quality of materials, surface preparation, and successful adhesion. In essence,

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Aircraft Composite Repair Technology the success of either a bonded or bolted repair relies heavily on the skill of the technician. 12.3.4 On-Aircraft Curing Options

In-situ composite repairs that require high-temperature cures can be problematic. Obtaining uniform cure temperatures is difficult due to heat sinks created by the airframe structures. In addition, when performing high-temperature in-situ repairs on the flight line with a fueled aircraft, many restrictions apply in terms of acceptable equipment and tools, and standard hot-bonding equipment is not allowed at the gate because of the risk of fire. Wichitech Industries (Baltimore, Md.) and HEATCON Composite Systems (Seattle, Wash.) offer hot bonders designed specifically for use on fueled aircraft. HEATCON’s design purges and pressurizes its bonders and uses arc-suppressing power connectors, while Wichitech hermetically seals the elements of its bonders and includes an internal vacuum pump to speed setup time.

For wet-layup, glass-reinforced repair processes, a rapid-cure resin system patented by TRI/Austin (Austin, Texas) and developed for the U.S. Air Force Research Laboratory (Wright Patterson AFB, Dayton, Ohio) reportedly cures in as little as 20 minutes, using UV light at low temperature. During the process, alternating layers of the acrylate-based resin system and woven fiberglass fabric are applied to fill the hole and form a UV-curable composition. Traditional vacuum bagging is then applied, and the patch is irradiated with UV light. The temperature peaks briefly at 60°C/140°F and then levels off at 30°C/86°F. In general, the width of the patch can be up to 2 ft/0.6m, and the depth can be as much as 200 mils/0.2 inch. Reportedly, glass-reinforced composites as thick as 120 mils/0.12 inch can be thoroughly cured. Although essentially a depot-level repair, the system can be deployed in the field when necessary to return an aircraft to service.

Another option for low-temperature curing of on- and off-aircraft composite repairs may be electron beam (EB) or X-ray curing, which is relatively fast and requires no heat input. Air Canada Maintenance Base (Winnipeg, Manitoba, Canada), with the help of Lockheed Martin Skunk Works (Palmdale, Calif.) and Atomic Energy of Canada Ltd. (Pinawa, Manitoba, Canada), studied the feasibility and economics of EB curing for repair by testing a variety of composite components, including the fairing from an Airbus A320 aircraft. Subsequently, Transport Canada issued a Repair Design Certificate for EB repair of fiberglass wing to box fairing panels. Acsion Industries (Pinawa, Manitoba, Canada) supplies EB-curable products and related services.

A deep-core curing method developed by Cornerstone Research Group Inc. (CRG, Dayton, Ohio) uses a photo-delivery system to cure certified aircraft repair adhesives without surface heating. Developed for military aircraft, CRG’s prototype photo-delivery system integrates a high-power optical energy source with an optical scrim embedded into the film adhesive. The technology reportedly provides uniform heat distribution within the bond line and achieves full cure at 250°F/121°C without backside heating.

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Aircraft Composite Repair Technology 12.3 Glass-Aluminum Reinforcement (GLARE) Laminate Repair

The Airbus A380 is one of the latest development of large size airliner that is using many innovate technology, advanced materials and user friendly system. One of the technology breakthrough is the use of glass reinforced plastic or known as GLARE. See Figure 160. The A380 is the first aircraft to use Glare, a hybrid material (fiber metal laminates), built-up from alternating layers of aluminum foils and unidirectional glass fibers, impregnated with an adhesive. Dutch-based Stork developed this new fiber laminate with Airbus. Its qualities are a compromise between conventional aluminum alloys and carbon-fiber reinforced plastic (CFRP).

Figure 160 Major monolithic CFRP and thermoplastic applications on A 380

Glare is lighter than aluminum and has very good mechanical properties. Fatigue and damage resistance are superior to those of aluminum. The energy required to create a dent (of a given depth) is much higher than it is for aluminum. Cracks initiate earlier but are stopped at each glass fiber layer. As a result, the propagation is so slow that a crack cannot reach a critical size in the entire aircraft's life in service. Due to its high resistance to impacts, Glare was chosen for the leading edges of the empennage. Corrosion, too, is stopped by the sandwich structure. Tests have shown that corrosion is stopped at the first glass fiber layer, whereas, in the same conditions, it goes through the aluminum's thickness.

Glare was chosen for the upper fuselage in the forward and aft sections because of its fatigue performance and damage tolerance. Aluminum was chosen for the center section because it is submitted to higher static loads, due to the presence of the wing. A butt strap, however, is made of Glare in the center upper fuselage. A German military A310 transport was fitted in 1999 with a Glare fuselage panel. Since then, it has accumulated 5,056 flight hours and 1,339 cycles.

Where Glare is used, CFRP was also an option. It is lighter and its performance, especially in terms of fatigue resistance, is very close. However, its production cost is higher. And, most importantly, it was not seen as mature enough for use on fuselage panels.

As Glare is a new material, an inspection program has been set up. It affects the first years of operation of the A380. These inspections will be added to the

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Aircraft Composite Repair Technology existing scheduled maintenance program. In the mid-term, Airbus is confident that it will require fewer inspections than conventional aluminum alloys. After a repair on a Glare panel, no specific inspection will be required.

After accidental damage on a Glare panel, a visual inspection is supposed to allow a first assessment of the damage extent. Therefore, this does not require special skills in addition to those required for visual inspections of aluminum panels. If necessary, in case of very large impact damage, ultrasonic inspection is possible as it is on aluminum panels.

Repair principles are the same on aluminum and Glare. Technicians use conventional tooling, for example, carbide tools for manual trimming and drilling. The steps they have to follow are exactly identical to those of an aluminum panel repair procedure. Of course, the repaired Glare panel's properties are at least as good as they were before the damage. Most frequently affected areas are where servicing trucks can hit the aircraft on the ground--namely the lower fuselage and the doors' surroundings. In general, after structural repair manual (SRM) consultation and as soon as damage is found above the ADL limit, maintenance engineers have to report to Airbus for repair advice before any further flight.

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REFERENCES

1. Black S., Strctural Health Monitoring: Composite Get Smart, High Performance Composites, September 2008.

2. Roach D., Smart Aircraft Structures: A Future Necessity, High Performance Composites, January 2007

3. Roach D., Assessing Conventional and Advanced NDI for composite Aircraft, High Performance Composites, June 2008

4. Wood K., In-situ Composite Repair Builds on Basics, High Performance Composites, November 2008

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Aircraft Composite Repair Technology CHAPTER 13: SAFETY IN REPAIR 13.1 General

Safety is one of the major concern in the composite industry because it

involves chemical, mechanical works, residues etc. Materials used in or during repair of the items manufactured with fiberglass reinforced plastics, are usually toxic, irritating or flammable products.

These practices must abide the rules and regulations set up by the authorities

in that particular place. This involves the government and local authorities. In Malaysia, the workers who are involved in this field are covered in the Occupational Safety and Hazard Act. Failure to follow the act will result

withdraw for any compensation or benefit covered in the insurance. In the shop a few safety consideration must be noticed or known before any

task is started. They are:- 1) Respirable Fiber and dust. For example fiber glass, aramid carbon

etc. 2) Fumes and Vapor. Especially from the resin, solvent, sealant,

coatings, fuel and acid vapor. 3) Exothermic reaction. Chemical reaction that is produce when certain

part are mixing. Certain chemicals are very dangerous if it is expose for too long. The limit of exposure are known as Threshold Limit Value (TLV). It is refer to airborne concentrations of substances and represent conditions under which is believed that nearly all personnel may be repeatedly exposed day after day without adverse effect. The American Conference of Governmental Industrial Hygienist (ACGIH) has establish the following three categories for TLV

Threshold Limit Value – Time-Weighted Average (TLV-TWA) The time weighted average for normal 8-hour work a day and 40-hour a

week, to which most workers may be exposed daily. Threshold Limit Value-Short Term Exposure Limit (TLV-STEL) The concentration to which a worker can be expose continuously for a

short period of tme (15min) without suffering any irritation, chronic or irreversible tissue damage etc.

Threshold Limit Value-Ceiling (TLV-C) The concentration that should not be exceed during any part of the work

day.

13.2 Precautions

In making the repair a few precautions need to be observed and followed in order to ensure good quality of repair and safety. Failure to follow the recommendation stated in the manual will result un-airworthy repair. Caution steps are:-

1) Fasteners installed in graphite composite structure must be bare or coated titanium coated material, or corrosion resistant steel. Cadmium plated corrosion resistant steel may also be used.

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Aircraft Composite Repair Technology 2) Aluminum alloy steel fasteners are not allowed in graphite composite

structures. When reinstalling aluminum fittings on graphite composite structure, ensure that the original corrosion protective treatment is maintained.

3) Install fittings with faying surface approved seal in fuel tank areas or all other areas.

4) Do not let carbon fiber dust go into electrical equipment. Carbon fiber dust is electrically conductive and it can cause short circuits if it goes into electrical equipment. Use a vacuum near the source of the dust to remove the dust from the air. If you do not obey, you can cause damage to the electrical equipment.

5) To get an acceptable quality repair, do as follows: § refer to the applicable component which you are going to repair for the

repair limits and material of the component before you use these repair instructions.

§ use the correct materials for the type of component and repair that is made.

§ use the materials that agree with the specifications in the SRM. § accurately follow the SRM procedures at each step of the repair.

6) If you do not do this, it can cause a repair that is not satisfactory and is not approved.

7) Room temperature/1500F (660C) cure repairs will not restore either the strength or the durability of the original 2500F (1210C) or 3500F (1770C) cure components.

8) For size and limits of such repairs, see applicable repair subject. 9) Room temperature/1500F (660C) repairs must not be used in stress critical

areas of primary structure components. Failure to comply would result in an inadequate repair.

10) Do not use chemical paint strippers to remove paint before making damage evaluations. Damage to the adhesive resin system will occur.

11) Do not immerse parts in trichloroethane solvent or allow standing solvent on part. Damage to part will occur.

12) Do not exceed immersion criteria given. Damage to core material will occur.

13) Do not immerse parts in trichloroethane solvent or allow standing solvent on part. Damage to part will occur.

14) The use of precured patches is not recommended. Precured patches bonded to the structure under vacuum pressure only and large precured patches bonded to contoured surfaces can result in porous or noncontinuous bondlines.

15) Ensure that parting film is removed from repair plies prior to layup and curing. Noncompliance will result in a ruined repair.

16) For repairs to hollow assemblies (e.g., elevators, rudders, or Ailerons) do not vacuum bag the entire part. Full atmospheric pressure must be maintained inside the hollow assembly or the assembly might collapse.

17) Surface temperature must not exceed 1700F (770C). Damage or distortion of structure may occur if temperature exceeds 1700F (770C).

18) Do not sand into original structure. Failure to comply will reduce the strength of the component.

19) Repairs made to control surfaces and/or adjacent structure must not interfere with the designed operation of the control surfaces. Damage to airplane structure may occur.

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Aircraft Composite Repair Technology 20) Repairs must not cover existing drain holes in panels. Water will

accumulate if drain holes are covered. 21) Do not sand into bond ply or core. Loss of structural strength of the

component will occur if this caution is not observed. Bond ply is ply adjacent to core.

22) On hybrid panels, ensure that graphite repair plies do not extend into areas of fastener holes. Electrolysis between the metal fasteners, support structure, and graphite may cause corrosion in the aluminum structure.

23) Do not sand into fibers in the area surrounding the damage. 24) The following repair is limited to lightning damage not penetrating fibers. If

fibers or core have been damaged, refer to specific component structural repair subject.

25) The following repair has inspection requirements and time limits. See specific component structural repair subject for repair limitations. Failure to comply could result in an unacceptable and unauthorized repair.

26) If skin is exposed to direct contact with uncured resins or curing agent, wash with warm water or soap. Avoid the use of solvents for cleaning the skin.

27) To prevent contamination of the resin, do not use waxed containers for mixing.

28) Do not sand into fibers in the area surrounding dent.

More precaution needs to be taken when Warning is written. The instructions are as follow:-

1) Flight safety demands that certain control surfaces be properly balanced at

all times once repair is made. 2) For accelerated cure, use heat curing equipment that is acceptable to local

fire protection authorities. Noncompliance could cause personnel injury. 3) For accelerated cure, use heat curing equipment that is acceptable to local

fire protection authorities. Failure to comply could cause personnel injury. 4) Sanding gives off a fine dust that may cause skin irritations. Breathing of

an excessive amount of this dust may be injurious. Observe precautions for skin and respiration protection.

5) Air-powered equipment must be used where the possibility of vapor ignition exists. Noncompliance could cause personnel injury.

6) Keep solvents away from sources of heat, fire, or sparks. Heat, fire, or sparks can cause an

7) Explosion. Avoid contact of solvent with skin, eyes, and clothing. 8) Wear eye protection and use mechanical ventilation or respiratory

protection when working in a confined space or area. Breathing vapors or allowing solvent to contact skin or eyes is hazardous.

9) Some equipment used for heat curing repairs may constitute a fire or explosion hazard when used in the vicinity of an airplane which contains, or has contained fuel. Consult the local fire department for authority to use specific equipment. Noncompliance could cause personnel injury.

10) Breathing vapors or allowing solvent to contact skin or eyes is hazardous. Wear neoprene gloves with cotton liners, protective clothing, and eye goggles. If chemical contact occurs, wash thoroughly with water. If chemical should splash into eyes, flush eyes with large quantities of water and seek medical aid.

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Aircraft Composite Repair Technology 11) Use mechanical ventilation or respiratory protection when working in a

confined space or area. 12) Do not breathe carbon fiber dust. Make sure that there is a good flow of air

where you do the work. 13) Use equipment to help you breathe when you work in a confined space.

Use a vacuum near the source of the dust to remove the dust from the air. If you do not obey, you can cause injury.

14) For machining, cutting and/or milling these materials, proper eye protection must be provided. 15) Any machines used must be handle with great care.

16) The area wherein products are mixed must be well ventilated and away from other area wherein open flames or electrical sparks exist.

17) Wear safety goggles and rubber gloves for handling solvents, resins and prepregs.

18) Keep working area clean and free from residue from previous work. 19) Do not delay cleaning of spillage of products used for mixture. 20) Oxidizing agents shall be stored in separated areas apart from other materials. 21) Make sure the hardener and the resin come from the same company and the

same product. Never mix them up. 22) Always wear proper safety attire. 23) Make sure read the instruction carefully when using chemical items. 24) Ensure the surrounding area are well ventilated. 25) Clean the shop every time after usage to prevent residue build up. 26) Do not throw chemical into the sink. 27) Wash hand after contact with any dangerous item such as the fibers, resin,

solvent etc. 28) Never mix accelerator directly to the catalyst, since explosion hazard is

involved. 29) Carefully washed hands after handling resins.

13.3 Emergency Action Procedures These steps are only a recommendation set by the writer. The most proper way to solve when a person make contact with the chemicals and hazardous material in the shop are to refer the MSDS or the instruction notes for the part. These treatment are only for temporary and later the personnel need to be send to hospital for further diagnose.

1) Skin Contact- Immediately remove liquids or pastes from the skin by wiping with disposable paper towels or remove the powder with bushing it off from the skin. Then cleanse the affected area with resin-removing cream, followed by washing with warm, soapy water.

2) Eye Contamination or Irritation- If eye contact occur, flush the affected eye with water immediately with an eyewash bottle or fountain for at least 15 minutes. The water should go away from another body parts.

3) Inhalation- Usually it comes from the powder, fumes, vapor mist or droplets. Action should put the personnel into rest until further help arrive.

4) Ingestion- Immediately rinse the mouth with water. If swallowing occurred, drink f water

5) Fires- Retard all fire with the proper fire extinguisher and wear proper safety attire to prevent any inhalation of fume, vapor etc.

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pg 155

Aircraft Composite Repair Technology CHAPTER 14: QUALITY CONTROL IN REPAIR 13.1 General Quality Assurance and Production shall be responsible for compliance with requirements herein, as well as these on Engineering drawings and applicable documents. 13.2 Control Procedure Any defects found by the appropriate personnel in bonding shops must be noted and documented in approved official document such as Material Review Board Discrepancy Sheet. Once the M.R.B. has determined the type of repair, materials, equipments and facilities and etc, if any, applicable copies of Discrepancy shall be distributed to Manufacturing and Control section, respectively, for accomplishment of the established repair. Quality Assurance will keep a record to contain the following data, as minimum:

§ Designation and part number. § Serial number § Summary of required actions to keep the part as serviceable. § Time, Temperature and Pressure used in repair

When applicable, accomplishment of test specimens, to assure some of the repair sequences and values, shall also be included in the control records. When using resins for polymerizing at room temperature the mixture excess shall be retained for subsequent examination. 13.3 Use Of Master Parts With Typical Defects Whenever particulars of items, either for appearance or work conditions cause the acceptance level to be different from those contained herein, it will be imperative to prepare master parts showing the defect in various extension stages, and specifying the acceptance level thereof. These masters shall be prepared by Manufacturing and evaluated joinly by Engineering and Quality Control.

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References 1. Niu Michael, C.Y. Composite Airframe Structure, Hong Kong Commilit Press

Ltd..1993. 2. Kumpulan Materi Training Composite Tooling, Pendidikan dan Latihan IPTN,

Bandung- Indonesia 1997. 3. C.A.S.A., Repair of wit Reinforced Plastic Manufacturing Elements, 1977. 4. ATA 51, Structural Repair Manual – Boeing 727, 2003

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