pg 1
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
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
8.11.4 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
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
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
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
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.
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
CHAPTER 2: FITTER & FINISH PROCESS
2.1 IntroductionFitter 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.
Aircraft Composite Repair Technology
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.
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.
Aircraft Composite Repair Technology
CHAPTER 3: INSERT INSTALLATION
3.1 InsertThe 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
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
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)
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.
Aircraft Composite Repair Technology
CHAPTER 4: COMPOSITES D AMAGE REPAIR
4.1 IntroductionRepair 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
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.
Aircraft Composite Repair Technology
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.
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.
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
CHAPTER 5: DAMAGE ASSESMENT
5.1 GeneralThe 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.
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
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
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 cmx 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
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.
Aircraft Composite Repair Technology
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.
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)
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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
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.
Aircraft Composite Repair Technology
CHAPTER 6: TYPES OF DEFECT & DAMAGE
6.1 IntroductionThe 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.
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)
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
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.
Aircraft Composite Repair Technology 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
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.
Aircraft Composite Repair Technology
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
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
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
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.
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.
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.
Aircraft Composite Repair Technology
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.
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.
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
CHAPTER 7: INSPECTI ON METHODOLOGY
7.1 Non Destructive InspectionIn 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.
Aircraft Composite Repair Technology
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
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.
Aircraft Composite Repair Technology
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
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
