Citation preview
CHAPTER 4 — FIBER REINFORCED COMPOSITE REPAIRS
FIBER REINFORCED COMPOSITE REPAIRS
Fiber
Composites...........................................................................................
4-27 4-2-7 Removal of Honeycomb Core in Glass or Carbon Fiber
Honeycomb
Reinforced Composite Parts
...............................................................................
4-54 4-3-1 Wet Layup Repairs —
General......................................................................
4-54 4-3-2 Approved Processes (Process Sheets)
......................................................... 4-55
4-3-3 Wet Layup Impregnation
Process..................................................................
4-56 4-3-4 Installation of Copper Wire Mesh in Wet Layup Repairs
............................... 4-61 4-3-5 Wet Layup Bagging
Process
.........................................................................
4-63 4-3-6 Curing Process for Wet Layup Using Epoxy Resin
....................................... 4-67 4-3-7 Honeycomb Core
Plug Installation and Splicing Using Epoxy Resin ............ 4-73
4-3-8 Preparing a Precured Wet Layup Doubler or Edge Filler
.............................. 4-76 4-4 Typical Wet Layup Repair
Procedures for Glass Reinforced Composites.......... 4-82 4-4-1
Glass Monolithic Laminate — Surface Damage
............................................ 4-84 4-4-2 Glass
Monolithic Laminate — Damage up to Half of Laminate
Structural
Plies
...............................................................................................................
4-90 4-4-3 Glass Monolithic Laminate — Damage Through Full
Thickness ................... 4-96 4-4-4 Glass Monolithic Laminate
— Laminate Edge Damage ................................ 4-107 4-4-5
Glass Honeycomb Panel — Surface
Damage............................................... 4-115 4-4-6
Glass Honeycomb Panel — Damage up to Half of Skin Structural Plies
...... 4-121 4-4-7 Glass Honeycomb Panel — Skin to Core
Disbond........................................ 4-127 4-4-8 Glass
Honeycomb Panel — Smooth
Dent..................................................... 4-137
4-4-9 Glass Honeycomb Panel — Single Skin and Core Damage
......................... 4-142 4-4-10 Glass Honeycomb Panel — Hole
Through Panel (Both Skins) ..................... 4-154 4-4-11 Glass
Honeycomb Panel — Core Bevel Damage
......................................... 4-169
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4-4-12 Glass Honeycomb Panel — Panel Edge Damage
........................................ 4-177 4-4-13 Aluminum
Facing Honeycomb Panel — Core Bevel Damage.......................
4-188 4-5 Typical Wet Layup Repair Procedures for Carbon Fiber
Composite Parts......... 4-193 4-5-1 Carbon Monolithic Laminate —
Damage Through First Ply .......................... 4-195 4-5-2
Carbon Monolithic Laminate — Damage up to Half of Laminate
Structural
Plies
...............................................................................................................
4-201 4-5-3 Carbon Monolithic Laminate — Damage Through Full
Thickness ................ 4-207 4-5-4 Carbon Monolithic Laminate —
Laminate Edge Damage.............................. 4-218 4-5-5
Carbon Honeycomb Panel — Surface Damage
............................................ 4-226 4-5-6 Carbon
Honeycomb Panel — Damage up to Half of Laminate Structural
Plies
...............................................................................................................
4-232 4-5-7 Carbon Honeycomb Panel — Skin to Core Disbond
..................................... 4-238 4-5-8 Carbon Honeycomb
Panel — Single Skin and Core Damage....................... 4-248
4-5-9 Carbon Honeycomb Panel — Hole Through Panel (Both Skins)
.................. 4-260 4-5-10 Carbon Honeycomb Panel — Core Bevel
Damage....................................... 4-275 4-5-11 Carbon
Honeycomb Panel — Panel Edge
Damage...................................... 4-283
FIGURES
4-1 Wet Layup Repair Flow
Chart.............................................................................
4-8 4-2 Damage Outlining
...............................................................................................
4-20 4-3 Acceptance Criteria for Protruding Head Fastener
Holes................................... 4-24 4-4 Acceptance
Criteria for Countersunk Fastener
Holes......................................... 4-26 4-5 Copper Wire
Mesh — Installation
.......................................................................
4-62 4-6 Epoxy Cure without Heat Blanket — Bagging Process
..................................... 4-64 4-7 High Temperature
Epoxy Cure with Heat Blanket — Bagging Process ............. 4-66
4-8 Core Plug — Installation
.....................................................................................
4-75 4-9 Precured Wet Layup Doubler — Fabrication
...................................................... 4-78 4-10
Precured Wet Layup Edge Filler —
Fabrication.................................................. 4-81
4-11 Damage through Surface Finish of Glass Laminate — Negligible
Damage ....... 4-85 4-12 Damage Through First Ply of Glass Laminate
— Repair.................................... 4-89 4-13 Damage
through Half of Glass Laminate Thickness — Negligible Damage.......
4-91 4-14 Damage Through Half of Glass Laminate Thickness — Repair
......................... 4-95 4-15 Small Puncture Through Glass
Laminate — One Sided Repair ......................... 4-98 4-16
Small Puncture Through Glass Laminate — Double Sided Repair
.................... 4-101 4-17 Large Puncture Through Glass
Laminate — Repair........................................... 4-106
4-18 Small Edge Damage in Glass Laminate — Repair
............................................. 4-109 4-19 Wide Edge
Damage in Glass Laminate —
Repair.............................................. 4-114 4-20
Damage through Surface Finish of Glass Panel Skin — Negligible
Damage..... 4-116 4-21 Damage Through First Ply of Glass Panel Skin
— Repair ................................. 4-120 4-22 Damage
through Half of Glass Panel Skin Thickness — Negligible Damage ....
4-122 4-23 Damage Through Half of Glass Panel skin Thickness —
Repair ....................... 4-126 4-24 Disbond between Glass
Panel Skin and Core — Negligible Damage ................ 4-128 4-25
Glass Panel Skin to Core Disbond — Repair
..................................................... 4-132
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Figure Page Number Title Number
4-26 Large Glass Panel Skin to Core Disbond — Repair
........................................... 4-136 4-27 Dent in
Glass Panel Skin — Negligible Damage
................................................ 4-138 4-28 Smooth
Dent in Glass Panel Skin —
Repair....................................................... 4-141
4-29 Small Puncture Affecting One Glass Panel Skin and Core —
Repair ................ 4-145 4-30 Puncture Affecting One Glass
Panel Skin and Core — Repair .......................... 4-149 4-31
Large Puncture Affecting One Glass Panel Skin and Core — Repair
................ 4-153 4-32 Small Puncture Affecting Both Glass
Panel Skins and Core — Repair .............. 4-157 4-33 Puncture
Affecting Both Glass Panel Skins and Core — Repair
........................ 4-162 4-34 Large Puncture Affecting Both
Glass Panel Skins and Core — Repair.............. 4-167 4-35 Small
Edge Damage Affecting One Glass Panel Skin and Core —
Repair........ 4-172 4-36 Large Edge Damage Affecting One Glass
Panel Skin and Core — Repair........ 4-176 4-37 Small Edge Damage
in Glass Panel —
Repair................................................... 4-179
4-38 Large Edge Damage Affecting Both Glass Panel Skins and Core —
Repair ..... 4-185 4-39 Small Edge Damage Affecting One Glass Panel
Skin and Core — Repair........ 4-192 4-40 Damage through Surface
Finish of Carbon Laminate — Negligible Damage..... 4-196 4-41
Damage Through First Ply of Carbon Laminate — Repair
................................. 4-200 4-42 Damage through Half of
Carbon Laminate Thickness — Negligible Damage .... 4-202 4-43
Damage Through Half of Carbon Laminate Thickness —
Repair....................... 4-206 4-44 Small Puncture Through
Carbon Laminate — One-sided Repair....................... 4-209
4-45 Small Puncture Through Carbon Laminate — Double-sided Repair
.................. 4-212 4-46 Large Puncture Through Carbon
Laminate — Repair ........................................ 4-217
4-47 Small Edge Damage in Carbon Laminate — Repair
.......................................... 4-220 4-48 Wide Edge
Damage in Carbon Laminate — Repair
........................................... 4-225 4-49 Damage
through Surface Finish of Carbon Panel Skin — Negligible Damage ..
4-227 4-50 Damage Through First Ply of Carbon Panel Skin —
Repair............................... 4-231 4-51 Damage through
Half of Carbon Panel Skin Thickness — Negligible Damage.. 4-233
4-52 Damage Through Half of Carbon Panel Skin Thickness — Repair
.................... 4-237 4-53 Disbond between Carbon Panel Skin
and Core — Negligible Damage ............. 4-239 4-54 Carbon Panel
Skin to Core Disbond —
Repair................................................... 4-243
4-55 Large Carbon Panel Skin to Core Disbond —
Repair......................................... 4-247 4-56 Small
Puncture Affecting One Carbon Panel Skin and Core —
Repair.............. 4-251 4-57 Puncture Affecting One Carbon Panel
Skin and Core — Repair........................ 4-255 4-58 Large
Puncture Affecting One Carbon Panel Skin and Core —
Repair.............. 4-259 4-59 Small Puncture Affecting Both
Carbon Panel Skins and Core — Repair ........... 4-263 4-60
Puncture Affecting Both Carbon Panel Skins and Core — Repair
..................... 4-268 4-61 Large Puncture Affecting Both
Carbon Panel Skins and Core — Repair ........... 4-273 4-62 Small
Edge Damage Affecting One Carbon Panel Skin and Core — Repair .....
4-278 4-63 Large Edge Damage Affecting One Carbon Panel Skin and
Core — Repair ..... 4-282 4-64 Small Edge Damage in Carbon Panel —
Repair ................................................ 4-285 4-65
Large Edge Damage Affecting Both Carbon Panel Skin and Core —
Repair .... 4-290
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Table Page Number Title Number
4-1 Repair Selection Guide for Rivet Pattern Discrepancy
....................................... 4-9 4-2 Repair Selection
Guide for Bonded Panels With Aluminum Face Sheets
and Glass
Edging................................................................................................
4-9 4-3 Repair Selection Guide for Glass or Carbon Monolithic
Part.............................. 4-10 4-4 Repair Selection Guide
for Bonded Panel
.......................................................... 4-11 4-5
Process
Sheets...................................................................................................
4-15 4-6 Drilling Damage Limits for Protruding Head Fastener Holes
.............................. 4-23 4-7 Drilling Damage Limits for
Countersink Side of Holes ........................................
4-25 4-8 Process
Sheets...................................................................................................
4-55 4-9 Number of Repair
Plies.......................................................................................
4-60 4-10 Standard Cure at Room Temperature
................................................................
4-69 4-11 Room Temperature Cure with Post-cure
............................................................ 4-71
4-12 High Temperature Cure
......................................................................................
4-72
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4-1. INTRODUCTION
Many of the repairs for the Bell Helicopter Textron products are a
function of the type of construction to be repaired. Some repairs
are in fact similar in nature and may be applicable to various
parts of the structure. Those repairs are therefore considered
typical for a certain type of construction. This chapter describes
fiber reinforced composite repairs using glass or carbon fiber
fabrics. This chapter provides “typical repairs” for the following
types of structure:
NOTE
A monolithic laminate is defined as a unit formed or composed of
material without joints or seams consisting of or constituting a
single unit.
• Monolithic laminate made from glass fibers and epoxy resin
• Honeycomb panel with skin(s) made from glass fibers and epoxy
resin (Refer to Chapter 3 for honeycomb panels made with aluminum
facings)
• Monolithic laminate made from carbon fibers and epoxy resin
• Honeycomb panel with skin(s) made from carbon fibers and epoxy
resin
4-1-1. CONSIDERATIONS PRIOR TO REPAIRING FIBER REINFORCED
COMPOSITES
In the past, the majority of the helicopter structure was made out
of metal. However, the use of fiber reinforced composite structure
is increasing and will continue to do so. The following are
characteristic differences between the composite and metallic
structures that need to be considered when evaluating damage to and
developing repairs for fiber reinforced composite structure.
1.0 DAMAGE
Fiber reinforced materials (composites) do not exhibit plastic
deformation prior to failure. Even if they resist impact loads,
internal damage such as matrix cracking or delamination may be
induced. Propagation of damage in composite structure occurs by
growth of a delamination rather than by through thickness cracks,
as normally observed with metals. Inspection methods shall be
appropriate to detect these damage.
2.0 LIGHTNING STRIKE PROTECTION
In most cases, carbon fiber reinforced composites require a
protective layer against potential lightning strikes. The most
common form of protective layer is a copper wire mesh that has been
applied, using adhesive, to the outboard side of the structure.
When a part with this protection has been damaged, the lightning
strike layer must be restored during the repair using instructions
detailed in paragraph 4-3-4.
3.0 MOISTURE ABSORPTION
Over time, fiber reinforced composite materials will absorb
moisture from the surrounding environment (typically about 1 to 2%
of their original weight). Removal of entrapped moisture, by drying
the structure, is critical prior to carrying out any repair that
requires a cure or post-cure at or above 200°F (93.3°C). Failure to
dry the laminate before performing a repair can lead to a reduction
in the strength of the repair or to additional damage in the
original panel. In a solid laminate, the moisture will tend to
migrate into the bondline adhesive and reduce its
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strength. In a honeycomb panel, the water entrapped in core cells
may turn to steam during the cure cycle and ‘blow the repair
doubler off’ or damage the honeycomb core due to the buildup of
internal pressure.
4.0 MATERIAL COMPATIBILITY
When carbon fiber composite structure is in contact with various
metals (aluminum and some steels), corrosion to the adjacent metal
can result (for a complete list of compatible materials, refer to
Appendix A-7). Whenever carbon fiber composite structure is
required to contact a susceptible material, a glass ply serving as
a barrier is used as the last ply. In addition to the glass barrier
ply and the epoxy polyamide primer, the faying surfaces (mating
surfaces) are coated with a sealant before the joint is assembled.
The material compatibility requirements also restrict the choice of
fastener type to be used. If no clear direction is provided in the
Structural Repair Manual (SRM), please contact Product Support
Engineering for the specification of compatible fasteners.
5.0 INSTALLATION OF MECHANICAL FASTENERS
The following are guidelines for mechanically fastened joints in
carbon fiber laminates:
5.1 The part has to be properly braced and a backup material (piece
of hard wood or equivalent material) has to be used to prevent
fiber breakout on the backside of the part during drilling of
fastener holes (paragraph 4-2-5).
5.2 Edge distance requirement is greater for composites than for
metals. Typically 2.5D is used for both protruding head and
countersunk fasteners (Chapter 3, paragraph 3-3-1).
5.3 Only fasteners that are acceptable for use in contact with
composite materials may be used. Refer to Appendix A-6 for a list
of approved fasteners for use with carbon fiber reinforced
composites.
5.4 Blind rivets are only allowed if there is a metal backing on
the tail side to avoid damaging the fibers.
5.5 Interference fit fasteners such as blind bolts and Hi-Lites are
to be used.
5.6 The use of fasteners that expand to fill the hole shall be
avoided as much as possible to prevent fiber breakout and
delamination at the periphery of the hole during the installation
of these fasteners.
5.7 In some cases it is permissible to squeeze solid rivets. Never
use a rivet gun and bucking bar to install rivets as the repetitive
impact may damage the composite material.
5.8 Wet installation of fasteners using a corrosion inhibitor
(primer or sealant) is generally required.
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4-1-2. WET LAYUP REPAIR FLOW CHART
Figure 4-1 provides a flow chart for the wet layup repair process.
The major steps are as follows:
WARNING
REFER TO YOUR MODEL-SPECIFIC STRUCTURAL REPAIR MANUAL FOR REPAIRS
OF RESTRICTED AREAS.
1.0 DAMAGE INSPECTION (Chapter 2, paragraph 2-13)
2.0 DAMAGE CLASSIFICATION AND EXTENT (Chapter 2, paragraph
2-13-1)
3.0 SELECTION OF TYPICAL REPAIR PROCEDURE
3.1 Glass fiber reinforced composites (paragraph 4-4).
3.2 Carbon fiber reinforced composites (paragraph 4-5).
4.0 VERIFICATION OF RESTRICTIONS
4.1 Specific Restrictions: Per typical repair procedure in this
manual and per part specific requirements in the model-specific
SRM.
4.2 General Restrictions: Per typical repair procedure in this
manual.
5.0 PERFORM REPAIR
5.1 Using process sheets (paragraph 3-2, paragraph 4-2 and
paragraph 4-3).
5.2 Using typical repair procedures
5.2.1 Glass fiber reinforced composites (paragraph 4-4).
5.2.2 Carbon fiber reinforced composites (paragraph 4-5).
6.0 INSPECTION OF REPAIR
BHT-ALL-SRM TC/FAA APPROVED
DAMAGE INSPECTION
– Section 2-13
DAMAGE CLASSIFICATION
– Section 2-13-1 – Type – Location – Size – Material (Glass or
Carbon) – Construction Type (monolithic or Sandwich)
SELECTION OF TYPICAL REPAIR PROCEDURES
– For Glass Composites: Section 4-4 – For Carbon Composites:
Section 4-5
VERIFICATION OF RESTRICTIONS
SUITABLE REPAIR FOUND?
YES NO
PERFORM REPAIR
– Use of Specific Repair Procedures: Model-Specific SRM – Use of
Typical Repair Procedures: Section 4-4 and Section 4-5 – Use of
Process Sheet(s): Section 4-2 and Section 4-3
ASSISTANCE BY PRODUCT SUPPORT ENGINEERING (PSE)
– Damage Evaluation
YES NO
TC/FAA APPROVED BHT-ALL-SRM
4-1-3. REPAIR SELECTION GUIDE
As this chapter contains repairs for various types of structures
with various repair methods based on damage type, size, and
location, a repair selection guide is included to facilitate the
identification of the appropriate repair. The following tables list
the different generic repairs given in this chapter. These repairs
are generic and apply to unrestricted repairable areas of the
helicopters. Refer to the model-specific Structural Repair Manual
(SRM) and to Chapter 1, paragraph 1-20 for a list of restricted
areas. If damage is in a restricted area or beyond the limitations
given in this chapter, the model-specific SRM may have a suitable
specific repair. If the damage cannot be repaired using
instructions given in this manual or the model-specific SRM,
contact Product Support Engineering for a disposition. Refer to
Chapter 2, paragraph 2-21 for the procedure to follow to request an
approved structural field repair from Product Support
Engineering.
The repair guide in this chapter is divided in several tables. Each
table provides a selection guide based on the type of structure or
material being repaired: Monolithic glass structures, glass
facing(s) on honeycomb panels, monolithic carbon fiber structures,
or carbon fiber facing(s) on honeycomb panels. Also included are
tables for rivet pattern discrepancies.
Table 4-1. Repair Selection Guide for Rivet Pattern
Discrepancy
Maximum Damage
Pitch(1) Repair Para. Appl.
Monolithic Glass or Carbon
50 for 1st oversize
25 for 2nd oversize
Low edge distance 1/5
50 for 1st oversize
25 for 2nd oversize
4D 4D
3-4-3 B
1) Top value is for protruding head fasteners and bottom value is
for flush head fasteners.
Table 4-2. Repair Selection Guide for Bonded Panels With Aluminum
Face Sheets and Glass Edging
Maximum Damage
Para. Appl.
Punctures Affecting Core Bevel
4-4-13 -
1) All dimensions are in inches. Values between parentheses are in
millimeters (mm).
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Table 4-3. Repair Selection Guide for Glass or Carbon Monolithic
Part
Maximum Damage
r Para.
3 0.050
(0.323) 1/2
3 0.050
(0.323) Full
1 2
Fasteners + ED
1.5 (38)
4-4-4 B
3 0.050
(0.323) 1/2
3 0.050
(0.323) Full
1 2
Fasteners + ED
4-5-4 B
1) All dimensions are in square inches. Values between parentheses
are square centimeter (cm2).
2) All dimensions are in inches. Values between parentheses are in
millimeters (mm).
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Maximum Damage
Glass
3 0.050
thickness
Deep Smooth Dent, Sharp Dent or Puncture Through a Skin
3 0.050
3 0.050
1 2 Fasteners
(12.7) 4-4-12 B
1) All dimensions are in square inches. Values between parentheses
are square centimeter (cm2).
2) All dimensions are in inches. Values between parentheses are in
millimeters (mm).
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3 0.050
0.25 (6.4)
4-5-6 A
per Model- specific SRM
Deep Smooth Dent, Sharp Dent or Puncture Through a Skin
3 0.050
3 0.050
Maximum Damage
Size(1) Depth(2) Length(2) Width(2) Repair Para. Appl.
1) All dimensions are in square inches. Values between parentheses
are square centimeter (cm2).
2) All dimensions are in inches. Values between parentheses are in
millimeters (mm).
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1 2 Fasteners
Maximum Damage
Size(1) Depth(2) Length(2) Width(2) Repair Para. Appl.
1) All dimensions are in square inches. Values between parentheses
are square centimeter (cm2).
2) All dimensions are in inches. Values between parentheses are in
millimeters (mm).
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4-2. COMMON PROCEDURES FOR GLASS OR CARBON FIBER REINFORCED
COMPOSITE PARTS
4-2-1. GLASS OR CARBON FIBER REINFORCED COMPOSITE PARTS —
GENERAL
CAUTION
NOTE THAT BOTH THE GENERAL AND SPECIFIC RESTRICTIONS MUST BE MET
BEFORE PERFORMING A COMPOSITE REPAIR. IF THE RESTRICTIONS CANNOT BE
MET, CONSULT PRODUCT SUPPORT ENGINEERING. IN CASE OF CONFLICT
BETWEEN A GENERAL AND A SPECIFIC RESTRICTION, THE SPECIFIC
RESTRICTION OVERRULES THE GENERAL RESTRICTION.
NEVER USE MEK, ACETONE OR ALIPHATIC NAPHTHA INSTEAD OF THE
INDICATED ALCOHOLS ON GLASS OR CARBON FIBER REINFORCED COMPOSITES.
MEK IS ONLY TO BE USED FOR CLEANING TOOLS.
WEAR APPROPRIATE SAFETY EQUIPMENT (GLOVES, GOWNS, RESPIRATORS,
ETC.) WHEN HANDLING OR WORKING WITH MATERIALS AND MAKING REPAIRS.
CONSULT MATERIAL SAFETY DATA SHEETS (MSDS) FOR POTENTIAL HAZARDS
AND FOLLOW ALL APPLICABLE SAFETY PROCEDURES.
NOTE
Determine material and tooling requirements and ensure materials
are at hand before proceeding with any repair.
• Consumable materials and standards: Materials needed to
accomplish a particular repair are listed in the “REQUIRED” section
of each repair procedure. Each item is accompanied by a description
of the material and a numerical code (C-xxx). This code references
a consumable item, which is further described in Chapter 13 of the
Standard Practices Manual (BHT-ALL-SPM).
• Required number of plies of repair doubler: Unless otherwise
indicated in a repair procedure, the number of plies of a glass
fiber repair doubler and their orientation shall be as defined in
paragraph 4-3-3. The number of plies of a carbon fiber repair
doubler and their orientation may vary and is specific to the
region being repaired, and shall be provided by Product Support
Engineering unless explicitly specified in the repair.
• Stop drilling: Although stop drilling relieves the stresses in
the extremity of a crack in sheet metal parts, Bell Helicopter
Textron does not permit stop drilling cracks in fiber reinforced
composite structures.
• Finishing: All repairs shall be sealed against moisture intrusion
and then finished in accordance with the original finish
specifications.
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4-2-2. APPROVED PROCESSES (PROCESS SHEETS)
This section describes the typical processes and repair procedures
involving glass or carbon fiber reinforced composite parts.
Composite repairs are similar in nature and may be applicable to
various parts of the structure. However, the restrictions on
repairability may vary from one composite part to another.
Therefore, the general and specific restrictions must be verified
for each composite part before using a typical repair
procedure.
This section describes the processes used in the repair of glass or
carbon fiber reinforced composites. These process sheets cover
topics such as machining, cleaning and surface preparation prior to
bonding. Repair procedures throughout this manual and the
model-specific Structural Repair Manual (SRM) make reference to the
applicable process sheets in the “REQUIRED” section for each
repair.
The applicable processes that are covered in this section are
listed in Table 4-5:
Table 4-5. Process Sheets
4-2-3 Page 4-16
Provides a method for sanding glass or carbon fiber
composites.
4-2-4 Page 4-18
Provides a method for Cutting/Routing Glass or Carbon fiber
composites.
4-2-5 Page 4-21
Provides a method for drilling glass or carbon fiber
composites.
4-2-6 Page 4-27
Removal of Paint, Primer and Sanding Surfacer on Glass or Carbon
Fiber Composites.
Provides a method of removing paint, primer and sanding surfacer
from composite parts.
4-2-7 Page 4-30
Removal of Honeycomb Core in Glass or Carbon Fiber Honeycomb
Panels.
Provides a method for removing honeycomb core in glass or carbon
fiber honeycomb panels.
4-2-8 Page 4-32
Removal of Copper Wire Mesh from Glass or Carbon Fiber
Composites.
Provides a method for removing honeycomb core in glass or carbon
fiber honeycomb panels.
4-2-9 Page 4-34
Surface Preparation for Bonding on Glass or Carbon Fiber
Composites.
Provides a method for removing copper wire mesh layer from glass or
carbon fiber composites.
4-2-10 Page 4-36
Drying Composites Parts Prior to Bonding.
Provides a method for drying glass or carbon fiber composite parts
prior to bonding.
4-2-11 Page 4-42
Finish Process Following a Composite Repair.
Provides a method for finishing a glass or carbon fiber composite
repair.
4-2-12 Page 4-46
Preparation of Molding Tool for Composite Repair.
Provides a method for fabricating molds to be used in repairs of
glass or carbon fiber reinforced composites.
4-2-13 Page 4-51
Preparation of Tooling Surfaces.
Provides a method for cleaning the surface of tools in preparation
for bonding wet layup repairs of glass or carbon fiber reinforced
composites.
4-2-14 Page 4-53
Preparation of Filled Epoxy Resin.
Provides a method for preparing filler material by mixing epoxy
resin with chopped glass fiber.
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4-2-3. SANDING GLASS OR CARBON FIBER COMPOSITES
This process sheet describes the different equipment and procedures
used to sand glass or carbon fiber reinforced composites.
CAUTION
WHEN SANDING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK, AND LONG
SLEEVED PROTECTIVE CLOTHING. GLASS OR CARBON FIBER DUST CAN CAUSE
IRRITATION TO EYES, SKIN, AND LUNGS.
1.0 EQUIPMENT (Use as required)
1.1 Pneumatic high speed sander capable of 20,000 RPM with 3M Roloc
sanding disc, 2.0 inches (51 mm) maximum diameter, 80 to 240 grit
silicon carbide abrasive.
1.2 Pneumatic vibrating sander capable of approximately 10,000 RPM
with self-adhesive sanding disc, 5.0 inches (127 mm) maximum
diameter, 220 to 400 grit silicon carbide abrasive.
1.3 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 80 to 400 grit.
3.0 PROCEDURE
3.1 Determine type of sanding to be performed (i.e., rough cuts,
edge smoothing, or surface finish).
CAUTION
DO NOT OVERHEAT COMPOSITES DURING MATERIAL SANDING BY USING ORBITAL
MOTION.
DO NOT USE WATER SOLUBLE COOLANT ON HONEYCOMB OR SYNTACTIC CORE
PANELS.
DO NOT USE ABRASIVE PAPER THAT HAS BEEN USED TO SAND METAL.
ALWAYS COVER WIRE BUNDLES TO PREVENT ENTRAPMENT OF
PARTICULES.
3.2 For rough cuts, use high speed sander specified in 1.1 with
appropriate grit abrasive. Finish by hand sanding with 400 grit
abrasive paper specified in 2.1.
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3.3 For edge smoothing, use abrasive paper of appropriate grit
specified in 2.1. Sand by hand in direction along long dimension,
not across edge of part, whenever possible. When sanding across
edge of part, use proper backups to prevent fiber breakout or
splintering.
3.4 For surface finish, use vibrating sander specified in 2.1 for
large surface or sand by hand using abrasive paper of fine grit.
Generally, finish with 400 grit abrasive specified in 2.1.
14 DEC 2010 4-17ECCN EAR99
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4-2-4. CUTTING/ROUTING GLASS OR CARBON FIBER COMPOSITES
This process sheet describes the different equipment and procedures
used to cut/rout out damaged area in glass or carbon fiber
reinforced composites.
CAUTION
WHEN CUTTING/ROUTING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK,
AND LONG SLEEVED PROTECTIVE CLOTHING. GLASS OR CARBON FIBER DUST
CAN CAUSE IRRITATION TO EYES, SKIN, AND LUNGS.
1.0 EQUIPMENT (Use as required)
1.1 Pneumatic high speed cutter capable of 20,000 RPM.
1.1.1 Reinforced cutting disc, 3 inch (76.2 mm) diameter, 60 to 80
grit abrasive, or
1.1.2 Diamond cutting wheel of appropriate size, 40 to 80 grit
abrasive.
1.2 Pneumatic router capable of 12,000 to 23,000 RPM for parts made
entirely of composite or 12,000 RPM maximum for composite parts
with metallic components or syntactic core embedded in the
laminate.
1.2.1 Diamond plated or solid carbide router bits, 0.1875 inch
(4.763 mm) diameter, 40 to 80 grit.
1.3 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Air vacuum or liquid coolant (lubricant (C-569)).
2.2 Process Sheet(s): Sanding Glass or Carbon Fiber Composites
(paragraph 4-2-3)
3.0 PROCEDURE
CAUTION
DO NOT MARK THE COMPOSITE SURFACE USING ANY METHOD THAT WILL INDENT
OR OTHERWISE DEFORM THE SURFACE.
ALWAYS COVER WIRE BUNDLES TO PREVENT ENTRAPMENT OF
PARTICULES.
3.1 Locate damaged area and mark section to be removed. This
section shall be at least 1/32 inch (0.79 mm) greater than existing
damage in all directions. For honeycomb panels, remove only damaged
skin(s), unless otherwise specified in applicable repair.
4-18 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
CAUTION
ENSURE THAT NO OTHER PART INTERFERES WITH SECTION TO BE CUT OR
ROUTED OUT.
DO NOT OVERHEAT COMPOSITES DURING MATERIAL CUTTING/ROUTING.
DO NOT USE WATER SOLUBLE COOLANT ON HONEYCOMB OR SYNTACTIC CORE
PANELS.
FOR SYNTACTIC CORE PANELS, REDUCE ROUTER SPEED TO 12,500 RPM.
DO NOT USE SAME ROUTING BIT FOR METAL AND COMPOSITE PARTS.
ROUTERS FITTED WITH DUST EXTRACTION ARE RECOMMENDED.
A TEMPLATE IS RECOMMENDED TO GUIDE ROUTER TO OBTAIN REQUIRED
SHAPE.
3.2 Crack damage: Rout out crack using router specified in 1.2
equipped with router bit specified in 1.2.1, and using a maximum
feed rate of 12.0 inches (305 mm) per minute.
3.3 All types of damage (including excessive crack damage): Cut out
damaged area using high speed cutter specified in 1.1 or router
specified in 1.2 equipped with router bit specified in 1.2.1.
Cutout radius to be a minimum of 0.50 inch (12.7 mm). Maintain a
maximum feed rate of 12.0 inches (305 mm) per minute.
3.4 If applicable, sand edges to required dimensions using
instructions detailed in paragraph 4-2-3.
14 DEC 2010 4-19ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
TC/FAA APPROVED BHT-ALL-SRM
4-2-5. DRILLING GLASS OR CARBON FIBER COMPOSITES
This process sheet describes the different equipment and procedures
used to drill glass or carbon fiber reinforced composites.
CAUTION
WHEN DRILLING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK, AND LONG
SLEEVED PROTECTIVE CLOTHING. GLASS OR CARBON FIBER DUST CAN CAUSE
IRRITATION TO EYES, SKIN, AND LUNGS.
1.0 EQUIPMENT (Use as required)
1.1 Pneumatic high speed drill or drill press capable of the RPM
specified below.
1.1.1 C2 carbide drill bits of appropriate diameter.
1.1.2 C2 carbide reamer of appropriate diameter.
1.1.3 C2 carbide dreamer (drill/reamer) of appropriate
diameter.
1.2 Locally fabricated metallic fixture, workaid or jig.
1.3 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 180 to 240
grit.
2.2 Lubricant (C-569).
2.3 Backup material (of same contour or shape), particle board or
phenolic (other than glass base) at least 3/16 inch (4.76 mm)
thick.
2.4 Wet layup adhesive (C-363).
2.5 Process Sheet(s): Preparing and Mixing Two-part Epoxy Resin by
Weight (paragraph 3-2-25) Sanding Glass or Carbon Fiber Composites
(paragraph 4-2-3) Curing Process for Epoxy Resin (paragraph
4-3-6)
14 DEC 2010 4-21ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
ENSURE THAT DRILLING DEPTH WILL NOT AFFECT OTHER PARTS.
ALWAYS COVER WIRE BUNDLES TO PREVENT ENTRAPMENT OF
PARTICULES.
3.1 Locate center of hole to be drilled and use metal fixture,
workaid, or jig to start hole.
3.2 To prevent splintering and delamination, use a backup workaid
such as a hard wood block against tool exit surface. Backup workaid
shall be held in intimate contact with workpiece at all time during
drilling process.
CAUTION
DO NOT OVERHEAT COMPOSITES DURING MATERIAL DRILLING PROCESS.
DO NOT USE LUBRICANT SPECIFIED IN 2.2 ON HONEYCOMB OR SYNTACTIC
CORE PANELS.
NOTE
Drill bits shall be changed often so that a sharp cutting edge is
maintained at all time. Life of sharp carbide drill bits are
limited to approximately 30 holes per bit. Bits that exceed this
limit must be either resharpened, or discarded and replaced.
Step 3.3 and step 3.4 can be performed in a single operation using
a dreamer (combined drill and reamer), and using a speed between
2400 and 2700 RPM.
3.3 Drill hole using a speed between 1600 and 2700 RPM. To prevent
fiber breakout, reduce feed rate near end of cut. For solid
laminate thicker than 1/16 inch (1.59 mm) use lubricant specified
in 2.2.
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TC/FAA APPROVED BHT-ALL-SRM
DO NOT OVERHEAT COMPOSITES DURING MATERIAL REAMING.
DO NOT USE LUBRICANT SPECIFIED IN 2.2 ON HONEYCOMB OR SYNTACTIC
CORE PANELS.
NOTE
Reamers are limited to approximately 10 holes. Reamers that exceed
this limit must be discarded and replaced.
3.4 Ream hole using a hand drill operating at a maximum speed of
250 RPM and a feed rate of 4.0 to 6.0 inches (102 to 152 mm) per
minute.
3.5 If required, countersink hole. Countersink and counter bore
shall be concentric within 0.003 inch (0.08 mm) and parallel to the
fastener hole direction.
3.6 Inspect hole for fiber breakouts, delamination and chipping.
Damage shall not exceed the following limits:
3.6.1 Minor/negligible damage is allowed on exit side of drill
only.
3.6.2 Acceptance criteria are provided in Table 4-6. Both
negligible and repairable limits are provided in second and third
column respectively as are repair procedures.
3.7 Deburr, if necessary, by hand sanding using silicon carbide
abrasive paper specified in 2.1. Refer to instructions detailed in
paragraph 4-2-3.
Table 4-6. Drilling Damage Limits for Protruding Head Fastener
Holes
MATERIAL TYPE
LIMIT(1)(2)(4)(5)(6) REPAIR PROCEDURE
Fabric Either • 0.16 times hole diameter, or • 0.03 inch (0.76
mm)
0.5 times hole diameter in any direction
• Prepare wet layup adhesive specified in 2.4 according to
paragraph 3-2-5. • Coat with wet layup adhesive. • Cure according
to paragraph 4-3-6. • Redrill hole as necessary • Sand smooth
according to paragraph 4-2-3. • No fiber damage allowed during
corrective action.
Unidirectional Tape
Either • 0.50 times hole diameter, or • 0.10 inch (2.54 mm)
whichever is smaller.
1.0 inch (25.4 mm) in length, no more than 90° of hole
periphery
NOTES:
2. Distance measured radially from edge of the hole.
3. No rework required.
4. Rework in accordance with procedures provided in fourth
column.
5. Damage must not extend within 0.5 inch (12.7 mm) of another
fastener hole.
6. Breakout from any two separate fastener holes shall not
meet.
14 DEC 2010 4-23ECCN EAR99
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4-24 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
Table 4-7. Drilling Damage Limits for Countersink Side of
Holes
MATERIAL TYPE
LIMIT(1)(2)(4)(5)(6) REPAIR PROCEDURE
Fabric Either • 0.16 times hole diameter, or • 0.030 inch (0.76
mm)
0.25 inch (6.4 mm) in length, no more than 60° of hole
periphery
• Prepare wet layup adhesive specified in 2.4 according to
paragraph 3-2-25. • Coat with wet layup adhesive. • Cure according
to paragraph 4-3-6. • Redrill hole as necessary. • Sand smooth
according to paragraph 4-2-3. • No fiber damage allowed during
corrective action.
Unidirectional Tape
Either • 0.50 times hole diameter, or • 0.100 inch (2.54 mm)
whichever is smaller.
NOTES:
2. Distance measured radially from edge of the hole.
3. No rework required.
4. Rework in accordance with procedures provided in fourth
column.
5. Damage must not extend within 0.50 inch (12.7 mm) of countersink
of another fastener hole.
6. Breakout from any two separate fastener holes shall not
meet.
14 DEC 2010 4-25ECCN EAR99
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4-2-6. REMOVAL OF PAINT, PRIMER, AND SANDING SURFACER ON GLASS OR
CARBON FIBER COMPOSITES
This process sheet describes the different equipment and procedures
used to remove paint, primer, and sanding surfacer on glass or
carbon fiber reinforced composites.
CAUTION
CHEMICAL PAINT STRIPPERS ARE NOT TO BE USED TO REMOVE PAINT
FINISHES ON BONDED PARTS OR PANELS. CHEMICAL PAINT STRIPPERS MAY
CONTRIBUTE TO CONTAMINATION OF CORE, DETERIORATE ADHESIVE BOND LINE
AND GLASS OR CARBON FIBER COMPOSITES. THIS MAY ULTIMATELY CAUSE
PART OR PANEL TO BE REJECTED.
MEK, ACETONE, TRICHLOROETHYLENE, AND VAPOR DEGREASERS ARE NOT TO BE
USED TO CLEAN OR STRIP SURFACES ADJACENT TO A DAMAGED AREA FOR THE
SAME REASONS.
METHYL ALCOHOL, ETHYL ALCOHOL AND ISOPROPYL ALCOHOL ARE ACCEPTABLE
SOLVENTS FOR REMOVAL OF PAINT FROM SOLID LAMINATE COMPOSITE PARTS
OR BONDED PANEL WITH COMPOSITE SKINS. HOWEVER EXCESSIVE APPLICATION
OF ALCOHOL MAY AFFECT ADHESIVES USED IN A BOND. IT IS PREFERABLE TO
WIPE SURFACE TO BE STRIPPED USING A MOISTENED CHEESECLOTH RATHER
THAN BY SOAKING.
BEFORE HANDLING A SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS.
KEEP PRODUCT AND ITS VAPORS AWAY FROM HEAT, SPARKS, AND FLAME.
DURING APPLICATION AND UNTIL VAPORS HAVE DISSIPATED, AVOID USING
SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES,
ETC. AVOID PROLONGED BREATHING OF VAPORS AND REPEATED CONTACT WITH
SKIN.
1.0 EQUIPMENT
1.1 Pneumatic high speed sander or drill equipped with a variable
pressure regulator and 3M Roloc white bristle brush, 3.0 inches (76
mm) maximum diameter, or material specified in 2.1 or 2.2.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 80 to 400 grit,
or
2.2 Nylon web abrasive pad (C-407), maroon and green.
2.3 Cleaner: ethyl alcohol (C-339), isopropyl alcohol (C-385), or
toluene (C-306).
2.4 Cheesecloth (C-486).
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3.0 PROCEDURE
3.1 Mask around area to be reworked using plastic film or Kraft
paper specified in 2.6, and masking tape specified in 2.5. If a
repair doubler is required, masked area shall be 1.00 inch (25.4
mm) further away from periphery of repair doubler.
CAUTION
DO NOT WET SAND.
DO NOT SAND INTO GLASS OR CARBON FIBERS. PAY SPECIAL ATTENTION NOT
TO DAMAGE FIBERS WHEN MULTIPLE SANDING OPERATIONS ARE TO BE
PERFORMED AS EACH STEP MAY SLIGHTLY DAMAGE FIBERS.
DO NOT OVERHEAT COMPOSITES DURING MATERIAL SANDING. USE OF ORBITAL
MOTION IS PREFERABLE.
DO NOT USE ABRASIVE PAPER THAT HAS BEEN USED TO SAND METAL.
DO NOT USE PNEUMATIC TOOL, OR ABRASIVE CLOTH OR PAPER ON THIN
COMPOSITE SKIN OF BONDED PANEL WHERE HEXAGONAL SHAPE OF CORE CELLS
ARE VISIBLE. USE NYLON WEB ABRASIVE PAD INSTEAD.
ALWAYS COVER WIRE BUNDLES TO PREVENT ENTRAPMENT OF
PARTICULES.
WHEN SANDING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK, AND LONG
SLEEVED PROTECTIVE CLOTHING. GLASS OR CARBON FIBER DUST CAN CAUSE
IRRITATION TO EYES, SKIN, AND LUNGS.
NOTE
Some primer residue may remain in recesses of peel ply and vacuum
bag impression on surface of laminate. This condition is
acceptable, but all efforts shall be made to keep it to a minimum
while taking care not to damage glass or carbon fibers.
3.2 Determine type of sanding to be performed (i.e., sanding with
pneumatic tool, by hand with abrasive cloth/paper, or with nylon
web abrasive pad).
3.3 For hand sanding with abrasive paper specified in 2.1, remove
paint, primer and sanding surfacer using orbital motion. When
primer is or becomes visible, use 180 to 240 grit or finer abrasive
paper to continue sanding operation. Vacuum area frequently to
reduce paint/primer residue buildup on abrasive paper.
3.4 For hand sanding with green or maroon nylon web abrasive pads
specified in 2.2, remove paint, primer, and sanding surfacer using
orbital motion. When primer is or becomes visible, use green pads
only to continue sanding operation. Vacuum area frequently to
reduce paint/primer residue buildup inside nylon web abrasive
pad.
4-28 14 DEC 2010 ECCN EAR99
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3.5 For pneumatic tool sanding using equipment specified in 1.1,
remove paint, primer, and sanding surfacer using orbital motion.
Always begin with lowest speed setting to prevent damaging the
fibers, and slowly increase speed until bristle brush stops
gripping. Maintain bristle brush perpendicular to surface to be
reworked. Vacuum area frequently to reduce paint/primer residue
buildup inside brush.
3.6 Accomplish surface cleaning by wiping with a clean cheesecloth
moistened with cleaner specified in 2.3. Change cheesecloth often.
Repeat operation until all evidence of residue is removed and wipe
dry using a clean cheesecloth.
14 DEC 2010 4-29ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
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4-2-7. REMOVAL OF HONEYCOMB CORE IN GLASS OR CARBON FIBER HONEYCOMB
PANELS
This process sheet describes the different equipment and procedures
used to remove honeycomb core from glass or carbon fiber reinforced
honeycomb panels.
CAUTION
WHEN REPAIRING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK, AND LONG
SLEEVED PROTECTIVE CLOTHING. NOMEX CORE DUST CAN CAUSE IRRITATION
TO EYES, SKIN, AND LUNGS.
1.0 EQUIPMENT (Use as required)
1.1 Pneumatic router capable of 12,000 to 23,000 RPM.
1.2 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED
2.1 Diamond plated or solid carbide router bits.
2.2 Process Sheet(s): Sanding Glass or Carbon Fiber Composites
(paragraph 4-2-3) Cutting/Routing Glass or Carbon Fiber Composites
(paragraph 4-2-4)
3.0 PROCEDURE
3.1 If not already accomplished, remove damaged section of skin(s)
using instructions detailed in paragraph 4-2-4.
CAUTION
UNLESS A SECTION OF OPPOSITE SKIN MUST ALSO BE REMOVED, ENSURE THAT
OPPOSITE SKIN WILL NOT BE DAMAGED.
3.2 Adjust depth of the cut as required.
CAUTION
DO NOT OVERHEAT COMPOSITES DURING CORE REMOVAL.
DO NOT USE WATER SOLUBLE COOLANT ON HONEYCOMB OR SYNTACTIC CORE
PANELS.
3.3 Rout out section of core to be removed using router specified
in 1.1 equipped with router bit specified in 2.1, and using a speed
between 12,000 and 23,000 RPM.
4-30 14 DEC 2010 ECCN EAR99
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3.4 Sand edges of skin(s) around cavity using instructions detailed
in paragraph 4-2-3.
14 DEC 2010 4-31ECCN EAR99
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4-2-8. REMOVAL OF COPPER WIRE MESH FROM GLASS OR CARBON FIBER
COMPOSITES
This process sheet describes the different equipment and procedures
used to remove the copper wire mesh layer from glass or carbon
fiber reinforced parts.
CAUTION
WHEN SANDING COMPOSITES, WEAR SAFETY GLASSES, DUST MASK, AND LONG
SLEEVED PROTECTIVE CLOTHING. GLASS OR CARBON FIBER DUST CAN CAUSE
IRRITATION TO EYES, SKIN, AND LUNGS.
1.0 EQUIPMENT
1.1 None.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 400 grit.
2.2 Process Sheet(s): Sanding Glass or Carbon Fiber Composites
(paragraph 4-2-3) Removal of Paint, Primer and Sanding Surfacer on
Glass or Carbon Fiber (paragraph 4-2-6)
3.0 PROCEDURE
3.1 If not already accomplished, remove paint, primer and sanding
surfacer from affected area using instructions detailed in
(paragraph 4-2-6).
CAUTION
DO NOT OVERHEAT COMPOSITES DURING MATERIAL SANDING.
DO NOT USE ABRASIVE PAPER THAT HAS BEEN USED TO SAND METAL.
3.2 Carefully sand off adhesive to expose copper wire mesh taking
care not to damage glass or carbon fibers.
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CAUTION
CARE MUST BE TAKEN TO PREVENT DAMAGING GLASS OR CARBON FIBERS WHILE
CUTTING COPPER WIRE MESH.
NOTE
Do not remove more copper wire mesh than required by applicable
repair.
3.3 Use a small knife blade to carefully cut copper wire mesh along
periphery of section to be removed.
3.4 Slowly peel off section of copper wire mesh to be
removed.
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4-2-9. SURFACE PREPARATION FOR BONDING ON GLASS OR CARBON FIBER
COMPOSITES
This process sheet describes the different equipment and procedures
used to prepare glass or carbon fiber reinforced composites
surfaces for bonding.
NOTE
In order to provide optimal conditions for bonding, it is
recommended that delays between surface preparation detailed in
this process sheet and bonding of repair plies be kept to a
minimum.
1.0 EQUIPMENT
1.1 None.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 100 grit or finer,
240 grit preferred.
2.2 Cleaner: ethyl alcohol (C-339), isopropyl alcohol (C-385), or
toluene (C-306).
2.3 Kraft paper (C-254).
2.4 Masking tape (C-426).
2.5 Cheesecloth (C-486).
2.6 Process Sheet(s): Sanding Glass or Carbon Fiber Composites
(paragraph 4-2-3) Removal of Paint, Primer, and Sanding Surfacer on
Glass or Carbon Fiber (paragraph 4-2-6) Removal of Copper Wire Mesh
from Glass or Carbon Fiber Composites (paragraph 4-2-8)
3.0 PROCEDURE
NOTE
When performing structural repair, surfacing film and copper wire
mesh must be removed to allow bonding of repair plies to existing
structural plies.
3.1 If not already accomplished, remove paint, primer, sanding
surfacer, and surfacing film (if required) using instructions
detailed in paragraph 4-2-6.
3.2 If required and not already accomplished, remove copper wire
mesh using instructions detailed in paragraph 4-2-8.
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CAUTION
BEFORE HANDLING A SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS.
KEEP PRODUCT AND ITS VAPORS AWAY FROM HEAT, SPARKS, AND FLAME.
DURING APPLICATION AND UNTIL VAPORS HAVE DISSIPATED, AVOID USING
SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES,
ETC. AVOID PROLONGED BREATHING OF VAPORS AND REPEATED CONTACT WITH
SKIN.
3.3 Wipe surface to be bonded with a clean cheesecloth moistened
with cleaner specified in 2.2.
CAUTION
GLASS OR CARBON FIBERS MUST BE SANDED UNTIL FIBERS ARE SLIGHTLY
ABRADED.
DO NOT SAND INTO GLASS OR CARBON FIBER PLIES.
3.4 Sand surface(s) to be bonded with silicon carbide abrasive
paper specified in 2.1. Refer to instructions detailed in paragraph
4-2-3.
3.5 Wipe surface with a clean cheesecloth moistened with cleaner
specified in 2.2. Change cheesecloth often. Repeat operation until
all evidence of residue is removed and wipe dry using a clean
cheesecloth.
3.6 Allow surface to air dry for at least 30 minutes before
bonding.
3.7 Use clean Kraft paper specified in 2.3 and tape specified in
2.4 to protect surface from contamination until ready for
bonding.
NOTE
If fibers have been exposed during a previous operation and this
process sheet must be performed again, substitute abrasive paper
specified in 2.1 with green web abrasive pad (C-407).
Surface must be bonded within 72 hours following abrading and
alcohol wipe. If not, surface preparation using this process sheet
must be performed again.
14 DEC 2010 4-35ECCN EAR99
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4-2-10. DRYING COMPOSITE PARTS PRIOR TO BONDING
This process sheet describes the different equipment and procedures
used to dry solid laminates or honeycomb panels.
NOTE
Laminates or honeycomb panels made with organic matrix material
such as epoxy must be dried before performing any repair that
requires a cure or a post-cure at or above 200°F (93.3°C). If
moisture is not removed from laminate or honeycomb panel, there is
a possibility of delamination or skin to core disbond due to steam
buildup.
APPLICATION A: DRYING OF SOLID LAMINATES
1.0 EQUIPMENT
1.1 Temperature monitoring equipment with a minimum of two
thermocouples.
1.2 Device capable of applying heat such as:
1.2.1 An oven (if part can be removed from helicopter).
1.2.2 A Bell Helicopter Textron approved hot-bonding unit, or
equivalent, in conjunction with a heat blanket capable of
maintaining a minimum of 250°F (121.1°C).
1.2.3 Heat lamps.
1.2.4 Heat guns.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Bagging film (C-257) or bagging film (C-564), if
required.
2.2 Pressure sensitive masking tape (C-260), if required.
2.3 Breather felt (C-258), if required.
2.4 Glass fabric (C-404), if required.
2.5 Vacuum bag sealant tape (C-259) or sealant tape (C-566), if
required.
2.6 Pressure sensitive masking tape (C-260) or tape (C-462)
respectively, if required.
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NOTE
Drying time depends on material type, part thickness, drying
temperature, and relative humidity inside part. Ensure that part is
dried before performing repair.
If repair is to be cured or post-cured in an oven, entire part must
be dried in an oven.
It is recommended to vacuum bag part or area to be dried using
equipment specified in 1.3, materials specified in 2.1, 2.3, 2.4,
and 2.5, and applying a minimum of 15 inches (381 mm) HG (vacuum
bagging process, paragraph 3-2-26).
3.1 Dry region to be repaired at 190 ±10°F (87.8 ±5.6°C) for carbon
fiber reinforced composite parts or 160 ±10°F (71.1 ±5.6°C) for
glass fiber reinforced composite parts for a minimum of 4 hours
using equipment specified in 1.2. Use a heat up rate and a cool
down rate between 1 to 5°F (0.6 to 2.8°C) per minute.
Examples:
3.2 A minimum of two thermocouples must constantly monitor
temperature. Thermocouples’ readings must fall within specified
temperature range at all times.
3.3 Repair must be performed within 8 hours after drying process is
performed. If not, drying procedure must be repeated. The 8 hours
limit between drying and repair must be met. It is possible to
protect dried part or repair area by covering it with bagging film
specified in 2.1 secured in place with tape specified in 2.6.
Carbon fabric: Starting point 75°F (23.9°C)
Ramp up 3°F/minute (1.7°C/min) for 38 minutes
Dwell time 4 hours at 190°F (87.8°C)
Cool down 3°F/minute (1.7°C/min) for 38 minutes
Total Elapsed Time: 316 minutes
Glass fabric: Starting point 75°F (23.9°C)
Ramp up 5°F/minute (2.8°C/min) for 17 minutes
Dwell time 6 hours at 160°F (71.1°C)
Cool down 5°F/minute (2.8°C/min) for 17 minutes
Total Elapsed Time: 394 minutes
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APPLICATION B: DRYING OF HONEYCOMB PANELS USING A LOCALIZED HEAT
SOURCE.
NOTE
Procedure for drying honeycomb panels depends on equipment used to
cure or post-cure repair. If repair is cured or post-cured using a
local heat source (heat blanket, heat lamps, or heat guns) use
Application B. If repair is cured or post-cured using an oven, use
Application C.
1.0 EQUIPMENT
1.1 Temperature monitoring equipment with a minimum of two
thermocouples.
1.2 Device capable of applying heat such as:
1.2.1 A Bell Helicopter Textron approved hot-bonding unit, or
equivalent, in conjunction with a heat blanket capable of
maintaining a minimum of 250°F (121.1°C).
1.2.2 Heat lamps.
1.2.3 Heat guns.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Bagging film (C-257) or bagging film (C-564), if
required.
2.2 Breather felt (C-258), if required.
2.3 Glass fabric (C-404) if required.
2.4 Vacuum bag sealant tape (C-259) or sealant tape (C-566), if
required.
2.5 Pressure sensitive masking tape (C-260) or tape (C-462), if
required.
2.6 Process Sheet(s): Drilling Glass or Carbon Fiber Composites
(paragraph 4-2-5)
4-38 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
NOTE
Before drying repair area, inspect for traces of water that has
migrated in core. If part has water in a repairable zone, drill
holes through skin in area containing water. Use a 0.063 inch (1.60
mm) diameter carbide drill bit, and drill using instructions
detailed in paragraph 4-2-5. For each square inch (645 mm2) of
water accumulation in repair area, drill two holes through skin to
be repaired. Drilled holes to be uniformly distributed over entire
area where water has accumulated in core.
Drying time depends on material type, part thickness, drying
temperature, and relative humidity inside part. Ensure that the
part is dried before performing repair.
It is recommended to vacuum bag part or area to be dried using
equipment specified in 1.3, material specified in 2.1, 2.2, 2.3,
and 2.4, and applying a minimum of 15 inches (381 mm) HG (vacuum
bagging process, paragraph 3-2-26).
3.1 Dry region to be repaired at 190 ±10°F (87.8 ±5.6°C) for carbon
fiber reinforced composite parts or 160 ±10°F (71.1 ±5.6°C) for
glass fiber reinforced composite parts for a minimum of 8 hours
using equipment specified in 1.2. Use a heat up rate and a cool
down rate between 1 to 5°F (0.6 to 2.8°C) per minute.
Examples:
3.2 A minimum of two thermocouples must monitor temperature.
Thermocouples’ readings must fall within specified temperature
range.
3.3 Repair must be performed within 8 hours after drying process is
performed. If not, drying procedure must be repeated. The 8 hours
limit between drying and repair must be met. It is possible to
protect dried repair area by covering it with bagging film
specified in 2.1 secured in place with tape specified in 2.5.
Carbon fabric: Starting point 75°F (23.9°C)
Ramp up 3°F/minute (1.7°C/min) for 38 minutes
Dwell time 8 hours at 190°F (87.8°C)
Cool down 3°F/minute (1.7°C/min) for 38 minutes
Total Elapsed Time: 556 minutes
Glass fabric: Starting point 75°F (23.9°C)
Ramp up 5°F/minute (2.8°C/min) for 17 minutes
Dwell time 8 hours at 160°F (71.1°C)
Cool down 5°F/minute (2.8°C/min) for 17 minutes
Total Elapsed Time: 634 minutes
14 DEC 2010 4-39ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
NOTE
Procedure for drying honeycomb panels depends on equipment used to
cure or post-cure repair. If repair is cured or post-cured using a
local heat source (heat blanket, heat lamps, or heat guns) use
Application B. If repair is cured or post-cured using an oven, use
Application C.
1.0 EQUIPMENT
1.1 Temperature monitoring equipment with a minimum of two
thermocouples.
1.2 Device capable of applying heat such as:
1.2.1 An oven (if part can be removed from helicopter).
1.3 Vacuum fitting (if required).
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Bagging film (C-257) or bagging film (C-564), if
required.
2.2 Breather felt (C-258), if required.
2.3 Glass fabric (C-404) if required.
2.4 Vacuum bag sealant tape (C-259) or sealant tape (C-566), if
required.
2.5 Pressure sensitive masking tape (C-260) or tape (C-462), if
required.
2.6 Process Sheet(s): Drilling Glass or Carbon Fiber Composites
(paragraph 4-2-5)
4-40 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
NOTE
Before drying repair area, inspect for traces of water that has
migrated in core. If part has water in a repairable zone, drill
holes through skin in area containing water. Use a 0.063 inch (1.60
mm) diameter carbide drill bit, and drill using instructions
detailed in paragraph 4-2-5. For each square inch (645 mm2) of
water accumulation in repair area, drill two holes through skin to
be repaired. Drilled holes to be uniformly distributed over entire
area where water has accumulated in core.
Drying time depends on material type, part thickness, drying
temperature, and relative humidity inside part. Ensure that part is
dried before performing repair.
It is recommended to vacuum bag part or area to be dried using the
equipment specified in 1.3, material specified in 2.1, 2.2, 2.3,
and 2.4, and applying a minimum of 15 inches (381 mm) HG (vacuum
bagging process, paragraph 3-2-26).
3.1 Dry bonded panel to be repaired in an oven at 190 ±10°F (87.8
±5.6°C) for carbon fiber reinforced composite parts or 160 ±10°F
(71.1 ±5.6°C) for glass fiber reinforced composite parts for a
minimum of 16 hours using equipment specified in 1.2. Use a heat up
rate and a cool down rate between 1 to 5°F (0.6 to 2.8°C) per
minute.
Examples:
3.2 A minimum of two thermocouples must monitor temperature.
Thermocouples’ readings must fall within specified temperature
range.
3.3 Repair must be performed within 8 hours after drying process is
performed. If not, drying procedure must be repeated. The 8 hours
limit between drying and repair must be met. It is possible to
protect dried part by covering it with bagging film specified in
2.1 secured in place with tape specified in 2.5.
Carbon fabric: Starting point 75°F (23.9°C)
Ramp up 3°F/minute (1.7°C/min) for 38 minutes
Dwell time 16 hours at 190°F (87.8°C)
Cool down 3°F/minute (1.7°C/min) for 38 minutes
Total Elapsed Time: 1036 minutes
Glass fabric: Starting point 75°F (23.9°C)
Ramp up 5°F/minute (2.8°C/min) for 17 minutes
Dwell time 20 hours at 160°F (71.1°C)
Cool down 5°F/minute (2.8°C/min) for 17 minutes
Total Elapsed Time: 1234 minutes
14 DEC 2010 4-41ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
4-2-11. FINISH PROCESS FOLLOWING A COMPOSITE REPAIR
This process sheet describes the different equipment and procedures
used to prepare the composite repair before paint
application.
APPLICATION A: INTERIOR SURFACES
CAUTION
BEFORE HANDLING A SOLVENT, PRIMER, AND SANDING SURFACER, EXTINGUISH
ALL FLAMES AND PILOT LIGHTS. KEEP PRODUCT AND ITS VAPORS AWAY FROM
HEAT, SPARKS, AND FLAME. DURING APPLICATION AND UNTIL VAPORS HAVE
DISSIPATED, AVOID USING SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH
AS SWITCHES, APPLIANCES, ETC. AVOID PROLONGED BREATHING OF VAPORS
AND REPEATED CONTACT WITH SKIN.
FINISH PROCESS MUST BE PERFORMED IN A CLEAN ENVIRONMENT, FREE OF
DUST, OIL, AND GREASE.
USE CARE WHEN APPLYING PRIMER AND SANDING SURFACER. FUMES FROM
THESE MATERIALS CAUSE BOTH HEALTH AND FIRE HAZARD. HAVE GOOD
VENTILATION AND BREATHING PROTECTION. WEAR PROTECTIVE CLOTHING AND
EYE SHIELD.
1.0 EQUIPMENT (Use as required)
1.1 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 320 to 360
grit.
2.2 Epoxy Polyamide Primer (C-204).
2.3 Cleaner: ethyl alcohol (C-339), isopropyl alcohol (C-385), or
toluene (C-306).
2.4 Cheesecloth (C-486).
4-42 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
DO NOT SAND INTO GLASS OR CARBON FIBERS.
DO NOT USE ABRASIVE PAPER THAT HAS BEEN USED TO SAND METAL.
3.1 Once composite repair is cured per instructions detailed in
paragraph 4-3-6, lightly sand surface with abrasive paper specified
in 2.1.
3.2 Wipe surface with a clean cheesecloth moistened with cleaner
specified in 2.3. Allow to air dry for 30 minutes.
3.3 Apply pinhole filler specified in 2.5 per manufacturer’s
instructions. Dry surface per manufacturer’s instructions.
3.4 Once pinhole filler is fully cured, lightly sand surface with
abrasive paper specified in 2.1.
3.5 Wipe surface with a clean cheesecloth moistened with cleaner
specified in 2.3. Allow to air dry for 30 minutes.
3.6 If required, sand to match surrounding structure and within 4
hours of sanding (step 3.4), prime with epoxy polyamide primer
specified in 2.2. Allow to dry.
14 DEC 2010 4-43ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
CAUTION
BEFORE HANDLING A SOLVENT, PRIMER, AND SANDING SURFACER, EXTINGUISH
ALL FLAMES AND PILOT LIGHTS. KEEP PRODUCT AND ITS VAPORS AWAY FROM
HEAT, SPARKS, AND FLAME. DURING APPLICATION AND UNTIL VAPORS HAVE
DISSIPATED, AVOID USING SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH
AS SWITCHES, APPLIANCES, ETC. AVOID PROLONGED BREATHING OF VAPORS
AND REPEATED CONTACT WITH SKIN.
FINISH PROCESS MUST BE PERFORMED IN A CLEAN ENVIRONMENT, FREE OF
DUST, OIL, AND GREASE.
USE CARE WHEN APPLYING PRIMER AND SANDING SURFACER. FUMES FROM
THESE MATERIALS CAUSE BOTH HEALTH AND FIRE HAZARDS. HAVE GOOD
VENTILATION AND BREATHING PROTECTION. WEAR PROTECTIVE CLOTHING AND
EYE SHIELD.
1.0 EQUIPMENT (Use as required)
1.1 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 320 to 360
grit.
2.2 Epoxy Polyamide Primer (C-204).
2.3 Sanding Surfacer (C-228).
2.4 Cleaner: ethyl alcohol (C-339), isopropyl alcohol (C-385), or
toluene (C-306).
2.5 Cheesecloth (C-486).
2.8 Process Sheet(s): Curing Process for Epoxy Resin (paragraph
4-3-6)
4-44 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
DO NOT SAND INTO GLASS OR CARBON FIBERS.
DO NOT USE ABRASIVE PAPER THAT HAS BEEN USED TO SAND METAL.
3.1 Once composite repair is cured per instructions detailed in
paragraph 4-3-6, lightly sand surface with abrasive paper specified
in 2.1.
3.2 Wipe surface with a clean cheesecloth moistened with cleaner
specified in 2.4. Allow to air dry for 30 minutes.
3.3 Apply fairing compound specified in 2.7, per manufacturer’s
instructions. Dry surface per manufacturer’s instructions.
3.4 Once fairing compound is fully cured, if required, lightly sand
surface with abrasive paper specified in 2.1.
3.5 If required, wipe surface with a clean cheesecloth moistened
with cleaner specified in 2.4. Allow to air dry for 30
minutes.
3.6 Apply pinhole filler specified in 2.6, per manufacturer’s
instructions. Dry surface per manufacturer’s instructions.
3.7 Once pinhole filler is fully cured, if required, lightly sand
surface with abrasive paper specified in 2.1.
3.8 If required, wipe surface with a clean cheesecloth moistened
with cleaner specified in 2.4. Allow to air dry for 30
minutes.
3.9 Apply sanding surfacer specified in 2.3. Allow to dry.
3.10 Sand to a smooth finish with abrasive paper specified in
2.1.
3.11 Wipe surface with a clean cheesecloth moistened with cleaner
specified in 2.4. Allow to air dry for 30 minutes.
3.12 Prime with epoxy polyamide primer specified in 2.2 before
topcoat is applied.
14 DEC 2010 4-45ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
4-2-12. PREPARATION OF MOLDING TOOL FOR COMPOSITE REPAIR
This process sheet describes the different equipment and procedures
used to fabricate a molding tool that is used to recreate the shape
of a part to be repaired.
APPLICATION A: FABRICATION OF MOLDING TOOL USING ADHESIVE ALUMINUM
TAPE
This procedure is recommended for damage up to 1.00 inch (25.4 mm)
diameter. For larger damage, procedure detailed in Application B is
recommended.
1.0 EQUIPMENT
1.1 None.
2.1 Adhesive aluminum tape (C-439).
2.2 release film (C-256).
2.3 Process Sheet(s):
Preparing and Mixing Two-part Epoxy Resin by Weight (paragraph
3-2-25) Sanding Glass or Carbon Fiber Composites (paragraph 4-2-3)
Removal of Paint, Primer, and Sanding Surfacer on Glass or Carbon
Fiber (paragraph 4-2-6) Wet Layup Impregnation Process (paragraph
4-3-3)
3.0 PROCEDURE
DO NOT SAND INTO GLASS OR CARBON FIBERS.
3.1 If required, sand surface to be molded smooth and free of
obstructions using instructions detailed in paragraph 4-2-3.
3.2 Clean surface to be molded using instructions detailed in
paragraph 4-2-6.
3.3 Prepare a piece of release film specified in 2.2 dimensioned to
same shape and size as cutout in damaged part.
3.4 Prepare a sufficient number of strips of adhesive aluminum tape
specified in 2.1 to cover beyond cutout in damaged part. Strips to
extend a minimum of 2.00 inches (50.4 mm) from edge of cutout.
Alternate tape direction to ensure maximum adhesion to damaged
part. Multiple layers of tape may be used to increase
stiffness.
3.5 Stick a piece of release film prepared in step 3.3 at the
center of sticky side of adhesive aluminum tape strips prepared in
step 3.4.
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3.6 Stick tape strips to surface of damaged part. Ensure strips
match contour of part and are sufficiently stiff to maintain
contour throughout repair. Additional tape strips may be used as
required.
3.7 Perform remaining instructions of applicable repair.
14 DEC 2010 4-47ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
APPLICATION B: FABRICATION OF MOLDING TOOL (SPLASH) USING GLASS
FABRIC
This process sheet describes the different equipment and procedures
used to fabricate a molding tool that is used to recreate the shape
of a part to be repaired.
CAUTION
PRIOR TO PREPARING A MOLDING TOOL, DAMAGED AREA MUST BE THOROUGHLY
INSPECTED AND EXTENT OF DAMAGE FULLY DEFINED THROUGH VISUAL
INSPECTION, TAP TEST, OR OTHER MEANS SUCH AS ULTRASONIC
INSPECTION.
MOLDING TOOL IS DIMENSIONED TO ACCOMMODATE EXPECTED REPAIR DOUBLER
SIZE AND BAGGING MATERIALS NECESSARY FOR CURE STEP. MOLDING TOOL
SHALL BE APPROXIMATELY 4.0 TO 8.0 INCHES (101.6 TO 203.2 MM) LARGER
(ALL AROUND) THAN LARGEST REPAIR PLY.
MOLDING TOOL SHALL BE CENTERED OVER REPAIR AREA AND ITS LOCATION
CLEARLY MARKED TO FACILITATE SUBSEQUENT PLACEMENT OF REPAIR DOUBLER
ONTO HIGHLY CONTOURED SURFACES WITHOUT ANY GAPING CONDITION.
IF IT IS NOT POSSIBLE TO TAKE MOLDING TOOL DIRECTLY FROM DAMAGED
COMPONENT, USE AN UNDAMAGED IDENTICAL COMPONENT.
1.0 EQUIPMENT (Use as required)
1.1 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Silicon carbide abrasive paper (C-423) of 100 grit or finer,
180 to 240 grit preferred.
2.2 Glass fabric (C-404), style 7781 or glass fabric (C-560), style
120.
2.3 General purpose bonding adhesive (C-317).
2.4 Wet layup adhesive (C-363).
2.5 Glass tape (C-157) or adhesive tape (C-460).
2.6 Process Sheet(s):
Preparing and Mixing Two-part Epoxy Resin by Weight (paragraph
3-2-25) Sanding Glass or Carbon Fiber Composites (paragraph 4-2-3)
Removal of Paint, Primer, and Sanding Surfacer on Glass or Carbon
Fiber (paragraph 4-2-6) Wet Layup Impregnation Process (paragraph
4-3-3)
4-48 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
DO NOT SAND INTO GLASS OR CARBON FIBERS.
3.1 If required, sand surface to be molded smooth and free of
obstructions using instructions detailed in paragraph 4-2-3.
3.2 Clean surface to be molded using instructions detailed in
paragraph 4-2-6.
3.3 Cover surface to be molded with adhesive backed Teflon coated
glass fabric specified in 2.5 to act as a release surface. Release
surface shall be 3.0 inches (76.2 mm) larger (all around) than size
of molding tool.
3.4 Build up a wax dam of suitable height to contain layup around
periphery of release surface material to prevent adhesive from
flowing away from release surface.
3.5 Prepare general purpose bonding adhesive specified in 2.3 using
instructions detailed in Chapter 3, paragraph 3-2-25.
3.6 Molding tool will be built up from alternating layers of
general purpose bonding adhesive specified in 2.3 and dry glass
fabric specified in 2.2 as follows:
NOTE
A balanced and symmetrical layup is composed of an even number of
plies. The plies are positioned in such a way that plies at a given
distance from the mid plane of the laminate (on both sides of the
mid plane) have the same fiber orientation, which helps avoid
warpage during the curing of the part. For example, a laminate
composed of 6 plies having the following orientation
(90/45/0/0/45/90) is balanced and symmetrical since the first ply
on both sides of the mid plane is at 0°, the second ply is at 45°
and the third ply is at 90°. The following laminate (90/45/0/90/
45/0) is not balanced and symmetrical since the orientation of the
first and last plies on both sides of the mid plane are not
identical.
3.6.1 Generously apply general purpose bonding adhesive specified
in 2.3 as required for a base coat.
3.6.2 Wet layup one glass fabric repair ply using instructions
detailed in paragraph 4-3-3 and wet layup adhesive specified in
2.4.
3.6.3 Repeat step 3.6.1 and step 3.6.2 as many times as required
until desired thickness is reached. First two and last two plies of
glass used may be style 120 in order to provide a smooth surface
and then style 7781 plies shall be added to provide strength and
stiffness to final tool. Final molding tool shall consist of at
least 10 plies of style 7781 fabric to handle vacuum bagging
pressure without distortion. Actual number of plies used shall take
into account size of molding tool. Plies must be oriented to
produce a balanced and symmetrical layup. If molding tool is found
to require greater strength and stiffness, a backing structure
(e.g., supporting ribs) can be fabricated onto tool using these
same materials.
14 DEC 2010 4-49ECCN EAR99
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3.7 Allow splash to cure at room temperature for 24 hours in order
to achieve handling strength.
3.8 Remove molding tool from surface of part, and allow to cure at
room temperature for an additional 48 hours.
3.9 Sand tool surface lightly with fine grit sandpaper to remove
all bubbles.
3.10 Inspect tool surface, any remaining defects must be filled and
faired using general purpose bonding adhesive specified in
2.3.
3.11 Verify fit of molding tool to damaged part.
3.12 Perform remaining instructions of applicable repair.
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4-2-13. PREPARATION OF TOOLING SURFACES
This process sheet describes the different equipment and procedures
used to clean tooling surfaces prior to layup and bonding.
CAUTION
BEFORE HANDLING SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS.
KEEP PRODUCT AND ITS VAPOURS AWAY FROM HEAT, SPARKS, AND FLAME.
DURING APPLICATION AND UNTIL VAPOURS HAVE DISIPATED, AVOID USING
SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES,
ETC.
CLEANING MUST BE PERFORMED IN A CLEAN ENVIRONMENT, FREE OF DUST,
OIL, AND GREASE.
HAVE GOOD VENTILATION AND BREATHING PROTECTION. WEAR PROTECTIVE
CLOTHING AND EYE SHIELD.
1.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
1.1 Cleaner: ethyl alcohol (C-339), isopropyl alcohol (C-385), or
toluene (C-306).
1.2 Cheesecloth (C-486).
1.4 Release agent (C-555).
1.5 Process Sheet(s):
Preparing and Mixing Two-part Epoxy Resin by Weight (paragraph
3-2-25) Sanding Glass or Carbon Fiber Composites (paragraph 4-2-3)
Preparation of Molding Tool for Composite Repair (paragraph 4-2-12)
Wet Layup Impregnation Process (paragraph 4-3-3)
2.0 CLEANING OF MOLD OR TOOL
2.1 Wipe mold or tool surface thoroughly with a clean cheesecloth
moistened with cleaner specified in 1.1. Change cheesecloth often.
Repeat operation until all evidence of residue is removed and wipe
dry using a clean cheesecloth. If scrubbing is necessary to remove
residue, use an abrasive pad specified in 1.3 wetted with cleaner
specified in 1.1, followed with wiping with cheesecloth.
3.0 APPLICATION OF RELEASE AGENT
3.1 Wipe two coats of release agent specified in 1.4 over entire
surface of mold or tool allowing 15 minutes air dry between coats.
Baking at 250 to 280°F (121.1 to 137.8°C) for 15 to 20 minutes is
recommended. Apply an additional coat of release agent and air dry
for 2 hours minimum before beginning layup.
14 DEC 2010 4-51ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
NOTE
If a problem with release of part from the tool is known or
anticipated, the following procedure shall be used instead:
3.2 Generously apply four coats of release agent specified in 1.4.
Air dry for 15 minutes between each coat.
3.3 Bake tool at 265 ±15°F (129.4 ±8.3°C) for 15 to 20
minutes.
3.4 Generously apply four coats of release agent specified in 1.4.
Air dry for 15 minutes between each coat.
3.5 Bake tool at 265 ±15°F (129.4 ±8.3°C) for 15 to 20
minutes.
3.6 Generously apply two coats of release agent specified in 1.4.
Air dry for 15 minutes between each coat.
3.7 Air dry for 2 hours before beginning layup.
3.8 Perform remaining instructions of applicable repair.
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4-2-14. PREPARATION OF FILLED EPOXY RESIN
This process sheet describes the different equipment and procedures
used to prepare epoxy resin filled with milled/chopped glass fibers
to be used to fabricate tools or fill core cavities.
1.0 EQUIPMENT (Use as required)
1.1 Personal protection equipment: safety glasses, dust mask, and
long sleeved clothing.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Glass fabric (C-404), style 7781.
2.2 Wet layup adhesive (C-363) as required in applicable
repair.
2.3 Process Sheet(s): Preparing and Mixing Two-part Epoxy Resin by
Weight (paragraph 3-2-25)
3.0 PROCEDURE
3.1 Chop glass fiber cloth specified in 2.1 to length of
approximately 0.125 inch (3.18 mm).
3.2 Prepare wet layup adhesive specified in 2.2 using instructions
detailed in Chapter 3, paragraph 3-2-25.
3.3 Mix 35 to 50% by weight of chopped glass fibers prepared in
step 3.1 to wet layup adhesive.
NOTE
Filled epoxy resin must be used within pot life of unfilled
adhesive. Refer to Table 3-16 for pot life.
14 DEC 2010 4-53ECCN EAR99
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4-3. COMMON PROCEDURES FOR WET LAYUP REPAIRS OF GLASS OR CARBON
FIBER REINFORCED COMPOSITE PARTS
4-3-1. WET LAYUP REPAIRS — GENERAL
CAUTION
NOTE THAT BOTH THE GENERAL AND SPECIFIC RESTRICTIONS MUST BE MET
BEFORE PERFORMING A COMPOSITE REPAIR. IF THE RESTRICTIONS CANNOT BE
MET, CONSULT PRODUCT SUPPORT ENGINEERING. IN CASE OF CONFLICT
BETWEEN A GENERAL AND A SPECIFIC RESTRICTION, THE SPECIFIC
RESTRICTION OVERRULES THE GENERAL RESTRICTION.
NEVER USE MEK, ACETONE, OR ALIPHATIC NAPHTHA INSTEAD OF THE
INDICATED ALCOHOLS ON GLASS OR CARBON FIBER REINFORCED COMPOSITES.
THOSE SOLVENTS ARE ONLY TO BE USED FOR CLEANING TOOLS.
WEAR APPROPRIATE SAFETY EQUIPMENT (GLOVES, GOWNS, RESPIRATORS,
ETC.) WHEN HANDLING OR WORKING WITH MATERIALS AND MAKING REPAIRS.
CONSULT MATERIAL SAFETY DATA SHEETS (MSDS) FOR POTENTIAL HAZARDS
AND FOLLOW ALL APPLICABLE SAFETY PROCEDURES.
NOTE
Determine material and tooling requirements and ensure materials
are at hand before proceeding with any repair.
Consumable Materials and Standards: Materials needed to accomplish
a particular repair are listed in the “REQUIRED” section of each
repair procedure. Each item is accompanied by a description of the
material and a numerical code (C-xxx). This code references a
consumable item, which is further described in Chapter 13 of the
Standard Practices Manual (BHT-ALL-SPM).
Required Number of Plies of Repair Doubler: Unless otherwise
indicated in a repair procedure, the number of plies of a glass
fiber repair doubler and their orientation shall be as defined in
paragraph 4-3-3. The number of plies of a glass fiber repair
doubler and their orientation may vary and is specific to the
region being repaired, but unless it is explicitly specified in the
repair, the 0° orientation is parallel to the longitudinal axis of
the helicopter (FWD/AFT direction). The number of plies of a carbon
fiber repair doubler and their orientation may vary and is specific
to the region being repaired, and shall be provided by Product
Support Engineering unless explicitly specified in the
repair.
Stop Drilling: Although stop drilling relieves the stresses in the
extremity of a crack in sheet metal parts, Bell Helicopter Textron
does not permit stop drilling cracks in fiber reinforced composite
structures.
Finishing: All repairs shall be sealed against moisture intrusion
and then finished in accordance with the original finish
specifications.
4-54 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
4-3-2. APPROVED PROCESSES (PROCESS SHEETS)
This section describes the processes used in wet layup repairs of
glass or carbon fiber reinforced composite parts. These process
sheets cover topics such as wet layup impregnation process, core
plug preparation and installation, wet layup vacuum bagging, and
curing procedures. Wet layup is a process in which dry woven glass
or carbon cloth is impregnated with a semi-liquid epoxy resin and
then cured to form a laminate. Repair procedures throughout this
manual and the model-specific Structural Repair Manual (SRM) make
reference to the applicable process sheets in the “REQUIRED”
section.
The following processes that are covered in this section are listed
in Table 4-8.
Table 4-8. Process Sheets
4-3-3 Page 4-56
Wet Layup Impregnation Process
Provides a method for impregnating dry carbon or glass fiber fabric
with epoxy resin to be used in a wet layup repair.
4-3-4 Page 4-61
Installation of Copper Wire Mesh in Wet Layup Repairs
Provides a method for repairing the copper wire mesh grounding
plane following a repair.
4-3-5 Page 4-63
Provides a method for vacuum bagging a wet layup repair.
4-3-6 Page 4-67
Curing Process for Epoxy Resin
Provides a method for curing carbon or glass fiber reinforced wet
layup repairs.
4-3-7 Page 4-73
Honeycomb Core Plug Installation and Splicing Using Epoxy
Resin
Provides a method for installing core plugs in composite faced
bonded panels.
4-3-8 Page 4-76
Preparing a Precured Wet Layup Doubler or Edge Filler
Provides a method for preparing a precured wet layup repair
doubler.
14 DEC 2010 4-55ECCN EAR99
BHT-ALL-SRM TC/FAA APPROVED
4-3-3. WET LAYUP IMPREGNATION PROCESS
This process sheet describes the different equipment and procedures
used to impregnate dry glass or carbon fabric with epoxy
resin.
CAUTION
WET LAYUP MUST BE PERFORMED IN A CLEAN ENVIRONMENT, FREE OF DUST,
OIL AND GREASE.
TO AVOID CONTAMINATION, MANIPULATE PARTS, EPOXY RESIN, AND FIBERS
USING CLEAN COTTON GLOVE.
BEFORE HANDLING A SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS.
KEEP PRODUCT AND ITS VAPOURS AWAY FROM HEAT, SPARKS, AND FLAME.
DURING APPLICATION AND UNTIL VAPOURS HAVE DISSIPATED, AVOID USING
SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES,
ETC.
1.0 PRELIMINARY REQUIREMENTS
CAUTION
AN EXCELLENT SURFACE PREPARATION IS ESSENTIAL TO ENSURE INTEGRITY
AND DURABILITY OF A BONDED JOINT. EXTRA CARE IS TO BE TAKEN TO
ENSURE THAT BONDING SURFACES ARE PROPERLY CLEANED AND PROTECTED
FROM CONTAMINATION DURING ALL PHASES OF A REPAIR.
1.1 Cleanliness is to be carefully controlled through all phases of
the preparation and bonding operations.
2.0 REQUIRED (Refer to BHT-ALL-SPM for C-xxx consumable
materials.)
2.1 Glass fabric (C-404) style 7781, glass fabric (C-560) style
120, carbon fabric (C-255), or other material as required by
applicable repair.
2.2 Wet layup adhesive as required in applicable repair.
2.3 Polyethylene film of 0.004 inch (0.10 mm) minimum thickness
(commercially available), release film (C-256), or bagging film
(C-257) or bagging film (C-564). Commercial polyethylene film is
inexpensive and is recommended.
2.4 Cleaner: MEK (C-309) or acetone (C-316).
2.5 Process Sheet(s): Preparing and Mixing Two-part Epoxy Resin by
Weight (paragraph 3-2-25) Installation of Copper Wire Mesh in Wet
Layup Repair (paragraph 4-3-4)
4-56 14 DEC 2010 ECCN EAR99
TC/FAA APPROVED BHT-ALL-SRM
3.0 PROCEDURE
3.1 Clean working table, equipment, and tools with cleaner
specified in 2.4.
NOTE
Refer to the model-specific SRM for the 0° orientation. If no 0°
orientation is specified in model-specific SRM for part to be
repaired, 0°orientation is parallel to longitudinal axis of
helicopter (FWD/AFT direction) for glass fiber repair. For carbon
fiber repair contact Product Support Engineering.
3.2 Fabricate templates using cardboard material for each ply of
repair doubler maintaining a minimum of 0.75 inch (19.1 mm)
overlap, or as specified in applicable repair, between end of
damage and first repair ply, and between end of each subsequent
repair ply. The 0° orientation and ply number shall be marked on
each template to facilitate lay up. Ensure no ply overlaps under
rivet, bolt, insert, or any other fastener. Number of plies to be
as shown in Table 4-9 unless otherwise specified in applicable
repair.
3.3 Get appropriate dry fiber reinforcement (i.e., glass fiber
fabric style 7781, glass fiber fabric style 120, carbon fiber
fabric, or other material as specified in 2.1).
3.4 Calculate amount of fabric required to fabricate all plies of
repair doubler by arranging cardboard templates on work surface
with 0° orientation properly aligned with approximately 1.0 inch
(25 mm) buffer material all around. Do not use the last 1.0 inch
(25 mm) on all sides of fabric since handling may have loosened
fibers or strands of fibers. Measure maximum length and width of
arranged templates. Cut length of fabric to required size taking
into account constraints stated above.
NOTE
For repairs with small repair plies and/or small number of plies,
it is possible to wet out all repair plies as a single piece of
fabric to be cut as individual plies after fabric is impregnated in
step 3.18. In this case, maximum size of fabric sheets should be
24.0 x 24.0 inches (610 x 610 mm) or a maximum weight of 0.265
pound (120 grams) of dry fabric, whichever results in smallest
sheet dimensions.
3.5 If required, cut fabric to more manageable dimensions by
cutting each individual ply in such way to maintain a minimum of
1.00 inch (25.4 mm) of bu