BHT-206-SRM-1 STRUCTURAL REPAIR MANUAL FOR BELL MODEL 206 SERIES HELICOPTERS THESE DATA ARE PROPRIETARY TO BELL HELICO
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BHT-206-SRM-1
STRUCTURAL REPAIR MANUAL FOR BELL MODEL 206 SERIES HELICOPTERS
THESE DATA ARE PROPRIETARY TO BELL HELICOPTER TEXTRON. DISCLOSURE, REPRODUCTION, OR USE OF THESE DATA FOR ANY PURPOSE OTHER THAN AS A GUIDE FOR REPAIR OF BELL PRODUCTS IS FORBIDDEN WITHOUT PRIOR WRITTEN AUTHORIZATION FROM BELL HELICOPTER TEXTRON. THIS MANUAL IS REPRINTED IN IT'S ENTIRETY AND INCLUDES ORIGINAL REISSUE DATED 18 FEBRUARY 1994 AS WELL AS REVISION 1 DATA. THIS MANUAL SUPERCEDES 206 SERIES STRUCTURAL REPAIR MANUAL DATED MARCH 1990 AND ALL PREVIOUS ISSUES.
Bell Helicopter TEXTRON A Subsidiary of Textron Inc
COPYRIGHT NOTICE COPYRIGHT 1995 AND BELL INC. BELL HELICOPTER HELICOPTER TEXTRON TEXTRON. AND A DIVISION OF TEXTRON CANADA LTD.
ALL RIGHTS RESERVED
POST OFFICE BOX 482 . FORT WORTH TEXAS 76101
18 FEBRUARY 1994 REVISION
1-
4 APRIL 1995
FAA APPROVED
REVISION 1
BHT-206-SRM-1
LIST OF FAA APPROVED PAGES REVISION
DATE
FEDERAL AVIATION ADMINISTRATION
APPPROVAL
MANAGER REISSUED
18 FEBRUARY 1994
REVISION 1
4 APRIL 1995
ROTORCRAFT CERTIFICATION OFFICE, FEDERAL AVIATION ADMINISTRATION, FT. WORTH, TX. 76193-0170. MANAGER ROTORCRAFT CERTIFICATION OFFICE, FEDERAL AVIATION ADMINISTRATION, FT. WORTH, TX. 76193-0170.
PAGE ...........
REPAIR ........
REVISION
PAGE...........
REPAIR .......
REVISION
NUMBER ........
NUMBER.......
NUMBER
NUMBER ........
NUMBER .......
NUMBER
1
3-31.........
N/A
1
3-32 .............
N/A
1
3-33........
TABLE 3-3 (SH1)
1
3-34 ........
TABLE 3-3 (SH2) .......
3-35.........
.
AP1 - AP2
........
N/A ............
AP3/AP4.
......... .....
N/A
3-1 .
.............
3-2 . 3-3
N/A
............. .............
...... ............
N/A ............ 3-2-3
......
3-4 .............
3-2-4...........
3-5 .
3-2-5...........
......... ...
3-6 - 3-7 ........
3-2-6
3-8 .............
3-2-7...........
3-9 .............
3-2-8
........
3-2-9
3-10 ............
. ..
3-36.......... 0
1 ...
O
TABLE 3-4
....... 1
.........
TABLE 3-6........
3-37.............
...........
........ ............
1
N/A............
O
3-38 ..........
TABLE 3-7
.........
O
3-39 ..........
TABLE 3-9
.........
O
3-40 .............
N/A
............
...........
0
3-41 ............
3-4-1A ...........
O
3-11 ............
3-2-10 ...........
1
3-42 - 3-43 ......
3-4-1B ...........
O
3-12 .............
N/A ............
1
3-44 - 3-45 ......
3-4-2 ...........
O O
3-13 .............
N/A ............
O
3-46 - 3-47 ......
3-4-3 ...........
3-14 ...........
3-2-12A
..........
0
3-48 - 3-49 ......
3-4-4 ...........
O
3-15 - 3-16 .....
3-2-12B
..........
O..
3-50 .
N/A ............
O
3-17 - 3-18 .....
3-2-12C
..........
0
3-51 - 3-52 ......
3-5-1A ...........
3-19 .............
............
O
N/A ............
0
3-53
.........
3-5-1B ..........
1
3-2-13A ..........
0
3-54
............
3-5-1B...........
0
3-21 ...........
3-2-13B
..........
0
3-55 - 3-56 ;.....
3-5-2 .........
0 0
3-22 ...........
3-2-13C
..........
0
3-57 - 3-58 ......
3-5-3
...........
0
3-23 ........
TABLE 3-2 (SH1)
.......
0
3-59 ............
3-5-4A ...........
0
3-24 ........
TABLE 3-2 (SH2) .......
0
3-60 - 3-61 ......
3-5-4B ...........
0
3-25 ........
TABLE 3-2 (SH3)
.......
0
3-62 .
N/A ............
0
3-26 ........
TABLE 3-2 (SH4)
.......
0
3-63 - 3-64 ......
3-27 ........
TABLE 3-2 (SH5)
.......
O
3-65- 3-66
0
3-67............
3-6-3A ..........
0
3-68 - 3-69 ......
3-6-3B ...........
3-20 .
.........
3-28............
3-2-14
3-29 - 3-30 ......
3-2-15 ...........
...........
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
.
......
3-6-1
...........
3-6-2
...........
0 0 0 0
AP 1
BHT-206-SRM-1 PAGE ...........
REVISION 1 REPAIR ........
NUMBER ........
FAA APPROVED
REVISION
PAGE ...........
NUMBER
NUMBER....... ......
REPAIR ........
REVISION
NUMBER .......
NUMBER
3-70 - 3-72 ......
3-6-3C ...........
0
4-59 - 4-64 ......
4-1-2D ...........
1
3-73 - 3-74 ......
3-6-4
...........
0
4-65 - 4-71 ......
4-1-2E ...........
1
3-75 ............
3-6-5A ...........
0
4-72 - 4-75 ......
4-1-2F ..........
1
3-76 - 3-77 ......
3-6-5B ...........
0
4-76 - 4-80 ......
4-1-2G ..........
1
N/A ............
1
4-81 - 4-86 ......
4-1-2H ..........
1
3-79 - 3-80 ......
3-7-1A ...........
O
4-87 - 4-92 ......
4-1-2J
...........
1
3-81 - 3-82 ......
3-7-1B ...........
0
4-93 - 4-98 ......
4-1-2K ...........
1
3-83 - 3-84 ......
3-7-2
4-99 - 4-101 .....
4-1-2L ...........
1
3-78 .............
...........
3-85 .............
N/A ............
0
4-102 - 4-109 ....
4-1-2M ...........
1
3-86 ............
3-8-1A ...........
1
4-110 - 4-117 ....
4-1-2N ...........
1
3-87 ............
3-8-1B ...........
4-1-2
1
3-88 ...........
N/A ............
3-89 ............
O
4-118 - 4-124 ....
0
4-125 - 4-132 ....
N/A ...........
1
3-9-1A ...........
0
4-133 - 4-135 ....
4-1-3A ...........
1
3-90 - 3-91 ......
3-9-1B ...........
0
4-136 - 4-142 ....
4-1-3B ...........
1
3-92 - 3-93 ......
3-9-2
...........
O
4-143 - 4-147 ....
4-1-3C ...........
1
3-94 - 3-96 ......
3-9-3A ...........
0
4-148 - 4-153 ....
4-1-3D ...........
1
3-97
3-9-3B ...........
1
4-154 - 4-158 ....
4-1-3E ...........
1
3-98 ............
3-9-3B ...........
0
4-159 - 4-167 ....
4-1-3F ...........
1
3-99 ............
3-9-3C ...........
1
4-168 - 4-175 ....
4-1-3G ...........
1
3-100 ...........
3-9-3C ...........
O
4-176 - 4-180 ....
4-1-3H ...........
1
3-101 ............
N/A ............
0
4-181 - 4-184 ....
4-1-3J ...........
1
...........
...........
3-102 - 3-103 ....
3-10-1 ...........
4-185 - 4-190 ....
4-1-3K ...........
1
3-104
3-10-2 ...........
1
4-191 - 4-198 ....
4-1-3L ...........
1
3-105 - 3-106 ....
3-10-2 ...........
0
4-199 - 4-205 ....
4-1-3M ...........
1
3-107
..........
3-10-3 ...........
1
4-206 - 4-208 ....
4-1-3N ...........
1
3-108 ...........
..........
3-10-3 ...........
0
4-209 - 4-215 ....
4-1-3P ...........
1
3-109
3-10-4 ...........
1
4-216 - 4-222 ....
4-1-3R ...........
1
3-110 ...........
3-10-4 ...........
0
4-223 - 4-229 ....
4-1-3S ...........
1
3-111 - 3-112 ....
3-10-5 ...........
1
4-230 - 4-235 .....
N/A ............
1
3-113 ...........
3-10-5 ...........
0
4-236 - 4-242 ....
4-1-4A ...........
1
3-114 ............
N/A ............
0
4-243 - 4-251 ....
4-1-4B ...........
1
..........
3-115 - 3-116 ....
3-11-1 ...........
4-252 - 4-259 ....
4-1-4C ...........
1
3-117 - 3-119 ....
3-11-2 ...........
0
4-260 - 4-265 ....
4-1-4D ...........
1
4-1 - 4-12 ........
N/A ............
1
4-266 - 4-271 ....
4-1-4E ...........
1
4-13 - 4-19 ......
4-1-1A ...........
1
4-272 - 4-278 ....
4-1-4F ...........
1
4-20 - 4-26 ......
4-1-1B ...........
1
4-279 - 4-285 ....
4-1-4G ...........
1
4-27 - 4-35 ......
4-1-1C
1
4-286 - 4-291 ....
4-1-4H ...........
1
- 4-301 ....
4-1-4J ...........
1
4-1-4K ...........
1
...........
4-36 - 4-43 .......
N/A ............
4-44 - 4-46 ......
4-1-2A ...........
1
4-302 - 4-305 ....
4-47 - 4-53 ......
4-1-2B ...........
1
4-306 - 4-311 .....
N/A ............
1
4-54 - 4-58 ......
4-1-2C ...........
1
4-312 - 4-316 ....
4-1-5A ...........
1
AP2
14-292
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
PAGE ........... NUMBER
.......
4-317 - 4-323 ....
BHT-206-SRM-1
REVISION 1
FAA APPROVED REPAIR ........ NUMBER
.......
4-1-5B ...........
REVISION
PAGE ...........
REPAIR ........
REVISION
NUMBER
NUMBER ........
NUMBER ......
NUMBER
1
5-28 - 5-33 ......
5-1-4B ...........
1 0
4-324 - 4-329 ....
4-1-5C ...........
1
5-34 - 5-38 ......
5-1-4B ...........
4-330 - 4-336 ....
4-1-5D ...........
1
5-39 - 5-41 ......
5-2-1A ...........
1
4-337 - 4-343 ....
4-1-5E ...........
1
5-42 - 5-43 ......
5-2-1A ...........
0
4-344 - 4-349 ....
4-1-5F ...........
1
5-44 - 5-49 ......
5-2-1B ...........
0
4-350 - 4-356 ....
4-1-5G ...........
1
5-50 . ...........
5-2-2
...........
0
4-357 - 4-363 ....
4-1-5H ...........
1
5-51 - 5-55 ......
5-2-3
...........
0
4-364 - 4-369 ....
4-1-5J ...........
1
5-56 - 5-56E .....
5-2-4 ...........
1
5-2-5
...........
1
4-370 - 4-376 ....
4-1-5K ...........
1
5-56F - 5-56H ....
4-377 - 4-384 ....
4-1-5L ...........
1
5-56J - 5-56K ....
5-3-1A ...........
1
4-385 ...........
4-2-1A ...........
1
5-57 ............
5-3-1A ...........
0
4-386 - 4-388 ....
4-2-10
1
5-58- 5-59
5-3-1B ...........
O
1
5-60 . ...........
5-3-1B ...........
1
5-3-1B ...........
1
5-1 ..............
...........
N/A ............
......
5-2 - 5-5 ........
5-1-1A
...........
0
5-60A- 5-60B ....
5-6 - 5-10 .......
5-1-18 ...........
0
5-61
5-3-1C ...........
1
5-3-1C ...........
1 1
. .. .......
5-11 ............
5-1-2
...........
0
5-62 - 5-63 ......
5-12 - 5-14 ......
5-1-3
...........
0
5-64 ............
5-3-1C ...........
5-1-4A ...........
0
5-65 ............
5-3-1D ...........
5-16 - 5-19 ......
5-1-4A ...........
1
5-66
5-3-1D ...........
1
5-20 ............
5-1-4A ...........
0
5-67 - 5-71 ......
5-3-1D ...........
0
5-21 ............
5-1-4A ...........
1
5-72 - 5-74 ......
5-22 - 5-26 ......
5-1-4A ...........
0
5-27 ............
5-1-4
...........
0
5-15
...........
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
.
........
5-3-2
...........
0
AP 3/AP 4
H
ELP
E
VALUATE
L
OG ISTICS
P
UBLICATIONS
Have you found something wrong with this manual -
an
error, an inconsistency, unclear instructions, etc? Although we strive for accuracy and clarity, we may make errors on occasion. If we do and you discover it, we would appreciate your telling us about it so that we can change whatever is incorrect or unclear. Please be as specific as possible. Your complaint or suggestion will be acknowledged and we will tell you what we intend to do. You may use the enclosed Customer Feedback form, as applicable, to inform us where we have erred. Your assistance is sincerely appreciated.
CUSTOMER FEEDBACK FAX TO:
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Bell Helicopter TEXTRON PRODUCT SUPPORT DIVISION 12,800 RUE DE L'AVENIR MIRABEL, QUEBEC CANADA J7J 1R4
FOLD ON DOTTED LINES AND TAPE.
TAPE HERE
BHT-206-SRM-1
REVISION 1
LIST OF EFFECTIVE PAGES Original
MARCH 1990
15
1
REVISION NUMBER
PAGE NUMBER
4 APRIL
Revision
1995
REVISION NUMBER
PAGE NUMBER
1
1
5-1 ...........................
1 1
5-2 - 5-15 ................. 5-16 - 5-19 ................
0
1 1
5-20 ...................... 5-21 ......................
0
1
5-22 - 5-27 ................ 5-28 - 5-33 ................
0
TC1 - TC8 .................
1
5-34-5-38 ................
......... TC9/TC10 ........ 1-1 - 1-18 ................. 2-1 - 2-8 ................. 2-9/2-10 ................... 3-1 - 3-2 .................
1 0
1
1 ........... 5-39 - 5-41 ..... 0 5-42 - 5-55 ................ 1 ... 5-56 ................... 1 5-56A - 5-56H .............. 1 5-56J/5-56K ................
1
5-57 - 5-59 ................ 5-60 ......................
1
5-60A - 5-60B ..............
1
......
COVER .............. TITLE ..................... AP1 - AP2 .................
............... AP3/AP4 .. A/B ......................
..
N1/N2 ..................... F1/F2 .....................
0
3-3- 3-7 .................. 3-8 ...................
.
0 ...
3-9 - 3-10 .................
1
3-11 - 3-12 ................ 3-13 - 3-33 ................
3-34 ...................... 3-35 ...................... 3-36 ...................... 3-37 - 3-52 ................
1
1
5-65 ......................
0
5-66
1
6-1 - 6-68 ................. A-1 - A-6 ................. B-1 - B-10 .................
.
...............
1 0 0 0
1
3-53 ................... 3-54 - 3-77 ................
0
1
3-78 ...................... 3-79 - 3-85 ................
3-86 ......
1
5-61 - 5-64 ................
1 0
1
........... ...
0
1
3-87 - 3-96 ................
3-97 ...................... 3-98 ...................... 3-99 ...................... 3-100- 3-103 .............. 3-104 ...................
1 0 1 O 1
3-105 - 3-106 ..............
0
3-107 3-108 3-109 3-110
................... ................ ................... .....................
1 0 1
3-111 - 3-112 .............. .............. 3-113 3-110 .............. 3-113---3-118
1 00
........... 3-119/3-120..... 4-1 - 4-388 ................
0 1
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
A/B
REVISION 1
BHT-206-SRM-1
NOTICES The information contained herein is provided subject to the pertinent provisions of Federal Aviation Regulations and statutes, and is limited to use for legitimate structural repairs and not rebuilding/remanufacture of aircraft or major portion thereof. The following acts are expressly forbidden by federal regulation or law: 1. No person, firm, or corporation other than the original manufacturer may rebuild any aircraft. FAR §43.3. 2. No person may install a data or identification plate on any aircraft, aircraft engine, rotor component, rotor blade, or rotor hub other than the one from which it was removed. FAR §45.13. 3. Persons, firms, or corporations which misidentify and aircraft rebuilt or remanufactured by that original manufacturer may be subject to civil suit for damages, injunction, seizure or destruction of the subject aircraft, or criminal prosecution with penalties up to 15 years in prison and $5,000,000 in fines. 15 U.S.C.A. §§ 1114, 1117, 1120, 1125; 18 U.S.C.A. § 2320. Bell Helicopter Textron DER approvals on FAA 8110 of repairs performed pursuant to the instructions in this manual are applicable only to repairs performed using parts/materials manufactured or approved by Bell Helicopter Textron. Repairs using parts other than parts manufactured or approved by Bell Helicopter Textron require approval by DERs other than those employed by Bell Helicopter Textron and may jeopardize Bell Helicopter's commercial warranty for any aircraft so repaired. Failure to obtain such approval may result in a violation of the terms and conditions of the Airworthiness Certificate.
FAA APPROVAL NOTICE The structural aspects only, and not the workmanship and installation, of repairs shown on FAA Approved pages in sections 3, 4 and 5 of this manual are approved by the Federal Aviation Administration for Bell Helicopter Textron model 206 series helicopters. The FAA approved repairs are listed on pages AP1 and AP2. Should conflict exist between between the indication of FAA approval on pages providing repairs and the list of of repairs on pages AP1 and AP2, the actual approval on pages AP1 and AP2 takes precedence. COMPLIANCE NOTICE The instructions set forth in this manual are applicable only to parts procured through sources approved by Bell Helicopter. The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins and other directions issued by Bell Helicopter Textron and airworthiness directives issued by the Federal Aviation Administration, must be strictly followed.
PROPRIETARY RIGHTS NOTICE These data are proprietary to Bell Helicopter Textron. Disclosure, reproduction, or use of these data for any purpose other than as a guide for helicopter repair is forbidden without written authorization from Bell Helicopter Textron.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
N1/N2
BHT-206-SRM-1
FOREWORD This structural repair manual is intended to facilitate field repair of Bell Helicopter Textron Model 206 Series airframe structure where damage is judged to fall within limits described in this manual. This manual supercedes 206 series Structural Repair Manual dated March 1990 and all previous issues. It is not possible to cover all repair eventualities in this type of publication. Repairs that deviate from those described in this manual, and in previous DER/FAA approved documents, must be analyzed to determine if the change is equivalent to a minor of major repair. Minor repairs may be accomplished using the data contained in FAA publication AC43-13-1A and the data contained in this manual. This manual does not restrict the use of FAA publication AC43-13-1A for minor repairs to cowlings, firewalls, fairings and non-structural doors. Furthermore data contained in the FAA publication may be used in other cases if all of the following conditions are met: The user has determined that (1) the data is appropriate to the part being repaired; (2) the data is directly applicable to the repair being made; and (3) the data is not contrary to manufacturer's data. Local aviation authority approval is required when using data other than supplied by the manufacturer. Major repairs that deviate from or are not covered specifically in this manual will require engineering approval. Approved sources of approval are: (1) manufacturer's data, (2) outside engineering approval (DER), or (3) FAA engineering approval.
Copies of this manual may be obtained from: Bell Helicopter Textron Customer Support and Service Division Commercial Publications Distribution Center P.O. Box 482 Fort Worth, TX 76101 This manual is published for you the operator. We welcome your comments concerning this manual. Address comments and requests for approvals as specified in other sections of this manual to: Bell Helicopter Textron Product Support Engineering 12800 Rue de L'Avenir Mirabel, Qc. J7J 1R4 Canada. Fax: (514) 433-0272 Tel: (514) 437-2862
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
REVISION 1
BHT-206-SRM-1
TABLE OF CONTENTS Section/
Title
.
Page Number
Paragraph
1-1 1-1 1-1 1-2 1-2 1-4 1-5
SECTION 1. 1-1 1-2 1-3 1-4 1-5 1-6
GENERAL INFORMATION Scope of Manual Description of Structure Types of Construction Description of Main Sub-assemblies Types of Repairs Reference Lines
SECTION 2. 2-1 2-2 2-3 2-4 2-5 2-6 2-7 2-8 2-9 2-10 2-11 2-12 2-13 2-14
2-1 DAMAGE EVALUATION 2-1 Definitions 2-2 Classification of Damage 2-2 Preliminary External Inspection 2-3 Special Inspections 2-3 Corrosion Types and Identification 2-4 Corrosion Control 2-5 Inspection of Honeycomb Panels 2-6 Requesting a Repair Procedure 2-6 Typical Request for a Repair 2-7 Gaining Access for Repair or Replacement of Parts 2-8 Requirements for a Fuselage Alignment Fixture Preparing the Helicopter for Installation into a BHT Approved Fuselage Fixture 2-9 2-9 Restricted Repair Areas 2-9 Maximum Damage Allowance
SECTION 3. 3-1 3-2 3-2-1 3-2-2 3-2-3 3-2-4 3-2-5 3-2-6 3-2-7 3-2-8 3-2-9 3-2-10 3-2-11 3-2-12 3-2-13 3-2-14 3-2-15 3-3 3-3-1 3-3-2
TYPICAL PROCESSES AND REPAIR PROCEDURES Introduction Common Procedures General Approved Processes (Process Sheets) Removal of Paints and Primers Cleaning of Honeycomb Core Cavity Preparation of Panel Bonding Surfaces Preparation of Core Plug Prior to Bonding Bonding of Flat Stock (Fillers and Doublers) Wet Lay-Up of Fiberglass Fiberglass Edging Replacement Core Splicing Potted Inserts- General Installation of Potted Inserts Removal of Potted Inserts Chemical Film Application Installation of Rivets Through Thin Honeycomb Panels Riveted Structures Rivet Replacement, General Rivet Substitution
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-1 3-1 3-1 3-1 3-2 3-3 3-4 3-5 3-6 3-8 3-9 3-10 3-11 3-12 3-12 3-19 3-28 3-29 3-31 3-31 3-32
TC 1
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS Section/ Paragraph 3-3-3 3-4 3-4-1 3-4-2 3-4-3 3-4-4 3-5 3-5-1 3-5-2 3-5-3 3-5-4 3-6 3-6-1 3-6-2 3-6-3 3-6-4 3-6-5 3-7 3-7-1 3-7-2 3-8 3-8-1 3-9 3-9-1 3-9-2 3-9-3 3-10 3-10-1 3-10-2 3-10-3 3-10-4 3-10-5 3-11 3-11-1 3-11-2
(CONT'D)
Title
Replacement of Close Tolerance Fasteners, General Rivet Pattern Discrepancies Rivet Short Edge Distance Cracked Rivet Holes Elongated, Mismatched, or Oversized Rivet Holes Mislocated Holes in Flanges Angle Section Repairs Flange Damage to Angles Extensive Damage to End of Angles Lengthwise Crack in Bend Radius of Angles Crack in Double-Formed Flange Web and Skin Repairs Oil Can Condition in Skin or Web Oil Can Condition in Bulkhead Edge Tears and Cracks in Skins and Webs Doubler Repair of Skins and Webs Edge Tears and Cracks in Bulkhead Lightening Holes Titanium Structure Repairs Flange Damage to Titanium Angles Doubler Repair to Titanium Skins or Webs Corrosion Repairs Surface Corrosion Honeycomb Panel Face Repairs Smooth Dents Sharp Dents Punctures Honeycomb Panel Edge Repairs Fiberglass Bevel Damage Fiberglass Bevel and Metal Edge Doubler Damage All Metal Bevel and Edge Doubler Damage All Fiberglass Panel Construction Damage (w or w/out Metal Edge Doubler) Outer Skin Damage on Metal Faced Panel Unsupported Composite Skin Repairs Fractured Plies Resulting from Impact Damage Puncture Damage
COMPOSITE STRUCTURE REPAIRS SECTION 4. 4-1. MAIN PANELS. (includes forward lower shell, aft lower shell and roof shell and other structural panels) Roof Shell. 206A/B, L series. 4-1-1 ROOF OUTER SECTION, 206 A/B series. Application A ROOF OUTER SECTION, 206L series Application B ROOF CENTER SECTION, 206L series Application C Forward Lower Shell, 206A/B series. 4-1-2 FWD EDGE, Damage to Edging Application A FWD EDGE, Damage to Inner Skin Around Pedals Application B FWD EDGE, Splice Repair of Console-Attaching Floor Angles Application C FWD EDGE, Damage to Corner Outer Skin Application D PANEL OPENINGS, Damage to Outer Skin/Base of Cyclic Support Application E
TC2
Page Number 3-37 3-40 3-41 3-44 3-46 3-48 3-50 3-51 3-55 3-57 3-59 3-62 3-63 3-65 3-67 3-73 3-75 3-78 3-79 3-83 3-85 3-86 3-88 3-89 3-92 3-95 3-101 3-102 3-104 3-107 3-109 3-111 3-114 3-115 3-118
4-1 4-13 4-20 4-27 4-44 4-47 4-54 4-59 4-65
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS Title
Section/ Paragraph Application Application Application Application Application Application Application Application Application 4-1-3 Application Application Application Application Application Application Application Application Application Application Application Application Application Application Application Application 4-1-4 Application Application Application Application Application Application Application Application Application Application 4-1-5 Application Application Application Application Application Application Application Application Application Application Application
(CONT'D)
F G H J K L M N P A B C D E F G H J K L M N P R S A B C D E F G H J K A B C D E F G H J K L
PANEL OPENINGS, Damage to Outer Skin/Antenna Provision PANEL OPENINGS, Damage to Outer Skin/Antenna Provision PANEL OPENINGS, Damage to Outer Skin/Cable Routing Provision PANEL OPENINGS, Damage to Outer Skin/Antenna Provision PANEL OPENINGS, Damage to Outer Skin/Antenna Provision AFT EDGE, Damage to Edging AFT EDGE, Damage to Outer Skin AFT EDGE, Damage to Outer Skin and Inner Doubler AFT EDGE, Damage to Inner Skin Forward Lower Shell, 206L Series. FWD EDGE, Damage to Edging FWD EDGE, Damage to Inner Skin Around Pedals FWD EDGE, Splice Repair of Console-Attaching Floor Angles FWD EDGE, Damage to Corner Outer Skin PANEL OPENING. Damage to Outer Skin/Fuel Drain Openings PANEL OPENING. Damage to Outer Skin/Position Light Openings PANEL OPENING, Damage to Outer Skin/Antenna Provision PANEL OPENING, Damage to Outer Skin/Antenna Provision PANEL OPENING, Damage to Outer Skin/Cargo Hook Provision PANEL OPENING, Damage to Outer Skin/Antenna Provision PANEL OPENING, Damage to Outer Skin/Antenna Provision PANEL SURFACE. Damage to Inner Skin/Fuel Transfer Lines AFT EDGE, Damage to Edging AFT EDGE, Damage to Outer Skin AFT EDGE, Damage to Outer Skin and Inner Doubler AFT EDGE, Damage to Inner Skin Aft Lower Shell, 206A/B series. FWD EDGE, Damage to Outer Skin FWD EDGE, Damage to Outer Skin and Inner Doubler FWD EDGE, Damage to Extensive Area of Inner Doubler PASSENGER DOOR SILL. Damage Affecting Outer Skin and Core PASSENGER DOOR FRAME, Pulled Inserts/Damage to Outer Skin... PASSENGER DOOR FRAME. Typical Seat Belt Damage... PASSENGER DOOR FRAME. Damage to Outer Skin Below and Aft... FUEL FILLER OPENING, Damage to Outer Skin Below Fuel Cap AFT EDGE, Damage to Outer Skin and Inner Doubler PANEL SURFACE, Damage to Fuel System Positioning Doublers Aft Lower Shell, 206L series. FWD EDGE, Damage to Outer Skin FWD EDGE. Damage to Outer Skin and Inner Doubler FWD EDGE, Damage to Extensive Area of Inner Doubler PASSENGER DOOR FRAME, Pulled Inserts/Damage to Outer Skin... FUEL FILLER OPENING, Damage to Outer Skin Below Fuel Cap PASSENGER DOOR SILL, Pulled Inserts/Damage to Outer Skin... PASSENGER DOOR SILL, Typical Seat Belt Damage... FUEL BOOST PUMP OPENING, Damage to Outer Skin/Inner Doubler FUEL BOOST PUMP OPENING, Damage to Outer Skin/Inner Doubler AFT EDGE, Damage to Outer Skin AFT EDGE, Damage to Outer Skin and Inner Doubler
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
Page Number 4-72 4-76 4-81 4-87 4-93 4-99 4-102 4-110 4-118 4-133 4-136 4-143 4-148 4-154 4-159 4-168 4-176 4-181 4-185 4-191 4-199 4-206 4-209 4-216 4-223 4-236 4-243 4-252 4-260 4-266 4-272 4-279 4-286 4-292 4-302 4-312 4-317 4-324 4-330 4-337 4-344 4-350 4-357 4-364 4-370 4-377
TC3
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS Section/ Paragraph
(CONT'D)
Title
Page Number
4-2.
SECONDARY PANELS. (includes nose panels, seat structures, baggage floor, fuselage fairings and oil cooler support). 4-2-1. Nose panel assembly. Application A PANEL SURFACE, Surface damage 4-385 Application B PANEL SURFACE, Replacement of damaged skin portion... 4-386
4-3. FLIGHT SURFACES. Not assigned at this time.
(includes horizontal stabilizer, auxiliary fins and vertical fin)
SECTION 5. 5.1 5.1.1. 5.1.2. 5.1.3. 5.1.4. 5.2. 5.2.1. 5.2.2. 5.2.3. 5.2.4 5.2.5 5.3. 5.3.1. 5.3.2. 5.4.
SHEET METAL STRUCTURE REPAIRS Forward fuselage. Controls tunnel side web repair. Controls tunnel vertical stiffeners repair. Addition of access panel to lower web of roof box beam. Forward engine attachment. Aft fuselage. Crosstube support structure. Engine pan. Longeron assembly. Engine mount half-frame splice/replacement Typical splice of aft fuselage skin Tailboom. Tailboom skin. Tailboom driveshaft support. Fairings, cowlings and cowl supports.
SECTION 6. 6.1. 6.1.1. 6.1.2. 6.1.3. 6.1.4. 6.1.5. 6.2. 6.2.1. 6.2.2. 6.2.3. 6.2.4. 6.2.5. 6.2.6. 6.2.7. 6.2.8. 6.2.9. 6.2.10. 6.2.11. 6.2.12. 6.2.13. 6.2.14.
REPLACEMENT INSTRUCTIONS Composite panels. Cabin Roof Shell Replacement. Forward or Aft Lower Shell Replacement. Passenger Seat Back Panel Replacement. Baggage Compartment Wall Replacement. Passenger Seat Bottom and Front Panels Replacement. Sheet metal structure. Instrument Console Assembly Replacement. Reserved. Crew Seat Bulkhead Assembly Replacement. Lower Shell Splice Stiffener Replacement. Engine Pan Assembly Replacement. Engine Pan Longeron Replacement. Roof Beam Assembly Replacement. Engine Mount Support Fitting Replacement. Frame Assembly Replacement. Engine Mount Support Channel and Frame Assembly Replacement. Passenger Compartment Bulkhead Frame Replacement. Landing Gear Aft Crosstube Support Structure Replacement. Fuselage Aft Bulkhead Replacement. Upper Tailboom Attachment Fitting Replacement.
TC4
5-2 5-11 5-12 5-15 5-39 5-50 5-51 5-56 5-56F 5-56J 5-72 Not assigned.
6-2 6-4 6-6 6-8 6-10 6-12 6-14 6-16 6-19 6-21 6-24 6-27 6-30 6-32 6-34 6-36 6-38 6-41
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS Section/ Paragraph
6.3. 6.3.1. 6.3.2 6.3.3. 6.3.4. 6.3.5. 6.3.6. 6.3.7. 6.3.8. 6-62 6.3.9. Appendix A. A-1. A-2. A-3. A-4. A-5. A-6 Appendix B. B-1 B-2 B-3
(CONT'D)
Title
Page Number
Tailboom / Empennage. Tailboom Attachment Fitting and Intercostal Replacement. Reserved. Reserved. Tailboom Skin Replacement. Driveshaft Support Replacement. Tailboom Control Tube Fairlead Replacement. Replacement of One-piece Cast Aluminum Support. Replacement of S. Metal T/R Gearbox Support with a Cast Aluminum Support.
6-43
6-45 6-55 6-56 6-59
Replacement of Vertical Stabilizer Leading and Trailing Edges.
6-65
CONSUMMABLE MATERIALS AND ORDERING INFORMATION General. Identifying Sheet Aluminum Alloys. General Procedure for Duplication of Parts. Ordering Consummable Materials. Ordering Honeycomb Material. Ordering Composite Bond Material.
A-2 A-2 A-3 A-3 A-3 A-3
WORKAIDS Fuselage Support. Tailboom Drill Plate. Tailrotor Gearbox Positioning Workaid.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
B-2 B-5 B-9
TC5
REVISION 1
BHT-206-SRM-1
TABLE OF CONTENTS
(CONT'D)
LIST OF FIGURES Title
Figure Number
206A/B Series General Arrangement 206L Series General Arrangement 206LT Series General Arrangement 206A/B Series Principal Dimensions 206L Series Principal Dimensions 206LT Series Principal Dimensions 206 Series Fuselage Assembly 206A/B Series Basic Lines Data 206L Series Basic Lines Data 206LT Series Basic Lines Data
1-1 1-2 1-3 1-4 1-5 1-6 1-7 1-8 1-9 1-10
Model Model Model Model Model Model Model Model Model Model
3-1 3-2 3-3 3-4 3-5 3-6 3-7 3-8 3-9 3-10 3-11 3-12 3-13 3-14 3-15 3-16 3-17 3-18 3-19 3-20 3-21 3-22 3-23 3-24 3-25 3-26 3-27 3-28 3-29 3-30 3-31 3-32 3-33
Preparation of Core Plug Prior to Bonding Fiberglass Edging Replacement Core Splicing Installation of Bell Standard 80-004/80-005 Insert Installation of Bell Standard 80-011/80-013 Insert Installation of Bell Standard 80-007 Insert Removal of Bell Standard 80-004/80-005 Insert Removal of Bell Standard 80-011/80-013 Insert Removal of Bell Standard 80-007 Insert Installation of Rivets Through Thin Honeycomb Panels Rivet Short Edge Distance (1 of 2) Rivet Short Edge Distance (2 of 2) Cracked Rivet Holes Elongated, Mismatched, or Oversized Holes Mislocated Holes in Flanges Limited Flange Damage to Angles Extensive Flange Damage to Angles Extensive Damage to End of Angles Lengthwise Crack in Bend Radius of Angles Crack in Double-Formed Flange (1 of 2) Crack in Double-Formed Flange (2 of 2) Oil Can Condition in Skin or Web Oil Can Condition in Bulkhead Edge Tear/Crack in Skins and Webs (1 of 3) Edge Tear/Crack in Skins and Webs (2 of 3) Edge Tear/Crack in Skins and Webs (3 of 3) Doubler Repair to Skins and Webs Tear/Crack in Bulkhead Lightening Hole (1 of 2) Tear/Crack in Bulkhead Lightening Hole (2 of 2) Limited Flange Damage to Titanium Angle Extensive Flange Damage to Titanium Angle Doubler Repair to Titanium Skin or Web Panel Surface- Smooth Dents
TC6
Page Number 1-6 1-7 1-8 1-9 1-10 1-11 1-12 1-16 1-17 1-18 3-6 3-10 3-11 3-14 3-15 3-17 3-20 3-21 3-22 3-29 3-41 3-42 3-44 3-46 3-48 3-51 3-53 3-55 3-57 3-59 3-60 3-63 3-65 3-67 3-68 3-70 3-73 3-75 3-76 3-79 3-81 3-83 3-90
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS
(CONT'D)
LIST OF FIGURES (CONT'D) Figure Number
Title
Page Number
3-34 3-35 3-36 3-37 3-38 3-39 3-40 3-41 3-109 3-42 3-43 3-44
3-92 Panel Surface- Sharp Dents 3-94 Puncture in Fiberglass or Metal Faced Panel (1 of 3) 3-97 Puncture in Metal Faced Panel (2 of 3) 3-99 Puncture in Fiberglass Faced Panel (3 of 3) 3-102 Fiberglass Panel Bevel Repair 3-104 Repair Fiberglass Edging and Metal Edge Doubler 3-107 Panel Edge-All Metal Construction Repair Repair Doubler) Metal Edge w/out or (with Construction Panel Edge-All Fiberglass Edge Repair Affecting Outer Skin of Metal Faced Panel Fractured Plies in Unsupported Composite Skin Panel Puncture in Unsupported Composite Skin Panel
3-111 3-115 3-117
4-1 4-2 4-3 4-4 4-5 4-6
Section 4, Guidelines Roof Shell 206A/B and L Series, Restrictions Roof Shell 206A/B and L Series, Insert Locations Roof Outer Section, 206A/B Series Roof Outer Section, 206L Series Roof Center Section, 206L Series
4-5 4-9 4-11 4-13 4-20 4-27
4-7 4-8 4-9 4-10 4-11 4-12 4-13 4-14 4-15 4-16 4-17 4-18 4-19 4-20 4-21 4-22
Forward Lower Shell 206A/B Series, Restrictions Forward Lower Shell 206A/B Series, Insert Locations Forward Edge, Edging Forward Edge, Inner Skin at Base of T/R Pedals Forward Edge, Splicing Console-Attaching Angle(s) Forward Edge, Corner Outer Skin Panel Openings, Base of Cyclic Support Panel Openings, Antenna Provision Panel Openings, Antenna Provision Panel Openings, Cable Routing Hole Panel Openings, Antenna Provision Panel Openings, Antenna Provision Aft Edge, Edging Aft Edge, Outer Skin Aft Edge, Outer Skin and Inner Doubler Aft Edge, Inner Skin
4-38 4-39 4-44 4-47 4-54 4-59 4-65 4-72 4-76 4-81 4-87 4-93 4-99 4-102 4-110 4-118
4-23 4-24 4-25 4-26 4-27 4-28 4-29 4-30 4-31 4-32
Forward Lower Shell, 206L Series-Restrictions Forward Lower Shell, 206L Series-Insert Locations Forward Edge, Edging Forward Edge, Inner Skin at Base of T/R Pedals Forward Edge, Splicing Console-Attaching Angle(s) Forward Edge, Corner Outer Skin Panel Openings, Forward Fuel Cell Drain Panel Openings, Position Light Panel Openings, Antenna Provision Panel Openings, Antenna Provision
4-127 4-128 4-133 4-136 4-143 4-148 4-154 4-159 4-168 4-176
USE OR DISCLOSURE OF-DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
TC7
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS
(CONT'D)
LIST OF FIGURES (CONT'D) Figure Number
Title
Page Number
4-33 4-34 4-35 4-36 4-37 4-38 4-39 4-40
Panel Openings, Cable Routing Hole Panel Openings, Antenna Provision Panel Openings, Antenna Provision Panel Surface, Fuel Transfer Lines Cover Aft Edge, Edging Aft Edge, Outer Skin Aft Edge, Outer Skin and Inner Doubler Aft Edge, Inner Skin
4-181 4-185 4-191 4-199 4-206 4-209 4-216 4-223
4-41 4-42 4-43 4-44 4-45 4-46 4-47 4-48 4-49 4-50 4-51 4-52
Aft Lower Shell, 206A/B Series-Restrictions Aft Lower Shell, 206A/B Series-Insert Locations Forward Edge, Outer Skin Forward Edge, Outer Skin and Inner Doubler Forward Edge, Inner Doubler Passenger Door Sill, Outer Skin Passenger Door Frame, Pulled Inserts Passenger Door Sill, Typical Seat Belt Damage Passenger Door Frame, Aft of Door Frame Fuel Filler Opening, Below Fuel Cap Aft Edge, Below Fuel Cap Panel Surface, Positioning Doublers
4-232 4-233 4-236 4-243 4-252 4-260 4-266 4-272 4-279 4-286 4-292 4-302
4-53 4-54 4-55 4-56 4-57 4-58 4-59 4-60 4-61 4-62 4-63 4-64 4-65
Aft Lower Shell, 206L Series-Restrictions Aft Lower Shell, 206L Series-Insert Locations Forward Edge, Outer Skin Forward Edge, Outer Skin and Inner Doubler Forward Edge, Inner Doubler Passenger Door Frame, Pulled Inserts Fuel Filler Opening, Below Fuel Cap Passenger Door Sill, Outer Skin Passenger Door Sill, Typical Seat Belt Damage Fuel Pump Opening, Outer Skin Fuel Pump Opening, Outer Skin and Inner Doubler Aft Edge, Outer Skin Aft Edge, Outer Skin and Inner Doubler
4-308 4-309 4-312 4-317 4-324 4-330 4-337 4-344 4-350 4-357 4-364 4-370 4-377
TC8
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
REVISION 1
TABLE OF CONTENTS
(CONT'D)
LIST OF TABLES Title
Table Number 3-1 3-2 3-2 3-2 3-2 3-2 3-3 3-3 3-4 3-5 3-6 3-7 3-8 3-9
Adhesive Data Insert 80-004 Installation Data (1 of 5) Insert 80-005 Installation Data (2 of 5) Insert 80-007 Installation Data (3 of 5) Insert 80-011 Installation Data (4 of 5) Insert 80-013 Installation Data (5 of 5) Substitution of Protruding Head Rivets (1 of 2) Substitution of Flush Head Rivets (2 of 2) Drill and Hole Size Limits for Solid and Blind Rivets Effective Grip Lengths for Blind Rivet Installation Solid Rivet Identification Substitution of Shear Type Fasteners Oversize Part Number Equivalencies for Hi-Lok Fasteners Selected Hole Diameters for Hi-Lok Fasteners
4-1 4-2 4-3 4-4 4-5
Repairs Repairs Repairs Repairs Repairs
A-1 A-2 A-3
Consummable Materials Composite Bond Sheet Material. Honeycomb Repair Material.
Covered Covered Covered Covered Covered
in in in in in
Section Section Section Section Section
4-1-1 4-1-2 4-1-3 4-1-4 4-1-5
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
Page Number 3-8 3-23 3-24 3-25 3-26 3-27 3-33 3-34 3-35 3-35 3-36 3-38 3-38 3-39 4-8 4-37 4-126 4-231 4-307 A-4 A-5 A-6
TC9/ TC 10
BHT-206-SRM-1
SECTION 1. GENERAL INFORMATION 1-1.
SCOPE OF THIS MANUAL.
This Structural Repair Manual provides information pertinent to the repair of the fuselage assembly. The manual has been divided into sections to facilitate retrieval of information and it's use to the operator. a. SECTION ONE. Offers general information about use of this manual and the general construction of the helicopters covered by this book. b.
SECTION TWO.
Provides useful information regarding damage evaluation.
c. SECTION THREE. Provides approved data on the most useful aspects of accomplishing repairs. It includes information about processes, techniques and typical repairs used every day. Typical repairs may be used in most cases. d. SECTIONS FOUR THRU SIX. These sections deal with instructions on specific known cases of repair for a particular area of the aircraft. These procedures deal with precise limits of repairability and the procedure for repair in cases when these limits exceed those provided in section 3. This data is compiled from cases reported to us over the last few years. These sections are intended to be expanded as new data is made available. SECTION FOUR. Provides repair instructions for composite structure. SECTION FIVE.
Provides repair instructions for sheetmetal structure.
SECTION SIX.
Provides replacement instructions for composite and sheetmetal structure.
e. APPENDICES. The appendices list information useful, but not essential, to accomplishing a repair. The information has been broken down as follows: APPENDIX A. Provides information on consummables and ordering of materials specific to structural repairs. APPENDIX B.
1-2.
Provides information on required workaids referenced in this manual.
DESCRIPTION OF STRUCTURE.
The 206 family of helicopters is composed of the 206A JetRanger and 206B JetRanger II and III (figure 11), 206L LongRanger, 206L-1 LongRanger II, 206L-3 LongRanger III, 206L-4 LongRanger IV (figure 1-2) and 206LT TwinRanger (figure 1-3). The structures on all these model helicopters are similar; the differences being changes required to accomodate different engines and the 25 inches longer cabin and 10.5 inches longer tailboom used on the 206L series. The structures of each of these models consists of three sections, distinct in the way they are constructed. Refer to figures 1-4, 1-5 and 1-6 for helicopter principal dimensions.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-1
BHT-206-SRM-1 1-3.
TYPES OF CONSTRUCTION.
a. FORWARD SECTION (CABIN): The forward section extends from the nose to the rear of the passenger compartment. It's basic structure consists of two honeycomb panels joined to form the bottom of the cabin or floor, while a sheetmetal roof beam assembly joined to the center of another honeycomb panel above form the roof. The forward floor and roof honeycomb panels consist of aluminum core material and aluminum faces or skins. The aft floor assembly consists of aluminum honeycomb core material with fiberglass skins. The box beam structure supports the transmission and forward section of the engine. Two bulkhead assemblies and a center post connect the floors and roof to form and integrated stucture. The bases of the forward and aft bulkheads support the landing gear. b. INTERMEDIATE SECTION (AFT FUSELAGE): The intermediate section extends from aft of the passenger compartment to the front of the tailboom. The structure is of semi-monocoque construction utilizing six bulkheads, four primary longerons, two secondary longerons and the fuselage skins. The upper part of the aft fuselage accomodates the engine pan assembly, which supports the engine assembly, and a support structure for the oil cooler. The fuselage skins and stiffeners transfer loads from the four main longerons to the forward section. c. EMPENNAGE SECTION (TAILBOOM): The tailboom extends from the aft fuselage to the vertical fin. It is of monocoque construction consisting of two aluminum skins joining an intercostal assembly at the forward end and a support structure for the tailrotor gearbox and vertical fin at the back. Support for the tailrotor driveshaft is accomplished with the use of partial frames attached to the inside of the upper skin. The tailboom assembly supports the vertical fin, horizontal stabilizer assembly and tailrotor drive components.
1-4.
DESCRIPTION OF MAIN SUB-ASSEMBLIES. (refer to figure 1-7)
a. NOSE SKIN AND INSTRUMENT CONSOLE (item 1). The nose skin consists of two panels on either side of the battery compartment, generally consisting of aluminum core material and fiberglass skins. However, the lower aft sides by the static ports consist of sheet aluminum core covered with fiberglass facing. The panels are stiffened by frames and attached to the structure by means of rivets. Aluminum retainer strips riveted to the nose panels support part of the winshields and lower windows. The instrument console (item 8) is constructed of aluminum alloy skins, stiffeners and bulkheads. Access doors, support shelves and a flight control support fitting are installed on the console. b. CABIN ROOF (item 2). The roof shell is constructed of aluminum honeycomb core, aluminum skins and fiberglass edging material. The aftmost 10 inches of the roof assembly behind the forward engine firewall is constructed of titanium alloy. The transmission, hydraulic servos and flight controls are mounted to the roof. c. ENGINE PAN (item 3). The engine pan is constructed of titanium sheet and aluminum longerons. It supports the engine and protects the surrounding structure from flammable liquids and will contain possible fires. It is located on the upper forward section of the aft fuselage. d. OIL COOLER SUPPORT (item 4). The oil cooler support is constructed of aluminum sheet and stiffeners (206A/B prior to S/N 2212) and aluminum honeycomb core, skins and fiberglass edging for all other model effectivities. It is mounted just behind, and is separated from the engine pan by the aft engine firewall. The oil cooler provides a mounting surface for the oil cooler assembly. e. COWLING SUPPORT RAILS (item 5). The sheet metal angles mounted to the upper portion of the fuselage, locate and attach the cowlings to the structure.
1-2
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1 f. FLOOR PANEL (item 7). The floor panel provides the floor and exterior surface of the cabin portion of the fuselage. The floor panel consists of two honeycomb panels; the forward floor panel and the aft floor panel. The front panel extends from the aft edge of the lower windows to the base of the rear passenger forward facing seat. It is constructed of aluminum honeycomb with aluminum inner and outer skins. The aft panel which houses the fuel cell extends aft from the forward lower shell to the rear of the cabin. It is constructed of aluminum honeycomb core with fiberglass inner and outer skins. The landing gear attaches to the floor panel. g. CREW SEAT STRUCTURE (item 9). The seat base is supported by a bulkhead on the forward end and support angles attached to the crew seat bulkhead on the aft end. Aluminum closure panels seal the outboard sides and center section between the two seats. Two configurations of seat base panel may be found depending on the model effectivity: early configurations used a flat honeycomb panel while current production uses a stamped sheetmetal base. h. CREW SEAT BULKHEAD (item 10). This bulkhead is situated directly aft of the crew seat structure and serves as the seat back. It is one of two bulkheads that join the roof to the floor and provides attachment for the forward crosstube supports, crew lap and shoulder seat belt attachment points and crew seat structure. It is constructed of aluminum frames and webs. i. CENTER POST (item 11). The center post is situated between the crew seats and just behind the crew seat bulkhead. It is consists of four angle sections and aluminum sheetmetal webs riveted on three sides. Access through the back of the centerpost is provided by a removable panel. The center post provides support to the front end of the roof beam and accomodates engine and flight controls installations. j. ROOF BEAM (item 12): The roof beam assembly extends from the crew seat bulkhead to the rear of the cabin. It consists of aluminum alloy forgings, bulkheads, longeron angles and skins. The beam provides the structural support for the transmission and rotor system loads. k. AFT CABIN CLOSURE BULKHEAD (item 13). This bulkhead is situated directly behind the passenger seat structure. The second of the two bulkhead assemblies that join the roof to the floor, it supports the aft end of the roof beam and provides attachment for the aft crosstube support structure. Along with the aft floor and aft cabin bulkhead it encapsulates the fuel cell compartment. I. AFT CABIN BULKHEAD (item 15). The aft cabin bulkhead is constructed of aluminum honeycomb and skins, aluminum and titanium frames, clips and stiffeners. It forms the aft section of the aft cabin closure and supports the aft end of the roof beam and takes landing gear loads. m. PASSENGER SEAT BOTTOM, SEAT BACK AND FUEL CELL TOP PANELS (item 16). These panels are constructed of aluminum honeycomb core and facings. The seat back forms the forward section of the aft cabin closure and provides attachment for the seat lap belt fittings. n. AFT FUSELAGE LONGERONS AND FRAMES (item 17). The semi-monocoque aft fuselage structure is comprised of four full-length or main longerons and two partial-length or secondary longerons, six bulkheads, a canted web assembly, a baggage compartment floor panel, aluminum alloy skins and aluminum faced fairings. The two upper longerons extend from the aft portion of the roof panel aft on either side of the engine pan. The lower longerons extend from the aft cabin closure bulkhead aft just above the baggage compartment and are joined in the middle by the canted web assembly. Two short longerons extend from the lower shoulder of the aft cabin bulkhead to the base of the fifth bulkhead and are joined by the baggage compartment floor panel. Equally spaced bulkheads span the distance from the aft fuselage bulkhead to the rear of the aft fuselage. o. HORIZONTAL STABILIZER, 206A/B SERIES (item 20). The stabilizer assembly consists of two reverse airfoil sections attached to a tubular spar. The airfoil sections are constructed of aluminum sheet metal ribs and skins riveted together. Navigation lights are mounted on the outboard ribs of the stabilizers.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-3
BHT-206-SRM-1 HORIZONTAL STABILIZER, 206L SERIES (item 20). The stabilizer assembly consists of a full-length inverted airfoil section with two tip mounted vertical fins on either end and a movable elevator mounted on it's trailing edge. Two extruded nylon slats are mounted to the leading edge lower surface on all but the model 206L helicopters. Construction of the components of the stabilizer assembly is of aluminum honeycomb and facings. p. VERTICAL FIN (item 21). The vertical fin is a flat aluminum honeycomb and facings panel mounted at a preset angle on the aft section of the tailboom. Preformed sheetmetal leading and trailing edges are bonded to the face of the panel. A tail skid and bumper assembly are bonded to the lower edge of the panel. The tips of the vertical fin are sealed with fiberglass covers. Formed angles bonded to the center of some vertical fins provides attachment for the tailrotor gearbox fairings. A stroboscopic light is monted on the upper tip of the vertical fin assembly.
1.5.
TYPES OF REPAIRS.
The FAA Approved repair data contained in this manual were developed following the requirements of the United States Department of Transportation, Federal Aviation Regulation number 43 and Bell Helicopter Textron Engineering. The structural engineering aspects only of the repairs described in this manual are FAA approved for parts obtained through sources approved by Bell Helicopter Textron . Refer to the FAA approved identification on the individual pages and the List of FAA approved pages and repairs at the front of the book. WARNING USE OF THIS DATA ON PARTS OBTAINED THROUGH SOURCES OTHER THAN BELL HELICOPTER TEXTRON IS FORBIDDEN. Various types of repair methods are available to the operator depending on the type of structure to be repaired. The methods mostly used throughout this manual are: This is one of the most common repair procedure used. The damaged DOUBLER REPAIR. a. portion of the part being repaired is removed and a doubler attached to it's surface. The doubler may be bonded, riveted or attached using a combination of both to ensure that the original strength is regained. INSERTION REPAIR. Insertion repairs are necessary when a smooth repaired surface is required. b. It consists of removing the damage section of material and then inserting a newly fabricated section of the same material and size as the original parent material. A doubler repair is then accomplished on one of the surfaces to add the necessary reinforcement. Riveted repairs as specified throughout this manual specify the materials c. RIVETED REPAIR. and the rivet pattern to be used. Refer to Section three for additional information. Bonded repairs or a combination of bonded and riveted repairs are given d. BONDED REPAIR. throughout this manual. Refer to Section three for additional information. Welded repairs are not approved or discussed in this manual. Limited e. WELDED REPAIR. welded repairs to parts such as skid shoes and exhaust ducts are permitted and procedures may be found in the applicable Maintenance or Component Repair manuals.
1-4
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BHT-206-SRM-1 1.6 REFERENCE LINES.
(refer to figures 1.7 and 1.8)
Reference lines are a useful tool used to identify or locate a precise location on the airframe. A location is given by specifying the intersection in space of various lines known as: a. DATUM. The datum is the origin from which the location or geometric characteristics of features of a part are established. This point may or may not be on the part or assembly. On the 206 Series for example, datum are all outside the actual structure. Refer to figures 1-5 (206A/B series) and 1-6 (206L series). b. FUSELAGE STATION LINES. Fuselage station lines (abbreviated FS) are vertical reference lines as viewed from the side of the helicopter. They express the distance in inches from th datum line of FS 0.0 of a given point. On the 206 Series, datum line FS 0.0 is located approximately 1.0 inch forward of the most forward nose contour. c. BOOM STATION LINES. Boom station lines (abbreviated BS) are similar to the FS but serve to distinguish station lines applicable to the tailboom. They express the distance in inches from th datum line of BS 0.0 of a given point. On the 206 Series, datum line BS 0.0 is located approximately 32.0 inches forward of the most forward surface of the tailboom d. WATER LINES. Water lines (abbreviated WL) are horizontal reference lines as viewed from the side or front of the helicopter. They express the distance in inches above a line of origin located approximately 20.0 inches below the outer surface of the floor panel. BUTTOCK LINES. Buttock lines (abbreviated BL) are vertical reference lines as viewed from the e. front of the helicopter. They express the distance in inches from the center line of the helicopter. Dimensions given left of the center line are refered to as left buttock line (LBL); dimensions given right of the center line are refered to as right buttock line (RBL).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-5
BHT-206-SRM-1
1. 2. 3. 4. 5. 6. 7. 8.
Cowling Installation Forward Section Landing Gear Assembly Intermediate Section Tailboom Assembly Stabilizer Installation, Horizontal Fairing Installation, Tail Rotor Gearbox Tail Fin Assembly
Figure 1-1. Model 206A/B Series General Arrangement.
1-6
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BHT-206-SRM-1
8
2
1. 2. 3. 4. 5. 6. 7. 8.
Cowling Installation Forward Section Landing Gear Assembly Intermediate Section Tailboom Assembly Stabilizer Installation, Horizontal Fairing Installation, Tail Rotor Gearbox Tail Fin Assembly
Figure 1-2. Model 206L Series General Arrangement. USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-7
BHT-206-SRM-1
6
Cowling Installation
1. 2. 3. 4. 5. 6.
Cowling Installation Fairing Installation, Tail Rotor Stabilizer Installation, Horizontal Tailboom Assembly Landing Gear Assembly Fuselage Structure Assembly
Figure 1-3. Model 206LT Series General Arrangement.
1-8
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BHT-206-SRM-1
NOTE: DIMENSIONS ARE IN FEET
4.3
38.8 8.8
6.5
Figure 1-4. Model 206A/B Series Principal Dimensions USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-9
BHT-206-SRM-1
NOTE: DIMENSIONS ARE IN FEET
4.33
42.39
1.05
. 4.33 37.00
Figure 1-5. Model 206.91Series Principal Dimensions
12.05
33.22
6.48
Figure 1-5. Model 206L Series Principal Dimensions USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
NO LOAD ON GEAR 7 FT 5 IN. 4000 LB ON GEAR 7 FT 8.1 N. POINT
RADIUS 23 FT 10 IN.
37 FT 0.0 IN.
9 FT 10.9 IN.
42 FT 6.2 IN.
2 15'PRECONE
11 FT 8.3 IN. 6 FT 2.7 IN.
20 09'
3 FT 2.3 IN.
WL 0.00 1 FT 3.0 IN. DOOR OPENING (APPROXIMATE) 40 IN. HEIGHT 60 IN. WIDTH
Figure 1-6. Model 206LT Series Principal Dimensions USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-11
BHT-206-SRM-1
206B
206L Figure 1-7. Model 206 Series Fuselage Assembly (Sheet 1 of 4)
1-12
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BHT-206-SRM-1
206 B
206L
Figure 1-7. Model 206 Series Fuselage Assembly (Sheet 2 of 4) USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1-13
BHT-206-SRM-1
21
206A/B/B
III
A206L/L1/L
III
Figure 1-7. Model 206 Series Fuselage Assembly (Sheet 3 of 4)
1-14
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BHT-206-SRM-1
BOTTOM FLAT
WL 93
ROOF FLAT
150
217.84 PLAN
76
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
1.17
BHT-206-SRM-1
STA
STA
STA
182.405
WL 72.00
WL
230.76 I
Figure 1-10. Model 206LT Series Basic Lines Data.
1-18
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BHT-206-SRM-1
SECTION 2. DAMAGE EVALUATION 2-1.
DEFINITIONS.
The term damage is used to describe the degradation of a part or assembly from the intended or new condition. Damage may be a result of any of a multitude of conditions varying from normal mechanical wear, exposure to the elements, be a result of an accident or modification to the original part. The following terms are commonly used to describe different types of damage. ABRASION. The wearing away of small amounts of material as a result of friction between mating parts. BLISTER. The raised portion of a surface, caused by separation of layers of material. BUCKLING. Large scale deformation of a part from the original shape, usually caused by high compressive loads (e.g. hard landing), excessive localized heating or a combination of these loads. BURN. Discoloration resulting from exposure to excessive heat. CORROSION. A surface chemical action resulting in surface discoloration, advanced stages, the removal of surface metal.
a layer of oxide, or in
CRACK. A fissure or break in material. DENT. A smooth, round-bottom depression. DISTORTION: FLAKING.
A change from the intended or original shape.
Loose particles of material on a surface or evidence of removal of surface covering.
FRETTING. Wearing away of surface material caused by repeated motion of adjacent surfaces. GOUGING. A removal of surface material typified by rough and deep depressions. NICK. A sharp-bottomed depression with rough outer edges. OIL CANNING. A characteristic of thin sheet material following a contour to be flexed so that the surface will snap-through and be either concave or convex. PITTING. A condition recognized by minute holes or cavities which occur on surface areas. PUNCTURE. A break in thin sheet material, typically caused by a foreign object contacting the surface of the material. SCORING. A form of wear characterized by a scratched, scuffed or dragged appearance with markings in the direction of sliding. SCRATCH. Narrow, shallow mark or line resulting from the movement of a metallic particle or sharp pointed object across a surface. SCUFFING. A dulling or moderate wear of a surface resulting from rubbing. USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
2-1
BHT-206-SRM-1 SPALLING. A surface or subsurface defect characterized by chips of material that spall or flake out leaving cavities of varying sizes and depths. STRESS FAILURE. Compression, tension, shear, torsion or shock which leads to material failure. TEAR. A shape linear rupture in sheet material, typically with the sheet bent away from it's original shape. WEAR. A condition resulting from a relatively slow removal of parent material. Frequently not visible to the unaided eye.
2-2.
CLASSIFICATIONS OF DAMAGE. a. NEGLIGIBLE DAMAGE. Negligible damage is that which can be permitted to remain as it is or made acceptable by simple procedure, such as removing dents, polishing, smoothing nicks, priming or spot painting, and without placing restrictions on flight. b. REPAIRABLE DAMAGE. Repairable damage is that exceeding the specified negligible damage limits but not so severe as to warrant replacement. Components damaged beyond the limits deemed repairable must be replaced. c. CRASH DAMAGE. In view of the many possible combinations resulting from crash damage, it is not possible to include specific repair schemes in this category. Crash damage must be evaluated for individual situations.
2-3. PRELIMINARY EXTERNAL INSPECTION. NOTE The following inspections are provided as a guide in damage evaluation. These instructions are not intended to replace the approved inspections specified in the applicable Maintenance or Overhaul Manual. a.
Inspect fuselage skins for dents, cracks, holes, wrinkles, buckles and missing or loose rivets.
b. Inspect cabin doors and inspection access panels for fit, alignment, obvious damage and security of mounting. c.
Inspect windows for cracks, scratches, holes and security of mounting.
d. Inspect composite panels for deformation,surface damage such as dents, scratches, voids and missing or loose rivets. e. Inspect skid landing landing gear and tailskid for deformation, dents, cracks, holes, loose or missing rivets and security of fasteners. f.
Inspect tailboom exterior surfaces and mounting surfaces for condition.
g. Inspect fuselage and tailboom internal structures for cracks, dents, deformation, missing or loose rivets and misalignment. h. Inspect structural mounting areas for condition, possible deformation and damage. If local damage is found, determine if damage extends to adjacent structure.
2-2
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BHT-206-SRM-1 i.
Inspect all areas above for corrosion.
2-4. SPECIAL INSPECTIONS. Helicopters that have been subjected to one or more of the following events must be inspected in accordance with the appropriate special inspection procedure in the applicable Maintenance Manual. a.
Hard Landing
b.
Sudden Stoppage
c.
Overspeed or Overtorque
d.
Lightning Strike
e. Fire Damage. Inspect areas that are suspected of having been subjected to excessive heat for signs of scorching, discoloration/blistering of the paint or other signs of heat damage. If damage is evident, verify that heat treatment has not been affected. Use a non-destructive hardness test. Honeycomb panels suspected of exposure to extreme heat should be removed from service.
2-5. CORROSION TYPES AND IDENTIFICATION. a. DIRECT SURFACE ATTACK. This form of corrosion is generally the least serious of the various forms of corrosion. It is the result of direct reaction of metal surfaces with oxygen in the air and occurs more readily when metal surfaces are exposed to salt-bearing or acid-laden air. Sulphur and chlorine compounds present in some industrial smoke and in aircraft engine exhaust gases also cause direct surface attack. Visual indications of this condition include etching of the surface underneath the area where the corrosion was present. If the metal is aluminum alloy with a coating of pure aluminum (Alclad), the effect on the strength and ductility of the metal is negligible. However, corrosion of the same degree on a non-protected (non-clad) surface may be considered serious as cracks may originate from the pits present. b. GALVANIC OR DISSIMILAR METAL CORROSION. This type of corrosion results from contact between dissimilar metals in the presence of a reactive solution such as salt spray. A true chemical cell is formed. c. PITTING. Pitting corrosion is a special form of galvanic corrosion that is usually localized. It occurs at a point of weakness in the alloy surface caused by mechanical working, faulty heat treatment or contamination that breaks down surface protection. These areas become chemical cells and corrosion products make the situation worse. Deep penetrating attack develops rather than general surface attack. d. INTERGRANULAR CORROSION. This type of corrosion progresses along the grain boundaries of metal alloys. Aluminum alloys which contain large amounts of copper and zinc and some stainless steels are susceptible to this type of corrosion. Hinges are an example of aluminum extrusions which are susceptible to intergranular corrosion. Lack of uniformity in the alloy structure caused by heat treating or localized overheating may result in intergranular corrosion. This corrosion may exist without visible evidence on exterior surfaces, and serious structural weakening may occur without detection. e. STRESS CORROSION. This type of corrosion affects metals that are highly stressed under corrosive conditions. Shrink fit parts and parts subjected to cold working are susceptible to stress corrosion cracking. Stressed metal in contact with stress-free metals may form a chemical cell. Galvanic corrosion occurs along the lines of stress and rapid failure of the part results. f. FATIGUE CORROSION. This type of corrosion is closely related to stress corrosion and appears in metals USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
2-3
BHT-206-SRM-1 subjected to cyclic stress in a corrosive surrounding. A jet engine blade is an example of a part susceptible to fatigue corrosion. The corrosion causes sharp. deep pits which become the origin of cracks that may result in failure of the part. It is difficult to detect this type of corrosion until cracking occurs. g. FRETTING CORROSION. This type of corrosion develops when two loaded surfaces in contact with each other are subject to vibratory motion. The rubbing contact removes small particles of metal from the surface which will oxidize to form abrasive materials. The continuing motion prevents formation of any protective oxide film and, in conjunction with the abrasive formed, creates an area for further corrosion to occur. Fretting is evident at an early stage by surface discoloration and the presence of corrosion products in any lubricant present. Continued fretting will ruin bearing surfaces and destroy critical dimensions. It may be serious enough to cause eventual cracking and fatigue failure of the part. Fretting may be controlled by preventing slippage of two surfaces or by lubricating the surfaces.
2-6. CORROSION CONTROL. If a condition is isolated before damage exceeds the limits stated in section 3-8, proper treatment can control the corrosion. Corrosion control, in general, consists of the following: a.
Removal of finish and corroded material from damaged area.
Corrosion products and damaged metal surfaces must be removed by the mildest method available to prevent additional damage to the part being repaired. Evaluate corroded parts before and after rework to determine the depth of the damage, size, location and the number of affected areas. This will determine the type of repair to be accomplished. b.
Surface treatment of corrosion site.
c.
Structural repair if required. NOTE Due to the limited material thickness available in sheetmetal details used on the Model 206 series helicopters, corrosion damage exceeding 10% of material thickness after cleanup will be treated as mechanical damage. Remove the entire corroded section and repair using the applicable sections of this manual or replace affected part(s).
d.
Refinish of refurbished structure.
e.
Continued monitoring and preventative maintenance of known areas prone to corrosion. NOTE Corrosion prevention may be accomplished using the recommendations contained in Corrosion Control Guide, CSSD-PSE-87-001.
2-4
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BHT-206-SRM-1 2-7. INSPECTION OF HONEYCOMB PANELS. 1. TYPES OF DEFECTS. NOTE Refer to paragraph 2-1 for additional definitions. a. VOID. An area this is supposed to be bonded which has come unbonded (also referred to as bond separation). b.
CRACK.
Fissure or break in the material.
c.
TEAR. Panel facing torn.
d. CORROSION. A chemical action on the panel face or in the core material resulting in discoloration, a layer of oxide or, in advanced states, removal of parent material. e. DENT. A smooth round bottomed depression. f.
PENETRATION.
g. SCRATCH. panel face.
Damage that extends through one or both panel facings into the core material.
Narrow shallow marks or lines resulting from contact with a sharp object across the
2. INSPECTION FOR DEFECTS. a.
Visually inspect panels for corrosion and evidence of mechanical damage.
b. Inspect panels for voids (bonding separation) by light tapping with a suitable sounding device. A light tap over the metal facing of a panel will produce a change to a flat or dead sound when crossing the edge of a void area. Outline the area with a grease pencil. NOTE Areas containing heavy doublers or potting under the surface of the panel skin may sound similar to that of a void. A large void area of regular shape may be an indication of the underlying structure. c. If a void has been detected, inspect panel core. Remove panel facing to expose core in void area by cutting a 0.5 inch diameter hole through skin with a hole saw. Inspect core for condition, paying particular attention to possible corrosion or contamination. CAUTION WHEN INTERNAL CORROSION OR CONTAMINATION (FUEL, OIL, WATER, ETC.) IS DISCOVERED, AFFECTED SECTIONS OF SKIN OR CORE MUST BE COMPLETELY REMOVED. FAILURE TO COMPLY WITH THIS REQUIREMENT MAY RESULT IN FAILURE OF THE REPAIR OR PROGRESSIVE CORE DEGENERATION. NOTE Every effort should be made toward maintaining the highest standards of cleanliness and following the recommended general instruction repair procedures. In all repairs,
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
2-5
BHT-206-SRM-1 the precautionary measures regarding inspection for water, fuel and oil contamination and resultant corrosion must be rigorously followed.
2-8. REQUESTING A REPAIR PROCEDURE. Determine the extent of damage and compare with the applicable section of this manual. Damage falling within the limits of the Structural Repair Manual (SRM) may be repaired using the applicable repair procedure. In cases where the damage is not covered in the SRM, or that the limits of the SRM are exceeded, operators are invited to submit a request for repair through Bell Helicopter Product Support Engineering (PSE) for review. An approved repair procedure will be issued for damages affecting original Bell Helicopter parts that are deemed repairable. To ensure prompt service, a minimum amount of information is required in order to initiate the repair approval process. Refer to the para. 2-9. Requests for repair will be taken on a first-come-first-served basis. While response will vary depending on the workload, allow five (5) to ten (10) working days for the preparation of an approved repair procedure. Bell Helicopter does not offer a service of customized modifications to the aircraft and cannot approve repairs previously accomplished on the aircraft or repairs to parts not procured through sources approved by Bell Helicopter Textron. 2-9. TYPICAL REQUEST FOR REPAIR Applications forwarded to Bell Helicopter for approved repair schemes should contain, as a minimum, the following data. Failure to provide sufficient information may result in delays obtaining Bell Helicopter approval. CHECKLIST ITEMS a. Aircraft nationality and registration number. b. Aircraft model and serial numbers. c. Total airframe time and/or component time in service. d. Correctly and precisely determine the extent of the damage; whether dents, punctures, corrosion, delamination, core contamination, etc. are present. Indicate the precise location of the damage on the aircraft on your sketch. Use the lines drawing references applicable to your model helicopter discussed in section 1. e. Accurately locate and describe any previous repairs in subject area that should be considered in the repair. f. Describe adjacent area to the damage that might affect the way the repair is to be carried out (i.e. Bell Helicopter approved kits, optional installations, antennae, etc.). Any equipment (kits, customizing, BHT/STC/STA) installed which may affect the way the repair will be designed. g. State cause of damage (accident, wear, corrosion,etc.).
2-6
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BHT-206-SRM-1 2-10. GAINING ACCESS FOR REPAIR OR REPLACEMENT OF PARTS. The equipment removed for access to a repair on the airframe is dependant upon the extent and location of the damage. Damages such as localized dent, puncture, crack or corrosion not requiring replacement of the damage part may require removal of adjacent components and equipment necessary to provide access. 1. CONSOLE REPLACEMENT. Remove the following parts and assemblies to gain access for removal of the instrument console. Refer to section 6 for replacement instructions. a.
Battery door, battery door hinge, APU connector, battery and battery relay.
b. Lower windows; left and right hand sides, landing light panel, access panel for pedal controls, battery drain and vent lines and pitot and static lines. c.
Pilot and co-pilot seats, seat bottom panels, access cover between crew seats.
d. Access panels on top and side of console and all equipment on electrical shelf. Ballast weights if installed. e. Control tubes, bellcranks and control cables mounted inside or on console. Cyclic control tube under crew seats. f.
Instruments, instrument piping, wiring, instrument panel face and glareshield.
2. FORWARD FLOOR PANEL (FORWARD LOWER SHELL) REPLACEMENT. Remove the following in preparation for removal of the forward floor panel. Support the helicopter using a holding workaid described in appendix B. CAUTION MODEL 206 SERIES ONLY, DEFUEL HELICOPTER IN ACCORDANCE WITH THE APPLICABLE MAINTENANCE MANUAL INSTRUCTIONS. a.
All seats, carpeting and interior trim.
b.
Tailboom assembly, complete.
c.
Transmission and pylon assembly, complete.
d.
Battery and heavy avionics in instrument console.
e. Controls supports and tubes as necessary. f.
Fuel forward cells and piping on 206L series.
3. AFT FLOOR PANEL (AFT LOWER SHELL) REPLACEMENT. Remove the following in preparation for removal of the aft floor panel. Support the helicopter using a holding workaid describe in appendix B.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
2-7
BHT-206-SRM-1 CAUTION DEFUEL HELICOPTER IN ACCORDANCE WITH THE APPLICABLE MAINTENANCE MANUAL INSTRUCTIONS. a. All seats, carpeting and interior trim. b.
Tailboom assembly, complete.
c.
Transmission and pylon assembly, complete.
d.
Engine assembly, complete.
e.
Oil cooler and blower assembly.
f.
Battery and heavy avionics in instrument console.
g.
Fuel cell.
4. AFT FUSELAGE ASSEMBLY. Removal and installation of the aft fuselage assembly affects the powertrain alignment and must be accomplished with the aid of a Bell Helicopter approved fuselage fixture. For replacement of the engine pan, oil cooler support panel, longerons and fittings, refer to chapter 6. Remove the following:
2-11.
a.
Tailboom assembly, complete.
b.
Engine assembly, complete.
c.
Oil cooler and blower assembly.
d.
Control tubes.
REQUIREMENTS FOR A FUSELAGE ALIGNMENT FIXTURE. Replacement of any one of the following structural sub-assemblies will affect the alignment of the powertrain and must be accomplished with the aid of a Bell helicopter approved fuselage fixture.
2-8
a.
Roof panel and roof beam assembly.
b.
Crew seat or aft cabin bulkhead assemblies.
c.
Engine pan assembly.
d.
Oil cooler blower support assembly.
e.
Aft fuselage assembly.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1 2-12. PREPARING THE HELICOPTER FOR INSTALLATION INTO A BELL HELICOPTER APPROVED FUSELAGE FIXTURE. The following items will require removal prior to fitting your aircraft to a Bell Helicopter approved fuselage fixture. CAUTION DEFUEL HELICOPTER IN ACCORDANCE WITH THE APPLICABLE MAINTENANCE MANUAL. a.
All seats, carpeting and interior trim.
b.
Tailboom assembly, complete.
c.
Transmission, pylon assembly and supports, complete.
d.
Engine assembly and firewalls, complete.
e. Roof mounted cowlings, controls and bellcranks, hydraulic systems, electrical harnesses, control supports and tubes. f.
Battery and heavy avionics in instrument console.
g.
Landing gear.
2-13.
RESTRICTED REPAIR AREAS. Areas of the structure may have restrictions imposed on their repairability for various reasons. These areas are clearly defined by shading in the specific repair procedures for individual parts. Do not attempt repair to areas identified as restricted.
2-14.
MAXIMUM DAMAGE ALLOWANCE. The maximum damage allowed on a part is that limit beyond which the part is deemed not-repairable and must be replaced. This limit may be exceeded by one damage or as a result of the number of individual repairs affecting more than the specified maximum area allowed. These limits are clearly defined in the repair procedures.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
2-9/2-10
FAA APPROVED
REVISION 1
BHT-206-SRM-1
SECTION 3. TYPICAL PROCESSES AND REPAIR PROCEDURES 3-1.
INTRODUCTION.
Many of the repairs for the 206 family of helicopters are a function of the type of construction of the area to be repaired. Some repairs in fact are 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 section provides "typical repairs" for the following types of structure: a.
Riveted structures, both aluminum and titanium,
b.
Resistance (spot) welded assemblies, and
c.
Bonded panels, both fiberglass or aluminum faced, of various core types.
3-2.
COMMON PROCEDURES.
3-2-1. GENERAL. a. Consumable material and standards: The material 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 or a numerical code. This code references a consumable item in table A-1 which can be found in Appendix A of this manual. NOTE Determine material and tooling requirements and ensure materials are at hand before proceeding with any repair. b. Repair material thickness: Unless otherwise indicated in a repair procedure, repair doubler thickness should be one gauge thicker and of the same material specification as the parent material. A filler is usually as thick as the material it replaces and of the same material specification. NOTE For field fabricated parts (ref. APPENDIX A), it is acceptable to substitute 2024-T3 sheet material, two (2) gauges thicker for repair of 7075-T6 aluminum alloy sheetmetal less than 0.040 inch thick. Fit and function of later installations may not be affected. No other material substitution may be accomplished without prior written approval from Product Support Engineering. c. Stop drilling: Stop drilling relieves the stresses in the extremity of a crack. The usual drill size for stop drilling sheet material is a #30 drill. All cracks should be dye penetrant inspected to determine extent of crack prior to stop drilling. Stop drill both ends of crack except when one extremity relieves itself at a lightening hole, a rivet hole or runs to edge of material. Accomplish another dye penetrant inspection after stop drilling to ensure that the end(s) of the crack does not extend past the stop drill hole. If necessary, enlarge hole to a maximum of 0.25 inch diameter.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-1
BHT-206-SRM-1
REVISION 1
FAA APPROVED
d. Refinishing: All repairs should be sealed against moisture intrusion then refinished in accordance with original finish specifications. 3-2-2. APPROVED PROCESSES (PROCESS SHEETS). This section provides instructions to the most common processes used in the repair procedures found in this manual. These process sheets cover Cleaning, Bonding, Wet Lay-up, Edging replacement, Core splice and Insert removal, replacement and installation. Repair procedures throughout this manual make reference to the applicable process sheets in the "required" section. CAUTION 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. The following processses are covered in this section: Para.
Process Sheet.
3-2-3
Removal of Paints and Primers: provides a method of removing commonly used finishes.
3-2-4 Cleaning of Honeycomb Core Cavity: provides a method of cleaning a core cavity in preparation for bonding or core splicing. 3-2-5 Preparation of Panel Bonding Surfaces: provides a method of cleaning faying surfaces of fillers, doublers and panel surfaces in preparation for bonding. 3-2-6 Preparation of Core Plug Prior to Bonding: provides a method of sealing replacement core cells prior to bonding in a panel. 3-2-7
Bonding of Flat Stock: provides a method of bonding flat stock such as fillers and doublers.
3-2-8
Wet Lay-up of Fiberglass: provides a method of laying and bonding fiberglass.
3-2-9
Fiberglass Edging Replacement: provides a method of replacing damaged fiberglass edge material.
3-2-10
Core Splicing: provides a method of bonding a core plug in a panel.
3-2-11
Potted Inserts- General: provides general information regarding inserts installation.
3-2-12
Installation of Inserts: provides a method of installing commonly used inserts in a panel.
3-2-13
Removal of Inserts: provides a method of removing commonly used inserts in a panel.
3-2-14
Chemical Film Application: provides a method of treating bare aluminum against corrosion.
3-2-15 Installation of Rivets Through Thin Panels: provides a method of installing mechanical rivets through thin honeycomb panels.
3-2
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-2-3. REMOVAL OF PAINTS AND PRIMERS. CAUTION
CHEMICAL PAINT STRIPPERS ARE NOT TO BE USED TO REMOVE PAINT FINISHES ON CHEMICAL PAINT STRIPPERS MAY CONTRIBUTE TO BONDED PANELS. CONTAMINATION OF CORE, DETERIORATE ADHESIVE BOND LINE AND FIBERGLASS EDGING. THIS MAY ULTIMATELY CAUSE PANEL TO BE REJECTED. TRICHLOROETHYLENE AND VAPOUR DEGREASERS ARE NOT TO BE USED TO CLEAN OR STRIP SURFACES ADJACENT TO A DAMAGED AREA FOR THE SAME REASONS. METHYL-ETHYL-KEYTONE (MEK) AND ACETONE ARE ACCEPTABLE SOLVENTS FOR REMOVAL OF PAINT FROM ALUMINUM OR FIBERGLASS SKINS AND PANEL EDGING. HOWEVER EXCESSIVE APPLICATION OF MEK OR ACETONE MAY AFFECT ADHESIVES USED IN A BOND. IT IS PREFERABLE TO WIPE SURFACE TO BE STRIPPED USING A MOIST RAG RATHER THAN BY SOAKING. CAUTION BEFORE HANDLING A SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS. KEEP DURING PRODUCT AND ITS VAPOURS AWAY FROM HEAT, SPARKS AND FLAME. APPLICATION AND UNTIL VAPOURS HAVE DISSIPATED, AVOID USING SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES, ETC.
CONVENTIONAL FINISHES (VARNISHES, ALKYD ENAMELS, ZINC CHROMATE PRIMER, ETC.).
A. 1. 2. 3. 4.
B.
Mask around area to be stripped. Brush apply MEK (item S309). Remove lifted paint with a stiff fiber bristle brush. Accomplish final surface cleaning by wiping with a clean cheesecloth moistened with MEK.
ACRYLIC FINISHES. 1. 2.
C.
Mask around area to be stripped. Remove paint by wiping with clean cheesecloth moistened with MEK (item S309).
EPOXY FINISHES. 1. 2.
Mask off area to be stripped. Remove paint using abrasive paper (item S423) of 240 grit or finer. CAUTION DO NOT SAND INTO FIBERGLASS.
3.
Wipe with clean cheesecloth moistened with MEK (item S309). Change cheesecloth often. Repeat operation until all evidence of residue is removed and wipe dry using a clean cheesecloth.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-3
FAA APPROVED
BHT-206-SRM-1 3-2-4. CLEANING OF HONEYCOMB CORE CAVITY. NOTE
Sections of core or skin found contaminated by fuel, oil, water, corrosion or debris must be removed. a.
Remove all loose debris from cavity. CAUTION BEFORE HANDLING A SOLVENT, EXTINGUISH ALL FLAMES AND PILOT LIGHTS. KEEP DURING PRODUCT AND ITS VAPOURS AWAY FROM HEAT, SPARKS AND FLAME. APPLICATION AND UNTIL VAPOURS HAVE DISSIPATED, AVOID USING SPARK PRODUCING ELECTRICAL EQUIPMENT SUCH AS SWITCHES, APPLIANCES, ETC.
b. Flush cavity using MEK (item S309), acetone (item S316) or naphtha (item S305). Dry immediately using dry, filtered compressed air. CAUTION WEAR EYE PROTECTION. AIR PRESSURE NOT TO EXCEED 30 PSI. c. Use clean wrapping paper to protect clean cavity from contamination until ready for subsequent operation.
3-4
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-2-5. PREPARATION OF PANEL BONDING SURFACES. a.
If not already accomplished, remove paint and primer. Refer to paragraph 3-2-3.
b. Sand surfaces to be bonded with non-silicon 240 or finer grit abrasive paper to provide a slightly dull surface finish. NOTE Prepare surface of Composite Bond Material (Bell Specification 150-021-XXX) by first removing peel ply and then sanding lightly with 200 grit paper to remove glaze. No further cleaning of the Composite bond is required. c.
Mask surrounding area. CAUTION 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.
d. Wipe surfaces to be bonded with a cheesecloth moistened with MEK (item S309). Change cheesecloth often. Repeat operation until all evidence of residue is removed and wipe dry using a clean cheesecloth. e.
Remove masking tape and protect clean surfaces from contamination until ready to bond. NOTE Do not handle clean parts with bare hands. The use of clean cotton gloves is recommended when handling parts.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-5
BHT-206-SRM-1
FAA APPROVED
RELEASE FILM ADHESIVE SEALS CELLS
CURED ADHESIVE CURED ADHESIVE CORE PLUG
(SEALED CELLS)
EXISTING PANEL NEW ADHESIVE INSTALL CORE PLUG IN PANEL WITH SEALED FACE OUTWARDS
Figure 3-1. 3-2-6. PREPARATION OF CORE PLUG PRIOR TO BONDING. NOTE The following procedure provides a means of sealing the core plug cells to provide a better bonding surface. In most repairs where repair of the panel is accomplished from one side, only one face of the plug need be sealed. However, if both panel skins are to be replaced, it is recommended that both sides of the core plug be sealed prior to installation in panel. a. Cut core plug, of same material and thickness as original core, to fit the section of core removed while allowing 0.1 to 0.2 inch gap for adhesive at the edges. Align core ribbon in the same direction as that of the existing core in panel. b.
Remove core plug from panel, flush with MEK (item S309) and dry immediately using dry filtered air. CAUTION AIR PRESSURE NOT TO EXCEED 30 PSI.
c. Lay a piece of release film such as Tedlar (item S477) of size equal to or slightly larger than core plug on a flat surface. Apply a uniform film (0.020 to 0.030 inch) of specified adhesive to surface of tedlar. d. Place the core plug on the adhesive covered release agent. Apply pressure of 5 psi maximum on core plug and allow adhesive to cure at room temperature for 24 hrs. Refer to table 3-1 for applicable cure time and temperature. e.
3-6
Trim excess adhesive from core plug.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED f. .
BHT-206-SRM-1
Wrap core plug in clean paper to protect until ready to use. When installing the core plug in panel, position with sealed cells facing away from existing surface as shown on figure 3-1.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-7
BHT-206-SRM-1
REVISION 1
FAA APPROVED
Table 3-1, Adhesive Data. CURING TIME AND TEMPERATURE TYPICAL ADHESIVE
POT LIFE (MINUTES) Room Temp.
HANDLING STRENGTH
FULL CURE Room Temp. (65-85°F)
ALTERNATE FULL CURE
MIXING RATIO RESIN/CATALYST BY WEIGHT
EA934NA
30 to 40
Room temp. @ 24hrs
7 days
180 deg.F @ 1hrs
100/33
EA956
30 to 40
Room temp. @ 24hrs
7 days
180 deg.F @ 1hr
100/58
EA9309NA
30 to 40
Room temp. @ 24hrs
7 days
180 deg.F @ 1hr
100/23
EA932 NA
25 to 30
Room temp. @ 24hrs
7 days
180 deg.F @ 1hr
100/19
Magnolia 6398
40
Room temp. @ 24hrs
7 days
180 deg.F @ 1 hr
100/27
Magnolia 6367
50
Room temp. @ 24hrs
7 days
180 deg.F @ 1hr
100/44
3-2-7. BONDING OF FLAT STOCK (FILLERS AND DOUBLERS). NOTE Do not handle clean parts with bare hands. The use of clean cotton gloves is recommended when handling parts. a.
Coat prepared surface(s) of the doubler and filler to be bonded with a thin layer (approximately 0.006 inch thick) of specified adhesive. NOTES Unless the repair procedure specifies otherwise: when using composite bond material (bell specification 150-021-xxx) on one face only (B code), position the composite bond surface to the existing structure. The sequence when bonding is to work from the existing structure outwards. Do not prebond fillers to doublers. When bonding bare 301 cres steel, surface to be bonded must be etched prior to bonding.
b.
Bond components in their respective positions. Install fasteners, if required, while adhesive is wet.
c.
Apply pressure of 0.5 to 1.0 psi. to surface of repair.
d.
Remove adhesive squeeze-out.
e.
Allow adhesive to cure. Refer to Table 3-1 for applicable cure time and temperature.
f.
Inspect for voids or unbonded areas. Voids shall not exceed 10% of total bonded area. No one void shall exceed 0.25 square inch.
3-8
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-2-8. WET LAY-UP OF FIBERGLASS.
a. Thoroughly impregnate fiberglass cloth with specified adhesive. The use of a glass roller or rubber squeegee will assist in attaining thorough wetting of the fibers. b. Position individual plies on repair. Use a roller or squeegee to remove trapped air and force excess adhesive out. Repeat operation for additional plies. c. Apply a uniform pressure of 0.5 to 1.0 psi to surface of repair and allow adhesive to cure a minimum of 24 hours at room temperature. d.
Remove surface irregularities using 380 to 400 grit non-silicon abrasive paper (S423). CAUTION DO NOT SAND INTO FIBERGLASS.
e.
Trim wet lay-up laminate edges as required and seal (item S392) against moisture intrusion.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-9
FAA APPROVED
BHT-206-SRM-1
A END OF DAMAGED EDGING AREA
2.00 INCHES OVERLAPPING IN ONE INCH INCREMENTS
1 00 IN 0.25 IN
CUT LINE NOT TO EXCEED THIS CORNER
-EOP OF NEW EDGING END OF EDGING CUT OUT1.00
IN.
EOP OF EXISTING EDGING SECTION
2.00 INCHES OVERLAPPING IN ONE INCH A INCREMENTS
A-A
Figure 3-2 3-2-9. FIBERGLASS EDGING REPLACEMENT. a.
Remove section of edging to be replaced plus a minimum of 0.25 inch beyond damage on all sides.
b. Inspect exposed surfaces of core for evidence of damage, corrosion or contamination. If damage exists, determine extent of damage and repair using the applicable section of this manual before proceeding with edging replacement. c.
Clean panel in preparation for bonding in accordance with paragraph 3-2-5.
d. Prepare replacement number of edging plies required, dimensioned to extend beyond section of edging removed in one inch increments. The number of plies in the replacement edging is determined by the number of plies removed plus one ply. e.
3-10
Position and bond new edging per paragraph 3-2-8.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
REVISION 1
FAA APPROVED
BHT-206-SRM-1
NEW CORE PLUG FILL GAP WITH BONDING ADHESIVE
EOPOFNEW CORE PLUG
0.10 TO 0.20 INCH GAP
EOP OF CORE CUTOUT
0.25 INCH OF EXPOSED CORE
SECTION A-A
EXISTIN
EOP OF SKIN
CORE (REF)
CUT OUT
Figure 3-3. 3-2-10.
CORE SPLICING.
a.
Prepare core plug(s) per paragraph 3-2-6.
b.
Clean panel in preparation for bonding in accordance with paragraph 3-2-5.
c. Apply a light film of specified adhesive to faying surfaces of parent structure or doubler/filler and core plug. e.
Bond core plug in place and fill gap between plug and parent core with specified adhesive.
g.
Proceed with remaining portion of repair while adhesive is wet.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-11
BHT-206-SRM-1 3-2-11.
REVISION 1
FAA APPROVED
POTTED INSERTS-GENERAL.
a. INSERT INSTALLATION TOLERANCES. Except when type 80-007, 80-011 and 80-013 are used, the faces of the inserts covered in this section shall be installed flush with the surface of the panel within + /0.020 inch. When these inserts are installed in rigidized skins, the face(s) shall be flush with the face of the adjacent diamond pattern of the rigidized skin within +/-0.020 inch. b. PRELIMINARY REQUIREMENTS. preparation and bonding operations. c.
PREPARATION OF INSERTS.
Cleanliness is to be carefully controlled through all phases of the
Prior to bonding in panel, inserts shall be prepared as follows:
1. Solvent clean by soaking in MEK (item S309). 2. Air dry for a minimum of 15 minutes. 3. Handle clean dry inserts with clean cotton gloves. Do not handle clean parts with bare hands. CAUTION 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. d.
HONEYCOMB PANEL PREPARATION. 1. Bonded honeycomb panels shall have the proper size hole drilled through the skin(s) into the core. Refer to table 3-2, sheets 1 thru 5, for applicable hole size(s). CAUTION DO NOT OVERHEAT SKIN SURFACE(S) DURING THE DRILLING OPERATION. 2. In the case of blind type inserts, the depth of the hole in the core will be such that it does not bottom or damage opposite skin. 3. The hole in the honeycomb core will be enlarged to a minimum of 0.50 inch larger than diameter of insert (0.25 inch undercut from edge of cutout). NOTE Clean cutting tools to prevent contamination of the insert hole during drilling and routing operations. 4. Clean drilled holes of debris resulting from cutting operation using dry filtered compressed air. No further surface preparation of the hole is required. CAUTION WEAR EYE PROTECTION. AIR PRESSURE NOT TO EXCEED 30 PSI. 5.
3-12
Protect holes from contamination until ready to install inserts.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED 3-2-12.
BHT-206-SRM-1
INSTALLATION OF POTTED INSERTS (REFER TO FIGURES 3-4 THROUGH 3-6 AND TABLE 3-2).
This page intentionally left blank.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-13
FAA APPROVED
BHT-206-SRM-1
TWO 0.070/0.093 INJECTION HOLES
INSTALLATION
0.125INCH TYP
180 APART 45
HOLE SIZE (TABLE 3.2)
POTTING ADHESIVE POTTING ADHESIVE INJECT ADHESIVE INTO ONE HOLE UNTIL IT COMES OUT OF OTHER HOLE.
INSTALLATION OF BLIND INSERT, BELL STANDARD 80-004 AND 80-005.
Figure 3-4. 3-2-12.
INSTALLATION OF POTTED INSERTS.
APPLICATION A. INSTALLATION OF BLIND TYPE POTTED INSERT, BELL STANDARD 80-004 AND 80-005. A.1 Determine position and mark center of insert on panel. Drill hole of appropriate size, for insert type and diameter required, through one skin only. Refer to table 3-2. A.2 Undercut core a minimum of 0.25 inch or as specified in the applicable repair procedure. A.3 If insert used does not have provisions in head for injecting adhesive, drill two opposed 0.078/0.093 inch injection holes through outer skin, at 45 deg. angle to the surface of the panel and 0.065 to 0.125 inch from edge of hole. A.4 Deburr holes. Remove debris and loose material from cavity and surface of panel. A.5 Fill cavity one-third full of adhesive. Press insert in place and secure in position with masking tape. Punch through tape at both injection hole locations. A.6 Using a small tipped syringe inject a slow steady flow of adhesive through one hole. Continue injecting until a steady flow of adhesive emerges out of the opposite hole. A.7 Remove excess adhesive and allow to cure. Refer to table 3-1 for adhesive curing time and temperature. A.8 Remove tape. Sand surface of insert and adjacent area with 360 grit or finer abrasive cloth.
3-14
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1 SYRINGE TWO 0.078/0.093 INJECTION HOLES 180 APART ADHESIVE INSTALLATION HOLE SIZE (TABLE 3-2)
ADHESIVE ADHESIVE
UNDERCUT CORE 0.25 MIN TYP
INJECT ADHESIVE INTO ONE HOLE UNTIL IT COMES OUT OF OTHER HOLE
INSTALLATION OF PLUG AND SLEEVE INSERT, BELL STANDARD 80-011 AND 80-013.
Figure 3-5. 3-2-12.
INSTALLATION OF POTTED INSERTS.
APPLICATION B.
Installation of plug and sleeve type potted inserts, Bell Standard 80-011 and 80-013.
B.1 Determine position and mark center of insert on panel. Drill hole of appropriate size, for insert type and diameter required, through panel. Refer to table 3-2. B.2 Undercut core a minimum of 0.25 inch or as specified in repair procedure. B.3 If insert used does not have provisions in head for injecting adhesive, drill two opposed 0.078/0.093 inch injection holes through outer skin, at 45 deg. angle to the surface of the panel and 0.065 to 0.125 inch from edge of hole.
B.4 Deburr holes. Remove debris and loose material from cavity and surface of panel. B.5 Coat mating flange and lip of plug and sleeve with a thin layer of adhesive. Also coat lip underneath heads of plug and sleeve with adhesive. Insert plug and sleeve from their respective side of panel and press halves together. Secure insert halves with a nut, washers and bolt. NOTE A light coat of grease on surfaces of nut, washers and bolt is recommended to prevent their adhering to insert and panel. B.6 Using a small tipped syringe inject a slow steady flow of adhesive through one hole. Continue injecting until a steady flow of adhesive is seen to emerge out of the opposite hole.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-15
BHT-206-SRM-1
FAA APPROVED
B.7 Clean excess adhesive around insert. Allow adhesive to cure. Refer to table 3-1 for adhesive curing time and temperature. B.8 Remove bolt and ensure that insert bore and threads are free from adhesive. Sand surface of insert and adjacent area, on both sides of panel, with 360 grit or finer abrasive cloth.
3-16
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
TWO 0.078/0.093 INJECTION HOLES HOLE DIA H (TABLE 3-2)
180°APART-
0.13
INCH TYP
UNDERCUT CORE 0.25 MIN TYP
TABLE
LENGTH
(TABLE 3-2)
OF INSERT
EQUALS THICKNESS OF CORE PLUS THICKNESS OF SKIN
INSTALLATION OF DOME HEAD STYLE INSERT, BELL STANDARD 80-007.
Figure 3-6. 3-2-12.
INSTALLATION OF POTTED INSERTS.
APPLICATION C.
Installation of dome head type potted inserts, Bell Standard 80-007.
C.1 Determine position and mark center of insert on panel. Drill different diameter holes (K and H) of appropriate size, for insert type and required diameter, through both skins. Refer to table 3-2. C.2 Undercut core a minimum of 0.25 inch more than larger hole or as specified in repair procedure. C.3 Drill two opposed 0.078/0.093 inch injection holes through outer skin, at 45 deg. angle to the surface of the panel and 0.065 to 0.125 inch from edge of hole. C.4 Deburr holes. Remove debris and loose material from cavity and surface of panel. C.5 Coat small extremity and underneath head of insert with a thin layer of adhesive. Push insert into panel from one side and press against opposite skin. Secure insert to panel with bolt and washer. NOTE A light coat of grease on surfaces of washer and bolt is recommended to prevent their adhering to insert or panel. C.6 Using a small tipped syringe inject a slow steady flow of adhesive through one hole. injecting until a steady flow of adhesive emerges out of the opposite hole.
Continue
C.7 Clean excess adhesive around insert. Allow adhesive to cure. Refer to table 3-1 for adhesive curing time and temperature.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-17
BHT-206-SRM-1
FAA APPROVED
C.8 Remove bolt and ensure that insert threads are free of adhesive. Sand surface of insert and adjacent area, on both sides of panel, with 360 grit or finer abrasive cloth.
3-18
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED 3-2-13.
BHT-206-SRM-1
REMOVAL OF POTTED INSERTS: (REFER TO FIGURES 3-7 THRU 3-9 AND TABLE 3-2).
The following tools are required to accomplish insert removal. 1) Piloted hole saw or counterbore of same diameter or slightly smaller than hole in face panel. Pilot drill should be of same size as threaded portion of insert. Refer to table 3-2 for applicable insert installation hole size. 2)
Routing tool for removal of potting. NOTE Routing tool may be locally fabricated from a section of allen key, long end fastened to drill motor.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-19
BHT-206-SRM-1
FAA APPROVED
INSTALLATION HOLE SIZE REF TABLE 3-2
1.
DRILL BODY OF INSERT
2.
CAUTION DO NOT ENLARGE EXISTING DIAMETER HOLE THROUGH FACE OF PANEL.
REMOVE REMAINING SECTIONS OF INSERT. REMOVE POTTING TO EXPOSE CORE.
Figure 3-7. 3-2-13.
REMOVAL OF POTTED INSERTS.
APPLICATION A. Removal of potted blind-type insert, Bell Standard 80-004 and 80-005. A.1 Drill out insert body using hole saw or counterbore. CAUTION DO NOT DRILL OR ROUT THROUGH OPPOSITE SKIN IN PANEL. CAUTION DO NOT OVERHEAT SKIN SURFACES. A.2 Remove remains of insert with use of a pick or by routing. Remove existing potting to expose core. A.3 Remove all debris and loose material. A.4 Prepare insert cavity for replacement insert in accordance with paragraph 3-2-11 d. A.5 Install blind insert in accordance with paragraph 3-2-12, application A.
3-20
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
1.
DRILL BODY OF INSERT.
2.
REMOVE REMAINING SECTION OF INSERT.
3.
REMOVE POTTING TO EXPOSE CORE.
CAUTION DO NOT ENLARGE EXISTING DIAMETER HOLES THROUGH FACES OF PANEL.
Figure 3-8. 3-2-13.
REMOVAL OF POTTED INSERTS.
APPLICATION B. Removal of plug and sleeve type potted insert, Bell Standard 80-011 and 80-013. B.1 Drill out insert body using hole saw or counterbore. CAUTION DO NOT OVERHEAT SKIN SURFACES. B.2 Remove remains of insert with use of a pick or by routing. Remove existing potting to expose core. B.3 Remove all debris and loose material. B.4 Prepare insert cavity for replacement insert in accordance with paragraph 3-2-11d. B.5 Install plug and sleeve in accordance with paragraph 3-2-12, application B.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-21
BHT-206-SRM-1
FAA APPROVED
TABLE 3-2
1.
PILOT DRILLTHROUGH HEAD FROM OPPOSITE SIDE OF PANEL.
2.
DRILL BODY OF INSERT.
3.
CAUTION DO NOT ENLARGE EXISTING DIAMETER HOLES THROUGH FACES OF PANEL.
REMOVE REMAINING SECTION OF INSERT. REMOVE POTTING TO EXPOSE CORE.
Figure 3-9. 3-2-13.
REMOVAL OF POTTED INSERTS.
APPLICATION C.
Removal of dome head type potted insert, Bell Standard 80-007.
C.1 Using drill of same diameter as threaded portion and using threads in insert to guide the drill, drill through insert head. C.2 From head side, using hole drilled in C-1 as pilot, drill out insert body using hole saw or counterbore.
CAUTION DO NOT DRILL OR ROUT THROUGH OPPOSITE SKIN IN PANEL. CAUTION DO NOT OVERHEAT SKIN SURFACES. C.3 Remove remains of insert with use of a pick or by routing. Remove existing potting to expose core. C.4 Remove all debris and loose material. C.5 Prepare insert cavity for replacement insert accordance with paragraph 3-2-11 d. C.6 Install dome head insert in accordance with paragraph 3-2-12, application C.
3-22
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
INFORMATION 0.093 DIA. POTTING HOLES OR SLOTS 180 APART
ONLY Insert Basic Code:
80-004
- 3
B
12
TAPERED AREA SERRATED Basic No BASE Thread Size Skin Thickness Length
THREAD L
THREAD LOCK
TABLE 3-2. INSERT INSTALLATION DATA. SHEET 1 OF 5.
POTTED STANDARD SELF LOCKING INSERT. BELL SPECIFICATION 80-004. 1st DASH No
THREAD SIZE
(thread size)
HEAD DIAMETER
INSTALLATION
2nd DASH No
(A)
HOLE SIZE
(insert length)
LENGTH (L)
-1
0.1640-32 UNJC
0.500
0.469/0.474
-4
0.220
-2
0.1900-32 UNJF
0.500
0.469/0.474
-5
0.285
-3
0.2500-28 UNJF
0.562
0.531/0.534
-6
0.335
-4
0.3125-24 UNJF
0.687
0.656/0.659
-7
0.395
-5
0.3750-24 UNJF
0.812
0.781/0.784
-8
0.455
-6
0.1120-40 UNJC
0.375
0.344/0.349
-10
0.565
-7
0.1380-32 UNJC
0.437
0.406/0.411
-12
0.690
-14
0.815
-16
0.935
-18
1.055
NOTE
-20
1.185
SKIN PANEL THICKNESS
CODE
0.010/0.019 0.020/0.029
B
THIS DATA NOT INTENDED FOR DESIGN PURPOSES. REFER TO APPLICABLE REPAIR
-22
1.305
0.030/0.039
C
PROCEDURE OR ORIGINAL DATA FOR CORRECT REPLACEMENT INSERT TYPE.
-24
1.430
0.040/0.049
D
0.050/0.059
E
0.060/0.069
F
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-23
FAA APPROVED
BHT-206-SRM-1
0.078 TO 0.093 DIA. OR SLOT POTTING HOLES .180° APART
INFORMATION ONLY Insert Basic Code: 80-005
TAPERED AREA SERRATED OR THREADED
3B
12
S
Basic Code
BASE CONFIGURATION
Thread Size Skin Thickness
OPTIONAL
Length Steel Housing
-
THREAD LOCK
L THREAD
TABLE 3-2. INSERTS INSTALLATION DATA, SHEET 2 OF 5.
POTTED FLOATING & SELF LOCKING INSERT. BELL SPECIFICATION 80-005. 1t DASH No (thread size)
THREAD SIZE
-1
HEAD DIAMETER (A)
INSTALLATIO N HOLE SIZE
2nd DASH No (insert length)
0.1640-32 UNJC
0.593
0.562/0.565
-6
0.340
-2
0.1900-32 UNJF
0.593
0.562/0.565
-8
0.455
-3
0.2500-28 UNJF
0.718
0.687/0.690
-10
0.565
-4
0.3125-24 UNJF
0.843
0.812/0.815
-12
0.690
-5
0.3750-24 UNJF
0.968
0.937/0.940
-14
0.815
-6
0.1120-40 UNJC
0.593
0.562/0.565
-16
0.935
-7
0.1380-32 UNJC
0.593
0.562/0.565
-18
1.060
-20
1.185
PANEL SKIN THICKNESS
LENGTH (L) (L)
____
CODE
0.010/0.019 0.020/0.029
B
0.030/0.039
C
0.040/0.049
D
0.050/0.059
E
0.060/0.069
F
3-24
THIS DATA NOT INTENDED FOR DESIGN PURPOSES. REFER TO APPLICABLE REPAIR PROCEDURE OR ORIGINAL DATA FOR CORRECT REPLACEMENT INSERT TYPE
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
HOLE K
HOLE H INFORMATION ONLY Insert Basic Code
80-007
- 06
- 3
Basic Code
INSTL DETAIL Length L
Material - = Carbon Steel C = Corr. Res Steel
TMIN
0.060
THREAD HEAD DIA LOCK STRAIGHT COARSE KNURL TABLE 3-2. INSERT INSTALLATION DATA. SHEET 3 OF 5.
DOMED HEAD INSERT. BELL SPECIFICATION 80-007.
1st DASH No (thread size)
THREAD TYPE & SIZE
HEAD
DIAMEER
INSTALLATION HOLE SIZES K H
2nd DASH No (insert length)
-1
0.1380-32 UNJC
0.625
0.323/0.329
0.145/0.150
-06
0.375
-2
0.1640-32 UNJC
0.625
0.323/0.329
0.168/0.173
-07
0.437
-3
0.1900-32 UNJF
0.687
0.386/0.392
0.193/0.198
-08
0.500
-4
0.2500-28 UNJF
0.687
0.386/0.392
0.256/0.262
-09
0.562
-4A
0.2500-28UNJF
0.750
0.453/0.459
0.256/0.262
-10
0.625
-5
0.3125-24 UNJF
0.750
0.453/0.459
0.316/0.322
-11
0.687
-12
0.750
-13
0.812
-14
0.875
-15
0.937
16
1.000
-17
1.062
-18
1.125
-19
1.187
-20
1.250
THIS DATA NOT INTENDED FOR DESIGN PURPOSES. REFER TO APPLICABLE REPAIR PROCEDURE OR ORIGINAL DATA FOR CORRECT REPLACEMENT INSERT TYPE
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-25
BHT-206-SRM-1
FAA APPROVED
INFORMATION ONLY Grommet Basic Code 80-011 - S
INSTALLATION HOLE SIZE PLUG
4
F 06
0
Basic Code P = Plug S = Sleeve Diameter Head Style
SLEEVE
SLEEVE
/
D = Flush Surface F = Protruding
Length (ref) Material
0
=
Aluminum
C
TABLE 3-2. INSERT INSTALLATION DATA, SHEET 4 OF 5.
GROMMET INSERT, NON-THREADED. BELL SPECIFICATION 80-011.
DASH No
THRU OPENING DIAMETER THRU OPENING DIAMETER
HEAD DIAMETER (C)
INSTALLATION HOLE SIZE
-4
0.113-0.119
0.375
0.228
-5
0.133-0.139
0.500
0.290
-6
0.139-0.147
0.500
0.290
-7
0.165-0.172
0.500
0.290
-8
0.165-0.172
0.500
0.290
-9
0.191-0.197
0.625
0.323
-10
0.191-0.197
0.625
0.323
-11
0.253-0.260
0.750
0.390
-12
0.253-0.260
0.750
0.390
-13
0.286-0.293
0.812
0.421
-14
0.315-0.321
0.875
0.484
-15
0.315-0.321
0.875
0.484
-16
0.378-0.384
1.000
0.640
(A)
THIS DATA NOT INTENDED FOR DESIGN PURPOSES. REFER TO APPLICABLE REPAIR PROCEDURE OR ORIGINAL DATA FOR CORRECT REPLACEMENT INSERT TYPE
3-26
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
INFORMATION
C
ONLY Insert Basic Code
80-013 - S
3
F
13
- 9
Basic Code SLEEVE PLUG SLEEVE
S = Sleeve P = Plug
(THREADED)
Size
PLUG (THREADED)
Head Style
D = Flush F Protruding
Length (ref) THIN PANEL TYPE
REGULAR TYPE
Material
0 = Aluminum 9= Steel
TABLE 3-2. INSERT INSTALLATION DATA, SHEET 5 OF 5.
PLUG AND SLEEVE INSERT, THREADED.
BELL SPECIFICATION 80-013.
DASH No
THREAD TYPE AND SIZE
DIAMETER (A)
HEAD DIAMETER (C)
INSTALLATION HOLE SIZE (B)
-1
0.1380-32 UNJC
0.139-0.147
0.500
0.323
-2
0.1640-32 UNJC
0.165-0.172
0.500
0.323
-3
0.1900-32 UNJF
0.191-0.197
0.625
0.358
-4
0.2500-28 UNJF
0.253-0.260
0.750
0.421
-5
0.3125-24 UNJF
0.315-0.321
0.875
0.515
THIS DATA NOT INTENDED FOR DESIGN PURPOSES. REFER TO APPLICABLE REPAIR PROCEDURE OR ORIGINAL DATA FOR CORRECT REPLACEMENT INSERT TYPE
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-27
BHT-206-SRM-1 3-2-14.
FAA APPROVED
CHEMICAL FILM APPLICATION.
The chemical film treatment of aluminum alloy is a chromate conversion coating that increases the material's resistance to corrosion and provides a base for organic finishes on all aluminum alloys. The resulting chem-film coating offers no significant resistance to abrasion. 1. PREPARATION. Protect the area surrounding the area to receive treatment by masking. Use masking tape and wrapping paper. 2.
CLEANING. a.
Remove oil and grease using a clean cloth moistened with MEK (item S309)
b. Scrub the area to be treated to bare metal using a nylon web abrasive pad (item S407) or fine aluminum wool (item S422). c. Scrub area with abrasive pad and cleaning compound (item S318) mixed to 10 to 20 percent by volume in clean water. d.
Rinse the surface thoroughly using clean water.
e. Inspect for a water break free surface. Repeat steps c. and d. as required to obtain a clean surface. 3.
DRYING. Allow surface to air dry or force dry using clean filtered compressed air and clean, dry cloth.
4.
Mix chemical film solution (item S100) as follows: To 256 parts of distilled or demineralized water, add 6 parts of chemical film material (item S121) and 1 part of nitric acid (item S432).
5. TREATMENT. Apply chem-film solution prepared in 4. liberally to the area to be treated. Keep the area wet with the solution for a period of one to three minutes then rinse thoroughly with clean water and allow to dry. NOTE The treatment should provide a golden irridescent to brown colour of continuous and uniform appearance. Streaks and mottled areas caused by surface condition of the aluminum are acceptable provided that there is chemical film coverage in these areas. 6.
3-28
PRIMER. Apply primer (item S204) to treated surfaces.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
RIVET
01.25
HOLE THROUGH OUTER SKIN ONLY.
UNDERCUT CORE MIN
0.125
DOUBLER
UNDERCUT CORE. FILL WITH ADHESIVE. ALLOW TO CURE AND SAND TO CONTOUR. 1 40-001 -1
HOLE THROUGH POTTINGAND INNER SKIN OF PANEL.
WASHER
Figure 3-10. 3-2-15.
INSTALLATION OF RIVETS THROUGH THIN HONEYCOMB PANELS.
APPLICATION. Installation of riveted doubler on thin panel (0.38 inch thick core or less) where installation of a blind fastener would damage the opposite skin. RESTRICTION.
Valid as specified by BHT Product Support Engineering only.
REQUIRED. 1. Fasteners: Rivets M7885/2-4, grip length to suit and washers 140-001-1; quantity as required. 2. Adhesive (item S317) 3. Abrasive paper (item S423) 4. Sealant (item S392) 5. Process sheet: Cleaning (refer to Para. 3-2-5) PROCEDURE. 1. Gain access to both sides of panel to be repaired. 2. Determine location of fastener holes from doubler or part to be installed to panel. 3. Drill hole through one face of panel only and core. Refer to Table 3-4 for correct hole size.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-29
BHT-206-SRM-1
FAA APPROVED NOTE Do not drill through opposite skin in panel at this time.
4.
Under cut core in panel a minimum of 0.125 in.
5.
Clean cavity of debris and loose material in preparation for bonding.
6.
Fill cavity with specified adhesive and allow to cure at room temperature for 24 hours.
7.
Fair adhesive to contour of panel.
8. Locate doubler to be installed on panel face and drill hole of correct size through adhesive and opposite skin in panel. Clean debris and loose material. 9.
Prepare panel surface in preparation for bonding. NOTE Installation of washer under shop head of rivet is not required when doublers are installed on both faces of panel.
10.
3-30
Bond and rivet detail in position. Install rivets with washer under shop head while adhesive is wet.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-3.
RIVETED STRUCTURES.
3-3-1.
RIVET REPLACEMENT, GENERAL.
Generaly, replacement fasteners should be of the same type and size as the fastener removed except when hole damage, repair alteration or field conditions (e.g. lack of heat treat capabilities) require a change within these guidelines. a.
Use of next-size solid rivet. It is acceptable to install same type oversize solid rivet providing sufficient material edge distance exists to accomodate the new fastener. Typically, the repair procedures in this manual require 0.38 inch edge distance. NOTE Except when Product Support Engineering or the repair procedure allows otherwise, universal head rivets require as a minimum 2D edge distance (ED). Countersunk rivets require as a minimum 2.5D edge distance be maintained.
b.
Edge distance. Fastener edge distance (ED) is a measure of the proximity of the fastener to edge of material. ED is measured from the center of the rivet to the edge of the material.
c.
Hole preparation. Inspect rivet holes to the requirements of table 3-4 to determine correct size fastener to be used. Use a go-no-go gauge or other suitable test instrument to determine hole size. Never attempt to install rivets in hole(s) that are too big for the fastener; rather, drill to the next available size rivet.
d.
Rivet lay-out. Where possible, lay out location of fasteners using the same pattern and spacing used in the immediate surrounding structure. When adding universal head rivets or increasing rivet size, a minimum edge distance (ED) of two times rivet diameter (2D) and minimum rivet spacing of four times rivet diameter (4D) must be maintained unless otherwise specified. In the case of countersunk rivets, minimum ED and minimum spacing to be 2.5D and 5D respectively unless otherwise specified. Rivet spacing should be between 0.9 to 1.3 inch apart, 1.12 inch being the desired value.
e.
Rivet identification. Solid rivets are easily recognisable by the distinct markings on the rivet head. Refer to table 3-6 for information on solid rivet identification. Blind rivets and close tolerance fasteners have their part number embossed on the rivet head.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-31
FAA APPROVED
BHT-206-SRM-1 3-3-2.
RIVET SUBSTITUTION.
Solid rivets are the prefered fastener for most repair conditions. However, blind rivets may be substituted for solid rivets if one or more of the conditions below prevail. NOTE Refer to table 3-3 for allowable rivet alternates and substitution data. 1.
Blind rivets are specified in the repair procedure, or
Access is not possible from both sides of the structure (in the case of sheet metal structure) for 2. installation of a solid rivet or fasteners attach to a composite panel.
3-32
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1 TABLE 3-3 SUBSTITUTION OF RIVETS (sheet 1 of 2).
APPLICATION A. PROTRUDING HEAD RIVETS. Use of this table is restricted to the following conditions: 1. Protruding head fasteners replacement- this table applies to rivet diameters of 1/8 inch (-4) thru 3/16 inch (-6). 2. Minimum edge distance (2D for protruding) must be respected by the new fastener. 3. Meets the thickness restrictions specified in note 1 below. Existing rivet type
Alternate Blind Rivet Cherry Max (oversize)
Bulbed Cherry Lock
MS20470AD(D)-(L)
M7885/6-(D)-(L) (note 1) or M7885/8-(D)-(L).
NAS1738MW(D)-(L)
MS20470B(D)-(L)
M7885/6-(D)-(L) (note 1) or M7885/8-(D)-(L).
NAS1738MW(D)-(L)
MS20470D(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
MS20470DD(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
MS20470E(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
MS20600AD(D)-(L)
M7885/6-(D)-(L)
NAS1738MW(D)-(L)
MS20600B(D)-(L)
M7885/6-(D)-(L)
NAS1738MW(D)-(L)
MS20600M(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
MS20615-(D)MP(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
NAS1398B-(D)-(L)
M7885/6-(D)-(L)
NAS1738MW(D)-(L)
NAS1398M-(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
NAS1738B-(D)-(L)
M7885/8-(D)-(L)
NAS1738MW(D)-(L)
NAS1738M-(D)-(L)
M7885/8-(D)-(L)
N/A
NOTE. 1. The following sheet thickness restrictions exist when using M7885/6 rivets. The numbers express the thinner of two skins fastened: 1/8 inch (-4) diameter rivets require minimum sheet thickness of 0.040 inch; in the case of 5/32 inch (-5) diameter rivets, a min. sheet thickness of 0.050 inch is required; for 3/16 inch (-6) diameter rivets this restriction is 0.063 inch. LEGEND.(D): indicates rivet diameter in 1/32 inch increments while (L): indicates length in 1/16 inch increments.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-33
FAA APPROVED
REVISION 1
BHT-206-SRM-1
TABLE 3-3 SUBSTITUTION OF RIVETS (sheet 2 of 2). APPLICATION B. FLUSH HEAD RIVETS. Use of this table is restricted to the following conditions: 1. Flush head fasteners replacement- this table applies to rivet diameters of 1/8 inch (-4) thru 3/16 inch (-6). 3. Minimum edge distance ( 2.5D for countersunk) must be respected by the new fastener. 4. Minimum sheet thickness required to accomodate countersink is as follows: 0.050 inch in the case of 1/8 inch (-4) rivet, 0.063 inch for 5/32 (-5) rivet and 0.071 if a 3/16 inch diameter (-6) rivet is used. Alternate Blind Rivet Existing rivet type
Bulbed Cherry Lock
MS20426AD(D)-(L)
M7885/7-(D)-(L).
NAS1739MW(D)-(L)
MS20426B(D)-(L)
M7885/7-(D)-(L).
NAS1739MW(D)-(L)
MS20426D(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
MS20426DD(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
MS20426E(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
MS20427M(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
MS20601 AD(D)-(L)
M7885/7-(D)-(L)
NAS1739MW(D)-(L)
MS20601 B(D)-(L)
M7885/7-(D)-(L)
NAS1739MW(D)-(L)
MS2060MM(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
NAS1399B-(D)-(L)
M7885/7-(D)-(L)
NAS1739MW(D)-(L)
NAS1399M-(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
NAS1739B-(D)-(L)
M7885/9-(D)-(L)
NAS1739MW(D)-(L)
NAS1739M-(D)-(L)
M7885/9-(D)-(L)
N/A
LEGEND. (D):
3-34
Cherry Max (oversize)
indicates rivet diameter in 1/32 inch increments while (L): indicates length in 1/16 inch increments.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
TABLE 3-4. DRILL and HOLE SIZE LIMITS for SOLID and BLIND RIVET INSTALLATION NOMINAL SIZE SOLID OR BLIND RIVET
OVERSIZE BLIND RIVET
Rivet Diameter
Drill Size
Hole Size Min/Max
Rivet Diameter
Drill Size
Hole Size Min/Max
1/8 (-4) 5/32 (-5) 3/16 (-6) 1/4 (-8)
#30 #20 #10 F
.129/.132 .160/.164 .192/.196 .256/.261
1/8 (-4) 5/32 (-5) 3/16 (-6) 1/4 (-8)
#27 #16 #5
.143/.146 .176/.180 .205/.209 .271/.275
|
TABLE 3-5. EFFECTIVE GRIP LENGTHS FOR BLIND RIVET INSTALLATION. MATERIAL THICKNESS (inches) RIVET GRIP
PROTRUDING HEAD
FLUSH HEAD
MINIMUM
MAXIMUM
MINIMUM
MAXIMUM
-1
(Note 1)
1/16 (0.062)
n/a
n/a
-2
1/16 (0.063)
1/8 (0.125)
(Note 2)
1/8 (0.125)
-3
1/8 (0.126)
3/16 (0.187)
1/8 (0.126)
3/16 (0.187)
-4
3/16 (0.188)
1/4 (0.250)
3/16 (0.188)
1/4 (0.250)
-5
1/4 (0.251)
5/16 (0.312)
1/4 (0.251)
5/16 (0.312)
-6
5/16 (0.313)
3/8 (0.375)
5/16 (0.313)
3/8 (0.375)
-7
3/8 (0.376)
7/16 (0.437)
3/8 (0.376)
7/16 (0.437)
-8
7/16 (0.438)
1/2 (0.500)
7/16 (0.438)
1/2 (0.500)
-9 (Note 3)
1/2 (0.501)
9/16 (0.562)
1/2 (0.501)
9/16 (0.562)
-10
9/16 (0.563)
5/8 (0.625)
9/16 (0.563)
5/8 (0.625)
-11 (Note 4)
5/8 (0.626)
11/16 (0.687)
5/8 (0.626)
11/16 (0.687)
-12 (Note 5)
11/16 (0.688)
3/4 (0.750)
11/16 (0.688)
3/4 (0.750)
Notes: 1. Do not use protruding head rivets M7885/all-4-01 below 0.025 sheet thickness, M7885/all-5-01 below 0.031 sheet thickness, M7885/all-6-01 below 0.037 sheet thickness. 2. Do not use flush head rivets M7885/all-4-02 (except /7 and/9) below 0.063 sheet thickness, M7885/all-5-02 below 0.065 sheet thickness, M7885/all-6-02 below 0.080 sheet thickness. In case of M7885/7-4-02 this value is 0.045, for M7885/7-5-02 this value is 0.063 and forM7885/7-6-02 this value is 0.073. 3. Maximum grip length for M7885/all-4 rivets. 4. Maximum grip length for M7885/all-5 rivets. 5. Maximum grip length for M7885/all-6 rivets.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-35
BHT-206-SRM-1
REVISION 1
FAA APPROVED
TABLE 3-6. SOLID RIVET IDENTIFICATION
HEAD CODE
Raised Dimpled
COUNTERSUNK RIVET
MS20426D(D)-(L) MS20426AD(D)-(L)
Raised Cross
MS20426B(D)-(L)
Raised Double
MS20426DD(D)-(L)
Dash
Green Raised
110-174-(D)-(L)
Raised
MS20426E(D)-(L)
Double Dimple
MS20427M(D)-(L)
Plain
MS20427F(D)-(L)
HEAD CODE
PROTRUDING HEAD
Raised
MS20470D(D)-(L)
Dimpled
MS20470AD(D)-(L)
Raised Cross
MS20470B(D)-(L)
Raised Double
MS2047 DD(D)-(L)
Dash Green, Raised
110-175:(D)-(L)
Raised
MS20470E(D)-(L)
Double Dimple
MS20615(D)MP(L)
Recessed Triangle
MS20613(D)P(L)
LEGEND: (D) expresses diameter in 1/32 inch increments while (L) expresses length in 1/16 inch increments
3-36
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-3-3.
REPLACEMENT OF CLOSE TOLERANCE FASTENERS, GENERAL.
When replacing a part that is attached with close tolerance (Hi-Lok and Hi-Shear) fasteners, it is essential that hole(s) drilled in the replacement part exactly match existing hole(s) in structure or mating part. This must be accomplished without damaging or enlarging the existing hole(s). 1.
Secure replacement part in position by clamping or other suitable means.
2. Place a drill bushing in one of the existing holes in mating part and drill a pilot hole through replacement part. Bushing should fit snuggly in existing hole to ensure correct positionning of the pilot hole. CAUTION ENSURE THAT EXISTING HOLES ARE NOT DAMAGED DURING THE PROCESS. 3. Remove drill bushing. Drill and ream hole to required size. Deburr and inspect for correct size and possible hole damage. NOTE Inspect Hi-Lok holes to the requirements of table 3-9 to determine correct size fastener to be used. Use a go-no-go gauge or other suitable test instrument to determine hole size. Never install a Hi-Lok in a hole that is too big. 4. Insert Hi-Lok in hole and temporarily install collar. Do not apply sufficient torque to fracture nut drive from collar as it will have to be removed. 5.
Install remaining Hi-Loks in pattern following steps 1 thru 4.
6.
Remove all fasteners. Reinspect all holes for correct size and possible damage.
7. Prepare parts for final assembly. Install all Hi-Lok fasteners in pattern using unreduced epoxy polyamide primer (item S204). Torque collars to full torque required to fracture the nut drive.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-37
BHT-206-SRM-1
FAA APPROVED TABLE 3-7 SUBSTITUTION OF SHEAR-TYPE FASTENERS.
The following considerations must be given when replacing a 5/32 inch diameter Hi-Shear fastener with same diameter Hi-Lok fastener due to differences in installation hole size and depth of countersink. In all other cases, Hi-Shear fasteners may be replaced with Hi-Lok fasteners in accordance with Table 3-8, Note 1. Fastener Type; protruding/countersunk
Installation Hole Size (inch)
Depth of Countersink (inch)
NAS1054-5-(L) /NAS1055-5-(L)
0.1560/0.1580
0.042 (NAS1055-5-(L) only)
100-048-5-(L) /100-076-5-(L)
0.1635/0.1655
0.042 (100-047-5-(L) only)
For replacement with oversize fastener refer to Tables 3-8 and 3-9. TABLE 3-8 OVERSIZE PART NUMBER EQUIVALENCIES FOR HI-LOK FASTENERS
TENSION HEAD FASTENERS (Note 1). Standard size Protruding/Flush Pin (Collar)
1/64 inch Oversize Protruding/Flush Pin (Collar)
1/32 inch Oversize Protruding/Flush Pin (Collar)
STAINLESS STEEL PIN Bell Specification No.
100-049 (30-017) 100-059 (30-017)
N/A
N/A
Hi-Lok No.
HL-646 (HL-73) HL-647 (HL73)
HL-36 (HL-273) HL-37 (HL-273)
HL-136 (HL-373) HL-137 (HL-373)
ALLOY STEEL PIN Bell Specification No.
100-048 (30-015) 100-047 (30-015)
N/A
N/A
Hi-Lok No.
HL-20 (HL-86) HL-21 (HL-86)
HL-64 (HL-87) HL-65 (HL-87)
HL-220 (HL-93) HL-221 (HL-93)
SHEAR HEAD FASTENERS (Note 1). Standard size Protruding/Flush Pin (Collar)
1/64 inch Oversize Protruding/Flush Pin (Collar)
1/32 inch Oversize Protruding/Flush Pin (Collar)
ALLOY STEEL PIN Bell Specification No.
100-076 (30-055)
N/A
N/A
Hi-Lok No.
HL-19 (HL-94)
HL-63 (HL-94)
HL219 (HL294)
NOTES: 1. The 100-048 and 100-076 fasteners listed in table 3-8 are direct replacements for NAS1054 and NAS1055 respectively, provided the limitations given in table 3-7 are met. NAS1054 and NAS1055 will in no case be used to replace a Hi-Lok fastener.
3-38
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1 TABLE 3-9. SELECTED HOLE DIAMETERS FOR HI-LOK FASTENERS
SHANK DIAMETER (inch)
REAMED HOLE DIAMETER (inch) Min./Max.
5/32 Standard (Note 1)
0.1635 /0.1655
3/16 standard
0.1895/0.1915
3/16 + 1/64 O.S.
0.2026/0.2046
3/16 + 1/32 O.S.
0.2182/0.2202
1/4 Standard
0.2495/0.2515
1/4 + 1/64 O.S.
0.2651/0.2671
1/4 + 1/32 O.S.
0.2807 /0.2827
NOTES: 1. No oversize fastener exists for -5 ( 5/32 inch diameter) Hi-Lok: use a -6 (3/16 inch diameter) Hi-Lok.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-39
FAA APPROVED
BHT-206-SRM-1 3-4.
RIVET PATTERN DISCREPANCIES.
There are cases in repairs when conditions occur that deviate from established standards for fastener installation. Some deviations are required to allow use of parts that otherwise would have to be replaced. This section covers such cases and their respective corrective action. CAUTION
THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS ON SECTIONS 4 AND 5 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE TAILBOOM, STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTIONS 4 AND 5. a. NEGLIGIBLE DAMAGE. 1. Rivet short edge distance b. REPAIRABLE DAMAGE. 1. Rivet short edge distance 2. Cracked rivet holes 3. Elongated, mismatched or oversize rivet holes
3-40
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
-
NO PART OF RIVET HEAD MAY EXTEND PAST EDGE OF PART
ONE RIVET DIAMETER MIN CONDITION "X" NOT PERMITTED
Figure 3-11. 3-4-1.
RIVET SHORT EDGE DISTANCE.
APPLICATION A. Replacement of part, previously installed with short edge distance. RESTRICTIONS A. 1.
Replacement part rivets shall have equal or greater edge distance than removed part.
2.
Maximum of one rivet in any five consecutive rivets may be affected.
3.
Condition "X" is not permitted.
4.
Not applicable to close tolerance fasteners (e.g. Hi-Lok, Hi-Shear, etc.).
5.
Not applicable to control tubes.
REQUIRED A. Same type rivets as used in immediate area. Refer to Table 3-6 for rivet identification. PROCEDURE A. Use as is.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-41
BHT-206-SRM-1
FAA APPROVED
ADDED PATTERN (ONE MORE THAN O.S. USED) EXISTING PATTERN 4D MIN OVERSIZE RIVETS PITCH (TYP) EXISTING PITCH
OVERSIZE RIVET SHORT EDGE DISTANCE -
2D EDGE DISTANCE MIN
Figure 3-12 3-4-1.
RIVET SHORT EDGE DISTANCE.
APPLICATION B. For short edge distance in exterior skin or interior structure, resulting from installation of oversize rivet. RESTRICTIONS B. 1. Not to exceed 10 consecutive holes or 50 percent of rivet pattern, whichever is least. 2. Added rivets to maintain standard edge distance in all components of structure to which they are common. Minimum distance between any two rivets to be four times the larger rivet diameter (4D). 3. Added rivets are not to interfere with later installations. 4. Not applicable to bonded panels, close tolerance fasteners or control tubes. REQUIRED B. 1. Same type rivets as used in immediate surrounding area. 2. Primer (item S204)
3-42
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
PROCEDURE B. 1. Drill for applicable rivet size, refer to Table 3-4. Interpitch added rivets of the same size as the original rivets in immediate area, maintaining edge distance shown. 2. Deburr all holes. 3. Apply primer to bare metal surfaces. Allow to dry.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-43
BHT-206-SRM-1
FAA APPROVED
0.38 TYPICAL
THICKNESS (T) PLUS ONE GAUGE 0.38 TYPICAL
KNESS (T)
RADIUS TO MATCH ANGLEE THICKNESS (T)
CRACKS IN SKINS
CRACKS IN FLANGES
Figure 3-13. 3-4-2. CRACKED RIVET HOLES. APPLICATION. For cracks in skins or flanged members extending from rivet hole to edge of material. RESTRICTIONS. 1. Maximum of one cracked hole allowed in any five consecutive holes. Not applicable to close tolerance fastener holes, or to cracks entering holes at a tangent. 2. Cracked holes may not exceed 5 percent of pattern. 3. Maximum material thickness 0.032 inch. 4. Minimum edge distance twice rivet diameter (2D) for protruding head rivet and (2.5D) for flush head rivet must be respected. REQUIRED. 1. Doubler of like composite bond material and one gauge thicker of sufficient length to extend a minimum of two rivets plus required edge distance beyond crack. Width to extend from bend tangent to edge of part and maintain proper edge distance. 2. Rivets of same type and size as installed in immediate area. Refer to Table 3-6 for rivet identification. 3. Adhesive (item S317) 4. Primer (Item S204).
3-44
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
5. Sealant (Item S392). 6. Process sheets:
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7)
PROCEDURE. 1. Position doubler about damaged rivet hole. Pick up at least two rivets on each side of damaged hole. 2. Drill holes and deburr. Refer to Table 3-4. 3. Prepare faying surfaces of doubler and structure for bonding. 4. Bond doubler in position and install rivets while adhesive is wet. 4. Prime bare metal surfaces. Allow to dry 5. Seal edges of repair.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-45
BHT-206-SRM-1
FAA APPROVED
ORIGINAL I RIVET EDGE DISTANCE 1.5D MINIMUM FOR 1/16" O.S. RIVET
I I
Figure 3-14. 3-4-3. ELONGATED, MISMATCHED, OR OVERSIZED RIVET HOLES. APPLICATION. For elongated, mismatched, or oversize holes that can be cleaned up by drilling through and using oversize rivets. RESTRICTIONS. 1. Maximum of 50 percent of the rivet pattern when using rivets 1/32 inch dia. larger than existing rivet pattern, for a maximum of five consecutive holes. NOTE If both rivet edge distance and pitch can be maintained in accordance with paragraphs 3-3-1 a and on all affected holes, through all components, then there are no restrictions to the number of fastener holes affected. 2. Maximum of 25 percent of the rivet pattern when using rivets 1/16 inch dia. larger than existing rivet pattern, for a maximum of five consecutive holes. 3. Valid for "AD", "B" and "Cherry" aluminum rivets only. 4. Minimum edge distance one and one-half oversized rivet diameter (1.5D) is allowed on replacement rivets. REQUIRED. 1. Rivets, 1/32 inch or 1/16 inch larger diameter, same type as those removed. Refer to Table 3-6 for rivet identification. 2. Primer (item S204). 3-46
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
PROCEDURE. 1. Drill through for the applicable rivet size and deburr holes. Refer to Table 3-4. 2. Install larger rivet. 3. Apply primer to bare metal surfaces. Allow to dry.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-47
BHT-206-SRM-1
A
>R
FAA APPROVED
+ 1/2 RIVET HEAD DIA
"T"
RADIUS BLOCK
THICKNESS
T=R
VIEW A-A
A
A
Figure 3-15. 3-4-4. MISLOCATED HOLES IN FLANGES. APPLICATION. For mislocated hole where fastener would normally ride or interfere with radius or dimension "A" is not met. RESTRICTION. 1. Not applicable to angles with radius greater than 0.1 88 inch in thickness without prior written approval from Product Support Engineering (PSE). 2. Rivet hole cannot fall in radius of angle. REQUIRED. 1. Radius block of like material. Minimum thickness (T) to be same as angle bend radius (R.) Length sufficient to pick up minimum of one rivet or fastener plus 0.38 inch each side of mislocated hole. Radius one corner to match angle radius (R). NOTE It is acceptable to use a metal radius block on fiberglass material. 2. Rivets or fasteners of same type as installed in immediate area. identification.
Refer to Table 3-6 for rivet
3. Adhesive (Item S317). 4. Primer (Item S204). 3-48
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 5. Sealant (Item S392). 6. Process sheets.
Cleaning (refer Para. 3-2-5) Bonding (refer Para. 3-2-7)
PROCEDURE. 1. Position radius block over mislocated hole. Pick up a minimum of one fastener on each side of mislocated hole. 2. Drill and deburr mounting holes. 3. Prepare faying surfaces of radius block and flange for bonding. 4. Bond and rivet block in position. Install rivets while adhesive is wet. 5. Apply primer to bare metal. Allow to dry. 6. Seal edges of repair. 7. Refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-49
FAA APPROVED
BHT-206-SRM-1 3-5.
ANGLE SECTION REPAIRS.
Angles are typically used as stiffeners or frame flanges. Refer to Repairs 3-4-1 through 3-4-4 for repairs. NOTE When repair specifies use of adhesive, the mating surfaces of the adhesive bond line must be prepared in accordance with Paragraph 3.2.7. CAUTION THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS IN SECTIONS 4 AND 5 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE TAILBOOM, STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTIONS 4 AND 5. a.
NEGLIGIBLE DAMAGE. 1. N/A
b.
REPAIRABLE DAMAGE. 1. Flange cracks 2. Radius cracks
3-50
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
1/3 FLANGE WIDTH MAX AFTER CLEAN-UP FILL HOLE WITH RIVET STEM RADIUS CUTOUT TO 0.13 N. MIN
1.0 INCH MAX
DOUBLER ONE GAUGE THICKER THAN ANGLE
Figure 3-16. 3-5-1. FLANGE DAMAGE TO ANGLES. APPLICATION A. For damage in angle member, not exceeding 1/3 of flange depth and 1.0 inch in length. RESTRICTIONS A. 1. Cleanup cannot impinge on rivet hole. 2. Repair will not interfere with subsequent installations. 3. Cleanup not to exceed limits shown. 4. Maximum flange thickness of 0.040. 5. Not applicable where close tolerance fasteners are used. REQUIRED A. 1. Doubler; from composite bond material, same material and one gauge thicker, of length sufficient to pick up 3 rivets on either side of damage plus 0.38 inch. 2. Rivets of same type and material and one size larger than rivets used in surrounding structure. 3. Adhesive (Item S317). 4. Primer (Item S204). 5. Sealant (Item S392). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-51
BHT-206-SRM-1
FAA APPROVED
6. Process sheets: Cleaning (refer to para. 3-2-5) Bonding (refer to para. 3-2-7) PROCEDURE A. 1. Blend out damage to a minimum radius of 0.13 inches. 2. Position strap doubler and drill for required fasteners. Deburr all holes. 3. Prepare faying surfaces of doubler and angle for bonding. 4. Bond and rivet doubler. Install rivets while adhesive is wet. 5. Apply primer to bare metal. Allow to dry. 6. Seal edges of repair.
3-52
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1
BHT-206-SRM-1
0.50 MIN EXISTING ANGLE
DAMAGED SECTION REMOVED
ATTACHING STRUCTURE
withrivets instald
FILLER ANGLE
0.38 MIN ANGLE DOUBLER TO OVERLAP DAMAGED AREA BY FOUR RIVETS ON EACH END
Figure 3-17. 3-5-1. EXTENSIVE FLANGE DAMAGE TO ANGLE. APPLICATION B. 3-5-1).
Extensively damaged or broken angle flange. (Greater damage than application repair
RESTRICTIONS B. 1. Limited to angles with rivets installed in one leg only and less than 0.032 in. thick. 2. Not applicable to extruded angles, or when close tolerance fasteners are used in pattern. 3. Vertical leg of original angle must be sufficiently wide to provide 1.5D inch minimum edge distance for added rivets. 4. Cannot interfere with later installations. REQUIRED B. 1. Filler: angle of same material and thickness as original to replace removed damaged section. 2. Doubler: angle of same material and one gauge thicker than original angle. As an alternate, composite bond material may be used. Length of doubler sufficient to overlap damaged area by a minimum of 4 rivets plus 0.38 inch in each flange, on each end of discrepant area. 3. Adhesive (Item S317). 4. Primer (Item S204).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-53
BHT-206-SRM-1
FAA APPROVED
5. Oversize rivets, M7885/6, or one size larger than original. 6. Sealant (Item S392). 7. Process sheets:
Cleaning (refer Para. 3-2-5) Bonding (refer Para. 3-2-7)
PROCEDURE B. 1. Cut out damaged section of angle. CAUTION DO NOT DAMAGE SURROUNDING STRUCTURE. 2. Replace damaged section with filler angle. 3. Position doubler on angle, maintaining minimum of 1.5D edge distance. Pick up a minimum of four rivets on each side of damage area on all flanges. Drill holes, remove filler and doubler, and deburr holes. 4. Clean filler angle, doubler and structure in preparation for bonding. 5. Bond filler angle to structure. Bond and rivet doubler to existing angle and filler angle. Install rivets while adhesive is wet. 6. Apply primer to bare metal surfaces. Allow to dry. 7. Seal edges of repair.
3-54
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
CUT OFF AT END OF JOGGLE
EXISTING ANGLE WITH DAMAGED END REMOVED 1.5D MIN
1.5D MIN
ANGLE DOUBLER FILLERS (IF REQD)
0.38
TYP
TO MATCH JOGGLE
EXISTING ANGLE WITH DAMAGED END REMOVED
FILLERS (IF REQUIRED) (ATTACH WITH TWO RIVETS MIN)
EXISTING RIVET PATTERN ANGLE DOUBLER
NO MORE THAN
STRAIGHT ANGLE
4 END RIVETS
JOGGLED ANGLE
Figure 3-18. 3-5-2. EXTENSIVE DAMAGE TO END OF ANGLES. APPLICATION.
Extensive damage to angle end, requiring a splice.
RESTRICTIONS. 1. Repair cannot interfere with later installations. 2. Not applicable when close tolerance fasteners are in splice pattern. 3. Not applicable to angles greater than 0.032 inch thick. REQUIRED. 1. Doubler: Angle section from same material and thickness as existing angle and of sufficient length to overlap existing angle a minimum of 5 existing rivets plus 0.38 inch in each flange. As an alternate composite bond material of same material and thickness may be used. 2. Filler blocks, if required, from same material as existing angle and of sufficient thickness to match existing joggle. As an alternate composite bond material of same material and thickness may be used. 3. Adhesive (Item S317). 4. Primer (Item S204). 5. Oversize rivets, M7885/6, or one size larger than original. 6. Sealant (Item S392). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-55
BHT-206-SRM-1 7. Process sheets:
FAA APPROVED Cleaning (refer Para. 3-2-5) Bonding (refer Para. 3-2-7)
PROCEDURE. 1. Cut out damaged angle end. CAUTION DO NOT DAMAGE SURROUNDING STRUCTURE. NOTE Install filler if required to restore in-plane conditions. Fillers to be bonded and riveted using a minimum of two rivets on end of straight angle and all splice rivets on end of joggled angle 2. Position replacement angle section and filler if required. Ensure there is minimum 1.5D edge distance for rivets. Drill holes, remove parts and deburr holes. 3. Clean all parts in preparation for bonding. 4. Bond and rivet angle doubler and filler. Install rivets while adhesive is wet. 5. Apply primer to bare metal surfaces. Allow to dry. 6. Seal edges of repair.
3-56
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
MAXIMUM LENGTH OF CRACK 25X FLANGE THICKNESS
0.38 MIN TYPICAL T=FLANGE THICKNESS
1.5D MIN ANGLE DOUBLER
0.38 MIN TYPICAL
DAMAGED MEMBER
Figure 3-19. 3-5-3.
LENGTHWISE CRACK IN BEND RADIUS OF ANGLES.
APPLICATION.
For crack 25 T or less along length of bend radius of formed member.
RESTRICTIONS. 1. Not applicable where the repair doubler would be closer than 3 inches from the end of the member, end of fitting, or end of splice fitting. 2. Not applicable to extruded shapes or if close tolerance fasteners are in pattern. 3. Maximum allowable thickness of damaged member to be 0.032 inch. 4. Crack length not to exceed 25 x flange thickness. 5. Cannot interfere with later installations. REQUIRED. 1. Doubler: Angle from composite bond material of same material and one gauge thicker than original. Length sufficient to overlap a minimum of two rivets plus 0.38 inch beyond the end of the crack on each side of damage. 2. Adhesive:(ltem S317). 3. Rivets of same type and size as installed in immediate area. 4. Primer (Item S204). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-57
FAA APPROVED
BHT-206-SRM-1 5. Sealant (Item S392). 6. Process sheets:
Cleaning (refer Para. 3-2-5) Bonding (refer Para. 3-2-7)
PROCEDURE. 1. Dye penetrant inspect to determine extent of crack. 2. Stop drill each end of crack with No. 30 drill. Carry out dye penetrant inspection as in step 1 to ensure crack does not protrude past stop drill. 3. Position angle doubler about damaged area. Pick up a minimum of two rivets beyond the end of the crack on each side of damaged area, on each flange, as well as all the rivets along the length of the crack. 4. Drill and deburr all holes in both flanges. 5. Clean parts in preparation for bonding. 6. Bond and rivet doubler angle to existing angle.
Install rivets while adhesive is wet.
7. Apply primer to bare metal surfaces. Allow to dry. 8. Seal edges of repair.
3-58
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
8T MAX CLEAN-UP T=FLANGE THICKNESS
Figure 3-20. 3-5-4. CRACK IN DOUBLE-FORMED FLANGE. APPLICATION A.
Repair of crack in the corner of a double-formed flange.
RESTRICTION A. Length of crack limited to 8 times the damaged material thickness (8T); cracks exceeding 8T can be repaired using Application B. REQUIRED A. 1. Primer (Item S204). PROCEDURE A. 1. Dye penetrant inspect to determine extent of crack(s). 2. Radius out crack with 1/8 inch diameter cutter. Cleanup not to exceed end of crack by more than the material thickness (1T). 3. Apply primer to bare metal. Allow to dry.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-59
FAA APPROVED
BHT-206-SRM-1
T = FLANGE THICKNESS
STOP DRILLED CRACK
DOUBLER
DAMAGED MEMBER 0.38 TYP
CORNER CRACKS IN DOUBLE FORMED FLANGE
Figure 3-21. 3-5-4. CRACK IN DOUBLE-FORMED FLANGE. APPLICATION B.
Crack in corner of double-formed flange exceeding limitations stated in Application A.
RESTRICTION B. Length of crack cannot exceed 25 times the material thickness (25T) or 1.0 inch whichever is the least. REQUIRED B. 1. Doubler: double formed flange from composite bond material of same material and one gauge thicker than original part. Length of flanges to allow a minimum of 0.38 inch edge distance for a minimum of two rivets. 2. Rivets of same type as those installed in immediate area. 3. Adhesive (item S317). 4. Primer (Item S204). 5. Sealant (Item S392). 6. Process sheet: Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) PROCEDURE B. 1. Dye penetrant inspect to determine extent of crack(s).
3-60
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
2. Stop drill end of crack with No. 30 drill. Carry out dye penetrant inspection as in step 1 to ensure crack does not protrude past stop drill. NOTE If no rivets are present in the damaged area, install type common to that detail. Size and pitch to be determined by surrounding structure. 3. Position repair doubler. Drill for two rivets in each flange and one rivet each side of stop drilled hole. Space rivets equally while maintaining standard edge distance. 4. Deburr holes. 5. Prepare faying surfaces of doubler and structure for bonding. 6. Bond and rivet doubler in position. Install rivets while adhesive is wet. 7. Apply primer to bare metal. Allow to dry. 8. Seal edges of repair.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-61
BHT-206-SRM-1
3-6.
FAA APPROVED
WEB AND SKIN REPAIRS.
Webs and skins are suceptible to a variety of damage, including cracks and punctures. Refer to repairs 3-61 through 3-6-5. CAUTION THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS IN SECTIONS 4 AND 5 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE TAILBOOM, STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTIONS 4 AND 5. a.
NEGLIGIBLE DAMAGE.
1. N/A b.
REPAIRABLE DAMAGE.
1. Oil canning 2. Cracks in webs and skins 3. Small punctures 4. Cracks at lightening holes
3-62
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
0.38 MIN TYP
OIL CAN CONDITION
0.9TO 1.3 1.12 DESIRED
/
T=THICKNESS 0.032 INCH MAX
ANGLE
8D TYP ATTACH CLIP TO SUPPORTING STRUCTURE
0.62 CLIP MINIMUM
Figure 3-22. 3-6-1. "OIL CAN" CONDITION IN SKIN OR WEB. APPLICATION.
For cases where unsupported skins or webs create false contour or "Oil Can" effect.
RESTRICTIONS. 1. Maximum skin thickness of 0.032 inch. 2. Damage cannot extend into stiffeners or other supporting structure. 3. Not applicable to the tailboom. REQUIRED. 1. Angle of same material and thickness as adjacent stringers. Length sufficient to span between bulkheads or stringers. Flange width sufficient to allow a minimum of 0.38 inch edge distance for rivets. 2. Clip of same material and thickness as stringers. Length and width to allow a minimum of two rivets per flange plus 0.38 inch edge distance. 3. Rivets of same type as those installed in the immediate area. 4. Primer (Item S204). 5. Sealant (Item S392).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-63
FAA APPROVED
BHT-206-SRM-1 PROCEDURE. 1. Position angle and clips through center of "oil can" area. 2. Drill for rivets. Remove parts and deburr holes. 3. Apply primer to bare metal. Allow to dry. 4. Rivet angle and clips in position. 5. Seal edges of repair.
3-64
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
EXISTING BULKHEAD
OIL CAN CONDITION
0.38 TYP
'
DOUBLER
1.5 MIN 0.9 T0 1.3 1.12 DESIRED
Figure 3-23. 3-6-2. "OIL CAN" CONDITION IN BULKHEAD. APPLICATION.
For cases where unsupported web of bulkhead creates false contour or "oil can".
RESTRICTIONS. 1. Minimum rivet edge distance and spacing are available with ends of doubler extending past damaged area a minimum of 1.5 inches. 2. Minimum edge distance is available to accommodate a single row of rivets around each lightening hole through which the doubler extends. 3. Flanges of bulkhead are not buckled. 4. Cannot interfere with later installations. 5. No more than two lightening holes in any one bulkhead can be covered by repair doubler. REQUIRED. 1. Doubler: using composite bond material of same material and one gauge thicker than bulkhead. Doubler to match contour of bulkhead and overlap oil can condition by a minimum of 1.5 inches each end plus 0.38 inch edge distance. If installed on flanged side of lightening holes, provide clearance for hole and bulkhead flange radii in doubler. 2. Rivets MS20470AD4, grip length and quantity to suit. 3. Adhesive (Item S317). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-65
BHT-206-SRM-1
FAA APPROVED
4. Primer (Item S204). 5. Sealant (Item S392). 6. Process sheets:
Cleaning (refer Para. 3-2-5) Bonding (refer Para. 3-2-7)
PROCEDURE. 1. Remove from bulkhead, clips, rivets, decals, etc. which may interfere with doubler installation. 2. Position doubler on bulkhead. Drill 0.125 inch diameter rivet pattern including single row around lightening holes. Maintain 8D spacing and 0.38 inch edge distance, remove doubler and deburr all holes. 3. Prepare faying surfaces of doubler and structure for bonding. 4. Rivet and bond doubler in position. Install rivets while adhesive is wet. 5. Apply primer to bare metal surfaces. Allow to dry. 6. Reinstall any details removed in step 1. above. 7. Seal edges of repair.
3-66
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
RADIUS TO 0.10 INCH MINIMUM
ONE HALF OF RIVET EDGE DISTANCE MAXIMUM
0.38 MIN TYPICAL
2D MINIMUM 2D MINIMUM
BLEND OVER A DISTANCE OFLENGTH 3 TO 4 TIMES TEAR/CRACK
Figure 3-24. 3-6-3. EDGE TEARS AND CRACKS IN SKINS AND WEBS. APPLICATION A.
Repair of edge tear or crack in skin or web.
RESTRICTIONS A. 1. Length of tear/crack maximum of one-half of rivet edge distance; for edge tear exceeding 1/2 edge distance, refer to Applications B and C. 2. One crack or tear allowed in any 6.0 inch length. 3. Cleanup to provide minimum of 2 times edge distance from nearest rivet. REQUIRED A. 1. Primer (item S204). PROCEDURE A. 1. Dye penetrant inspect to determine extent of tear/crack. 2. Blend out damage over a distance of 3 to 4 times the length of tear/crack. Provide a minimum of 0.10 inch radius. 3. Apply primer to bare metal surfaces. Allow to dry.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-67
BHT-206-SRM-1
FAA APPROVED
ADD ONE 3/32 RIVET ON EACH SIDE OF DRILLED HOLE 0.142 INCH MIN. EDGE DISTANCE
STOP DRILLED CRACK
Figure 3-25. 3-6-3. EDGE TEARS AND CRACKS IN SKINS AND WEBS. APPLICATION B. Edge tear/crack in skin or web not exceeding three times rivet diameter (3D) in length. For tear exceeding 3D in length, refer to application C. RESTRICTIONS B. WARNING USE OF THIS REPAIR REQUIRES PRIOR APPROVAL FROM PRODUCT SUPPORT ENGINEERING (PSE). 1. Length of tear/crack maximum of 3 times the rivet diameter (3D). 2. Multiple cracks or tears must be separated by a minimum of 6.0 inches. 3. Not applicable to skin/web thickness greater than 0.032 inch. REQUIRED B. 1. Rivets of 3/32 inch diameter and of same type as installed in immediate area. Refer to table 3-1 for rivet identification. 2. Primer (Item S204).
3-68
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
PROCEDURE B. CAUTION ENSURE THAT SUBSTRUCTURE IS PROTECTED FROM DAMAGE WHILE DRILLING. 1. Dye penetrant inspect to determine extent of tear/crack. 2. Stop drill end of crack with No. 30 drill. Carry out dye penetrant inspection as in step 1 to ensure crack does not protrude past stop drill. 3. Drill for two 3/32 inch diameter rivets through web and faying structure, one each side of stop drill hole. Maintain minimum of 0.38 inch edge distance. Deburr holes. 4. Apply primer to bare metal surfaces. Allow to dry. 5. Install rivets.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-69
BHT-206-SRM-1
FAA APPROVED
DOUBLER
STOP DRILL END OF CRACK
0.9
TO 1.3 1.12 DESIRED
MIN
RIVET CRACK
0.38 TYP DOUBLER STOP DRILL END OF CRACK
1.12
0.9 TO 1.3 DESIRED
4.0 TWO ROWS TYPICAL
0.50
3 RIVETS MIN
MIN
RIVET
CRACK
Figure 3-26. 3-70
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
3-6-3. EDGE TEARS AND CRACKS IN SKINS AND WEBS. APPLICATION C.
Edge crack or tear greater than Application B (refer to Fig. 3-26).
RESTRICTIONS C. 1. Tear or crack not to exceed 4.00 inches long. 2. Maximum skin thickness of 0.032 inch. 3. Maximum of one tear per skin at a minimum of 60 degrees to edge. 4. Repair cannot interfere with later installations. REQUIRED C. 1. Doubler: from composite bond material of same material and one gauge thicker than existing skin. Length and width to allow minimum of two full rows of rivets on all sides of crack plus 0.38 inch edge distance. 2. Sealant (Item S392). 3. Adhesive (Item S317). 4. Primer (Item S204). 5. Rivets of same type as those installed in immediate area. Refer to Table 3-1. 6. Process sheets.
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7)
PROCEDURE C. 1. Dye penetrant inspect to determine extent of crack. CAUTION TAKE CARE TO INSURE THAT SUBSTRUCTURE IS PROTECTED FROM DAMAGE WHILE DRILLING. 2. Stop drill end of crack with a No. 30 drill. Carry out dye penetrant inspection as in step 1 to ensure crack does not extend past stop drill. 3. Fabricate repair doubler. edge distance.
Doubler must pick up three or more rivets each side of crack plus required
4. Remove existing rivets in repair area and record type and size rivets removed. Refer to Table 3-1 for rivet identification. 5. Backdrill existing rivet pattern and add rivets in any area where rivets are more than 2.5 inches apart. Rivet spacing to be 0.9 to 1.3 inch apart, 1.12 in. prefered. 6. Remove doubler and deburr all holes. 7. Clean parts in preparation for bonding. USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-71
FAA APPROVED
BHT-206-SRM-1 8. Bond and rivet doubler in position. Install rivets while adhesive is wet. 9. Apply primer to bare metal. Allow to dry. 10. Seal edges of repair.
3-72
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
2.00 N. MIN.
FROM EDGE OF STRUCTURE -
0.38 MIN TYP
DAMAGED
AREA CLEANUP +
0.5 MIN RADIUS
0.9 TO 1.3 1.12 DESIRED
TYP
0.38 TYP
0.38 MIN
RADIUS
TYP
Figure 3-27. 3-6-4. DOUBLER REPAIR OF SKINS AND WEBS. APPLICATION.
For repair of damaged aluminum webs and skins.
RESTRICTIONS. 1. Edge of damage must be a minimum of 2.00 inches from nearest adjacent structure, after clean-up. 2. Minimum distance between repair cleanup areas is 4.0 inches. 3. Damage limited to a maximum of 20 percent of the skin area after cleanup. 4. Overlapping of repair doublers is not permitted. REQUIRED. 1. Doubler: from composite bond material of same material and one gauge thicker than damaged material. Length and width to allow a minimum of two rows of rivets plus 0.38 inch edge distance. Minimum thickness of doubler to be 0.020 inch. Doubler may be installed on either side of skin. NOTE Filler is required if doing internal doubler repair and optional with external doublers. Use a minimum of two rivets to attach filler one inch or less in diameter; three rivets for filler greater than one inch but less than 1.5 inches in diameter; and four rivets for filler 1.5 inches to 2.0 inches in diameter. 2. Filler same gauge and material as web, size to match damage cutout (optional). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-73
BHT-206-SRM-1
FAA APPROVED
3. Sealant (Item S392). 4. Rivets of same type as those installed in the immediate area and 0.125 inch minimum diameter. Refer to Table 3-1 for rivet identification. 5. Primer (Item S204). 6. Adhesive (Item S317). 7. Process sheets.
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7)
PROCEDURE. 1. Rout out damaged area, removing minimum material, to provide a smooth rectangular, oval or circular cleanup area. Minimum cleanup corner radius is 0.5 inch. 2. Deburr and refinish damaged area. 3. Position doubler over repair area. Drill and deburr rivet holes. 4. Clean mating surfaces in preparation for bonding. 5. Bond and rivet doubler in position. Install rivets while adhesive is wet. 6. Apply primer to bare metal surfaces. Allow to dry. 7. Seal edges of repair.
3-74
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
FLANGE DEPTH (SHOWN UP)
1/3 FLANGE DEPTH MAX
0.10 RADIUS MIN 0.10 INCH MIN 1/2 FLANGE DEPTH MAX
Figure 3-28 3-6-5. EDGE TEARS AND CRACKS IN BULKHEAD LIGHTENING HOLES. APPLICATION A.
For repair of torn or cracked lightening hole flange in bulkheads.
RESTRICTION A. Damage not to exceed 1/3 of flange depth. Length of damage not to exceed 1/2 of flange depth after cleanup. For damage exceeding 1/2 of flange depth after cleanup, refer to Application B. REQUIRED A. 1. Primer (Item S204). PROCEDURE A. 1. Dye penetrant inspect to determine extent of crack. 2. Remove crack by blending. Blend radius to be 0.10 inch minimum. 3. Apply primer to bare metal surfaces. Allow to dry.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-75
BHT-206-SRM-1
FAA APPROVED
FLANGE WIDTH (SHOWN UP)
0.9 TO 1.3 1.12 DESIRED 0.38TYP
STOP DRILL WITH NO. 30 DRILL
CRACK 10.50 MIN
CONTOUR TO MATCH FLANGE
Figure 3-29. 3-6-5. TEARS AND CRACKS IN BULKHEAD LIGHTENING HOLES. APPLICATION B. A.
For repair of torn or cracked bulkhead in lightening hole flange exceeding Application
RESTRICTIONS B. 1. Crack length not to exceed flange width. 2. Maximum of two cracks per lightening hole. 3. Not applicable to material greater than 0.040 in. thick. REQUIRED B. 1. Doubler: from composite bond material of same material and one gauge thicker than cracked web. Length and width to allow two rows of rivets plus 0.38 inch edge distance. 2. Rivets of same type as installed in the immediate area; 0.125 inch minimum diameter. 3. Adhesive (item S317). 4. Primer (Item S204). 5. Sealant (Item S392). 6. Process sheets.
3-76
Cleaning (refer to Para 3-2-5) Bonding (refer to Para. 3-2-7) USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
PROCEDURE B. 1. Dye penetrant inspect to determine extent of crack. 2. Stop drill crack with a No. 30 drill. Carry out dye penetrant inspection as in step 1 to ensure crack does not extend past stop drill. 3. Locate doubler on either surface of web. Drill and deburr rivet holes. 4. Prepare doubler and faying surface of structure for bonding. 5. Bond and rivet doubler in position. Install rivets while adhesive is wet. 5. Apply primer to bare metal. Allow to dry. 6. Seal edges of repair.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-77
BHT-206-SRM-1
REVISION 1
FAA APPROVED
3-7 TITANIUM STRUCTURE REPAIRS The physical characteristics of titanium and its primary use in the water and fume tight areas of the engine and fuel systems make titanium structures a special case of sheet metal structures. Typical rivet repairs apply to titanium structures (monel rivets are required) except that formed sections, webs and skins require special considerations. Refer to repairs 3-7-1 through 3-7-2. NEGLIGIBLE DAMAGE.
a.
1. Surface scratches less than 10% of skin thickness and small dents on skin surfaces that do not penetrate the skin or interfere with mounting surfaces. 2. Surface scratches less than 10% of skin thickness and smooth contoured dents less than 0.025 inch deep on angles. Dents must clear spot welds and rivets and be free of cracks or sharp gouges. b.
REPAIRABLE DAMAGE.
1. Dents that show evidence of cracks. 2. Cracks, holes, and tears. CAUTION ANY DAMAGE THAT PENETRATES THE FIREWALL, SKINS OR ANY OPENINGS CREATED BY REPAIRS OR BROKEN SPOT WELDS SHALL BE REPAIRED AND SEALED TO ENSURE LIQUIDS AND FUME TIGHTNESS. c.
MATERIAL SUBSTITUTION. It is acceptable to substitute corrosion resistant steel per MIL-S-5059 1/2 hard, one gauge thicker, for repair of titanium in the following repairs.
3-78
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED MAX DEPTH OF DAMAGE 0.20 INCH
MAX WIDTH OF DAMAGE 0.10 INCH WIDTH OF BLENDED CUTOUT TO BE 0.30 MIN, 0.60 MAX
0.38 TYPICAL MAX DEPTH OF CUTOUT 0.25 INCH
DAMAGEDANGLE TYP
REPAIR ANGLE
PICK UP EXISTING RIVET PATTERN IF POSSIBLE. IF SPOT WELDS WERE USED IN LIEU OF RIVETS, ATTACHMENT RIVETS SHOULD BE SPACED EVERY 0.75 INCH.
Figure 3-30. 3-7-1. FLANGE DAMAGE TO TITANIUM ANGLES. For damages in titanium angle not exceeding 0.20 inch depth and 0.10 width. Refer APPLICATION A. to Application B for damage exceeding this application. RESTRICTIONS A. 1. Vertical flange must be wide enough to provide 1.5D edge distance for added rivets. 2. Not applicable to angle thicker than 0.032 inch. 3. Maximum of two repairs separated by 2 inches are permitted; repairs may not overlap or interfere with later installations. 4. Cutout to be within 0.30 to 0.60 inch in length and maximum of 0.25 inch in depth. 5. Repair angle to be located inside existing angle except when part interference requires external installation. 6. Maximum width of damage 0.10 inch before cleanup. REQUIRED A. 1. Doubler: Angle of the same material and one gauge thicker than original angle. Length sufficient to pick up a minimum of two rivets plus 0.38 inch edge distance per flange on each side of damaged area and one rivet opposite damage.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-79
BHT-206-SRM-1 2. Sealant
FAA APPROVED General purpose (Item S308). Firewall (item S353)
3. MS20615-(MP() or M7885/4 rivets of same diameter as previously installed or MS20615-4MP() or M7885/4-4 rivets if previously spot welded, grip length to suit. 4. Solvent (Item S309) 5. Abrasive cloth (Item S423). 6. Process sheet. Cleaning (refer to Para. 3-2-5) PROCEDURE A. 1. Blend out damaged area to limits shown, including any cracks resulting from damage. 2. Position repair angle. Drill and deburr rivet holes, picking up existing rivet pattern if rivets were used originally to attach angle. Use 0.75 inch rivet spacing if parts were originally spot welded. 3. Clean paint and dirt from repair area. Apply general purpose sealant to faying surfaces in repair area. 4. Install rivets of same size as originally used or if spot welding was used originally, use MS20615-4MP or M7885/4-4 rivets. 5. Seal surface of repair using firewall sealant.
3-80
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
DAMAGE 0.50 MIN 0.75 MAX 0.75 MAX 0.50 MIN CUTOUT PICK UP EXISTING
FILLER
RIVET PATTERN IF RIVETS WERE USED ORIGINALLY. IF SPOT WELDS WERE USED IN LIEU OF RIVETS. MS20615-4MP RIVETS SHOULD BE USED WITH A SPACING OF 0.75 MIN. NOTE: MIN OF 4 RIVETS IN EACH FLANGE. TYPICAL EACH END. 0.38 TYP DOUBLER ANGLE
Figure 3-31. 3-7-1.
FLANGE DAMAGE TO TITANIUM ANGLES.
APPLICATION B.
Damage greater than 0.20 inch depth and 0.10 inch length.
RESTRICTIONS B. 1. Vertical flange must be wide enough to provide 1.5D minimum inch edge distance for added rivets. 2. Not applicable to any angle thickness greater than 0.032 inch. 3. Maximum of two repairs separated by 8 inches permitted; repairs may not overlap or interfere with later installations. 4. Cutout to be within 0.50 to 0.75 inch wide. 5. Repair angle to be nested inside existing angle except when part interference requires external installation. REQUIRED B. 1. Filler: angle of same material and thickness as original angle. 2. Doubler: angle of same material and one gauge thicker than original angle. Length sufficient to pick up a minimum of four (4) rivets plus 0.38 inch in each flange on each side of filler. 3. MS20615-(MP( or M7885/4 rivets of same diameter as previously installed or MS20615-4MP( M7885/4-4 rivets if previously spot welded, grip length to suit.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
or
3-81
BHT-206-SRM-1 4. Sealant:
FAA APPROVED General purpose (Item S308) Firewall (item S353).
5. Solvent (Item S309). 6. Abrasive cloth (Item S423). 7. Process sheet.
Cleaning (refer to Para. 3-2-5).
PROCEDURE B. 1. Cut out damaged area (including any cracks resulting from damage) to limits shown. 2. Position filler and doubler on damage area. Drill and deburr rivet holes, picking up existing rivet locations and at least one rivet in each flange of filler. Use 0.75 inch rivet spacing if parts were originally spot welded. 3. Clean paint and dirt from repair area. Apply general purpose sealant to faying surfaces in repair area. 4. Install rivets of same size as originally used or if spot welding was used originally, use M7885/4-4 rivets or MS20615-4MP rivets, grip length as required. 5. Seal edges of repair using firewall sealant.
3-82
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 0.9 TO 1.3 1.12 DESIRED DOUBLER
(OPTIONAL) MS20615-4MP ITYP)
0.50 INCH
'
ROWS
BETWEEN STAGGERED (TYP)
RADIUS
/
CORNERS
DAMAGE 5.0 INCHES MAX
0.38 MIN 0.38 (TYP) RADIUS CORNERS 0.5 MIN
DAMAGE CUTOUT
5.0 INCHES MAX
WEB
Figure 3-32. 3-7-2. DOUBLER REPAIR TO TITANIUM SKIN OR WEB. APPLICATION. Cracks, dents, holes, and tears that can be completely removed with a cutout of not more than 5.00 inches square. Maximum 25 square inches. RESTRICTIONS. 1. Edge of damage a minimum of 2.00 inches from adjacent structure, after cleanup. 2. Maximum length of cleanup 5.00 inches on long side. 3. Maximum of two repairs per skin, separated by five inches, after cleanup. REQUIRED. 1. Doubler: of same material and one gauge thicker than web or skin, large enough to accommodate two rows of rivets plus 0.38 inch edge distance. NOTE Filler is required if doing internal doubler repair and optional with external doublers. Use a minimum of two rivets to attach plugs one inch or less in diameter; three rivets for plugs greater than one inch but less than 1.5 inches in diameter; and four rivets for plugs 1.5 inches to 2.0 inches in diameter. 2. Filler: of same gauge and material as web, size to match damage cutout (optional).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-83
FAA APPROVED
BHT-206-SRM-1 3. Sealant: General purpose (Item S308). Firewall (item S353)
4. Rivets, MS20615-(MP() of same diameter as previously installed. As an alternate, M7885/4-4 rivets may be used if spot welds were used or area is not accessible from the back, grip length to suit. 5. Solvent (Item S309). 6. Abrasive cloth (Item S423). 7. Sealant (Item S353). 8. Process sheet. Cleaning (refer to Para. 3-2-5) PROCEDURE. 1. Rout out damaged area, removing minimum material, to provide a smooth rectangular, oval or circular cleanup area. Minimum cleanup corner radius is 0.5 inch. 2. Deburr and refinish damaged area. 3. Position doubler over repair area. Drill and deburr rivet holes. 4. Clean mating surfaces in preparation for assembly. 5. Apply general purpose sealant to faying surfaces in repair area and rivet doubler in position. Install rivets while sealant is wet. 6. Seal surface of repair with firewall sealant.
3-84
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED 3-8.
BHT-206-SRM-1
CORROSION REPAIRS.
GENERAL. Corrosion products and damaged metal surfaces must be removed by the mildest method available to prevent additional damage to the part being repaired. Evaluate corroded parts before and after rework to determine the depth of the damage, size, location and the number of affected areas. This will determine the type of repair to be accomplished. WARNING CORROSION DAMAGE ON 7000 SERIES ALUMINUM ALLOY PARTS THAT HAVE HAD SPECIAL TREATMENTS TO PRECLUDE STRESS AND FATIGUE CORROSION, SUCH AS SHOTPEENING OF SURFACES AND STRESS RELIEVING, CANNOT BE REPAIRED EXCEPT AS SHOWN IN THE APPLICABLE MAINTENANCE/OVERHAUL MANUAL.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-85
BHT-206-SRM-1
REVISION 1
FAA APPROVED
3-8-1. SURFACE CORROSION APPLICATION A.
Removal of surface corrosion from sheetmetal parts and extrusions.
RESTRICTIONS A. 1. Not applicable to horizontal stabilizer (206L), stabilizer spar (206B) and vertical fin unless directed to do so by a specific repair in this manual. 2. Maximum affected area 20% of skin surface area. 3. Maximum depth of repair 10% of skin thickness after cleanup. 4. No pitting and no penetration allowed. REQUIRED A. 1. 2. 3. 4.
Abrasive cloth or paper (item S423), 320 grit. Scotchbrite (item S407), grade A, very fine. Alcoholic Phosphoric Acid (item S344). Mix one part of APA to three parts de-mineralized water. Primer (item S204).
PROCEDURE A. 1. Remove corroded aluminum by hand sanding with abrasive cloth. Remove sanding residue with a solvent or detergent. 2. Clean area by scrubbing using Scotchbrite and acid. Rinse surface of skin with water within 30 minutes of acid solution application and dry using clean dry compressed air. CAUTION AIR PRESSURE NOT TO EXCEED 30 PSI. 3. Inspect depth of reworked surface to ensure that it does not exceed the restrictions stated above. Damage exceeding limitations may be repairable (refer to Procedure B). 4. Apply primer to bare metal. Allow to dry. 5. Refinish as required.
3-86
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
3-8-1. SURFACE CORROSION. APPLICATION B. Removal of corrosion in excess of limits stated in Application A. NOTE Due to the limited material thickness available in sheetmetal details used on the Model 206 series helicopters, corrosion damage exceeding 10% of material thickness after cleanup will be treated as mechanical damage. Remove the entire corroded section and repair using the applicable sections of this manual or replace affected part(s). RESTRICTIONS B. As directed by the applicable Procedure in this manual. REQUIRED B. As directed by the applicable Procedure in this manual. PROCEDURE B. As directed by the applicable Procedure in this manual.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-87
BHT-206-SRM-1 3-9.
FAA APPROVED
HONEYCOMB PANEL FACE REPAIRS.
Refer to repairs 3-9-1 through 3-9-3. Honeycomb panels are constructed of aluminum or fiberglass face sheets bonded to aluminum or composite honeycomb core. Doublers and thicker core sections are used in areas of load concentration. All honeycomb repairs must be adequately sealed (item S308) and primed (Item S204). CAUTION THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS ON SECTIONS 4 AND 5 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE HORIZONTAL STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTION 4 a.
NEGLIGIBLE DAMAGE. 1. Small dents/voids
b.
REPAIRABLE DAMAGE. 1. Large dents 2. Small punctures
3-88
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
3-9-1. PANEL SURFACE- SMOOTH DENTS. APPLICATION A.
Small smooth dents and/or small voids in skin.
RESTRICTIONS A. 1. No more than three dents allowed in panel. considered as one dent.
Dents closer than one inch (edge to edge) are to be
2. Dents may not exceed 1/2 inch in diameter and may not exceed 10 percent of panel thickness or a maximum of 0.050 inch in depth. 3. Edge of any dent a minimum of one inch from any attachment point or insert attaching a structural member, fitting, control support, cutout or the panel edge bevel. 4.
Voids (bond failures) may not exceed 0.25 square inch, (i.e. 1/2 inch X 1/2 inch).
5. No more than two skin voids allowed within a four inch diameter circle. Voids closer than 1.0 inch (edge to edge) are to be considered one void. 6. The edge of any void must be at least one inch away from any attachment point or insert attaching a structural member, cutout or panel edge bevel. 7.
No edge separation (delamination) allowed.
REQUIRED A. Not applicable. PROCEDURE A. 1. Voids or bonding separations can be detected by tapping. A dead or flat sound will be produced if a void exists. Outline affected area with a grease pencil. 2.
No repair required. Application B may be used if desired.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-89
BHT-206-SRM-1
FAA APPROVED
1.75 REMOVE FINISH 1.75 INCHES BEYOND DAMAGE AREA. SAND SURFACE WITH 360 GRIT PAPER.
REMOVE FINISH A
400MAX
10 T OR 0.05 MAX 1.75
SECTION A-A
APPLICATION B:
Figure 3-33. 3-9-1. PANEL SURFACE- SMOOTH DENTS. APPLICATION B.
Large, smooth dents.
RESTRICTIONS B. 1. Smooth dents in one skin only, a maximum depth of 10 percent of panel thickness (not to exceed .050 deep) and a maximum of 4.00 inch diameter. 2. Edge of dent at least one inch from any attachment point or insert attaching a structural member, fitting, control support, a cutout or the panel edge bevel. 3.
No other damage within four inches of edge of dent.
4.
No other damage permitted to dented portion of skin.
REQUIRED B.
3-90
1.
Solvent (item S309).
2.
Adhesives: Filler (Item S317), wet lay-up (Item S363).
3.
Fiberglass cloth (Item S404) patch, 0.125 inch smaller than area of removed finish.
4.
Primer (Item S204).
5.
Abrasive cloth (Item S423).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED 6.
Process sheets.
BHT-206-SRM-1 Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Wet lay-up (refer to Para. 3-2-8)
PROCEDURE B. 1.
Clean and remove finish from damaged area, 1.75 inches beyond edge of dent.
2.
Fill dent above contour with filler adhesive and allow to cure. CAUTION DO NOT SAND INTO FIBERGLASS. REMOVED WITH SOLVENT.
DO NOT CLEAN AREA FROM WHICH FINISH WAS
4.
Sand cured adhesive to contour and prepare area for bonding.
5.
Prepare fiberglass doubler.
6.
Bond doubler in position.
7.
Prime repair area and refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-91
BHT-206-SRM-1
FAA APPROVED DAMAGE
Figure 3-34 3-9-2. PANEL SURFACE - SHARP DENTS. APPLICATION. Smooth dents more than 10 percent of panel thickness in depth, affecting a single skin and core; sharp dents, creases, or small punctures affecting a single skin and core. RESTRICTIONS. 1.
Diameter of hole after cleanup does not exceed 0.50 inch diameter.
2.
No more than two damaged areas can be contained within a four inch diameter circle.
3. Damage areas closer than one inch (edge to edge) are to be considered one damage. Refer to repair 3-9-3. 4. Edge of the cleanup at least three inches from any attachment point or insert attaching a structural member, fitting, control support, a cutout, or the panel edge bevel. REQUIRED.
3-92
1.
Solvent (item S309).
2.
Adhesive: Filler (Item S317).
3.
Primer (Item S204).
4.
Abrasive cloth (Item S423).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
5.
Process sheets.
BHT-206-SRM-1
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7)
PROCEDURE. 1. Cut 0.50 inch maximum diameter hole through skin (centered on damaged area), one-half of core thickness deep. 2. Undercut core from skin 0.25 inch. A high speed burr or cutting tool made from an alien wrench may be used. 3.
Deburr hole, remove chips and loose material.
4.
Fill hole above contour with adhesive and allow to cure. CAUTION DO NOT SAND INTO FIBERGLASS. DO NOT CLEAN AREA FROM WHICH FINISH WAS REMOVED WITH SOLVENT.
5.
Sand cured adhesive to contour.
6.
Prime repair and refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-93
FAA APPROVED
BHT-206-SRM-1
HOLE THROUGH PANEL
HOLE THROUGH SKIN
HOLE TRIMMED AND CORE REMOVED
HOLE TRIMMED AND CORE REMOVED c
1.0 .MAX MAX
UNDERCUT 0.25 TYPICAL
UNDERCUT
0.25 TYPICAL
DAMAGE CLEAN-UP
FIBERGLASS CLOTH (2 PLIES MIN) ALUMINUM DOUBLER
FIBERGLASS CLOTH (2 PLIES MIN) DOUBLER ALUMINUM OR SEAL EDGES OF
HOLE FILLED WITH ADHESIVE
DOUBLER
FLUSH WITH OUTER SKINS
SEAL EDGES OF DOUBLER
OVERLAP
HOLE FILLEDWITH ADHESIVE FLUSH WITH OUTER SKIN
HOLE THROUGH ONE SURFACE
MAX
OVERLAP OVERLAP
HOLE THROUGH BOTH SURFACES
Figure 3-35. 3-94
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
3-9-3. PANEL SURFACE PUNCTURES, FIBERGLASS OR METAL FACED. APPLICATION A. cleanup.
Hole through one or both surfaces of a panel, a maximum of 1.00 inch in diameter after
RESTRICTIONS A. 1. Maximum of two repairs per panel. The edges of each repair must be separated by a minimum of five inches. 2. Edge of cleanup a minimum of three inches from any attachment point or insert attaching a structural member, fitting, control support, a cutout or the panel edge bevel. 3.
Skin repair cutout shall not exceed 1.00 inch diameter.
4. When damage affects a metal skin, the panel shall be a minimum of 0.38 in thick to allow installation of blind fasteners. REQUIRED A. 1.
Adhesives, general purpose (Item S317), wet lay-up (Item S363).
2. Doublers: from fiberglass cloth (Item S404) consisting of one ply more than existing skin if panel is faced of fiberglass or composite bond material of same material and one gauge thicker than panel skin(s) when repairing a metal faced panel. NOTE Use doubler of 0.020 in min. for panel skins less than 0.020 in thick. 3.
Blind fasteners M7885/6-4-(), (for fastening metal doubler only).
4.
Sealant (Item S392).
5.
Abrasive paper (Item S423).
6.
Primer (item S204)
7.
Process sheets.
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Wet lay-up (refer to Para. 3-2-8)
PROCEDURE A. 1. Cut out damaged area. When all cracks, tears and rough edges are removed, repair area must not exceed damage limits listed above. CAUTION DO NOT SAND INTO FIBERGLASS. 2. Undercut core approximately 0.25 Inch. A high speed burr or cutting tool made from an allen wrench may be used. 3.
Deburr skins. Remove all chips and loose material.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-95
BHT-206-SRM-1 4.
FAA APPROVED
Remove finish from damaged area and 1.5 inch beyond edge of cutout (using abrasive paper).
5. Fill prepared area with filler adhesive. Allow to cure. Sand smooth to contour. Do not sand into fiberglass. 6.
(a) For repair of fiberglass skin, prepare doubler consisting of one ply more than existing skin. Prepare panel surface for bonding and bond doubler. (b) For repair of metal skin, prepare doubler and position on repair. Drill doubler and panel for eight equally spaced fasteners. Provide 0.38 inch minimum edge distance between fastener and edge of doubler or cutout. Remove doubler and deburr all holes. Prepare surfaces of doubler and panel for bonding. Bond and rivet doubler in position on panel. Install rivets while adhesive is wet.
3-96
7.
Seal edges of repair.
8.
Prime surface of repair. Allow to dry.
9.
Refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1 2.00 TYP
BHT-206-SRM-1
0.9TO 1.3 1.12 DESIRED 0.38 RAD TYP
+
+
+
+
+
0.38 TYP MAX AREA 12.0 SQ. IN.
+
0.50RADMINTYP +
+
1.5 MIN. OVERLAP--
+
+ COMPOSITE BOND DOUBLER (ORIGINAL SKIN THICKNESS PLUS ONE GAUGE)
EXPOSED CORE 0.250TYP
OVERLAP
METAL FILLER
SEAL EDGES OF REPAIR
TO 0.20 INCH GAP BETWEEN CORE -0.10 PLUG AND EXISTING CORE TYPICAL
HONEYCOMB CORE PLUG
Figure 3-36. 3-9-3.
PANEL SURFACE PUNCTURE, DELAMINATION OR CORROSION - METAL FACED PANEL.
APPLICATION B. For damage to metal faced panels, greater than one inch in diameter. damage to fiberglass faced panels, refer to Application C.
For repair of
RESTRICTIONS B. 1.
Maximum length of cleanup of five inches in any direction.
2. Maximum of two repairs per panel which must be separated by a minimum of five inches between edges of cleanup. 3. Total damage a maximum of 12.0 square inches or 20% of total panel area, whichever is smaller, when only one skin and core is affected, and 10.0 square inches or 15% of total panel area, whichever is smaller, when both skins are affected. 4. Edge of cleanup must be a minimum of three inches from any attachment point or insert attaching a structural member, fitting, control support, a cutout or a panel bevel edge. 5.
The panel shall be a minimum of 0.38 in thick to allow installation of blind fasteners.
REQUIRED B. 1.
Core plug same material and thickness as core used in panel.
2. Filler(s): from composite bond material of same material and gauge as skin(s). Dimensions to fit in skin cutout(s).
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-97
FAA APPROVED
BHT-206-SRM-1
3. Doubler: from composite bond material of same material and one gauge thicker as existing skin(s). Doubler(s) to extend 2.00 inches beyond repair cutout on all sides. NOTE Use doubler of 0.020 in min. for panel skins less than 0.020 in thick. 4.
Blind fasteners, M7885/2-4-0, grip length to suit.
5.
Adhesive (Item S317).
6.
Solvent (Item S309).
7.
Primer (Item S204).
8.
Sealant (Item S392).
9.
Process sheets.
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Core splice (refer to Para. 3-2-10)
PROCEDURE B. 1. Cut out skin(s) in damaged area allowing for a minimum of 0.25 inch undamaged core to protrude from skin cutout all around. Ensure core is not damaged while cutting out skin(s). 2. Cut out damaged core leaving a minimum of 0.25 inch of core protruding from the skin cutout all around. 3.
Deburr skin(s) and remove loose material.
4.
Prepare core plug. Align the plug ribbon direction with existing core in panel.
5. Position doubler(s) on repair location and drill holes for blind fasteners. Provide 0.38 edge distance minimum from edge of doubler and edge of repair cutout. Space holes equally at 0.9 to 1.3 inches, 1.12 desired. 6.
Remove doublers. Deburr all parts, remove chips and loose material.
7.
Remove paint and dirt from repair and clean in preparation for bonding. CAUTION DO NOT ALLOW ANY SOLVENT TO CONTAMINATE CORE PORTION OR EDGE OF HONEYCOMB PANEL.
8.
Bond core plug in position. Ensure core ribbon direction matches existing core.
9.
Bond filler(s) in position.
10. Bond and rivet doubler(s) in position. Install rivets while adhesive is wet. 11.
Apply primer to bare metal surfaces. Allow to dry.
12. Seal edges of repair and refinish as required.
3-98
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1
OVERLAP OVERLAP EXPOSED CORE
0.250 TYP
BHT-206-SRM-1
FIBERGLASS CLOTH DOUBLERS (NUMBER EQUAL TO NUMBER OF SKIN PLIES, PLUS ONE)
FILLER (OPTIONAL)
CORE SPLICE ADHESIVE.20 HONEYCOMB CORE PLUG
SEAL EDGE OF
DOUBLER
INCH GAP
BETWEEN CORE PLUG AND EXISTING CORE TYPICAL
Figure 3-37. 3-9-3.
PANEL SURFACE PUNCTURE, FIBERGLASS FACED PANEL.
APPLICATION C.
For repairs of delamination, corrosion and punctures in fiberglass faced panel, greater than one inch in diameter.
RESTRICTIONS C. 1.
Maximum length of cleanup of five inches in any direction.
2. Maximum of two repairs per panel which must be separated by a minimum of five inches between edges of cleanup. 3. Total damage a maximum of 12.0 square inches or 20% of total panel area, whichever is smaller, when only one skin and core is affected, and 10.0 square inches or 15% of total panel area, whichever is smaller, when both skins are affected. 4. Edge of cleanup must be a minimum of three inches from any attachment point or insert attaching a structural member, fitting, control support, a cutout or a panel bevel edge. REQUIRED C. 1.
Core plug same material and thickness as core used in panel.
2. Fiberglass cloth doubler(s) (Item S404). Dimensions to suit repair.
Same number of plies as skin plies removed, plus one.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-99
BHT-206-SRM-1
FAA APPROVED
3.
Adhesives:
Core splice (Item S317) Wet lay-up (itemS363).
4.
Primer (Item S204).
5.
Process sheets:
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Wet lay-up (refer to Para. 3-2-8) Core splice (refer to Para. 3-2-10)
PROCEDURE C. 1.
Same as Procedure B, steps (1) thru (4). CAUTION DO NOT SAND INTO FIBERGLASS. DO NOT CLEAN AREA FROM WHICH FINISH WAS REMOVED WITH SOLVENT.
2. Sand surface of panel with 360 or finer grit abrasive paper. Clean surface of panel with dry, clean shop air. Do not use solvent. 3. Cut fiberglass doublers; first doubler overlaps cutout by 1.0 inches and each additional doubler overlaps preceding doubler by 1.0 inch. 4.
Bond doublers in position.
5.
Seal edges and prime repair area. Allow to dry.
6.
Refinish as required.
3-100
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 3-10
HONEYCOMB PANEL EDGE REPAIR.
Most honeycomb panels will taper near the edges to provide a thin flat surface that permits a solid attachment to frames and beams. The tapered edge of core is usually covered with metal or fiberglass cloth to seal it. Refer to repairs 3-9-1 thru 3-9-5. WARNING WEAR BREATHING MASK, FACE SHIELD, AND PERSONAL PROTECTION EQUIPMENT WHEN WORKING ON COMPOSITES. FUMES AND RESIDUE CAN CAUSE IRRITATION TO THE EYES, SKIN, AND LUNGS. CAUTION THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS ON SECTION 4 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTION 4. a.
NEGLIGIBLE DAMAGE. 1. N/A
b.
REPAIRABLE DAMAGE. 1. Core bevel damage, various panel constructions. 2. Damage to skin and edging, various panel constructions.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-101
BHT-206-SRM-1
FAA APPROVED
CORE REMOVED EDGE REPLACEMENT PLIES MIN
DAMAGE
EXISTING EDGING FILLER ADHESIVE
REPAIR
Figure 3-38. 3-10-1.
PANEL EDGE-FIBERGLASS BEVEL DAMAGE.
APPLICATION.
Punctures, dents, or creases in beveled fiberglass edging of panel.
RESTRICTIONS. 1.
Damage restricted to fiberglass edging and core (no damage allowed to metal skins or doubler).
2.
Damage to the core a maximum of 0.50 inch inside the inboard edge of the bevel (top) after cleanup.
3.
Length of damage a maximum of 2.0 inches after cleanup.
4. Maximum of three repairs per panel separated by a minimum of 2.0 inches between edges after cleanup. REQUIRED. 1. Fiberglass edging doubler plies (Item S404). Quantity equal to number of plies removed plus one. Dimensions to suit repair. 2.
Adhesives.
3.
Primer (Item S204).
4.
Abrasive paper (Item S423).
5.
Sealant (Item S392).
3-102
Filler (Item S317) Wet lay-up (Item S363)
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED 5.
Sealant (Item S392).
6.
Process sheets.
Cleaning (refer to Para. 3-2-5) Wet lay-up (refer to Para. 3-2-8)
PROCEDURE. 1.
Remove damaged portion of fiberglass edging and core. Remove debris and loose material.
2.
Fill hole above contour with filler adhesive and allow to cure. CAUTION DO NOT SAND INTO FIBERGLASS. DO NOT CLEAN AREA FROM WHICH FINISH WAS REMOVED WITH SOLVENT.
3.
Sand adhesive to contour. Clean surface of panel with dry, clean shop air. Do not use solvent.
4. Prepare fiberglass doubler plies; first ply overlaps cutout by 1.0 inches and each additional ply overlaps preceding ply by 1.0 inch. 5.
Bond doubler plies in position.
6.
Seal edges and prime repair area. Allow to dry.
7.
Refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-103
BHT-206-SRM-1
REVISION 1
FAA APPROVED 0.90 TO 1.3 1.12 DESIRED
0.38R
TYP
+
DAMAGE 0.38 TYP ORIGINAL METAL DOUBLER
1.0
OVERLAP 0.5 MAX.
FIBERGLASS DOUBLER ADHESIVE ORIGINAL EDGING
0.5 MIN RADIUSS 0.38 MIN. EXISTING FASTENERS
0.25 UNDERCUT 4 M788512-4 RIVETS RIVETS METAL SKINS
COMPOSITE BOND DOUBLER
ADDED FASTENERS
EXISTING SKIN EXISTING DOUBLER
Figure 3-39. 3-10-2.
PANEL EDGE- FIBERGLASS BEVEL AND METAL EDGE DOUBLER DAMAGE.
APPLICATION.
Edge damage affecting metal skin, doubler, and fiberglass edging.
RESTRICTIONS. 1. Damage to the skins, core, and doubler does not extend more than 0.5 inch inside the inboard edge of the bevel (top) after clean-up. 2.
Damage length a maximum of two fasteners plus edge distance (1.25 inch maximum) after clean up.
3. Maximum of two repairs per panel, separated by a minimum of 5.00 inches between edges after clean-up. 4.
Repair cannot interfere with later installations.
5.
Panel minimum core thickness of 0.38 inch to allow use of blind fasteners.
REQUIRED. 1. Doubler: from composite bond material of same material and one gauge thicker than combined thickness of existing skin plus doubler. Doubler to extend from edge of panel to two inches beyond repair cutout on the three remaining sides. 2. Fiberglass doubler plies (Item S404). Quantity of plies equal to number of plies in damaged edging plus one. Dimensions to suit repair. 3.
3-104
Blind fasteners, M7885/2-4-(). USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
4.
Adhesives.
5.
Solvent (Item S309).
6
Primer (Item S204).
7.
Abrasive paper (Item S423).
8.
Sealant (Item S392).
9.
Process sheets.
Filler and metal doubler bond (Item S317) Wet lay-up (Item S363)
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Wet lay-up (refer to Para. 3-2-8) Edging (refer to Para. 3-2-9)
PROCEDURE. 1.
Provide unobstructed access to damaged edge.
2. Cutout damaged area. When all cracks, tears, and rough edges are removed, repair area shall not exceed limits listed above. NOTE Do not cut undamaged skin when only one side of panel is affected. 3.
Undercut core a minimum of 0.25 inch.
4. Center composite bond doubler on repair cutout. Drill for fasteners. Provide 0.38 inch minimum edge distance from edge of doubler to edge of repair cutout. Space rivets equally 0.9 to 1.3 inches apart, 1.12 desired. 5.
Deburr all parts and remove chips and loose material.
6. Cut fiberglass doubler plies as follows; first ply overlaps end of original edging by 1.0 inch and additional plie(s) overlaps previous ply by 1.0 inch. 7.
Fill hole above edging cutout with filler adhesive and allow to cure. CAUTION DO NOT SAND INTO FIBERGLASS.
DO NOT CLEAN WITH SOLVENT.
8. Sand to contour with abrasive paper. Clean panel surface with dry, clean shop air. Do not use solvent. CAUTION DO NOT USE SOLVENT ON FIBERGLASS. REMOVE PAINT BY SANDING. OR FINER GRIT PAPER AND CLEAN WITH DRY, CLEAN SHOP AIR. 9.
USE 360
Clean panel in preparation for bonding.
10. Bond and rivet metal doubler in position. Install rivets while adhesive is wet. USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-105
BHT-206-SRM-1 11.
FAA APPROVED
Bond fiberglass doublers in position.
12. Apply primer to bare metal surfaces. Allow to dry 13. Seal edges of repair area. 14. Refinish as required.
3-106
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1
BHT-206-SRM-1 0.9 TO 1.3 1.12 DESIRED
0.38 R TYP
DAMAGE 0.38 TYP INNER METAL SKIN ORIGINAL METAL DOUBLER COMPOSITE BOND EDGING
0.5 MAX 0.50 RADIUS TYP
--
0.38
TYP
0.25 CORE UNDERCUT M7885/2-4 RIVET TYPICAL
0.38MIN.-
COMPOSITE BOND DOUBLER
+
OUTER METAL SKIN
EXISTING FASTENERS ADDED FASTENERS
Figure 3-40. 3-10-3.
PANEL EDGE- ALL METAL BEVEL AND EDGE DOUBLER DAMAGE.
APPLICATION.
Edge damage to panels with all-metal construction.
RESTRICTIONS. 1. Damage to the skins, core, and doubler does not extend more than 0.50 inch inside the inboard edge of the bevel (top) after cleanup. 2.
Damage length a maximum of two fasteners plus edge distance (1.25 inch maximum) after clean up.
3.
Maximum of two repairs separated by a minimum of 5.0 inches between edges after cleanup.
4.
Repair not to interfere with later installations.
5.
Panel minimum core thickness of 0.38 inch to allow use of blind fasteners.
REQUIRED. 1. Composite bond doubler of same material and one gauge thicker than combined thickness of basic outer skin plus existing metal doubler. Doubler to extend from edge of panel to 2.0 inches beyond repair cutout on the three remaining sides. NOTE Minimum doubler/edging thickness shall be 0.020 inch for original skin/edging thickness of less than 0.020 inch.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-107
FAA APPROVED
BHT-206-SRM-1
2. Composite bond edging of same material and one gauge thicker than inner skin. Edging to extend from edge of panel to 2.0 inches beyond repair cutout on the three remaining sides. 3.
Blind fasteners, M7885/2-4-(,
4.
Adhesive (Item S317).
5.
Abrasive paper (Item S423).
6.
Primer (Item S204).
7.
Sealant (Item S392).
8.
Process sheets.
grip length to suit.
Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7)
PROCEDURE. 1.
Provide unobstructed access to damaged edge.
2. Cutout damaged area. When all cracks, tears, and rough edges are removed, repair area shall not exceed limits listed above. NOTE Do not cut out undamaged skin when only one side of panel is affected. 3.
Undercut core a minimum of 0.25 inch beyond skin cutout.
4.
Fabricate doubler(s) dimensioned as required, position on repair.
5. Drill holes for rivets providing a minimum of 0.38 inch edge distance from edge of doubler and edge of repair cutout. Space rivets equally 0.9 to 1.3 inch apart, 1.12 in. desired. Stagger rivet rows on opposite skin surfaces to prevent interference. 6.
Remove doublers, deburr all parts, remove chips and loose materials.
7.
Prepare panel and doubler surfaces in preparation for bonding.
8.
Bond and rivet one of the doublers in position. Install rivets while adhesive is wet.
9. Fill hole above edging with adhesive. contour with abrasive paper.
Allow to cure 24 hours at room temperature then sand to
10. Prepare panel and doubler surfaces in preparation for bonding. 11. Bond and rivet remaining doubler in position. Install rivets while adhesive is wet. Apply 0.5 to 1.0 psi pressure to panel while adhesive cures. Clean excess adhesive squeeze-out. 12. Seal edges of repair. 13. Apply primer to bare metal surfaces. Allow to dry. 14. Refinish as required.
3-108
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1
BHT-206-SRM-1 EDGE OF FIBERGLASS DOUBLERS
EDGE OF METAL 0.38 RA
TYP
DOUBLER IF0.9 APPLICABLE TO 13
DAMAGE EXISTING INNER DOUBLER 1.0 OVERLAP
12 DESIRED +
UNDERCUT CORE 0.25 MINIMUM
0.38 TYP
0 38 TYP
0.50 MAX CORE DAMAGE CLEAN UP
REPLACEMENT EDGING FILLER ADHESIVE DAMAGED EDGING
0.50 RADIUS TYP
I
EDGEOF CUTOUT + (IF APPLICABLE) FIBERGLASS
+
EXISTING FASTENERS 0.38 MIN. ADDED FASTENERS
Figure 3-41. 3-10-4.
PANEL EDGE- ALL FIBERGLASS PANEL CONSTRUCTION
(WITH OR WITHOUT METAL EDGE
DOUBLER). APPLICATION.
Edge damage to fiberglass skin, metal or fiberglass doubler and fiberglass edging.
RESTRICTIONS. 1. Damage to the skins, core, and doubler does not extend further than 0.500 inch inside the inboard edge of the bevel (top) after cleanup. 2.
Damage length a maximum of two fasteners plus edge distance AFTER CLEAN UP.
3.
Maximum of two repairs separated by a minimum of 5.00 inches between edges after cleanup.
4.
Repair not to interfere with later installations.
5.
Panel minimum core thickness of 0.50 inch to allow use of blind fasteners.
REQUIRED. 1. Doubler: from composite bond material of same material and one gauge thicker than existing inner doubler (if applicable). Doubler to extend from edge of panel to 2.0 inches beyond repair cutout on the three remaining sides. NOTE Use doubler of 0.020 in min. for panel skins less than 0.020 in thick.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-109
BHT-206-SRM-1
FAA APPROVED
2. Doubler: from fiberglass cloth (Item S404) doubler plies equal to the number of plies of the damaged panel skin(s) and fiberglass doubler combination, plus one. Dimensions to suit. 3.
Blind fastener M7885/2-4-(, grip length to suit (if metal doubler is used).
4.
Adhesives.
5.
Primer (Item S204).
6.
Abrasive paper (Item S423).
7.
Sealant (Item S392).
8.
Process sheets.
General purpose (Item S317) Wet lay-up (Item S363)
Cleaning: (refer to Para. 3-2-5) Bonding: (refer to Para. 3-2-7) Edging: (refer to Para. 3-2-9) Wet lay-up: (refer to Para. 3-2-14)
PROCEDURE. 1.
Provide unobstructed access to damaged edge.
2. Cut out damaged area. When all cracks, tears, and rough edges are removed, repair is not to exceed limits listed above. NOTE Do not cutout undamaged skin when only one side of panel is affected. 3.
Undercut core approximately 0.25. inch beyond skin and doubler cutout.
4. Prepare composite bond doubler as shown and position on repair. Drill for fasteners. Provide a minimum of 0.38 edge distance from edge of doubler and edge of repair cutout. Space rivets equally 0.9 to 1.3 inches, 1.12 desired. 5. 6.
Deburr all parts. Remove chips and loose material. Prepare fiberglass doubler plies: first ply overlaps end of original edging by 1.0 inch and additional plies overlap previous ply by 1.0 inch. CAUTION DO NOT USE SOLVENT ON FIBERGLASS. CLEAN WITH DRY, CLEAN SHOP AIR.
7.
Prepare surfaces of doubler and panel in preparation for bonding.
8.
Bond and rivet composite bond doubler in position. Install rivets while adhesive is wet.
9.
Fill hole above edging cutout with filler adhesive. Allow to cure.
10. Bond fiberglass doublers in position. 11.
Seal edges and prime repair area. Allow to dry.
12. Refinish as required. 3-110
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
USE OR DISCLOSURE
rr
REVISION 1
OF DATA CONTAINED ON THIS PAGE IS SUBJECT
BHT-206-SRM-1
REVISION 1
BHT-206-SRM-1 3-10-5
FAA APPROVED
PANEL EDGE- OUTER SKIN AND INNER DOUBLER DAMAGE ON METAL-FACED PANEL
APPLICATION.
Edge damage in a metal panel, consisting of skin to doubler separation and/or corrosion.
RESTRICTIONS. 1. Applicable to panel skins 0.025 inch thick or less. 2. Separation shall not extend into panel further than extent of existing doubler. 3. No corrosion permitted in core or between skin or doubler(s) and core and no corrosion on skin with bevel. 4. Doubler to pick up a minimum of 5 existing rivets in pattern. 5. Length of cutout not to exceed 20% of length of side being repaired. 6. Maximum of one repair per side but no more than two repairs per panel. 7. Repairs cannot overlap. REQUIRED: 1. Doubler: from composite bond material of same material and one gauge thicker than combined thickness of existing doubler and skin. Doubler to extend from edge of panel to 2.0 inches beyond cutout on three remaining sides. 2. Skin Filler: from composite bond material of same material and gauge as existing skin. 3. Doubler Filler: from composite bond material of same material and gauge as existing doubler. 4. Adhesive (item S317). 5. Fasteners: M7885/2-4-( ) rivets, grip length to suit. 6. Solvent (item S309). 7. Sealant (item S392). 8. Primer (item S204). 9. Process sheets: Cleaning (refer to Para. 3-2-5) Bonding (refer to Para. 3-2-7) Fiberglass edging replacement (refer to Para. 3-2-9)
3-112
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
PROCEDURE: 1. Cut and remove damaged section of skin along the lines drawn. Inspect internal doubler for corrosion and remove damaged section. Ensure that damage does not exceed restrictions. 2. Adjust skin cutout to suit amount of inner doubler removed. 3. Prepare doubler filler and skin filler of same size as section of inner doubler and skin removed. skin doubler.
Prepare
4. Position fillers and doubler on repair. Drill for fasteners providing spacing and edge distance as shown on figure 3-42. Remove parts, deburr all parts. Remove chips and loose material. 5. Prepare faying surfaces of fillers, doubler and panel for bonding. 6. Bond and rivet fillers and doubler in position. Install fasteners while adhesive is wet. 7. If applicable, repair damaged edging. 8. Apply primer to bare metal surfaces and allow to dry. 9. Seal edges of repair. 10. Refinish as required.
0
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3 113
BHT-206-SRM-1
FAA APPROVED
3-11. UNSUPPORTED COMPOSITE SKIN REPAIRS. Composite skins are comonly used on non-structural components such as doors and fairings. They offer a lighter weight alternative to comparable metal parts. Refer to repairs 3-11-1 and 3-11-2. CAUTION THE TYPICAL REPAIRS IN THIS SECTION ARE SUBJECT TO THE LIMITATIONS OF SPECIFIED REPAIRS IN SECTION 4 OF THIS MANUAL. THESE TYPICAL REPAIRS MAY NOT BE APPLIED TO THE STABILIZER, OR VERTICAL FINS UNLESS SPECIFICALLY DIRECTED IN SECTION 5. a.
NEGLIGIBLE DAMAGE.
1. N/A b.
REPAIRABLE DAMAGE.
1. Delaminations. 2. Punctures
3-114
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
IMPACT DAMAGE
9.0 SQ IN.
PLY FRACTURED/SEPARATED (IF MORE THAN 2 PLIES AFFECTED,
MAX AREA1.00 TYP
REMOVE DAMAGED MATERIAL AND BEVEL EDGES
1.00
TREAT AS PUNCTURE) FILLER PLY
1.00
REPAIR PLIES
Figure 3-43. 3-11-1.
UNSUPPORTED COMPOSITE SKIN- FRACTURED PLIES RESULTING FROM IMPACT DAMAGE.
APPLICATION.
Repair of fractured or separated inner plies in an unsupported composite skin (fiberglass).
RESTRICTIONS. 1. Damage not to exceed 5 inches on long side or 25% of part width after cleanup, whichever is least, for a maximum area of 9.0 sq. inches after cleanup. Not more than two internal plies may be damaged. If the outer surface ply is damaged, repair must 2. be treated as a puncture (see Repair 3-11-2). REQUIRED. 1.
Fiberglass cloth (Item S404).
2.
Solvent (Item S305).
3.
Adhesive (Item S363).
4.
Abrasive paper (Item S423).
5.
Primer (Item S204).
6.
Process sheets.
Cleaning (refer to Para. 3-2-5) Wet layup (refer to Para. 3-2-7)
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-115
BHT-206-SRM-1
FAA APPROVED
PROCEDURE. 1.
Remove damaged material as follows:. (a)
Back up repair area so that it will not flex while removing damaged material. WARNING WEAR BREATHING MASK, FACE SHIELD, AND PERSONAL PROTECTION EQUIPMENT WHEN WORKING ON COMPOSITES. FUMES AND RESIDUE CAN CAUSE IRRITATION TO THE EYES, SKIN, AND LUNGS. CAUTION DO NOT OVERHEAT COMPOSITES DURING MATERIAL REMOVAL.
(b) Cut out damaged material. Use a low speed hole saw (tungsten carbide saw) or a very high speed burr with coolant. (c)
Deburr skin edges, remove debris and loose material.
(d)
Remove abrasion products and dust using solvent.
2. Prepare glass fabric filler and doubler plies as shown on figure 3-42. Number of repair plies to be equal to number of damaged plies plus. 3.
Clean inner surface in preparation for bonding.
4.
Bond skin fillers and doublers in position.
5.
Prime repair. Allow to dry.
6.
Seal edges of repair.
7.
Refinish as required.
3-116
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
BHT-206-SRM-1
FAA APPROVED
CUT OUT DAMAGED MATERIAL, VOIDS AND DELAMINATIONS OUTER SURFACE
LINNER SURFACE
DAMAGE REMOVAL
OUTER SURFACE
ALUMINUM SHEET FOR
2.0 IN MIN
RELEASE FILM
TEMPORARY SUPPORT
INNER SURFACE
ALUMINUM SHEET INSTALLATION
OUTER SURFACE
1.0 MIN YP
DOUBLER PLIES OVERLAP IN ONE INCH INCREMENTS
LINNER SURFACE
INNER SURFACE REPAIR
SURFACE PREPARATION 1.75 MIN
ONE PLY 1.0 IN MIN
OUTER SURFACE
INNER SURFACE
OUTER SURFACE REPAIR
Figure 3-44.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-117
BHT-206-SRM-1
3-11-2.
FAA APPROVED
UNSUPPORTED COMPOSITE SKIN-PUNCTURE DAMAGE.
APPLICATION.
Hole through woven fabric/epoxy skin (refer to figure 3-44).
RESTRICTIONS. 1. Damage not to exceed 5 inches on long side or 25% of part width after cleanup, whichever is least, for a maximum area of 15 sq. inches after cleanup. 2. Minimum edge distance for damage must not be less than 1.5 inches for first ply plus 1.0 inch for each additional ply required. 3.
Maximum of two repairs per skin.
4.
Applicable to fiberglass panels only.
5. Access from both sides of panel is required. access.
It may be necessary to remove the part to provide
REQUIRED. 1.
Fiberglass cloth (Item S404).
2.
Solvent (Item S305).
3.
Adhesive (Item S363).
4.
Abrasive paper (Item S423).
5.
Primer (Item S204).
6.
Process sheets.
Cleaning (refer to Para. 3-2-5) Wet layup (refer to Para. 3-2-7)
PROCEDURE. 1.
Remove damaged material as follows. (a)
Provide support to back of repair area so that it will not flex while removing damaged material. WARNING WEAR BREATHING MASK, FACE SHIELD, AND PERSONAL PROTECTION EQUIPMENT WHEN WORKING ON COMPOSITES. FUMES AND RESIDUE CAN CAUSE IRRITATION TO THE EYES, SKIN, AND LUNGS.
3-118
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
BHT-206-SRM-1
CAUTION DO NOT OVERHEAT COMPOSITES DURING MATERIAL REMOVAL. (b) Cut out damaged material. Use a low speed hole saw (tungsten carbide saw) or a very high speed burr with coolant. (c)
Deburr skin edges, remove debris and loose material.
(d)
Remove abrasion products and dust using solvent.
2. Prepare fiberglass inner and outer surface doubler plies as shown in figure 3-44. Number of inner surface repair plies to be equal to number of plies in panel plus one. 3.
Prepare fiberglass filler. Number of plies to be equal to number of plies in panel.
4.
Clean inner surface in preparation for bonding, approximately 0.75 inches beyon largest doubler ply.
5.
Position a sheet of aluminum, 0.025 inch or thicker over the hole on the outer surface. NOTE Use release film (item S477) between aluminum sheet and panel.
6.
Bond filler plies in repair cutout to provide flush surface for doubler plies.
7.
Bond doubler in position. Allow adhesive to cure.
8. Remove aluminum sheet installed in step 4 and prepare surface for bonding. Apply one ply of glass fabric impregnated with adhesive on outer surface. Allow adhesive to cure. 9.
Prime repair. Allow to dry.
10. Seal edges of repair. 11.
Refinish as required.
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
3-119/3-120
FAA APPROVED
REVISION 1
BHT-206-SRM-1
SECTION 4. REPAIRS TO COMPOSITE STRUCTURE INTRODUCTION. Section 4 presents repair procedures for various bonded panel assemblies. It covers repairs to the most commonly reported damages. The format followed in this section is to group panels according to their function as listed in the subsections below. Additional repairs will be included in their applicable subsection as they become available. Refer to Tables 4-1 thru 4-5 at the beginning of each subsection for additionnal information on each repair. TABLE OF CONTENTS: 1. HOW TO USE THIS SECTION 2. GUIDELINES 3. AREA SPECIFIC REPAIRS 4-1.
MAIN PANELS.
(includes forward lower shell, aft lower shell and roof shell and other structural panels). 4-1-1 ROOF SHELL, 206A/B, L SERIES. Application A ROOF OUTER SECTION, 206 A/B series. Application B ROOF OUTER SECTION, 206L series Application C ROOF CENTER SECTION, 206L series 4-1-2 FORWARD LOWER SHELL, 206A/B SERIES Application A FWD EDGE, Damage to Edging Application B FWD EDGE, Damage to Inner Skin Around Pedals Application C FWD EDGE, Splice Repair of Console-Attaching Floor Angles Application D FWD EDGE, Damage to Corner Outer Skin Application E PANEL OPENINGS, Damage to Outer Skin/Base of Cyclic Support Application F PANEL OPENINGS, Damage to Outer Skin/Antenna Provision Application G PANEL OPENINGS, Damage to Outer Skin/Antenna Provision Application H PANEL OPENINGS, Damage to Outer Skin/Cable Routing Provision Application J PANEL OPENINGS, Damage to Outer Skin/Antenna Provision Application K PANEL OPENINGS, Damage to Outer Skin/Antenna Provision Application L AFT EDGE, Damage to Edging Application M AFT EDGE, Damage to Outer Skin Application N AFT EDGE, Damage to Outer Skin and Inner Doubler Application P AFT EDGE, Damage to Inner Skin 4-1-3 FORWARD LOWER SHELL, 206L SERIES Application A FWD EDGE, Damage to Edging Application B FWD EDGE, Damage to Inner Skin Around Pedals Application C FWD EDGE, Splice Repair of Console-Attaching Floor Angles Application D FWD EDGE, Damage to Corner Outer Skin Application E PANEL OPENING, Damage to Outer Skin/Fuel Drain Openings Application F PANEL OPENING. Damage to Outer Skin/Position Light Openings Application G PANEL OPENING, Damage to Outer Skin/Antenna Provision Application H PANEL OPENING, Damage to Outer Skin/Antenna Provision Application J PANEL OPENING, Damage to Outer Skin/Cargo Hook Provision Application K PANEL OPENING, Damage to Outer Skin/Antenna Provision Application L PANEL OPENING, Damage to Outer Skin/Antenna Provision
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
4-1
BHT-206-SRM-1
REVISION 1
Application M Application N Application P Application R Application S 4-1-4 AFT LOWER Application A Application B Application C Application D Application E Application F Application G Application H Application J Application K 4-1-5 AFT LOWER Application A Application B Application C Application D Application E Application F Application G Application H Application J Application K Application L
FAA APPROVED
PANEL SURFACE, Damage to Inner Skin/Fuel Transfer Lines AFT EDGE, Damage to Edging AFT EDGE, Damage to Outer Skin AFT EDGE, Damage to Outer Skin and Inner Doubler AFT EDGE, Damage to Inner Skin SHELL, 206A/B SERIES FWD EDGE, Damage to Outer Skin FWD EDGE, Damage to Outer Skin and Inner Doubler FWD EDGE, Damage to Extensive Area of Inner Doubler PASSENGER DOOR SILL, Damage Affecting Outer Skin and Core PASSENGER DOOR FRAME, Pulled Inserts/Damage to Outer Skin... PASSENGER DOOR FRAME, Typical Seat Belt Damage... PASSENGER DOOR FRAME, Damage to Outer Skin Below and Aft... FUEL FILLER OPENING, Damage to Outer Skin Below Fuel Cap AFT EDGE, Damage to Outer Skin and Inner Doubler PANEL SURFACE, Damage to Fuel System Positioning Doublers SHELL, 206L SERIES FWD EDGE, Damage to Outer Skin FWD EDGE, Damage to Outer Skin and Inner Doubler FWD EDGE, Damage to Extensive Area of Inner Doubler PASSENGER DOOR FRAME, Pulled Inserts/Damage to Outer Skin... FUEL FILLER OPENING, Damage to Outer Skin Below Fuel Cap PASSENGER DOOR SILL, Pulled Inserts/Damage to Outer Skin... PASSENGER DOOR SILL, Typical Seat Belt Damage... FUEL BOOST PUMP OPENING, Damage to Outer Skin/,nner Doubler FUEL BOOST PUMP OPENING, Damage to Outer Skin/Inner Doubler AFT EDGE, Damage to Outer Skin AFT EDGE, Damage to Outer Skin and Inner Doubler
4-2. SECONDARY PANELS.
(includes nose panels, seat structures, baggage floor, fuselage fairings and oil cooler support). 4-2-1 NOSE SKIN REPAIRS, 206A/B and L SERIES. Application A PANEL SURFACE, Surface damage Application B PANEL SURFACE, Replacement of damaged skin portion...
4-3. FLIGHT SURFACES.
(includes horizontal stabilizer, auxiliary fins and vertical fin).
Not assigned at this time. NOTE TO PREVENT THE POSSIBILITY OF CONFUSION BETWEEN CERTAIN LETTERS AND NUMBERS, THE FOLLOWING LETTERS ARE NOT USED TO DENOTE AN APPLICATION: LETTERS I, Oand Q.
4-2
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
REVISION 1
BHT-206-SRM-1
HOW TO USE THIS SECTION. A figure titled RESTRICTIONS at the beginning of each subsection guides the operator in defining the applicable repair for his situation. This same figure identifies the type of repair that may be applied to areas of the panel as follows: 1. Areas that are double cross-hatched are restricted/critical areas where repair using this manual is not allowed, unless covered as part of a SPECIFIC REPAIR procedure. NOTE A SPECIFIC REPAIR MAY AFFECT PART OF, OR AN ENTIRE RESTRICTED AREA OF A PANEL. BECAUSE OF THE CRITICAL NATURE OF THESE AREAS, IT IS IMPORTANT THAT ALL CONDITIONS OF THE SPECIFIC REPAIR ARE MET BEFORE ATTEMPTING A REPAIR IN THOSE AREAS. IF IN DOUBT AS TO WHETHER YOUR SITUATION MEETS ALL CONDITIONS OF ELIGIBILITY FOR REPAIR, CONTACT PRODUCT SUPPORT ENGINEERING. 2. Areas that are single cross-hatched are reparable using a specific repair covered in this section. The operator then refers to the application number for the repair procedure which is listed beside the area of interest to him. 3. Areas that are free of cross-hatch represent those areas that can be repaired using the applicable repair procedure in section 3. Repair material data for repairs falling within the scope of section 3 is specified in this figure. Another figure entitled INSERT LOCATIONS will provide all information about inserts that may fall within a repair covered in this section. Inserts added as part of a kit or later installation may not be identified. Refer to the Service Instruction covering this kit. GUIDELINES. Unless a specific repair states otherwise, the following notes apply to this section, (refer to FIGURE 4-1, GUIDELINES). NOTE ALL RESTRICTIONS SPECIFIC TO EACH APPLICATION SHALL BE OBSERVED. MULTIPLE REPAIRS: 1. Individual cutouts separated by less than 1.5 inches, edge to edge, after cleanup shall be considered a single damage and repaired as such, (refer to FIGURE 4-1A). 2. Cutouts separated by 1.5 to 5.0 inches, edge to edge, after cleanup are considered distinct damages in as much as separate core plugs, fillers and inner doublers may be used. However, a single doubler shall cover all damaged areas and a double row of fasteners will be installed through doubler and existing, undisturbed portion of panel between cutouts, (refer to FIGURE 4-1 B). 3. Skin cutouts separated by more than 5 inches, edge to edge, after cleanup are considered distinct damages, (refer to FIGURE 4-1 C). NOTE WHEN DAMAGE TO BE REPAIRED IS CLOSER THAN TWO (2) INCHES FROM EDGE OF AN EXISTING REPAIR DOUBLER, REMOVE EXISTING DOUBLER AND TREAT BOTH NEW AND EXISTING DAMAGES AS ONE. USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
4-3
BHT-206-SRM-1
REVISION 1
FAA APPROVED
CUTLINE RESTRICTION (skin and/or core): When inserts are included in skin or core cutout, extend cutline a minimum of 0.75 inch past center of insert (refer to FIGURE 4-1D). DOUBLERS: Metal doublers will be bonded and riveted to surface of panel. Two rows of fasteners will be used around entire perimeter of doubler and between cutouts when applicable as specified in FIGURE 4-1 B. When laying out rivet pattern on doubler and panel skin, ensure that rivets do not interfere with later installations or opposite skin potted inserts. Edge of doubler to extend a minimum of 1.5 inches past skin cutline except when inserts are in pattern as specified in FIGURE 4-1 D. CORE REPLACEMENT: Core replacement procedure, preparation and bonding in accordance with process sheets, section 3. NOTE DEPENDING ON AMOUNT OF CORE REMOVED, IT IS POSSIBLE TO REPLACE SMALL AMOUNTS OF CORE USING ONLY POTTING ADHESIVE, REFER TO SECTION 3. Core to be replaced with same density core as originally used (high density replaces high density and low density replaces low density). In cases where damage affects both high and low density cores, the following considerations apply. 1. If the splice line for new core is 2.0 inches or less from original high density core line, replace both cores using the higher density core material as shown in FIGURE 4-1 E. 2. If the splice line for new core is more than 2.0 from the original high density core line, replace each core using original density core material as shown in FIGURE 4-1 F. INSERT REPLACEMENT: Refer to applicable process sheet in section 3 for insert removal and installation instructions. The applicable repair procedure will list the specific inserts used in this repair, in the form of a table, in the REQUIRED section.
4-4
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
FAA APPROVED
FIGURE 4-1 GUIDELINES
REVISION 1
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
o
L
BHT-206-SRM-1
MULTIPLE REPAIRS SHEET 1 OF 3 4-5
BHT-206-SRM-1
REVISION 1
0
FAA APPROVED
)
FAA APPROVED
FIGURE 4-1 GUIDELINES
Z
z
)
REVISION 1
USE OR DISCLOSURE OF DATA CONTAINED ON THIS PAGE IS SUBJECT TO THE RESTRICTIONS ON THE TITLE PAGE OF THIS DOCUMENT.
L
C
I