# AC 43.13-1A: Acceptable Methods, Techniques, and Practices - Aircraft Inspection and Repair - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Historical - Version in force: [AC 43.13-1B CHG 1 Ed Update](https://openairdocs.org/ac/43.13-1), approved September 27, 2001. This is a historical version - Approved: January 1, 1972 - Last reviewed: June 11, 2024 - Family: AC 43.13-1 - Cancels: [AC 43.13-1](https://openairdocs.org/ac/43.13-1) - Cancelled by: [AC 43.13-1A CHG 1](https://openairdocs.org/ac/43.13-1A-CHG-1) - 14 CFR parts: 43 - Pages: 356 - PDF: https://openairdocs.org/files/sha256/0f/d9/0fd9bbcd87e53a409f3f8839fc4344b1f6594018336474c760482ac988254b79 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/49E5FBEAF25E12F9862577C800658986.0001 - DRS record last changed: November 21, 2025 - Page: https://openairdocs.org/ac/43.13-1A ## Versions - [AC 43.13-1B CHG 1 Ed Update](https://openairdocs.org/ac/43.13-1) (Current), approved Sep 27, 2001, 756 pages. Editorial Update to AC 43.13-1B CHG 1. - [AC 43.13-1B CHG 1](https://openairdocs.org/ac/43.13-1B-CHG-1) (Historical), approved Sep 9, 1998, 646 pages - [AC 43.13-1B](https://openairdocs.org/ac/43.13-1B) (Historical), approved Sep 8, 1998, 696 pages - [AC 43.13-1A CHG 2](https://openairdocs.org/ac/43.13-1A-CHG-2) (Historical), approved Dec 22, 1976, 358 pages - [AC 43.13-1A CHG 1](https://openairdocs.org/ac/43.13-1A-CHG-1) (Historical), approved May 12, 1975, 356 pages - [AC 43.13-1A](https://openairdocs.org/ac/43.13-1A) (Historical), approved Jan 1, 1972, 356 pages - [AC 43.13-1](https://openairdocs.org/ac/43.13-1?v=E562265B3D7C20FF86257983004DB8A8) (Historical), approved Jan 1, 1965, 346 pages. AC 43.13-1 with all changes incorporated. ## ADs that cite it (95, newest 50) - [AD 2024-19-08](https://openairdocs.org/ad/2024-19-08): Nose/Tail Landing Gear · Diamond Aircraft Industries DA 40 and DA 40F. Effective Nov 29, 2024 - [AD 2024-12-10](https://openairdocs.org/ad/2024-12-10): Main Rotor Drive System · Centerpointe Aerospace S-58BT, S-58DT, S-58ET and 3 more. Effective Aug 26, 2024 - [AD 2023-09-09](https://openairdocs.org/ad/2023-09-09): Exhaust Turbine System (Recip) · Aerostar PA-60-600 (Aerostar 600), PA-60-601 (Aerostar 601), PA-60-601P (Aerostar 601P) and 292 more. Effective Jul 17, 2023 - [AD 2021-04-12](https://openairdocs.org/ad/2021-04-12): Tail Rotor Drive Shaft · Robinson Helicopter R66. Effective Apr 12, 2021 - [AD 2020-16-15](https://openairdocs.org/ad/2020-16-15): Fuselage · Viking Air DHC-2 Mk.I and DHC-2 Mk.III with STC SA00299NY or SA945NE. Effective Sep 21, 2020 - [AD 2015-08-04](https://openairdocs.org/ad/2015-08-04): Wings · FS 2000 L-14, AE-1, HE-1 and 77 more. Effective Jun 3, 2015 - [AD 2015-02-15](https://openairdocs.org/ad/2015-02-15): Elevator skins · Quest Aircraft Design Kodiak 100. Effective Feb 19, 2015 - [AD 99-26-19 R1](https://openairdocs.org/ad/99-26-19-R1): Wings · Piper J-2. Effective Jan 21, 2014 - [AD 99-01-05 R1](https://openairdocs.org/ad/99-01-05-R1) (historical): Wings · FS 2000 L-14, AE-1, HE-1 and 73 more. Effective Jan 14, 2014 - [AD 2013-10-04](https://openairdocs.org/ad/2013-10-04): Engine Exhaust · Piper PA-31, PA-31-325 and PA-31-350. Effective Jul 17, 2013 - [AD 2012-17-06](https://openairdocs.org/ad/2012-17-06): Flight Controls · Piper PA-24, PA-24-250 and PA-24-260. Effective Oct 22, 2012 - [AD 2012-04-10](https://openairdocs.org/ad/2012-04-10): Wings · Rogers, Burl A. 15AC and S15AC. Effective Apr 17, 2012 - [AD 2011-15-11](https://openairdocs.org/ad/2011-15-11): Wings · Nineteen and a half(19.5 U.S.) gallons wing tip fuel tank (Flint Aero STC SA5090NM). Effective Sep 6, 2011 - [AD 2011-05-07](https://openairdocs.org/ad/2011-05-07): Stabilizers · Allied Ag Cat Productions G-164, G-164A, G-164B and 7 more. Effective Mar 17, 2011 - [AD 2010-21-18](https://openairdocs.org/ad/2010-21-18): Wings · Wing tip extension fuel tanks (Aviation Enterprises STC SA02055AT or 3 others). Effective Nov 3, 2010 - [AD 2009-26-11](https://openairdocs.org/ad/2009-26-11): Spar caps · Thrush 600 S-2D, S-2R, S2R-G1 and 13 more. Effective Feb 24, 2010 - [AD 2009-12-15](https://openairdocs.org/ad/2009-12-15): Electric Power · Grob-Werke G120A. Effective Jul 1, 2009 - [AD 2008-26-12](https://openairdocs.org/ad/2008-26-12): Control bridge · BLANIK L 23 Super - Blanik. Effective Jan 28, 2009 - [AD 2008-22-21](https://openairdocs.org/ad/2008-22-21) (historical): Main tubular spar · Allied Ag Cat Productions G-164, G-164A, G-164B and 7 more. Effective Dec 19, 2008 - [AD 2008-09-18](https://openairdocs.org/ad/2008-09-18): Wing strut attach fittings · Taylorcraft A, BF, BF-60 and 34 more. Effective Jun 6, 2008 - [AD 2007-25-01](https://openairdocs.org/ad/2007-25-01): forward and aft control sticks · BLANIK Blanik L-13. Effective Jan 7, 2008 - [AD 2004-23-01](https://openairdocs.org/ad/2004-23-01): Fuel Booster pumps · Pilatus PC-7. Effective Dec 27, 2004 - [AD 2003-24-07](https://openairdocs.org/ad/2003-24-07): Rudder torque tube · Piper PA-31, PA-31-300, PA-31-325 and 7 more. Effective Feb 9, 2004 - [AD 2003-17-16](https://openairdocs.org/ad/2003-17-16): Tailplane Attachment Brackets and Engine Mount · Rust, Robert E. DH.C1 Chipmunk 21, DH.C1 Chipmunk 22 and DH.C1 Chipmunk 22A. Effective Oct 10, 2003 - [AD 2002-19-13](https://openairdocs.org/ad/2002-19-13): Left and Right lower wing surface · Aero Union SP-2H (P2V-7), SP-2H (P2V-7), SP-2H (P2V-7) and 4 more. Effective Oct 1, 2002 - [AD 2001-23-03](https://openairdocs.org/ad/2001-23-03): Forward Door Post · Cessna 172N, 172P, 172RG and 5 more. Effective Dec 27, 2001 - [AD 2000-25-02 R1](https://openairdocs.org/ad/2000-25-02-R1): Wood Spars · American Champion 11AC, 11BC, 11CC and 27 more. Effective Jul 13, 2001 - [AD 2000-25-02](https://openairdocs.org/ad/2000-25-02) (historical): Wood Wing Spars · American Champion 11AC, 11BC, 11CC and 15 more. Effective Jan 19, 2001 - [AD 2000-07-03](https://openairdocs.org/ad/2000-07-03): Fuel Line Assembly · Robinson Helicopter R44. Effective May 11, 2000 - [AD 99-10-07](https://openairdocs.org/ad/99-10-07): Elevator Control Cable · Textron 65-90, 65-A90, 65-A90-1 and 9 more. Effective Jun 8, 1999 - [AD 98-19-14](https://openairdocs.org/ad/98-19-14): Airbrake Control System · Centrair 101, 101A, 101AP and 1 more. Effective Nov 9, 1998 - [AD 98-09-10](https://openairdocs.org/ad/98-09-10): Rudder Control Cables · Extra Flugzeugproduktions- und Vertriebs- EA-300 and EA-300S. Effective Jun 7, 1998 - [AD 96-16-08](https://openairdocs.org/ad/96-16-08): Main Deck Cargo Door · Boeing 727-100 Series and 727-200 Series with STC SA1797SO. Effective Aug 27, 1996 - [AD 96-02-04](https://openairdocs.org/ad/96-02-04): 80 Octane Avgas · Wytwornia Sprzetu Komunikacyjnego 6A4-150-B3 and 6A4-165-B3. Effective Feb 13, 1996 - [AD 94-14-10](https://openairdocs.org/ad/94-14-10): Aviation Gasoline · Wright Aeronautical Division, Curtiss-Wright Cyclone R-1820F-2A, Cyclone R-1820F-3A, Cyclone R-1820F-56 and 6 more. Effective Feb 13, 1996 - [AD 95-12-01](https://openairdocs.org/ad/95-12-01): Forward Wing Spars · Lavia Argentina PA-25, PA-25-235 and PA-25-260. Effective Jul 7, 1995 - [AD 92-15-13](https://openairdocs.org/ad/92-15-13): Wing Structure · Textron 99, 99A, 99A (FACH) and 4 more. Effective Aug 24, 1992 - [AD 90-04-06 R1](https://openairdocs.org/ad/90-04-06-R1): Propeller Governor External Oil Line · Lycoming Engines AEIO-320-D1B, AEIO-320-D2B, AEIO-320-E1A and 427 more. Effective May 28, 1991 - [AD 90-15-07](https://openairdocs.org/ad/90-15-07): Air Conditioning System · Short Brothers SD3-60 with STC SA3749SW. Effective Jul 31, 1990 - [AD 89-18-06](https://openairdocs.org/ad/89-18-06): Front Folding Seats · American Champion 7ECA, 7GCAA, 7GCBC and 3 more. Effective Sep 21, 1989 - [AD 78-26-10 R1](https://openairdocs.org/ad/78-26-10-R1): Cockpit Heater System · Great Lakes 2T-1A-1 and 2T-1A-2. Effective Dec 17, 1986 - [AD 86-21-08](https://openairdocs.org/ad/86-21-08): Battery Tray · SOCATA TB 10 and TB 20. Effective Oct 24, 1986 - [AD 85-04-01 R1](https://openairdocs.org/ad/85-04-01-R1): Hydraulic And Oxygen Line Inspection · M7 Aerospace SA226-AT, SA226-T, SA226-TC and 1 more. Effective Oct 28, 1985 - [AD 85-10-02](https://openairdocs.org/ad/85-10-02): Engine Induction Airbox · Cessna 206, 207, 207A and 24 more. Effective Jun 20, 1985 - [AD 78-23-14 R1](https://openairdocs.org/ad/78-23-14-R1): Tow Release System · BLANIK Blanik L-13 and Blanik L-13 AC. Effective Aug 17, 1984 - [AD 84-15-01](https://openairdocs.org/ad/84-15-01): Placard -Tethered Flight · Avian Balloon Falcon II, Skyhawk and Sparrow. Effective Jul 31, 1984 - [AD 84-13-01](https://openairdocs.org/ad/84-13-01): Generator Control J-Box · M7 Aerospace SA226-TC, SA227-AC (C-26A), SA227-AT and 1 more. Effective Jul 5, 1984 - [AD 84-10-04](https://openairdocs.org/ad/84-10-04): APU Feeder Cable · Boeing 737-100 Series, 737-200 Series, 737-200C Series and 6 more. Effective Jun 4, 1984 - [AD 80-19-12 R1](https://openairdocs.org/ad/80-19-12-R1): Engine Mount Structure · Textron 76. Effective Jun 14, 1982 - [AD 81-23-01 R1](https://openairdocs.org/ad/81-23-01-R1): Beech Kit or Supplemental Type Certificate · Textron 100, 65, 65-88 and 20 more. Effective Jan 4, 1982 All 95: https://openairdocs.org/ac/43.13-1/ads.md ## Text Each page's text from the PDF; open the PDF at a page with #page=N. This is OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. ### Page 1 (OCR) CHANGE REFERENCE. AC NO: 48.18-1A CHG 1 DATE: 5/12/75 **ENT OF FRA SPORT** **ADVISORY** **FO STATES OF CIRCULAR** DEPARTMENT OF TRANSPORTATION DEPARTMENT OF TRANSPORTATION OST-FOB-10A, TAD-494.7 800 INDEPENDENCE AVE S.H. **FEDERAL AVIATION ADMINISTRATION WASH., D.C. 20590-** ACCEPTABLE METHODS, TECHNIQUES, AND SUBJECT: PRACTICES-AIRCRAFT INSPECTION AND RE, PAIR **1. PURPOSE.** This change transmits new and revised material for Advisory Circular 48.13-1A. **2. PRINCIPAL CHANGES.** a. Corrections have been made to paragraph 318 and Figure 10.1. b. Chapter 2, Section 3, has been revised to add Selective Plating in aircraft maintenance. Chapter 4, Section 6 has been revised to add handling and care of aircraft recovered from water immersion. c. A new Section 5, Cabin Interior Materials- Fire Protection Qualities, has been added to Chapter T. d. New information on tire slippage and maintenance has been added to Section 1 of Chapter 8. **3. HOW TO GET THIS PUBLICATION.** Copies of this change to AC 43.18-1A may be obtained by ordering from: Superintendent of Documents U.S. Government Printing Office Washington, D.C. 20402 For up-to-date information on how to order, and the price of this publication, check with the closest FAA office or obtain the latest copy of Advisory Circular 00-2, Advisory Circular Checklist and Status of the Federal Aviation Regulations. This checklist may be obtained without charge by requesting from: Department of Transportation Publications Section, TAD 443.1 Washington, D.C. 20580 Initlated by: AFS-830 PoP Stally R. P. SKULLY, Director Flight Standards Service Initiated by: ### Page 2 (OCR) PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated i thru iv 1972 i 1972 ii thru iv 1975 73 and 74 1972 73 thru 73-3 1975 74 1972 127 and 128 1972 127 1972 128 thru 128-3 1975 189 and 140 1872 189 1972 140 thru 140-1 1975 149 and 150 1972 149 1972 150 1975 165 and 166 1972 165 1975 166 1972 ### Page 3 (OCR) AC 43.13-1A Page 1 1972 TABLE OF CONTENTS Chapter 1. AIRCRAFT WOOD STRUCTURES • Section 1. MATERIALS AND PRACTICES Paragraph Page Materials 2. Replacement of Drainholes and Skin Stiffeners Control Surface Flutter Precautions Gluing Precautions. - Б. Preparation of Wood Surfaces for Gluing . 6. Glues 7. Mixing of Resin Glues 4 8. Spreading of Glue 9. Assembly Time in Gluing 10. Gluing Pressure 11 Scarf Joints 12 Splicing of Spars 18. Splicing of Box Spar Webs 14. Replacing Solid-Type Spars with Laminated Type 15. Spar Longitudinal Cracks and Local Damage 16. Elongated Holes in Spars 17. Rib Repairs 18. Plywood Skin Repair 9 19. Repairs to Single Curvature Plywood Skin 10 20. Repairs to Double Curvature Plywood Skin 10 21. Splayed Patch 10 22. Surface Patch 16 23. Scarf Patch 17 24. Plug Patches 17 25. Fabric Patch 28 26.-86. Reserved 28 Section 2. FINISHING WOOD STRUCTURES 87. General 24 88. Precautions to be Observed 24 89. Finishing of Interior Surfaces 24 40. Finishing of Exterior Surfaces 24 41. Finishing of End Grain Surfaces 25 42. Finishing with Fabric or Tape 25 48.-58. Reserved - 25 Chapter 2. AIRCRAFT METAL STRUCTURES Section 1. REFERENCES AND PRECAUTIONARY MEASURES 54. References 27 55. Flutter and Vibration Precautions 27 66. Brazing 28 07.-67. Reserved 28 Section 2. WELDING 68. General 29 69. Steel Parts Not to be Welded . 82 70. Repair of Tubular Members 82 ### Page 4 (OCR) Page lI AC 43.13-1A CHG 1 1975 TABLE OF CONTENTS-Continued Paragraph Page 71. Repalr by Welded Sleeve 35 72. Repair by Bolted Sleeve 35 79. Welded-Patch Hepair 35 74. Splicing Tubing by Inner Sleeve Method 36 75. Splicing Tubing by Outer Sleeve Method 38 78. Splicing Using Larger Diameter Replacement Tubes 39 77. Repairs at Bullt-In Fuselage Fittings 40 78. Engine Mount Repairs 40 79. Landing Gear Repairs 40 80. Built-Up Tubular Wing or Tall Surface Spars 47 81. Wing and Tail Brace Struts 47 82. Repairs to Welded Parts 47 83. Stainless Steel Structure 47 81.-94. Reserved 48 Section 3. METAL REPAIR PROCEDURES 95. Riveted or Bolted Steel Truss-Type Structures 49 96. Aluminum Alloy Structures 48 97. Selection of Aluminum for Replacement Parts 50 98. Heat Treatment of Aluminum Alloy Parts 50 09. Riveting 51 100. Repair Methods and Precautions for Aluminum Structure 54 101. Repairing Cracked Members -_-- 102. Steel and Aluminum Fittings -- 103. Castings * 104. Selective Plating in Alrcraft Maintenance 73 * 105.-114. Reserved 73-3 Section 4. REPAIR OF LAMINATE STRUCTURES 115. General 74 110.-128. Reserved 78 Chapter 3. FABRIC COVERING Section 1. PRACTICES AND PRECAUTIONS 127. Textile Materials 81 128. Covering Practices 81 129. Preparation of the Structure for Covering . 130. Fabric Seams -- 181. Covering Methods 84 132. Relnforcing Tape 86 133. Lacing 86 134. Stitch Spacing 81 185, Surface Tape (Finishing Tape) - 87 136. Special Fasteners 137.-147. Reserved 91 Section 2. DOPING 148. Application of Aircraft Dope, Epoxy, and Resins 92 149. Blushing and Use of Blush-Retarding Thinner 92 150. Number of Coats . 92 151. Technique - 93 152. Rejuvenation of Fabric 93 153. Common Dope Troubles 94 154.-164. Reserved 94 ### Page 5 (OCR) AC 43.13-1A CHO 1 Page Hi 1975 TABLE OF CONTENTS-Continued Paragraph Page Section 3. REPAIRS TO FABRIC COVERING 105. General 95 166. Repair of Tears In Fabric 95 167. Sewed Patch Repair --- 96 168. Unsewed (Doped-On) Repairs 90 160. Repair by a Doped-In Panel 90 170.-180. Reserved 97 Section 4. FABRIC TESTING 181, Testing of Fabric Covering 98 182.-102. Reserved 98 Chapter 4. CONTROL CABLES AND TERMINALS Section 1. INSPECTION AND REPAIR 193. General 184. Cable Specifications 105. Replacement of Cables 190. Mechanically Fabricated Cable Assemblies 187. Hand Fabricated Cable Assembles 188. Cable System Inspection 199. Cable Tension Adjustment 110 200. Corrosion and Rust Prevention . 110 201.-211. Reserved 111 Section 2. SAFETY METHODS FOR TURNBUCKLES 212. General 213. Double Wrap Method 214. Single Wrap Method 114 215. Special Locking Devices 119 216.-226. Reserved 114 Chapter 5. AIRCRAFT HARDWARE Section 1. IDENTIFICATION AND USE OF AIRCRAFT HARDWARE 227. Bolta 115 228. Screws 117 229. Taper Pins (AN-385 and AN-386) - 119 230. Nuts 110 231. Washers - 120 292. Types of Rivets 233. Material Applications 284. Fasteners (Cowl and Fairing) 285. Unconventional Attachments 286.-248. Reserved 122 Chapter 6. CORROSION PROTECTION 247. Types of Corrosion 248. Corrosion Protection Measures for Basic Materials 249. Construction of Corrosion Resistant Structure 250. Corrosion Removal 251. Corrosion Proofing of Landplanes and Seaplanes 127 252. Corrosion Proofing of Landplanes Converted to Seaplanes 127 258. Cleaners, Polishes, Brighteners 128 * 255-284. Reserved 254. Handling and Care of Aircraft Recovered from Water Immersion 128 * 128-8 ### Page 6 (OCR) Page Iv AC 43.13-1A CHG 1 1975 TABLE OF CONTENTS--Continued Chapter 7. IDENTIFICATION, TESTING, AND INSPECTION OF MATERIALS Paragraph Page Section 1. IDENTIFICATION OF METALS 265. IdentIfication of Steel Stock 129 208. Interchangeability of Steel Tubing - 120 287. Identification of Aluminum 120 288.-278. Reserved 181 Section 2. TESTING OF METALS 279. Hardness Testing 132 280. Brinell Hardness Test 132 281. Rockwell Hardness Test 182 282. Vickers Hardness Test 182 283. Testing Aluminum 182 284.-294. Reserved 132 Section 3. NONDESTRUCTIVE TESTING 205. General 186 206. Inspection by Magnifying Glass After Welding 136 207. Magnetic Particle Inspection 186 208. X-Ray or Radiographic Inspection 137 299. Fluorescent Penetrant 188 300. Dye Penetrants 188 801. Ultrasonle Flaw Detection 139 802.-312. Reserved 189 Section 4. IDENTIFICATION OF FABRICS AND PLASTICS 818. Identification of Fabric 140 * 814. Identification of Plastics 140 315.-320. Reserved 140 * Section 5. CABIN INTERIOR MATERIALS, FIRE PROTECTION QUALITIES 321. Requirements for Fire Protection 140-1 922.-325. Reserved 140-1* Chapter 8. AIRCRAFT EQUIPMENT Section 1. LANDING GEAR EQUIPMENT 926. General 141 927. Inspection 141 328. Cleaning and Lubricating 145 329. Emergency Systems 145 330. Special Inspections 145 381. Retraction Tests 145 332. Tires 145 * 383. Tire Slippage - 150 834. Tire Maintenance 150 * 385.-348. Reserved 150 Section 2. LIFE PRESERVERS AND LIFERAFTS 344. General 151 345. Inspection Procedure for Life Preservers 151 346. Repair of Life Preservers 151 847. Inspection Procedure for Liferafts 151 348. Repair of Liferafts 152 849.-359. Reserved 152 ### Page 7 (OCR) AC 43.13-1A CHG 1 Page iv- 1975 TABLE OF CONTENTS-Continued Paragraph Page Section 3. MISCELLANEOUS EQUIPMENT 360. Parachutes 153 361. Safety Belta 158 382. Flares 154 863. Oxygen Systems 151 364.-374. Reserved 158 ### Page 8 (OCR) AC 43.13-1A CHG 1 1975 all highly stressed main fittings, as set forth in Chapter 7, and if necessary, take corrosion prevention measures as recommended in Chapter 6. (3) Replace torn, kinked, or cracked fittings. (4) Elongated or worn boltholes in fittings which were designed without bushings are not to be reamed oversize. Replace such fittings, unless the method of repair is approved by a representative of the FAA. Do not fill holes with welding rod. Acceptable methods of repairing elongated or worn boltholes in landing gear, stabilizer, interplane, or cabane-strut ends only, not originally equipped with pin plates, are shown in figure 2.36. (Also see figure 2.11 on longeron repair at fitting.) b. Aluminum and Aluminum Alloy Fittings. (1) Replace damaged fittings with new parts, having the same material specifications. (2) Repairs may be made in accordance with data furnished by the aircraft manufacturer or data substantiating the method of repair may be submitted to the FAA for review. 103. CASTINGS. Damaged castings are to be replaced and not repaired unless the method of repair is specifically approved by the aircraft manufacturer or substantiating data for the repair has been reviewed by the FAA. * 104. SELECTIVE PLATING IN AIRCRAFT MAINTE- NANCE. Selective plating is a method of depositing metal from an electrolyte to the selected area. The electrolyte is held in an absorbent material attached to an inert anode. Plating contact is made by brushing or swabbing the part (cathode) with the electrolyte-bearing anode. a. Selective Plating Uses. This process can be utilized for any of the following reasons: (1) To prevent or minimize disassembly, reassembly, or masking costs. (2) Resizing worn components (plate to size). (3) Filling in damaged or corroded areas. (4) To plate small areas of extremely large parts. (5) To plate electrical contacts. (6) To plate parts too large for existing baths. Chap 2 Page 73 (7) To supplement conventional plating. (8) To plate components which become contaminated if immersed in a plating bath. (9) To cadmium-plate ultra high strength steels without hydrogen embrittlement. (10) On-site plating. 111) Reverse current applications (e.g., stain removal, deburring, etching, dynamic balancing). b. Specifications. Selective plating (electrodeposition), when properly applied, will meet the following specifications and standards: QQ-C-320 Chromium Plating QQ-N-290 Nickel Plating QQ-P-416 Plating, Cadmium QQ-5-365 Silver Plating 22-1-325 Zinc Plating MIL-T-10727 Tin Plating MIL-C-14550 Copper Plating MIL-G-45204 Gold Plating c. General Requirements. (1) Areas to be repaired by this process should be limited to reasonably small areas of large parts, although it may be desirable to plate small parts, particularly electrical or electronic parts, in their entirety. (2) All solutions should be kept clean and free from contamination. Care should be taken to insure that the solutions are not contaminated by used anodes or other plating solutions. Brush plating solutions are not designed to remove large amounts of scale, oil, or grease. Mechanical or chemical methods should be used to remove large amounts of scale or oxide. Use solvents to remove grease or oil. (3) Brush plating solutions are five to fifty times as concentrated as tank solutions. The current densities used range from 500 to 4,000 amps/feet?. The voltages listed on the solution bottles have been precalculated to give proper current densities. Too high a current density burns the plating, while too low a current density produces stressed deposits and low efficiencies. Agitation is provided by anode/cathode motion. Too fast a motion results in low efficiencies and stressed deposits; too slow a motion causes burning. A dry tool results in burnt plate, coarse grain structure, and unsound deposits. The tool* Par 102 ### Page 9 (OCR) *cannot be too wel. Solution temperatures of 110° to 120° F. are reached during operation. (4) Materials such as stainless steel, aluminum, chromium, and nickel, which have a passive surface, will require an activating operation to remove the passive surface, During the activating process, do not use solutions that have been previously used with reverse current because of solution contamination. d. Equipment. The power source should operate on either 110 or 220 volt alternating current (AC), 60 Hertz, single phase input. It should have a capability to produce direct current having smooth characteristics with controlled ripple and be able to output a current of at least 25 amperes at 0 to 25 volts. Minimum instrumentation of the power source should include a voltmeter, ammeter, and ampere-hour meter. (1) The ammeter should provide a full scale reading equal to the maximum capacity of the power source with an accuracy of +5 percent of the current being measured. (2) The voltmeter should have sufficient capacity to provide a full scale reading equal to the maximum capacity of the power source and an accuracy of +1.0 volt. (3) An ampere-hour meter should be readable to 0.001 ampere-hour and have an accuracy of 10.01 ampere-hour. (4) The stylus should be designed for rapid cooling and to hold anodes of various sizes and configurations. For safety, the anode holder should be insulated. (5) The containers for holding and catching runoff solutions should be designed to the proper configuration and be inert to the specific solution. (6) The mechanical cleaning equipment and materials should be designed and selected to prevent contamination of the parts to be cleaned. e. Materlals. The anodes should be of highpurity dense graphite or platinum-iridium alloys. Do not mix solutions from different suppliers. This could result in contamination. f. Detail Requirements. On large parts, no area greater than approximately 10 percent of the total area of the part should be plated by this process. Small parts may be partially or completely plated. Special cases exceeding these Par 104 AC 43.13-1A CHG 1 1975 or oil. Alloys. (1) Electroclean the area using forward (direct) current until water does not break on the surface. This electroclean process should be accomplished at 10 to 15 volts, using the appropriate electroclean solution. (2) Rinse the area in cold clean tap water. (3) Activate the area with reverse current, 7 to 10 volts, in conjunction with the proper activating solution until a uniform gray to black surface is obtained. (4) Rinse thoroughly in cold, clean tap water. (5) Immediately electroplate to color while the area is still wet, using the appropriate nickel solution. (6) Rinse thoroughly. (7) Immediately continue plating with any other solution to desired thickness. (8) Rinse and dry. k. Plating on Copper and Copper Base Alloys. (1) Electroclean the area using forward (direct) current until water does not break on the surface. The electroclean process should be accomplished at 8 to 12 volts using the appropriate electroclean solution. * Chap 2 ### Page 10 (OCR) AC 43.13-1A CHG 1 1975 * (2) Rinse the area in cold, clean tap water. (3) Immediately electroplate the area with any of the plating solutions except silver. Silver requires an undercoat. (4) Rinse and dry. I. Plating on 300 and 400 Series Stainless Steels, Nickel Base Alloys, Chrome Base Alloys, High Nickel Ferrous Alloys, Cobalt Base Allays, Nickel Plate, and Chrome Plate. (1) Electroclean the area using forward (direct) current until water does not break on the surface. This electroclean process should be accomplished at 12 to 20 volts using the appropriate electrocleaning solution. (2) Rinse the aren in cold, clean tap water. (3) Activate the surface using forward (direct) current for 1 to 2 minutes using the activating solution and accomplish at 6 to 20 volts. (4) Do not rinse. (5) Immediately nickel flash the surface to a thickness of 0.00005 to 0.0001 inch using the appropriate nickel solution. (6) Rinse thoroughly. (7) Immediately continue plating with any other solution to desired thickness. (8) Rinse and dry. m. Plating on Low-Carbon Steels (Heat Treated to 180,000 psi). (1) Electroclean the area using forward Break an (direct) current until water does not break on the surface. This electroclean process should be accomplished at 12 to 20 volts using the appropriate electrocleaning solution. (2) Rinse the area in cold, clean tap water. (3) Reverse current etch at 8 to 10 volts, using the appropriate activating solution, until a uniform gray surface is obtained. (4) Rinse thoroughly. (5) Immediately electroplate the part using any solutions except copper acid or silver. Both these require undercoats. (6) Rinse and dry. n. Plating on Cast Iron and High-Carbon Steels (Steels Heat Treated to 180,000 psi). (1) Electroclean the area using forward (direct) current until water does not break on Chap 2 Page 73-2 the surface. This electroclean process should be accomplished at 12 to 20 volts using the appropriate electrocleaning solution. (2) Rinse the area thoroughly in cold, clean tap water. (3) Reverse current etch at & to 10 volts, using the appropriate etching solution, until a uniform gray is obtained. (4) Rinse thoroughly. (5) Remove surface smut with 15 to 25 volts using the appropriate activating solution. (6) Rinse thoroughly. (7) Electroplate immediately using any of the solutions except copper or ailver (both these require undercoats). (8) Rinse and dry. o. Plating on Ultra High Strength Steels (Heat Treated Above 180,000 psi). (1) Electroclean the area using REVERSE current until water does not break on the surface. This electroclean process should be accomplished at 8 to 12 volts using the appropriate electroclean solution. (2) Rinse the area thoroughly in cold, clean tap water. (3) Immediately electroplate the part, using either nickel, chromium, gold, or cadmium. Other metals require an undercoat of one of the above. Plate initially at the highest voltage recommended for the solution so as to develop an initial barrier layer. Then reduce to standard voltage. (4) Rinse and dry. (5) Bake the part for 4 hours at 375° +25° F. NoTE 1: Where the solution vendor provides substantiating data that hydrogen embrittlement will not result from plating with a particular solution, then a post bake is not required. This substantlating data can be in the form of alrcraft industry manufacturers' process specifications, military specifications, or other sultable data. NoTE 2: Acid etching should be avolded, if posible. Where etching is absolutely necessary, it should always be done with reverse current. alkallne solutions for initial deposita. p. Dissimilar Metals and Changing Base. As a general rule, when plating two dissimilar metals, follow the plating procedure for the one with the most steps or activation. If activating* Par 104 ### Page 11 (OCR) * steps have to be mixed, use reverse current activation steps prior to forward (direct) current activation steps. q. Plating Solution Selection. (l) Alkaline and neutral solutions are to be used on porous base metals, white metals, highstrength steel, and for improved coating ability. Acid solutions are to be used for rapid buildup and as a laminating structure material in conjunction with alkaline type solutions. (2) Chrome brush plating solutions do not yield as hard a deposit as bath plating solutions. The hardness is about 600 Brinell as compared to 1,000 Brinell for hard chrome deposited from a tank. (3) Silver immersion deposits will form with no current flowing on most base metals from the silver brush plating solutions; such deposits have poor adhesion to the base metal. Consequently, a flash or a more noble metal should be deposited prior to silver plating to develop a good bond. (4) In general, brush plating gives less hydrogen embrittlement and a lower fatigue strength loss than does equivalent tank deposits. However, all brush-plated, ultra high strength steel parts (heat treated above 180,000 psi) Par 104 AC 43.13-1A CHG 1 1975 should be baked as mentioned in paragraph 104.0.(5), unless it is specifically known that embrittlement is not a factor. r. Qualification Tests. All brush plated surfaces shall be tested for adhesion of the electrodeposit. Apply a 1-inch wide strip of Minnesota Mining and Manufacturing tape code 250, or an approved equal, with the adhesive side to the freshly plated surface. Apply the tape with heavy hand pressure and remove it with one quick motion perpendicular to the plated surface. Any plating adhering to the tape shall be cause for rejection. s. Personnel Training for Quality Control. Manufacturers of selective plating equipment provide training in application techniques at their facilities. Personnel performing selective plating must have adequate knowledge of the methods, techniques, and practices involved. These personnel should be products of those training programs and certified as qualifed operators by the manufacturers of the products used, as well as by local quality control depart* ments. **105.-114. RESERVED.** Chap 2 ### Page 12 (OCR) AC 43.13-1A CHG 1 Page 73-4 1975 ### Page 13 (OCR) Page 74 AC 43.13-1A 1972 Section 4. REPAIR OF LAMINATE STRUCTURES 115. GENERAL. There is a wide variation in the composition and structural application of laminates, and it is essential that these factors be given major consideration when any restoration activities are undertaken. To a similar extent, there also exist many types of laminate structure repairs that may or may not be suitable for a given condition. For this reason, it is important that the aircraft or component manufacturer's repair data be reviewed when determining what specific type of repair is permissible and appropriate for the damage at The materials used in the repair of laminate structures must preserve the strength, weight, aerodynamic characteristics, or electrical properties of the original part or assembly. This can best be accomplished by replacing damaged material with material of identical chemical composition or a substitute approved by the manufacturer. In order to eliminate dangerous stress concentrations, avoid abrupt changes in cross-sectional areas. Whenever possible, for scarf joints and facings, make small patches round or oval-shaped, and round the corners of large repairs. Smooth and properly contour aerodynamic surfaces. It is recommended that test specimens be prepared at the same time that the actual repair is accomplished. These can then be subjected to a destructive test to establish the quality of the adhesive bond in the repaired part. To make this determination valid, the specimens must be assembled with the same adhesive batch mixture and subjected to curing pressure, temperature, and time identical with those in the actual repair. a. Fiberglass Laminate Repairs. The following repairs are applicable to fiberglass laminate used for fairings, covers, cowlings, honeycomb panel facings, etc. Prior to undertaking the repair, clean the repair area thoroughly with a castile soap and warm water. Remove any paint by wet or dry sanding methods. Seed blasting may be used but caution must be exercised to not abrade the surfaces excessively. Superficial scars, scratches, surface abrasion, or rain erosion can generally be repaired by applying one or more coats of a suitable resin, catalyzed to cure at room temperature, to the abraded surface. The number of coats required will depend upon the type of resin and severity of the damage. Damage not exceeding the first layer or ply of fiberglass laminate can be repaired by filling with a putty consisting of a compatible roomtemperature-setting resin and clean short glass fibers. Before the resin sets, apply a sheet of cellophane over the repair area and work out any bubbles and excess resin. After the resin has cured, sand off any excess and prepare the area for refinishing. Damage deep enough to seriously affect the strength of the laminate (usually more than the first ply or layer of fabric) may be repaired as illustrated in figure 2.37. Coat the - DAMAGE TO FACING **DAMAGED MATERIAL** **REMOVED BY SANDING , REPLACED LAMINATIONS** SANDED TO CONTOUR FIGURE 2.37.-Typical laminate (facing) repair. Par 115 Chop 2 ### Page 14 (OCR) AC 43.13-1A 1972 in any cleaning process to avoid unnecessary breaking of the protective film, particularly at the edges of the aluminum sheet. Chromic acid and other inhibitive treatments tend to restore the oxide film. d. More severe cleaning is necessary with intergranular corrosion (attack along grain boundaries). The mechanical removal of all corrosion products and visible delaminated metal layers must be accomplished in order to determine the extent of the destruction and to evaluate the remaining structural strength of the component. Inspection with a 5- to 10- power magnifying glass or the use of dye penetrant will assist in determining if all unsound metal and corrosion products have been removed. Grinding to blend or fair out the edges of damaged areas can best be accomplished by using aluminum oxide-impregnated rubberbase wheels. Chemically inhibit the exposed surfaces and restore chemical surface films or paint in the same manner as for other aluminum surfaces. e. Magnesium is the most chemically active of the metals used in aircraft construction and is therefore the most difficult to protect. The prompt and complete correction of the coating failure is imperative if serious structural damage is to be avoided. Treat the corroded area with 10 percent chromic acid solution to which has been added approximately 20 drops of battery electrolyte per gallon, in the same manner as for aluminum alloys. 251. CORROSION PROOFING OF LANDPLANES AND SEAPLANES. In the repair or alteration of aircraft, use corrosion proofing materials the same as, or equivalent to, that originally applied unless the repair or alteration would result in increased susceptibility to corrosion, in which case, employ additional corrosion protection measures. 252. CORROSION PROOFING OF LANDPLANES CONVERTED TO SEAPLANES. A special problem is encountered in the conversion of landplanes to seaplanes. In general, landplanes do not receive corrosion proofing to the same extent as do seaplanes manufactured as such. Corro- Chap 6 Page 127 sion-proofing standards for landplanes converted to seaplanes are divided into two classes: (1) Necessary minimum precautions; and (2) Recommended precautions. Regardless of such precautions, it is imperative that the exterior surfaces of seaplanes be washed with clear fresh water immediately following extended water operation, or at least once a day when operated in salty or brackish water. Wash interior surfaces of seaplanes exposed to spray, taking care to prevent damage to electrical circuits or other items subject to injury. a. Necessary Minimum Precautions. The following procedures are considered the minimum to safeguard the airworthiness of the converted aircraft and are not in themselves intended to maintain airworthiness for an indefinite period. (1) Unless already protected, treat exposed fittings or fittings which can be reached through inspection openings with two coats of zinc chromate primer, paralketone, nonwatersoluble heavy grease, or comparable materials. This applies to items such as wing-root fittings, wing-strut fittings, control-surface hinges, horns, mating edges of fittings, and attach bolts, etc. (2) Coat nonstainless control cables with grease or paralketone or other comparable protective coating, if not replaced with corrosion-resistant cables. (3) Inspect all accessible sections of aircraft structure. Clean structural parts showing corrosion and refinish if corrosion attack is superficial. If a part is severely corroded, replace with an adequately corrosion-proofed part. b. Recommended Precautions. The recommended precautions are those which are suggested as a means of maintaining such aircraft in condition for safe operation over extended periods of time. (1) Provide additional inspection openings to assist in detecting corrosion. Experience has shown openings to allow inspection of lower and rearward portion of the fuselage to be particularly desirable. (2) Incorporate additional provisions for free drainage and ventilation of all interiors to Par 250 ### Page 15 (OCR) prevent collection of moisture (scoop-type drain grommets). (3) Protect the interior of structural steel tubing. This may be done by air and watertight sealing or by flushing with hot linseed oil and plugging the openings. Inspect tubing for miss. ing sealing screws, presence of entrapped water, local corrosion around sealing screws, welded clusters, and bolted fittings which may be indicative of entrapped moisture. 14) Shit the fabric of fabric-covered aircraft longitudinally on the bottom of the fuselage and tail structure for access to these sections. Coat the lower structural members with zinc chromate primer (two coats); follow by a coat of dopeproof paint or wrap with cellophane tape and rejoin the fabric. This precaution is advisable within a few months after start of operation as a seaplane. (5) Spray the interior of metal-covered wings and fuselages with an adherent corrosion inhibitor. (6) Place bags of potassium or sodium dichromate in the bottom of floats and boat hulls to inhibit corrosion. (7) Prevent the entry of water by sealing, as completely as possible, all openings in wings, fuselage, control-surface members, openings for control cables, tail-wheel wells, etc. 253. CLEANERS, POLISHES, BRIGHTENERS. It is important that aircraft be kept thoroughly clean of deposits containing contaminating substances such as oil, grease, dirt, and other foreign materials. a. Materials. Avoid damage to aircraft by not using harmful cleaning, polishing, brightening, or paint-removing materials. Use only those compounds which conform to existing government or established industry specifications or products that have been specifically recommended by the aircraft manufacturer as being satisfactory for the intended application. Observe the product manufacturer's recommendations concerning use of his agent. b. Chemical Cleaners. Chemical cleaners must be used with great care in cleaning assembled aircraft. The danger of entrapping corrosive materials in faying surfaces and crevices coun- Par 252 AC 43.13-1A CHG 1 1975 taracts any advantages in their speed and effectiveness. Use materials which are relatively neutral and easy to remove. c. Removal of Spilled Battery Acid. In order to neutralize spilled battery acid, use sodium bicarbonate (baking soda), or sodium borate (borax) 20 percent by weight dissolved in water. After neutralization, remove alkali salt completely with copious quantities of water to prevent corrosion. An application of acidproof paint to the structure surrounding the battery may be an effective control for this type of cor- * 254. HANDLING AND CARE OF AIRCRAFT RECOV- ERED FROM WATER IMMERSION. Aircraft which were recovered from partial or total immersion in water, including flash floods, have been allowed to air dry, in certain instances, with no safety precautions other than a cursory inspection of the aircraft exterior. The lack of an adequate cleanup of water-immersed areas may subsequently adversely affect the safety of the aircraft. That is, water immersion increases the probability of corrosive attack, the removal of lubrication, the deterioration of aircraft materials, and/or degradation of electrical and avionic equipment. Sea water, because of salt content, is more corrosive than surface fresh water. However, fresh water may contain varying amounts of salt and, as drying occurs, the salt concentration is increased and corrosive attack accelerated. The most important factor following recovery of an aircraft from sea or fresh water immersion is prompt action. Components of the aircraft which have been water immersed, such as the powerplant, accessories, airframe sections, actuating mechanisms, screws, bearings, working surfaces, fuel and oil systems, wiring, radio, and radar should be disassembled, to the extent considered necessary, so that the contaminants can be completely removed. a. Initial Fresh Water/Detergent Wash. As soon as possible after the aircraft is recovered from water immersion, thoroughly wash contaminated internal and external areas of the aircraft using a water/detergent solution as follows: (1) Mix liquid detergent (MIL-D-16791, Type I) and isopropyl alcohol (TT-I-735) in* Chap 6 ### Page 16 (OCR) AC 43.13-1A CHG 1 1975 *ratio of eight parts detergent to 20 parts of alcohol. Add the detergent/alcohol mixture to 72 parts of tap water and mix thoroughly. For use, add one part of the foregoing concentrate to nine parts of tap water (warm water if available) and mix thoroughly. (2) If the above specified detergent/alcohol materials are not available, use water emulsion cleaning compound (MIL-C-43613). Add one part compound to nine parts water. If the MIL cleaning compound is not available, use any available mild household detergent solution with fresh tap water. b. Safety Precautions. The following safety precautions should be observed: (1) Electrically ground the aircraft. Attach the ground lead to the aircraft at a point which is outside the area that could contain explosive vapors. (2) If the landing gear of land planes is used as a supporting mechanism, install a spreader bar, jury strut, landing gear downlocks or other suitable devices to insure that the gear will not collapse. If the landing gear is not serviceable, insure that the aircraft is solidly supported to prevent hazardous movement. (3) Disconnect and remove wet- and/or drycell batteries and isolate aircraft from all sources of electricity or other spark-producing devices. Spark-producing static electricity is genernted at compressed air hose outlets, so this method should not be used for ventilating or purging fuel vapors. (4) Remove all fuel, oil, and hydraulic fluid. 15) Flush all fuel and oil cells with clean fresh water. (6) Deflate tires, especially on magnesium wheels. Depressurize landing genr struts, pneumatic systems, and hydraulic accumulators. c. Reciprocating Engines and Propellers. The propeller should be removed from the engine and the engine from the aircraft. The exterior of the engine and propeller should be washed with stenm, or hot or cold fresh water. The major accessories, engine parts, etc., should be removed and all surfaces flushed with fresh water, preferably hot. If facilities are available, the removed parts, size permitting, should be Chap 6 Page 128-1 immersed in hot water or hot oil, 180° F., for a short period of time. Soft water is preferred and should be changed frequently. All parts must be completely dried by air blast or other means. If no heat drying facility is available, wipe the cleaned parts with suitable drying cloths. Constant speed propeller mechanism should be disassembled, as required, to permit complete decontamination. Clean parts with steam or hot or cold fresh water. Dry the cleaned parts in an oven, but if a heat drying facility is not available, wipe the cleaned parts with suitable drying cloths. d. Gas Turbine, Turboprop and Turboshaft Engines. The engine exhaust shield, insulation blankets, separate exhaust collectors, compressor housing, pinion cowl and/or upper pinion housing should be removed. The engine accessories, outer housing exhaust shields, and other exposed parts should be steam cleaned. The steamcleaned parts should be dried in an oven at approximately 200° F., or with hot air from a portable engine heater, or with clean wiping cloths. Immerse remainder of engine in 10-20 percent water solution of sodium dichromate, or hot, fresh water and apply sufficient agitation to provide complete flushing. If immersion cannot be accomplished, flood the lower section of the compressor housing and thoroughly flush rotor blades with sodium dichromate solution or warm fresh water. Seal off combustion chamber openings and alternately fill and drain the combustion chamber with fresh water. Dry clean parts in an oven, or use dry air from a portable engine heater or wipe with suitable drying cloths. e. Airframe. 'The salvageable components of the fuselage, wings, empennage, seaplane and amphibian hulls and floats, and movable surfaces should be processed as follows: (1) The fabric from fabric-covered surfaces should be removed and replaced. (2) The aircraft interior and exterior should be cleaned using steam under pressure with steam-cleaning compound. The steam should be directed into all seams and crevices where corrosive water may have penetrated. Avoid steam * Par 254 ### Page 17 (OCR) *cleaning electrical equipment, such as terminal boards and relays. (3) Areas that have been steam cleaned should be rinsed immediately with either hot or cold fresh water. (4) Touch up all scratches und scars on painted surfaces using zinc chromate primer or preservative. (5) Undrained hollow spaces or fluid entrapment areas should be provided temporary draining facilities by drilling out rivets at lowest point. Install new rivets after drainnge. (6) All leather, fabric upholstery, and insulation should be removed and replaced. Plastic or rubber foam which cannot be clenned of all corrosive water should be replaced. (7) All drain plugs or drive screws in tubular struotures should be removed and the structure blown out with compressed air. If corrosive water has reached the tubular interiors, carefully flush with hot, fresh water and blow out water with compressed air. Roll the structure as necessary to remove water from pockets. Fill the tubes with hot linseed oil at upproximately 180° F. Drain oil and replace drain plugs or drive screws. (8) Clean sealed wood, metalite, and other nonmetallic areas, excluding acrylic plastics, with warm water. Wood, metalite, and other porous materials exposed to water immersion should be replaced, unless surfaces are adequately sealed to prevent penetration by corrosive water. Virtually all solvents and phenolic type cleaning agents are detrimental to acrylics and will either soften the plastic or cause crazing. f. Helicopter Rotor Dynamic Components. (1) All evidence of corrosive water should be removed from the exterior of transmissions and gear boxes by flushing with clean hot or cold fresh water. (2) Where it is possible that corrosive water has reached the interior of the transmission or gear boxes, remove plugs and/or covers and drain completely. Flush interior of part with hot or cold fresh water. Drain residual water, replace plugs and/or covers, and reapply proper lubricant. g. Helicopter Blades. Except for blades with wooden or other nonmetallic constructions, treat Par 254 AC 43.13-1A CHG 1 1975 helicopter blades the same as propellers for reciprocating engines. Clean nonmetallie blades by hand, using warm water. Dry with wiping cloths or a warm air blast. When cleaning helicopter rotor blades, insure that nonsealed hollow members, such as the spars and blade pockets, are cleaned and dried. h. Fuel and Oil Systems. Contaminated fuel and oil systems should be processed as follows: (1) Flush oil system, inchiding lines, using water-displacing preservative. (2) Open fuel systems. Purge cells and fuel lines. Check effectiveness of the purging with combustible gas indicator. Remove bladder and self-sealing type fuel cells and all cavity liners or pads. (3) Use clean, fresh water to wash fuel cells. After drying, spray interior walls with oil. (4) Chemically treat bare metal surfaces of cleaned aluminum tanks with 10-20 percent water solution of sodium dichromate. (5) Flush fuel lines with hot water (150° F. maximum). Dry, using clean, dry, compressed air. i. Landing Gear. Process salvageable components of the landing gear, wheels, and brakes as follows: (1) Remove tires and wheels. (2) Steam clean wheels, rinse in fresh water, and dry. (3) Remove wheel bearings and clean, using dry cleaning solvent. (4) Immerse bearings in methyl alcohol and dry the cleaned brarings in air blast. Do not permit bearings to rotate during air blast drying. Reapply proper lubricant. (5) Remove brakes, steam clean, and rinse with fresh water and dry. j. Electrical Equipment. (1) Wet cell batteries. The risk involved in using wet cell batteries that have been immersed in sea water may outweigh any economic advantage and should be replaced. (2) Because of posable flight hazards and later defects caused by progressive corrosive attack, all electrical wiring immersed in corrosive water should be replaced. If wiring is* Chap 6 ### Page 18 (OCR) AC 43.13-1A CHG 1 Poge 128-3 1975 * merely splashed or sprayed with corrosive water, flush thoroughly with clean, fresh water and dry, using compressed air. Following compressed air drying, coat with water dispensing preservative (MIL-C-81309, Grade A). k. Miscellaneous Equipment. The following equipment should be thoroughly washed to remove dirt, salt, and other contaminants. Dry with air blast or other means, and, if required, reapply proper lubricant: (1) Wiring and fuselage hydraulic units. (2) Electric landing gears. (3) Actuators. (4) Cables. (5) Accumulators. (6) Hydraulic reservoirs. (7) Flight control. (8) Torque tubes and bell cranks. (9) Heating units, ducts, etc. * **255.-264. RESERVED.** Chap 6 Par 254 ### Page 19 (OCR) AC 43.13-1A 1972 essary to remove anodic films from parts to be inspected, since the dye readily penetrates such films. Special procedures for removing the excess dye should be followed. b. Application of Penetrant. The penetrant is applied by brushing, spraying, or by dipping and allowed to stand for a minimum of 2 minutes. Dwell time may be extended up to 15 minutes, depending upon the temperature of part and fineness of the defect or when the surface being inspected is ground. Parts being inspected should be dry and heated to at least 70° F., but not over 130° F. Very small indications require increased penetration periods. c. Removal of Dye Penetrant, Surplus penetrant is usually removed by application of a special cleaner or remover, or by washing with plain water and the part allowed to dry. Water rinse may also be used in conjunction with the remover, subject to the manufacturer's recommendations. d. Application of Developer. A light and even coat of developer is applied by spraying, brushing, or dipping. When dipping, avoid excess accumulation. Penetrant which has penetrated into cracks or other openings in the surface of the material will be drawn out by the developer resulting in a bright red indication. Some Chap 7 Page 139 idea of the size of the defect may be obtained after experience by watching the size and rate of growth of the indication. 301. ULTRASONIC FLAW DETECTION. Ultrasonic flaw detection equipment has made it possible to locate defects in all types of materials without damaging the material being inspected. Very small cracks, checks, and voids, too small to be seen by X-ray, are located by means of ultrasonic inspection. An ultrasonic test instrument requires access to only one surface of the material to be inspected, and can be used with either straight line or angle beam testing techniques. The instrument electronically generates ultrasonic vibrations and sends them in a pulsed beam through the part to be tested. Any discontinuity within the part, or the opposite end, will reflect the vibration back to the instrument, which measures the elapsed time between the initial pulse and the return of all reflections and indicates such time lapse on a cathode ray indicator or paper recorder. Ultrasonic inspection requires a skilled operator who is familiar with the equipment being used as well as the inspection method to be used for the many different parts being tested. **302.-312. RESERVED.** Par 300 ### Page 20 (OCR) Page 140 AC 43.13-1A CHG 1 1975 Section 4. IDENTIFICATION OF FABRICS AND PLASTICS 313. IDENTIFICATION OF FABRIC. Cotton fabric is often used as covering for wing, fuselage, and control surfaces of aircraft. Acceptable grades of fabric for use on civil aircraft are listed in Chapter 3. In general, the fabric can be readily identified by a continuous marking to show the manufacturer's name or trademark and specification number. This marking may be found stamped along the selvage edge. The specification number for grade "A" fabric is AMS-3806, and for the intermediate grade AMS-3804. The corresponding FAA Technical Standard Order * Numbers for these materials are TSO-C15 and 1SO-C14, respectively. Increasing interest in* the use of linen and certain synthetic fabrics in lieu of cotton has been noted. Identity of such materials should always be verified by the user. 314. IDENTIFICATION OF PLASTICS, Plastics cover a broad field of organic synthetic resins and may be divided into two main classifications -thermoplastic and thermosetting plustics. a. Thermoplastics. Thermoplastics may be softened by heat and can be dissolved in various organic solvents. Two kinds of transparent thermoplastic materials are commonly employed in windows, canopies, etc. These materials are known as acrylic plastics and cellulose acetate plastics. These two plastics may be distinguished by the absence of color, the greater transparency, and the greater stiffness of the acrylic as compared to the slight yellow tint, lower transparency, and greater flexibility of cellulose acetate. b. Thermosetting Plastics. Thermosetting plastics do not soften appreciably under heat but may char and blister at temperatures of 204° to 260° C. (400° to 500° F.). Most of the molded products of synthetic resin composition, such as phenolic, urea-formaldehyde, and melamineformaldehyde resins, belong to the thermosetting group. **315.-320. RESERVED.** Par 313 Chap 7 ### Page 21 (OCR) AC 43.13-1A CHG 1 Page 140-1 1975 * Section 5. CABIN INTERIOR MATERIALS, FIRE PROTECTION QUALITIES 321. REQUIREMENTS FOR FIRE PROTECTION. The cleaning, repairing. and/or replacement of cabin interior materials necessitates continuing compliance with the fire protection requirements of the FARs. A recent field survey on this matter revealed a wide variance in the procedures used for maintaining cabin interior materials with respeot to the fire protection qualities of these materials. a. Interior Materials. The methods, techniques, practices, and materials used in the cleaning, repair, and replacement of cabin materials must meet the requirements of Section 43.13 of the FARS. b. Fire Protection. The requirements relative to fire protection qualities of cabin interior materials are specified in: (1) Section 121.312, for aircraft operated under Parts 121, 123, and 135. (2) Section 127.91, for helicopters used in passenger service under Part 127. c. Source of Information. If information regarding the original or properly altered fire protection qualities of certain cabin interior materials is not available, requests for this information should be made to the aircraft manufacturers or the local FAA regional office, specifying the model aircraft, the aircraft manufacturer, the date the aircraft was manufactured or the serial number, and the FAR Part under which the aircraft is operated (i.e., Part 91, Part 121, etc.). **322.-325. RESERVED.** Chap 7 Par 321 ### Page 22 (OCR) AC 43.13-1A CHG 1 1975 cycles. The dynamometer landings will consist of 50 Test A, load-speed-time cycles and 50 Test B, energy cycles. (a) Test A, Load-speed-time. 1. Speed cycle. Land the tire against a dynamometer flywheel rotating at a peripheral speed of S, m.p.h. Immediately thereafter, de. crease the flywheel speed at an uverage deceleration rate of D ft./sec./sec. until a value of Sz is attained. No specific rate of deceleration is required after the flywheel's peripheral speed reaches a value of S2. Decrease the speed of the flywheel in the ubove manner until a roll distance of RD feet has been covered, at which time the tire is unlanded. 2. Load cycle. After landing, increase the load from zero to L, pounds within 1, seconds. Linearly increase the load with time to a value of Is pounds within I; seconds after landing, or at the moment of unlanding, whichever occurs first. If it is necessary to continue the roll after T, seconds to complete the required roll distance (RD), maintain the load at La pounds. 3. Symbol deflnitions. Determine the numerical values, which are used for the following symbols, from the applicable airplane load-speed-time data: S, = Initial dynamometer test speed. Sz=Speed at which the average deceleration between S, and S, does not exceed the specified values. D=Constant rate of deceleration between Si and S2 speeds. RD= Roll distance in feet. L, = Initial tire load. L½=Maximum rated static load of the tire. T, =Time for applying I, load. A T, tolerance of +1 second is acceptable. S1- VS, -2D (RD) D T, is the elapsed time for applying the I load. A T½ tolerance of +10% is acceptable. When T, is calculated by the aforementioned formula, S, may be ignored and D is assumed constant throughout roll distance (RD). Chap 8 Page 149 4. Test load adjustment. If the test load curve results in loads at a given speed being less than those dictated by the applicable aircraft data, eliminate the condition by making adjustments in T., L1, and/or T, (b) Test B, Energy. 1. Kinetic energy. Calculate and adjust the kinetic energy of the flywheel for the rated maximum static load of the tire. In the event that the correct number of flywheel plates cannot be used to obtain the calculated kinetic energy value or proper flywheel width, select a greater number of plates and adjust the dynamometer speed to achieve the required kinetic energy. 2. Kinetic energy computation. Compute kinetic energy as follows: KE=OWV: Where KE = Kinetic energy, ft.-1b. C=0.011 W=tire load, pounds V = 120 m.p.h. 3. Speed cycle. Land the test tire at 90 m.p.h. and unland at 0 m.p.h. Decrease the landing speed as necessary to assure that 56% of the calculated kinetic energy is absorbed by the tire. 4. Load cycle. Upon landing, and during the entire roll test, force the tire against the flywheel at its rated static load. (c) Taxi Test. 1. Test parameters. Conduct a minimum of three dynamometer taxi tests under the following conditions: Speed =35 m.p.h. Tire Load = Maximum static rating Roll Distance =35,000 feet. 2. Tire temperature. Heat the test tire to a temperature of not less than 120° F. at the start of each of the three taxi test cycles. Rolling the tire on the dynamometer is acceptable in obtaining this minimum tire temperature. Make no adjustments in the tire inflation pressure to compensate for increases due to temperature rise. (9) Alternate Dynamometer Tests. (a) Variable Loading. An alternate dynamometer test which more realistically simulates Par 332 ### Page 23 (OCR) actual airplane performance on the runway may be used in lien of the deceleration load-speedtime schedule. An acceleration load-speed-time schedule, wherein the dynamometer flywheel is accelerated to the applicable conditions, is accept- (b) Alternate Procedure for Reinforced-tread Tires. Qualification of a high-speed tire with a given ply rating and reinforced tread will automatically qualify a lesser ply rating reinforced tread tire of the same size and skid depth, provided : 1. The test conditions S1, RD, S2, I. and T, are no less severe than those which are applicable to the lesser ply rating tire. 2. The ratio of the test loads, Ly to I is not less than that applicable to the lesser ply rating tire. Make any necessary adjustment in this ratio by increasing L.. (10) Optional Dynamometer Equipment. Dynamic tests may be conducted on any dynanometer test equipment which will provide the load, speed, time, and roll distance parameters of the tire. c. Tubes. Punctured tubes may be repaired by the use of cemented or vulcanized patches. (1) The procedure for making such repairs is substantially identical to that used in connection with repair of automobile tires: (a) Keep the size of the patch to a minimum and avoid use of an excessive number of patches, particularly in one area, as the weight of the material may contribute to excessive wheel vibration due to the tube being out-of-balance. (b) The use of vulcanized patches is recommended because they are considered more reliable. (c) Reinstalled tires should be inflated, deflated, and again inflated to insure that the inner- Par 362 AC 43.13-1A CHG 1 1975 tube is not pinched. A pinched tube will chafe against the walls of the carcass and a thin spot will result in the rubber. In time, the tube wall will leak at this point. The pinching generally is due to the sticking of the tube to the carcass wall during the first inflation and the failure of the carcass to properly seat against the flange. (d) The tube is then confined to a smaller space and wrinkling (pinching) of the tube results. Complete deflation followed by inflation allows the tube to properly accommodate itself to the carcass which should now seat itself tightly against the flanges. *333. TIRE SUIPPAGE. To reduce the possibility of tire and tube failure due to slippage, and to provide a means of detecting tire slippage, tires should be marked and indexed with the wheel rim. Marking should be accomplished by painting a mark 1 inch in width and 2 inches in length across the tire sidewall and wheel rim. The paint used should be of a permanent type and contrasting color, such as white, red, or orange. Preflight inspection should include a check of slippage marks for alignment. If the slippage mark is not in alignment, the aircraft should not be operated until a detailed inspection is made. the reason determined, and if necessary the condition corrected. 334. TIRE MAINTENANCE. A program of tire maintenance can minimize tire failures and increase tire service life. Overinflation wears the center of the tread excessively, and reduces a tire's resistance to bruising, strains the tires beads, reduces traction and skid resistance. Underinflation increases deflection and may cause breakdown of the tire sidewalls. The manufacturer's recommendations should he followed to obtain maximum tire service life. 345.-349. RESERVED. * Chap 8 ### Page 24 (OCR) AC 43.13-1A CHG 1 .. Page 165 1975 Chapter 10. HYDRAULIC AND PNEUMATIC SYSTEMS Section 1. HYDRAULIC SYSTEMS 392. GENERAL. Maintain, service, and adjust connection fittings: AN-818 nut and AN-819 airplane hydraulic systems in accordance with sleeve. manufacturers' maintenance manuals and pergeneral principles of maintenance and repair tinent component maintenance manuals. Certain which apply are outlined below. Tube O.D. for tightening tube nuts Wrench torque rauge (inch pounds) Minimum bend radil * Dimension in inches. * measured at inalde of bend. (Inches) 393. HYDRAULIC LINES AND FITTINGS. Carefully 1100-H14, Alum alloy Alum. alloy 1100-114, inspect all lines and fittings at regular intervals 5052-0 Steel 5052-0 Steel to insure airworthiness. Investigate any evidence of fuid loss or leaks. Check metal lines for 80-70 leaks, loose anchorages, scratches, kinks, or other 40-05 60-30 50-90 70-120 2132 damage, Inspect fittings and connections for 75-125 90-150 leakage, looseness, cracks, burrs, or other damage. 150-250 155-250 1¾G Replace or repair defective elements. 200-350 300-400 1¾ 300-500 430-575 210 a. Replacement of Metal Lines. When inspec 500-700 550-750 28 tion shows a line to be damaged or defective, 600-900 replace the entire line or if the damaged section 600-900 3¾ is localized, a repair section may be inserted. 5 7 In replacing lines, always use tubing of the same 8 7 size and material as the original line. Use the old tubing as a template in bending the new line, FIGURE 10.1.-Tube data. unless it is too greatly damaged, in which case a template can be made from soft iron wire. In forming fares, cut the tube ends square, file Soft aluminum tubing (1100, 3003, or 5052) smooth, remove all burrs and sharp edges, and under ¼-inch outside dinmeter muy be bent by thoroughly clean. The tubing is then flared hand. For all other tubing usa an acceptable using the correct 37° aviation flare forming tool hand or power tube bending tool. Bend tubing for the size of tubing and type of fitting. A carefully to avoid excessive flattening, kinking, double flare is used on soft aluminum tubing or wrinkling. Minimum bend radii values are ⅛-inch outside diameter and under, and a single shown in figure 10.1. A small amount of flatten- flare on all other tubing. In making the connecing in bends is acceptable but do not exceed an tions, use hydraulic fluid as a lubricant and then amount such that the small diameter of the flat- tighten. Overtightening will damage the tube tened portion is less than 75 percent of the origi- or fitting, which may cause a failure; undernal outside diameter. When installing the tightening may cause leakage which could result replacement tubing. line it up correctly with the in a system failure. mating part so that it is not forced into line by means of the coupling nuts. CAUTION Mistaken use of 45° automotive fare forming tools will b. Tube Connections. Many tubing connections result In improper tubing flare shape and angle causing are made using fured tube ends, and standard misfit, stress and strain, and probable system fallure. Chap 10 Par 392 ### Page 25 (OCR) Page 166 AC 43.13-1A CHG 1 1975 Any excessive If the of connection fittings. fittings. e. O-Ring Seals. A thorough understanding of O-ring seal applications is necessary to determine when replacement must be made. The simplest application is where the O-ring merely serves as a gasket when it is compressed within a recessed area by applying pressure with a packing nut or screw cap. Leakage is not normally acceptable in this type of installation. In other installations the O- ring seals depend primarily upon their resiliency to accomplish their sealing action. When moving parts are involved, minor seepage may be normal and acceptable. A moist surface found on moving parts of hydraulic units is an indication the seal is being properly lubricated. In pneumatic systems, seal lubrication is provided by the installation of a grease-impreg- Par 393 * U.S.. GOVERNMENT PRINTING OFFICE: 1975-631-513/928 3-1 Chap TO ### Page 26 (OCR) CHANGE AC NO: 48.18-1A CHG 1 DATE: 5/12/75 **NONENT DE TRANSO ADVISORY** **TED STATES OF ARE CIRCULAR** **DEPARTMENT OF TRANSPORTATION** **FEDERAL AVIATION ADMINISTRATION** ACCEPTABLE METHODS, TECHNIQUES, AND SUBJECT: PRACTICES-AIRCRAFT INSPECTION AND RE- PAIR **1. PURPOSE.** This change transmits new and revised material for Advisory Circular 43.13-1A. **2. PRINCIPAL CHANGES.** a. Correotions have been made to paragraph 818 and Figure 10.1. b. Chapter 2, Section 3, has been revised to add Selective Plating in aircraft maintenance. Chapter 4, Section 6 has been revised to add handling and care of aircraft recovered from water immersion. c. A new Section 5, Cabin Interior Materiale Fire Protection Qualities, has been added to Chapter ?. d. New information on tire slippage and maintenance has been added to Section 1 of Chapter 8. **3. HOW TO GET THIS PUBLICATION.** Copies of this change to AC 13.13-1A may be obtained by ordering from: Superintendent of Documents U.S. Government Printing Office Washington, D.C. 20402 For up-to-date information on how to order, and the price of this publication, check with the closest FAA office or obtain the latest copy of Advisory Circular 00-2, Advisory Circular Checklist and Status of the Federal Aviation Regulations. This checklist may be obtained without charge by requesting from: Department of Transportation Publications Section, TAD 443.1 Washington, D.C. 20590 Initlated by: AFS-B30 PoP Sally R. P. SKULLY, Director Flight Standards Service Initiated by: ### Page 27 (OCR) PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated i thru iv 1972 1972 ii thru iv 1975 78 and 74 1972 73 thru 73-3 1975 1972 127 and 128 1972 1972 128 thru 128-3 1975 139 and 140 1972 1972 140 thru 140-1 • 1975 149 and 150 1972 1972 1975 165 and 166 1972 1975 166 1972 ### Page 28 (OCR) AC 43.13-1A Page I 1972 TABLE OF CONTENTS Chapter 1. AIRCRAFT WOOD STRUCTURES Section 1. MATERIALS AND PRACTICES Paragraph Pape Materials Replacement of Drainholes and Skin Stiffeners 3. Control Surface Flutter Precautions 4. Gluing Precautions Б. Preparation of Wood Surfaces for Gluing .. 6. Glues 7. Mixing of Resin Glues 8. Spreading of Glue 9. Assembly Time in Gluing 10. Gluing Pressure 11. Scarf Joints 12. Splicing of Spars 13. Splicing of Box Spar Webs 14. Replacing Solid-Type Spars with Laminated Type 15. Spar Longitudinal Cracks and Local Damage 16. Elongated Holes in Spars 17. Rib Repairs 18. Plywood Skin Repair 19. Repaira to Single Curvature Plywood Skin 10 20. Repairs to Double Curvature Plywood Skin 10 21. Splayed Patch 10 22. Surface Patch 16 28. Scarf Patch 24. Plug Patches 17 25. Fabric Patch 28 26.-86. Reserved 28 Section 2. FINISHING WOOD STRUCTURES 87. General 24 88. Precautions to be Observed 89. Finiching of Interior Surfaces 40. Finishing of Exterior Surfaces 41. Finishing of End Grain Surfaces 42. Finishing with Fabric or Tape 43.-53. Reserved Chapter 2. AIRCRAFT METAL STRUCTURES Section 1. REFERENCES AND PRECAUTIONARY MEASURES 64. References 55. Flutter and Vibration Precautions 56. Brazing 57.-67. Reserved Section 2. WELDING 68. General 69. Steel Parts Not to be Welded 70. Repair of Tubular Membera ### Page 29 (OCR) Page I1 AC 43.13-1A CHG 1 1975 TABLE OF CONTENTS-Continued Paragraph Page 71. Repair by Welded Sleeve 35 72. Repalr by Bolted Sleeve 85 73. Welded-Patch Repair 35 74. Splicing Tubing by Inner Sleeve Method 36 75. Splicing Tubing by Outer Sleeve Method 38 76. Splicing Using Larger Diameter Replacement Tubes 89 77. Repalrs at Built-In Fuselage Fittings .. 40 78. Engine Mount Repairs 40 79. Landing Gear Repairs 40 80. Built-Up Tubular Wing or Tall Surface Spars 47 81. Wing and Tall Brace Struts 47 82. Repairs to Welded Parts 47 83. Stainless Steel Structure 47 84.-04. Reserved 18 Section 3. METAL REPAIR PROCEDURES 95. Riveted or Bolted Steel Truas-Type Structureg 49 96. Aluminum Alloy Structures 97. Selection of Aluminum for Replacement Parts 98. Heat Treatment of Aluminum Alloy Parts 99. Riveting 51 100. Repnir Methods and Precautions for Aluminum Structure 54 101. Repairing Cracked Members 68 102. Steel and Aluminum Fittings 66 103. Castings 79 * 104. Selective Plating In Alrcraft Maintenance 73* 105.-114. Reserved 73-3 Section 4. REPAIR OF LAMINATE STRUCTURES 115. General 74 116.-126, Reserved 79 Chapter 3. FABRIC COVERING Section 1. PRACTICES AND PRECAUTIONS 127. Textile Materlals 128. Covering Practices 129. Preparation of the Structure for Covering 180. Fabric Seams - 131. Covering Methods 182. Reinforcing Tape 133. Lacing 134. Stitch Spneing so bo to 185. Surface Tape (Finishing Tape) - 180. Special Fasteners 137.-147. Reserved 91 Section 2. DOPING 148. Application of Alreraft Dope, Epoxy, and Resina 149, Blushing and Use of Blush-Retarding Thinner 150. Number of Coats 151. Technique -- 152. Rejuvenation of Rahrie 158. Common Dope Troubles 151-164. Reserved ### Page 30 (OCR) AC 43.13-1A CHG 1 Page Ill 1975 TABLE OF CONTENTS-Continued Paragraph Page Section 3. REPAIRS TO FABRIC COVERING 165. General 95 166. Repair of Tears in Fabric 95 167. Sewed Patch Repalr -- 168. Unsewed (Doped-On) Hepairs - 90 169. Repair by a Doped-In Panel 98 170.-180. Reserved 87 Section 4. FABRIC TESTING 181. Testing of Fabric Covering - 98 182.-192. Reserved 08 Chapter 4. CONTROL CABLES AND TERMINALS Section 1. INSPECTION AND REPAIR 193. General 104. Cable Specifcations 195. Replacement of Cables 196. Mechanically Fabricated Cable Assemblies 107. Hand Fabricated Cable Assemblles 108. Cable System Inspection 107 190. Cable Tension Adjustment 110 200. Corrosion and Rust Prevention 110 201.-211. Reserved 111 Sechion 2. SAFETY METHODS FOR TURNBUCKLES 212. General 218. Double Wrap Method 214. Single Wrap Method 114 215. Special Locking Devices 114 216.-220. Reserved 114 Chapter 5. AIRCRAFT HARDWARE Section 1. IDENTIFICATION AND USE OF AIRCRAFT HARDWARE Bolte 115 Screws 117 Taper Ping (AN-385 and AN-880) 118 Nuts 119 Washers - 120 282. Types of Rivets 121 233. Material Applications 121 284. Fasteners (Cowl and Falring) 122 235. Unconventional Attachments 122 Reserved 122 Chapter 6. CORROSION PROTECTION 247. Types of Corrosion -. 123 248. Corrosion Protection Measures for Basic Materials 124 249. Construction of Corrosion Resistant Structure 126 250. Corrosion Removal - 126 251. Corroslon Proofing of Landplanes and Seaplanes 127 252. Corrosion Proofing of Landplanes Converted to Seaplanes 127 258. Cleaners, Polishes, Brighteners 128 *255-284. Regerved 254. Handling and Care of Alrcraft Recovered from Water Immersion 128-3 128* ### Page 31 (OCR) Page iv AC 43.13-1A CHG 1 1975 TABLE OF CONTENTS-Continued Chapter 7. IDENTIFICATION, TESTING, AND INSPECTION OF MATERIALS Paragraph Page Section 1. IDENTIFICATION OF METALS 205. Identificatlon of Steel Stock - 129 266. Interchangeability of Steel Tubing 129 267. Identification of Aluminum --- 129 268.-278. Reserved 181 Section 2. TESTING OF METALS 279. Hardness Testing 182 Brinell Hardness Test 182 281. Rockwell Hardness Test 182 Vickers Hardness Test 182 283. Testing Aluminum 182 284.-294. Reserved 182 Section 3. NONDESTRUCTIVE TESTING 205. General 186 298. Inspection by Magnifying Glass After Welding 136 297. Magnetic Particle Inspection 186 298. X-Ray or Radiographic Inspection 187 209. Fluorescent Penetrant 138 300. Dye Penetrants 138 801. Ultrasonic Flaw Detection 139 802.-312. . Reserved 189 Section 4. IDENTIFICATION OF FABRICS AND PLASTICS 318. Identification of Fabric 140 * 314. Identification of Plastics 140 815.-320. Reserved 140 * Section 5. CABIN INTERIOR MATERIALS, FIRE PROTECTION QUALITIES 321. Requirements for Fire Protection 140-1 322,-325. Reserved 140-1* Chapter 8. AIRCRAFT EQUIPMENT Section 1. LANDING GEAR EQUIPMENT 826. General 141 827. Inspection 141 328. Clenning and Lubricating 145 329. Emergency Systems 145 830. Speclal Inspections 145 331. Retraction Tests 145 332. Tires 145 * 333. Tire Slippage 334. Tire Maintenance 335.-343. Reserved 150 Section 2. LIFE PRESERVERS AND LIFERAFTS 844. General 151 345. Inspection Procedure for Life Preservers 151 346. Repair of Hife Preservers 151 397. Inspection Procedure for Liferafts. 151 348. Repair of Liferafts - 152 349.-359. Reserved 152 ### Page 32 (OCR) AC 43.13-1A CHG 1 Page iv-I 1975 TABLE OF CONTENTS-Continued Paragraph Page Section 3. MISCELLANEOUS EQUIPMENT 360. Parachutes 153 881. Safety Belts 158 862. Flares 154 363. Oxygen Systems 154 361.-874. Reserved 158 ### Page 33 (OCR) C 43.13-1A Page v 1972 TABLE OF CONTENTS-Continued Chapter 9. WINDSHIELDS, ENCLOSURES, AND EXITS Section 1. PLASTIC WINDSHIELDS AND ENCLOSURES Paragraph Page 975. General 159 376. Replacement Panels 159 877. Installation Procedures 159 378. Repair of Plastics 160 879. Cleaning and Polishing Transparent Plastic 161 980. Emergency Exits 161 881.-891. Reserved 164 Chapter 10. HYDRAULIC AND PNEUMATIC SYSTEMS Section 1. HYDRAULIC SYSTEMS 392. General 165 394. Hydraulic Components 393. Hydraulic Lines and Fittings 165 172 395.-405. Reserved 172 Chapter 11. ELECTRICAL SYSTEMS Section 1. CARE OF ELECTRICAL SYSTEMS 406. General 178 407. Inspection and Operation Checka . 178 408. Cleaning and Preservation 178 409. Adjustment 174 410. Insulation of Electrical Equipment . 174 411. Bus Bar Maintenance 174 412. Junction Boxes 174 418.-423, Reserved 174 Section 2. EQUIPMENT INSTALLATION 424. Electrical Load Limits 176 425. Generators 175 426. Acceptable Means of Controlling or Monitoring the Electrical Load 175 427. Inverters, Alternators, and Similar Aircraft Electrical Power Sources 176 428. Determination of Electrical Load - 178 429. Circuit Protection Devices 176 430. Switches 176 481.-441. Reserved 178 Section 3. ELECTRIC WIRE 442. General 179 448. Aircraft Electric Wire 180 444. Instructions for Use of Electric Wire Chart 181 445. Splices in Electric Wire 186 446. Open Wiring 186 447. Heat Precautions 187 448. Protection Against Chafing 187 449. Stripping Insulation 187 450. Terminals 188 451. Attachment of Terminals to Studs 188 452. Bonding Jumper Installations 188 458.-463. Reserved 198 ### Page 34 (OCR) Pope vi AC 43.13- TABLE OF CONTENTS-Continued Section 4. WIRE MARKING Paragraph Page 464. Wire Identification 194 465. Sleeve Selection for Identification 194 466. Identification of Wire Bundles and Harnesses 194 467-477. Reserved 195 Section 5. CONNECTORS 478. General Purpose Connectors 196 479. Types of Connectors 198 480. AN Connector Identification 196 481. Special Purpose Connector 198 482. Potting Compounds 198 483. Potting Connectors 199 484. Through Bolts 200 485. Terminal Strips 200 486.-496. Reserved 200 Section 6. CONDUITS 497. General 201 498. Size of Conduit 201 499. Conduit Fittings 201 500. Conduit Installation 201 501. Rigid Conduit 201 502. Flexible Conduit 201 508.-518. Reserved 202 Section 7. ROUTING, TYING, LACING, AND CLAMPING 614. General 208 515. Wire Bend Radii 208 516. Slack 208 517. Routing Wires 208 518. Clamping 205 619. Ties and Lacing 205 520.-580. Reserved 205 Section 8. STORAGE BATTERIES 581. General 210 582. Battery Overheat Failure 210 588. Battery Freezing 210 584. Temperature Correction 210 685. Checking Lead-Acid Batteries 210 586. Checking Nickel-Cadmium Batteries 211 587. Electrolyte Corrosion 211 588. Noxious Fumes 211 589. Installation Practices 211 540.-552. Reserved 212 Chapter 12. PROPELLERS, ROTORS, AND ASSOCIATED EQUIPMENT Section 1. INSPECTION OF PROPELLERS 568. Wood or Composition Propellers and Blades 218 554. Metal Propellers and Blades 214 555. Lubrication 214 556.-566. Reserved 214 ### Page 35 (OCR) : 43.13-1A Page vil 72 TABLE OF CONTENTS- Continued Section 2. REPAIR OF WOOD AND COMPOSITION PROPELLERS Paragraph Page 567. General 215 668. Repair of Wood Propellers 215 569. Repair of Composition Blade 220 570. Refinishing 222 571. Propeller Balancing 222 572.-582. Reserved 222 Section 3. REPAIR OF METAL PROPELLERS 228 Steel Blades 228 585. Aluminum Propeller Repaire 228 Repair Tolerances 229 Propeller Balance 228 588. Steel Hubs and Hub Parts 280 Propeller Hub and Flange Repair 280 590. Control Systema 280 591. Deicing Systems 280 592.-602. Reserved 280 Section 4. OVERHAUL OF PROPELLERS 608. Hub and Hub Parts 282 604. Plating for Hubs and Parts 282 605. Inspection of Aluminum Propellers and Blades 282 606. Assembly 288 607.-617. Reserved 288 Section 5. ASSEMBLY OF PROPELLER TO ENGINE 618. Fixed-Pitch Propellers 284 619.-629. Reserved 284 Section 6. PRECAUTIONS 680. Blade-Tip Identification 285 631. Ground-Handling Practices 285 682. Composition Propeller Service Limits 285 688.-648. Reserved 285 Section 7. MAINTENANCE OF ROTOR BLADES 644. General 286 645.-655. Reserved 241 Chapter 13. WEIGHT AND BALANCE 656. General 248 657. Terminology 248 658. Weighing Procedure 250 659. Weight and Balance Computations 258 660. Weight and Balance Extreme Conditions 258 661. Loading Conditions and/or Placards 257 662. Equipment List 257 668. Sample Weight and Balance Report 257 664. Installation of Ballast 258 665. Loading Schedule 265 666.-676. Reserved 265 ### Page 36 (OCR) Page vill AC 43.13- 19 TABLE OF CONTENTS- Continued Chapter 14. ENGINES AND FUEL SYSTEMS Section 1. ENGINES Paragraph Page 677. General 267 678. inspection 267 679. Powerplant Sudden Stoppage - 267 680. Crankshafts - 269 681. Replacement Parts in Certificated Engines 289 682. Cylinder Holddown Nuts and Capscrews 270 688. Reuse of Safetying Devices 270 684. Self-Locking Nuts for Aircraft Engines and Accessories 270 685. Welding of Highly Stressed Engine Parts 271 686. Welding of Minor Engine Parte 271 687. Metallizing 271 688. Plating 271 689. Engine Accessories 272 690. Corrosion 272 691. Engine Run-In 272 692. Compression Testing of Aircraft Engine Cylinders 212 698. Spark Plugs 274 694. Ignition Harnesses 278 695. Magneto Inspection 278 696. Magneto-to-Engine Timing 280 697-707. Reserved 280 Section 2. FUEL SYSTEMS 708. General 281 709. Fuel Lines and Fittings 281 710. Fuel Tanks and Cella 282 711. Fuel Tank Caps, Vents, and Overflow Lines 282 712. Fuel Crossfeed, Firewall Shutoff, and Tank Selector Valves 288 718. Fuel Pumps 288 714. Fuel Filters, Strainers, and Drains 288 716. Indicator Systems 288 716. Turbine Fuel Systems 288 717.-727. Reserved 284 Section 3. EXHAUST SYSTEMS 728. General 729. Muffer/Heat Exchanger Failures 780. Manifold/Stack Failures 285 781. Internal Muffler Failures 285 782- Inspection 288 788. Repairs 288 784. Turbosupercharger 290 785. Augmentor Systems 290 786.-746. Reserved 290 Chapter 15. RADIO AND ELECTRONIC SYSTEMS Section 1. MAINTENANCE OF RADIO AND ELECTRONIC SYSTEMS 747. General 291 748. Maintenance of Equipment 291 749. Replacement of Components 291 760. Inspection of Installations 291 ### Page 37 (OCR) AC 43,13-14 Page ix 1972 TABLE OF CONTENTS- Continued Paragraph Page 751. Cleaning 291 752. Contact Cleaners 291 758. Operational Checks 292 754.-765. Reserved 292 Section 2. MAINTENANCE OF WIRING AND CABLES 766. Wiring and Cable Integrity 767- Protection from Fluids 768. Wire and Cable Clamps 769. Conduit 770. Plastic Wire Sleeving 771. Electrical Tape 772.-782. Reserved Section 3. JUNCTION BOXES AND TERMINAL STRIPS 788. Junction Boxes 784. Internal Arrangement 785. Terminal Strips 786. Pluge and Receptacles 787.-797. Reserved Section 4. POWER SOURCES AND DISTRIBUTION SYSTEMS 798. Battery Electrolyte Corrosion 285 799. Generators and Alternators 295 800. Alternator Diodes 295 801. Diode Protection 295 802. Alternator/Battery Connections 295 808. Static Electrical Power Convertera 295 804. Cleaning and Preservation 295 805. Miscellaneous Check Items 295 806. Adjustment and Repair 296 807. Electrical Switch Inspection 296 808. Radio Equipment Operation 296 809.-819. Reserved 296 Section 5. NOISE SUPPRESSION 820. General 297 821. Bonding 297 822. Shielding 297 828. Spark Plugs 297 824. Filters 297 825. Precipitation Static 297 826. Static Dischargers 297 827. Maintenance of Static Dischargers 297 828.-888. Reserved 298 Section 6. ANTENNAS 889. Fixed and Trailing Wire Antennas 299 840. Whip and Blade Antennas 299 841. Flush Mounted Antennas 299 842. Coaxial Cables and Fittings 299 843. Radomes 299 844. Radome Inspection 299 845. Radome Maintenance and Repair 800 846. Loop Antennas 800 847. Loop Quadrantal Error 800 ### Page 38 (OCR) Page x AC 43.13 1A 1972 TABLE OF CONTENTS Continued Paragraph Pane 848. Compensation for Quadrantal Error 800 849. Ground Calibration 801 850. Air Calibration 801 851. Calibration Points 801 852. Calibration Procedure 801 858. Calibration Data 801 854.-868. Reserved 801 Chapter 16. INSTRUMENTS Section 1. MAINTENANCE OF INSTRUMENTS General 808 865. Definition 808 866. Maintenance of Instruments 808 867. Test/Adjustment of Instruments 808 868. Replacement of Components 804 869.-879. Reserved 804 Section 2. INSPECTION AND CHECKS 880. General 805 881. Periodic Inapection 805 882. Additional Inspectiona 805 888.-899. Reserved 805 Section 3. INSTRUMENT POWER SOURCES 894. Electrical Compatibility 806 895. Vacuum Systems 806 896.-906. Reserved 806 Section 4. PITOT-STATIC SYSTEMS 907. System Components 807 908. Pitot-Static Tubes and Lines 807 909. Pressure Porta or Vents 807 910. Cleaning of System 807 911. Heater Elements 807 912. System Leak Testing 807 918. Static System Test 807 914. Pitot System Test 808 915. Precautions in Teating 308 916.-926. Reserved 808 Section 5. MECHANICAL ADJUSTMENT OF MAGNETIC DIRECTION **INDICATORS** 927. Correction for Errors 809 928. Ground Swinging 809 929. Remote Gyro Compass Systems 809 980.-940. Reserved 809 Section 6. PRESERVATION AND PACKAGING 941. Preservation 810 942. Storage 810 943. Shipping 810 924-954. Reserved 810 Appendix 1. Cross Reference Table of Paragraph Number Changes 811 Appendix 2. Cross Reference Table of Illustration Number Changes 815 ### Page 39 (OCR) C 43,13-1A Page xi 972 LIST OF ILLUSTRATIONS Pigure Page 1.1 Tapering of Face Plates .. 1 1.2 Selection and Properties of Aircraft Wood 1.8 Consideration of Grain Direction When Making Scarf Joints 6 Method of Splicing Solid or Laminated Rectangular Spars 1.5 Method of Splicing Solid "I" Spars -- 1.6 Repaira to Bullt-Up "I' Spar .. 9 1,7 Method of Splicing Box-Spar Flanges (Plate Method) 10 1.8 1.9 Method of Splicing Box-Spar Webs Method of Reinforcing a LongitudInal Crack and/or Local Damage 11 in a Solld or Internally Routed Spar .. 12 1.10 Repair of Wood Ribs 18 1.11 Typical Wing Compression Rib Repairs 14 1.12 Minimum Recommended Bend Radii for Aircraft Plywood 15 1.18 Splayed Patch 16 1.1₫ Surface Patches 18 1.15 Scarf Patches Back of Skin Accessible 19 1.16 Scarf Patches-Back of Skin Not Accessible 20 1.17 Oval Plug Patch Assembly 21 1.18 Round Plug Patch Assembly 22 2.1 Current and Polarity Selection for Inert Gas Welding 29 2.2 Specifications Guide for Welding 81 2.8 Strengths of Welded Joints 82 2,4 Members Dented at a Cluster 88 2.5 Members Dented in a Bay-Repairs by Welded Sleeve 84 2,6 Welded Patch Repair 85 2.7 Splicing by Inner Sleeve Method 86 2.8 Splicing by Outer Sleeve Method-Replacement by Welded Outside Sleeve 87 2.9 Tube Replacement at a Station by Welded Outer Sleeves 88 2.10 Splicing Using Larger Diameter Replacement Tube 89 2.11 Repairs at Built-In Fuselage Fittings 41 2.12 Streamline Tube Splice Using Round Tube (Applicable to Landing Gears) 42 2.18 Streamline Tube Splice Using Split Sleeve (Applicable to Wing and Tail Surface Brace Struts and Other Members) 48 2.14 Streamline Tube Splice Using Split Insert (Applicable to Landing Gears) 44 2.15 Streamline Tube Splice Using Plates (Applicable to Landing Gears) 45 2.16 Representative Types of Repairable and Nonrepairable Axle Assemblies 46 2.17 Recommended Radii for 90° Bends in Aluminum Alloys 50 2.18 Rivet Hole Spacing and Edge Distance for Single-Lap Sheet Splices 52 2.19 Riveting Practice and Rivet Imperfections 58 2.20 Typical Repair Method for Tubular Members of Aluminum Alloy 55 2.21 Typical Repair for Buckled or Cracked Formed Metal Wing Rib Capstripa 56 Typical Metal Rib Repaire (Usually Found on Small and Medium-Size Aircraft) 57 2.28 Typical Repairs of Trailing Edges .. 58 2-24 Typical Repairs of Stressed Sheet Metal Covering 59 2.25 Typical Stringer and Flange Splices 61 2.26 Example of Stringer Splice (Material-2017 Alloy) ₴2 ### Page 40 (OCR) Page xil AC 43.13-1 197 LIST OF ILLUSTRATIONS- Continued Figure Page 2.27 Application of Typical Flange Splices and Reinforcement 68 2.28 Number of Rivets Required for Splices (Single-Lap Joint) in Bare 2014-T6, 8024-18, 2024-T86, and 7075-T6 Sheet, Clad 2014-T6, 2024-T8, 2024-T96 and 7075-T6 Sheet, 2024-T4, and 7075-T6 Plate, Bar, Rod, Tube, and Extrusions, 2014-T6 Extrusions 64 2.29 Number of Rivete Required for Splices (Single-Lap Joint) in 2017, 2017 Alclad, 2024-T86 and 2024-T86 Alclad Sheet, Plate, Bar, Rod, Tube, and Extrusions - 65 2.80 Number of Rivets Required for Splices (Single-Lap Joint) in 5052 (All Hardnesses) Sheet 66 2.81 Example of Splice of Intermediate Frame (Material 2017 AL Alloy) 67 2.82 Typical Methods of Repalring Cracked Leading and Trailing Edges and Rib Intersections - 68 2.88 Typical Methods of Replacing Cracked Members at Fittinga 69 2.84 Typical Methods of Repairing Cracked Frame and Stiffener Combinations 70 2.85 Typical Repairs to Rudder and to Fuselage at Tail Post 71 2.86 Typical Methods of Repairing Elongated or Worn Bolt Holes 72 2.87 Typical Laminate (Facing) Repair 74 2,88 Typical Core and Facing Repair - 75 2.89 Typical Stepped Joint Repair 75 2.40 Typical Scarfed Joint Repair 78 2.41 Typical One Side Blind Repair 76 2.42 Typical Metal Facing Patch 2.48 Typical Types of Core Replacement 78 2.44 Typical Undercut Core Material Cavity 78 8.1 Textile Fabric Used in Aircraft Covering 88 3.2 Miscellaneous Textile Materials 84 8.8 Typical Methods of Attaching Fabric 85 3.4 Standard Knot for Rib Lacing and Terminating a Sewed Seam (Modified Seine Knot) 86 8.6 Splice Knot 87 3.6 Starting Stitch for Rib Stitching 88 8.7 Standard Double Loop Lacing 89 8.8 Standard Knot for Double Loop Lacing 90 8.9 Fabric Attachment Spacing 91 3.10 Repair of Tears in Fabric 95 4.1 Cable Cross Section 99 4.2 Strength of Steel Cable 100 4.8 Straight-Shank Terminal Dimensiona 101 4.4 Insertion of Cable into Terminal 101 4.5 Gauging Terminal Shank After Swaging 101 4.6 Typical Cable Splices 102 4.7 Typical Terminal Gauge 102 4.8 Copper Oval Sleeve Data 108 4.9 Typical Thimble-Eye Splice 108 4.10 Copper Stop Sleeve Data 104 4.11 Typical Terminal Gauge 104 4.12 Preparation of a Woven Cable Splice 106 4.18 Preparation of a Wrapped Soldered Terminal 108 4.14 Cable Inspection Technique 107 4.15 Cable Wear Patterns .- 107 4.16 Worn Cable (Replacement Necessary) 108 4.17 Worn Cable (Replacement Recommended) 108 4.18 Internal Cable Wear 108 4.19 Corrosion 109 4.20 Manufacturer's Wire Splice 109 4.21 Pulley Wear Patterns 110 ### Page 41 (OCR) AC 43.13-1A Page xill 1972 LIST OF ILLUSTRATIONS- Continued Figure Page 4,22 Turnbuckle Safetying Guide 112 4.28 Safetying Turnbuckles 118 4.24 Clip Type Locking Device 114 5.1 Bolt Indentification 116 5.2 Recommended Torque Values For Nut-Bolt Combinations-(without lubrication) . 118 5.8 Torque Wrench With Various Adapters 119 7.1 Numerical System for Steel Identification 129 7.2 Basic Temper Designationa and Subdivisions for Aluminum Alloys 180 7.8 Rockwell C Scale Steel Hardneas Numbers Comparison Table 188 7,4 Hardness Values for Aluminum Alloys 182 7.5 Maximum Recommended X-Ray Kilovolts 188 8.1 Typical Bolt Cracks 141 8.2 Typical Cracke Near Boltholes 142 8.8 Typical Bolthole Cracka 142 8.4 Typical Rod-End Cracks 142 8.5 Typical Torque Tube Bolthole Elongation 142 8,6 Aircraft Tire Types and Ratings 146 8.7 Oxygen Cylinder Damage 155 8.8 Cylinder Brackets and Clamps 156 9.1 Stop-Drilling Cracks 160 9.2 Surface Patches 160 9.8 Plug Patch Repair 161 10.1 Tube Data 165 10.2 Hose Assembly Instructions 167 10.8 Aircraft Hose Specificationa 168 10.4 Proper Hose Installations 169 10.5 Minimum Bend Radif 170 10.6 Ball Diameters for Testing Hose Restrictions or Kinking 171 10.7 Suggested Handling of Preformed Hose 171 10.8 Minimum Bend Radil-Teflon Hose 172 11.1 Wire and Circuit Protector Chart 177 11.2 Switch Derating Factors 177 11.8 Examples Determining Voltage Drop from Electric Wire Chart (Fig. 11.7) 179 11.4 Examples Determining Voltage Drop from Electric Wire Chart (Fig. 11.7) 179 11.5 Copper Electric Wire Current Carrying Capacity 180 11.8 Aluminum Electric Wire Current Carrying Capacity 180 11.7 Electric Wire Chart 182 11.78 Wire Used in Aircraft Installations 188 11.8 Staggered Splices in Wire Bundle 186 11.9 Cable Clamp at Bulkhead Hole 187 11.10 Cable Clamp and Grommet at Bulkhead Hole 187 11.11 Allowable Nicked or Broken Strands .. 188 11.12 Stud Bonding or Grounding to Flat Surface 189 11.18 Plate Nut Bonding or Grounding to Flat Surface 190 11.14 Bolt and Nut Bonding or Grounding to Flat Surface 191 11.15 Copper Jumper Connection to Tubular Structure 192 11.16 Bonding Conduit to Structure 192 11.17 Aluminum Jumper Connection to Tubular Structure 192 11.18 Sizes of Identification Sleeving 104 11.19 Identification of Wire Bundles and Harnesses 195 11.20 AN Connectors 107 11.21 Classes of AN Connectors 187 11.22 AN Connector Marking .. 198 11.28 Spare Wires for Potting Connector 199 11.24 Bend Radil for Rigid Conduit 201 ### Page 42 (OCR) Page xiv AC 43.19-1 197 LIST OF ILLUSTRATIONS- Continued Pioure Page 11.25 Minimum Bending Radii for Flexible Aluminum or Brass Conduit 202 11.26 Group and Bundle Ties 208 11.27 Comb for Straightening Wires in Bundles 208 11.28 Slack Between Supports 204 11.29 Drainage Hole in Low Point of Tubing 204 11.80 Separation of Wires from Plumbing Lines 205 11.81 Safe Angle for Cable Clamps .. 206 11.82 Typical Mounting Hardware for MS-21919 Cable Clamps 206 11.88 Installing Cable Clamp to Structure - 207 11.84 Installing Cable Clamps to Tubular Structure 207 11.85 Single Cord Lacing 208 11.86 Double Cord Lacing 209 11.87 Making Ties.. 209 11.88 Lead-Acid Battery Electrolyte Freezing Points 210 11.89 Sulfuric Acid Temperature Correction 211 12.1 Propeller Repair by Addition of Small Inlay 217 12.2 Method of Repairing Surface Cracks, Nicks, etc., on Aluminum Alloy Propellers 224 12.8 Correct and Incorrect Method of Reworking Leading Edge of Aluminum Alloy Propellers 225 12.4 Repair Limits to Blade Section Width and Thickness for Aluminum Alloy Propellers 226 12.5 Method of Repairing Damaged Tip of Aluminum Alloy Propellers 227 12.6 Protractor and Method of Measuring Angle of Bend in Aluminum Alloy Propellers 228 12.7 Repair of Fixed-Pitch Hub and Propeller With Elongated or Damaged Boltholes 281 12.8 Blade Leading Edge and Trim Tab 287 12.9 Blade Bonding and Crack Location 288 12.10 Blade Damage Location 288 12.11 Blade Cross-Section 289 12.12 Blade Retention and Link Attach Points 240 18.1 Typical Datum Locations 244 18.2 Illustration of Arm (or Moment Arm) 245 18.8 Example of Moment Computation . 246 18.4 Empty Weight Center-of-Gravity Formulas 247 18.5 Empty Welght and Empty Weight Center-of-Gravity-Tail-Wheel Type Aircraft - 248 18,6 Empty Weight and Empty Weight Center-of-Gravity--Nose-Wheel Type Aircraft 249 18.7 Operating Center-of-Gravity Range 250 18.8 Weighing Point Centerline 251 18.9 Empty Weight and Empty Weight Center-of-Gravity When Aircraft is Weighed With Oil 252 18.10 Example of Check of Most Forward Weight and Balance Extreme 254 18.11 Example of Check of Most Rearward Weight and Balance Extreme 255 18.12 Loading Conditions: Determination of the Number of Passengers and Baggage Permissible with Full Fuel 256 18.18 Loading Conditions: Determination of the Fuel and Baggage Permissible with Maximum Passengers 258 18.14 Loading Conditions: Determination of the Fuel and the Number and Location of Passengers Permissible with Maximum Baggage 259 18.15 Effects of the Addition of Equipment Items on Balance 260 18.16 Example of Moment and Weight Changes Resulting from Equipment Changes 261 18.17 Sample Welght and Balance Report To Determine Empty Weight Center-of-Gravity 262 ### Page 43 (OCR) C 43.13-14 Page xY 973 LIST OF ILLUSTRATIONS- Continued Pinurs Page 18.18 Sample Weight and Balance Report Including an Equipment Change for Aircraft Fully Loaded 268 13.19 Permanent Ballast Computation Formula 264 14.1 Schematic of Typical Differential Pressure Compression Tester 278 14.2 Spark Plug Heat Ranges 214 14.8 Spark Plug Reach 275 14.4 Method of Checking Spark Plug Gap . 275 14.58 Typical Carbon Fouled Spark Plug . 276 14.5b Typical Lead Fouled Spark Plug 276 14.бе Typical Oil Fouled Spark Plug 276 14.6а Typical Eroded Spark Plug 277 14.6b Typical Spark Plug With Copper Runout 277 14.7 Typical Spark Plug With Bridged Electrodes 277 14.8 Spark Plug Well Flashover 278 14.9 Typical Method of Clamping Leads 278 14.10 Normal Contact Point -- 279 18.11 Point With Minor Irregularities 279 14.12 Point With Well-Defined Mound 279 14.18 Location of Clamps at Tube Bends 281 14.14 Typical Muffler Wall Fatigue Failure at Exhaust Outlet 285 14.15 Typical Muffer Wall Failure .. 286 14.16 Typical Muffler Wall Fatigue Failure 286 14.17 Typical Fatigue Failure of Muffler End Plate at Stack Inlet 287 14.18 Section of Muffler Showing Typical Internal Baffle Fallure 287 14.19 Loose Pieces of Failed Internal Baffle .. 287 14,20 Failed Internal Baffle Partially Obstructing Muffler Outlet 288 14.21 Failed Internal Baffle Completely Obstructing Muffer Outlet 288 14.22a Example of Exhaust Outlet Guard ... 288 14.22b Example of Exhaust Outlet Guard Installed 289 14.28 Effect of Improperly Positioned Exhaust Pipe/Muffler Clamp 289 14.24 Primary Inspection Areas 290 ### Page 44 (OCR) AC 43.13-1A Page 1 1972 Chapter 1. AIRCRAFT WOOD STRUCTURES Section 1. MATERIALS AND PRACTICES 1. MATERIALS. Three forms of wood are commonly used in aircraft: solid wood, plywood, and laminated wood. Although several kinds of modified wood are sometimes used for special purposes, these three forms constitute the bulk of all wood aircraft construction materials. a. Quality of Wood. All wood and plywood used in the repair of aircraft structures should be of aircraft quality. Figure 1.2 fists the permissible variations in characteristics and properties of aircraft wood. bracing and checking or splitting of wood members. A few suggestions for minimizing these shrinkage effects are: (1) Use bushings that are slightly short so that when the wood member shrinks the bushings do not protrude and the fittings may be tightened firmly against the member. (2) Gradually drop off plywood face plates either by feathering or by shaping as shown in figure 1.1. b. Species Substitution. The species used to repair a part should be the same as that of the original whenever possible; however, permissible substitutes are given in figure 1.2. c. Effects of Shrinkage. When the moisture content of wood is lowered, its dimensions decrease. The dimensional change is greatest in a tangential direction (across the fibers and parallel to the growth rings), somewhat less in a radial direction (across the fibers and perpendicular to the growth rings), and is negligible in a longitudinal direction (parallel to the fibers). These dimensional changes can have several detrimental effects upon a wood structure, such as loosening of fittings and wire 2. REPLACEMENT OF DRAINHOLES AND SKIN STIF- FENERS. Whenever repairs are made that require replacing a portion that includes drainholes, skin stiffeners, or any other items, the repaired portion must be provided with similar drainholes, skin stiffeners, or items of the same dimensions in the same location. Reinforcing, under skin repairs, that interferes with the flow of water from some source, such as inspection holes, is to be provided with drainholes at the lowest points. **3. CONTROL SURFACE FLUTTER PRECAUTIONS.** When repairing control surfaces, especially on high-performance airplanes, care must be exer- FEATHERED END/ SPADED END- FIGURE 1.1.-Tapering of face plates. Chap 1 Par 1 ### Page 45 (OCR) Page 2 AC 43.13-1A 1972 FIGURE 1.2.—Selection and properties of aircraft wood. Maximum utrength permisaible Species of wood properties as compared to ation (alope Frain devi- Remarks spruce of grain) Spruce (Picea) Sitka (P. Bit- 100% 1:15 Excellent for all us. Considered as standchensle) Red (P. Bubra) ard for this table. White (P. Glauce). Douglas Fir (Paeudotsuga Taxi- Exceeds spruce 1:16 May be need as substitute for spruce in same folia). sizes or in alightly reduced sizes providing reductions are substantiated. Difficult to work with hand tools. Some tendency to split and splinter during fabrication and considerable more care in manufacture is necessary. Large solid pleces should be avoided due to Inspection difficulties. Gluing satisfactory. Noble Fir (Ables Noblies) - Slightly exceede spruce 1:15 Satlafactory characteristics with respect to except 8 percent workability, warping, and splitting. May be deficient in shear. used as direct substitute for spruce In same sizes providing shear does not become critical. Hardness somewbat less then spruce. Gluing satisfactory. Western Hemlock (Truga Het- Slightly exceeds spruce 1:15 Less uniform in texture than spruce. Afay be eropbylia). used as direct substitute for spruce. Upland growth superior to lowland growth. Gluing satisfactory. Pine, Northern White (Pinus Properties hetween 85 1:15 Excellent working qualities and uniform in Strobus). percent and 96 per- propertles but somewhat low in hardness and cent those of spruce. abock-resisting capacity. Cannot be used as substitute for spruce without increase in sizes to compensate for lesser strength. Gluing satisfactory. White Cedar, Port Oxford Exceeds spruce - (Charaecyparla Lawsoniana). 1:15 Nay be used as substitute for spruce In same sizes or in slightly reduced sizes proriding reductione are substantiated. Easy to work with band toola. Gluing difficult but satistactory joints can be obtained if suitable precau tlone are taken. Poplar, Yellow (Lirlodendron Slightly leas than 1:15 Excellent working qualities. Should not be Tulipifera). spruce except in used as a direct aubstitute for spruce without compression (crush- carefully accounting for slightly reduced ing) and shear. strength properties. Somewhat low in shockresisting capacity. Gluing satisfactory. (Bee notes follosoing table.) Notes for Agure 1.2. 1. Defeots Permitted. a. Oross grain. Spiral grain, diagonal grain, or a combination of the two is acceptable providing the grain does Par 3 ### Page 46 (OCR) AC 43.13-1A Page 3 1972 not diverge from the longitudinal axis of the material more than specifed in column 8. A check of all four faces of the board is necessary to determine the amount of divergence. The direction of free-flowing ink will frequently assist in determining grain direction. b. Wacy, ourly, and interlocked grain. Acceptable, if local irregularities do not exceed limitations specifed for spiral and diagonal grain. c. Eard knots. Sound bard knots up to ¼-inch In maximum diameter are acceptable prorlding: (1) they are not in projecting portions of I-beams, niong the edges of rectangular or beveled unrouted beams, or along the edges of manges of box beams (except in lowly stressed portions); (2) they do not cause grain divergence at the edges of the board or in the fanges of a beam more than specifed in column 8; and (B) they are in the center third of the beam and are not closer than 20 inches to another knot or other defect (pertains to %-inch knots-smaller knots may be proportionately closer). Knote greater than ¾-inch must be used with caution. d. Pin knot olusters. Small clusters are acceptable providing they produce only a small effect on grain direction. e. Pitoh pookots. Acceptable, in center portion of a beam providing they are at least 14 inches apart when they lie in the same growth ring and do not exceed 1½ inch length by %-inch width by ⅛-inch depth and providing they are not along the projecting portions of I-beams, along the edges of rectangular or beveled unronted beams, or along the edges of the flanges of box beams. f. Afineral streaks. Acceptable, providing careful inspection falls to reveal any decay. 2. Defects Not Permitted. a. Cross grain. Not acceptable, unless within limitations noted in 1a. D. Wavy, curly, and interlocked grain. Not acceptable, unless within limitatlons noted in 1b. c. Hard knots. Not acceptable, unless within limitations noted in 1c. d. Pin knot oluaters. Not acceptable, If they produce large effect on grain direction. e. Spike knots. These are knots running completely through the depth of a beam perpendicular to the annual ringe and appear most frequently in quartersawed lumber. Reject wood containing this defect. 1. Pitch pockets. Not acceptable, unless within limitations noted in le. g. fineral streaks. Not acceptable. If accompanied by decay (see 1f). b. Checks, shakes, and aplits. Checks are longitudinal cracks extending, in general, across the annunl rings. Shakes are longitudinal cracks usually between two annual rings. Splits are longitudinal cracks induced by artificially induced stress. Reject wood containing these defects. 1. Compression cood. This defect is very detrimental to strength and is difficult to recognize readily. It la characterized by high specific gravits; has the appenrance of an excessive growth of summer wood; and in inost species shows but little contrast in color between epring wood and summer wood. In doubtful cases reject the material, or subject samples to a tonghness machine test to establish the quality of the wood. Reject all materlal containing compression mood. j. Compressio» failures. This defect ia caused from the wood belng overstressed in compresslon due to natural forces during the growth of the tree, felling trees on rough or irregular ground, or rough handling of logs or lnmber. Compression failures are choracterized by a buckling of the fbers that appear as streaks on the surface of the piece substantially at right angles to the grain, and rary from pronounced fallures to very fine hairlines that require close inspection to detect. Reject wood containing obvious failures. In doubtful cases reject the wood, or make a further inspection in the form of microscopic examination or toughness test, the latter means being the more rellable. k. Decay. Examine all stains and discolorations carefully to determine whether or not ther are harmless, or in a stage of preliminary or advanced decay. All pleces must he free from rot, dote, red heart, purple heart, and all other cised that the repairs do not involve the addition of weight aft of the hinge line. Such a procedure may adversely disturb the dynamic and static balance of the surface to a degree which would induce flutter. As a general rule, it will be necessary to repair control surfaces in such a manner that the structure is identical to the original so that the weight distribution and resulting mass balance are not affected in any way. 4. GLUING PRECAUTIONS. Satisfactory glue Chap 1 joints in aircraft will develop the full strength of wood under all conditions of stress. To produce this result, the gluing operation must be carefully controlled so as to obtain a continuous, thin, uniform film of solid glue in the joint with adequate adhesion to both surfaces of the wood. Some of the more important conditions involve: o. Properly prepared wood surfaces. b. Glue of good quality, properly prepared. c. Good gluing technique. Par 3 ### Page 47 (OCR) Pago 4 5. PREPARATION OF WOOD SURFACES FOR So more than GLUING, It is recommended that no more than 8 hours be permitted to elapse between final surfacing and gluing. The gluing surfaces should be machined smooth and true with planers, jointers, or special miter saws. Planer marks, chipped or loosened grain, and other surface irregularities will not be permitted. Sandpaper must never be used to smooth softwood surfaces that are to be glued. Sawed surfaces must approach well-planed surfaces in uniformity, smoothness, and freedom from crushed fibers. a. Tooth-planing, or other means of roughening smooth well-planed surfaces of normal wood before gluing is not recommended. Such treatment of well-planed wood surfaces may result in local irregularities and objectionable rounding of edges. While sanding of planed surfaces is not recommended for softwoods, sanding is a valuable aid in improving the gluing characteristics of some hard plywood surfaces, wood that has been compressed through exposure to high pressure and temperatures, resin-impregnated wood (impreg and compreg), and laminated paper plastic (papreg). b. Wood surfaces for gluing should be free from oil, wax, varnish, shellac, lacquer, enamel, dope, sealers, paint, dust, dirt, oil, glue, crayon marks, and other extraneous materials. c. Wetting tests are useful as a means of detecting the presence of wax. Drops of water placed on the surface of wax-coated wood do not spread or wet the wood. At present, preliminary gluing tests appear to be the only positive meana of actually determining the gluing characteristics of plywood surfaces. 6. GLUES. Glues used in aircraft repair fall into two general groups: casein glues and resin glues. Any glue that meets the performance requirements of applicable United States military specifications or has previously been accepted by the FAA is satisfactory for use in certificated civil aircraft. In all cases, glues are to be used strictly in accordance with the glue manufacturer's recommendations. Par 5 AC 43,13-1A 1972 a. Caseln Glues. Casein glues have been widely used in wood aircraft repair work. The forms, characteristics, and properties of water-resistant casein glues have remained substantially the same for many years, except for the addition of preservatives. Casein glues for use in aircraft should contain suitable preservatives such as the chlorinated phenols and their sodium salts to increase their resistance to organic deterioration under high-humidity exposures. Most casein glues are sold in powder form ready to be mixed with water at ordinary room temperatures. b. Synthetic Resin Glues. Synthetic resin glues for wood are outstanding in that they retain their strength and durability under moist conditions and after exposure to water. The best known and most commonly used synthetic resin glues are the phenol-formaldehyde, resorcinolformaldehyde, and urea-formaldehyde types. The resorcinol-formaldehyde type glue is recommended for wood aircraft applications. Materials, such as walnut-shell flour, are often added by the glue manufacturer to the resin glues to give better working characteristics and joint-forming properties. The suitable curing temperatures for both urea-formaldehyde and resorcinol glues are from 70° F. up. At the 70° F. minimum temperature it may take as long as 1 week for the glueline in a spar splice to cure to full strength. Thinner pieces of wood and/or higher curing temperatures shorten curing time considerably. The strength of a joint cannot be depended upon if assembled and cured at temperatures below 70° F. 7. MIXING OF RESIN GLUES. Liquid resin glues may come ready for use or in a form which requires only the addition of a hardener. In all cases, the mixing, glue consistency, assembly time, etc., should comply with the glue manufacturers' recommendations and instructions. Cold-setting synthetic-resin glue, when prepared for use, is limited in working life, and care should be taken to discard the glue and clean the equipment before the end of the working life period. In very warm weather it may be found advisable to keep the gluepot in a bath of cool water, approximately 70° F., to prolong the working life of the mixture. Chap 1 ### Page 48 (OCR) AC 43.13-1A Page 7 1972 1 6A RECOMMENDED •- DA MINIMUM ‡ - 12 A RECOMMENDED- 6A RECOMMENDEO 10A MINIMUM 5A MINIMUM NO FITTINGS WITHIN THESE LIMITS 1/4 A + **DIRECTION OF GRAIN IF SPRUCE** OR OUTER FACE GRAIN IF PLYWOOD REINFORCEMENT PLATES TO BE SPRUCE OR PLYWOOD AND SHALL BE GLUED ONLY. SOLID SPARS MAY BE REPLAGED WITH LAMINATED ONES OR VICE VERSA, PROVIDED THE MATERIAL IS OF THE SAME HIGH QUALITY. FIoURE 1.4.-Method of splicing solid or laminated rectangular spars. but not necessarily, at a bolthole or cutout and spread in each direction until, in most cases, they extend a short distance beyond the ends of the plates where the resistance to spar shrinkage disappears. Other factors which have been found conducive to the formation of cracks, due to spar shrinkage in the region of plywood plates, are poor protective finishes, large cutouts, and metal fittings which utilize two lines of large diameter bolts. The presence of cracks does not necessarily mean that the spar must be discarded. If the crack is not too long or too close to either edge and can be reinforced properly, it will probably be more economical and satisfactory to effect a repair than to install a new spar or section. However, a generally acceptable proce dure suitable for all airplane models cannot be Chap 1 described here. It is recommended the manufacturer or the Federal Aviation Administration be contacted for specific instructions before making repairs not in accordance with the manufacturer's approved instructions or the recommendations of this advisory circular, because of the possibility of strength deficiencies. 16. ELONGATED HOLES IN SPARS. In cases of elongated boltholes in a spar or cracks in the vicinity of boltholes, splice in a new section of spar, or replace the spar entirely, unless the method of repair is specifically approved by a representative of the HAA. In many cases, it has been found advantageous to laminate the new section of the spar (using aircraft plywood for the outer faces, particularly if the spar roots are being replaced. Par 15 ### Page 49 (OCR) Page 8 АС 43.13-1А 1972 IF SPLICE IS MADE WHERE ROUTIN 18 FEATHERED TO FULL WIDTH OF SPAR, TAPERED PLATES OONFORMING TO THE CONTOUR OF THE ROUTINS, SHOULD FIRST BE ADDED. **OTHERWISE THE SPLICE N THE SAME AS SHOWN.** БА - 6А - • 6А - 12A RECOMMENDED - **IDA MINIMUM** **NO FITTING WITHIN THESE LIMITS** **DIRECTION OF GRAIN** REINFORCEMENT PLATES MAY BE OF PLYWOOD, THE SAME MATERIAL AS THE SPAR, OR MATERIAL OF EQUAL OR HIGHER QUALITY AND SHALL BE GLUED ONLY. FIGURE 1.5.-Method of splicing solid "I spars. 17. RIB REPAIRS. Make replacement ribs from & manufacturer's approved drawing, or from a drawing made by the repair agency, and certified by the manufacturer as correct. The original rib may be used as a pattern in making the new rib if it is not too seriously damaged to permit comparison. Wood ribs should not be attached to wood spars by nails driven through the rib capstrips, as this weakens the rib materially. The attachment should be by means of glue with cement-coated, barbed, or spiraled nails driven through the vertical rib members on each side of the spar. The drawing or pattern should be retained by the repair agency for use by the FAA inspector when making his inspection. (Acceptable methods of repairing damaged ribs are shown in figure 1.10.) a. Compression Ribs. Acceptable methods of repairing damaged compression ribs are shown in figure 1.11. (A) illustrates the repair of a compression rib of the "I" section type; i.e., wide, shallow capstrips, a center plywood web with a rectangular compression member on Par 17 each side of the web. The rib is assumed to be cracked through capstrips, web member, and compression member. Cut the compression member as shown in (D), remove and replace the shortest section, adding the reinforcement blocks as also shown in figure 1.11(D). Cut and replace the aft portion of the capstrips and reinforce as shown in figure 1.10 excep that the reinforcement blocks are split in the vertical direction to straddle the center web. The plywood sideplates, as indicated in figure 1.11(A), are glued on. These plates are added to reinforce the damaged web. (B) illustrates a compression rib of the type that is basically a standard rib with rectangular compression members added to one side and plywood web to the other side. The method used in this repair is essentially the same as in (A) except that the plywood reinforcement plate shown in solid black in section B-B is continued the full distance between spars. (C) illustrates a com pression rib of the "I" type with a rectangular vertical member each side of the web. Th Chap ### Page 50 (OCR) I€ 43.13-1A Page 9 972 10 д **NO FITTING WITHIN THESE LIMITS** - DIRECTION OF GRAIN IN - 1/28 **PLYWOOD REINFORCEMENT** **PLATES TO BE SAME AS ORIGINAL WEB.** **PLYWOOD** **A-SOLID WOOD FILLER BLOOKS** SECTION A-A FIGURE 1.6.-Repairs to built-up "I spar. method of repair is essentially the same as in (A) except the plywood reinforcement plates on each side shown in solid black in section -C are continued, as in (C), the full distance between spars. 18. PLYWOOD SKIN REPAIR. Make extensive repairs to damaged stressed skin plywood structures in accordance with apecific recommendations from the manufacturer. It is recommended hat repairs be made by replacing the entire panel from one structural member to the hext, if damage is very extensive. When damaged plywood skin is repaired, carefully inspect the adjacent internal structure for possible hidden damage. Repair any defective frame members prior to making skin repairs. a. Types of Patches. Four types of patcheshe surface or overlay patch, the splayed atch, the plug patch, and the scarf patch-are acceptable for repairing plywood skins. Surface patches should not be used on skins over 1/8 inch thick. Splayed patches should not be used on skins over 1/10 inch thick. There are no skin thickness limitations for the use of scarf patches and plug patches. b. Determination of Single or Double Curvature, Much of the outside surface of plywood aircraft is curved. On such areas, plywood used for repairs to the skin must be similarly curved. Curved skins are either of single curvature or of double (compound) curvature. A simple test to determine which type of curvature exists may be made by laying a sheet of heavy paper on the surface in question. If the sheet can be made to fit the surface without wrinkling, the surface is either flat or has single curvature. If, however, the sheet cannot be made to fit the surface without wrinkling, the surface is of double curvature. op 1 Par 17 ### Page 51 (OCR) Page 10 AC 43.13-1A 1972 **ANGLE OF FACE GRAIN SAME AS IN ORIGINAL WEB** NEW WEB STAGGER WEB SPLICES (See Fig. 1.8 for method of splicing plywood webs). • 10 B C **FOR CLARITY IOA - 6A- NEW WEB NOT INOLUDED IN THIS VIEW.** - SPRUCE REINFORCEMENT PLATES. A, 8,0, D = ORIGINAL DIMENSIONS. FoURs 1.7.-Method of splicing box-spar flanges (plate method). 19. REPAIRS TO SINGLE CURVATURE PLYWOOD SKIN. Repairs to skins of single curvature may usually be formed from flat plywood, either by bending it dry or after soaking it in hot water. The degree of curvature to which a piece of plywood can be bent will depend upon the direction of the grain and the thickness. Figure 1.12 is presented as a guide in determining which process of bending should be used for the curvature being considered. Plywood, after softening, may be bent on a cold ventilated form, or it may be bent over the leading edge near the part being patched if space permits. In either method it should be allowed to dry completely on the form. When bending plywood over a leading edge, drying may be hastened by laying a piece of coarse burlap over the leading edge before using it as a bending form. A fan to circulate the air over the bent piece will speed the drying. In bending pieces of small radii, or to speed up the bending of a large number of parts of the same curvature, it Par 19 may be necessary to use a heated bending form. The surface temperature of this form may be as high as 149° C. (800° F.), if necessary, without danger of damage to the ply. wood. The plywood should be left on the form, however, only long enough to dry to room conditions. 20. REPAIRS TO DOUBLE CURVATURE PLYWOOD SKIN. The molded plywood necessary for a repair to a damaged plywood skin of double curvature cannot be made from flat plywood unless the area to be repaired is very small, or is of exceedingly slight double curvature; there fore, molded plywood of the proper curvature must be on hand before the repair can be made. If molded plywood of the proper curvature is available, the repair may be made following the recommended procedures. 21. SPLAYED PATCH. Small holes with largest dimensions not over 15 times the skin thickness in skins not more than 1/10 inch in thickness Chop ### Page 52 (OCR) AC 43.13-1A Pogo 11 972 JOAT - INSERTED WEB SECTION 46- (1/6 FLANGE DEPTH) PLYWOOD COVER STRIPS (Same thickness as "A"). FILLER BLOCKS I. AFTER INSERTED WEB HAS BEEN GLUED AND NAILED IN PLACE,GLUE AND NAIL COVER STRIP OVER ENTIRE LENGTH OF SPLICE JOINTS. 2. SECTIONAL SHAPE OF FILLER BLOCKS MUST CONFORM EXACTLY TO TAPER OF SPAR. THEY MUST NOT BE TOO TIGHTLY FITTED OR WEDGING ACTION WILL LOOSEN EXISTING GLUE JOINTS OF WEBS TO FLANGES. IF TOO LOOSELY FITTED, CRUSHING OF WEB WILL OCCUR WHEN CLAMPING. SIRIT TO SAME AS ORIGINAL PLY OF NEW PLYWOOD WEB AND COVEF FIGURE 1.8.-Method of splicing box-spar webs. Par 21 ### Page 53 (OCR) **NO FITTING WITHIN THESE LIMITS 5 TO 1 SLOPE** - LONGITUDINAL CRACK **AND LOCAL DAMAGE SPRUCE BLOCKY** **DIRECTION OF GRAIN IF** SPRUCE OR FAGE GRAIN IF PLYWOOD FIGURE 1.9.—Method of reinforcing a longitudinal crack and/or local damage in a solid or internally routed apar. AC 43.13-1A 1972 ### Page 54 (OCR) $3.1$- Page 13 • 12F RECOMMENDED -1 1OF MINIMUM **DAMAGED AREA GUSSET - ORIGINAL Y - 8** PLYWOOD- C== **SPRUCE OR** BASSWOOD AT A SPAR *4А† = C= •Од- ЗА AT TRAILING EDGE -10 A A, B,O, D, E, F, 0 • ORIGINAL DIMENSIONS• **REINFORCEMENT PLATES** **SHALL BE PLYWOOD GLUED AND NAILED.** **AT A JOINT TO TRUSS MEMBERS THE DIRECTION OF THE** **FACE GRAIN MUST BE** **THE SAME AS ORIGINAL GUSSET.** T +3А→ - 10 A- FIGURE 1.10.-Repair of wood ribs. ар 1 Par 21 ### Page 55 (OCR) Page 14 AC 43.13-1А 1972 SEE FIG. 1.10 FOR REPAIR OF CAPSTRIPS. , REPAIR DAMAGE- SEE (D) (A) -ЗА - ЗА SEE FIG. 1.10 FOR REPAIR OF CAPSTRIPS. - REPAIR DAMAGE- (B) SEE FIG. 1.10 FOR REPAIR OF CAPSTRIPS. → REPAIR DAMAGE (G) * 12 A RECOMMENDED→ IOA MINIMUM 6 A- (D) А-А C-0 **PLYWOOD** **REINFORCEMENT E SAME THICKNESS AND FACE GRAIN DIRECTION AS ORIGINAL.** FIGURE 1.11.-Typical wing compression rib repairs. Par 21 Chap 1 ### Page 56 (OCR) Ju r'ercent moisture content, Thoroughly soaked in hot water 1972 AC 43.13-1A bent on cold mandrels and bent on cold mandrels PLywood At 900 to At 0° or 450 At 900 to At 00 or 450 thickness face grain to face grain face grain to face grain ·→ (I) (2) (3) (4) (5) (6) Inch No. plies Inches Inches Inches Inches 3 0.5 0.1 3 105 ole 3 2.6 •8 3 3.8 1.2 3 5.3 1,8 3 761 2.6 5 17 6 2 5 21 7 3 5 27 10 4 5 31 20 12 5 5 43 28 16 7 5 54 36 21 10 Columns (1) and (3) may also be used for determining the maximun thickness of single laminations for curvel members. FIGURE 1.12.—Minimum recommended bend radii for aircraft plywood. ### Page 57 (OCR) may be repaired by using a circular splayed patch as illustrated in figure 1.18. The term "splayed" is used to denote that the edges of the patch are tapered, but the slope is steeper than is allowed in scarfing operations. The steps shown in figure 1.18 should be taken in making a splayed patch. a. Lay out the patch according to figure 1.18. Tack a small piece of plywood over the hole for a center point, and draw two circles, the inner one to be the size of the hole, and the outer one marking the limits of the taper. The difference between the radii is 5T (5 times the thickness of the skin). If one leg of the dividers has been sharpened to a chisel edge, the dividers may be used to cut the inner circle completely through. b. Taper the hole evenly to the outer mark with a chisel, knife, or rasp. c. Prepare a circular tapered patch to fit the prepared hole and glue the patch into place with face-grain direction matching that of the original surface. d. Use waxed paper between the patch and a AC 43.13-1А 1972 plywood pressure plate cut to the exact size of the patch. This prevents extruded glue from binding patch and plate together. Center the plate carefully over the patch. •. Apply pressure. As there is no reinforcement behind this patch, care must be used so that pressure is not great enough to crack the skin. On horizontal surfaces, weights or sandbags will be sufficient. On patches too far in for the use of standard hand clamps, jaws of greater length may be improvised. f. Fill, sand, and refinish the patch. 22. SURFACE PATCH. Plywood skins that are damaged between or along framing members may be repaired by surface or overlay patches as shown in figure 1.14. Trim the damaged skin to a rectangular or triangular shape and round the corners. The radius of rounded corners should be at least 5 times the skin thicknesa. Cover surface patches with fabric before finishing. Fabric should overlap the original skin at least 2 inches. Surface patches located TRIM TO CIRCULAR SHAPE (15 T MAXIMUM DIAMETER. MINIMUM DISTANCE TO FRAME = 15 T.) -FACE GRAIN OF PATCH PARALLEL TO FACE GRAIN OF SKIN. INCH WEIGHTS OR CLAMP OR 1585 PRESSURE PLATE fOR "PLYWOOD **PAPER OR CELLOPHANE** - PLYWOOD SKIN FlOURe 1.18.-Splayed patch. Par 21 Chap 1 ### Page 58 (OCR) AC 43.13-1A 11972 entirely aft of the 10 percent chordline, or which wrap around the leading edge and terminate aft of the 10 percent chordine are permissible. Bevel forward edges of patches located entirely aft of the 10 percent chordline to 4 times the skin thickness. Surface patches may have as much as a 50-inch perimeter and may cover as much as 1 frame (or rib) space. The face-grain direction must be the same as the original skin. 23. SCARF PATCH. A properly prepared and inserted scarf patch is the best repair for damaged plywood skins and is preferred for most skin repairs. Figure 1.15 shows the details and dimensions to be used when installing typical scarf skin patches when the back of the skin is accessible. Follow figure 1.16 when the back of the skin is not accessible. The scarf slope of 1 in 12 shown in both figures is the steepest slope permitted for all species of plywood. If the radius of curvature of the skin at all points on the trimmed opening is greater than 100 times the skin thickness, a scarf patch may be installed. Scarf cuts in plywood may be made by hand plane, spoke shave, scraper, or accurate sandpaper block, Rasped surfaces, except at the corners of scarf patches and sawed surfaces, are not recommended as they are likely to be rough or inaccurate. Nail-strip gluing is often the only method avallable for gluing scarf joints in plywood when used in repair work; therefore, it is esential that all scarf joints in plywood be wacked with plywood or solid wood to provide adequate nailholding capacity. The face grain direction of the plywood patch should be the same as that of the original skin. a. Scarf Patches (Back of Skin Accessible), When he back of a damaged plywood skin is accessible (such as a fuselage skin), it should be repaired with scarf patches following the details shown in figure 1.15. Whenever possible, the edge of the patch should be supported as shown in section C-C. When the damage folows or extends to a framing member, the carf may be supported as shown in section hap 1 Page 17 Damages that do not exceed 25 times the skin thickness in diameter after being trimmed to a circular shape, and if the trimmed opening is not nearer than 15 times the skin thickness to a framing member, may be repaired as shown in figure 1.15, section D-D. The backing block is carefully shaped from solid wood and fitted to the inside surface of the skin, and is temporarily held in place with nails. A hole, the exact size of the inside circle of the scarf patch, is made in the block, and is centered over the trimmed area of damage. The block is removed, after the glue on the patch has set, and leaves a flush surface to the repaired skin. b. Steps in Making Scarf Patch (Back of Skin Not Accessible). (1) After removing damaged sections, install backing strips, as shown in figure 1.16, along all edges that are not fully backed by a rib or a spar. To prevent warping of the skin, backing strips should be made of a soft-textured plywood, such as yellow poplar or spruce rather than solid wood. All junctions between backing atrips and ribs or spars should have the end of the backing strip supported by a saddle gusset of plywood. (2) If needed, nail and glue the new gusset plate to rib. It may be necessary to remove and replace the old gusset plate by a new saddle gusset; or, it may be necessary to nail a saddle gusset over the original. (3) Attach nailing strips to hold backing strips in place while the glue sets. Use bucking bar where necessary to provide support for nailing. Unlike the smaller patches made in a continuous process, work on the airplane must wait while the glue, holding the backing strips, sets. After setting, complete finishing in usual manner. 24. PLUG PATCHES. Two types of plug patches, oval and round, may be used on plywood sking provided the damage can be covered by the patches whose dimensions are given in figures 1.17 and 1.18. As the plug patch is strictly a skin repair, it should be used only for damage that does not involve the supporting structure under the skin. Oval patches must be prepared Far 12 ### Page 59 (OCR) Page 18 AC 43.13- JL _ JL **FRONT SPAR** •TRIMMED OPENING -MINIMUM RADIUS ST - RIBS • DAMAGE →Д SADDLE GUSSET -BEAB SPAR 30 т ￾ PATCHI -12 T-→ PLYWOOD SKIN (I" MINIMUM 8T ST It MIN PLYWOOD SADDLE GUSSET RIB CAP MINIMUM THICKNESS • T. **NAILED - GLUED IN PLACE.** SECTION A-A 12 T. - PATCH /ІГРАТОН *4T UNSUPPORTED -RIB. CAP LAP SECTION B-B SECTION C-C FIGURE 1.14.-Surface patches. Par 24 Chap ### Page 60 (OCR) Saddle Gusset Hailing Strips Maximum Diameter 251 Minimum Distance 15T • Temporary Backing 30T- Patch - 12T- Minimum Plywood Saddle Gusset Minlmum Thickness =T Nail-Glued in Place SECTION A-A 12T ~ - Patch - 121--- - Patch Backing 31 Clamp and Glue Backing Clamp and Glue Backing to Frame and Skin Framing Member to Frame and Skin **SECTION B-B SECTION C-C** • Hailing Strips- - 12T - 7 LEA Tomporary Backing Block Shape to fet Skir Cellophane or Paper SECTION D-D FIGURE 1.15.-Scarf patches. back of akin accessible. Per 24 ### Page 61 (OCR) Page 20 AC 43.1: FRONT SPI SADDLE GUSSET+ **REAR SPAR** • 30 T -12T- - PLYWOOD SKIN T 3T (· MINIMUM)- 8T (1" MINIMUM) RIb CAP **SECTION A - A PLYWOOD SADDLE GUSSET** PLYWOOD (MINIMUM THICKNESS SKIN- PATCH -12 T-.- PATCH RIB CAP • 3T "SPAR" PLYWOOD OR SPRUCE- (* MINIMUM) **PLYWOOD SKIN** SECTION B - B SECTION C-C FIGURE 1.16.-Scarf patches--back of akin not accessible. Par 24 Chap ### Page 62 (OCR) AC 43.13-1A Page 21 1972 - OUTER EDGE OF DOUBLER - BUTT JOINT OF PATCH TO SKIN -INNER EDGE OF DOUBLER - NAIL HOLES - **SCREW HOLES -** **TO BE FILLED BEFORE FINISHING.** PATCH DIMENSIONS _G SMALL 27112 LARGE 5 7 **TWO ROWS OF SCREWS AND** **NAILS REQUIRED ON LARGE** • BUTT JOINT OF PATCH TO SKIN PATCH. SPLUG PATCH (GRAIN PARALLEL TO SKIN) а шатика **PLYWOOD DOUBLER (GRAIN PARALLEL TO SKIN)** PLYWOOD SKIN FIOURE 1.17.-Oval plug patch assembly. with the face grain carefully oriented to the same direction as the original skin. Orientation of the face grain direction of the round plug patch to that of the skin surface is no problem, as the round patch may be rotated until grain Hirections match. a. Steps in Making Oval Plug Patch. (1) Explore the area about the hole to be sure it lies at least the width of the oval doubler from a rib or a spar. Refer to figure 1.17 'or repair details. (2) Lay a previously prepared oval plug ratch over the damage and trace the patch. Saw to the line and trim the hole edges with a knife and sandpaper. (3) Mark the exact size of the patch on one surface of the oval doubler and apply glue to the area outside the line. The oval doubler hould be made of some soft-textured plywood, Chap 1 such as yellow poplar or spruce. Insert doubler through the hole and bring it, glue side up, to the underside of the skin with its pencil outline of the patch matching the edges of the hole. If the curvature of the surface to be repaired is greater than a rise of 1/8 inch in 6 inches, the doubler should be preformed by hot water or steam bending to the approximate curvature. (4) Apply nailing strips outlining the hole to apply glue pressure between doubler and akin. Use bucking bar to provide support for nailing. When two rows of nails are used, stagger nail spacing. (5) Apply glue to remaining surface and to an equivalent surface on the patch. (6) Lay the patch in position over the doubler, and screw the pressure plate to the patch assembly using a small nail to line up the holes Par 24 ### Page 63 (OCR) Page 22 AC 43.13-' 19: **GRAIN DIRECTION OF SKIN, PATCH, AND DOUBLER.** **SAW OUT IN DOUBLER** **BUTT JOINT OF PATCH TO SKIN.** **INNER EDGE OF** **DOUBLER** **OUTER EDGE OF** **DOUBLER** **NAIL HOLES -** **SCREW HOLES -** **TO BE FILLED BEFORE FINISHING** • BUTT JOINT OF PATCH TO SKIN E PLUG PATCH PLYWOOD SKIN • SAW CUT IN DOUBLER **PLYWOOD** **DOUBLER** DIMENSIONS 8 SMALL CIRCULAR PLUG PATCH 2 LARGE CIRCULAR PLUG PATCH 3· 3 (TWO ROWS OF SCREWS AND NAILS REQUIRED FOR LARGE PATCH.) FIGURE 1.18.-Round plug patch assembly. Par 24 ### Page 64 (OCR) AC 43.13-1A 1972 that have been previously made with patch and plate matching. No. 4 roundhead screws are used. Lead holes in the plywood doubler are not necessary. Waxed paper or cellophane between the plate and patch prevents glue from sealing the plate to the patch. No clamps or further pressure need be applied as the nailing strips and screws exert ample pressure. Hot sandbags, however, may be laid over the patch to speed the setting of the glue. Finish in the usual manner. b. Round Plug Patch. The steps in making a round plug patch shown in figure 1.18 are identical with those for making the oval patch except the insertion of the doubler. In using the round patch, where access if from only one side, | Chap 1 Page 23 the round doubler cannot be inserted unless it has been split. 25. FABRIC PATCH. Small holes not exceeding 1 inch in diameter, after being trimmed to a smooth outline, may be repaired by doping a fabric patch on the outside of the plywood skin. The edges of the trimmed hole should first be sealed and the fabric patch should overlap the plywood akin by at least 1 inch. Holes nearer than 1 inch to any frame member, or in the leading edge or frontal area of the fuselage, should not be repaired with fabric patches. **26.-36. RESERVED.** Par 24 ### Page 65 (OCR) Page 24 AC 43.13-1 197 Section 2. FINISHING WOOD STRUCTURES 37. GENERAL, Any repair to spars, ribs, skin surfaces, or other structural parts of the airframe involves finishing as the final step of the job. The time and effort spent during the preparatory phase of the refinishing process will be reflected in the appearance of the finished surface. Adherence to the instructions issued by the finish manufacturer is necessary to obtain the appearance desired and protective characteristics for the product used. 38. PRECAUTIONS TO BE OBSERVED. When making repairs, avoid excessive contamination of surfaces with glue squeezeout at joints and on all surfaces. Excess glue should always be removed before applying finish. Because paints and glues are incompatible, even a slight amount of glue underneath the finish may cause premature deterioration. a. Soiling substances, such as oil and grease, should be removed as completely as possible. Naphtha may be used to sponge off oil and grease. Markings that are made by grease pencils or lumber crayons containing wax are harmful and should be removed, but marks made by ordinary soft graphite pencils and nonblotting stamp pad inks may be safely finished over. All dust, dirt, and other solid particles should be cleaned off. b. Sawdust, shavings, and chips must be removed from enclosed spaces before they are sealed off by replacement of akin. A vacuum cleaner is useful for such cleaning. 6. Since no satisfactory gluable sealer has yet been developed, it is necessary to avoid applying sealer over the areas where glue will be applied. Mark off areas to receive glue, allowing an additional 1/4 inch on each side of the glue area to provide for misalinement when mating the parts. It is preferable to leave some unsealed areas rather than risk weakening the glue joint by accidental overlap of the seale into the glued areas. d. Finish is likely to crack when applied oven fush-driven nails and screws. To avoid this, a strip of tape may be applied over the heads finish is applied. after application of sealer and before the finely e. Fill all holes left from nail-strip gluing or countersunk nails and screws with a wood filler before finishing the surface. It may be necessary to cover with a patching putty the slight depressions left after applying filler, if a completely smooth surface is desired; but, as rule, patching putty may be dispensed wit safety. f. Treat surfaces which are likely to come i contact with fabric during the doping proces with a dope-proof paint, cellophane tape, etc to protect them against the action of the so vents in the dope. 39. FINISHING OF INTERIOR SURFACES. Finig repaired ribs, spars, interior of plywood ski and other internal members, including areas c contact between metal and wood, by applyi at least two coats of spar varnish. Prot built-up box spars and similar closed stru tures on the interior by at least one heavy cor of spar varnish or lionoil. Where better protection is required, as on the surfaces of wheel wells and the bottoms of hulls below the foo boards, an additional coat of aluminized sea consisting of 12 to 16 ounces of alumir paste per gallon of sealer may be applied. 40. FINISHING OF EXTERIOR SURFACES. Exteri surfaces should first be sealed with at least t coats of sealer or spar varnish. The surfa finish should then be completed by the applid tion of enamel, aluminized varnish, or oth special finish as required to duplicate the orig Par 37 Chap ### Page 66 (OCR) AC 43.13-1A 1972 nal finish. If dope or lacquer is used to complete the finish, the sealer coats should be dope-proof. Spar varnish or sealer conforming to Specification MIL-V-6894 is satisfactory. 41. FINISHING OF END GRAIN SURFACES. End grain surfaces, such as edges of plywood skins and holes in spars and other primary structural members, require careful protection. Sand these surfaces smooth. Apply two coats of a highly pigmented sealer, or one coat of wood filler, and one coat of clear sealer to end grain interior surfaces and cut holes. Exterior end grain surfaces (except those covered with doped fabric) require an additional (third) coat of clear sealer. A final coat of aluminized varnish may be applied to end grain surfaces. If the surfaces are to be finished with dope or lacquer, a dope-proof sealer similar to Specification MIL-V-6894 should be used. Exposed end grain includes such surfaces as those around ventholes, inspection holes, fittings, and exposed scarfed or tapered surfaces such as those of tapered blocking. 42. FINISHING WITH FABRIC OR TAPE. To refinish with fabric or tape, it is first necessary to insure that paint has been removed from an Chap 1 Page 25 area greater than that to be covered by the fabric. a. Apply two brush coats of a dope-proof sealer similar to Specification MIL-V-6894, allowing the first coat to dry 2 hours and the second coat at least 6 hours. Follow with one coat of clear dope, and allow it to dry 45 minutes. Apply a second coat of clear dope and lay into the wet film a piece of pinked-edge airplane cloth. All air bubbles should be worked out by brushing to insure maximum adherence. Allow this to dry 45 minutes. Apply one brush coat to insure proper penetration, and at least one spray coat of clear dope, allowing each to dry 45 minutes. The dried spray coat may be sanded with fine sandpaper to obtain a smoother finish. Complete the refinishing of the surface by application of lacquer, enamel, or aluminized varnish as required to match the adjacent area. b. The size of the fabric patch should be such as to extend at least 1/2 inch on each side of any crack or group of cracks, at least 1 inch on each side of & scarfed joint glue line, and at least 2 inches beyond any edge of a skin patch to insure proper adhesion. **43.-53. RESERVED.** Par 40 ### Page 67 (OCR) C 43.13-1A Page 27 972 Chapter 2. AIRCRAFT METAL STRUCTURES Section 1. REFERENCES AND PRECAUTIONARY MEASURES I. REFERENCES. The following chapters of this aC should be referred to when accomplishing repairs to aircraft metal structures: a. Identification and Inspection of Materials. (dentification and inspection of materials should be conducted in accordance with Chaper 7. b. Corrosion Protection. Corrosion protection reatment, cleaners, and paint removing should e accomplished in accordance with Chapter 6. c. Aircraft Hardware. Acceptable means of ttachment are listed in Chapter 5. b. Materials and Construction Techniques. The development of new materials and techniques has made possible the use of control surfaces of less mass weight for a given area than some aircraft of older design. The effect of repair or weight change on the balance and center of gravity is proportionately greater on lighter surfaces than on the older designs. Since control surfaces on some models are balanced for flutter-free operation up to maximum speed for which the aircraft was originally designed, special attention, therefore, must be given to such surfaces relative to the effects of structural repairs and rework on their balance con- 5j. FLUTTER AND VIBRATION PRECAUTIONS. To revent the occurrence of severe vibration or utter of flight control surfaces during flight, recautions must be taken to stay within the esign balance limitations when performing aintenance or repair. a. Balance Changes. The importance of retainr the proper balance and rigidity of aircraft entrol surfaces cannot be underestimated. As reneral rule, repair the control surface in ch a manner that the structure is identical to : original so that the weight distribution is t affected in any way. In order to preclude he occurrence of flutter of the control surface n flight, a degree of static and/or dynamic balnce is established for each model of aircraft. certain conditions, counter-balance rht is added forward of the hinge line to tain balance. Remove or add balance ught only when necessary in accordance h the manufacturer's instructions, or obtain A approval. Flight testing may be required. filure to check and retain control surface bale within the original or maximum allowavalue could result in a serious flight haz- directed to the effect indiscriminate application of extra coats of dope or paint has on the balance of control surfaces. Proper maintenance of control surface balance may require removal of dope or paint, down to the base coat, prior to application of finish coats. Consult the aircraft manufacturer's instructions relative to finishing and balance of control sur- have occurred from unbalanced conditions caused by the collection of water or ice within the surface. Therefore, ventilation and drainage provisions must be checked and retained when maintenance is being done. Certain construction designs do not provide for ventilation and may collect moisture through condensation which will affect balance. In the event this condition is found, refer to the manufacturer's instructions for moisture removal. e. Trim Tab Maintenance. In addition to unbalanced control surface, loose or vibrating trim tabs will increase wear of actuating mechanisms and hinge points which may de- Par 54 ### Page 68 (OCR) velop serious flutter conditions. Most trim tabs are not balanced separately from the control surface. Minimum tab flutter is maintained through rigid actuating mechanisms. Trim tabs and their actuating mechanism are constructed as lightly as possible to keep the weight aft of the hinge line of the control surface as low as possible. Actuating mechanisms are highly susceptible to wear, deformation, and fatigue failures because of the buffeting nature of airflow over the tab mechanism. Trailing-edge play of the tab may increase, through wear of the mechanism, to an unsafe condition. Careful inspection of the tab and its mechanism should be conducted during overhaul and inspection periods. Compared to other systems on the aircraft, only a minor amount of tab-mechanism wear can be tolerated. In the absence of a specified limit for tab trailing edge movement in the manufacturer's manual, it is acceptable to use a limit of 2 1/2% of the chord, measured at the trailing edge of the tab. For example, a tab that has a chord of 4" would have a maximum permissible free play of 4" x .025 or 0.100 inches (total motion up and down) measured at the trailing edge. Correct any free play in excess of this amount. Care must also be exercised during repair or rework to prevent stress concentration points or areas which could increase the fatigue susceptibility of the trim tab system. Par 55 AC 43.13. 56. BRAZING, Brazing may be used for repa to primary aircraft structures only if brazi was originally approved for the particular & plication. Brazing is not suitable for repair welds in steel structures due to lower streng values of the brazed joint as compared welded joints. Brazing may be used in the 1 pair of secondary structures. Due to the large number of brazing allo used, it is difficult to be certain that the mai rial selected for repairing a brazed joint w result in a joint having the same streng characteristics as the original. In cases whe it is necessary to apply copper alloy brazil material, more than once on a steel surfa‹ and particularly if temperatures over 2,000° are reached, there is a possibility that brazi! metal may penetrate between the grains in t steel to an extent that may cause crackin Copper brazing of steel is normally accor plished in a special furnace having a controlk atmosphere, and at a temperature so high th field repairs are seldom feasible. If copp brazing is attempted without a controlled a mosphere, the copper will probably not con pletely flow and fill the joint. Therefore, col per brazing in any other than appropriatel controlled conditions is not recommended. **57-67. RESERVED.** Chap ### Page 69 (OCR) C 43.13-1A Page 29 1972 Section 2. WELDING 8. GENERAL. This section covers weld repairs aircraft and component parts thereof with he exception of welding on powerplants and propellers which is provided in chapters 12 and 14 respectively. Observe the following when using such equipment as tungsten inert gas (TIG), metal inert gas (MIG), plasma arc, shield carbon arc, and oxygen-acetylene gas. a. Equipment Selection. Use the welding equipment manufacturers' information to determine if the equipment will satsify the requirements lor the type of welding operation being underaken. Disregarding such detailed operating intructions may cause substandard welds. For xample, when using TIG equipment, a weld an be contaminated with tungsten if the proper size electrode is not used when welding with lirect current reverse polarity. Another example, the depletion of the inert gas supply below the critical level causes a reduction in the gas flow and will increase the danger of atmospheric contamination. (1) Electric welding equipment versatility requires careful selection of the type current and polarity to be used. Since the composition and thickness of metals are deciding factors, the selection may vary with each specific application. Metals having refractory surface oxide films, i.e., magnesium alloys and aluminum and its alloys, are generally welded with A.C., while D.C. is used for carbon, low alloy, noncorrodible, and heat-resisting steels, copper, etc. General recommendations covering current and polarity are shown in figure 2.1. (2) Oxygen-acetylene gas equipment is suitable for welding most metals. It is not, how- FIGURE 2.1.-Current and polarity selection for inert gas welding. **ALTERNATING DIREO** CURRENT CURREN' With High- Frequency STRAIGHT REVERSE MATERIAL Stabilization Polarity Polarity dagnesium up to 1/8 in. thick 1 N.R. 2 ngnesium above 8/16 in. thick 1 N.R. N.R. agnesium Castings N.R. 2 uminum up to 8/82 in. thick N.R. luminum over 8/82 in. thick N.R. N.R. luminum Castings N.R. N.R. Stainless Steel N.R. Brass Alloys N.R. Silicon Copper N.R. ilver OUUN: N.R. ver Cladding N.R. N.R. d-facing N.R. st Iron N.R. w Carbon Steel, 0.015 to 0.080 in. 2** N.R. w Carbon Steel, 0.030 to 0.125 in. N.R. N.R. ch Carbon Steel, 0.015 to 0.030 in. 2 * 1 N.R. eh Carbon Steel, 0.030 in. and up 2 * 1 N.R. Recommended 2. Acceptable N.R. Not Recommended Where A.C. is recommended as second choice, use approximately 25% higher current than is recommended for DCSP. Do not use A.C. on tightly jigged part. ap 2 Par 68 ### Page 70 (OCR) ever, the best method to use on such materials as stainless steels, magnesium, and aluminum alloys because of base metal oxidization, distortion, and loss of ductility. NoTE: When oxyacetylene is used, all flux must be removed, as it may cause corrosion. b. Accurately Identify the Type of Material to be Repaired. Reference to chapter 7 may provide this information. If positive identification is not possible, contact the aircraft manufacturer or subject the item to a metallurgical laboratory analysis. Before any welding is attempted, carefully consider the weldability of the alloy since all alloys are not readily weldable. The following steels are readily weldable: plain carbon of the 1000 series, nickel steel of the SAE 2800 series, chrome-nickel alloys of the SAE 8100 series, chrome-molybdenum steels of the SAF 4100 series, and low nickel-chromemolybdenum steel of the SAL 8600 series. c. Preparation for Welding. Hold elements to be welded in a welding jig or fixture which is sufficiently rigid to prevent misalinement due to expansion and contraction of the heated material and which positively and accurately position the pieces to be welded. d. Cleaning Prior to Welding. Clean parts to be welded with a wire brush or other suitable methods. When cleaning with a wire brush, do not use a brush of dissimilar metal; for example, brass or bronze on steel. The small deposit left by a brass or bronze brush will materially weaken the weld and may cause cracking or subsequent failure of the weld. In case members were metallized, the surface metal may be removed by careful sandblasting followed by a light buffing with emery cloth. •. Condition of Complete Weld. Make sure that: (1) The finished weld has a smooth seam and is uniform in thickness; (2) The weld metal is tapered smoothly into the base metal; (3) No oxide has formed on the base metal at a distance of more than 1/2 inch from the weld; Par 60 AC 43.13-1 197 (4) The weld shows no signs of blowholes porosity, or projecting globules; (5) The base metal shows no signs of pit ting, burning, cracking, or distortion; (6) The depth of penetration insures fusion of base metal and filler rod; and (7) Welding scale is removed by wire brush ing or sandblasting. f. Practices to Guard Against. Do not file weldi in an effort to make a smooth-appearing job, a such treatment causes a loss of strength. D not fill welds with solder, brazing metal, ol any other filler. When it is necessary to rewel a joint which was previously welded, remov all old weld material before rewelding. Avoic welding over a weld because reheating may cause the material to lose its strength and be come brittle. Never weld a joint which hai been previously brazed. g. Torch Size (Oxyacetylene Welding). The sizi of the torch tip depends upon the thickness of the material to be welded. Commonly user sizes proved satisfactory by experience are: Thickness of steel Diameter of Drill size (in Inches) hole in tip 0.015 to 0.031 0.026 71 0.031 to 0.065 .031 68 0.065 to 0.125 .037 63 0.125 to 0.188 .042 58 0.188 to 0.250 .055 Б4 0.250 to 0.875 .067 51 h. Welding Rods and Electrodes. Use weldin rods and electrodes that are compatible wit the materials to be welded. Welding rods anc electrodes for various applications have special properties suitable for the application intended. Figure 2.8 shows the allowable strength for the weld metal. Figure 2.2 lists specifications for corresponding metals, ar the applicable specification will identify t use of A.C. or D.C., straight or reverse polal 1. Rosette Welds. Rosette welds are generali employed to fuse an inner reinforcing tube (liner) with the outer member. Where a ro sette weld is used, drill the hole, in the outside tube only, of a sufficient size to insure fusion o Chap ### Page 71 (OCR) FIGURE 2.2.-Specifications guide for welding metals. Material To Electrie Tungaten Inert Metal Inert Shield Be Welded Welding Gas (TIG) Ga• (HIG) Carbon ARC Acetviene Medium & High MIL-1-22200/1 MIL-E-29765/1B MIL-E-23765/1B MIL-R-908A Tensile Steel. MIL-9018 Class Class NONE Class 1 Stress Relieved MIL-E-705-2 MIL-E-705-4 High Carbon Steel MIL-E-7018 MIL-E-23765/1B MIL-D-23765/1B MIL-R-CUS1-A MIL-R-908A MIL-E-22200 Class Class Class 1 MIL-E-705-2 MIL-E-705-6 Low and Medium MIL-E-15599-C: MIL-E-23765/1B MIL-E-23765/1B MIL-R-CUSI-A Carbon Steel Class Class MIL-R-5632 MIL-E-705-2 MIL-E-705-6 Class 1 Aluminum and MIL-E-15997A QQR-566 QQR-566 QQR-566° QQR-566 * Aluminum Alloy MIL-E-16053-K' MIL-E-16053-K' Stainless Alloys MIL-E-6844 MIL-R-5031B' MIL-R-5031B* MIL-R-5081B* MIL-R-5081B* M-L-E-22200• Copper and Nickel NONE QQR-571A * MIL-E-21659* QQR-571A * QQR-571A • Alloys Magnesium NONE MIL-R-6944 MIL-W-18326 NONE MIL-R-6944 Alloys Specify Alloy Specify Alloy Specify Alloy Hard Surface MIL-E-19141 MIL-E-19141 MIL-E-19141 MIL-E-19141 MIL-19141 Filler (Refer to A.W.S (Refer to A.W.S (Refer to A.W.S. (Refer to A.W.S. (Refer to A.W.S. A5.18 or ASTM-A899) А5.18 or А5.18 ог Аб.13 ог Аб.18 от ASTM-A399) ASTM-A399) ASTM-A899) ASTM-A899) ' Specify MIL Type 6010, 6011, 6012, 6013, 6020, 6024, 6027-Same as A.W.S. Number for type. A.W.S. American Welding Society * Specify Aluminum Alloy by Aluminum Assoc. Number e.g.: Type 1100, 5083, etc. ASTM American Society for Testing Materials • Specify Alloy by A.IS.I. Number e.g.: 304, 316, 842, etc. A.I.S.I.. American Iron and Steel Institute (SAE). * Specify type per A.W.S. A5.6-66 or ASTM B-225-66. ### Page 72 (OCR) Page 32 AC 43.13-14 197" FIGURE 2.3.-Strengths of welded joints. Materle! Heat treatment Welding rod Ultimate Ultimate subsequent to or electrode atress 11 tenaile welding shear × 1000 atresa X 1000 Carbon and MIL-R-5632, Class I 82 51 Alloy steels None MIL-E-15599, Classes E-6010 & E-6013 32 61 Alloy steels None MIL-R-5632, Class 2 48 72 Alloy steels Stress relieved MIL-E-6848, Class 10018 50 85 MIL-E-18088, Class E-10015 & E-10016 Alloy steels Stress relieved MIL-E-18038, Class 60 100 E-12015 & E-12016 the inner tube. A hole diameter of approximately one-fourth the tube diameter of the outer tube serves adequately for this purpose. In cases of tight-fitting sleeves or inner liners, the rosettes may be omitted. I. Heat-Treated Members. Certain structural parts may be heat-treated and therefore could require special handling. In general, the more responsive an alloy steel is to heat treatment, the less suitable it is for welding because of its tendency to become brittle and lose its ductility in the welded area. Weld the members which depend on heat treatment for their original physical properties, using a welding rod suitable for producing heat-treated values comparable to those of the original members (see paragraph 68h). After welding, reheat-treat such members to the manufacturer's specifica- **69. STEEL PARTS NOT TO BE WELDED.** a. Brace Wires and Cables. Do not weld airplane parts whose proper function depends upon strength properties developed by coldworking. Among parts in this classification are streamlined wires and cables. b. Brazed and Soldered Parts. Do not weld brazed or soldered parts as the brazing mixture or solder will penetrate the hot steel and weaken it. c. Alloy Steel Parts. Do not weld alloy steel parts such as aircraft bolts, turnbuckle ends, etc., which have been heat-treated to improve their mechanical properties. Par 68 **70. REPAIR OF TUBULAR MEMBERS.** a. Inspection. Prior to repairing tubular members, carefully examine the structure surrounding any visible damage to insure that n secondary damage remains undetected. Second. ary damage may be produced in some structur remote from the location of the primary dam age by the transmission of the damaging loa along the tube. Damage of this nature usuall occurs where the most abrupt change in dire tion of load travel is experienced. If this dam. age remains undetected, subsequent norma loads may cause failure of the part. b. Location and Alinement of Welds. Unles otherwise noted, welded steel tubing may spliced or repaired at any joint along the leng of the tube. Pay particular attention to th proper fit and alinement to avoid distortion. 6. Members Dented at a Cluster. Repair der at a steel-tube cluster-joint by welding a sp cially formed steel patch plate over the dented area and surrounding tubes, as shown in figure 2.4. To prepare the patch plate, cut a section of steel sheet of the same material and thickne as the heaviest tube damaged. Trim the re forcement plate so that the fingers extend o the tubes a minimum of 1.5 times the resp tive tube diameter as shown in the figu Remove all the existing finish on the damag cluster-joint area to be covered by the reinfor ment plate. The reinforcement plate may formed before any welding is attempted, or may be cut and tack-welded to one or more the tubes in the cluster-joint, then heated ar Chap ### Page 73 (OCR) AC 43.13-1A Page 33 1972 LONGERON DENTED AT A STATION. Lite 706 **THICKNESS OF PATCH-PLATE** **SAME AS LONGERON** **THICKNESS, PATCH-PLATE BEFORE FORMING AND WELDING.** **PATCH-PLATE FORMED AND WELDED TO TUBES.** FIOURE 2.4.-Members dented at a cluater. Chap 2 Par 70 ### Page 74 (OCR) Page 34 AC 43.13-1 197 DENTED OR BENT TUBE. **CRACKED TUBE** REINFORCEMENT TUBE SPLIT NOTE: **LOCALLY DENTED OR** **TINES MEMBERORSHOULD 15811 IN CLAMP.** **REINFORCEMENT LEAST THE SAME GAUGE AS TUBE BEING REPAIRED.** **WELD** **AS ALTERNATIVE TO SPLIT** A TWO PIECE REIN- FORCEMENT ‚MAY **FORMED FROM STEEL** **SHEET OF THE SAME MAT-** **ERIAL AND AT LEAST THE** **SAME GAUGE AS AGED TUBE FOUR ENDS AND THE DAM- ROSETTE WELDS AS SHOWN.** FIGURE 2.0.-Members dented in a bay--repairs by welded sleeve. Par 70 Chap ### Page 75 (OCR) AC 43.13-1A 1972 formed around the joint to produce a smooth contour. Apply sufficient heat to the plate while forming so that there is generally a gap of no more than 1/16 inch from the contour of the joint to the plate. In this operation avoid unnecessary heating, and exercise care to prevent damage at the point of the angle formed by any two adjacent fingers of the plate. After the plate is formed and tack-welded to the clusterjoint, weld all the plate edges to the clusterjoint. d. Members Dented in a Bay. Repair dented, bent, cracked, or otherwise damaged tubular members by using a split-sleeve reinforcement; carefully straighten the damaged member; and in the case of cracks, drill No. 40 (0.098) stopholes at the ends of the crack. 71. REPAIR BY WELDED SLEEVE. This repair is outlined in figure 2.5. Select a length of steel tube sleeve having an inside diameter approximately equal to the outside diameter of the damaged tube and of the same material, and at least the same wall thickness. Diagonally cut Ithe sleeve reinforcement at a 30° angle on both ends so that the minimum distance of the sleeve from the edge of the crack or dent is not less than 11/2 times the diameter of the damaged tube. Cut through the entire length of the reinforcement sleeve, and separate the halfections of the sleeve. Clamp the two sleeve sections to the proper positions on the affected reas of the original tube. Weld the reinforce- Page 35 ment sleeve along the length of the two sides, and weld both ends of the sleeve to the damaged tube as shown in the figure. The filling of dents or cracks with welding rod in lieu of reinforcing the member is not acceptable. 72. REPAIR BY BOLTED SLEEVE, Do not use bolted sleeve repairs on welded steel tube structure unless specifically authorized by the manufacturer or the FAA. The tube area removed by the boltholes in this types of repair may prove critical. 73. WELDED-PATCH REPAIR. Dents or holes in tubing may be repaired by a welded patch of the same material and one gauge thicker, as shown in figure 2.6 provided: a. Dented Tubing. (1) Dents are not deeper than 1/10 of tube diameter, do not involve more than 1/4 of the tube circumference, and are not longer than tube diameter. (2) Dents are free from cracks, abrasions, and sharp corners. (3) The dented tubing can be substantially reformed without cracking before application of the patch. b. Punctured Tubing. Holes are not longer than tube diameter and involve not more than 1/4 of tube circumference. * TUBE CIRCUM. -NOT GREATER THAN "A" (MAX.) • 2A - - DAMAGE FIGURE 2.6.-Welded patch repair. Chap 2 Far 70 ### Page 76 (OCR) AC 43.13-14 1971 ALLOW 1/8" GAP BETWEEN ORIGINAL AND ROSETTE WELD REPLAGEMENT TUBES FOR WELDING. **INSIDE** **ORIGINAL TUBE ISLEEVE TUBE** - UZA **REPLACEMENT TUBE** **ROSETTE WELD FORVIPELVISAP** **INSIDE SLEEVE TUBE** ORIGINAL TUBE FIGURE 2.7.-Splicing by inner sleeve method. **74. SPLICING TUBING BY INNER SLEEVE METHOD.** If the damage to a structural tube is such that a partial replacement of the tube is necessary, the inner sleeve splice shown in figure 2.7 is recommended, especially where a smooth tube surface is desired. Make a diagonal cut when removing the damaged portion of the tube and remove the burr from the edges of the cut by filing or similar means. Diagonally cut a replacement steel tube of the same material and diameter, and at least the same wall thickness, to match the length of the removed portion of the damaged tube. At each end of the replacement tube allow a 1/8-inch gap from the diago- Par 74 nal cuts to the stubs of the original tube. Select a length of steel tubing of the same material, and at least the same wall thickness, and of an outside diameter, equal to the inside diameter of the damaged tube. Fit this innersleeve tube material snugly within the original tube, wit a maximum diameter difference of 1/16 inc From this inner-sleeve tube material cut 2 se tions of tubing, each of such a length that th ends of the inner sleeve will be a minimum dis tance of 1 1/2-tube diameters from the neares end of the diagonal cut. If the inner sleeve fits very tightly in the replacement tube, chill the sleeve with dry ice or Chap 3 ### Page 77 (OCR) I€ 43.13-1A Page 37 972 ,30° SLEEVE **ROSETTE WELDS MAY - WELD** **BE OMITTED WHEN SLEEVES FIT TIGHTLY.** - REPLAGEMENT TUBE **ORIGINAL TUBE** **FISH-MOUTH SLEEVE** - UZA **ROSETTE WELDS MAY WELD** **BE OMITTED WHEN SLEEVES FIT TIGHTLY.** **LREPLACEMENT TUBE** **ORIGINAL TUBE** **FOUR ROSETTE - WELD HERE FIRST WELDS** ORIGINAL TUBE -1/8" GAP FOR WELDING ALTERNATIVE SPLIT SLEEVE SPLICE ~ IF OUTSIDE DIAMETER OF ORIGINAL TUBE IS LESS THAN I INCH, SPLIT SLEEVE MAY BE MADE FROM STEEL TUBE OR SHEET STEEL. USE SAME MATERIAL OF AT LEAST THE SAME GAUGE, FOR ORIGINAL TUBE DIAMETERS OF 1 INCH AND OVER, USE SHEET STEEL ONLY. FIGURE 2.8.-Splicing by outer sleeve method-replacement by welded outside sleeve. Par 74 ### Page 78 (OCR) in cold water. If this is insufficient, polish down the diameter of the sleeve with emery cloth. Weld the inner sleeve to the tube stubs through the 1/8-inch gap, forming a weld bead over the gap. **75. SPLICING TUBING BY OUTER SLEEVE METHOD.** If partial replacement of a tube is necessary, make the outer sleeve splice using a replacement tube of the same diameter. Since the outer sleeve splice requires the greatest amount of welding, it should be used only when the other splicing methods are not suitable. Information on the replacement by use of the welded outside sleeve method is given in figures 2.8 and 2.9. Remove the damaged section of a tube utilizing a 90° cut. Cut a replacement steel tube of the same material, diameter, and at least the same wall thickness to match the length of the removed portion of the damaged tube. This replacement tube must bear against the stubs of the original tube with a total tolerance not to AC 43.13- 19 exceed 1/82 inch. The outer sleeve tube mat rial selected must be of the same material an at least the same wall thickness as the origin tube. The clearance between inside diameter c the sleeve and the outside diameter of the orig inal tube may not exceed 1/16 inch. From thi outer sleeve tube material, cut diagonally (o fishmouth) 2 sections of tubing, each of suc length that the nearest end of the outer sleev is a minimum distance of 1 1/2-tube diameter from the end of the cut on the original tube Use a fishmouth sleeve wherever possible. Re move the burr from the edges of the sleeves replacement tube, and the original tube stubs Slip the two sleeves over the replacement tube aline the replacement tube with the origina tube stubs, and slip the sleeves out over thi center of each joint. Adjust the sleeves to sui the area and to provide maximum reinforce ment. Tackweld the two sleeves to the replace ment tube in two places before welding. Apply a uniform weld around both ends of one of the reinforcement sleeves and allow the weld t cool; then, weld around both ends of the r < ORIGINAL TUBE **SLEEVE SLEEVE** **TUBE TUBE** ROSETTE ORIGINAL TUBE WELD WELD ORIGINAL TUBE 1/2 A SLEEVE Ve" GAF ALLOW REPLACEMENT TUBE SLEEVE TUBE **TUBE BETWEEN SLEEVES** FOR WELDING, FIGURE 2.9.-Tube replacement at a station by welded outer sleeves. Far 74 ### Page 79 (OCR) АС 43.73-1А Page 41 1972 WELD SPACER APPROXIMATELY "A" LONG WELDED ON **INSIDE TUBE** •ЗА~ (А) (B) GENERAL NOTE:- **ALL ROSETTE WELDS (C) TO BE & DIAMETER OF THE ORIGINAL TUBES.** GAP 1" (D) FIGURE 2.11.-Repairs at built-in fuselage fittings. Par 79 ### Page 80 (OCR) Page 42 АС 43.13-1А 1972 GAP I" 30 WELD mh ABC **FORM INSIDE TUBE TO FIT. AT LEAST - 3C - AT LEAST** **A - SLOT WIDTH (ORIGINAL TUBE) 8 - OUTSIDE DIAMETER (INSERT TUBE)** C - STREAMLINE TUBE LENGTH OF MAJOR AXIS S.L. SIZE D 9/16 1.340" • 496" 1- 4 11/16 1.670 • 619 1-2 2.005 . 743 1- 4 2.339 .867 2 2,670 991 2-¼ 3.008 1.115 г- ½ ı- ₴ 3.342 1.239 **ROUND INSERT TUBE (B) SHOULD BE AT LEAST OF SAME** **MATERIAL AND ONE GAUGE THICKER THAN ORIGINAL STREAMLINE TUBE (C).** FIGURE 2.12.-Streamline tube splice using round tube (applicable to landing gears). Par 79 ### Page 81 (OCR) 1972 AC 43.13-TA **OPEN OUT OPEN OUT TO FIT ORIGIN ORIGINAL TUBE AND WELD.** OUTSIDE SLEEVE MAY BE OF SAME SL. TUBING AS ORIGINAL OR USE AT LEAST S.L. TUBING OF AT LEAST THE C- SAME GAUGE. **AT LEAST WELD** **A - MINIMUM LENGTH OF SLEEVE 9- STREAMLINE TUBE LENGTH OF MINOR AXIS** - STREAMLINE TUBE LENGTH Of MAJOR AXIS S. L. SIZE 7.324" .572" 1.340" 1-¼ 9.128 .714 1.670 1-½ 10.960 .858 2.005 1- ¾ 12.784 1.000 2.339 2 14.594 1.144 2.670 2-4% 16.442 1.286 3.008 2 - ½ 18.268 1.430 3.342 FIGURE 2.13.—Streamline tube splice using split aleeve (applicable to wing and tail surface brace struts and other members). ### Page 82 (OCR) Page 44 AC 43.13-1A 1972 *B-* WELD GAP $" SIZE OF ROSETTES = 14G. DRILL OUTSIDE **TUBES ONLY B C** SEA-SEE * B- -2A-→ IŁA L = MAXIMUM INSERT LENGTH SAW .08 C OFF OF T.E. AND WELD INSERT TUBE IS OF SAME STREAMLINE TUBING AS ORIGINAL. **A IS E B 8 IS MINOR AXIS LENGTH OF ORIGINAL STREAMLINE TUBE** C IS MAJOR AXIS LENGTH OF ORIGINAL STREAMLINE TUBE S. L. SIZE .382 _ 572 1.340 5.16 1¼ ,476 • 714 1.670 6.43 1/½ . 572 1858 2.005 7.72 1 ¾ .667 1.000 2.339 9.00 2 .763 1.144 2.670 10.30 2 '¼ .858 I.286 3.008 11.58 2 V/½ ,954 1.430 3.342 12.88 FIGURE 2.14.-Streamline tube splice using split insert (applicable to landing gears). Par 79 ### Page 83 (OCR) kc 43.13-14 1972 c. Axle Assemblies. Representative types of repairable and nonrepairable landing gear axle assemblies are shown in figure 2.16. The types as shown in A, B, and C of this figure are formed from steel tubing and may be repaired by the applicable method shown in figures 2.5 thru 2.15 in this Advisory Circular. However, it will always be necessary to ascertain whether or not the members are heat treated. The axle assembly as shown in figure 2.16(D) is, in general, of a nonrepairable type for the following reasons: (1) The axle stub is usually made from a highly heat-treated nickel alloy steel and carefully machined to close tolerances. These stubs are usually replaceable and should be replaced if damaged. (2) The oleo portion of the structure is generally heat-treated after welding and is perfectly machined to assure proper functioning of the shock absorber. These parts would be istorted by welding after machining. J. BUILT-UP TUBULAR WING OR TAIL SURFACE ARS. Repair built-up tubular wing or tail surce spars by using any of the applicable plices and methods of repair shown in figures 5 through 2.15 provided the spars are not eat-treated. In the case of heat-treated spars, he entire spar assembly would have to be reeat-treated to the manufacturer's specificaons after completion of the repair. In genal, this will be found less practicable than placing the spar with one furnished by the anufacturer. WING AND TAIL BRACE STRUTS. In general, 'ill be found advantageous to replace damwing-brace struts made either from ded or streamlined tubing with new mempurchased from the original manufac- However, there is no objection from an rthiness point of view to repairing such ers in a proper manner. An acceptable d, in case streamlined tubing is used, will nd in figure 2.18. Repair similar memnade of round tubes using a standard as shown in figures 2.5, 2.7, or 2.8. a. Location of Splices. Steel brace struts may be spliced at any point along the length of the strut, provided the splice does not overlap part of an end fitting. The jury strut attachment is not considered an end fitting; therefore, a splice may be made at this point. The repair procedure and workmanship should be such as to minimize distortion due to welding and the necessity for subsequent straightening operations. Observe every repaired strut carefully during initial flights to ascertain that the vibration characteristics of the strut and attaching components are not adversely affected by the repair. A wide range of speed and enginepower combination must be covered during this check. (1) Fit and Alignment. When making repairs to wing and tail surface brace members, pay particular attention to proper fit and alinement to avoid distortion. 82, REPAIRS TO WELDED PARTS. Repairs to welded assemblies may be made by: a. Repiacing welded joints-cutting out the welded joint and replacing it with one properly gusseted, or b. Replacing weld deposit-chipping out the metal deposited by the weld process and rewelding after properly reinforcing the joint by means of inserts or external gussets. 83. STAINLESS STEEL STRUCTURE. Repair structural components made from stainless steel, particularly the "18-8" variety (18 percent chromiun, 8 percent nickel), joined by spot welding in accordance with the instructions furnished by the manufacturer. Substitution of bolted or riveted connections for spot-welded joints are to be specifically approved by an authorized representative of the FAA. a. Secondary Structural and Nonstructural Elements. Repair such elements as tip-bows or leading and trailing edge tip-strips of wing-andcontrol surfaces by soldering with a 50-50 lead-tin solder or a 60-40 alloy of these metals. For best results use a flux of phosphoric acid (syrup). Since the purpose of flux is to attack the metal so that the soldering will be effec- Par 79 ### Page 84 (OCR) Page 48 AC 43.13-: 19: tive, remove excess flux by washing the joint. welded seams in boat hulls, fuel tanks, etc., in Due to the high-heat conductivity of the stain- similar manner. less steel, use a soldering iron large enough to do the work properly. Repair leaky spot- 84.~94. RESERVED. Par 83 ### Page 85 (OCR) lAC 43.13-1A Page 49 1972 Section 3. METAL REPAIR PROCEDURES 95. RIVETED OR BOLTED STEEL TRUSS-TYPE STRUC- TURES. Repairs to riveted or bolted steel trusstype structures should be made, employing the general principles outlined in the following paragraphs on aluminum alloy structures. Methods for repair of vital members should be specifically approved by a representative of the Federal Aviation Administration. 96. ALUMINUM ALLOY STRUCTURES. Extensive repairs to damaged stressed skin on monocoque types of aluminum alloy structures should preferably be made in accordance with specific recommendations of the manufacturer of the aircraft. In many cases, repair parts, joints, or reinforcements can be designed, and proof of adequate strength shown, without the calculaion of the design loads and stresses, by properly considering the material and dimensions f the original parts and the riveted attachments. Examples illustrating the principles of this method as applied to typical repairs are given in this handbook, or may be found in extbooks on metal structures. An important point to bear in mind in making repairs on nonocoque structures is that a repaired part must be as strong as the original with respect to all types of loads and general rigidity. a. Use of Annealed Alloys for Structural Parts. he use of annealed 2017 or 2024 for any ructural repair of an aircraft is not consided satisfactory because of its poor corrosionisting properties. Hygroscopic Materials Improperly Moisturefed. The use of hygroscopic materials imberly moisture-proofed, such as impregd fabrics, or leather, and the like, to effect r-tight joints and seams is not an acceptacactice. rilling Oversized Holes. Avoid drilling zed holes or otherwise decreasing the effective tensile areas of wing-spar capstrips, wing, fuselage, or fin longitudinal stringers, or other highly stressed tensile members. Make all repairs or reinforcements to such members in accordance with factory recommendations or with the specific approval or a representative of the Federal Aviation Administration. d. Disassembly Prior to Repairing. If the parts to be removed are essential to the rigidity of the complete structure, support the remaining structure prior to disassembly in such a manner as to prevent distortion and permanent damage to the remainder of the structure. When rivets are to be removed, weaken the rivet head by drilling. Use a drill of the same size as the rivet. Drilling must be exact center and to the base of the head only. After drilling, break off the head with a pin punch and carefully drive out the shank. Removal of rivet heads with a cold chisel and hammer is not recommended because skin damage and distorted rivet holes will probably result. Care must also be taken whenever screws must be removed for disassembly or removal of stress plates, access plates, fillets, etc., to avoid damage to adjoining structure. When properly used, impact wrenches can be effective tools for removal of screws; however, damage to adjoining structure may result from excessive vertical loads being applied through the screw axis. Excessive loads are usually related to improperly adjusted impact tools or attempting to remove screws that have seized from corrosion. Remove seized screw by drilling and use of a screw extractor. Structural cracks may appear in the doubler or tang that runs parallel to the line of anchor or plate nuts installed for securing access doors or plates. Inspect rivet joints adjacent to damaged structure for partial failure (slippage) by removing one or more rivets to see if holes are elongated or the rivets have started to shear. Por 95 ### Page 86 (OCR) 97. SELECTION OF ALUMINUM FOR REPLACEMENT PARTS. In selecting the alloy, it is usually satisfactory to use 2024 in place of 2017 since the former is stronger. Hence, it will not be permissible to replace 2024 by 2017 unless the deficiency in strength of the latter material is compensated by an increase in material thickness, or the structural strength is substantiated by tests or analysis. Information on the comparative strength properties of these alloys, as well as 2014, 6061, 7075, etc., is contained in MIL-HDBK-5, Metallic Materials and Elements for Flight Vehicle Structure. The choice of temper depends upon the severity of the subsequent forming operations. Parts having single curvature and straight bend lines with a large bend radius may be advantageously formed from heat-treated material, while a part, such as a fuselage frame, would АС 43.13-11 197: have to be formed from soft, annealed sheet and heat-treated after forming. Make sure sheet metal parts which are to be left unpainted are made of clad (aluminum coated) material. Make sure all sheet material and finished parts are free from cracks, scratches, kinks, tool marks, corrosion pits, and other defects which may be factors in subsequent failure. a. Forming Sheet Metal Parts. Bend lines should preferably be made to lie at an angle to the grain of the metal (preferably 90°). Before bending, smooth all rough edges, remove burrs, and drill relief holes at the ends of bend bure, and lines and at corners, to prevent cracks from starting. For material in the heat-treated condition, the bend radius should be large. (See figure 2.17 for recommended bend radii.) FIGURE 2.17.-Recommended radii for 90° bends in aluminum alloys. Approximate sheet thickness (t) (inch) Alloy and temper 0.016 0.082 0.064 0.128 0.182 0.268 2024-01 0-1t O-1t 0-1t 0-1t 2024-T8** 133t-St 2t 4t 8t-5t 4t-6t 4t-8t 2024-T6' 2t 4t 8t-5t 3t-5t 4t-6t 5t-7t 5052-0 O-It 0-1t O-It 5052-H82 ½t-It ½ t-1½t est-1%t 5052-H84 3st-1½t 1½t-2½t 1½t-2½t 5052-H86 0-1t % t-1½t 1t-2t 1½t-St 2t-4t 5052-F88 36t-1t It-2t 1½t-3t 2t-1t 8₺-бt 6061-0 0-1t O-1t 0-1t O-1t 6061-T4 0-It 0-1t 1t-2t 1½t-St 6061-T6 0-1t ½t-1%t 1t-2t 1½t-St 2t-4t 7075-0 0-It 0-1t ½t-1½t 1t-21 7075-T6' 2t-4t 3t-5t 4t-6t St-7t ' Alclad sheet may be bent over slight smaller radif than the corresponding tempers of uncoated alloy. ' Immediately after quenching, this alloy may be formed over appreciably smaller radi.. 98. HEAT TREATMENT OF ALUMINUM ALLOY PARTS. All structural aluminum alloy parts are to be heat-treated in accordance with the heattreatment instructions issued by the manufacturers of the materials. If the heat-treatment produces warping, straighten the parts immediately after quenching. Heat-treat riveted parts before riveting, to preclude warping and corrosion. When riveted assemblies are heated in a salt bath, the salt cannot be entirely Par 97 washed out of the crevices, thus causing cor sion. a. Quenching in Hot Water or Air. Quench terial from the solution heat-treating temp ture as rapidly as possible, with a minir delay after removal from the furnace. Que ing in cold water is preferred, although drastic chilling (hot or boiling water blast) is sometimes employed for bulk ### Page 87 (OCR) AC 43.13-1A 1972 tions, such as forgings, to minimize quenching stresses. b. Transferring Too Slowly From Heat-Treatment Medium to Quench Tank, Transfer of 2017 alloys from the heat-treatment medium to the quenchtank should be accomplished as quickly as possible. An elapsed time of 10 to 15 seconds will, in many cases, result in noticeably impaired corrosion resistance. c. Keheating af Temperatures Above Boiling Wafer. Reheating of 2017 and 2024 alloys at temperatures above that of boiling water after heat treatment, and the baking of primers at temperatures above that of boiling water, will not be considered acceptable without subsequent complete and correct heat treatment, as such practice tends to impair the original heat treatment. **99. RIVETING.** a. Identification of Rivet Material. Identification of rivet material is contained in Chapter b. Replacement of Aluminum Alloy Rivets. All rotruding head rivets (roundhead, flathead, nd brazier head) may be replaced by rivets of he same type or by AN-470 Universal head ivets. Use flushhead rivets to replace flushbead rivets. c. Replacement Rivet Size and Strength. Relace rivets with those of the same size and trength whenever possible. If the rivet hole pecomes enlarged, deformed, or otherwise damaged, drill or ream the hole for the next large ize rivet; however, make sure that the edge stance and spacings are not less than miniams listed in the next paragraph. Rivets may be replaced by a type having lower ength properties, unless the lower strength dequately compensated by an increase in or a greater number of rivets. Replacement Rivet-Edge Distances and ngs for Sheet Joints. Rivet-edge distance is ed as the distance from the center of the hole to the nearest edge of the sheet. spacing is the distance from the center rivet hole to the center of the adjacent Page 51 rivet hole. The following prescribes the minimum edge distance and spacing: (1) Single row-edge distance not less than 2 times the diameter of the rivet and spacing not less than 3 times the diameter of the rivet. (2) Double row-edge distance and spacing not less than the minimums shown in figure 2.18. (3) Triple or multiple rows-edge distance and spacing not less than the minimums shown in figure 2.18. •. Use of 2117-T3 Aluminum Alloy Replacement Rivets. It is acceptable to replace 2017-18 rivets of 8/16-inch diameter or less, and 2024-T4 rivets of 5/82-inch diameter or less with 2117-T8 rivets for general repairs, provided the replacement rivets are 1/82 inch greater in diameter than the rivets they replace and the edge distances and spacing are not less than the minimums listed in the preceding paragraph. f. Driving of Rivels. The 2117 rivets may be driven in the condition received, but 2017 rivets above 8/16 inch in diameter and all 2024 rivets are to be kept either refrigerated in the "quenched" condition until driven, or be reheat-treated just prior to driving, as they would otherwise be too hard for satisfactory riveting. Dimensions for formed flat rivet heads are shown in figure 2.19, together with commonly found rivet imperfections. g. Blind-Type and Hollow Rivets. Do not substitute hollow rivets for solid rivets in loadcarrying members without specific approval of the application by a representative of the Federal Aviation Administration. Blind rivets may be used in blind locations in accordance with the conditions listed in Chapter 5, provided the edge distances and spacings are not less than the minimum listed in paragraph 99d. h. New and Revised Rivet Patterns. Design a new or revised rivet pattern for the strength required in accordance with the specific instructions in paragraphs 100e and 100j. A general rule for the diameter of rivets used to join dural sheets is to use a diameter ap- Par 98 ### Page 88 (OCR) Page 52 AC 43.13-14 1972 STRENGTH THROUGH THIS SECTION IS 78% OF SHEET WITHOUT HOLES. D* DIAMETER OF 4D RIVET (PREFER- ABLY ABOUT 3X **THICKNESS OF** THICKER SHEET) 2D LADN →20 (A) DOUBLE ROW STRENGTH THROUGH THIS SECTION IS 83% OF SHEET WITHOUT HOLES. 6D 30 20- *20 (B) TRIPLE OR MULTIPLE ROWS FIOURE 2.18.-Rivet hole spacing and edge distance for single-lap sheet splices. Par 99 ### Page 89 (OCR) AC 43.13-1A Page 53 1972 -1žD 650 (ID MIN.) + 1¿0→ (MIN.) (gO MIN.) (8) DIMENSIONS FOR FORMED RIVET HEADS, **INCORRECT CORRECT SHAPE FOR** **RIVET OR SEE SET.** **BUCKING BAR** (6) RIVETING TOOLS OR DOLLY. **RIVET DRIVEN AT SLANT RIVET DRIVEN RIVET FLAT ON ONE** **CORRECTLY, DOLLY SIDE OR DOLLY HELD HEAD AT SLANT. FLAT.** BODY OF RIVET TOO RIVET NOT PULLED RIVET TIGHT, PLATES SHORT. CLOSING HEAD TIGHT, CLINCHES BULGED ON ACCOUNT SHAPED TOO MUCH BETWEENPLATES, OF POOR FIT. **WITH SNAP DIE. CLOSING HEAD TOO FLAT.** **RIVETED TOO MUCH. RIVET BODY CLINCHED TOO MUCH, HEAD CRACKED.** **PLATES CLINCHED AT RIVET MATERIAL TOO AND DRIVEN APART. HARD WHEN RIVETING TOOL FORMED.** DAMAGED PLATE 1C) RIVET, IMPERFECTIONS: FIGURE 2.19.-Riveting practice and rivet imperfections. Par 99 ### Page 90 (OCR) proximately 8 times the thickness of the thicker sheet. Do not use rivets where they would be placed in tension tending to pull the heads off. Back up a lap joint of thin sheets by a stiffening section. 100. REPAIR METHODS AND PRECAUTIONS FOR ALUMINUM STRUCTURE. Carefully examine all adjacent rivets after the repair or alteration is finished to ascertain that they have not been harmed by operations in adjacent areas. Drill rivet holes round, straight, and free from cracks. The rivet-set used in driving the rivets must be cupped slightly flatter than the rivethead as shown in figure 2.19. Rivets are to be driven straight and tight, but not overdriven or driven while too hard, since the finished rivet must be free from cracks. Information on special methods of riveting, such as flush riveting, usually may be obtained from manufacturer's service manuals. a. Splicing of Tubes. Round or streamline tubular members may be repaired by splicing as shown in figure 2.20. Splices in struts that overlap fittings are not acceptable. When solid rivets go completely through hollow tubes, their diameter must be at least one-eighth of the outside diameter of the outer tube. Rivets which are loaded in shear should be hammered only enough to form a small head, and no attempt made to form the standard roundhead. The amount of hammering required to form the standard roundhead often causes the rivet to buckle inside the tube. Satisfactory rivetheads may be produced in such installations by spinning, if the proper equipment is available. (Correct and incorrect examples of this type of rivet applications are incorporated in figure 2.20.) b. Repairs to Aluminum Alloy Members, Make repairs to aluminum alloy members with the same material or with suitable material of higher strength. The 7075 alloy has greater tensile strength than other commonly used aluminum alloys such as 2014 and 2024, but is subject to somewhat greater notch sensitivity. In order to take advantage of its strength characteristics, pay particular attention to de- Par 99 AC 43.13-11 1971 sign of parts to avoid notches, small radil large or rapid changes in cross sectional areas In fabrication, exercise caution to avoid proc essing and handling defects, such as machine marks, nicka, dents, burrs, scratches, and forming cracks. Cold straightening or forming of 7075-16 can cause cracking; hence, it may be advisable to limit this processing to minor cold straightening. c. Wing and Tail Surface Ribs. Damaged aluminum alloy ribs either of the stamped sheetmetal type or the built-up type employing special sections, square or round tubing, may be repaired by the addition of suitable reinforcement. (Acceptable methods of repair are shown in figures 2.21 and 2.22.) These examples deal with types of ribs commonly found in small and medium aircraft. Repair schemes developed by the aircraft manufacturer are acceptable, and any other methods of reinforcement will be specifically approved by a representative of the FAA. (1) Trailing and Leading Edges and Tip Strips Repairs to wing and control surface trailin and leading edges and tip strips should made by properly executed and reinforce splices. Acceptable methods of trailing edge r pairs are shown in figure 2.23. d. Repair of Damaged Skin. In case metal ski is damaged extensively, make repairs replacir an entire sheet panel from one structural mer ber to the next. The repair seams are to along stiffening members, bulkheads, etc., ai each seam must be made exactly the same i regard to rivet size, splicing, and rivet patter as the manufactured seams at the edges of th original sheet. If the two manufactured sear are different, the stronger one will be copi (See figure 2.24 for typical acceptable meth of repairs.) (1) Patching of Small Holes. Small hole skin panels which do not involve damage to stiffening members may be patched by c ing the hole with a patch plate in the me shown in figure 2.24. Flush patches also may be install stressed skin-type construction. An acce and easy flush patch may be made by trit ### Page 91 (OCR) AC 43.13-1A 1972 maining reinforcement tube. Allow one sleeve weld to cool before welding the remaining tube to prevent undue warping. 76. SPLICING USING LARGER DIAMETER REPLACE- MENT TUBES. The method of splicing structural tubes, as shown in figure 2.10, requires the least amount of cutting and welding. However, this splicing method cannot be used where the damaged tube is cut too near the adjacent cluster-joints, or where bracket-mounting provisions make it necessary to maintain the same replacement tube diameter as the original. As an aid in installing the replacement tube, squarely cut the original damaged tube leaving a minimum short stub equal to 2 1/2-tube diameters on one end and a minimum long stub equal to 41/2-tube diameters on the other end. Select a length of steel tube of the same material and at least the same wall thickness, having an inside diameter approximately equal to he outside diameter of the damaged tube. Fit Page 39 this replacement tube material snugly about the original tube with a maximum diameter difference of 1/16 inch. From this replacement tube material, cut a section of tubing diagonally (or fishmouth) of such a length that each end of the tube is a minimum distance of 11/2-tube diameters from the end of the cut on the original tube. Use a fishmouth cut replacement tube wherever possible. Remove the burr from the edges of the replacement tube and original tube stubs. If a fishmouth cut is used, file out the sharp radius of the cut with a small round file. Spring the long stub of the original tube from the normal position, slip the replacement tube over the long stub, then back over the short stub. Center the replacement tube between the stubs of the original tube. In several places tack-weld one end of the replacement tube, then weld completely around the end. In order to prevent distortion, allow the weld to cool completely, then weld the remaining end of the replacement tube to the original tube, ZZA- REPLACEMENT TUBE ORIGINAL → ZA TUBE - ІзА-→ **FREPLACEMENT TUBE 47 A** **ORIGINAL TUBE** •А- FIGURE 2.10. Splicing using larger diameter replacement tube. Par 75 ### Page 92 (OCR) **77. REPAIRS AT BUILT-IN FUSELAGE FITTINGS.** Make splices in accordance with the methods described in paragraphs 70 through 75. Repair built-in fuselage fittings in the manner shown in figure 2.11. The following paragraphs outline the different methods as shown in the figure: a. Tube of Larger Diameter Than Original. A tube (sleeve) of larger diameter than original is used in the method shown in figure 2.11(A). This necessitates reaming the fitting holes (at longeron) to a large diameter. The forward splice is to be a 80° scarf splice. Cut the rear longeron (right) approximately 4 inches from the centerline of the joint and a spacer 1 inch long fitted over the longeron. Edge-weld this spacer and longeron; make a tapered "V" cut approximately 2 inches long in the aft end of the outer sleeve and swage the end of the outer sleeve to fit the longeron and weld. b. Tube of Same Diameter as Original. In the method shown in figure 2.11(B) the new section is the same size as the longeron forward (left) of the fitting. The rear end (right) of the tube is cut at 80° and forms the outside sleeve of the scarf splice. A sleeve is centered over the forward joint as indicated. c. Simple Sleeve. It is assumed the longeron is the same size on each side of the fitting in this case, figure 2.11(C), and it is repaired by a simple sleeve of larger diameter than the longeron. d. Large Difference in Longeron Diameter Each Side of Fitting. Figure 2.11(D) assumes that there is 1/4 inch difference in the diameter of the longeron on the two sides of the fitting. The section of longeron forward (left) of the fitting is cut at 30°, and a section of tubing of the same size as the tube and of such length as to extend well to the rear (right) of the fitting is slipped through it. One end is cut at 30° to fit the 30° scarf at left and the other end fishmouthed as shown. This makes it possible to insert a tube of such diameter as to form an inside sleeve for the tube on the left of the fitting and an outside sleeve for the tube on the right of the fitting. Par 77 AC 43.13-1 19 78. ENGINE MOUNT REPAIRS. All welding on a engine mount should be of the highest qualit since vibration tends to accentuate any minc defect. Engine-mount members should prefers bly be repaired by using a larger diameter r placement tube telescoped over the stub of th original member and using fishmouth and r‹ sette welds. However, 30° scarf welds in plac of the fishmouth welds will be considered ac ceptable for engine-mount repair work. a. Check of Alignment. Repaired engin mounts must be checked for accurate align ment. When tubes are used to replace bent o damaged ones, the original alignment of th structure must be maintained. This can b done by measuring the distance betwee points of corresponding members that have no been distorted, and by reference to the manu facturer's drawings. b. Cause for Rejection. If all members are out of alinement, reject the engine mount and re place by one supplied by the manufacturer one which was built to conform to the man facturer's drawings. The method of checki, the alignment of the fuselage or nacelle poir should be requested from the manufacturer. c. Engine Mount Ring Damage. Repair mino damage, such as a crack adjacent to an engir attachment lug, by rewelding the ring and etending & gusset or a mounting lug past th damaged area. Engine mount rings which a extensively damaged must not be repaired, u less the method of repair is specifically a proved by an authorized representative of t FAA, or the repair is accomplished in acco! ance with FAA approved instructions nished by the aircraft manufacturer. **79. LANDING GEAR REPAIR.** a. Round Tube Construction. Repair lar gears made of round tubing using standa pairs and splices as shown in figures 2. 2.11. b. Streamline Tube Construction. Repaiz ing gears made of streamlined tubing by one of the methods shown in figures 2. 2.15. ### Page 93 (OCR) lAC 49.13-1A Page 45 1978 - GAP}" • 6А ЗА ￾2t 2t ЗА ЗА 6A - 30° ттт A- STREAMLINE TUBE LENGTH OF MINOR AXIS, PLATE WIDTHS. **B- DISTANCE OF FIRST PLATE FROM LEADING EDGE, /3 A.** **C- STREAMLINE TUBE LENGTH OF MAJOR AXIS.** S.L. SIZE ,572 „382 1,340 3,43 .714 ,476 1,670 4.28 1-/₴ .858 .572 2.005 5.15 1-4 1,000 ,667 2,339 6,00 . 2 1,144 ,762 2,670 6.86 2-¼ 1,286 .858 3.008 7,72 2-2 1.430 .954 3.342 8.58 FIGURE 2.15.-Streamline tube splice using plates (applicable to landing gears). Par 79 ### Page 94 (OCR) Page 46 AG 43.131 (A) (8) (C) (A), (B), AND (C) ARE TYPES OF REPAIRABLE AXLE **ASSEMBLIES, ASSEMBLIES ORIGINALLY HEAT-TREATED MUST BE REHEAT-TREATED AFTER WELDING.** (D) (D) IS A TYPE OF AXLE ASSEMBLY WHICH IS GENERALLY NONREPAIRABLE. FIGURE 2.16.-Representative types of repairable and nonrepairable axle assemblies. Par 79 ### Page 95 (OCR) L= SPAGES SPAGES d EACH TOTAL NO. OF RIVETS = 18(9 EACH SIDE OF SPLI A= ORIGINAL TUBE C = REP B= SLEEVE D USE AI7S ALUM. ALL **RIVETS, SMALL HEAD** **HAMMER STANDARD SPLICES MAY BE USED ON LIGHTLY- ROUND HEAD LONGERONS OR WEB MEMBERS.** APPROX. THICKNESS CORRECT WAY OF BRAZIER HEAD. INCORRECT WAY A, C * 3/4 7/8 065 065 7/8 1-178 065 1-1/8 1-1/411-3/81-1/2 065 11-1/41-3/8 058 .058 d MUST NOT BE LESS THAN D/8. ALL .058 THICK R= 3 d RIVET DIA 5 / 32 5/32 3/16 3/16 3/16 3 / 16 0000.0000 L 19-3/8|9-3/8111-1/4/11-1/4/11-1/4|11-1/4 0000: 0000 * INCLUDES ALL THICKNESSES UP TO AND IN MEACH. (USE BASIC ROUND SIZES IN TABLE STREAMLINE NOTE: USE SAME MATERIAL AS ORIGINAL FOR SLEE FOR SIZE OF RIVETS ETC.) FIGURE 2.20.-Typical repair method for tubular members of aluminum allo; ### Page 96 (OCR) AC 43.13 D CRACK 30 30 3020203030301 - D NOTE: FOR MINIMUM NUMBER OF RIVETS REQUIRED, SEE PARAGRAPH 100e. AND SUBSEQUENT. **AT LEAST AS THICK** **AS ORIGINAL** ORIGINAL TER • REINFORCEMENT MATERIAL - DURAL. OR 2117-AD RIVETS ALUMINUM ALLOY USED IN 3/32" DIA. MIN. **ORIGINAL CONSTRUCTION.** **SCALE - TWICE SIZE SECTION A-A** SCALE - TWICE SIZE 2017-T3 10000 SQ. ROD 3D 30 3D 3D _30 30 _ 3D • 2117-AD RIVETS 1/8" DIA: MIN FIGURE 2.21.-Typical repair for buckled or cracked formed metal wing rib capstrips. Par 100 ### Page 97 (OCR) - 1A **SEE (C) SEE (D) 0000 OO10 (A)** • SEE (E) (B) D * RIVET DIA. 2 1.5 t (APPROX.) (MINIMUM 3/32") **BEND RADI TO FIT** **ORIGINAL RIB.** (G) /I6PLYWOOD TO STIFFEr BENT RIBS. - MACHINE SCREW D OR RIVET. 1.5 t l APPROX. - PEEN (E) NOTE: FOR MINIMUM NUMBER OF RIVETS REQUIRED, SEE PARAGRAPH 100e. AND 3030 30 20) (D) SUBSEQUENT. FIGURE 2.22.—Typical metal rib repalrs (usually found on small and medium-size aircraft). Par 100 ### Page 98 (OCR) Pago 58 AC "А" •040" 2017-T3 SPACER 20, 30, 30,2020,30 3D 20 "B" 2024-T3 SPLICE BAR "Д" "в" 2117-AD RIVETS 3/32" DIA, MIN. CRACK 2117-AD RIVETS 3/32i DIA. MIN. - ЗА - ЗА 1. STRAIGHTEN CRIMPED OVER PORTION. • CRACK **2. INSERT HARDWOOD (ASH) INSERT SHAPED.** **TO CONFORM TO T.E. PIECE.** 2117-AD RIVETS 20 2D 2D 3/32" DIA. MIN. FLATTEN HEADS (APPROX. 1.5t ) OUTSIDE I. STRAIGHTEN CRIMPED PORTION. •CRACK **2. USE SAME AL. ALLOY AS IN ORIGINAL.** NOTE: FOR MINIMUM NUMBER OF RIVETS REQUIRED, SEE PARAGRAPH 100e. AND SUBSEQUENT. FroURE 2.28.- Typical repairs of trailing edges. Per 100 ### Page 99 (OCR) (A) ADDED STIFFEN" **ERS TO BACK UP SEAM OF PATCH.** 2 EACH PATOH PLATE TO 24-5/32" REQ'D BE OF SAME MAT- 12117-74 ¡RIVETS BY SAME BACKED ERIAL AS ORIGIN- AL AND SAME OR NEXT HEAVIER GAUGE. RIVET 3" HOLE MAX. DIA. SIZE, SPACING, **MAX DIA MATERIAL AND** **PATTERN SAME 16-5/32 AS ORIGINAL AB.** **PII7-T4 RIVETS DAMAGED SKIN** **SHOULD BE** CLEANED OUT "G", "H", AND "K" SMOOTHLY AS **ARE INDICATED BY CIRCULAR PATCH PLATES. COEF.** DESIGNATION DIE SIONATION OF CIR. PATCH LATES H "G DIAMETER (OUTSIDE) 12-1/2" 7-1/2" DIA. INNER RIVET CIRCLE 1-3/4° DIA. OUTER RIVET CIRCLE 16-1/2 **SAME RIVET SIZE, SPACING, MATERIAL** • PATTERN AS BO.. SAME RIVET SIZE, SPAGING, MATERIAL & PATTERN AS AD. , PATGH PLATE 8-5/32"2112 RIVETS. MAX DIA I' HOLE TO BE OF SAME MATERIAL AS ORIGINAL AND SAME OR NEXT HEAVIER BAUGE "K" - GRACA -NOT LESS THAN 2 TIMES, _PORTION OF SKIN REPLACED OF GRACK.- - ORIGINAL SKIN SHEET. 4 RIVETS RECD EACH SIDE. **RIVET SPACING AT HOLES DRILLED TO LEAST 3 X RIVET DIA PREVENT CRACK SPREADING.** SKIN GAUGE = ,028 (B) RIVET DIA. "=" CRACK 2,5" LONG" FroURE 2.24-Typical repairs of stressed sheet metal covering. Par 100 ### Page 100 (OCR) Раде 60 out the damaged area and then installing a conventional patch on the underneath side or back of the sheet being repaired. A plug patch plate of the same size as the opening may then be inserted and riveted to the patch plate. Other types of flush patches similar to those used for patching plywood may be used (ref. Chapter 1). The riveting pattern used, however, must follow standard practice to maintain satisfactory strength in the sheet. In general, patches in metal skin are not restricted as to size or shape; however, those of rectangular, circular, square, oval, and rectangular with round ends usually are more desirable because of appearance and ease of installation. e. Splicing of Sheets. The method of copying the seams at the edges of a sheet may not always be satisfactory; for example, when the sheet has cutouts, or doubler-plates at an edge-seam, or when other members transmit loads into the sheet, the splice must be designed as illustrated in the following exam- (1) Material: Clad 2024 sheet, 0.032-inch thickness. Width of sheet (i.e., length at splice) - "W" = 10 inches. (2) Determine rivet size and pattern for a single-lap joint similar to figure 2.18. (a) Use rivet diameter of approximately three times the sheet thickness, 8 × 0.032 = 0.096-inch. Use 1/8-inch 2117 AD rivets (5/82- inch 2117-AD would be satisfactory). (b) Use the number of rivets required per inch of width "W" from figure 2.29. (Number per inch 4.9 × .75 = 8.7 or the total number of rivets required = 10 × 8.7 or 87 rivets.) (e) Lay out rivet pattern with spacing not less than shown in figure 2,18. Referring to figure 2.18(A), it seems that a double row pattern with the minimum spacing will give a total of 40 rivets. However, as only 87 rivets are required, two rows of 19 rivets each equally spaced over the 10 inches will result in a satisfactory splice. f. Straightening of Stringers or Intermediate Framer. Members which are slightly bent may be straightened cold and examined with a mag- Far 100 AC 43.13-1 197 nifying glass for injury to the material. Rein force the straightened parts to an extent, de pending upon the condition of the material an the magnitude of any remaining kinks or buc kles. If any strain cracks are apparent, mak complete reinforcement in sound metal beyon the damaged portion. (1) Local Heating. Do not apply local heat ing to facilitate bending, swaging, flattening or expanding operations of heat-treated alumi num alloy members, as it is difficult to contro the temperatures closely enough to prevent possible damage to the metal, and it may im pair its corrosion resistance. g. Splicing of Stringers and Flanges. Make splices in accordance with the manufacturer's recommendations, which are usually contained in a repair manual. Typical splices for various shapes of sections are shown in figures 2.25 and 2.26. Design splices to carry both tension and compression and use the splice shown in figure 2.26 as an ex ample illustrating the following principles: (1) Statement of Principles: (a) To avoid eccentric loading and conse quent buckling in compression, place splicin or reinforcing parts as symmetrically as poss ble about the centerline of the member, and a tach to as many elements as necessary to pr vent bending in any direction; (b) To avoid reducing the strength tension of the original bulb angle the rive holes at the ends of the splice are made sma (no larger than the original skin attaching riv ets), and the second row of holes (thos through the bulbed leg) are staggered bi from the ends. In general arrange the rivet the splice so that the design tensile load for member and spliceplate can be carried into splice without failing the member at the most rivet holes; (c) To avoid concentration of loa end rivet and consequent tendency tow gressive rivet failure, the splice is ta the ends, in this case, by tapering the angle and by making it shorter than th bar (ref. figure 2.26) ; and (d) The preceding principles are ### Page 101 (OCR) 3.13-1A Page 61 00000000000000000 NOTE: FOR MINIMUM NUMBER TAPERED OF RIVETS REQUIRED, SEE SPLICE PARAGRAPH 100e. AND SUBSEQUENT. MEMBER NOTE: UNSHADED SECTIONS ARE ORIGINAL AND/OR REPLACEMENT SECTIONS. SHADED SECTIONS ARE CONNECTING OR REINFORCING SECTIONS. per W, a Wz FIGURE 2.25.-Typical stringer and flange splices. Par 100 ### Page 102 (OCR) Page 62 AC 43 CUT IN STRINGER TAPERED SPLICE MEMBER o 0-0 **ON LOWER SURFACE OF WING, MAKE END RIVET SAME SIZE AS SKIN ATTACHING RIVETS.** A → 2d2,MINIMUM - ід* 3 de, MINIMUM - THE NUMBER OF BOLTS REQUIRED IN THE "BULB" LEG OF THE STRINGER IS DETERMINED FROM FIG. 2.29 FOR 2017-T3 AS FOLLOWS: += 1/16" (Use .064") de= ,190" (AN3 bolt) NO. OF BOLTS PER INCH OF WIDTH FROM TABLE = 3.3. ACTUAL NO. OF BOLTS ON EACH SIDE OF CUI = WX 3.3 * (1.0)x 3.3 = 3.3 BOLTS USE 4BOLTS **WHEN BOTH RIVETS AND BOLTS** **ARE USED, BOLT HOLES SHOULD BE REAMED TO SIZE.** We (1.O) SPLICE ANGLE "A" SHADED AREA GREATER THAN Wexte. **SECTION A-A -KEEP LARGE HOLES BACK FROM END OF SPLICE.** **THE NUMBER OF RIVETS REQUIRED IN** THE UPPER LEG OF THE STRINGER IS DETERMINED FROM FIG. 2.29 FOR 2017-T3 AS FOLLOWS: t, = 1/16" (Use "064") d,= 1/8" (2117-AD rivet) NO. OF RIVETS PER INCH OF WIDTH FROM TABLE = 9.9 ACTUAL NO. OF RIVETS ON EACH SIDE OF CUT = WIx 9.9 = (.5),× 9.9 :4.95 RIVETS. USE SRIVETS. d, (1/8), RIVET DIA. PREFERABLY 2tTO 3t, BUT NOT GREATER THAN W/4. "B"SPLICE BAR. SHADED AREA GREATER THAN W, Xt.. **STRINGER** AREA REPRESENTING BULB FIOURE 2.26.--Example of stringer splice (material-2017 alloy). Par 100 ### Page 103 (OCR) NOTE: FOR MINIMUM NUMBER OF RIVETS REQUIRED, SEE PARAGRAPH 100e. AND SUBSEQUENT. NOTE: STRENGTH INVESTIGATION USUALLY REQUIRED FOR THIS TYPE OF REPAIR. FIGURE 2.27.-Application of typical flange splices and reinforcement. Par 100 ### Page 104 (OCR) Pae 64 AC 43.13 cially important in splicing stringers on the lower surface of stressed skin wings, where high tension stresses may exist, When several adjacent stringers are spliced, stagger the splices if possible. h. Size of Splicing Members. When the same material is used for the splicing members as for the original member, the next cross-section area (i.e., the shaded areas in figure 2.25) of the splicing member will be greater than the area of the section element which it splices. The area of a section element (e.g., each leg of an angle or channel) is equal to the width multipled by the thickness, For example, the bar "B" in figure 2.26 is assumed to splice the upper leg of the stringer, and the angle "A" to splice the bulbed leg of the stringer. Since the splice bar "B" is not as wide as the adjace leg, and since the rivet diameter is also at tracted from the width, the bar is made tw as thick in order to obtain sufficient net area. I. The Diameter of Rivels in Stringers. The i ameter of rivets in stringers might preferal be between 2 and 8 times the thickness "+" the leg, but must not be more than 1/4 t width "W" of the leg. Thus, 1/8-inch rivets a chosen in the example, figure 2.26. If t splices were in the lower surface of a wing, t end rivets would be made the same size as t skin-attaching rivets or 8/32 inch. I. The Number of Rivets. The number of rive required on each side of the cut in & stringer flange may be determined from standard tex FIGURE 2.28.- -Number of rivets required for splices (single-lap joint) in bare 2011-T6, 2024-T8, 2024-T96, al 7075-T6 sheet, clad 2014-T6, 2024-T8, 2024-T86, and 7075-T6 sheet, 2024-T4, and 7075-T6 plate, bar, rod, tut and extrusions, 2014-T6 extrusiona. No. of 3117-AD protruding bead No. of Thicknese reets required per inch of width "py" botta in inches 8/82 1/8 5/82 8/16 1/4 A7-8 0.016 4.0 .020 4.9 4.9 0.2 .086 7.0 .040 18 7.3 .051 9.8 „084 12.8 .081 .OĐ1 .102 .128 NOTES : a. For stringers in the upper surface of a wing, or in a fuselage, 80 percent of the number of rivets whown io table may be used. b. For intermediate frames, 60 percent of the number shown may be used. c. For single lap sheet Joints, 75 percent of the number shown may be used. EMoINEEBING NOTES: The above table was computed as follows: 1. The load per inch of width of material was calculated by assuming a strip one inch wide in tension. 2. Number of rivets required was calculated for 2117-AD rivets, baaed on a rivet allowable shear stre 40 percent of the sheet alloweble tensile stress, and a sheet allowable bearing stress aqual to 160 percent o allowable tensile stress, using nominal hole diameters for riveta. 3. Combinations of sheet thickness and rivet size above the heavy line are critical in (1. e., will fail by) the sheet; those below are critical in shearing of the rivets. 4. The number of AN-8 bolts required below the heavy line was calculated based on a sheet affowable ton of 70,000 p.s.l. and a bolt allowable single shear load of 2,128 pounds. Par 100 ### Page 105 (OCR) 43.13-1A Раде 65 oks on aircraft structures, or may be found figures 2.28, 2.29, and 2.30. In determining c number of rivets required in the example, ure 2.26, for attaching the splice bar "B" to e upper leg, the thickness "+" of the element area being spliced is 1/16 inch (use 0.064), e rivet size is 1/8 inch, and figure 2.29 shows at 9.9 rivets are required per inch of width. nce the width "W" is 1/2 inch, the actual imber of rivets required to attach the spliceir to the upper leg on each side of the cut is 9 (rivets per inch) × 0.5 (inch width) = 95 (use & rivets). or the bulbed leg of the stringer "+" = 1/16 ch (use 0.064); AN-3 bolts are chosen and e number of bolts required per inch of width 3.8. The width "W" for this leg, however, is inch and the actual number of bolts required 1 each side of the cut is 1 × 3.8 = 8.8 (use 4 ›lts). When both rivets and bolts are used in the same splice, the boltholes must be accurately reamed to size. It is preferable to use only one type of attachment, but in the above example, the dimensions of the legs of the bulb angle indicated rivets for the upper leg and bolts for the bulb leg. 1) Splicing of Intermediate Frames. The same principles used for stringer-splicing may be applied to intermediate frames, when the following point is considered: (a) Conventional frames of channel or Z section are relatively deep and thin compared to stringers, and usually fail by twisting or by buckling of the free flange. Reinforce the splice-joint against this type of failure by using a splice plate heavier than the frame and by splicing the free flange of the frame with a flange of the spliceplate, as illustrated in figure 2.31. Since a frame is likely to be subjected to bending loads, make the length of spliceplate |GURE 2.29.-Number of rivets required for splices (single-lap joint) in 2017, 2017 ALCLAD, 2021-T86 and 2021- T36 ALCLAD sheet, plate, bar, rod, tube, and extrusions. No. of 2117-AD protruding bead rivets required per inch of width "W" No. of bolts kness "t" In inches AN-3 6. 5 4. 9 6. 5 < 9 3. 9 6. 9 4. 9 3. 9 8. 9 4. 9 3. 9 10. 0 5. 6 11. 1 8. 2 7.0 coco so co Tea co co co co 9. 9 12. 5 10. 3 7.1 3. 9 12. 9 8. 9 1 9 3. 3 or stringers in the upper surface of a wing, or in a fuselage, 80 percent of the number of rivets shown in the be used r intermediate framee, 60 percent of the number shown may be used. ingle lap sheet joints, 75 percent of the number shown may be used. NOTES: The above table was computed as follows: id per inch of width of material was calculated by assuming a strip one inch wide in tension. er of rivets required was caloulated for 2117-AD rivete, based on a rivet allowable shear stress equal to i the sheet allowable tensile etress, and a sheet allowable bearing stress equal to 160 percent of the sheet usile stress, using nominal hole diametera for rivets. binations of sheet thickness and rivet size above the heavy line are critical in (i.e., will fail by) bearing on hose below are critical in shearing of the rivets. › number of AN-3 bolts required below the heavy line was caloulated based on s sheet allowable tensile stress .s.i. and a bolt allowable single shear load of 2,126 pounds. Per 100 ### Page 106 (OCR) Page 66 АС 43.13- "I" more than twice the width "W," and the rivets spread out to cover the plate. 101. REPAIRING CRACKED MEMBERS. Acceptable methods of repairing various types of cracks in structural elements are shown in figures 2.82 to 2.35. The following general procedures apply in repairing such defecta: a. Drill small holes 3/82 inch (or 1/8 inch) at the extreme ends of the cracks to minimize the possibility of their spreading further. b. Add reinforcement to carry the stresses across the damaged portion and to stiffen the joints (as shown in figures 2.82 to 2.35). The condition causing cracks to develop at a particular point is stress concentration at that point in conjunction with repetition of stress, such as produced by vibration of the structure. The stress concentration may be due to the q sign or to defects such as nicks, scratches, to marks, and initial stresses or cracks fro forming or heat-treating operations. It shou be noted, that an increase in sheet thickne alone is usually beneficial, but does not nece sarily remedy the conditions leading to crac ing. **102. STEEL AND ALUMINUM FITTINGS.** a. Steel Fittings-Inspection for Defects. (1) Finings are to be free from scratche vise and nibbler marks, and sharp bends ‹ edges. A careful examination of the fitting wit a medium power (at least 10 power) magnif ing glass is acceptable as an inspection. (2) When repairing aircraft after an accide or in the course of a major overhaul, inspe FIGURE 2.30.-Number of rivets required for splices (single-lap joint) in 5052 (all hardnesses) sheet. Thictness "t" in inches No. of 2117-AD protruding head rivets required per inch of width "p" No. of bolt AN-31 0.016-. 6. 3 1, 7 .020_. В. з 4.7 3. 8 .025__ 6. 8 3. 8 .032_. 6. 3 4, 7 3. 8 3. 2 .036__. 7.1 4.7 8. 8 3. 2 2.4 .040__ 7. 0 4. 7 3. 8 3. 2 2.4 .051-. 10. 1 5. 6 3. 8 3. 2 2,4 .064_ 12.7 7.0 46 3. 2 .081_. & 9 5. 8 2 .091-- 10. 0 6. 5 .102_ 11. 2 7. 3 .128_ 9. 2 NOTES: a. For stringers in the upper surface of a wing, or in a fuselage, 80 percent of the number of rivets shor table may be used. b. For intermediate frames, 60 percent of the number shown may be used. c. For single lap sheet jointe, 75 percent of the number shown may be used. ENGINEERINO NOTES: The above table was compuled as follows: 1. The load per inch of width of material was calculated by assuming a strip one inch wide in tension. 2. Number of rivets required was calculated for 2117-AD rivets, based on & rivet allowable shear stre. 70 percent of the sheet allowable tensile stress, and a sheet allowable bearing stress equal to 165 percent of allowable tensile stress, using nominal hole diameters for rivete 3. Combinations of sheet thickness and rivet size above the heavy line are critical in (l.e., will fall by) be the sheet; those below are critical in shearing of the rivets. Par 100 ### Page 107 (OCR) 13.13-14 Page 6? THE NUMBER OF RIVETS REQUIRED IN EACH LEG ON EACH SIDE OF THE CUT AND THE NET DIAMETER TH USING HIG. 22) IN DE ANNER SIMILAR TO THAT FOR STRINGERS IN FIG. 2.26. NOTE b. IN FIG. 2.29 INDICATES THAT ONLY 60% OF THE NUMBER OF RIVETS SO CALCULATED NEED BE USED IN SPLICES IN INTERMEDIATE FRAMES EXAMPLE (For 2017-T3 aluminum alloy frame) **FLANGE LEG** • = „040" (2.0) W, & W. = .6 inch d = 1/8" 2117-AD rivel ￾6) NO. OF RIVETS PER INCH OF WIDTH FROM FIG. 2.29 = 6.2 No. of rivets required = W × 6.2 = •6 × 6.2 = 3.72 or 4 rivets. 60% of 4 rivets = 2.4 rivets. **USE 3 RIVETS ON EACH SIDE OF THE CUT IN EACH FLANGE LEG.** WEB OF ZEE (OR CHANNEL) + = „040" d = 1/8" 2117-AD rivet W = 2.0 inches NO. OF RIVETS PER INCH OF WIDTH FROM FIG. 2.29 = 6.2 No. of rivets required = W x 6.2 = 2.0 x 6.2 =12.4 or 13 rivets. 60% of 13 rivets = 7.8 rivets. **USE 8 RIVETS ON EACH SIDE OF CUT IN THE WEB OF ZEE (OR CHANNEL).** SHOULD BE MORE THAN TWICE We hickness of splice plate to be greater than that of the frame to be spliced. FIGURE 2.81.—Example of splice of intermediate frame (material 2017 AL alloy). Par 102 ### Page 108 (OCR) Page 68 AC 43.13 **CHANNELS** **TRAILING** **FORWARD EDGE** **HORIZONTAL** **RIBS SKIN** **OVESTARL CRACKS** **CRACKS' ANCHOR** **NUTS FOR RIB** **ATTACH-** **MENT OF** **FIN TO SKIN** **STABILIZER** **AND TIP TO FIN CRACK** **TRAILING EDGE** **REAR SHEAR** **FRONT BEAM SPAR CRACKS** LEADING EDGE AILERO: NOTE: ALL REINFORCING PLATES TO BE OF REAR SAME ALLOY & APPROX. 1.5 THICKNESS OF ORIGINAL. FiGURE 2.32. -Typical methoda of repairing cracked leading and trailing edges and rib intersections. Par 102 ### Page 109 (OCR) **END RIE REAR SPAR** **CONTROL** **STIFFENING CHANNEL TO SURFACE** **FIT UNDER END OF RIB, FITTING** REINFORCEMENT AND UNDER ANCHOR NUT. MATERIAL DRILL REL " DIA.) **SAME ALLOY AS ORIGINAL, APPROX. HOLES 1/8** **1.5 THICKNESS OF ORIGINAL OR GREATER.** **RIVETS JOINING SPAR WEB WITH FRONT** **AND REAR CHANNEL TYPE REINFORCEMENT** **NOTCHED MEMBER** **CRACKS AT** **FITTING ANCHOR NUTS** **TRAILING EDGE** **PORTION OF RIB TORILL RELIEF HOLE** **SHEAR REAR PEAR** **REINFORCING PLATE - BEAM** **WITH FILLETED NOTCH** MATERIAL OF REINFORCING PLATES SAME ALLOY AS ORIGINAL AND 1.5 (OR MORE) × THIGKNESS OF ORIGINAL. FLAP HINGE FIoURE 2.88.-Typical methods of replacing cracked members at fittings. Par 103 ### Page 110 (OCR) Page 70 AC 43 **ORACKS IN SKIN AND** **TRANSVERSE STIFFENER** **AT DRAG RIB JUNCTION** IEPAIR SY OUTSIDE GUSSE TRANSVERSE STIFFENeR LONGITUI _ORAOK DUE **DRAG I REINFORCE-)** **MENT TO DE REPAIR OF GRACKED** **SAME MATERIAL STIFFENER** **AND AT LEAST** **SAME AREA AS** **TRANSVERSE** **STIFFENER** **FAILURES** SECTION A-A FIGURE 2.34.-Typical methods of repairing cracked frame and stiffener combinat Par 102 ### Page 111 (OCR) 1.13-1A Page 71 FRONT SPAR HORN RIB BOTTOM **FUSELAGE (TOP)** **FAILURES** **FUSELAGE (SIDE) USE SAME MATERIAL, NEXT HEAVIER GAUGE FOR REINFORGEMENT.** FIoURE 2.86.-Typical repairs to rudder and to fuselage at tail post. Par 102 ### Page 112 (OCR) Page 72 AC 43 +→#‡ BUSHING **ELONGATED OR WORN.** TAPER WASHER OF 4130 ADDED. THICKNESS NOT LESS THAN 1.5 t OR 167 GREATER THAN 2t. **ELONGATED** **HOLE IN** ORIGINAL REMOVE STRUT. ORIGINAL WASHER 2 NOTES: 16 R=M - iG I. METHODS A OR B MAY BE USED IN REPAIRIN TYPE C STRUT ENDS. **2. METHOD B SHOULD BE USED IN REPAIRIN TYPE D & E STRUT ENDS.** **3. MEMBERS ORIGINALLY HEAT TREATED M** BE REHEAT-TREATED AFTER WELDING, FIGURE 2.36.-Typical methods of repairing elongated or worn boltholes. Par 102 ### Page 113 (OCR) 43.12-1A CHG 1 highly stressed main fittings, as set forth in apter 7, and if necessary, take corrosion vention measures as recommended in Chap- (3) Replace torn, kinked, or cracked fittings. (4) Elongated or worn boltholes in fittings ich were designed without bushings are not be reamed oversize. Replace such fittings, ess the method of repair is approved by a resentative of the FAA. Do not fill holes h welding rod. Acceptable methods of rering elongated or worn boltholes in landing r, stabilizer, interplane, or cabane-strut ends y, not originally equipped with pin plates, are wn in figure 2.36. (Also see figure 2.11 on geron repair at fitting.) . Aluminum and Aluminum Alloy Fittings. (1) Replace damaged fittings with new parts, ring the same material specifications. (2) Repairs may be made in accordance with a furnished by the aircraft manufacturer or a substantiating the method of repair may be mitted to the FAA for review. 1. CASTINGS. Damaged castings are to be reced and not repaired unless the method of air is specifically approved by the aircraft nufacturer or substantiating data for the rer has been reviewed by the FAA. I. SELECTIVE PLATING IN AIRCRAFT MAINTE- NCE. Selective plating is a method of depositmetal from an electrolyte to the selected The electrolyte is held in an absorbent terial attached to an inert anode. Plating tact is made by brushing or swabbing the * (cathode) with the electrolyte-bearing Selective Plating Uses. This process can be I for any of the following reasons: Co prevent or minimize disassembly, reor masking costs. sizing worn components (plate to size). ing in damaged or corroded areas. plate small areas of extremely large plate electrical contacts. plate parts too large for existing Page 73 (7) To supplement conventional plating. (8) To plate components which become contaminated if immersed in a plating bath. (9) To cadmium-plate ultra high strength steels without hydrogen embrittlement. (10) On-site plating. (11) Reverse current applications (e.g., stain removal, deburring, etching, dynamic balancing). b. Specifications. Selective plating (electrodeposition), when properly applied, will meet the following specifications and standards: 2Q-C-320 Chromium Plating QQ-N-290 Nickel Plating QQ-P-416 Plating, Cadmium QQ-5-865 Silver Plating QQ-Z-325 Zinc Plating MIL-T-10727 Tin Plating MIL-C-14550 Copper Plating MII-G-45204 Gold Plating c. General Requirements. (1) Areas to be repaired by this process should be limited to reasonably small areas of large parts, although it may be desirable to plate small parts, particularly electrical or electronic parts, in their entirety. (2) All solutions should be kept clean and free from contamination. Care should be taken to insure that the solutions are not contaminated by used anodes or other plating solutions. Brush plating solutions are not designed to remove large amounts of scale, oil, or grease. Mechanical or chemical methods should be used to remove large amounts of scale or oxide. Use solvents to remove grease or oil. (3) Brush plating solutions are five to fifty times as concentrated as tank solutions. The current densities used range from 500 to 4,000 amps/feet?. The voltages listed on the solution bottles have been precalculated to give proper current densities. Too high a current density burns the plating, while too low a current density produces stressed deposits and low efficiencies. Agitation is provided by anode/cathode motion. Too fast a motion results in low efficiencies and stressed deposits; too slow a motion causes burning. A dry tool results in burnt plate, coarse grain structure, and unsound deposits. The tool* Par 102 ### Page 114 (OCR) * cannot be too wet. Solution temperatures of 110° to 120° F. are reached during operation. (4) Materials such as stainless steel, aluminum, chromium, and nickel, which have a passive surface, will require an activating operation to remove the passive surface. During the activating process, do not use solutions that have been previously used with reverse current because of solution contamination. d. Equipment. The power source should operate on either 110 or 220 volt alternating current (AC), 60 Hertz, single phase input. It should have a capability to produce direct current having smooth characteristics with controlled ripple and be able to output a current of at least 25 amperes at 0 to 25 volts. Minimum instrumentation of the power source should include a voltmeter, ammeter, and ampere-hour meter. (1) The ammeter should provide a full scale reading equal to the maximum capacity of the power source with an accuracy of +5 percent of the current being measured. (2) The voltmeter should have sufficient capacity to provide a full scale reading equal to the maximum capacity of the power source and an accuracy of +1.0 volt. (3) An ampere-hour meter should be readable to 0.001 ampere-hour and have an accuracy of 10.01 ampers-hour. (4) The stylus should be designed for rapid cooling and to hold anodes of various sizes and configurations. For safety, the anode holder should be insulated. (5) The containers for holding and catching runoff solutions should be designed to the proper configuration and be inert to the specific solution. (6) The mechanical cleaning equipment and materials should be designed and selected to prevent contamination of the parts to be cleaned. e. Materials. The anodes should be of highpurity dense graphite or platinum-iridium alloys. Do not mix solutions from different suppliers. This could result in contamination. f. Detail Requirements. On large parts, no area greater than approximately 10 percent of the total area of the part should be plated by this process. Small parts may be partially or completely plated. Special cases exceeding these Par 104 AC 43.13-1A CI limitations should be coordinated with the ma facturer of the plating equipment being used : his recommendations should be followed. g. Anode Selection. As a general guide, contact area of the anode should be appr mately one-third the size of the area to be plat When selecting the anode, the configuration the part will dictate the shape of the anode, h. Required Ampere-Hour Calculation. selected plating solution has a factor which equal to the ampere-hours required to dep 0.0001 inch on one square inch of surface. : termine the thickness of plating desired or certain area and multiply the solution fac times the plating thickness times the area square inches to determine the ampere-hours quired. This factor may vary because of te I. Cleaning. Remove corrosion, scale, oxi and unacceptable plating prior to processi Use a suitable solvent or cleaner to remove gre or oil. I. Plating on Aluminum and Aluminum Bi Alloys. (1) Electroclean the area using forw (direct) current until water does not break the surface. This electroclean process should accomplished at 10 to 15 volts, using the app priate electroclean solution. (2) Rinse the area in cold clean tap water. (3) Activate the area with reverse curre 7 to 10 volts, in conjunction with the proj activating solution until a uniform gray to bla surface is obtained. (4) Rinse thoroughly in cold, clean tap wat (5) Immediately electroplate to color wh the area is still wet, using the appropriate nic solution. (6) Rinse thoroughly. (7) Immediately continue plating other solution to desired thickness. (8) Rinse and dry. k. Plating on Copper and Copper i (1) Electroclean the area usin (direct) current until water does no the surface. The electroclean process accomplished at 8 to 12 volts using t. priate electroclean solution. ### Page 115 (OCR) **43.13-1A CHG 1** (2) Rinse the area in cold, clean tap water. (3) Immediately electroplate the area with y of the plating solutions except silver. Silver quires an undercoat. (4) Rinse and dry. I. Plating on 300 and 400 Series Stainless Steels, ckel Base Alloys, Chrome Base Alloys, High ckel Ferrous Alloys, Cobalt Base Alloys, Nickel ate, and Chrome Plate. (1) Electroclean the area using forward irect) current until water does not break on e surface. This electroclean process should be complished at 12 to 20 volts using the approiate electrocleaning solution. (2) Rinse the area in cold, clean tap water. (3) Activate the surface using forward lirect) current for 1 to 2 minutes using the tivating solution and accomplish at 6 to 20 Its. (4) Do not rinse. (5) Immediately nickel flash the surface to thickness of 0.00005 to 0.0001 inch using the propriate nickel solution. (6) Rinse thoroughly. (7) Immediately continue plating with any her solution to desired thickness. (8) Rinse and dry. m. Plating on Low-Carbon Steels (Heat Treated 180,000 psi). (1) Electroclean the area using forward irect) current until water does not break on e surface. This electroclean process should be complished at 12 to 20 volts using the approiate electrocleaning solution. (2) Rinse the aren in cold, clean tap water. (3) Reverse current etch at & to 10 volts, the appropriate activating solution, until form gray surface is obtained. Rinse thoroughly. mediately electroplate the part using ns except copper acid or silver. Both re undercoats. se and dry. ig on Cast Iron and High-Carbon Steels at Treated to 180,000 psi). flectroclean the area using forward current until water does not break on Page 73-2 the surface. This electroclean process should be accomplished at 12 to 20 volts using the appropriate electrocleaning solution. (2) Rinse the area thoroughly in cold, clean tap water. (3) Reverse current etch at & to 10 volts, using the appropriate etching solution, until a uniform gray is obtained. (4) Rinse thoroughly. (5) Remove surface smut with 15 to 25 volts using the appropriate activating solution. (6) Rinse thoroughly. (7) Electroplate immediately using any of the solutions except copper or silver (both these require undercoats). (8) Rinse and dry, o. Plating on Ultra High Strength Steels (Heat Treated Above 180,000 psi). (1) Electroclean the area using REVERSE current until water does not break on the surface. This electroclean process should be accomplished at 8 to 12 volts using the appropriate electroclean solution. (2) Rinse the area thoroughly in cold, clean tap water. (3) Immediately electroplate the part, using either nickel, chromium, gold, or cadmium. Other metals require an undercoat of one of the above. Plate initially at the highest voltage recommended for the solution so as to develop an initial barrier layer. Then reduce to standard voltage. (4) Rinse and dry. (5) Bake the part for 4 hours at 375° +25° F. NoTF 1: Where the solution vendor provides sub stantiating data that hydrogen embrittlement will not result from plating with a particular solution, then a post bake Is not required. This substantlating data can be in the form of aircraft Industry manufacturers' process specifications, military specifications, or other suitable data. NoTE 2: Acid etching should be avoided, if pos sible. Where etching is absolutely necessary, It should always be done with reverse current. Use alkaline solutions for initial deposita. p. Dissimilar Metals and Changing Base. a general rule, when plating two dissimilar metals, follow the plating procedure for the one with the most steps or activation. If activating* Par 104 ### Page 116 (OCR) * steps have to be mixed, use reverse current activation steps prior to forward (direct) current. activation steps. q. Plating Solution Selection. (1) Alkaline and neutral solutions are to be used on porous base metals, white metals, highstrength steel, and for improved coating ability. Acid solutions are to be used for rapid buildup and as a laminating structure material in conjunotion with alkaline type solutions. (2) Chrome brush plating solutions do not yield as hard a deposit as bath plating solutions. The hardness is about 600 Brinell as compared to 1,000 Brinell for hard chrome deposited from a tank. (3) Silver immersion deposits will form with no current flowing on most base metals from the silver brush plating solutions; such deposits have poor adhesion to the base metal. Consequently, a flash or a more noble metal should be deposited prior to silver plating to develop a good bond. (4) In general, brush plating gives less hydrogen embrittlement and a lower fatigue strength loss than does equivalent tank deposits. However, all brush-plated, ultra high strength steel parts (heat treated above 180,000 psi) Par 104 AC 43.13-1A CI should be baked as mentioned in paragr 104.0.(5), unless it is specifically known t embrittlement is not a factor. r. Qualification Tests. All brush plated t faces shall be tested for adhesion of the elec deposit. Apply a 1-inch wide strip of Minnes Mining and Manufacturing tape code 250, or approved equal, with the adhesive side to freshly plated surface. Apply the tape v heavy hand pressure and remove it with quick motion perpendicular to the plated t face. Any plating adhering to the tape shall cause for rejection. s. Personnel Training for Quality Cont Manufacturers of selective plating equipm provide training in application techniques their facilities. Personnel performing select plating must have adequate knowledge of methods, techniques, and practices invol These personnel should be products of th training programs and certified as qualified erators by the manufacturers of the prod used, as well as by local quality control deps **105.-114. RESERVED.** ### Page 117 (OCR) Page 74 AC 43.13 Section 4. REPAIR OF LAMINATE STRUCTURES 115. GENERAL. There is a wide variation in the composition and structural application of laminates, and it is essential that these factors be given major consideration when any restoration activities are undertaken. To a similar extent, there also exist many types of laminate structure repairs that may or may not be suitable for a given condition. For this reason, it is important that the aircraft or component manufacturer's repair data be reviewed when determining what specific type of repair is permissible and appropriate for the damage at The materials used in the repair of laminate structures must preserve the strength, weight, aerodynamic characteristics, or electrical properties of the original part or assembly. This . can best be accomplished by replacing damaged material with material of identical chemical composition or a substitute approved by the manufacturer. In order to eliminate dangerous stress concentrations, avoid abrupt changes in cross-sectional areas. Whenever possible, for scari joints and jacings, make small patches round or oval-shaped, and round the corners of large repairs. Smooth and properly contour aerodynamic surfaces. It is recommended that test specimens be prepared at the same time that the actual repair is accomplished. These can then be subjected to a destructive test to establish the quality of the adhesive bond in the repaired part. To make this determination valid, the specimens must be assembled with the same adhesive batch mixture and subjected to curing pressure, temperature, and time identical with those in the actual repair. a. Fiberglass Laminate Repairs. The following repairs are applicable to fiberglass laminate used for fairings, covers, cowlings, honeycomb panel facings, etc. Prior to undertaking the re- Par 115 pair, clean the repair area thoroughly with castile soap and warm water. Remove a paint by wet or dry sanding methods. Se blasting may be used but caution must be ex cised to not abrade the surfaces excessively. Superficial scars, scratches, surface abi sion, or rain erosion can generally be repair by applying one or more coats of a suital resin, catalyzed to cure at room temperatu to the abraded surface. The number of co required will depend upon the type of re and severity of the damage. Damage not exceeding the first layer or of fiberglass laminate can be repaired by fil with a putty consisting of a compatible ro temperature-setting resin and clean short gl fibers. Before the resin sets, apply a sheet cellophane over the repair area and work c any bubbles and excess resin. After the res has cured, sand off any excess and prepare t area for refinishing. Damage deep enough to seriously affect t strength of the laminate (usually more th the first ply or layer of fabric) may be 1 paired as illustrated in figure 2.87. Coat t **DAMAGE TO FACING** **DAMAGED MATERIAL /REMOVED BY SANDIN . REPLACED LAMINATIONS** SANDED TO CONTOUR FIGURE 2.37.-Typical laminate (facing) repair. Chaj ### Page 118 (OCR) I-IA Page 75 I area with room-temperature-setting DAMAGE TO FACHG ANO COME and apply contoured pieces of glass fabaked in resin. Apply a cellophane sheet he repair and work out any bubbles and resin. After the resin has cured, scrape excess resin and sand the surface of the to the original contour. OOT DAMASED COME MACHO' SCAMPED POR MATC REMOVEU, nage that extends completely through cing and into the core requires the rerent of the damaged core and facing. A d for accomplishing this type of repair is in figure 2.88. An alternate method for ling the facing is shown in figure 2.89. imaged portion is carefully trimmed out rcular or oval shape and the core matemoved completely to the opposite facing. se caution so as not to damage the oppoeing or to start delamination between ngs and the core around the damage. FIGURE 2.88.-Typical core and facing repair. CAUTION iding fiberglass laminates gives off a fne dust it may cause skin and/or respiratory irritation less suitable skin and respiration protection is ed. › replacement core stock of the same maand density as the original (or an acceptsubstitute) and cut it to fit snugly in the ned hole. When all pieces of replacement YOLE TRIMMED, "STEP-PEELED COME NEMOVEL r laminations are cut and soaked in resin, AND FACING all surfaces of the hole and the scarfed with resin. Then coat all surfaces of the replacement with resin and insert it into ole. After all of the pieces of resin-imated glass-fabric facing are in place, WET-LAMINATED PATCH WITH the entire area with a piece of cellophane EXTRA LAYER OF THIN CLOTH 'arefully work down the layers of fabric move any air bubbles and excess resin. y light pressure by means of sand bags or cuum blanket. When the resin has cured, the repair to match the original contour efinish the surface. TEMPORARY BLOCK OR MOLO 'TEMPORARY SHIM mage that is completely through a lamisandwich structure may be repaired as ilated in either figure 2.39 or 2.40. The ed joint method, figure 2.40, is normally on small punctures up to 8 or 4 inches in mum dimension and in facings which are e of thin fabric that is difficult to peel. The ›ed joint, figure 2.39, is most often used on **COMPLETE BY STEP-PEELING AND WET-LAMINATING OPPOSITE** FACING IN A SIMILAR OPERATION FIGURE 2.88.-Typical stepped Joint repair. larger repairs to facings composed of thick fabrics. When access for repair is possible from only one side, blind repair procedures are employed. Per 116 ### Page 119 (OCR) Page 76 АС • OLE Repairs of this type have been made by t glass-fiber-reinforced plastic backing pli proximately 0.060-inch thick. The b plate can be fabricated from layers of fabric and room-temperature-curing re sandwiching the resin-impregnated lay‹ - 100т-. HOLE AND TRIMMED, CORE REMOVED, FACING SCARFED tween two sheets of cellophane. Work t cess resin out and allow the plate to cur flat surface or a cellophane-covered sur! the proper curvature near the damaged or at the same location on a comparable u aged part. The technique for bondir backing plate in place is similar to tha WET-LAMINATED PATCH trated in figure 2.41 for metal-faced la repairs. After the backing plate is bor place with resin, complete the repair as in figure 2.88. b. Metal-Faced Laminate Repairs. Rel TEMPORARY MOLDLOCK TEMPORARY SHIM laminate with magnesium, titanium, o OR less steel facings require procedures w COMPLETE BY SCARFING AND WET-LAMINATING OPPOSITE specially devised and not included in FACING IN A SIMILAR OPERATION lowing methods of repair. Aluminum FIGURE 2.40.-Typical scarfed joint repair. such as 7075-T6, 2024-T8, and 2014-1 **WOOD OR SELF-TAPPING** **METAL SCREWS STEP 1- PLYWOOD CAUL** **BACKING PLATE --CELLOPHANE** **ANO COPE** **BEING GLUED** **IN PLACE** **OLUE SOUEEZE-OUT** REPLACE PLAT 10064" **PLYWODO CAUL** **CUSHION TAPERED COVER STEP 8- -CELLOPHANE** **PLATE BEING** **GLUED IN PLACE** **JAPERED COVER PLATE (SAME THICKNESS AS FACING)** **FILLED SCREW HOLES -OVEALAP 125T-** COMPLETED REPAIR FIGURE 2.41.-Typical one side blind repair. Par 115 ### Page 120 (OCR) 13-1A only used for the repair of facings for ated structural parts having aluminum gs. For maximum corrosion resistance, nly clad aluminum for repairs to clad alum alloy facings. hts, scratches, or fractures, not exceeding ourth inch in largest dimension, in alumifacings may be repaired with a suitable such as viscous epoxy resin. Thoroughly the repair area with fine sandpaper and he before applying the filler. After the has partially cured, remove any excess with a sharp scraper or chisel. When stely cured, sand to original contour. If amage included a fracture, reclean the round the filled hole and apply a surface as shown in figure 2.42, by means of resin and a similar curing cycle. tures or punctures in one facing and damage to the core of an aluminumaminate may be repaired by several difmethods. The technique used will deupon the size of the damage and the igth, aerodynamic, and sonic fatigue rence requirements of the area involved. If repair requires aerodynamic smoothness, 'acing surrounding the repair core cavity -DAMAGE TO METAL FACING **HONEYCOMB CORE 1 CELLOPHANE** PLYWOOD PRESSURE PLATE CUSHION: **ELWOOD OR SELF-TAPPING SCREWS.** **OVERLAP TAPERED COVER PLATE** •FILLER FIGURE 2.42.- -Typical metal facing patch. Page 77 may have to be step cut to one-half its thickness. This can be done by using a router with an end mill bit and a template. Damage that extends completely through the core and both facings may be repaired using the same general techniques as those used for repairing fiberglass laminates when both facings are accessible (see figures 2.39 and 2.40). When the inner facing is not accessible, a procedure such as that shown in figure 2.41 may be used. The outer surface patch could then be installed in a manner similar to the repair shown in figure 2.42. After locating the extent of the total damaged area by tapping or other nondestructive test methods, remove the damaged facing and that portion of the core material which is also affected. The depth to which the core must be removed will depend upon the type of core material and method of repair. If the replacement core material is the same as the original, fabricate it to shape keeping the same core ribbon or grain direction. When a substitution is permissible, wood or glass-fabric honeycomb cores are sometimes used in the repair of aluminum honeycomb cores as they are generally easier to shape. Typical types of core replacementa are shown in figure 2.43. Resin fills can be used to replace the core and facing where smaller core damage exists. Phenolic microballoons, low-density insulating materials and/or other ingredients, are added to lower the density and give greater flexibility, thus lowering stress concentrations in the repair area. For the repair of larger holes in which it is inconvenient to use a face patch because of aerodynamic smoothness requirements in that area, both the core and facing are sometimes replaced with glass-fiber fabric discs and resin. Undercut the core, as shown in figure 2.44, in order to obtain a better bonding of the fill with the facing. Fill the core cavity with accurately shaped resin-saturated glass cloth discs, and press each ply down to remove any air bubbles. Special care should be taken that the final plies fit well against the underside of the top facing. When the core cavity is filled, close the cutout in the facing with resin-impregnated glassfiber fabric discs that have been precut to size. There are five general types of facing Par 115 ### Page 121 (OCR) Page 78 AC 4 A. REPLACEMENT WITH CORE PLUG OF SAME OR OTHER TYPES OF CORE MATERIAL patches used after replacement of the cor been completed and they are described k COME PLUG Each of these patches meet certain servi COME IEYCOMB. AOUTED AREA ¡ EPOXY-RESIN quirements that must be considered bef ADHE SIVE decision can be reached as to which is a priate for a specific application. (1) A Small, Flush Plug-type. A plug ( same material and thickness as the faci B. RESIN FILL bonded in place after the core has bee EPOXY-RESIN 'FILL. ORIGINAL ROUTED AREA- placed. This type of patch gives aerody smoothness but does not replace any o HONEYCOMB. COME PARTIAL ON COMPLETE DEPTH OF CORE strength initially lost in the damaged facin 12) An Overlap Type Metal Facing Patc overlap surface patch is bonded in place the facing and replaced core. With this t patch, and good bonding techniques, i C. RESIN FILL, UNDERCUT CORE cally full strength of the facing is reg EPOSY-RESIN FILL. MOUTED AREA Overlaps of less than 125 times the thi HONEYCOMB ORIGINAL DEPTH OF COME ATIAL OR COMPLETE of the facing can be used in repairs wh COME CORE UNDERCUT skins are not highly stressed. (3) A Step-type Facing Patch. The ste facing patch retains aerodynamic smoot but has only a maximum of 50 percent c D. GLASS-CLOTH-RESIN LAMINATE, UNDERCUT CORE original facing tensile strength. (4) A Scarf-type Facing Patch. This is a OUTED AREA bination of the overlap and the ster CORE which results in the same aerody! smoothness as the step-type but gr strength efficiency. (5) An Overlap Type Resin-glass Facing E. REPLACEMENT WITH CORE PLUG OF SAME An overlap patch of glass-fiber fabric TYPE OF SANDWICH CONSTRUCTION resin has fair aerodynamic smoothness a SANDWICH HOUTED AREA easy to fabricate with a minimum of e ORIGIMAL МОЖЕУСОНА EPOXY-RESIN ment. It fits easily in transition curvature CORE ADHESIVE results in a lower facing strength thai original skin. c. Finishing. The type of finish coating al FIGURE 2.43.-Typical types of core replacement. to a laminate repair will normally be **CORE ORIGINAL CORE TO SKIN ADHESIVE** - JAGGED EDGES **SECTION SHOWING UNDERCUT** • FINISHED UNDERCUT AREA FIGURE 2.44.-Typical undercut core material cavity. Par 115 ### Page 122 (OCR) „13-1A d by the facing material and the applicaof the part or assembly. Rain erosion of ic parts, the need for electrical or dielecproperties and/or the necessity for anticorn coatings must be considered when the e of finish is made. Plastic-faced parts as radomes are finished primarily for rain on while aluminum- or other metal-faced nates are finished for corrosion protection. coatings to perform their function prop- Page 79 erly, it is essential that they be applied to surfaces that are clean, free of voids, and exceedingly smooth. The edges of all parts not protected by a bonding of aluminum or glass-fabric laminate must be sealed to reduce the rate of moisture absorption. The wood-bonded edges of plywood, mahogany, spruce, or balsa may be given two coats of aluminized spar varnish. **116.-126. RESERVED.** Par 115 ### Page 123 (OCR) 13-1A Page 81 Chapter 3. FABRIC COVERING Section 1. PRACTICES AND PRECAUTIONS TEXTILE MATERIALS. All fabric, surface reinforcing tape, machine thread, lacing etc., used for re-covering or repairing an sit structure must be high-grade aircraft e material of at least as good quality and alent strength as those described in subcraphs a through g. Aircraft Fabric. Acceptable fabrics such as and linen for covering wings, control yes, and fuselages are listed in figure 3.1. es conforming to the Aeronautical Mate- 'pecifications incorporate a continuous ung showing the specification number to it identification of the fabric in the field. Re-covering Alrcraft. Re-cover or repair airwith fabric of at least as good quality equivalent strength as that originally used he aircraft. However, in re-covering air- : which were originally covered with low igth or so-called "glider cloth," it is coned more desirable to use Grade A or "inediate fabric conforming to AMS 3806 or , as amended, respectively. Certain sync and fiberglass fabrics have been develthat are acceptable alternates to AMS or AMS 8804 fabric, providing the Supental Type Certificate (STC) installation uctions furnished with the material are wed. Specification MIL-C-9084, MIL-Y- C, and MIL-G-1140 materials in the uned condition have equivalent strength char- •istics to TSO-C15 material specifications. Reinforcing Tope. Acceptable reinforcing is listed in figure 8.2. Use reinforcing tape imilar quality as the fabric and at least alf the strength of that conforming to :fication MIL-T-5661. Surface Tape. Use surface tape (also finishape) having approximately the same props as the fabric used. See figure 8.2. •. Lacing Cord. Use lacing cord having a strength of at least 80 pounds double or 40 pounds single strand. See figure 8.2. f. Machine Thread. Use machine thread having a strength of at least 5 pounds single strand (figure 8.2). g. Hand-Sowing Thread. Use hand-sewing thread having a strength of at least 14 pounds single strand (figure 8.2). 128. COVERING PRACTICES, The method of fabric attachment should be identical, as far as strength and reliability are concerned, to the method used by the manufacturer of the airplane to be re-covered or repaired. Fabric may be applied so that either the warp or fillthreads are parallel to the line of flight. Either the envelope method or blanket method of covering is acceptable. a. Flutter Precautions. When re-covering or repairing control surfaces, especially on high performance airplanes, make sure that dynamic and static balances are not adversely affected. Weight distribution and mass balance must be considered to preclude the possibility of induced flutter. 129. PREPARATION OF THE STRUCTURE FOR COVERING. One of the most important items in covering aircraft is proper preparation of the structure. Dopeproofing, covering edges which are likely to wear the fabric, preparation of plywood surfaces, and similar operations, if properly done, will do much toward insuring an attractive and long-lasting job. o. Dopeproofing. Treat all parts of the structure which come in contact with doped fabric with a protective coating such as aluminum foil, dopeproof paint or cellulose tape. Clad aluminum and stainless steel parts need not be dopeproofed. Par 127 ### Page 124 (OCR) b. Chafe Points. Cover all points of the structure, such as sharp edges, boltheads, etc., which are likely to chafe or wear the covering with doped-on fabric strips or cover with an adhesive tape. After the cover has been installed, reinforce the chafe points of the fabric by doping on fabric patches. Where a stronger reinforcement is required, apply a cotton duck or leather patch sewed to a fabric patch and then dope in place. Reinforce all portions of the fabric pierced by wires, bolts, or other projections. c. Inter-Rib Bracing. Conventional wing ribs, which do not have permanent inter-rib bracing should be bound in position by means of cotton tape running parallel to the beams. Apply the tape diagonally between the top and bottom capstrips of each successive rib approximately halfway between the front and rear beams. Apply the tape continuously from the butt rib to the tip rib with one turn of tape around each intermediate rib capstrip. d. Preparation of Plywood Surfaces for Covering. Prior to covering plywood surfaces with fabric, prepare the surface by cleaning and applying sealer and dope. (1) Cleaning. Sand all surface areas which have been smeared with glue in order to expose a clean wood surface. Remove loose deposits such as woodchips and sawdust. Remove oil or grease spots by carefully washing with naphtha. (2) Application of Sealer and Dope, Apply one brush coat or two dip coats (wiped) of a dopeproof sealer such as that conforming to Specification MIL-V-6894 thinned to 80 percent nonvolatile content and allow to dry 2 to 4 hours. Finally, before covering, apply two ng thy hat brush coats of clear dope allowing the first coat of dope to dry approximately 45 minutes before applying the second coat. 130. FABRIC SEAMS. Seams parallel to the line of flight are preferable; however, spanwise seams are acceptable. a. Sewed Seams. (1) Machine-sewn seama (parts D, E, and F of figure 8.8) should be of the folded fell or French fell types. Where selvage edges or Par 129 AC 4 pinked edges are joined, a plain lap se satisfactory. (2) Begin hand sewing or tacking : point where machine sewing or uncut fal again reached. Lock hand sewing at int of 6 inches, and finish the seam with ; stitch and a knot (figure 8.4). At the where the hand sewing or permanent to is necessary, cut the fabric so that it c doubled under before sewing or permanent ing is performed (figure 3.8 (C)). After sewing has been completed, remove the t rary tacks. In hand sewing, use a minim four stitches per inch. (3) Cover a sewed spanwise seam metal- or wood-covered leading edge pinked-edge surface tape at least 4 wide. (4) Cover a sewed spanwise seam trailing edge with pinked-edge surface u least 3 inches wide. For aircraft with 1 exceed speeds in excess of 200 m.p.h notches at least 1 inch in depth and 1 ir width in both edges of the surface tape used to cover spanwise seams on trailing ‹ especially the trailing edges of control sur Space notches at intervals not exceedin inches. On tape less than 8 inches wid notches should be 1/3 the tape width. I event that the surface tape begins to ser because of poor adhesion or other cause tape will tear at a notched section, thus venting progressive loosening of the length of the tape which could seriously the controlling of the aircraft. (5) Cover a double-stiched lap joint pinked-edge surface tape at least 4 i wide. (6) Make sewed spanwise seams or upper or lower surface in a manner tha amount of protuberance is minimum, Cove seam with pinked-edge tape at least 3 i wide. (7) Sewed seams parallel to the lir flight (chordwise) may be located over however, place the seam on the rib so tha lacing will not penetrate through the seam. ### Page 125 (OCR) 13.13-1A Page 63 Minimum Minimum Minimum ten- Materials Specification tensile tearing sile strength Thread count Use and remarks atrength new strength new deter lorated per inch (undoped) (undoned) (undoped) ›lane cloth Society Auto- 80 pounda 5 pounds 56 pounds 80 mini- For use on all aircraft. ercerized motive En- per inch warp and per inch. mum, 84 Required on aircraft ›tton gineera AMB warp fll. maxi- with wing loadings Brade "A"). 3806 (T80- and fill. mum greater than 9 p.s.f. Oló refer- warp Required on aircraft ences thia and fill. with placarded neverspec.). exceed speed greater than 160 m.p.h. MIL-C-5646 Alternate to AMB 8806. ›lane cloth MIL-C-5013 Alternate to MIL-C-5646 allulose or AM8 3806 (unitrate pre- doped). Finish with oped. cellulose nitrate dope. ›lane cloth MIL-C-5642. Alternate to MIL-C-5646 ellulose or AM8 8806 (unmetate doped). Finish with utyrate, cellulose acetate redoped. butyrate dope. plane cloth Society Auto- 65 pounds 4 pounds 48 pounds 80 mini- For use on aircraft with tercerized motive En- per inoh warp per inoh. mum, 04 wing loadings of 9 p.s.f. ston. gineera AMS warp and fill. maxi- or less, provided never- 8804 (T80- and fill. mum exceed speed is 160 C14 refer- warp m.p.h. or less. ences this and fill. врес.). ›lane cloth Society Auto- 50 pounds 8 pounde 35 pounde 110 maxi- For use on gliders with Lercerized motive En- per inch warp per inch. mum wing loading of 8 p.s.f. itton. gineers AM8 warp and fill. warp or less, provided the 3802. and fll. and fill. placarded neverexceed speed is 185 m.p.h. or less. bric A. A. F. No. 55 pounds 4 pounds 30 pounds 80 mini- Alternate to AMS 8802- 16128. per inch warp per inch. mum A. AMB 3802. warp and fill. warp and fill. and fill. British 7F1_- This material meets the minimum strength requirements of TSO- C15. FIGURE 8.1.-Textile fabric used in aircraft covering. Par 130 ### Page 126 (OCR) Page 84 AC 4 Materiala Specification Yarn alse Minimum ten- aile strength Yarda per pound Use and remarks Reinforcing tape, MII-1-6661 150 pounda Used as reinforcing tape on cotton. per one- and under rib lacing half-inch Strength of other widths a width. in proportion. Lacing cord, pre- MII-0-5649 80 pounds 810 mini- Lacing fabric to structure waxed braided double. mum. less already waxed, mu cotton. lightly waxed before and Lacing cord, special U.8. Army No. 20/8/8/8 -- 85 pounde cotton. 8-27. double. Lacing cord, braided MIL-C-5648 80 pounda 170 minicotton. single. mum. Lacing cord thread; MIL-T-6779 a ply 59 pounds 820 minilinen and linen- single. mum. hemp. 11 ply 70 pounds 510 mininingie. mtim. Lacing cord thread; MIL-T-5060 Ticket No. 62 pounds 480 minihigh-tenacity cot- 10 single. mum. ton. Machine thread cot- Federal V-T- 20/4 ply - pounds 5,000 nor- Use for all machine sewing ton. 278b. singie. mal. Hand sewing thread V-T-278b. 8/4 ply 14 pounds 1,650 nor- Use for all hand sewing. Use cotton. single. mal. waxed thread. Burface tape cotton MIL-1-5083 80 1bs/in. - Use over seams, leading (made from AN- trailing edges, outer edge 0-121). ribs, pinked, scalloped or st edges. Surface tape cotton. Same as fabric Same as Alternate to MIL-T-5088. used. fabric used. FIGURE 8.2.-Miscellaneous textile materials. b. Doped Seams. (1) For a lapped and doped spanwise seam on a metal- or wood-covered leading edge, lap the fabric at least 4 inches and cover with pinked-edge surface tape at least 4 inches (2) For a lapped and doped spanwise seam at the trailing edge, lap the fabric at least 4 Par 130 inches and cover with pinked-edge sur at least 3 inches wide. **131. COVERING METHODS.** a. The Envelope Method. The enve of covering is accomplished by sewin widths of fabric cut to specified and machine sewn to form an envel ### Page 127 (OCR) PIN & TEMPORARY WRAPPING• **TO HOLD TENSION GIVEN TO** **TOP COVER** **TOP COVER (D) PLAIN OVERLAP SEAM.** **FORMER (A) ATTACHING FABRIC AT AILERON CUTOUT.** (E) FRENCH FELL SEAM. (F) FOLDED FELL SEAM. **TOP COVER FINISH TOP COVER** **WITH TAPE** SPA CANE CLOSER (B) ATTACHING FABRIC AT METAL WING BUTT. (C) ATTACHING FABRIC AT WOODEN WING BUTT. **TACKS SHOULD BE** **TINNED IRON OR BRASS, STAINLESS STEEL NO.18 B. W. G.** FIGURE 8.3.-Typical methods of attaching fabric. Par 131 ### Page 128 (OCR) **PULL NO. 2-** **COMPLETED KNOT** TO TIGHTEN KNOT IO LB STITCH AFTER PULL NO.I- HOLD STITCH TENSION* **BY PRESSING HERE DURING PULL NO.2 OPERATION NO. 2 -EDGE OF CAPSTRIP** DIRECTION OF LACING OPERATION NO. I FIGURE 3.4.-Standard knot for rib laring and terminating a sewed seam (modified spine knot). can be drawn over the frame. The trailing and outer edges of the covering should be machine sewn unless the component is not favorably shaped for sewing, in which case, the fabric should be joined by hand sewing. b. The Blanket Method. The blanket method of covering is accomplished by swing together widths of fabric of sufficient lengths to form a blanket over the surfaces of the frame. Join the trailing and outer edges of the covering by a plain overthrow or baseball stitch. For airplanes with placard never-exceed speed of 150 miles per hour or less, the blanket may be lapped at least 1 inch and doped to the frame or the blanket, lapped at least 4 inches at the nose of metal- or wood-covered leading edges, doped, and finished with pinked-edge surface tape at least 4 inches wide. When fabricating both the envelope and blanket coverings, cut the fabric in lengths sufficient to pass com- Par 131 pletely around the frame, starting at the t ing edge and returning to the trailing edge. 132. REINFORCING TAPE. Place reinforcing ; of at least the width of the capstrips unde lacing. In the case of wings with plywoo metal leading edge covering, the reinfor tape need be brought only to the front spa the upper and lower surfaces. a. Use of Antitear Strips. On aircraft never-exceed speed in excess of 250 mile hour, antitear strips are recommended the reinforcing tape on the upper surfa wings, and the bottom surface of that pa the wing in the slipstream. Where the anti strip is used on both the top and bottom faces, pass it continuously up to and aro the leading edges and back to the trailing e Where the strip is used only on the top i face, carry it up to and around the leac edge and back on the lower surface as far as the front beam. For this purpose the s stream should be considered as being eque the propeller diameter plus one extra rib si on each side. Cut antitear strips from the same materia used for covering and wide enough to ext beyond the reinforcing tape on each side to engage the lacing cord. Attach the str applying dope to that part of the fabric covered by the strip, and apply dope f over the strip. **133. LACING.** a. Securely fasten both surfaces of fa covering on wings and control surfaces t ribs by lacing cord or any other metho nally approved for the aircraft. Care sho taken to insure that all sharp edges a which the lacing cord may bear are pr by tape in order to prevent abrasion cord. Join separate lengths of lacing the splice knot shown in figure 3.5. the common square knot for this p ercise the utmost care to assure sion and security of all stitches. or starting stitch use a double li trated in figure 3.6. Make all stitches using a single loop tied o standard knot for rib lacing (mor ### Page 129 (OCR) 1.13-1A shown in figure 8.4). The spacing between starting stitch and the next stitch should ne-half the normal stitch spacing. Final loon of the knot depends upon the original tion selected by the manufacturer. If such rmation is not available, consider positionthe knot where it will have the least effect he aerodynamics of the airfoil. The seine t admits a possibility of improper tightenresulting in a false (slip) form with tly reduced efficiency and must not be used stitch tie-offs. Lock the tie-off knot for the stitch by an additional half-hitch. Where ning ends, as at the rear beam and at the ling edge, space the last two stitches at half normal spacing. Under no circumaces pull tie-off knots back through the lacholes. . The double-loop lacing illustrated in figure represents a method for obtaining higher ngths than possible with the standard sinlacing, When using the double-loop lacing, the tie-off knot shown in figure 3.8. . Fuselage Lacing. Fabric lacing is also necesy in the case of deep fuselages, and on fuseas where former strips and ribs shape the ric to a curvature. In the latter case, lace fabric at intervals to the formers. Attachof the fabric to fuselages must be so acplished as to be at least the equivalent in ength and reliability to that used by the ufacturer of the airplane. • STITCH SPACING. The stitch spacing should **ILL TO TIGHTEN PULL TO TIGHTEN** **KNOT FORMED BUT NOT TIGHTENED** • LOAD KNOT COMPLETED FIGURE 8.5.-Splice knot. Page 87 not exceed the spacing approved on the original aircraft. In case the spacing cannot be ascertained due to destruction of the covering, acceptable rib-stitch spacing is specified in figure 8.9. Place the lacing holes as near to the capstrip as possible in order to minimize the tendency of the cord to tear the fabric. Lightly wax all lacing cords with beeswax for protection. In case waxed-braided cord is used, this procedure is unnecessary. (See figure 3.2 for acceptable lacing cords.) 135. SURFACE TAPE (FINISHING TAPE). Cover all lacing with tape of at least the quality and width used on the original airplane. This tape should not be applied until the first coat of dope has dried. Replace all inspection openings in the covering, and reinforce the fabric around them and along leading edges with tape. Wear or friction, induced by moving parts or fittings, can be repaired by sewing a leather patch on a fabric patch and doping in place. Pinked surface tape is sometimes applied over the trailing edges of control surfaces and airfoils. For sure application, the tape must be at least 3 inches in width, and if the aircraft never-exceed speed is greater than 200 miles per hour, notch the tape at intervals not exceeding 18 inches. If separation of the tape from the trailing edge begins, it will tear at a notched section and thereby prevent loosening of the entire strip which could seriously affect the controllability of the aircraft. 136. SPECIAL FASTENERS. When repairs are made to fabric surfaces attached by special mechanical methods, duplicate the original type of fastener. When self-tapping screws are used for the attachment of fabric to the rib structure, observe the following procedure: a. Redrill the holes where necessary due to wear, distortion, etc., and in such cases, use a screw one size larger as a replacement. b. Extend the length of the screw beyond the rib capstrip at least two threads of the grip (threaded part). c. Install a thin washer, preferably celluloid, under the heads of screws and dope pinkededge tape over each screw head. Par 133 ### Page 130 (OCR) Pae 68 AC 4: STANDARD RIB LACING KNOT LEAD "A" HOOPS HITON A LEAD STITON WITH LEAD **BEFORE PROGEEDING TO SECOND STITCH.** LOCK SQUARE KNOT WITH / HALF HITCH AROUND STITCH LOOPS C WITH END "B" CUT OFF SURPLUS OF END "B" - CAPSTRIP -LEAD "A" END "8" **TIE SQUARE KNOT ON CENTER OF CAPSTRIP.** STITON LOOPS "O" **SATIN AH FOR** **CAPSTRIP -** **TIE SQUARE KNOT ON CENTER OF CAPSTRIP.** 151 8190 TARTING STITCI FIGURE 8.6.-Starting stitch for rib stitching. Par 136 ### Page 131 (OCR) 3.13-1A Page 89 **DOUBLE LOOP LACING, SEE FIG. 3.8 FOR DETAILS OF KNOT.** — CAPSTRIP **CAPSTRIP** **STARTING STITCH, SEE FIG. 3.6 FOR DETAILS.** 9* NORMAL STITCH SPACING FIGURE 3.7.-Standard double loop lacing. пр 3 Par 136 ### Page 132 (OCR) Page 90 AC • - PULL NO.4 PULL NO. 4- 10 L8. **TO COMPLETE KNOT.** **EDGE OF** **CAPSTRIP** **DIRECTION OF LACING PULL NO. 2 - 6 6B** **TO TIGHTEN STITCH** **COMPLETED KNOT** **EDGE OF CAPSTRIP PULL NO.3** **PULL SNUG OPERATION NO. 2 -STITCH LOOPS** **AROUND CAPSTRIP PULL NO. 1 - 10 LB. FROM PREGEEDING** TO TIGHTEN LOOPS STITCH OPERATION NO. I FIGURE 8.8.-Standard knot for double loop lacing. Par 136 ### Page 133 (OCR) 43.13~1A MAXIMUM SPACING OF RIB LACING - INCHES I. IF ORIGINAL RIB STITCH SPACING CANNOT BE DETERMINED, NOTES USE SPACING INDICATED IN THESE CURVES. **2. LACING TO BE CARRIED TO LEADING EDGE WHEN VELOCITY EXCEEDS 275 M.P.H.** • SPACING OTHER THAN IN SLIPSTREAM- SPACING IN SLIPSTREAM- PLACARD NEVER EXCEEDS SPEEDS-M.P.H. (INDICATED) (CURVES PRESUME LEADING EDGE SUPPORT REINFORCEMENT SUCH AS PLYWOOD, METAL) FIGURE 8.9.-Fabric attachment spacing. **37.-147. RESERVED.** 1г 136 Chap 3 ### Page 134 (OCR) Page 92 AC 43,13 Section 2. DOPING 148. APPLICATION OF AIRCRAFT DOPE, EPOXY, it be done properly. Since local atmosphe AND RESINS. Determine that dope and fabric conditions affect the doping process, determ materials are compatible by consulting the the amount of thinner necessary at the ti product manufacturer's instructions before the finishing material is to be applied by fi applying finish to aircraft surfaces. Compati- using it on experimental panels. bility of products may also be determined by wetting samples of the fabric with the impreg- 149. BLUSHING AND USE OF BLUSH-RETARDI nating materials and thorough examination of THINNER. Blushing of dopes is very comm the material after it has dried. The following when doping is accomplished under humid c military specifications, or later revisions ditions. The condition is caused by the ral thereof, apply to aircraft dope, epoxy, and poly- evaporation of thinners and solvents, wh ester resins: lowers the témperature on the surface, causi MIL-D-5549A-1 condensation of moisture and producing 1 white appearance known as "blush. Blushi MIL-D-7850 tendencies are also increased, if strong ci rents of air flow over the surface when app MIL-D-6550A-1 ing dope or immediately thereafter. MIL-D-5551A-2 A blushed finish has very little protective tautening value. When the relative humidity MIL-D-5558A-2 such that only a small amount of blushing MIL~D-5552A-1 encountered, the condition may be eliminat MIL-D-5554A-1 by thinning the dope with a blush-retardi MIL-D-5555-1 MIL-R-9800A thinner and slightly increasing the room te perature. If it is not possible to correct humi MIL-R-25042A ity conditions in the dope room, suspend dopi operations until more favorable atmosphe MIL-R-7575B conditions prevail. The use of large amounts blush-retarding thinner is not advisable 1 cause of the undesirable drying properties ¿ companying the use of this material. 150. NUMBER OF COATS. Apply as many coats dope as are necessary to result in a taut al well-filled finish job. A guide for finishing fa ric-covered aircraft follows: a. Two coats of clear dope, brushed on al sanded after the second coat, b. One coat of clear dope, either brushed sprayed, and sanded. c. Two coats of aluminum pigmented dop sanded after each coat. Par 148 Chap ### Page 135 (OCR) 43.13-1A Page 95 72 Section 3. REPAIRS TO FABRIC COVERING 55. GENERAL, Make repairs to fabric-covered irfaces in a manner that will return the origial strength and tautness to the fabric. Sewed spairs and unsewed (doped-on patches or pan- ato a surface which contains aluminum or ¡her color coats. Whenever it is necessary to id fabric reinforcement, remove the old dope ther by softening and scraping or by sanding own to the point where the base coat or clear ›at is exposed. Use clear dope in doping the abric to the surface. After reinforcement is lade, normal finishing procedures may be fol- ›wed. 66. REPAIR OF TEARS IN FABRIC. Repair tears & shown in figure 8.10 by sewing the torn edges together using a baseball stitch and doping a piece of pinked-edge fabric over the tear. If the tear is a straight rip, the sewing is started at one end so that, as the seam is made, the edges will be drawn tightly together throughout its entire length. If the openings are cut in wings to inspect the internal structure, start the sewing at the corner or point so that the edges of the cover will be held in place while the seams are being made. The sewing is done with a curved needle and well-waxed thread. Clean the surface to be covered by the patch by rubbing the surface with a rag dipped in dope, wiping dry with a clean rag, or by scraping the surface with a putty knife after it has been softened with fresh dope. Dope solvent or acetone may be used for the same pur- 1" MIN. LOCK EVERY 8 OR IO STITCHES → 1.51 АЛАЛ - START SEWING **AT THE POINT** **LOCK STITCHING** **AT ENDS WITH** **1.5 MODIFIED SEINE KNOT** MATERIALS SHOULD →1.5 BE AT LEAST AS GOOD AS ORIGINAL FIGURE 8.10.-Repair of tears in fabric. hap 3 Par 165 ### Page 136 (OCR) pose, but care should be taken that it does not drop through on the inside of the opposite surface, causing the dope to blister. Cut a patch of sufficient size from airplane cloth to cover the tear and extend at least 1 1/2 inches beyond the tear in all directions. The edges of the patch should either be pinked similar to surface tape or frayed out about 1/4 inch on all edges. 167. SEWED PATCH REPAIR. When the damage is wing than go in such that it will not permit sewing the edges together, a sewed-in repair patch may be used if the damage is not longer than 16 inches in any one direction (see figure 8.10). Cut out the damaged section, making a round or ovalshaped opening trimmed to a smooth contour. Clean the area of the old fabric to be doped as indicated in paragraph 166. Turn the edges of the patch 1/2 inch and sew to the edges of the opening. Before sewing, fasten the patch at several points with a few temporary stitches to facilitate sewing the seams. After the sewing is completed, clean the area of the old fabric to be doped as indicated for small repairs and then dope the patch in the regular manner. Apply surface tape over the seam with the second coat of dope. If the opening extends over or closer than 1 inch to a rib or other laced member, extend the patch 3 inches beyond the member. After sewing has been completed, rib lace the patch to the rib over a new section of reinforcing tape, using the method explained in paragraph 133. Do not remove the old rib lacing and reinforcing tape. 168. UNSEWED (DOPED-ON) REPAIRS. Unsewed (doped-on) repairs may be made on all aircraft fabric-covered surfaces provided the never-exceed speed is not greater than 150 miles per hour. A doped patch repair may be used if the damage does not exceed 16 inches in any direction. Cut out the damaged section making a round or oval shaped opening, trimmed to a smooth contour. Clean the edges of the opening which are to be covered by the patch with a grease solvent. Sand or wash off the dope from the area around the patch with dope thinner. Support the fabric from underneath while sanding. Par 166 AC 43.1: For holes up to 8 inches in size, make fabric patch of sufficient size to provide a of at least 2 inches around the hole. On h over 8 inches in size, make the overlap of fabric around the hole at least 1/4 the hole ameter with a maximum limit of lap o inches. If the hole extends over a rib or cl than the required overlap to a rib or ot laced member, extend the patch at leas inches beyond the rib. In this case, after edges of the patch have been doped in pl and the dope has dried, lace the patch to rib over a new section of reinforcing tape the usual manner. Do not remove the old lacing and reinforcing tape. All patches sho have pinked edges or, if smooth, should be: ished with pinked-edge surface tape. 169. REPAIR BY A DOPED-IN PANEL. When 1 damage exceeds 16 inches in any directi make the repair by doping in a new panel. T. type of repair may be extended to cover bi the upper and lower surfaces and to cover s eral rib bays if necessary. Lace the panel to the ribs covered, and dope or sew as in t blanket method. a. Remove the surface tape from the ribs i jacent to the damaged area and from the tri ing and leading edges of the section being paired. Leave the old reinforcing tape a lacing in place. Next, cut the fabric along a li approximately 1 inch from the ribs on t sides nearest the injury and continue the ci to completely remove the damaged section. ! not remove the old fabric from the leading a trailing edges, unless both upper and low surfaces are being re-covered. b. Cut a patch to run around the trailing ed 1 inch and to extend from the trailing edge to and around the leading edge and back a proximately to the front beam. Extend t patch approximately 3 inches beyond the ri adjacent to the damage. As an alternative attachment on metalwood-covered leading edges, the patch may lapped over the old fabric at least 4 inches the nose of the leading edge, doped, and fi ished with at least 4 inches of pinked-edge su face tape. Chap ### Page 137 (OCR) : 43.13-1A ￾72 c. Clean the area of the old fabric that is to be vered by the patch and apply a generous coat ' dope to this area. Put the new panel in ace, pull as taut as possible, and apply a coat I dope to that portion of the panel which overps the old fabric. After this coat has dried, pply a second coat of dope to the overlapped rea and let dry. d. Place reinforcing tape over the ribs under Chap 3 Page 97 moderate tension and lace down in the approved manner. e. Give the panel a coat of clear dope and allow to dry. Install surface tape with the second coat of dope over the reinforcing tape and over the edges of the panel. Finish the dope scheme, using the regular doping procedure. **170.-180. RESERVED.** Par 169 ### Page 138 (OCR) Page 98 AC 43.13- Section 4. FABRIC TESTING 181. TESTING OF FABRIC COVERING. Field test ber of tests in 30 percent. Fabric which h instruments that are commonly used to test the less than 70 percent of the original requir tensile strength of aircraft fabric covering tensile strength would not be considered a: give only approximate indications of the fabric worthy. Figure 3.1 contains the minimum te condition. Since the accuracy of field test in- sile strength values for deteriorated fabric struments is affected by climatic and environ- tested in the undoped condition. mental conditions, a laboratory test is recom- (3) Grade A fabric may be used whe mended when aircraft fabric covering is found only intermediate fabric is required. Wh to be marginal by field test methods. Labora- testing for deteriorated condition, 46 poun tory test procedures are set forth in Federal (70 percent of original requirements for inte Specification CCC-T-191B, methods 5122, mediate fabric) is considered airworthy. olo2, 5184, or 5136; American Society of Test. (4) Failures may occur in fiberglass cove ing Materials (ASTM) Method D89-61 or ing where rib stitching has worn through tl D39-49, and others. In all cases, test fabric reinforcing tape and covering material witho specimens in the undoped condition. Use ace- being detected through visual inspection. Su tone, dope thinner, or other appropriate thin- failures can be located by using a suitable su ning agents for the removal of finishing mate- tion cup and lifting the fabric in the r stitched area. If the fabric pulls away from tl a. Strength Criteria for Aircraft Fabric. (1) Present minimum strength values for new aircraft fabric covering are contained in figure 8.1. ribs, new stitching will need to be applie using additional reinforcing tape and doublin the number of stitches throughout the affecte area. Five particular attention to the are: within the propeller slipstream area. (2) The maximum permissible deterioration for used aircraft fabric based on a large num- 182.-192. RESERVED. Par 181 Chap 3 ### Page 139 (OCR) : 43.13-1A Page 99 172 Chapter 4. CONTROL CABLES AND TERMINALS Section 1. INSPECTION AND REPAIR 93. GENERAL. Aircraft control cables are genrally fabricated from carbon steel or corroion-resistant steel wire and may consist of eiher flexible or nonflexible type construction. Contents of this section may be used for conrol cable installations pertaining to both prinary and secondary system applications. a. Cable Definitions. Construction features of arious cables are shown in figure 4.1. The folowing terms define components used in airraft control cables. (1) Wire-Each individual cylindrical steel 'od or thread. (2) Strand-Each group of wires helically Avisted or laid. (3) Core Strand--The central strand about which the remaining strands of the cable are elically laid. (4) Coble-A group of stranda helically ¡wisted or laid about a central core. (5) Preformed Cable-Cable in which the wires and strands are shaped prior to fabrication of the cable. (6) Diameter-The diameter of cable is the diameter of the circumscribed circle. (7) Lay or Twist-The helical form taken by the wires and strands in a cable. A cable is said to have a right-hand lay if the wires and strands twist in the same direction as the thread on a right-hand screw. (8) Pitch-The distance in which a strand or wire makes one complete revolution about the axis of the cable or strand respectively. 194. CABLE SPECIFICATIONS. Cable size and strength data are given in figure 4.2. These values are acceptable for repair and modification of civil aircraft. a. Cable Proof Loads. Cable terminals and Chap 4 T **CHAMETER DIAMETER** **NONFLEXIBLE 7X7 7X19** FLEXIBLE EXTRA FLEXIBLE FIGURE 4.1.--Cable cross section. splices should be tested for proper strength prior to installation. Gradually apply a test load equal to 60 percent of the cable breaking strengths given in figure 4.2 for a period of 3 minutes. Place a suitable guard over the cable during the test to prevent injury to personnel in the event of cable failure. 195. REPLACEMENT OF CABLES. Replace control cables when they become worn, distorted, corroded, or otherwise injured. If spare cables are not available, prepare exact duplicates of the damaged cable. Use materials of the same size and quality as the original. Standard swaged cable terminals develop the full cable strength and may be substituted for the original terminals wherever practical. However, if facilities and supplies are limited and immediate corrective action is necessary, repairs may be made by using cable bushings, eye splices, and the proper combination of turnbuckles in place of the original installation. (See figure 4.6c.) a. Location of Splices. Locate splices so that no portion of the splice comes closer than two Par 193 ### Page 140 (OCR) Page 100 AC 43.13-14 1971 1X7 and 1X10 7X7, 7X19, and 6X19 (1WRC) Nonflexible, carbon Corrosion resisting Flexible, carbon Flexible, corrosion resisting Diameter (Inch) MIL-W-6910 MIL-C-5699 MIL-W-1511 MIL-C-5424 Weight, Breaking Weight, Breaking Weight, Breaking Weight, Breaking pound: strength, pounds strength, pounds strength, pounds strength, per 100 feet- pounds per 100 feet pounds per 100 feet pounds per 100 feet pounds ½a- 0. 25 185 0. 25 150 564. • 65 375 • 55 875 Ko- . 85 500 - 85 500 0. 75 480 0. 75 480 561- 1. 40 800 1. 40 800 2. 00 1, 200 2, 00 1, 200 I. 60 920 1. 60 020 2.70 1, 600 2.70 1, 600 3. 50 2, 100 3. 50 2, 100 2. 90 2, 000 2. 90 1, 760 5. 50 3, 300 5. 50 3, 300 4. 50 2, 800 4. 50 2, 400 7. 70 4, 700 7. 70 4, 700 6. 50 4, 200 6. 50 3, 700 10. 20 6, 300 10. 20 6, 300 8. 60 5, 600 8. 60 5, 000 ¼ 13. 50 8, 200 18. 50 8, 200 11. 00 7, 000 11. 00 6, 400 Ysa 13. 90 8, 000 13. 90 7, 800 21. 00 12, 500 21. 00 12, 500 17. 30 9, 800 17. 80 9, 000 20. 70 12, 500 21. 30 14, 400 24. 30 12, 000 35. 60 17, 600 35. 60 16, 300 45. 80 22, 800 45.80 22, 800 *The strength values listed were obtained from straight tension tests and do not include the effects of wrapped ends. FIGURE 4.2.-Strength of steel cable. inches to any fair-lead or pulley. Locate connections at points where jamming cannot occur during any portion of the travel of either the loaded cable or the slack cable in the deflected position. b. Cutting and Heating. Cut cables to length by mechanical means. The use of a torch in any manner is not permitted. Do not subject wires and cables to excessive temperature. Soldering bonding braid to control cable will not be considered satisfactory. c. Ball-and-Socket Type Terminals. Do not use pall-and-socket type terminals or other types for general replacement that do not positively prevent cable untwisting, except where they were utilized on the original installation by the aircraft manufacturer. d. Substitution of Cable. Substitution of cable for hard or streamlined wires will not be acceptable unless specifically approved by a representative of the Federal Aviation Administration. Par 195 **196. MECHANICALLY FABRICATED CABLE AS- SEMBLIES.** a. Swage Type Terminals. Swage type terminals, manufactured in accordance with Air Force-Navy Aeronautical Standard Specifications, are suitable for use in civil aircraft up to and including maximum cable loads. When swaging tools are used, it is important that all the manufacturers' instructions, including "go and no-go" dimensions, be followed in detail to avoid defective and inferior swaging. Observ. ance of all instructions should result in a terminal developing the full rated strength of the cable. Critical dimensions, both before and after swaging, are shown in figure 4.8. (1) Terminals. When swaging terminals onto cable ends, observe the following procedure: (a) Cut the cable to the proper length, allowing for growth during swaging. Apply a preservative compound to the cable ends before insertion into the terminal barrel. NorE: Never solder cable ends to prevent fraying Chap 4 ### Page 141 (OCR) AC 43,13-1A Page 101 1972 Before swaging After swaging Cable size (Inches) Wire strands Minimum Outside Bore Bore Swaging breaking Shank diameter diameter length length strength diameter* (роипав) 7×7 0. 160 0. 078 1. 042 0. 969 480 0. 188 352 7X7 • 218 - 109 1. 261 1. 188 920 . 190 7×19 . 250 . 141 1. 511 1. 488 2, 000 - 219 7×10 . 297 . 172 1. 761 1. 688 2, 800 , 250 7×10 359 • 208 2. 011 1. 988 4, 200 . 813 7×10 427 • 234 2. 261 2. 188 8, 600 - 875 7X10 494 . 265 2.511 2. 438 7, 000 438 7×19 563 . 297 2. 761 2.688 8, 000 500 F10-. 7×10 635 - 328 8. 011 2. 938 9, 800 563 7×19 4 708 - 890 3. 510 3. 438 14, 400 . 625 * Use gages in kit for checking diameters. FIGURE 4.8.-Straight-shank terminal dimensions (AN-666, 687, 668, and 669). since the presence of the solder will greatly increase the tendency of the cable to pull out of the terminal. (b) Insert the cable into the terminal aproximately one inch, and bend toward the terminal; straighten the cable back to normal position and then push the cable end entirely into the terminal barrel. The bending action puts a kink or bend in the cable end and provides enough friction to hold the terminal in place until the swaging operation can be performed. Bending also tends to separate the stranda inside the barrel, thereby reducing the strain on them. NorE: If the terminal ia drilled completely through, push the cable into the terminal until it reaches the approximate position shown in figure 4.4. If the hole is FIGURE 4.5.-Gauging terminal shank after swaging. **BEND CABLE THEN PUSH** IN TO SWAGING POSITION FIGURE 4.4.-Insertion of cable into terminal. Chap 4 not drilled through, insert the cable until the end rests against the bottom of the hole. (c) Accomplish the swaging operation in accordance with the instructions furnished by the manufacturer of the swaging equipment. (d) Inspect the terminal after swaging to determine that it is free from die marks and splits, and is not out-of-round. Check for cable slippage in the terminal and for cut or broken wire strands. lel Using a "go no-go" gauge or a mi- Par 196 ### Page 142 (OCR) crometer, check the terminal barrel diameter as shown in figure 4.5. (f) Test the cable by proof-loading it to 60 percent of its rated breaking strength. (2) Splicing, Completely severed cables, or those badly damaged in a localized area, may be repaired by the use of an eye terminal bolted to a clevis terminal. (See figure 4.6a.) However, this type of splice can only be used in free lengths of cable which do not pass over pulleys or through fair-leads. (3) Swaged Ball Terminals. On some aircraft cables, swaged ball terminals are used for attaching cables to quadrants and special connections where space is limited. Single shank terminals are generally used at the cable ends, and double shank fittings may be used at either the end or in the center of the cable. Dies are supplied with the swaging machines for attaching these terminals to cables in the following manner: (a) The steel balls and shanks have a hole through the center, and are slipped over the cable and positioned in the desired location. (b) Perform the swaging operation in accordance with the instructions furnished by the manufacturer of the swaging equipment. (c) Check the swaged fitting with a "go no-go" gauge to see that the fitting is properly AM 668 TERMINAL (c) FIGURE 4.6.- Typical cable splices. Par 196 AC 43.13-1 197 **MEASURE SHANK** - AFTER SWAGING 1/16 CABLE 3/32 MEASURE SHANK **BEFORE SWAGING** **SLEEVES** **MEASURE BALL** • BEFORE SWAGING **MEASURE BALL** AFTER SWAGING FIGURE 4.7.-Typical terminal gauge. compressed. (See figure 4.7.) Also inspect the physical condition of the finished terminal. (4) Cable Slippage in Terminal. Ensure that the cable is properly inserted in the terminal after the swaging operation is completed. Instances have been noted wherein only 1/4 inch of the cable was swaged in the terminal. Observance of the following precautions should minimize this possibility: (a) Measure the length of the terming end of the fitting to determine the prop length of cable to be inserted into the barrel of the fitting. (b) Lay off this length at the end of the cable and mark with masking tape. Since the tape will not slip, it will provide a positive marking during the swaging process. (c) After swaging, check the tape marker to make certain that the cable did not slip during the swaging operation. (d) Remove the tape and, using red paint, paint the junction of the swaged fitting and cable. (e) At all subsequent service inspections of the swaged fittings, check for a gap in the painted section to see if cable slippage has occurred. b. Nicopress Process. A patented process using copper sleeves may be used up to the full rated strength of the cable when the cable is looped around a thimble. This process may also be used in place of the five-tuck splice on cables up to and including 8/8-inch diameter. The use of sleeves that are fabricated of materials other than copper will require engineering approval of the specific application by a representative of the Federal Aviation Administration. Chap 4 ### Page 143 (OCR) 43.13-1A Page 103 72 Copper oval sleeve stock No. Bleeve length Sleove length Cable size Manuni tool No. before com- after com- Number of Tested strength pression (approx.) pression (approx.) presses (pounds) Plain Plated * (Inches) unches 18-11-B4 28-11-B4 51-B4-887 340 18-1-C 28-1-C 51-C-887 550 18-2-G 28-2-G 51-G-887 1, 180 18-3-M 28-3-M 51-M-850 2,300 18-4-P 28-4-P 51-P-850 3, 050 18-0-X 28-6-X 51-X-850 1% 4, 350 18-8-F2 28-8-F2 51-F2-850 1}1⅞ 4 5, 790 18-10-F6 28-10-F6 3-F6-950 1% 1% 7, 180 18-13-G9 28-18-G9 3-G9-950 1⅜ 11, 130 No. 035 HTy. draulle tool dies 18-23-H5 28-23-H5 Oval H5 1% 16, 800 18-21-J8 28-21-J8 Oval J8 2⅛ 19, 700 18-25-K8 28-25-K8 Oval K8 2½ 25, 200 18-27-M1 | 28-27-M1 Oval M1 2% 31, 025 18-28-N5 28-28-N5 Oval N5 3% 39,200 Tequired on stainless rable. FIGURE 4.8.- Copper oval sleeve data. Before undertaking a nicopress splice, determine the proper tool and sleeve for the cable to se used. Refer to figures 4.8 and 4.10 for details on sleeves, tools, and the number of presses required for the various sizes of aircraft cable. The tool must be in good working condition and properly adjusted to assure a satisfactory splice. To compress a sleeve, have it well centered in the tool groove with the major axis of the sleeve at right angles to the tool. If the sleeve appears to be out of line after the press is started, open the tool, re-center the sleeve, and complete the press. (1) Thimble-Eye Splice. Initially position the able so that the end will extend slightly berond the sleeve, as the sleeve will elongate omewhat when it is compressed. If the cable and is inside the sleeve, the splice may not hold he full strength of the cable. It is desirable hat the oval sleeve be placed in close proximty to the thimble points, so that when compressed the sleeve will contact the thimble as shown in figure 4.9. The sharp ends of the thimble may be cut off before being used; how- Chap 4 312 FIGURE 4.9.-Typical thimble-eye splice. ever, make certain the thimble is firmly secured in the cable loop after the splice has been completed. When using a sleeve requiring three compressions, make the center compression first, the compression next to the thimble second, and the one farthest from the thimble (2) Lap Splice. Lap or running splices may also be made with copper oval sleeves. When making such splices, it is usually necessary to Par 196 ### Page 144 (OCR) Page 104 AC 43.13-1 191 use two sleeves to develop the full strength of the cable. The sleeves should be positioned as shown in figure 4.6b, and the compressions made in the order shown. As in the case of eye sleeve stocksplices, it is desirable to have the cable ends number is extend beyond the sleeves sufficiently to allow atamped here sleeves. for the increased length of the compressed Gauge (3) Stop Sleeves. Stop sleeves may be used for special cable end and intermediate fittings and they are installed in the same manner as Nicopress oval sleeves. NoTE: All stop sleeves are plain copper-certain sizes are colored for identification. (4) Terminal Gauge. To make a satisfactory copper sleeve installation, it is important that the amount of sleeve pressure be kept uniform. The completed sleeves should be checked periodically with the proper gauge. Hold the gauge so that it contacts the major axis of the sleeve. The compressed portion at the center of the sleeve should enter the gauge opening with very little clearance, as shown in figure 4.11. If it does not, the tool must be adjusted accordingly. (5) Other Applications. The preceding information regarding copper oval sleeves and stop sleeves is based on tests made with flexible aircraft cable. The sleeves may also be used on Cable Sleeve Sleeve Tested (Inch) sizo Sleeve No. Tool No. length (Inch) (Inch) O.D. strength (pounds FIGURE 4.11.-Typical terminal gauge. wire ropes of other constructon if each specifi type of cable is proof tested initially. Becaus of variation in rope strengths, grades, con struction, and actual diameters, the test is nec essary to insure proper selection of materials the correct pressing procedure, and an ade quate margin of safety for the intended use. **197. HAND FABRICATED CABLE ASSEMBLIES.** a. Woven Splice Terminal. The 5-tuck wove splice may be utilized on 7×7 flexible and ×19 extra-flexible cables of 8/82 inch dian ter or greater; however, this type of termina will only develop 75 percent of the cable strength. It should not be used to replace high efficiency terminals unless it is definitely determined the design load for the cable is not greater than 75 percent of the cable minimum breaking strength. In some cases it will be necessary to splice one end of the cable on assembly. For this reason, investigate the original installation for pulleys and fair-leads that might restrict the passage of the splice. The procedure for the fab- 871-12-B4 51-B4-887 7⅞a 280 rication of a woven splice is as follows. ⅓½: 871-1-C 871-17-J 51-C-887 51-MJ 525 600 (Refer to figure 4.12 for the designation of (Yellow) numbers and letters referred to in this se- 871-18-J 51-MJ Sio 800 quence of operations.) ½a 871-19-M 1% 871-20-M (Red) (Black) 51-MJ 51-MJ Sio 3%¾ 3%% 1, 600 1, 200 thimble, by means of a splicing clamp in a vise, with the free end to the left of the standing (1) Secure the cables around a bushing or ½2 871-22-M 51-MJ | 2, 300 wire and away from the operator. If a thimble 871-26-F6 871-23-16 3-F6-950 3-F6-950 3, 800 3, 500 is used as the end fitting, turn to point outward approximately 45°. NoTE: All stop sleeves are plain copper--certain sizes are colored for identification. FIGURE 4.10.-Copper stop sleeve data. 12) Select the free strand (1) nearest the standing length at the end of the fitting, and free this strand from the rest of the free ends. Next, insert a marlinspike under the first three Par 196 Chap 4 ### Page 145 (OCR) -TEMPORARY SERVING CORD - SEAM TIGEONE AND WATERPAOOF - 5 FULL TUCKS - → ‡= FIGURE 4.12.--Preparation of a woven cable splice. strands (A, B, and C) of the standing length nearest the separated strand of the free end und separate them momentarily by twisting he marlinspike. Insert the free strand (1) inder the three separated strands, through the ›pening created by the marlinspike. Pull the free end taut by means of pliers. (3) Unlay a second strand (2), located to the left of the first strand tucked, and insert this second strand under the first two standing strands (A, B). Loosen the third free length (8), located to the left of the first two, and insert it under the first standing strand (A) of the original three (section AA). (4) Remove the center or core strand (7) from the free end and insert it under the same standing strands (A, B). Temporarily secure the core strand to the body of the standing cable. Loosen the last free strand (6) located just to the right of the first (1) and tuck it under the last two strands (E, F) of the standing cable. Tuck the free strand (5) around standing strand (E). Tuck the free end (4) around the sixth standing strand (F) (see figure 4.12 section AA). Pull all strands snug Chap 4 Page 105 toward the end fitting with the pliers, This completes the first tuck. (5) Begin with the firat free strand (1) and work in a counterclockwise direction, tucking free strands under every other strand. After the completion of every tuck, pull the strands tight with pliers toward the end fitting. After the completion of the third complete tuck, cut in half the number of wires in each free strand. Make another complete tuck with the wires remaining. At the completion of the fourth tuck, again halve the number of wires in the free strands and make one final tuck with the wires remaining. Cut off all protruding strands and pound the splice with a wooden or rawhide mallet to relieve the strands in the wires. (6) Serve the splice with waxed linen cord. Start 1/4 inch from the end of the splice and carry the wrapping over the loose end of the cord and along the tapered splice to a point between the second and third tucks. Insert the end of the cord back through the last five wrappings and pull snug. Cut off the end, and if a thimble is used as an end fitting, bend down the points. Apply two coats of waterproofing to the cord, allowing two hours between coats. Carefully inspect the cable strands and splices for local failure. Weakness in a woven splice is made evident by a separation of the strand of serving cord. b. Wrap-Soldered Terminal. The wrap-soldered splice terminal shown in figure 4.13 may be utilized on flexible cables less than 8/32 inch in diameter and on nonflexible single strand (19 wire) cable. This type of terminal will develop only 90 percent of the cable strength and should not be used to replace high efficiency terminals, unless it is definitely known that the design load for the cable is not greater than 90 percent of the cable minimum breaking strength. The method of making the wrapped and soldered splice is as follows. (l) Use serving or wrapping wire made of commercial soft-annealed steel wire or commercial soft iron wire, thoroughly and smoothly tinned or galvanized. (2) Use half tin and half lead solder conforming to Federal Specification QQ-S-571. Par 197 ### Page 146 (OCR) Page 106 AC 43.13 The melting point of this solder varies from 320° to 890° F., and the tensile strength is approximately 5,700 pounds per square inch. (3) Use solder flux consisting of stearic acid (there should be no mineral acid present) and resin, with a composition of 25 to 50 percent resin. A warming gluepot to keep the flux in fluid state is desirable. (4) Before the cable is cut, solder the wires to prevent slipping. The preferred process is to tin and solder the cable thoroughly two to three inches by placing in a solder trough, finishing smooth with a soldering tool. The cable may be cut diagonally to conform to the required taper finish. (5) After being soldered and cut, the cable is securely bent around the proper size thimble, and clamped so that the cables lie close and fiat and the taper end for finish lies on the outside. If it is necessary to trim the taper at this point in the process, it is preferable that it be done by nipping. Grinding is permissible, provided a steel guard at least 3 inches long and 1/82 inch thick is placed between the taper end and the main cable during the operation; and that the heat generated from 1 grinding does not melt the solder and loos the wires. (6) Serving may be done by hand or n chine, but in either case each serving convo tion must touch the adjoining one and pulled tightly against the cable, with spa for permitting a free flow of solder and insp tion. (See figure 4.13a.) (7) Prevent drawing of the temper of a cable resulting from excessive temperature duration of applied heat. Use a soldering fil consisting of stearic acid and resin. The use, a flux, of sal-ammoniac or any other compou: having a corrosive effect is not acceptable. (8) Soldering is accomplished by immersi the terminal alternately in the flux and in t solder bath, repeating the operation until tho ough tinning and filling with solder under tl serving wire and thimble is obtained. The ter perature of the solder bath and place whe: terminal is withdrawn should not be abo 450° I. A soldering iron may be used in th final operation to give a secure and good-a] pearing terminal. Assure that the solder con pletely fills the space under the serving wil and thimble. A slightly hollowed cast-irc block to support the splice during solderin may help in securing the best results. The us of abrasive wheels or files for removing exce: solder is not recommended. (9) As an alternative process for makin terminals for nonflexible cable, the oxyacety (a) Soldered de arrested cable lene cutting method and the presolderin method (soldering before wrapping) are ac ceptable, but only under the following condi tions: (a) that the process of cutting securel; (b) Soldered & wripped table terestal with burblay welds all wires together; D (b) that the annealing of the cable doe 3/33 3-1/4 3/4 Plus or minus 1/39" 5/8 1/8 .020 i Thimble (A) Bashing (B) AN-100-3 Specification No A24-111-3 the end, not extend more than one cable diameter fron 3/4 3/4 1/8 0035 025 58 AN-100-1 AN-100-9 AN-1114 AN-111-5 (c) that no filing is done either before o 1/23 35/9 1-9/0 1-1/0 1-7/4 1-1/₴ 3710 025 005 109" AN-100-0 AH-100-7 AN-111-8 AN-111-7 after soldering; 41/2 1-1/2 1-1/4 1-14 035 159- AN-1008 AN-111-3 3/16 5-1/4 13/4 170 1/4 050 195 AN-100-10 AN-111-10 (d) that for protection during the opera 21/4 AN-100-18 AN-111-18 212 050 AN-100-14 tion of grinding the tapered end of the cable, & 25701 DISO AH-100 16 steel guard at least three inches in length and FIGURE 4.18.-Preparation of a wrapped soldered 1/82 inch thick should be placed between the terminal. taper and the main cable; Pur 197 Chap 4 ### Page 147 (OCR) AC 43,13-1A 1972 (e) that heat from grinding does not draw the temper of the cable. (10) Do not use wrap-soldered splice terminals ahead of the firewall, or in other fre zones, or in other locations where they might be subjected to high temperature. 198. CABLE SYSTEM INSPECTION. Aircraft cable aystems are subject to a variety of environmental conditions and forms of deterioration that ultimately may be easy to recognize as wire/strand breakage or the not-so-readily visible types of wear, corrosion, and/or distortion. The following data will aid in detecting the presence of these conditions: a. Cable Damage. Critical areas for wire breakage are those sections of the cable which pass through fairleads and around pulleys. Examine cables for broken wires by passing a cloth along the length of the cable. This will lean the cable for a visual inspection, and decet broken wires if the cloth snags on the able. When snags are found, closely examine e cable to determine the full extent of the mage. he absence of snags is not positive evidence at broken wires do not exist. Figure 4.14a wows a cable with broken wires that were not detected by wiping, but were found during a visual inspection. The damage became readily FIGURE 4.14.- Cable inspection technique. Chap 4 Page 107 apparent (figure 4.14b) when the cable was removed and bent using the technique depicted in figure 4.14c. NOTE: Tests by various aeronautical agencies have indicated that a few broken wires spread over the length of a cable will not result in a critical loss of strength. Obtain specific information regarding acceptable wire breakage limits from the manufacturer of the aircraft involved. (I) External Wear Patterns. Wear will normally extend along the cable equal to the distance the cable moves at that location and may occur on one side of the cable only or on its entire circumference. Replace flexible and nonflexible cables when the individual wires in each strand appear to blend together (outer INDIVIDUAL OUTER WIRES WORN MORE THAN 50% INDIVIDUAL OUTER WIRES WORN 40-50% (NOTE BLENDING OF WORN AREAS) INDIVIDUAL OUTER WIRES WORN LESS THAN 40% (WORN AREAS INDIVIDUALLY DISTINGUISHABLE ) FIGURE 4.15.-Cable wear patterns. Par 197 ### Page 148 (OCR) FIGURE 4.16.-Worn cable (replacement necessary). wires worn 40-50 percent) as depicted in figure 4.15. Actual instances of cable wear beyond the recommended replacement point are shown in figures 4.16 and 4.17. (2) Internal Cable Wear. As wear is taking place on the exterior surface of a cable, the same condition is taking place internally, particularly in the sections of the cable which pass over pulleys and quadrants. This condition (shown in figure 4.18) is not easily detected unless the strands of the cable are separated. Wear of this type is a result of the relative motion between inner wire surfaces. Under certain conditions the rate of this type wear can be greater than that occurring on the surface. (3) Corrosion. Carefully examine any cable FIGURE 4.17.- Worn cable (replacement recommended). Par 198 AC 43.13-1A 1972 FIGURE 4.18.--Internal cable wear. for corrosion that has a broken wire in a section not in contact with wear producing airframe components such as pulleys, fairleads, etc. It may be necessary to remove and bend the cable to properly inspect it for internal strand corrosion as this condition is usually not evident on the outer surface of the cable. Replace cable segments if internal strand rust or corrosion is found. Areas especially conducive to cable corrosion are battery compartments, lavatories, wheel wells, etc., where concentrations of corrosive fumes, vapors, and liquids can accumulate. NOTE: Check all exposed sections of cable for corrosion after a cleaning and/or metal-brightening operation has been accomplished in that area. An example of cable corrosion, attributable to battery acid, is shown in figure 4.19. b. Wire Splices. Standard manufacturing splices have been mistaken for defects in the cable because individual wire end splices were Chap 4 ### Page 149 (OCR) AC 43.13-1A 1972 FIGURE 4.19.-Corrosion. **WIRE SF** DIE MARKS FIGURE 4.20.--Manufacturer's wire splice. visible after assembly of a finished cable length. In some instances, the process of twisting outer strands around the core strand may also slightly flatten individual outer wires, particularly in the area of a wire splice. This flattening is the result of die sizing the cable and does not affect the strength of the cable. These conditions, as shown in figure 4.20, are normal and are not a cause for cable rejection. c. Cable Maintenance. Frequent inspections and preservation measures such as rust prevention treatments for bare cable areas will help to extend cable service life. Where cables pass through fairleads, pressure seals, or over pulleys, remove accumulated heavy coatings of corrosion prevention compound. Provide corrosion protection for these cable sections by lubricating with a light coat of graphite grease or general purpose, low-temperature oil. Chap 4 Page 109 CAUTION Avoid the use of vapor degreasing, steam cleaning, methylethylketone (MEK) or other solvents to remove corrosion-preventative compounds, as these methods will also remove the cable internal lubricant. d. Routing. Examine cable runs for incorrect routing, fraying, twisting, or wear at fairleads, pulleys, anti-abrasion strips, and guards. Look for interference with adjacent structure, equipment, wiring, plumbing, and other controls. Inspect cable systems for binding, full travel, and security of attaching hardware. Check for slack in the cable system by attempting to move the control column and/or pedals while the gust locks are installed on the control surfaces. With the gust locks removed, actuate the controls and check for friction or hard movement. These are indications that excessive cable tension exists. NoTE: If the control movement is stiff after maintenance operations are performed on control surfaces, check for parallel cables twisted around each other or cables connected in reverse. e. Cable Fittings. Check swaged treminal reference marks for an indication of cable slippage within the fitting. Inspect the fitting assembly for distortion and/or broken strands at the terminal. Assure that all bearings and swivel fittings (bolted or pinned) pivot freely to prevent binding and subsequent failure. Check turnbuckles for proper thread exposure and broken or missing safety wires/clips. f. Pulleys. Inspect pulleys for roughness, sharp edges, and presence of foreign material embedded in the grooves. Examine pulley bearings to assure proper lubrication, smooth rotation, freedom from flat spots, dirt, and paint spray. Periodically rotate pulleys, which turn through a small arc, to provide a new bearing surface for the cable. Maintain pulley alignment to prevent the cable from riding on the flanges and chafing against guards, covers, or adjacent structure. Check all pulley brackets and guards for damage, alignment, and security. (1) Pulley Wear Patterns. Various cable system malfunctions may be detected by analyzing pulley conditions. These include such discrepan- Par 198 ### Page 150 (OCR) Page 110 AC 43,13-1/ 1971 **EXCESSIVE CABLE TENSION PULLEY MISALIGNMENT** **PULLEY TOO LARGE FOR CABLE CABLE MISALIGNMENT** FROZEN BEARING MORMAL CONDITION FIGURE 4.21.-Pulley wear patterns. cies as too much tension, misalignment, pulley bearing problems, and size mismatches between cables and pulleys. Examples of these conditions are shown in figure 4.21. g. Fairleads, Guides, Anti-Abrasion Strips. Inspect fairleads for wear, breakage, alignment, cleanness, and security. Examine cable routing at fairleads to assure that deflection angles are no greater than 3° maximum. Determine that all guides and anti-abrasion strips are secure and in good condition. h. Pressure Seals and Seal Guards, Examine pressure seals for wear and/or material deterioration. Determine that the seal guards are positioned to prevent jamming of a pulley if a pressure seal fails and pieces slide along the 199. CABLE TENSION ADJUSTMENT. Carefully adjust control cable tension in accordance with the airframe manufacturer's recommendations. Par 198 On large aircraft, take the temperature of the immediate area into consideration when using a tensiometer. For long cable sections, use the average of two or three temperature readings to obtain accurate tension values. If necessary, compensate for extreme surface temperature variations that may be encountered if the aircraft is operated primarily in unusual geographic or climatic conditions such as arctic, arid, or tropic locations. Use rigging pins and gust locks as necessary to assure satsifactory results. At the completion of rigging operations, check turnbuckle adjustment and safetying in accordance with Section 2 of this chapter. 200. CORROSION AND RUST PREVENTION. To insure a satisfactory service life for aircraft control cables, use a cable lubricant to reduce internal friction and prevent corrosion. Loose rust and surface corrosion may be removed with a stainless steel brush, being careful not Chap & ### Page 151 (OCR) AC 43.19-1А Page 111 1972 to damage the cable. Care should be taken to remove all residue from the cable strands prior to rust prevention treatment. If the cable is made from tinned steel, coat the cable with rust preventive oil and wipe off any excess. It should be noted that corrosion-resistant steel cable does not require this treatment for rust prevention. **201.-211. RESERVED.** Chap 4 Par 200 ### Page 152 (OCR) Page 112 AC 43.13-1A 1972 Section 2. SAFETY METHODS FOR TURNBUCKLES 212. GENERAL. Safety all turnbuckles with safety wire using either the double or single wrap method, or with any appropriately approved special safetying device complying with the requirements of FAA Technical Standard Order TSO-C21. The swaged and unswaged turnbuckle assemblies are covered by AN Standard Drawings. For safety wire sizes and materials, refer to figure 4.22. Do not reuse safety wire. Adjust the turnbuckle to the correct cable tension so that no more than three threads are exposed on either side of the turnbuckle barrel. Do not lubricate turnbuckles. 213. DOUBLE WRAP METHOD. O1 the methods using safety wire for safetying turnbuckles, the method described here is preferred, although either of the other methods described is satisfactory. The method of double wrap safetying is shown in figure 4.23(A). Use two separate lengths of the proper wire (see figure 4.22). Run one end of the wire through the hole in the barrel of the turnbuckle and bend the end of the wire towards opposite ends of the turnbuckle. Then pass the second length of the wire into the hole in the barrel and bend the ends along the barrel on the side opposite the first. Spiral the two wires in opposite directions around the barrel to cross each other twice between the center hole and the ends. Then paas the wires at the end of the turnbuckle in opposite directions through the holes in the turnbuckle eyes or between the jaws of the turnbuckle fork, as applicable, laying one wire along the barrel and wrapping the other at least four times around the shank of the turnbuckle and binding the laid wires in place be fore cutting the wrapped wire off. Wrap th remaining length of safety wire at least for turns around the shank and cut it off. Repe. the procedure at the opposite end of the turi buckle. When a swaged terminal is being safetied, pass Cable size Type of wrap Diameter of safety wire Material (ennes ed condition) the ends of both wires, if possible, through the hole provided in the terminal for this purpose and wrap both ends around the shank as de- 710- - Single_ 0.040_. Copper, brass.' Single.. 0.040. Copper, brass.! scribed above. When the hole in the terminal is %- - Single.. 0.040_. Stainless steel, not large enough to accommodate the ends of Monel and both wires, the hole may be enlarged in accord- 'K" Monel. ance with note 2 of figure 4.22 and the safety- %- %- - Single. Double. 0.057 min 0.040- Copper, brass,! Copper, brass.! ing completed as described above. If the hole is %e and Double 0.040_ Stainless steel, not large enough to allow passage of both greater. Monel and wires, pass the wire through the hole and loop 'K" Monel,' it over the free end of the other wire, and then %% and Singie_. 0.057 min.. Stainless steel, wrap both ends around the shank as described. greater. Monel or "K" Monel.' %% and Double_ 0.051 * Copper, braas. a. Another satisfactory double wrap method greater. is similar to the above, except that the spiraling of the wires is omitted as shown in figure I Galvanized or tinned steel, or soft tron wires are also acceptable. 4.28 (B). • The safety wire holes in ⅜s-inch diameter and larger turnbuokle terminals for swaging may be drilled auficiently to accommodate the double 0.051-Inch b. The wrapping procedures described and dlameter copper or brass wires when used. shown on MS 38591 may be used in lieu of the FIGURE 4.22-Turnbuckle safetying guide. safetying method shown herein. Par 212 Chap & ### Page 153 (OCR) 43.13-1A Page 113 71 4 TURN (MIN) **THIS APPLIES TO ALL** TURNBUCKLE WRAPPINGS (A) DOUBLE WRAP (SPIRAL) (B) DOUBLE WRAP SMENKE (C) SINGLE WRAP (SPIRAL) (D) SINGLE WRAP FIGURE 4.23.--Safetying turnbuckles. Chap 4 Par 219 ### Page 154 (OCR) 214. SINGLE WRAP METHOD, The single wrap methods described in the following paragraphs and as illustrated in figures 4.28 (C) and (D) are acceptable but are not the equal of the double wrap methods. a. Pass a single length of wire through the cable eye or fork, or through the hole in the swaged terminal at either end of the turnbuckle assembly. Spiral each of the wire ends in opposite directions around the first half of the turnbuckle barrel so as to cross each other twice. Thread both wire ends through the hole in the middle of the barrel so that the third crossing of the wire ends is in the hole. Again, spiral the two wire ends in opposite directions around the remaining half of the turnbuckle, crossing them twice. Then, pass one wire end through the cable eye or fork or through the hole in the swaged terminals, in the manner described above, wrap both wire ends around the shank for at least four turns each, cutting off ====: AC 43.13-1 197 excess wire. This method is shown in figure 4.28 (C). b. Pass one length of wire through the center hole of the turnbuckle and bend the wire ends toward opposite ends of the turnbuckle. Then pass each wire end through the cable eye or fork, or through the hole in the swaged terminal and wrap each wire end around the shank for at least four turns, cutting off excess wire. This method is shown in figure 4.23 (D). After safetying, no more than three threads of the turnbuckle threaded terminal should be exposed. 215. SPECIAL LOCKING DEVICES. Several turnbuckle locking devices are available for securing turnbuckle barrels. Persons intending to use a special device must assure the turnbuckle assembly has been designed to accommodate such device. A typical unit is shown in figure 4.24. When special locking devices are nof readily available, the use of safety wire is acceptable. E== **TURNBUCKLE BODY -LOCKING CLIP MS21251 MS2I256** CLIP TYPE LOCKING DEVICE FIGURE 4.24.- Clip type locking device. **216.-226. RESERVED.** Par 214 Chap & ### Page 155 (OCR) C 43.13-1A Page 115 972 Chapter 5. AIRCRAFT HARDWARE Section 1. IDENTIFICATION AND USE OF AIRCRAFT HARDWARE 27. BOLTS. Most bolts used in aircraft strucures are either general purpose AN bolts, or IAS (National Aircraft Standard) internal frenching or close-tolerance bolts. In certain ases, aircraft manufacturers make up special olts for a particular application and it is necesary to use them or their equivalent in relacement. a. Identification. AN-type aircraft bolts can be dentified by the code markings on the bolteads. The markings generally denote the bolt nanufacturer, the material of which the bolt is nade, and whether the bolt is a standard AN- ype or a special purpose bolt. AN standard teel bolts are marked with either a raised ash or asterisk, corrosion-resistant steel is inlicated by a single raised dash, and AN alumium alloy bolts are marked with two raised ashes. The strength and dimensional details f AN bolts are specified on the Army/Navy Aeronautical Standard Drawings. Special purpose bolts include the hightrength and low-strength types, close-tolerance ypes; such bolts are normally inspected y magnetic, fluorescent or equivalent inspecmethods, Typical markings include 'SPEC (usually highly heat treated), an ircraft manufacturer's part number stamped n the head, or plain heads (low strength). Close-tolerance National Aircraft Standards (NAS) bolts are marked with either a raised r recessed triangle. The material markings or NAS bolts are the same as for AN bolts, xcept that they may be either raised or reessed. Bolts inspected magnetically (Magnalux) or by fluorescent means (Zyglo) are idenified by means of colored lacquer, or a head narking of a distinctive type. Figure 5.1 shows he typical coding used on aircraft boltheads. b. Grip Length. In general, bolt-grip lengths should equal the material thickness. However, bolts of slightly greater grip length may be used provided washers are placed under the nut or the bolthead. In the case of plate nuts, add shims under the plate. For proper washers, refer to paragraph 231. c. Locking or Safetying of Bolts. Lock or safety all bolts and/or nuts, except self-locking nuts. Do not reuse cotter pins and safety wire. d. Bolt Fit. Many boltholes, particularly those in primary connecting elements, have close tolerances. Generally, it is permissible to use the frst lettered drill size larger than the normal bolt diameter, except where the AN hexagon bolts are used in light-drive fit (reamed) applications and where NAS close-tolerance bolts or AN clevis bolts are used. Boltholes are to be normal to the surface involved to provide full bearing surface for the bolthead and nut, and not be oversized or elongated. In case of oversized or elongated holes in critical members, obtain advice from an FAA inspector or engineer, or the aircraft manufacturer before drilling or reaming the hole to take the next larger bolt, as usually, items such as edge distance, clearance, etc., must be considered. e. Torques. The importance of correct applicacation can not be overemphasized. Undertorque can result in unnecessary wear of nuts and bolts as well as the parts they are holding together. When insufficient pressures are applied, uneven loads will be transmitted throughout the assembly which may result in excessive wear or premature failure due to fatigue. Overtorque can be equally damaging because of failure of a bolt or nut from overstressing the threaded areas. There are a few simple, but very important, procedures that should be followed to assure that correct torque is applied: hap S Par 227 ### Page 156 (OCR) Page 116 AC 43.13- AIR E дО}> ~ M •cIt **AN STANDARD AN STANDARD AN STANDARD AN STANDARD AN STANDARD** **STEEL BOLT STEEL BOLT STEEL BOLT STEEL BOLT STEEL BOLT (CORROSION RESISTANT) SPEC** **AN STANDARD AN STANDARD AN STANDARD AN STANDARD SPECIAL BOLT** STEEL BOLT STEEL BOLT STEEL BOLT STEEL BOLT 63-59131 SPECIAL BOLT DRILLED HEAD BOLT SPECIAL BOLT TOLERANCE NA S CLOSEOLT (2024) BOLT ALUMINUM ALLOY MAGNETICALLY ORANGE- DYED CLEVIS BOLT REWORKED BOLT LOW STRENGTH INSPECTED MAGNETIC INS. MATERIAL BOLT FIGURE 5.1.-Bolt identification. (1) Calibrate the torque wrench periodically to assure accuracy; and re-check frequently. (2) Be sure that bolt and nut threads are clean and dry (unless otherwise specified by the manufacturer). (8) Run nut down to near contact with the washer or bearing surface and check &friction drag torque" required to turn the nut. (4) Add the friction drag torque to the desired torque recommended by the manufacturer, or obtain desired torque as shown in figure 5.2. This is referred to as final torque Par 227 which should register on the indicator or the setting for a snapover type wrench. (b) Apply a smooth even pull when applying torque pressure. If chattering or a jerking motion occurs during final torque, back off and re-torque. (6) When installing a castle nut, start alignment with the cotter pin hole at minimum recommended torque, plus friction drag torque, and do not exceed maximum plus friction drag. If the hole and nut castellation do not align, change washers and try again. Exceeding the maximum recommended torque is not recommended. Chap 5 ### Page 157 (OCR) - 43.13~1A 172 (7) If torque is applied to capscrews or theads, apply recommended torque plus fricon drag torque as determined in step (8). (8) If special adapters are used which ill change the effective length of the torque rench, the final torque indication or wrench etting must be adjusted accordingly. Deterine the torque wrench indication or setting ith adapter installed as shown in figure 5.8. igure 5.2 is a composite chart of recomended torque to be used when specific orque is not recommended by the manufacurer. The chart includes standard nut and bolt ombinations currently used in aviation main- ex-head aircraft bolt is an all-purpose strucural bolt used for general applications involvug tension or shear loads. Alloy steel bolts maller than No. 10-32 and aluminum alloy olts smaller than 1/4-inch diameter are not to e used in primary structure. Do not use aluinum alloy bolts and nuts where they will be epeatedly removed for purposes of mainteance and inspection. Aluminum alloy nuts nay be used with cadmiumplated steel bolts oaded in shear on land airplanes, but are not o be used on seaplanes due to the possibility of lissimilar metals corrosion. g. Close-Tolerance Bolts (AN-173 through AN- 86 (Hex-Head), NAS-80 through NAS-86 (100° countersunk)). Close-tolerance bolts are used in ugh-performance aircraft in applications here the bolted joint is subject to severe load eversals and vibration. The standard AN ex-head bolts may be used for the same appliations provided a light-drive fit is accomlished. h. Internal Wrenching Bolts (MS-20004 through S-20024 or NAS-495). These bolts are suitable or use both in tension and shear applications. n steel parts, countersink the bolthole to seat he large radius of the shank at the head or, as n aluminum alloys, use a special heat-treated vasher (NAS-143C) that fits the head to rovide adequate bearing area. A special heatreated plain washer (NAS-143) is used under he nut. Use special high-strength nuts on hap 5 Page 117 these bolts. (Refer to paragraph 280c(7).) Replace all internal wrenching bolts by another internal wrenching bolt. Standard AN hexhead bolts and washers cannot be substituted for them, as they do not have the required strength. I. Drilled-Head Bolts (AN-73). The AN drilledhead bolt is similar to the standard hex-bolt, but has a deeper head which is drilled to receive wire for safetying. The AN-8 and the AN-78 series of bolts are interchangeable for all practical purposes from the standpoint of tension and shear strengths. 228. SCREWS. In general, screws differ from bolts by the following characteristics: Usually lower material strength, a looser thread fit (No. 2), head shapes formed to engage a screwdriver, and the shank threaded along its entire length without a clearly defined grip. However, several types of structural screws are available that differ from the standard structural bolts only in the type of head. The material is equivalent and a definite grip is provided. The AN-525 washerhead screws, the AN 509-100° countersunk structural screws, and the NAS-204 through NAS-235 are such parts. The material markings are the same as those used on AN standard bolts. a. Structural Screws (NAS-204 through NAS- 235, AN-509 and AN-525). This type of screw, when made of alloy steel such as SAB-4180, ND-8680, or equivalent, and heat-treated from 125,000 p.s.i., may be used for structural assembly in shear applications similar to structural bolts. b. Self-Tapping Screws. The AN-504 and AN- 506 screws are used for attaching minor removable parts such as nameplates and the like. AN-580 and AN-581 are used in blind applications for the temporary attachment of sheet metal for riveting and the permanent assembly of nonstructural assemblies. AN-585 is a plain head self-tapping screw used in the attachment of nameplates or in sealing drainholes in corrosion-proofing tubular structures, and is not intended to be removed after installation. Never use self-tapping screws to replace standard Par 227 ### Page 158 (OCR) Page 118 AC 43.1 QuI Ble sion BOLT cel Tonsk tamina AN 3 thru AN 20 MB 20004 thru MB 20024 AN 3DD thru AN 20DD AN 42 thru AN 49 NAB 144 thru NAB 158 AN 173DD thru AN 186DD AN 78 thru AN 81 NAS 333 thru NAB 340 AN 509DD AN 173 thru AN 186 NAS 583 thru NAB 590 AN 525D M8 20033 thru M8 20046 NA8 624 thru NA8 644 MB 27039D MS 20073 NAB 1803 thru NAB 1320 M8 24694DD M8 20074 NAS 172 AN 500 NK9 NAS 174 M8 24694 NAS 517 AN 523 NK525 Steel shear boll M8 27030 NAH 106 NUTS NUTE NUTE Bleel Tension Steel Shear Steel Tenalon Steel Shear Aluminam Tension Aluminum Shear AN 310 AN 820 AN 810 AN 320 AN 365D AN 320D AN 315 AN 364 AN 315 AN 364 AN 310D AN 364D AN 363 NAS 1022 AN 383 NAS 1022 NA8 1021D NAS 1022D AN 365 MB 17826 AN 365 M8 17826 NAS 1021 MS 20384 M8 17825 M8 20364 MS 17825 M8 20365 MB 21045 M8 21045 MS 20365 NAS 1021 MB 20500 NA8 679 NAS 079 NAS 1291 FINE THREAD SERIES lorque Limit Torque Limit Torque Limita Torque Limits Forgus Limit Nat-bolt n.-lb in.-lbe fn.-Ibe in.-ibs sizo Min. Max. MiD. Mas. Min. Max. Min. MaI. Min. Max. Min. MAI. 8 -36 9 10-32 25 30 15 20 10 6-28 80 100 50 60 80 120 145 70 40 36-21 160 200 250 150 70 3r20 520 630 400 1-20 690 770 950 550 Kr-18 1, 000 1, 100 1, 300 800 % -18 1, 300 1, 250 1, 550 050 % -16 500 3, 650 3, 200 1, 900 76-14 3, 000 8, 550 4, 350 2, 600 1 -14 4, 500 4, 500 6, 500 3, 300 1% -12 7, 000 6, 000 7, 300 4, 400 1% -12 11, 000 11, 000 13, 400 8, 000 COARSE THREAD SERIES 8 -82 12 15 7 10 -24 20 12 15 * -20 40 25 80 ·-18 80 48 55 % -18 160 185 95 110 Kr14 285 255 140 155 ½ - 13 400 480 240 290 Kг12 500 700 300 420 % -11 700 900 420 540 4-10 1, 150 1, 600 700 950 36 - 9 2, 200 3, 000 1, 300 1, 800 1 - 8 8, 700 5, 000 2, 200 3, 000 1% - 8 5, 500 6, 500 8, 300 4,000 1¼ - 8 6, 500 8, 000 4, 000 5, 000 FIGURE 5.2.-Recommended torque values for nut-bolt combinations-- (without lubrication). Par 228 Chay ### Page 159 (OCR) 43.13-1A Page 119 **NOTE - LEGEND-** WHEN USING A TORQUE WRENCH ADAPTER WHICH T - ACTUAL (DESIREDI TORQUE CHANGES THE DISTANCE FROM THE TORQUE WRENCH Y= APPARENT (INDICATED) TORQUE DRIVE TO THE ADAPTER DRIVE, APPLY THE L. - EFFECTIVE LENOTH LEVER FOLLOWING FORMULAS TO OBTAIN THE CORRECTED = ENFECTIVE LENGTH OF EXTENSION TORQUE READING: Hort Open End Open-End Wrench Adapter CENTERLINE ADAPTER DRIVE WRENCH TORQUE WRENCH DRIVE TORQUE WRENCH HANDGRIP CENTERLINE Adapter **WRENCH CENTERLINE ADAPTER (PREDETE MINED) DRIVE** CENTERLINE CENTERINE Set Screw Flare Nut Wrench Adapter Adapter EXAMPLE: (WITH "E" AS PLUS DIMENSIONI EXAMPLE: (ATTH "E" AS HINUS DUHENSION) T - 135 LE. IN. 8= 10715-18-11790 T - 135 LD. 04. r- 123-18 - 125-150.82 Hose Clamp Y - UNKNOWN Y - UNKNOWN Spanner Wrench Adapter L - 10.0 IN. 1 - 10.0 MN. Adapter € = 1.5 IN. Y - 117 LB. AN. 1 - 1.5 HN. Y - 159 18. 04. FIGURE 5.3.-Torque wrench with various adapters. crews, nuts, bolts, or rivets in the original pructure. 129. TAPER PINS (AN-385 AND AN-386), Plain und threaded taper pins are used in joints which carry shear loads and where absence of ›lay is essential. The plain taper pin is drilled and usually safetied with wire. The threaded aper pin is used with a taper-pin washer (AN-975) and shear nut (safetied with cotter ›in) or self-locking nut. a. The Flathead Pin (MS-20392), Commonly 'alled a clevis pin, the flathead pin is used in conjunction with tie rod terminals and in sec- ›ndary controls which are not subject to coninuous operation. The pin is customarily intalled with the head up so that if the cotter ›in fails or works out, the pin will remain in lace. b. The AN-380 Cotter Pin. This is used for safetying bolts, screws, nuts, other pins, and in arious applications where such safetying is recessary. Use AN-381 cotter pins in locations where nonmagnetic material is required or in ocations where resistance to corrosion is deired. hap 3 **230. NUTS.** a. Self-Locking Nuts. Self-locking nuts are acceptable for use on certificated aircraft subject to the restrictions on the pertinent manufacturer's recommended practice sheets. Selflocking nuts are used on aircraft to provide tight connections which will not shake loose under severe vibration. Two types of selflocking nuts are currently in use, the all-metal type and the fiber or nylon lock type. Do not use self-locking nuts at joints which subject either the nut or bolt to rotation. They may be used with antifriction bearings and control pulleys, provided the inner race of the bearing is clamped to the supporting structure by the nut and bolt. Attach nuts to the structure in a positive manner to eliminate rotation or misalignment when tightening the bolts or screws. (1) All-metal locknuts are constructed with either the threads in the locking insert out-ofphase with the load-carrying section, or with a saw-cut insert with a pinched-in thread in the locking section. The locking action of the allmetal nut depends upon the resiliency of the metal when the locking section and loadcarrying section are engaged by screw threads. Par 228 ### Page 160 (OCR) (2) Fiber or nylon locknuts are constructed with an unthreaded fiber-locking insert held securely in place. The fiber or nylon has a smaller diameter than the nut, and when a bolt or screw is entered, it taps into the insert, producing a locking action. After the nut has been tightened, make sure the rounded or chamfered end bolts, studs, or screws extend at least the full round or chamfer through the nut. Flat end bolts, studs, or screws should extend at least 1/32 inch through the nut. When fiber-type self-locking nuts are reused, check the fiber carefully to make sure it has not lost its locking friction or become brittle. Do not reuse locknuts if they can be run up fingertight. Bolts 5/16-inch diameter and over with cotter pinholes may be used with self-locking nuts but only if free from burrs around the holes. Bolts with damaged threads and rough ends are not acceptable. Do not tap the fiberlocking insert. (3) Self-locking nut bases are made in a number of forms and materials for riveting and welding to aircraft structure or parts. Certain applications require the installation of self-locking nuts in channels, an arrangement which permits the attachment of many nuts with only a few rivets. These channels are track-like bases with regularly spaced nuts which are either removable or nonremovable. The removable type carries a floating nut, which can be snapped in or out of the channel, thus making possible the ready removal of damaged nuts. Nuts such as the clinch-type and spline-type which depend on friction for their anchorage are not acceptable for use in aircraft structures. (4) Self-locking nuts may be used on aircraft engines and accessories when their use is specified by the engine manufacturer in his bulletins or manuals. b. Aircraft Castle Nut (AN-310). The castle nut is used with drilled-shank AN hex-head bolts, clevis bolts, eye bolts, drilled-head bolts or studs, and is designed to accommodate a cotter pin or lockwire as a means of safetying. c. Miscellaneous Aircraft Nuts. (1) The plain nut (AN-315 and AN-385) has limited use on aircraft structures and re- Par 230 AC 43.13 quires an auxiliary locking device such as checknut or lockwasher. (2) Light hex-nuts (AN-840 and AN-345 are used in miscellaneous applications an must be locked by an auxiliary device. (3) The checknut (AN-316) is used as locking device for plain nuts, screws, threade rod ends, and other devices. (4) The castellated shear nut (AN-320) i designed for use with clevis bolts and threade taper pins, which are normally subjected t shearing stress only. (5) Wing nuts (AN-350) are intended fo we were the desiren i battes is onditiont obtained by the use of the fingers or han tools. (6) Sheet spring nuts, such as speed nuts are used with standard and sheet metal self tapping screws in nonstructural locations They find various uses in supporting lin clamps, conduit clamps, electrical equipment access doors, and the like, and are available i several types. (7) Two commercial types of high-strengt internal or external wrenching nuts are availa ble, the internal and external wrenching elastic stop nut and the Unbrako internal and externa wrenching nut. Both are of the self-lockin type, are heat-treated, and are capable of carry ing the high-strength bolt-tension load. 231. WASHERS. The types of washers used i aircraft structure are: plain washers, lock washers, and special washers. a. Plain washers (AN-960 and AN-970) ar widely used under hex nuts to provide smoothbearing surface, to act as a shim and t adjust holes in bolts. Use plain washers unde lockwashers to prevent damage to surfaces Cadmiumplated steel washers are recom mended for use under boltheads or nuts on alu minum alloy or magnesium structures wher corrosion, if it occurs, will then be between the washer and the steel. The AN-970 stee washer provides a greater bearing area than th plain type, and is used in wooden structure: under both boltheads and nuts to prevent loca crushing of the surface. Chap ! ### Page 161 (OCR) |43.13-1A b. Lock washers (AN-935 and AN-936) may be ed with machine screws or bolts whenever e self-locking or castellated type of nut is t applicable. They are not to be used as fasnings to primary or secondary structures, or here subject to frequent removal or corrosive nditions. c. Ball-socket and seat-washers (AN-950 and V-955) are used in special applications where e bolt is installed at an angle to the surface, ' when perfect alinement with the surface is quired at all times. These washers are used gether. d. Taper-pin washers (AN-975) are used with le threaded taper pin. e. NAS-143 washers are used with NAS inernal wrenching bolts and internal wrenching uts. Type "C" is countersunk to seat the boltead shank radius and a plain-type washer is sed under the nut. Both of these washers are eat treated from 125,000 to 145,000 p.s.i. **32. TYPES OF RIVETS.** a. Standard solid-shank rivets and the univeral head rivets (AN-470) are used in aircraft onstruction in both interior and exterior locaions. b. Roundhead rivets (AN-430) are used in the nterior of aircraft except where clearance is equired for adjacent members. c. Flathead rivets (AN-442) are used in the inerior of the aircraft where interference of adacent members does not permit the use of 'oundhead rivets. d. Brazierhead rivets (AN-455 and AN-456) tre used on the exterior surfaces of aircraft where flush riveting is not essential. e. All protruding head rivets may be replaced y MS-20470 (supersedes AN-470) rivets. This has been adopted as the standard for proruding head rivets in this country. f. Countersunk head rivets MS-20426 (supersedes AN-426 100° ) are used on the exterior surfaces of aircraft to provide a smooth aerodynamic surface, and in other applications where a smooth finish is desired. The 100° Chap S Page 121 countersunk head has been adopted as the standard in this country. **233, MATERIAL APPLICATIONS.** a, 2117-13 is the most commonly used rivet material utilized in aluminum alloy structures. Its main advantage lies in the fact that it may be used in the condition received without further treatment. b. The 2017-T3, 2017-731, and 2024-74 rivets are used in aluminum alloy structures where strength higher than that of the 2117-T8 rivet is needed. See Handbook MIL-HDBK-5 for differences between the two types of 17ST rivets specified here. c. The 1100 rivets of pure aluminum are used for riveting nonstructural parts fabricated from the softer aluminum alloys, such as 1100, 3003, and 5052. d. When riveting magnesium alloy structures, 5056 rivets are used exclusively due to their corrosion-resistant qualities in combination with the magnesium alloys. •. Mild steel rivets are used primarily in riveting steel parts. Do not use galvanized rivets on steel parts subjected to high heat. f. Corrosion-resistant steel rivets are used primarily in riveting corrosion-resistant steel parts such as firewalls, exhaust stack bracket attachments, and similar structures. g. Monel rivets are used in special cases for riveting high-nickel steel alloys and nickel alloys. They may be used interchangeably with stainless steel rivets as they are more easily driven. However, it is preferable to use stainless steel rivets in stainless steel parts. h. Copper rivets are used for riveting copper alloys, leather, and other nonmetallic materials. This rivet has only limited usage in aircraft. I. Hi-shear rivets are sometimes used in connections where the shearing loads are the primary design consideration. Its use is restricted to such connections. It should be noted that hishear rivet patterns are not to be used for the installation of control surface hinges and hinge Par 231 ### Page 162 (OCR) brackets. Do not paint the rivets prior to assembly, even where dissimilar metals are being joined. However, it is advisable to touch up each end of the driven rivet with zinc chromate primer to allow the later application of the general airplane finish. i. Blind rivets in the MS-20600 through MS- 20603 series rivets and the mechanically-locked stem NAS 1898, 1399, 1738, and 1739 rivets may be substituted for solid rivets in accordance with the blind rivet manufacturer's recommendations. They should not be used where the looseness or failure of a few rivets will impair the airworthiness of the aircraft. Design allowables for blind rivets are specified in MIL- HDBK-5, Metallic Materials and Elements for Flight Vehicle Structures. Specific structural applications are outlined in MS-33522. Non- Par 233 AC 43.13- structural applications for such blind rivets MS-20604 and MS-20605 are contained in MS 33557. 234. FASTENERS (COWL AND FAIRING). A nun ber of patented fasteners are in use on aircraf A variety of these fasteners are commel cially available and the manufacturer's recon mendations concerning the proper use of thes types of fasteners should always be considere in other than replacement application. 235. UNCONVENTIONAL ATTACHMENTS. Do no use unconventional or new attachment device in the primary structure unless approved by representative of the Federal Aviation Admin istration. **236.-246. RESERVED.** Chap s ### Page 163 (OCR) 43.13-1A Page 133 Chapter 6. CORROSION PROTECTION 7. TYPES OF CORROSION. Almost all metals ed in aircraft are subject to corrosion. The tack may take place over an entire metal surce, or it may be penetrating in nature, formg deep pits. It may follow grain boundaries its attack on metallic surfaces or it may penrate a surface at random. It may be accenated by stresses from external loads or existg in the metallic structure from lack of mogeneity or improper heat treatment. It is romoted by contact of the metals with marials that absorb water, such as wood, sponge abber, felt, dirt, surface film, etc. Corrosion is ferred to as the following types: a. Direct Surface Atfack, The most common pe of general surface corrosion results from rect reaction of metal surface with oxygen in e air. Unless properly protected, steel will ist and aluminum and magnesium will form rrosion products. The attack may be accelerted by salt spray or salt-bearing air, by inustrial gases, or by the aircraft engine exaust gases. b. Dissimilar Metals Corrosion. When two dismilar metals are in contact and are connected y an electrolyte (continuous liquid or gas ath-salt spray, exhaust gas, condensate) acelerated corrosion of one of the metals may ccur. The most easily oxidized surface beomes the anode and corrodes. The less active ember of the couple becomes the cathode of he galvanic cell. The degree of attack depends n the relative activity of the two surfaces; the reater the difference in activity, the more seere the attack. The materials listed in Group I re quite active and corrode easily. They reuire maximum protection. Group IV materials re the least active and therefore require minium protection. Except as noted below, whenver metals from two different groups are in ontact with each other, special protection is equired to assure that dissimilar metal corrosion does not occur. Although aluminum alloys and tin are in different groups than magnesium, tin and the 5000 and 6000 series aluminum alloys may each be used in contact with magnesium without such protection. Tin may also be used with all aluminum alloys without special protection. (1) Group I. Magnesium and its alloys. (2) Group II. All aluminum alloys, cadmium, zine. (a) Subgroup A. 1100, 3003, 5052, 6061, 220, 855, 856. All clad alloys. (b) Subgroup B. 2014, 2017, 2023, 7075, 195. Under severe corrosive conditions, the subgroups should be considered as dissimilar metals insofar as corrosion protection is concerned. This is particularly true when a large area of an alloy of Subgroup B is in contact with a small area of Subgroup A. Severe corrosion of the alloy from Subgroup A may be expected. (3) Group III. Iron, lead, and tin and their alloys (except stainless steels). (4) Group IV. Stainless steels, titanium, chromium, nickel, and copper and their alloys, graphite (including dry film lubricants containing graphite). c. Pitting. While pitting may occur in any metal, it is particularly characteristic of passive materials such as the alloys of aluminum, nickel, and chromium. It is usually a localized breakdown of protection and may be due to a lack of homogeneity in the alloy surface, either from mechanical working or faulty heat treatment. It may also be due to an inclusion or rough spot in the metal surface or from localized contamination that breaks down the surface protection. Pitting takes place at random with no selective attack along grain boundaries. Isolated areas become anodic to the rest of the surface. Corrosion products formed accentuate the anodic characteristics in the pit hap 6 ### Page 164 (OCR) area, and deep penetrating attack develops rather than a general surface attack. d. Intergranular Corrosion. Selective attack along the grain boundaries of metal alloys is referred to as intergranular corrosion. It results from lack of uniformity in the alloy structure. It ia particularly characteristic of precipitation hardened alloys of aluminum and some stainless steels. Aluminum alloys 2024 and 7075 which contain appreciable amounts of copper and zinc respectively are highly vulnerable to this type of attack if not quenched rapidly during heat treatment or given other special treatment such as the T78 temper condition for the 7075 alloys. Aluminum extrusions and forgings in general may contain nonuniform areas, which in turn may result in galvanic attack along the grain boundaries. This type of corrosion is difficult to detect in its original stage although ultrasonic and eddy current inspection methods are being used. When attack is well advanced, the metal may blister or delaminate. This is referred to as "exfoliation." e. Stress Corrosion. This results from the combined effect of static tensile stresses applied to a surface over a period of time under corrosive conditions. In general, cracking susceptibility increases with stress, particularly at stresses approaching the yield point, and with increasing temperature, exposure time, and concentration oi corrosive ingredients in the surrounding environment. Aluminum alloy bellcranks employing pressed-in taper pins, landing gear shock struts with pipe thread-type grease fittings, clevis joints, and shrink fits are examples of parts which are susceptible to stress corrosion cracking. f. Corrosion Fatigue. Corrosion fatigue is a type of stress corrosion resulting from cyclic stresses on a metal in corrosive surroundings. Corrosion may start at the bottom of a shallow pit in the stressed area. Once attack begins, the continuous flexing prevents the repair of protective surface coating or oxide films and additional corrosion takes place in the area of stress. It is difficult to detect this type of attack in advance except as cracking develops. Par 247 AC 43.1 g. Fretting, Fretting corrosion is a lim type of attack that develops when relative tion of small amplitude takes place betw close fitting components. The rubbing con destroys any protective film that may be p ent on the metallic surface and addition removes small particles of virgin metal f the surface. These particles act as an abra: and prevent the formation of any protec oxide film and exposes fresh active metal to atmosphere. If the contact areas are small. sharp, deep grooves resembling brinnell me ings or pressure indentations may be worr the rubbing surface. As a result, this type corrosion has also been called false brinnell when developed on bearing surfaces. 248. CORROSION PROTECTION MEASURES I BASIC MATERIALS. In the repair or alteratior aircraft, apply corrosion proofing of the st type or equivalent to that originally appl unless the repair or alteration would resul increased susceptibility to corrosion, in wh case use additional corrosion protection me ures. Aluminum and magnesium alloys n be protected by a variety of surface tre ments. Most aluminum alloy structural surfa have been electrically anodized in chromicfuric acid tanks before fabrication. Magnesi sheet has received cold nitric-chromic a treatment; magnesium castings have usua received hot tank processing in fluoride dich mate mixtures. Steel surfaces may have b oxidized or plated during manufacture. a. Anodizing and Related Processes. In a dizing, aluminum alloys are placed in an el trolytic bath causing a thin film of alumin oxide to form on the surface of the alumin: This is resistant to corrosion and affords a g‹ paint base. Other processes which do not p vide as good a corrosive protection as anodiz are, however, good paint bases. These p cesses are: (1) Alkaline cleaning followed by chror acid dip. (2) Alcoholic phosphoric acid cleaner. (3) Alkaline dichromate treatment. b. Plating. Steels are commonly plated w other metale to prevent corrosion. Plating Chal ### Page 165 (OCR) 3.13-1A mplished by placing the article in an elecytic bath and metal from the plating solu- I is deposited on it. Various metals used in ting vary in the corrosion protection they ord steel. For instance, cadmium and zinc rode before the steel; hence, slight breaks cracks through the plating of these metals I not result in rusting of the exposed steel, ce the surface metal is corroded and prots the steel. Chromium does not protect steel this method, as steel will corrode before the omium and thus depends on the tightness of plating for its protection. The process of tplate bake treatment to relieve hydrogen brittlement is a necessary part of replating ›cedures for high-strength steel parts. gh-strength nuts and bolts are highly sus- ›tible to failure from hydrogen embrittlent, at the many stress risers and notches in ir design, through normal usage of susned and dynamic stresses. Because of the poitial failures of embrittled parts, careful trol over the heat treatment, grinding, preite cleaning, plating, and postplate baking of rh-strength parts is necessary. c. Sherardizing. Steel parts may be sherardd by heating them in an atmosphere of zine wder; however, this is not considered to be effective against corrosion as zinc or cadum plating. d. Phosphate Rust-Proofing, This process is mmercially known as Parkerizing, Bonderiz- 3, Granodizing, etc. The coating placed on e part is used to protect steel parts after chining and before painting. e. Chrome-Pickle Treatment. Magnesium parts lich have been immersed or brushed with a lution of nitric acid and sodium dichromate Il be protected for temporary storage. The ating also will serve as a bond for subseent organic finishes. Sealed chrome-pickle eatment is used on magnesium parts for long rm protection. Diluted chromic acid is a uch-up treatment. It is less critical to apply d can be applied over previously applied thin romate films. f. Dichromate Treatment. The dichromate treatent consists of boiling magnesium parts in a ap 6 Page 125 solution of sodium dichromate. This treatment provides good paint base and protective qualities on all standard wrought magnesium alloys except the magnesium-thorium alloys HK 31A, HM 21A, and HM 31A. No coating forms on these alloys. Acid pickling of the magnesium surface prior to application of the dichromate treatment is required if maximum corrosion resistance of the finish is expected. g. Stannate Immersion Treatment. This treatment deposits a layer of tin. It is a protective paint base for magnesium alloy parts which contain inserts and fasteners of a dissimilar metal such as brass, copper, or steel. This treatment cannot be used with parts containing aluminum inserts or fasteners because the high alkalinity of the bath attacks the aluminum. h. Galvanic Anodizing Treatment. This is an electrolytic process used to provide a paint base and corrosion preventive film on magnesium alloys containing manganese. I. Cladding, Aluminum alloys which are susceptible to corrosion are frequently clad with pure aluminum. Slight pits, scratches, or other detects through the cladding material will not result in corrosion of the core, since the pure aluminum on the edges of the defect will be preferentially corroded, protecting the core. I. Metal Spraying. Metal is melted and sprayed on the surface to be protected. The surface must be properly prepared and thoroughly cleaned to prevent peeling of the sprayed coat. k. Shot-Peening. Shot-peening and other treatments by which the surface can be placed in compression are effective in preventing stress corrosion. 1. Organic Coatings. Zinc chromate primer, enamels, chlorinated rubber compounds, etc., are organic coatings commonly used to protect metals. m. Dopeproofing. When doped fabrics are applied over an organic finished metal structure, the dope will have a tendency to loosen the finish on the metal. For this reason, organic coatings on the metal are usually covered Par 248 ### Page 166 (OCR) with a dopeproof paint, with metal foil, or with cellulose tape to prevent the dope from striking through. n. Tube Interiors. Protect the interiors of structural steel and aluminum tubing against corrosion. A small amount of water entrapped in a tube can corrode entirely through the tube thickness in a short period of time. Coat the tube interior by flushing with hot linseed oil, paralketone, or other corrosion inhibitor. The flushing liquid is usually introduced through small holes drilled in the tubing. Allow the flushing liquid to drain and plug the holes with a screw or by other means to prevent entry of moisture. Air and watertight sealing of the tubing will also give adequate protection against corrosion if the tubing is internally dry before being sealed. 249. CONSTRUCTION OF CORROSION RESISTANT STRUCTURE. During fabrication, the use of such practices as listed below will reduce the sources of corrosion initiation. o. Use combinations of metals which are as close together as possible in the galvanic series -do not use small aluminum clips on large stainless steel webs; use stainless steel fasteners in stainless steel assemblies. b. Insulate dissimilar metals but do not use materials which absorb moisture or conduct it in "wick" fashion. c. Paint cut edges. d. Seal edges of faying surfaces on butt joints. e. Provide adequate drainholes. 250. CORROSION REMOVAL. When corrosion is detected, remove it as quickly as possible and protect the surface from further corrosion. In those cases where the corrosion has progressed extensively, the strength of the part may be in jeopardy and the part may need to be replaced. a. For steel parts, except for highly stressed steel or stainless steel surfaces, the use of abrasive papers, buffers, hand wire brushing, and steel wool are acceptable cleanup procedures. However, it should be recognized that it Par 248 AC 43 18 practically impossible to remove a.. c sion products from the bottom of small and crevices. As a result, a part once rusty ally corrodes again, more easily than it di first time. Removing corrosion products highly stressed steel parta requires ca handling. b. For aluminum parts, treatment include mechanical removal of as much of the c sion products as practicable, the inhibitic residual material by chemical means, folk by the restoration of permanent surface ings. Do not use steel wool, emery cloth, (except stainless steel) wire brushes, or st abrasive materials because particles of steel wool or emery cloth will become en ded in the soft material and cause corro Hand polishing corroded areas with house abrasives or with metal polish available u Specification MIL-P-6888 is acceptable fo: on clad aluminum but must not be used or odized aluminum since it is severe enoug actually remove the protective anodized If the surface is particularly difficult clean. Type II material under Specifica MIL-C-5410 mixed 50-50 percent with sol or mineral spirits may be used. Treat any perficial corrosion with a 10 percent solutic sodium dichromate to which one percen chromium trioxide has been added or material available under Specification N C-5541. Allow these solutions to remair the corroded area for 5 to 20 minutes then wipe the surface dry. A more se cleaning procedure consists of using a 10 cent solution of chromic acid to which has added approximately 20 drops of battery trolyte per gallon. Thorough brushing wi stiff fiber brush should loosen or remove ! existing corrosion and assure complete I tration of the inhibitor into crevices and Allow the chromic acid to remain in place at least 5 minutes, then remove the exces. flushing with water or wiping with a cloth. Apply a protective coating the same that the corrosion treatment is accomplis c. When an anodized surface coating is d aged, it can only be partially restored by chi cal surface treatment. Therefore, exercise CH ### Page 167 (OCR) aning process to avoid unnecessary f the protective film, particularly at of the aluminum sheet. Chromic acid inhibitive treatments tend to restore ilm. severe cleaning is necessary with inr corrosion (attack along grain s). The mechanical removal of all products and visible delaminated rs must be accomplished in order to the extent of the destruction and to he remaining structural strength of onent. Inspection with a 5- to 10- gnifying glass or the use of dye peneassist in determining if all unsound I corrosion products have been reinding to blend or fair out the edges ed areas can best be accomplished by minum oxide-impregnated rubberels. Chemically inhibit the exposed und restore chemical surface films or he same manner as for other alumiices. esium is the most chemically active of s used in aircraft construction and is the most difficult to protect. The ad complete correction of the coating imperative if serious structural dambe avoided. Treat the corroded area ercent chromic acid solution to which added approximately 20 drops of batrolyte per gallon, in the same manaluminum alloys. ROSION PROOFING OF LANDPLANES LANES. In the repair or alteration of use corrosion proofing materials the or equivalent to, that originally apss the repair or alteration would recreased susceptibility to corrosion, in e, employ additional corrosion protecures. ROSION PROOFING OF LANDPLANES D TO SEAPLANES. A special problem tered in the conversion of landplanes es. In general, landplanes do not reosion proofing to the same extent as nes manufactured as such. Corro- Page 127 sion-proofing standards for landplanes converted to seaplanes are divided into two classes: (1) Necessary minimum precautions; and (2) Recommended precautions. Regardless of such precautions, it is imperative that the exterior surfaces of seaplanes be washed with clear fresh water immediately following extended water operation, or at least once a day when operated in salty or brackish water. Wash interior surfaces of seaplanes exposed to spray, taking care to prevent damage to electrical circuits or other items subject to injury. a. Necessary Minimum Precautions. The following procedures are considered the minimum to safeguard the airworthiness of the converted aircraft and are not in themselves intended to maintain airworthiness for an indefinite period. (1) Unless already protected, treat exposed fittings or fittings which can be reached through inspection openings with two coats of zinc chromate primer, paralketone, nonwatersoluble heavy grease, or comparable materials. This applies to items such as wing-root fitwing-strut fittings, control-surface hinges, horns, mating edges of fittings, and attach bolts, etc. (2) Coat nonstainless control cables with grease or paralketone or other comparable protective coating, if not replaced with corrosion-resistant cables. (3) Inspect all accessible sections of aircraft structure. Clean structural parts showing corrosion and refinish if corrosion attack is superficial. If a part is severely corroded, replace with an adequately corrosion-proofed part. b. Recommended Precautions. The recommended precautions are those which are suggested as a means of maintaining such aircraft in condition for safe operation over extended periods of time. (1) Provide additional inspection openings to assist in detecting corrosion. Experience has shown openings to allow inspection of lower and rearward portion of the fuselage to be particularly desirable. (2) Incorporate additional provisions for free drainage and ventilation of all interiors to Par 250 ### Page 168 (OCR) Page 128 AC 43.1: prevent collection of moisture (scoop type drain grommets). (3) Protect the interior of structural steel tubing. This may be done by air and watertight sealing or by flushing with hot linseed oil and plugging the openings. Inspect tubing for missing sealing screws, presence of entrapped water. local corrosion around sealing screws, welded clusters, and bolted fittings which may be indicative of entrapped moisture. (4) Slit the fabric of fabric covered aircraft longitudinally on the bottom of the fuselage and tail structure for access to these sections. Cont teracts any advantages in their speec fectiveness. Ise materials which are neutral and easy to remove. c. Removal of Spilled Battery Acid. to neutralize spilled battery acid, us bicarbonate (baking soda). or sodiu (borax) 20 percent by weight dissolved After neutralization, remove alkali : pletely with copious quantities of wate vent corrosion. An application of paint to the structure surrounding th may be an effective control for this tyl rosion. the lower structural members with zine chromate primer (two coats) ; follow by a coat of dope proof paint or wrap with cellophane tape and rejoin the fabric. This precaution is advisable within a few nonths after start of operation as a seaplane. (5) Spray the interior of metal-covered wings and fuselages with an adherent corrosion inhibitor. (6) Place bags of potassium or sodium dichromate in the bottom of floats and boat hulls to inhibit corrosion. * 254. HANDLING AND CARE OF AIRCRAF ERED FROM WATER IMMERSION. Aircr were recovered from partial or total i in water, including flash floods, have lowed to air dry, in certain instances. safety precantions other than a cursor tion of the aircraft exterior. The la adequate cleanup of water-immersed a subsequently adversely affect the safet aircraft. That is, water immersion inc probability of corrosive attack, the re lubrication, the deterioration of airc (7) Prevent the entry of water by sealing. as completely as possible, all openings in wings, fuselage, control-surface members, openings for control cables, tail-wherl wells, ete. 253. CLEANERS, POLISHES, BRIGHTENERS, 11 iS important that aircraft be kept thoroughly clean terials, and/or degradation of electi avionic equipment. Sea water, because of salt content, corrosive than surface fresh water. fresh water may contain varying amou and, as drying occurs, the salt concen increased and corrosive attack accelerat of deposits containing contaminating substances such as oil, grease, dirt, and other foreign materials. a. Materials. Avoid damage to aircraft by not using harmful cleaning, polishing, brightening, or paint-removing naterials. Use only those compounds whieh conform to existing government or established industry specifications or The most important factor following of an aircraft from sea or fresh water i is prompt action. Components of the which have been water immersed, suc poverplant, accessories, airframe section ing mechanisms, screws, bearings, worl faces, fuel and oil systens, wiring, ra radar should be disassembled. to the ex product- that have been specifically recommended by be a reraft manufacturer as being satisfactory for the intended application. Observe the product manufacturer's recommendations concerning de of his agent. b. Cherical Cleaners. Chemical cleaners must be used with preat care in cleaning assembled sidered necessary, so that the contamir be completely removed. a. Initial Fresh Water/Detergent Wc soon as possible after the aircraft is 1 from water immersion, thoroughly w taminated internal and external areas of craft using a water/detergent solution as aireraft. The danger of entrapping corrosive 01) Mix liquid detergent (MI- materials in faving surfaces and crevices cou- Type I) and isopropyl alcohol (TT-I Por 252 ### Page 169 (OCR) CHG T ight parts detergent to 20 parts of dd the detergent/alcohol mixture to tap water and mix thoroughly. For ne part of the foregoing concentrate ts of tap water (warm water if availnix thoroughly. the above specified detergent/alcohol re not available, use water emulsion ›mpound (MIL-C-43613). Add one und to nine parts water. If the MIL ompound is not available, use any ild household detergent solution with 'ater. Precautions. The following safety i should be observed: ctrically ground the aircraft. Attach lead to the aircraft at a point which the area that could contain explosive the landing gear of land planes is • supporting mechanism, install a ir, jury strut, landing gear downlocks itable devices to insure that the gear ollapse. If the landing gear is not insure that the aircraft is solidly to prevent hazardous movement. connect and remove wet- and/or dryes and isolate aircraft from all sources ty or other spark-producing devices. lucing static electricity is generated at air hose outlets, so this method should ed for ventilating or purging fuel nove all fuel, oil, and hydraulic fluid. sh all fuel and oil cells with clean late tires, especially on magnesium epressurize landing gear struts, pneums, and hydraulic accumulators. scating Engines and Propellers. The would be removed from the engine and from the aircraft. The exterior of the propeller should be washed with ot or cold fresh water. ›r accessories, engine parts, etc., should I and all surfaces flushed with fresh erably hot. If facilities are available, ed parts, size permitting, should be Page 128-1 immersed in hot water or hot oil, 180° F., for a short period of time. Soft water is preferred and should be changed frequently. All parts must be completely dried by air blast or other means. If no heat drying facility is available, wipe the cleaned parts with suitable drying cloths. Constant speed propeller mechanism should be disassembled, as required, to permit complete decontamination. Clean parts with steam or hot or cold fresh water. Dry the cleaned parts in an oven, but if a heat drying facility is not available, wipe the cleaned parts with suitable drying d. Gas Turbine, Turboprop and Turboshaft En- The engine exhaust shield, insulation blankets, separate exhaust collectors, compressor housing, pinion cowl and/or upper pinion housing should be removed. The engine accessories, outer housing exhaust shields, and other exposed parts should be steam cleaned. The steamcleaned parts should be dried in an oven at approxímately 200° F., or with hot air from a portable engine heater, or with clean wiping cloths. Immerse remainder of engine in 10-20 percent water solution of sodium dichromate, or hot, fresh water and apply sufficient agitation to provide complete flushing. If immersion cannot be accomplished, flood the lower section of the compressor housing and thoroughly flush rotor blades with sodium dichromate solution or warm fresh water. Seal off combustion chamber openings and alternately fill and drain the combustion chamber with fresh water. Dry clean parts in an oven, or use dry air from a portable engine heater or wipe with suitable drying cloths. e. Airframe. The salvageable components of the fuselage, wings, empennage, seaplane and amphibian hulls and floats, and movable surfaces should be processed as follows: (1) The fabric from fabric-covered surfaces should be removed and replaced. (2) The sircraft interior and exterior should be cleaned using steam under pressure with steam-cleaning compound. The steam should be directed into all seams and crevices where corrosive water may have penetrated. Avoid steam * Par 254 ### Page 170 (OCR) *cleaning electrical equipment, such as terminal boards and relays. (3) Areas that have been steam cleaned should be rinsed immediately with either hot or cold fresh water. (4) Touch up all scratches and scars on painted surfaces using zinc chromate primer or preservative. (5) Undrained hollow spaces or fluid entrapment areas should be provided temporary draining facilities by drilling out rivets at lowest point. Install new rivets after drainage. (6) All leather, fabric upholstery, and insulation should be removed and replaced. Plastic or rubber foam which cannot be cleaned of all to cleanes corrosive water should be replaced. lar structures should be removed and the structure blown out with compressed air. If corrosive water has reached the tubular interiors, carefully flush with hot, fresh water and blow out water with compressed air. Roll the structure as necessary to remove water from pockets. Fill the tubes with hot linseed oil at approximately 180° F. Drain oil and replace drain plugs or drive screws. (8) Clean sealed wood, metalite, and other nonmetallic areas, excluding acrylic plastics, with warm water. Wood, metalite, and other porous materials exposed to water immersion should be replaced, unless surfaces are adequately sealed to prevent penetration by corrosive water. Virtually all solvents and phenolic type cleaning agents are detrimental to acrylics and will either f. Helicopter Rofor Dynamic Components. (1) All evidence of corrosive water should be removed from the exterior of transmissions and gear boxes by flushing with clean hot or cold fresh water. (2) Where it is possible that corrosive water has reached the interior of the transmission or gear boxes, remove plugs and/or covers and drain completely. Flush interior of part with hot or cold fresh water. Drain residual water, replace plugs and/or covers, and reapply proper lubricant. g. Helicopter Blades. Except for blades with wooden or other nonmetallic constructions, treat Par 254 AC 43.1: helicopter blades the same as propelle ciprocating engines. Clean nonmetal by hand, using warm water. Dry wi cloths or a warm air blast. When clea copter rotor blades, insure that nonsen members, such as the spars and blad are cleaned and dried. h. Fuel and Oil Systems. Contamir and oil systems should be processed as : (1) Flush oil system, including li water-displacing preservative. (2) Open fuel systems. Purge cell lines. Check effectiveness of the pur combustible gas indicator. Remove bli self-sealing type fuel cells and all cal or pads. (3) Use clean, fresh water to wash After drying, spray interior walls witł (4) Chemically treat bare metal si cleaned aluminum tanks with 10-21 water solution of sodium dichromate. (5) Flush fuel lines with hot water maximum). Dry, using clean, dry, a air. 1. Landing Gear. Process salvageab nents of the landing gear, wheels, and follows: (1) Remove tires and wheels. (2) Steam clean wheels, rinse in fr and dry. (3) Remove wheel bearings and cl dry cleaning solvent. (4) Immerse bearings in methyl al dry the cleaned bearings in air blast. permit bearings to rotate during air bla Reapply proper lubricant. (5) Remove brakes, steam clean, : with fresh water and dry. I. Electrical Equipment. (1) Wet cell batteries. The risk in using wet cell batteries that have been in sea water may outweigh any ecor vantage and should be replaced. (2) Because of possible flight haz later defects caused by progressive cor tack, all electrical wiring immersed sive water should be replaced. If ### Page 171 (OCR) CHG 1 Page 128-3 shed or sprayed with corrosive water, ughly with clean, fresh water and dry, ressed air. Following compressed air it with water dispensing preservative 1309, Grade A). laneous Equipment. The following should be thoroughly washed to resalt, and other contaminants. Dry last or other means, and, if required, oper lubricant: ring and fuselage hydraulic units. (2) Electric landing gears. (3) Actuators. (4) Cables. (5) Aceumulators. (6) Hydraulic reservoirs. (7) Flight control. (8) Torque tubes and bell cranks. (9) Heating units, ducts, etc. * **255.-264. RESERVED.** Par 254 ### Page 172 (OCR) 13-14 Page 129 Chapter 7. IDENTIFICATION, TESTING, AND INSPECTION OF MATERIALS Section 1. IDENTIFICATION OF METALS IDENTIFICATION OF STEEL STOCK, The So- Numerala y of Automotive Engineers (SAE) and the Type of steels and digita erican Iron and Steel Institute (A.I.S.I.) a numerical index system to identify the Carbon Steels ---- 1xxx position of various steels. The numbers as- Plain Carbon Steels 10Xx Free Cutting Steels 11xx ed in the combined listing of standard Manganese Steels (Manganese 1.60 to 1.90%) 18xx la issued by these groups represent the Nickel Steels 2xXX ! of steel and make it possible to readily 8.50% nickel 28xx tify the principal elements in the material. 5.00% nickel 25xx he first digit of the four number designa- Nickel Chromium Steels indicates the type to which the steel be- 8.70% nickel, 0.07% chromium g. Thus "1" indicates a carbon steel, "2" a 1.25% nickel, 0.60% chromium 1.75% nickel, 2.00% chromium el steel, "¿" a nickel chromium steel, etc. 8.50% nickel, 1.50% chromium he case of simple alloy steels, the second Corrosion and heat resisting -. 80xXX t indicates the approximate percentage of Molybdenum Steels --. 40xx predominant alloying element. The last two Chromfum Molybdenum Steels -. SIZI ›on content. Thus the symbol "1020" indi- ts usually indicate the mean of the range of Nickel Chromium Molybdenum Steels Nickel Molybdenum Steels 1.75% nickel, 0.25% molybdenum 48xx 1ůxx s a plain carbon steel lacking a principal 8.50% nickel, 1.50% chromium 48×x ving element and containing an average of Chromium Steele БxXX percent (0.18 to 0.28) carbon. The symbol Low chromium 51xx 30" indicates a nickel steel of approxi- Medium chromium Corrosion and heat resisting 521XX 61xxx ely 8 percent (8.25 to 8.75) nickel and an Chromium Vanadium Steels - 6XІX *age of 0.30 percent, (0.28 to 0.88) carbon 1.00% chromium -- 61xx ent. The symbol "4180" indicates a chro- National Emergency Steels 8xxx m-molybdenum steel of approximately 1 Silicon Manganese Steels 9xxx 'ent (0.80 to 1.10) chromium, 0.20 percent 2.00% silicon -- 82xI 5 to 0.25) molybdenum, and 0.80 percent 8 to 0.83) carbon. The basic numbers for FIGURE 7.1.-Numerical system for steel identification. four digit series of the carbon and alloy Is may be found in figure 7.1. forming to Specification MIL-1-6786 are in- INTERCHANGEABILITY OF STEEL TUBING. terchangeable. "1025" welded tubing as per Specification -T-5066 and "1025" seamless tubing conning to Specification MIL-T-5066A are inhangeable. c. NE-8630 welded tubing conforming to Specification MIL-T-6784, and NE-8680 seamless tubing conforming to Specification MIL-T-6782 are interchangeable. , "4130" welded tubing as per Specification 267. IDENTIFICATION OF ALUMINUM. To pro- ~T-6781, and "4180" seamless tubing con- vide a visual means for identifying the various Par 265 ### Page 173 (OCR) Page 130 AC 4 grades of aluminum and aluminum alloys, such metals are usually marked with symbols such as Government Specification Number, the temper or condition furnished, or the commercial code marking. Plate and sheet are usually marked with specification numbers or code markings in rows approximately 5 inches Nonbeat-treatable Alloye Temper designation Defnition Annealed recrystallized (wrought products only) applies to softest temper of wrought producta. -H12 Strain-hardened one-quarter-hard temper. -H14 Strain-hardened half-hard temper. -H16 Strain-hardened three-quarters• hard temper. -H18 Strain-hardened full-bard temper, -H22 Strain-bardened and partially annealed to one-quarter-bard temper. -H24 Strain-hardened and partially annealed to balf-hard temper. -H26 Strain-hardened and partially annealed to three-quarters-hard temper. -1128 Strain-bardened and partially annealed to full-hard temper. -H82 Straln-hardened and then lized. Final temper la one-quarter -H84 Strain-hardened and then stabilized. Final temper is one-half -9186 Strain-bardened and then stabilized. Final temper is three quarters hard. -#88 Strain-bardened and then stabllized. final temper is full-hard. -H112 As fabricated; with specified mechanical property limita. For wrought alloys; as fabricated. No mechanical properties limits. For cast alloys; as cast. apart. Tubes, bars, rods, and extruded s are marked with specification numbers or markings at intervals of 8 to 5 feet alon length of each piece. The commercial marking consists of a number which iden the particular composition of the alloy. Ir dition, letter suffixes designate the follov Heat-trentable Alloya Temper designation Dennition -0 Annealed recryatallized producta only) applies to a temper of wrought products. -12 Annealed (castinge only). -T8 Solution beat-treated worked by the Battening atraightening operation. -T86 Solution heat-treated and worked by reduction of 6 per -T4 Solution heat-treated. -T42 Solutlon heat-treated by the regardieas of prior temper plicable only to 2014 and -T5 Artifcially aged only (cas only). -T6 Solution heat-treated and artific aged. -782 Solution heat-treated and age user regardless of prior te (applicable only to 2014 and alloys). -T851, -T451, Solution heat-treated and stres -T8510, -78511 lieved by stretching to produ -T4510, -T4511. permanent set of 1 to 8 per depending on the product. -T651, -T851, Solution heat-treated, stress rel! -76510, -T8510, by stretching to produce a I -78511, -78511. anent set of 1 to 8 percent, artificially aged. -T852 Solution heat-treated, compress produce a permanent set and artificially aged. -T81 Solution heat-treated, cold-wo by the fattening or straighte operation, and then artific aged. -T86 Solution heat-treated, cold-wo by reduction of 6 percent, and artificially aged. For wrought alloys; as fabric No mechanical properties Ill For cast alloys; as cast. FIGURE 7.2.-Basic temper designations and subdivisions for aluminum alloys. Par 267 ### Page 174 (OCR) 3.13-1A Clad aluminum alloys have surface layers of e aluminum or corrosion-resistant alumin alloy bonded to the core material to inpit corrosion. Presence of such a coating may determined under a magnifying glass by exination of the edge surface which will show ree distinct layers. b. Test for Distinguishing Heat-Treatable and inheat-Treatable Aluminum Alloys. If for any ason the identification mark of the alloy is t on the material, it is possible to distinguish tween some heat-treatable alloys and some nheat-treatable alloys by immersing a same of the material in a 10 percent solution of ustic soda (sodium hydroxide). Those heateated alloys containing several percent of ap 7 Page 131 copper (2014, 2017, and 2024) will turn black due to the copper content. High copper alloys when clad will not turn black on the surface, but the edges will turn black at the center of the sheet where the core is exposed. If the alloy does not turn black in the caustic soda solution it is not evidence that the alloy is not heat-treatable, as various high strength heattreatable alloys are not based primarily on the use of copper as an alloying agent. These include among others 6058, 6061, and 7075 alloys. The composition and heat-treatability of alloys which do not turn black in a caustic soda solution can be established only by chemical or spectro-analysis. **268.-278. RESERVED.** Par 267 ### Page 175 (OCR) Page 132 AC 43. Section 2. TESTING OF METALS 279. HARDNESS TESTING. Hardness testing provides a convenient means for determining, within reasonable limits, the tensile strength of steel. It has several limitations in that it is not suitable for very soft or very hard steels. In hardness testing, the thickness of the specimen being tested and the edge distance should be such that distortion of the metal due to these factors is eliminated. Several readings should be taken and the results averaged. In general, the higher the tensile strength, the greater is its hardness. Common methods of hardness testing are outlined in the following paragraphs. These tests are suitable for determining the tensile properties resulting from the heat treatment of steel. Care should be taken to have case hardened, corroded, pitted, decarburized, or otherwise nonuniform surfaces removed to a sufficient depth. Also, exercise caution not to cold-work and consequently harden the steel during removal of the surface. The relationship between tensile strength and hardness is indicated in figure 7.3. 282. VICKERS HARDNESS TEST. In this test small pyramidal diamond is pressed into metal. The hardness number is the ratio of load to the surface area of indentation. 283. TESTING ALUMINUM. Hardness tests : useful for testing aluminum alloy chiefly a means of distinguishing between anneal cold-worked, heat-treated, and heat-trea and aged material. It is of little value in in cating the strength or quality of heat tre ment. Typical hardness values for alumin alloys are shown in figure 7.4 FIGURE 7.4.-Hardness values for aluminum alloy! Brinell numb Material commercial Eardnesa 500 kg. lom designation temper 10 mm. ball 1100 A18 8008 H16 280. BRINELL HARDNESS TEST. In this test a 2014 standard load is applied to a smooth surface of T8 185 2017 45 metal through a hardened steel ball, one centi- 18 105 meter in diameter. The numerical value of Bri- 2024 47 nell hardness is equal to the load divided by T· 120 the surface area of the resulting spherical im- 2025 T8 110 pression. 8151 100 5052 47 H86 78 281. ROCKWELL HARDNESS TEST. In this test, a 6061 80 standard minor load is applied to seat a hard- 65 ened steel ball or a diamond cone in the surface 95 of the metal, followed by the application of a 7075 185 standard major load. The hardness is meas- 7079 185 T6 185 75 ured by depth of penetration. Rockwell super- 220 T4 ficial hardness tests are made using light 0855 T6 minor and major loads and a more sensitive A856 system for measuring depth of indentation. It is useful for thinner sections, very small parts, **284.-294. RESERVED.** Par 279 Char ### Page 176 (OCR) Rockwell, superficial bardness Brinell bardness number Rockwell hardness number number, superfielal brale 10-mm. ball, 3000-kg. load penetrator Tenaile Diamond Shore etrengths Hockwell pyramid A-scale, B-scale, D-scale, sclero- (approxi- Rockwell bardness numbers C-scale (Vickers) hardness number ard ball Stand- Kult- gren ball penetra- 60-kz. load, brale 1/16-1n. 100-kg load, diam. penetra- 100-kg. lond, brale 15-kg. 15-N acale, load 30-kg. scale, 30-N load 45-kg. 45-N load scale, hardness number всоре mate) in P. 8. 1. 1000 bardness number® C-scale tor ball tor Col. 1 Col. 2 Col. 3 Col. 4 Col. 5 Col. 6 Col. 7 Col. 8 Col. 9 Col. 10 Col. 11 Col. 12 Col. 13 Col. 14 68 -- 940 85.6 76.9 93.2 84.4 75.4 97 68 67 900 85.0 76.1 92.9 83.6 74.2 95 67 66 865 84.5 75.4 92.5 82.8 73.3 92 66 65 832 739 83.9 74.5 92.2 81.9 72.0 91 65 64 800 722 83.4 73.8 91.8 81.1 71.0 88 64 63 772 705 82.8 73.0 914 80.1 69.9 87 63 62 746 688 82.3 72.2 91.1 79.3 68.8 62 61 720 670 81.8 71.5 90.7 78.4 67.7 83 61 60 697 613 654 T18 70.7 90.2 77.5 66.6 81 60 59 674 599 $£9 1°09 6*69 8°68 76.6 65.5 80 326 59 58 653 587 919 I°08 769 £*68 75.7 64.3 78 315 58 57 633 575 969 9'62 5*89 6*88 S72 63.2 76 305 57 56 613 561 LIS 0°62 L'19 73.9 62.0 75 295 56 55 595 546 099 S81 6°99 87.9 73.0 60.9 74 287 55 54 577 534 543 0°82 1°99 #°48 72.0 59.8 72 278 54 53 560 519 £ZS F°LL 6°98 71.2 58.6 71 269 53 52 544 500 508 ZIS 8°94 9°F9 1°98 70.2 57.4 69 262 52 51 528 487 494 96₺ £94 8°89 698 69.4 56.1 68 253 51 50 513 475 481 IS₺ 6°9L T°£9 68.5 55.0 67 245 50 49 498 464 469 69₺ 1°29 0°S8 67.6 53.8 66 239 49 48 484 451 455 SS₺ LYL £₺8 66.7 52.5 64 232 48 47 471 442 443 £FT I°72 8°09 6°&8 8"99 51.4 68 225 47 46 458 432 432 ZET 9EL 0°09 S£8 8*79 50.3 62 219 46 See footnotes at end of table ### Page 177 (OCR) FIGURE 7.3.-Rockwell C scale steel hardness numbers comparison table-continued. Brinell hardness number Rockwell, superficial hardness 10-mm. ball, 8000-kg. load Rockwell bardness numbers number, superficial brale penetrator Tenalle Rockwell bardnede number* C-scale (Vickera) hardness Diamond pyramid number ard ball Stand- Hult- gren ball carbide Tung- sten- ball penetra- A-scale, 60-kg. load, brale B-scale, 1/16-10. 100-kg. load, diam. D-scale, 100-kg. penetra- load, brale 15-kg. scale, 15-N load 80-kg. scale. 30-N load 45-58. scale, 45-N load hardness number Shore scope sciero- strengtha (approxi- mate) in D. e. I. 1000 hardness Rockwell numbers C-scale tor ball tor Col. 1 Col. 2 Col. 3 Col. 4 Col. 5 Col. 6 Col. 1 Col. 8 Col. 9 Col. 10 Col. 11 Col. 12 Col. 13 Col. 14 446 421 421 421 731 59.2 83.0 64.0 49.0 60 212 45 434 409 409 409 725 58.5 82,5 63.1 47.8 58 206 44 423 400 400 400 72.0 57.7 82.0 62.2 46.7 57 201 43 412 390 390 390 71.5 56.9 81.5 61.3 45.5 56 196 42 402 381 381 381 70.9 56.2 80.9 60.4 44.3 55 191 41 392 371 371 371 70.4 55.4 80.4 59.5 43.1 54 186 40 382 362 362 362 69.9 54.6 79.9 58.6 41.9 52 181 39 372 353 353 353 69.4 538 79.4 57.7 40.8 51 176 38 363 344 344 344 68.9 53.1 78.8 56.8 39.6 50 172 37 354 336 336 336 68.4 (109.0) 52.3 78.3 55.9 38.4 49 168 36 345 327 327 327 67.9 (108.5) 51.5 77.7 55.0 37.2 48 168 35 336 319 319 319 67.4 (108.0) 50.8 77.2 54.2 36.1 47 159 34 327 311 311 311 66.8 (920I) 50.0 76.6 53,3 34.9 46 154 33 318 301 301 301 66.3 (107.0) 49.2 76.1 52.1 33.7 44 150 32 310 294 294 294 65.8 (106.0) 48.4 75.6 51.3 32.5 43 146 31 302 286 286 286 65.3 (105.5) 47.7 75.0 50.4 31.3 42 142 30 294 279 279 279 64.7 (104.5) 47.0 74.5 49.5 30.1 41 188 29 286 271 271 271 64.3 (104.0) 46.1 73.9 48.6 28.9 41 184 28 279 264 264 264 63.8 (103.0) 45.2 73.3 47.7 27.8 40 131 272 258 258 258 63.9 (102.5) 44.6 72.8 46.8 26.7 38 197 26 266 253 253 253 62.8 (101.5) 43.8 72.2 45.9 25.5 38 124 25 260 247 247 247 62.4 (101.0) 49.1 71.6 45.0 24.3 37 121 24 254 243 243 243 62,0 100.0 42.1 71.0 44.0 23.1 36 118 23 248 237 237 237 61.5 99.0 41.6 70.5 43.2 22.0 35 115 22 243 231 231 231 61.0 98.5 40.9 69.9 42.3 20.7 35 113 21 ### Page 178 (OCR) PIGUKE TO -доскжен ~ осас і 20_ 238 226 226 226 60.5 97. 8 40. 1 69-4 $1.5 19. 6 110 20 3.13-14 (18) - 230 219 219 219 96. 7 (18) (16) 222 212 212 212 95. 5 (14) 213 203 203 203 93. 9 (12) 204 194 194 194 92. 3 (10) 196 187 187 187 90. 7 188 179 179 179 89. 5 180 171 171 171 87. 1 (4) 173 165 165 165 85. 5 (2) 166 158 158 158 83. 5 (0) - 160 152 152 152 81. 7 : The values in this table shown in bold-face type correspond to the values shown in the corresponding joint SAE-ABM-ASTH Committee on Hardoess Conversions as printed in ABTH E4S, table 2. • Values in () are beyond normal range and are given for information only. * It is possible that steels of various compositions and processing historles will deviate in hardness-tensile strength relationship from the data pres nted in this table. Above the level of Rockwell C43, deviation increases with increasing hardness and the table shall not be used above Bo 48 ercept in the absence of other data specifcally approved by the procuring agency. ### Page 179 (OCR) Page 136 AC 43. Section 3. NONDESTRUCTIVE TESTING 295. GENERAL, The field of NDT (nondestructive testing) and inspection is too varied to be covered in detail in this handbook. This section provides a brief description of the various methods that are available for use in aircraft maintenance. The effectiveness of any particular method of NDT inspection depends upon the skill, experience and training of the persona applying the process. Each process is limited in its usefulness as an inspection tool through its adaptability to the particular component to be inspected. Consult the aircraft or product manufacturer for specific instructions regarding NDT inspection of their product. 296. INSPECTION BY MAGNIFYING GLASS AFTER WELDING. Careful examination of all joints with a medium-power magnifying glass (at least 10-power), after first removing all scale, is considered an acceptable method of inspection for repaired structures. The practice of filling steel tubular structures with hot linseed or petroleum base oils, under pressure, in order to coat the inside surface and inhibit corrosion, assists in the detection of weld cracks, as the hot oil will seep through cracks invisible to the eye. This practice, though not justifiable in all cases, is suggested where a very large portion of the structure has been rewelded. 297. MAGNETIC PARTICLE INSPECTION. Magnetic particle inspection of Magnaflux (a patented trade name used by the Magnaflux Corp.) can be used only on magnetic materials; l.e., iron and steel. Most stainless or high chromium nickel and manganese alloy steels, being nonmagnetic, cannot be inspected by this method. The method consists essentially of detection of discontinuities (cracks, voids, defects, pits, subsurface holes, etc.) by means of accumulation of magnetic particles on the discontinuities when the part has been magnetized. The magnetic particles are applied either dry as a powder or suspended in light oil. F complete magnetic inspection, both circu and longitudinal, magnetization should be e ployed. CAUTION Improper operation of Magnaflux equipment be cause of faulty equipment or by untrained persons can Jeopardize the airworthinesa of parta bein inspected. Minute electrical arc burns caused durin inspection by improper operation of Magnaflu equipment, can result in eventual fallure of thi a. Circular magnetization is produced by tral mitting an electric current directly throu the article being tested, or through a centi conductor placed through the part, in whi case defects parallel to the flow of current m be detected. As an example, circular magne zation of a round steel bar would be produc by placing the ends of the steel bar betwe the heads of the magnetic inspection machi and passing a current through the bars. Ma netic particles applied either during or aft passage of the current, or after passage of t current in magnetically retentive steels, wou disclose discontinuities parallel to the axis the bar. b. Longitudinal magnetization is induced in part by placing the part in a strong magne! field, such as the center of a coil. Thus, longit dinal magnetization of a round steel bar wou be produced by placing the ends of the bar t tween the heads of a magnetic-inspecti machine and placing the D.C. solenoid arou the bar. After application of the magnet particles, either during or subsequent to ma netization, discontinuities perpendicular to t axis of the bar would be disclosed. c. Red, black, and sometimes gray particles & used in the wet or dry methods. In the case wet inspection, a fluorescent magnetic partic may also be used. This process is commercial known as Magnaglo. Articles inspected usi Par 295 Chap ### Page 180 (OCR) atter medium are illuminated by so-called it light, and the magnetic particles glow by pescence causing any defects or indications pe easily visible. The wet inspection procee provides better control and standardiza- 1 of the concentration of magnetic particles, ier application to complex shapes, and indiions that are easier to interpret. This is due the difficulty of obtaining efficient distribu- 1 of the dry powder during magnetization. e dry procedure is particularly suitable for ecting subsurface defects, such as, when pecting heavy welds, forgings, castings, etc. e wet continuous process is recommended ' most aircraft work. 1. The presence of accumulations of magnetic ticles in magnetic inspection does not necesily mean that a defect exists. Changes in tion of the part, particularly where the ange in section is very sharp, and also holes illed through a part, will frequently cause fications. Surface defects are most easily deted, however, since a crack will cause a arp line of magnetic particles to appear. surface defects are less easily detected, ice only a general collection of magnetic rticles will be observed. e. After magnetic inspection carefully demagtize and clean the parts. Examine for possie evidence of electrical arc burns that may ‚ve occurred during inspection. All metal irticles must be removed and the serviceable irts coated with a suitable preservative. f. Portable type magnetic particle inspection ruipment such as Sonoflux has been developed ›r use in the detection of surface or slightly ibsurface discontinuities in ferromagnetic rcraft materials and parts. This type equipent usually gives better results when the wet ispension type of indicator, such as that con- ›rming to Specification MIL-1-6868, is used ith the powder supplied by the equipment anufacturer. As in other inspection methods, ›llow the manufacturers' recommendations oncerning use of the procedure. **98. X-RAY OR RADIOGRAPHIC INSPECTION.** -ray may be used on either magnetic or nonagnetic materials for detecting subsurface lap 7 Page 137 voids such as open cracks, blowholes, etc. When a photographic film or plate is used to record the X-ray (in a similar manner to ex- Posine posing a photographic film), the process is known as radiography. When the X-rays are prosen, t th projected through the part onto a fluorescent she echnig screen, the process is known as fluoroscopy. The technique used for radiography should be capable of indicating the presence of defects having a dimension parallel to the X-ray beam of 3 percent of the thickness of the part being radiographed for magnesium alloys, and 2 percent for all other metals and alloys. Inspection using a fluoroscopic screen is much less sensitive. Consequently, the radiographic method is usually used for inspection and the fluoroscopic method is used for culling. a. Radiographic inspection is extensively used in the aircraft industry for the inspection of all types of castings including sand castings, permanent-mold castings, die castings, etc. X- ray is particularly useful for this application, since it is capable of disclosing defects which exist below the surface, and also since the open types of defects which may occur in castings (shrinks, blowholes, dross inclusions, etc.) are readily disclosed by proper use of X-rays. In the inspection of forged or wrought metals, on the other hand, X-ray inspection is not used so extensively. This is due to the fact that the process of forging or working may cause defects which originally existed in the metal to become tightwalled cracks. Such defects are somewhat difficult to disclose by X-rays. If doubt exists as to the suitability of the X-ray examination, consult a laboratory familiar with the X-ray examination of aircraft parts. b. In radiography, values of peak kilovoltage, radiographic density range and penetrometer characteristics are often selected that produce less than optimum radiological data. This selection of high kilovoltage is made in order to reduce the exposure time. The use of too high a kilovoltage reduces the resolvable detail recorded on the radiographic film. As the kilovoltage is increased, X-rays of shorter wave length and greater penetrating power are produced. This presents a sound argument for increased kilovoltages but does not take into Par 297 ### Page 181 (OCR) account the effects of scatter both within the sample and the radiographic film, which in reality, reduces the resolvability of details recorded on the X-rays. c. If for some reason a short exposure time is required, a faster film is normally used with a higher kilovoltage; however, this has the effect of increasing the granularity and reducing the resolution on the radiographic material. d. The use of low voltages results in improved radiographic signal-to-noise ratio and improved resolution. The recommended kilovoltages are shown in figure 7.5. FIGURE 7.5.-Maximum recommended X-ray kilovolts, Material Atomio No. Man. KV Beryllium 25 Carbon 40 Magnesium 12 50 Alumium 18 75 Titanium 22 120 Steel 22-20 150 Steel 20-28 200 Silver 47 800 Lead 82 1-2 (megavolta) 299. FLUORESCENT PENETRANT. In this method of inspection (Zyglo), the article, which may be of metal, plastic material, etc., is first carefully cleaned to permit the fluorescent material to penetrate cracks and defects. It should be noted that cleaning of aluminum may necessitate stripping of any anodizing, since the anodized film, if formed after the defect, could prevent penetration of the fluorescent material and an anodized film tends to hold penetrants which may obscure defect indications. After the article is cleaned, it is either sprayed, painted, or immersed in a bath of fluorescent penetrant. The penetrant is a light oil which has the property of fluorescing or emitting visible light when excited by invisible radiation in the near ultra-violet range (so-called black light). It is important that the penetrant be given sufficient time to penetrate cracks and defects, and for fatigue cracks a minimum of 80 minutes is stipulated by MIL-I-6866. Heat may also be applied to facilitate entry of the penetrant. After the penetrant has had sufficient time to enter any defects, the excess on Par 298 AC 4: the surface of the article is washed o water spray. This washing should be ch by inspection with black light, by which r any penetrant left on the surface may I tected. After washing, a developer is us bring out the indication. This developer m in a liquid form or may be a light powde absorbs the penetrant as it oozes from c and defects in the part. The development also be aided by application of heat to the After the indications have been develope‹ part is inspected under black light. Any ices into which the fluorescent material has etrated will show as luminous areas. Indications which appear are us checked by close inspection with a magni glass, by etching with a suitable acid or ca solution, or it may be necessary to cross-se the part, a procedure which, of course stroys its usefulness. Usually a skilled ope can determine whether an indication act shows a defect or whether it is a false in tion. Also, the internal extent of the defec sometimes be estimated with fair accurat should be noted that this process of inspec like all others, has its limitations. If the orescent material for any reason is not at penetrate into a defect, such a defect cann detected. 300. DYE PENETRANTS. Several dye penet type inspection kits are now available w will reveal the presence of surface crack defects and subsurface flaws which exter the surface of the part being inspected. I penetrant type inspection methods are co: ered acceptable, provided the part being spected has been thoroughly cleaned, all a are readily accessible for viewing, and manufacturer's recommendations as to me of application are closely followed. a. Cleaning. An inspection is initiated by cleaning the surface to be inspected of loose scale, oil, and grease. Precleaning usually be accomplished by vapor degrea or with volatile cleaners. Use a volatile cle as it will evaporate from the defects be applying the penetrant dye. Sand blastin not desirable as a cleaning method, since face indications may be obscured. It is not ### Page 182 (OCR) 43.13-1A ary to remove anodic films from parts to be spected, since the dye readily penetrates such ms. Special procedures for removing the exsa dye should he followed. b. Application of Panefrant. The penetrant is plied by brushing, spraying, or by dipping d allowed to stand for a minimum of 2 mines. Dwell time may be extended up to 15 mines, depending upon the temperature of part d fineness of the defect or when the surface ing inspected is ground. Parts being inected should be dry and heated to at least ° F., but not over 130° F. Very small indicains require increased penetration periods. t. Removal of Dye Peneirant. Surplus peneint is usually removed by application of a acial cleaner or remover, or by washing with in water and the part allowed to dry. Water he may also be used in conjunction with the over, subject to the manufacturer's recom- •dations. Application of Developer. A light and even of developer is applied by spraying, brushor dipping. When dipping, avoid excess aclation. Penetrant which has penetrated cracks or other openings in the surface of naterial will be drawn out by the develresulting in a bright red indication. Some Page 139 idea of the size of the defect may be obtained after experience by watching the size and rate of growth of the indication. 301. ULTRASONIC FLAW DETECTION. Ultrasonic flaw detection equipment has made it possible to locate defects in all types of materials without damaging the material being inspected. Very small cracks, checks, and voids, too small to be seen by X-ray, are located by means of ultrasonic inspection. An ultrasonic test instrument requires access to only one surface of the material to be inspected, and can be used with either straight line or angle beam testing techniques. The instrument electronically generates ultrasonic vibrations and sends them in a pulsed beam through the part to be tested. Any discontinuity within the part, or the opposite end, will reflect the vibration back to the instrument, which measures the elapsed time between the initial pulse and the return of all reflections and indicates such time lapse on a cathode ray indicator or paper recorder. Ultrasonic inspection requires a skilled operator who is familiar with the equipment being used as well as the inspection method to be used for the many different parts being tested. **302.-312. RESERVED.** Par 300 ### Page 183 (OCR) Page 140 AC 43.13-1A CH Section 4. IDENTIFICATION OF FABRICS AND PLASTICS 313. IDENTIFICATION OF FABRIC. Cotton fabric is often used as covering for wing, fuselage, and control surfaces of aircraft. Acceptable grades of fabric for use on civil aircraft are listed in Chapter 3. In general, the fabric can be readily identified by a continuous marking to show the manufacturer's name or trademark and specification number. This marking may be found stamped along the selvage edge. The specification number for grade "A" fabric is AMS-3806, and for the intermediate grade AMS-3804. The corresponding FAA Technical Standard Order * Numbers for these materials are TSO-C15 and 150-C14, respectively. Increasing interest in* the use of linen and certain synthetic fabrics in lieu of cotton has been noted. Identity of such materials should always be verified by the user. 314. IDENTIFICATION OF PLASTICS. Plastics cover a broad field of organic synthetic resins and may be divided into two main classifications thermoplastic and thermosetting plastics. a. Thermoplostics. Thermoplastics softened by heat and can be dissolved in varic organic solvents. Two kinds of transpare thermoplastic materials are commonly employ in windows, canopies, etc. These materials s known as acrylic plastics and cellulose acets plastics. These two plastics may be distinguish by the absence of color, the greater transparen and the greater stiffness of the acrylic as co pared to the slight yellow tint, lower transp ency, and greater flexibility of cellulose aceta b. Thermosetting Plastics. Thermosetting pl tics do not soften appreciably under hent may char and blister at temperatures of 204 260° C. (400° to 500° F.). Most of the mo products of synthetic resin composition, suc phenolic, urea-formaldehyde, and melan formaldehyde resins, belong to the thermoset group. **315.-320. RESERVED.** Par 313 ### Page 184 (OCR) C 43.13-1A CHG 1 Page 140~1 75 Section 5. CABIN INTERIOR MATERIALS, FIRE PROTECTION QUALITIES 21. REQUIREMENTS FOR FIRE PROTECTION. The leaning, repairing: and/or replacement of cabin nterior materials necessitates continuing complince with the fire protection requirements of the ARs. A recent field survey on this matter reor maintaining cabin interior materials with ealed a wide variance in the procedures used spect to the fire protection qualities of these aterials. a. Inferior Materials. The methods, techniques, actices, and materials used in the cleaning, reir, and replacement of cabin materials must et the requirements of Section 43.13 of the • Fire Protection. The requirements relative ire protection qualities of cabin interior maas are specified in: (1) Section 121.312, for aircraft operated under Parts 121, 123, and 135. (2) Section 127.91, for helicopters used in passenger service under Part 127. c. Source of Information. If information regarding the original or properly altered fire protection qualities of certain cabin interior materials is not available, requests for this information should be made to the aircraft manufacturers or the local FAA regional office, specifying the model aircraft, the aircraft manufacturer, the date the aircraft was manufactured or the serial number, and the FAR Part under which the aircraft is operated (i.e., Part 91, Part 121, etc.). **322.-325. RESERVED.** Par 321 ### Page 185 (OCR) AC 43.13-1А Page 141 1972 Chapter 8. AIRCRAFT EQUIPMENT Section 1. LANDING GEAR EQUIPMENT 326. GENERAL. Many factors affect the scope and frequency of landing gear aystem inspection and maintenance. The possibility of system malfunction increases with severe operating conditions, such as pilot training and agricultural operations. Cracks are the foremost area of concern, followed closely by worn, sheared, bent, or broken bolts, and elongated boltholes, Although these defects occur during normal aircraft usage, adverse operating conditions accelerate the problem. Improper adjustment or lubrication can also cause malfunctioning, especially in landing gear retraction mechanisms. Excessive lubrication or the wrong type of lubricant can often be as bad or worse than none at all. With regard to lubrication, closely follow the manufacturer's service and maintenance instructions. residue around bolts, rivets, and pins is a good indication of looseness or wear, Thoroughly clean and re-inspect the landing gear to determine the extent of any damage or wear. Some components may require removal and complete disassembly for detailed inspection. Others may require a specific check using an inspection process such as dye penetrant, magnetic particle, radiographic, ultrasonic, or eddy current (see Chapter 7). The frequency, degree of thoroughness, and selection of inspection methods are dependent upon the age, use, and general condition of the aircraft. Inspect the aircraft structure surrounding any visible damage to insure that no secondary damage remains undetected. Forces can be transmitted along the affected member to remote areas where subsequent normal loads can cause failure at a later date. CAUTION Inspect grease fittings after applying high pressure lubricant to determine that the spring-loaded ball has reseated, Failure of the ball to reseat is an indication of an internal fallure of the fitting which can result in serious damage to the part 327. INSPECTION. Inspect the aircraft landing gear and all associated hardware closely for cracks, cleanliness, lubrication, leaks, deformation, excessive wear, and security of attachments. Before removing any accumulated dirt, cloaely observe the area being inspected while the wingtips are gently rocked up and down. Excessive motion between normally close fitting components may indicate wear, cracks, or improper adjustment. If a crack exists, it will generally be indicated by dirt or metallic part!- cles which tend to outline the fault. Seepage of rust inhibiting ofa, used to coat internal surfaces of steel tuber, also assists in the early detection of cracks. in addition, a sooty, olly a. Types of Malfunctions. (1) Cracks. Príme locations for cracks on any landing gear are bolts, boltholes, pins, rivets, and welds. The following are typical locations where cracks may develop: (a) Bolts. Most suaceptible areas are at the radius between the head and the shank, and in the location where the threads join the shank, as shown in figure 8.1. (b) Boltholes and Pinholes, Cracks primarily occur at the edge of boltholes or pinholes on FIOURE 8.1.Typical bolt cracks. 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