# AC 43.13-2A CHG 2: Acceptable Methods, Techniques, and Practices -- Aircraft Alterations - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Historical - Version in force: [AC 43.13-2C](https://openairdocs.org/ac/43.13-2), approved August 10, 2026. This is a historical version - Approved: October 30, 1989 - Last reviewed: June 12, 2024 - Family: AC 43.13-2 - Cancels: AC 43.13-2A CHG 1 - Cancelled by: [AC 43.13-2B](https://openairdocs.org/ac/43.13-2B) - 14 CFR parts: 43, 23, 27, 65, 91, 133, 45, 21, 29, 25 - 14 CFR sections: [43.13](https://www.ecfr.gov/current/title-14/section-43.13), [43.16](https://www.ecfr.gov/current/title-14/section-43.16), [43.3](https://www.ecfr.gov/current/title-14/section-43.3), [43.9](https://www.ecfr.gov/current/title-14/section-43.9), [43.5](https://www.ecfr.gov/current/title-14/section-43.5), [23.303](https://www.ecfr.gov/current/title-14/section-23.303), [23.562](https://www.ecfr.gov/current/title-14/section-23.562), [23.561](https://www.ecfr.gov/current/title-14/section-23.561), [23.341](https://www.ecfr.gov/current/title-14/section-23.341), [23.473](https://www.ecfr.gov/current/title-14/section-23.473), [23.301](https://www.ecfr.gov/current/title-14/section-23.301), [23.337](https://www.ecfr.gov/current/title-14/section-23.337), [23.321](https://www.ecfr.gov/current/title-14/section-23.321), [23.471](https://www.ecfr.gov/current/title-14/section-23.471), [23.307](https://www.ecfr.gov/current/title-14/section-23.307), [23.621](https://www.ecfr.gov/current/title-14/section-23.621), [23.625](https://www.ecfr.gov/current/title-14/section-23.625), [23.785](https://www.ecfr.gov/current/title-14/section-23.785), [23.787](https://www.ecfr.gov/current/title-14/section-23.787), [23.613](https://www.ecfr.gov/current/title-14/section-23.613), [23.611](https://www.ecfr.gov/current/title-14/section-23.611), [23.1401](https://www.ecfr.gov/current/title-14/section-23.1401), [23.1397](https://www.ecfr.gov/current/title-14/section-23.1397), [23.1321](https://www.ecfr.gov/current/title-14/section-23.1321), [23.1311](https://www.ecfr.gov/current/title-14/section-23.1311), [27.1401](https://www.ecfr.gov/current/title-14/section-27.1401), [27.1397](https://www.ecfr.gov/current/title-14/section-27.1397), [91.205](https://www.ecfr.gov/current/title-14/section-91.205), [91.407](https://www.ecfr.gov/current/title-14/section-91.407), [91.409](https://www.ecfr.gov/current/title-14/section-91.409), [133.43](https://www.ecfr.gov/current/title-14/section-133.43), [133.47](https://www.ecfr.gov/current/title-14/section-133.47), [45.15](https://www.ecfr.gov/current/title-14/section-45.15), [21.607](https://www.ecfr.gov/current/title-14/section-21.607), [21.303](https://www.ecfr.gov/current/title-14/section-21.303), [21.305](https://www.ecfr.gov/current/title-14/section-21.305), [27.321](https://www.ecfr.gov/current/title-14/section-27.321), [27.471](https://www.ecfr.gov/current/title-14/section-27.471), [27.561](https://www.ecfr.gov/current/title-14/section-27.561), [27.307](https://www.ecfr.gov/current/title-14/section-27.307), [27.785](https://www.ecfr.gov/current/title-14/section-27.785), [27.787](https://www.ecfr.gov/current/title-14/section-27.787), [27.613](https://www.ecfr.gov/current/title-14/section-27.613), [29.785](https://www.ecfr.gov/current/title-14/section-29.785), [29.787](https://www.ecfr.gov/current/title-14/section-29.787), [29.613](https://www.ecfr.gov/current/title-14/section-29.613), [25.307](https://www.ecfr.gov/current/title-14/section-25.307), [25.561](https://www.ecfr.gov/current/title-14/section-25.561), [25.613](https://www.ecfr.gov/current/title-14/section-25.613), [29.307](https://www.ecfr.gov/current/title-14/section-29.307), [29.561](https://www.ecfr.gov/current/title-14/section-29.561) - Pages: 103 - PDF: https://openairdocs.org/files/sha256/88/1f/881fbb4f3fa6c9060bdc0a65e00e4e3819173350085ce3f70bcf5700ffe96d71 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/DRSDOCID179920183920240612153158.0001 - DRS record last changed: November 21, 2025 - Page: https://openairdocs.org/ac/43.13-2A-CHG-2 ## Versions - [AC 43.13-2C](https://openairdocs.org/ac/43.13-2) (Current), approved Aug 10, 2026, 196 pages - [AC 43.13-2B](https://openairdocs.org/ac/43.13-2B) (Historical), approved Mar 3, 2008, 137 pages - [AC 43.13-2A CHG 2](https://openairdocs.org/ac/43.13-2A-CHG-2) (Historical), approved Oct 30, 1989, 103 pages ## ADs that cite it (11) - [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 2011-27-04](https://openairdocs.org/ad/2011-27-04): Airspeed Indicator · Two Teledyne Continental Motors model IO-550-C engines… (Mike Jones Aircraft Sales STC SA1762SO). Effective Dec 29, 2011 - [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 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 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 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 80-21-03](https://openairdocs.org/ad/80-21-03): Roll Axis Flight Control · King Model KFC-200 two axis flight control system (Honeywell International STC SA1202CE or SA1487CE). Effective Oct 23, 1980 - [AD 80-08-04](https://openairdocs.org/ad/80-08-04): Auxiliary Fuel System · Chadwick C-20 auxiliary fuel system (Phoenix Air-Atlanta STC SH129WE or SH139WE). Effective Apr 17, 1980 - [AD 73-20-02](https://openairdocs.org/ad/73-20-02): Front Spar Lower Cap Center Section · Transcontinental and Western Air Army B-17F and Army B-17G. Effective Jan 26, 1979 - [AD 75-11-02](https://openairdocs.org/ad/75-11-02): Fuel Lines · Maule Aerospace Technology M-4-210, M-4-210C and M-5-210C. Effective May 19, 1975 - [AD 70-05-05](https://openairdocs.org/ad/70-05-05): Muffler Inspection · True Flight Holdings AA-1. Effective Mar 13, 1970 ## 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) 8200 4124170 HAC - 640 Advisory U.S. Department of Transportation Circular Federal Aviation Administration Subject: ACCEPTABLE METHODS, Date: 10/30/89 AC No: AC 43.13-2A TECHNIQUES, AND PRACTICES -- Initiated by: AFS-340 Change: 2 **AIRCRAFT ALTERATIONS** 1. PURPOSE. This Change corrects an editing error in issuance of "Change 1." The Change number and date of the changed material are carried at the top of the page. Pages having no changes retain the same heading information. 2. PRINCIPAL CHANGES. Change 1, dated June 1, 1979, deleted paragraph 133. Paragraph 133 continued onto page 61, but this page was inadvertently left out in the page control chart to be removed. Therefore, this change corrects that error. PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated 61 Rev. 1977 61 10/30/89 62 Rev. 1977 62 Rev. 1977 DEBeaudette D. C. Beaudette Acting Director, Flight Standards Service U.S. GOVERNMENT PRINTING OFFICE: 1989 0-942-127 ### Page 2 (OCR) feled 4124190 886-444 AC 43.13-2A CHANGE 1 DATE 6/1/79 **ADVISORY CIRCULAR** DEPARTMENT OF TRANSPORTATION CHANGE Federal Aviation Administration Washington, D.C. FAR GUIDANCE MATERIAL Subject: ACCEPTABLE METHODS, TECHNIQUES AND PRACTICES-AIRCRAFT ALTERATIONS-Towline Information 1. PURPOSE. This change transmits revised material for Advisory Circular 43.13-2A. Changed material is indicated by asterisks, and the change number and date are carried at the top of each page. 2. PRINCIPAL CHANGES. This change deletes operational practices from Chapter 8, section 1, dealing with towplane performance considerations. The change also deletes paragraph 133 dealing with towlines. This information is not pertinent to this publication since it is primarily an operational consideration and not germane to altering of aircraft. The change also clarifies the placarding limits of the towline in terms of the maximum breaking strength of the towrope to be used. The old procedure was to limit the weight of gliders to be towed. Page Control Chart Remove Pages Dated Insert Pages Dated MI-VI Rev. 77 III 6/1/79 IV-V Rev. 77 VI 6/1/79 57-60 Rev. 77 57-60 6/1/79 Came r. Sins 1079 JAMES M. VINES Acting Director Flight Standards Service Initiated by: AFS-830 ### Page 3 (OCR) AL NU: 43.13-2A DATE: Revised 1977 **ADVISORY** **CIRCULAR** **DEPARTMENT OF TRANSPORTATION** FEDERAL AVIATION ADMINISTRATION SUBJECT: ACCEPACT AL MERATIONS TECHNIQUES, AND PRACTICES- **1. PURPOSE.** This advisory circular contains methods, techniques, and practices acceptable to the Administrator for use in altering civil aircraft. **2. CANCELLATION.** Advisory Circular 43.13-2 dated 1965 is cancelled. **3. REFERENCE.** Part 43 of the Federal Aviation Regulations requires that methods, techniques, and practices acceptable to the Administrator must be used when altering civil aircraft. Methods, techniques, and practices other than those prescribed in this advisory circular may be used provided they are acceptable to the Administrator. FAA Inspectors are prepared to answer questions that may arise in this regard. Persons engaged in alteration of civil aircraft should be familiar with FAR Part 43, and subparts A, D, and E of FAR Part 65, and the applicable airworthiness requirements under which the aircraft was type certificated. Acting Director, Flight Standards Service Initiated by: AFS-580/AFS-830 ### Page 4 (OCR) Rev. 1977 AC 43.13-2A CONTENTS Chapter 1. STRUCTURAL DATA Paragraph Page 1. General - 2. Static Loads -. .... 3. Static Tests 4. Materials --. 5. Fabrication 6. Fasteners -- 7. Protection Against Deterioration 8. Provisions for Inspection -- 9. Effects on Weight and Balance 10. Effects on Safe Operation -. 11. Controls and Indicators -- 12. Placarding 18.-20. [Reserved] Chapter 2. RADIO INSTALLATIONS 21. Installation 22. Equipment Manufacturer's Instructions 23. Instrument Panel Mounting 24. Other Mounting Areas -- 25. Fabrication of Supporting Brackets for Attachment to Structure Other Than Flooring - 26. Supporting Structure Reinforcement 27. Electrical Systems 28. Emergency Locator Transmitter (ELT) Installations 29.-35. [Reserved] Chapter 3. ANTENNA INSTALLATIONS 36. Performance 37. VHF Antenna-Whip 38. VHF Antenna-Rigid - 39. VHF Navigation Receiving Antennas 40. Range Receiving Antennas - 41. Marker Receiving Antenna - 42. ATC Radar Beacon (Transponder) and Distance Measuring Equipment (DME) Antennas - 43. Direction Finding Antennas 44. Antenna Installation on Pressurized Aircraft 45.-50. [Reserved] 24 Chapter 4. ANTICOLLISION AND SUPPLEMENTARY LIGHT INSTALLATIONS 51. Anticollision and Supplementary Light Systems 25 52. Interference 53. Markings and Placards 54. Electrical Installation - 55. Alteration of Structure 56. Guidelines for Installations - 57.-60. [Reserved] ### Page 5 (OCR) AC 43.13-2A Rev. 1977 CONTENTS-Continued Chapter 5. SKI INSTALLATIONS Paragraph Section 1. SELECTION OF SKIS Page 61. Determining Eligibility of Aircraft 33 62. Identification of Approved Model Skis 63. Maximum Limit Load Rating 64. Landing Gear Moment Reactions 34 65. [Reserved] - 34 Section 2. CONVERSION AND INSTALLATION 66. Hub-Axle Clearance --. 35 67. Crust-Cutter Cables ------ 35 68. Cable and Shock Cord Attachment and Attachment Fittings - 35 69. Provisions for Inspection -. 36 70. [Reserved] 36 Section 3. RIGGING OF SKIS 71. Location of Attach Fittings on Fuselage or Landing Gear 72. Main Ski Incidence Angles 73. Tension Required in Main Ski Shock Cords 74. Nose Ski Installation - 75. Tail Ski Installation 76.-78. [Reserved] 40 Section 4. OPERATION 79. Performance Information - 80. Flight and Handling Operational Checks 81. Interchanging of Skis and Wheels 82.-85. [Reserved] Chapter 6. OXYGEN SYSTEM INSTALLATIONS IN NONPRESSURIZED AIRCRAFT 86. System Requirements --- 43 87. Installation 43 88. Location and Mounting - 41 89. Thread Compound 90. Functional Test - 91. Operating Instructions 46 92.-95. [Reserved] 47 Chapter 7. ROTORCRAFT EXTERNAL-LOAD DEVICE INSTALLATIONS Section 1. CARGO SLINGS 96. General -- 97. Quick-Release Device 98. Location of Cargo Release in Relation to the Rotorcraft's C. G. Limits 99. Maximum External Load 100. Static Test 101. Sling-Leg Angles of Cable-Supported Slings 102. Minimum Sling-Leg Cable Strength 50 103. Sling Installation 50 104. Installation of Release Controls 51 52 105. Functional Test -- 106. Supplemental Flight Information 52 107.-110. [Reserved] 53 Section 2. CARGO RACKS 111. General 112. Fabrication of Cargo Racks 113. Static Test ------ 114. Supplemental Flight Information 115.-120. [Reserved] ### Page 6 (OCR) CONTENTS-Continued Chapter 8. GLIDER AND BANNER TOW-HITCH INSTALLATIONS Paragraph Section 1. TOWPLANE CONSIDERATIONS Page * 121.-125. [Deleted] Change 1 57 Section 2. TOW-HITCH INSTALLATIONS 126. Structural Requirements. 58 127. Structural Testing 59 128. Attachment Points _ 59 129. Angies of Tow .. 59 130. Placards. 59 131. Weight and Balance 59 132. Tow Release Mechanism _ 59 * 133. [Deleted] Change 1 60 134.-145 [Reserved] _ 63 Chapter 9. SHOULDER HARNESS INSTALLATIONS Section 1. RESTRAINT SYSTEMS 146. General 65 147. Types of Restraint Systems 148. Advantages of Different Harness Types 149. Mounting Configurations ... 150. Standards 151. Materials 152.-155. [Reserved] _ Section 2. EFFECTIVE RESTRAINT ANGLES 156. Relationship of the Harness Assembly to the Occupant . 157. Attachment Areas for Shoulder Harness 158. Area and Angle Deviations - 159.-160. [Reserved] __. Section 3. ATTACHMENT METHODS 161. Structural Attachments 162. Floor and Seat Attachments 163. Airframe Attachments _ 164. Structural Repair Kits_. 165. Flexible Attachments 166. Simplified Installation Criteria 167. Engineering Approval 168.-175. [Reserved]. Chapter 10. BATTERY INSTALLATIONS Section 1. GENERAL 176. General 177. Location Requirements 178. Duplication of the Manufacturer's Installation 179. Other Installations . 180.-185. [Reserved] 78 Section 2. INSTALLATION 186. Securing the Battery 79 187. Venting-_ 79 188. Drains .. 79 189. Electrical Installation 80 190.-195. [Reserved] _ 80 III ### Page 7 (OCR) AC 43.13-2A Rev. 1977 CONTENTS-Continued Paragraph Page Section 3. REPLACEMENT OF LEAD-ACID BATTERIES WITH NICKEL-CADMIUM BATTERIES 196. General 81 197. Electrical Analysis ào 00 198. Maintenance Considerations 199. Structural Requirements --- 81 200. Isolation of Battery Case -- 81 201. Ventilation 82 202. Preinstallation Requirements 82 203. Securing The Battery 83 204. Voltage and Current Regulation 83 205. Weight and Balance --- 83 206. Restoration of Lead-Acid Batteries 83 207.-210. [Reserved] 83 Chapter 11. ADDING OR RELOCATING INSTRUMENTS 211. General 212. Preparation 213. Installation 214. Testing, Marking and Placarding 215.-240. [Reserved] Chapter 12. LITTER, BERTH, AND CARGO TIEDOWN DEVICE INSTALLATIONS 241. General 89 242. Installation Considerations - 243. Fabrication and Installation 244. Structural Attachment 245. Load Factors 95 246. Static Tests 247. Operating Limitations, Loading Instructions, and Placards 248.-260. [Reserved] Chapter 13. PENETRATION THROUGH PRESSURIZED STRUCTURE Section 1. ELECTRICAL WIRE BUNDLES AND COAXIAL CABLE FEED THROUGH PRESSURIZED STRUCTURE 261. General 97 262. Cleaning 263. Application of Sealant 264.-290. [Reserved] 97 LIST OF ILLUSTRATIONS Figure Page 1.1 Hypothetical Example of Determining Static Test Loads 2.1 Typical Radio Installations in Stationary Instrument Panels 2.2 Typical Panel Mount 2.3 Typical Shockmounted Base --. 2.4 Typical Shockmounted Base --- 2.5 Typical Rigid Baseplate Mount 2.6 Typical Underseat Installation 2.7 Typical Remote Unit Mounting Base-Vertical or Horizontal 2.8 Typical Shelf Installation 2.9 Typical Rail Platform Installation-Aluminum Alloy Structure - 2.10 Typical Attachment of Support Structure to Tubular Frame --. 2.11 Typical Mounting Plates for Remote Location of Radio Units 3.1 Typical Whip Antenna Installation IV ### Page 8 (OCR) Rev. 1977 AC 43.13-2A CONTENTS-Continued LIST OF ILLUSTRATIONS Continued Figure Page 3.2 Antenna Ground Plane for Nonmetallic Aircraft 14 3.3 Typical Shockmounted Antenna Installation 3.4 Typical Rigid Antennas _. 3.5 Typical Antenna Mounting Template 15 3.6 Typical Antenna Installation on a Skin Panel 15 3.7 Preferred VOR Antenna Location for Maximum Signal Pickup With Minimum Ignition Interference -- 16 3.8 Typical Antenna Installation Involving a Cut Stringer 16 3.9 Typical Dipole Antenna Assembly 16 3.10 Typical VOR Antenna Balun -- 3.11 Coaxial Cable Preparation -- 3.12 Coaxial Cable Tinning Center Lead - 3.13 Coaxial Cable Soldering Center Lead 3.14 Coaxial Cable Install Connector Body 3.15 Typical Wire Antenna Locations 3.16 Typical Marker Beacon Receiving Antenna 18 3.17 Typical Wing Tip Installations 19 3.18 Typical Fin Tip Installations - 19 3.19 Typical Wire Antenna Installations 20 3.20 Preferred Position for Distance Measuring Antennas and/or ATC Radar Beacon Antennas 20 3.21 Typical ADF Antenna Installations 21 3.22 Typical Antenna Installation --- 22 3.23 Faying Surface Sealing 22 3.24 Fastener Sealing 22 3.25 Injection Sealing 23 3.26 Blade Antenna Installation 23 3.27 ATC Flush Mounted Antenna -- 23 3.28 Marker Beacon Flush Mounted Antenna 24 3.29 VOR Flush Mounted Antenna -- 24 4.1 Typical Anticollision or Supplementary Light Installation in a Skin Panel (Unpressurized) -. 26 4.2 Typical Anticollision or Supplementary Light Installation Involving a Cut Stringer (Unpressurized) ---- 26 4.3 Typical Anticollision or Supplementary Light Installation in a Fin Tip 27 4.4 Typical Rotating Type Anticollision Light Installation for Pressurized Aircrait 27 4.5 Typical Oscillating Type Anticollision Light Installation for Pressurized Aircraft _- 4.6 Ramp Shadow Pattern 4.7 Conversion to Vertical vs. Horizontal Degrees 4.8 Conversion to True Square Degrees-1st Quadrant 4.9 Scale Drawings _ 4.10 Solid Angle Blockage1st Quadrant 31 5.1 Typical Ski Installation -- 33 5.2 Typical Hub Installation 35 5.3 Typical Fuselage Fitting 36 5.4 Main Ski at Maximum Positive Incidence (Check Cable Tight) - 37 5.5 Main Ski at Maximum Negative Incidence (Safety Cable Tight) 38 5.6 Main Ski Incidence Angles - 38 5.7 Typical Nose Ski Installation -- 39 5.8 Typical Tail Ski Installation --. 40 6.1 Typical Oxygen Duration Table . 6.2 Typical Floor Mounting ------- 6.3 Typical Baggage Compartment Mounting 46 ### Page 9 (OCR) AC 43.13-2A, CHG. 1 6/1/79 CONTENTS-Continued LIST OF ILLUSTRATIONS-Continued Page Typical Oxygen Installation in Light Twin Aircraft_. 47 Location of Cargo Release in Relation to the Rotorcraft's Center of Gravity 49 Cable Load Table _ 50 Typical Cargo Sling Wiring Diagram - 51 Typical Cargo Sling Installation No. 1 52 Typical Cargo Sling Installation No. 1 (Showing Fuselage Attachment 52 Treas Cargo sling Installation No. 1 (Showing Fore and Aft Limiting 52 7.7 Typical Cargo Sling Installation No. 2 (Cargo Hook Attached Directly to Underside of Fuselage). 52 Typical Cargo Sling Installation No. 3(4-leg, Cable Suspended). 53 Typical Cargo Sling Installation No. 3 (Showing Cable Sling Leg Attachment to Landing Gear Cross-Tube Fitting) 53 7.10 Typical Cargo Sling Installation No. 3 (Showing Cargo Sling in Stowed 53 7.11 Typical Cargo Rack Installation No. 1 54 7.12 Typical Cargo Rack Installation No. 1 (Showing Attachment Detail) - 54 7.13 Typical Cargo Rack Installation No. 2 - 55 7.14 Typical Cargo Rack Installation No. 2 (Showing Rack Partially Installed) 55 7.15 Typical Cargo Rack Installation No. 3 - 55 Glider Tow Angle 58 8.2 Banner Tow Angle 58 8.3 Typical Tow-Hitch Release Handle 59 8.4 Representative Steel Cable Qualities. 60 * 8.5 [Deleted] Change 1 - 60 8.6 Tricycle Gear Aircraft 62 8.7 Conventional Gear Aircraft-Leaf Spring Type Tail wheel 62 8.8 Conventional Gear Aircraft-Shock Strut Type Tailwheel 63 8.9 Conventional Gear Aircraft-Tubular Spring Type Tailwheel 63 9.1 Single Diagonal Type Harness _ 66 9.2 Double Over-the-Shoulder Type Harness 66 9.3 Side Mounted-Single Diagonal Type Harness.. 67 9.4 Ceiling Mounted Inertia Reel-Single Diagonal Type Harness 68 9.5 Acceptable Mounting Area-Single Diagonal Type Harness 69 9.6 Ceiling Mounted Inertia Reel-Double Over-the-Shoulder Type Harness - 69 9.7 Acceptable Mounting Areas- Double Over-the-Shoulder Type Harness 69 9.8 Floor Mounted Inertia Reel-Double Over-the-Shoulder Type Harness 71 9.9 Typical Floor Mounted Inertia Reel Installation. 72 9.10 Typical Seat Mounted Inertia Reel Installation 72 9.11 Typical Bulkhead Reinforcement Installation 73 9.12 Typical Wing Carry-Through Member Installation. 73 9.13 Typical Hat Section Reinforcement Installation. 74 9.14 Typical Stringer Section Reinforcement Installation 74 9.15 Typical Stringer Section Reinforcement Installation 74 9.16 Typical Attachment to Tubular Members - 75 Typical Battery Box Installation in Aft Fuselage Area... 77 Typical Battery Box Installation in Aft Fuselage Area, Below Cabin Floorboards .. 77 10.3 Typical Battery Box Installation in Aft Fuselage Area 10.4 Typical Battery Box Installation Suspended From Cabin Floorboard Section 10.5 Battery with Integral Vent Nipples _ 10.6 Typical Battery Cutoff and Generator/Alternator Control Circuit 10.7 Change in Cantilever Support Battery Overhang Moment. 11.1 Pressure Drop Data for Smooth Tubing VI ### Page 10 (OCR) Rev. 1977 AC 43.13-2A CONTENTS-Continued LIST OF ILLUSTRATIONS- Continued Figure Page 12.1 Extruded Track, Anchor Plates, and Associated Fittings 90 12.2 Miscellaneous Litter, Berth, and Cargo Tiedown Fittings 91 12.3 Typical Attachment of Fittings to Honeycomb Structures 92 12.4 Installation of Underfloor Support for Extruded Track - 93 12.5 Typical Cargo Tiedown Straps and Cargo Nets - 94 13.1 Sealant Application 98 13.2 Fair-lead Feed Through 99 13.3 Typical Connector 99 13.4 Coaxial Connectors 100 LIST OF TABLES Table Page 11.1 Instrument Air Consumption ------------- 86 11.2 Equivalent Straight Tube Line Drops for 90° Elbows ------------------- 86 11.3 Differential Pressure Across Instrument Inlet and Outlet --------------- 86 VII ### Page 11 (OCR) Rev. 1977 AC 43.13-2A Chapter 1. STRUCTURAL DATA 1. GENERAL. The minimum airworthiness requirements are those under which the aircraft was type certificated. Addition or removal of equipment involving changes in weight could affect the structural integrity, weight, balance, flight characteristics, or performance of ** Due to differences among various aircraft designs in flight and ground load factors, contact the aircraft manufacturer for the load factors required for a given model and location. In lieu of specific information, the factors used for FAR 23 utility category are acceptable for aircraft with never exceed speed of 250 knots or less and the factors used for FAR 23 acrobatic category for all other transport aircraft. 2. STATIC LOADS. Utilize equipment supporting The following is an example of determining structure and attachments that are capable of the static test loads for a 7-pound piece of equipwithstanding the additional inertia forces ("g." ment to be installed in a utility category alrload factors) imposed by weight of equipment craft (FAR Part 23). installed. Load factors are defined as follows: a. Limit Load Factors are the maximum load Load Factors Static Test Loads factors which may be expected during service (From the above table) (Load factor × 7 pounds) (the maneuvering, gust, or ground load factors Sideward 1.5g 10.5 pounds established by the manufacturer for type cer- Upward Downward 6.6g 3.0g 21.0 pounds 46.2 pounds tification). Forward 9.0g 63.0 pounds b. Ultimate Load Factors are the limit load factors multiplied by a prescribed factor of safety. Certain loads, such as the minimum ultimate inertia forces prescribed for emergency landing conditions, are given directly in terms of ultimate loads. When an additional load is to be added to structure already supporting previously installed equipment, determine the capability of the structure to support the total load (previous load plus added load). c. Static Test Load Factors are the ultimate load 3. STATIC TESTS. factors multiplied by prescribed casting, fitting, Caution: The aircraft and/or equipment can bearing, and/or other special factors. Where no be damaged in applying static loads, particularly special factors apply, the static test load factors if careless or improper procedure is used. direction (up, down, side, fore, and aft). greater of the maneuvering, gust, ground, and inertia load static test load factors for each are equal to the ultimate load factors. d. Critical Static Test Load Factors are the related aircraft structure. Static test loads may exceed the yield limits of the assemblies being It is recommended, whenever practicable, that static testing be conducted on a duplicate installa- tion in a jig or mockup which simulates the Static tests using the following load factors are substantiated and can result in partially sheared acceptable for equipment installations: fasteners, elongated holes, or other damage which Normal- may not be visible unless the structure is dis- Direction of Utility Acrobatic Transport • Rotorcraft assembled. If the structure is materially weak- Force Applied (CAR 3) FAR 23 (CAR 3) PAR 23 (CAR 4b) FAR 25 FAR 27, 29 (CAR 6, 7) ened during testing, it may fail at a later date. Sideward 1.5g 1.5g 1.5g 2.0g Riveted sheet metal and composite laminate con- Upward 3.0g 4.5g 1.5g struction methods especially do not lend them- Forward* Downward 9.0g 6.6g 9.0g 9.0g 9.0g 4.0g 4.0g selves to easy detection of such damage. To conduct static tests: * When equipment mounting is located externally to one side, or forward of occupants, a forward load factor a. Determine the weight and center of gravity of 2.0g is sufficient. position of the equipment item. Chap 1 ### Page 12 (OCR) AC 43.13-2A b. Make actual or simulated installation of attachment in the aircraft or preferably on a jig using the applicable static test load factors. c. Determine the critical ultimate load factors for the up, down, side, fore, and aft directions. A hypothetical example which follows steps (1) through (4) below is shown in figure 1.1. (1) Convert the gust, maneuvering, and ground load factors obtained from the manufacturer or FAA engineering to ultimate load factors. Unless otherwise specified in the airworthiness standards applicable to the aircraft, ultimate load factors are limit load factors multiplied by a 1.5 safety factor. (See columns 1, 2, and 3 for items A, B, and C of fig. 1.1.) (2) Determine the ultimate inertia load forces for the emergency landing conditions as prescribed in the applicable airworthiness standards. (See items D and E, column 3, of fig. 1.1.) (3) Determine what additional load factors are applicable to the specific seat, litter, berth, or cargo tiedown device installation. The ultimate load factors are then multiplied by these factors to obtain the static test factors. (To simplify this example, only the seat, litter, berth, and safety belt attachment factor of 1.33 was assumed to be applicable. See Item E, column 4, of fig. 1.1.) (4) Select the highest static test load factors obtained in Steps 1, 2, and/or 3 for each direction (up, down, side, fore, and aft). These factors are the critical static test load factors used to compute the static test load. (See column 6 of fig. 1.1.) d. Apply load at center of gravity position (of equipment item or dummy) by any suitable means that will demonstrate that the attachment and structure are capable of supporting the required loads. When no damage or permanent deformation occurs after 3 seconds of applied static load, the structure and attachments are acceptable. Should permanent deformation occur after 3 seconds, repair or replace the deformed structure to return it to its normal configuration and strength. Additional load testing is not necessary. Rev. 1977 4. MATERIALS. Use materials conforming to an accepted standard such as AN, NAS, TSO, or MIL-SPEC. 5. FABRICATION. When a fabrication process which requires close control is used, employ methods which produce consistently sound structure that is compatible with the aircraft structure. 6. FASTENERS. Use hardware conforming to an accepted standard such as AN, NAS, TSO, or MIL-SPEC. Attach equipment so as to prevent loosening in service due to vibration. 7. PROTECTION AGAINST DETERIORATION. Provide protection against deterioration or loss of strength due to corrosion, abrasion, electrolytic action, or other causes. 8. PROVISIONS FOR INSPECTION. Provide adequate provisions to permit close examination of equipment or adjacent parts of the aircraft that regularly require inspection, adjustment, lubrication, etc. 9. EFFECTS ON WEIGHT AND BALANCE. Assure that the altered aircraft can be operated within the weight and center of gravity ranges listed in the FAA Type Certificate (T.C.) Data Sheet or Aircraft Listing. Determine that the altered aircraft will not exceed maximum gross weight. (If applicable, correct the loading schedule to reflect the current loading procedure.) Consult Advisory Circular 43.13-1A, Acceptable Methods, Techniques, and Practices--Aircraft Inspection and Repairi for Weight and Balance Computation Procedures. 10. EFFECTS ON SAFE OPERATION. Install equipment in a manner that will not interfere with or adversely affect the safe operation of the aircraft (controls, navigation equipment operation, etc.). 11. CONTROLS AND INDICATORS. Locate and identify equipment controls and indicators so they can be operated and read from the appropriate crewmember position. 12. PLACARDING. Label equipment requiring identification and, if necessary, placard operational instructions. Amend weight and balance information as required. 13.-20. [RESERVED] Chap 1 ### Page 13 (OCR) AC 43.13-2A Static Test 9.3g 5.7g *9. 6g 2. 4g 9. 9g **4.5g 12. 0g Static = Test 9.3g 5.7g 1. 5g 9. 0g 9.6g 4. 2g 9. 9g 6. 0g 1.5g - 3. 0g 12. 0g 2. 0g - 4. 0g LOAD FACTORS X Special — Ultimate 9.30g 5.7g 1. 5g 9. 0g 9. 6g 2. 4g 4. 2g 9. 9g 6. 0g 9. 0g 4. 5g (None) 1. 5g - 3. 0g (None) 1.33 9. 0g (None) 1.33 1. 5g 1.33 - 3. 0g (None) FIGURE 1.1-Hypothetical example of determining static test loads. Utility Category Aircraft (FAR 23) X Safety 1.5 1.5 1.5 1.5 1.5 1. 5 1.5 1.5 1.5 Already Prescribed as Ultimate Limit (None) 6. 2g (None) - 3.8g 1. 0g (None) 6. 0g 6. 4g 1. 6g - 2. 8g (None) 6. 6g 4. 0g *Asterisks denote special load conditions for the situation shown. Direction Fwd_ Down Side- Up-- Aft... Fwd. Down. Down* 'Side_ Up_. Aft_. Fwd. Down Fwd. Fwd.** Down. Side. Up_ Aft_ Fwd Down Side Up_ Aft _ TYPE OF LOAD Maneuvering. Gust (— 30 FPS @ KVc) *For locations aft of fuselage Sta. 73.85. Ground_ Ultimate Inertia Forces for Emer- gency Landing Condition (FAR 23.561). **For Separate cargo compartments. Ultimate Inertia Forces for Emer- gency Landing Condition For Seat, Litter, & Berth Attachment to Air- craft Structure (FAR 23.785). E. Chap 1 ### Page 14 (OCR) Rev. 1977 AC 43.13-2A Chapter 2. RADIO INSTALLATIONS 21. INSTALLATION. When installing radio equipment, first consider areas or locations designated by the airframe manufacturer and use factory supplied brackets or racks. Follow the aircraft manufacturer's installation instructions. When this information is not available, use locations in the aircraft of known load carrying capabilities. Baggage compartments and cabins or cockpit floors are good mounting platforms providing the floor attachments meet the strength requirements. Another method is to fabricate support racks, brackets, or shelves, and attach them to the aircraft structure to provide a mounting that will withstand the inertia forces stipulated in chapter 1 of this handbook. a. Determine that the location and installation of radio equipment provides: (1) Sufficient air circulation to avoid overheating. (2) Sufficient clearance between high temperature areas of equipment and readily flammable (3) Protection from potentially hazardous fluids and/or fumes; e.g., water (condensation), fuel, hydraulic fluid, or oxygen units. (4) Protection against damage from baggage or by seat deflection. (5) Sufficient clearance to avoid equipment striking adjacent parts of the aircraft or other equipment. (6) Minimum interference to other installed navigational equipment from the emission of radar/pulse frequencies or from electromagnetic and is applicable to the installation of radio units in instrument panels. a. Stationary Instrument Panels-Nonstructural and Structural. The stationary instrument panel in some aircraft is primary structure. Prior to making any additional "cutouts" or enlargement of an existing "cutout" determine if the panel is primary structure. tural, make additional "cutouts" or the enlargement of existing "cutouts" in accordance with the aircraft manufacturer's instructions, or substantiate the structural integrity of the altered panel in a manner acceptable to the Administrator. Radius all corners and remove all burrs from "cutout" and drilled holes. b. Added Equipment - Stationary Instrument Panel. When radio equipment is to be installed in a stationary panel already supporting instruments, glove compartments, etc., determine the capability of the panel to support the total load. c. Case Support. To minimize the load on a stationary instrument panel, whenever practicable, install a support between the rear (or side) surface of the radio case and a nearby structural member of the aircraft. d. Added Equipment-Shockmounted Panels. When installing radio equipment designed for use in shockmounted panels, total accumulated weight of equipment installed must not exceed the weight carrying capabilities of the shockmounts. Determine that the structure to which the shockmounts are connected is satisfactory for the added weight. 22. EQUIPMENT MANUFACTURER'S INSTRUC- TIONS. Installation instructions provided by the aircraft radio manufacturer are acceptable guidelines when adapted to the aircraft in accordance with data contained in this chapter. 23. INSTRUMENT PANEL MOUNTING. Data in this paragraph is supplemented by chapter 2 of AC 43.13-1A, 'Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repair" e. Existing Factory Fasteners. When possible, utilize existing plate nuts and machine screws provided by the aircraft manufacturer for attachment of the radio case or rack. If additional fastening is required, use machine screws and elastic stop nuts (preferably plate nuts). f. Magnetic Direction Indicator. As a function of the radio installation, determine if it is necessary to compensate the compass. Install a suitable placard which indicates the compass error Chap 2 ### Page 15 (OCR) AC 43.13-2A Rev. 1977 with the radios) on and off. The receiver (s) should be tuned through the low, middle and higher frequencies to cover all contingencies involving the operation of relays which would cause electromagnetic induction to the magnetic compass. When inverters are installed, deter- Typical Compass Calibration Card FOR N 30 60 120 150 mine what effect their operation has on the mag. netic compass. Maximum acceptable deviation in level flight is 10° on any heading. The following is an example of a typical compass calibration card. 24. OTHER MOUNTING AREAS. The following are acceptable methods for installing radio equipment at other than instrument panel locations. a. Shockmounted Units. Radio On. (1) Wood or Composition Flooring. Secure the Steer 4° 35° 63° 93° 123° 154° shockmounted base assembly (suitable to radio Radio Off. unit) directly to the floor using machine screws. Steer 358° 27° 58° 88° 118° | 148° Add a doubler to the bottom of the floor thereby sandwiching the composition floor between each FOR S 210 240 300 330 shockmount foot and the doubler. Subsequent removal and reinstallation of the shockmount Radio On. foot will be facilitated if plate nuts are secured Steer 183° 214° 224° 274° 304° 337° to the doubler. Where practicable, use small Radio Off. Steer 178° 208° 238° 268° 293° 327° retaining screws to keep the doubler in position. Install a ground strap between the radio rack and metal structure of the aircraft. Typical Locations of Rodio Equipment FIGURE 2.1.-Typical radio installations in stationary instrument panels. 3.0g (up) Wood 1.5g (side) Flooring MACHINE SCREW **NORMAL-UTILITY** **CATEGORY** 2.0 (fwd) 6.6g (down) Doubler 4 Plate Nur FIGURE 2.3.--Typical shockmounted base. Machine screws and self-locking nuts. (2) Metal Flooring. Secure the shockmounted base assembly directly to the floor using machine screws, washers, and self-locking nuts. Floor area under and around the radio mounting bases may require installation of doublers or other reinforcement to prevent flexing and cracking. Installation of plate nuts on the floor or doubler Rivets or Machine screws will facilitate removal and installation of the and self-locking nuts. shockmounts. Install a ground strap between FIGURE 2.2.-Typical panel mount. the shockmount foot and the radio rack. Chap 2 ### Page 16 (OCR) Rev. 1977 Moral Flooring FIGURE 2.4.-Typical shockmounted base. b. Rigid-Mounted Unit Base. Secure radio mounting base plate (s) to floor (wood, composite, or metal) using machine screws as shown in figure 2.5. Use a reinforcing plate or large area washers or equivalent under wood or composite flooring. When mounting base is secured to wood or composite material, install a ground strap between the base and aircraft metal structure. **FLOOR** **MACHINE SCREWS** **MACHINE SCREWS AND** SELF-LOCKING NUTS FIGURE 2.5.-Typical rigid baseplate mount. **25. FABRICATION OF SUPPORTING BRACKETS** **FOR ATTACHMENT TO STRUCTURE OTHER THAN FLOORING.** a. Typical supporting brackets usually consist of a shelf or platform upon which the radio unit mounting base assembly can be installed in the same manner as described in applicable paragraph 24. b. Fabricate bracket in accordance with good aircraft design, layout, assembly practices, and workmanship to obtain results compatible with the airframe structure. Generally the thickness of bracket material will depend on the size or area of the platform and load it must sustain in accordance with provisions set forth in chapter 1 of this handbook. Chap 2 AC 43.13-2A c. Use a rivet size and pattern compatible with the aircraft structure to provide the strength needed to assure support of the loads imposed under all flight and landing conditions. I" clearance with seat occupied and subjected to maximum seat spring deflection (ú,6 x 170 pounde) downward (load). • Floor Unit a Plywood floor AN 970 flat washer a t Machine scrow and To increase the strangth at floor attachment points, metal reinforcement may be installed as needed. FIGURE 2.6.-Typical underseat installation. **26. SUPPORTING STRUCTURE REINFORCEMENT.** a. Attach equipment supporting structure to the aircraft so that its supported load will be transmitted to aircraft structural members. If direct attachment to existing structure (bulkheads, horizontal stringers, etc.) is not feasible, add the necessary stringers, doublers, bulkhead flange reinforcements, etc., to provide adequate support and assure load transfer to primary structure. o. Placard. Fasten on the shelf or bracket a permanent placard (as the example below) stating the design load which the installed structure has been determined to be capable of supporting. Shelf load not to exceed Ibs. ### Page 17 (OCR) AC 43.13-2A Rev. 1977 **FUSELAGE TO SUIT 2T** • 7 T (min) EQUIP. (min) • 0 S31V7d InN- BULB ANGLE- FIGURE 2.7.-Typical remote unit mounting base-vertical or horizontal. Chap 2 ### Page 18 (OCR) Rev. 1977 BULKHEAD Use standard aircraft practices and procedures for fabrication and attachment of shelf. Reinforce fore and aft corners with gussets or bulb angle. FIGURE 2.8.-Typical shelf installation. Fabricate platform using 2017T4(17ST) or equivalent. Apply standard aircraft practices for fabrication and installation. The equipment manufacturer mounting bases that meet load requirements and can be utilized are acceptable. 27. ELECTRICAL SYSTEMS. The following data in addition to that shown in chapter 11 in AC 43.13-1A is applicable to radio installations. a. Installation of Wiring. (1) Use a type and design satisfactory for the purpose intended. (2) Install in a manner to be suitably protected from fuel, oil, water, other detrimental substances and abrasion damages. b. Power Sources. (1) Connect radio electrical systems to the aircraft electrical system at a terminal strip, or use a plug and receptacle connection. (2) Radio electrical systems must function properly whether connected in combination or independently. Chap 2 AC 43.13-2A c. Protective Devices. (1) Incorporate a "trip free" resettable type circuit breaker or a fuse in the power supply from the bus. Mount in a manner accessible to a crewmember during flight for circuit breaker resetting or fuse replacement. (2) Select circuit breakers or fuses that will provide adequate protection against overloading of the radio system circuits. (3) Connect all leads in such a manner that the master switch of the aircraft will interrupt the circuit when the master switch is opened. (4) Radio system controls are to provide independent operation of each system installed and are to be clearly placarded to identify their function relative to the unit of equipment they operate. d. Available Power Supply. To preclude overloading the electric power system of the aircraft when additional equipment is added, make an L **1. BULB ANGLE** **2. STIFFENING FLANGE OR ANGLE** **AT ENDS OF PLATFORM** 3. REINFORCEMENT ANGLE FOR BULKHEAD FIGURE 2.9.-Typical rail platform installationaluminum alloy structure. ### Page 19 (OCR) AC 43.13-2A Rev. 1977 NOTE: AVOID PLACING ADDITIONAL **CLAMPS IN CENTER THIRD** **OF TUBE MEMBERS** **LOCATE CLAMPS CLOSE TO** **WELD CLUSTERS** **AN - BOLT** - PLYWOOD AN970- WASHERS AN742-CLAMP -AN 365-NUT CLAMP DETAIL FrauRE 2.10.-Typical attachment of support structure to tubular frame. electrical load analysis to determine whether the available power is adequate. Radio equipment must operate satisfactorily throughout the voltage range of the aircraft electrical system under taxi, takeoff, slow cruise, normal cruise, and landing operating conditions. If night and instrument flight is contemplated, compute the electrical load analysis for the above flight regimes under the most adverse operating conditions. e. Wire Bundle Separation from Flammable Fluid Lines. (1) Physically separate radio electric wire bundles from lines or equipment containing oil, fuel, hydraulic fluid, alcohol, or oxygen. (2) Mount radio electrical wire bundles above flammable fluid lines and securely clamp to structure. (In no case must radio electrical wire bundles be clamped to lines containing flammable fluids.) f. Cable Attachment to Shockmounted Units. (1) Route and support electrical wire bundles and mechanical cables in a manner that will allow normal motion of equipment without strain or damage to the wire bundles or mechanical cables. g. Radio Bonding. It is advisable to bond radio equipment to the aircraft in order to provide a low impedance ground and to minimize radio interference from static electrical charges. When bonding is used, observe the following: (1) Keep bonding jumpers as short as (2) Prepare bonded surfaces for best contact (resistance of connections should not exceed .003 ohm). (3) Avoid use of solder to attach bonding jumpers. Clamps and screws are preferred. (4) For bonding aluminum alloy, use aluminum alloy or tinned or cadmium-plated copper jumpers. Use brass or bronze jumpers on steel (5) When contact between dissimilar metals cannot be avoided, put a protective coating over the finished connection to minimize corrosion. 10 Chap 2 ### Page 20 (OCR) Rev. 1977 2017 T4 Bulb Angle or equivalent tubing. Mounting Base - 2017 T4 or equivalent Plate Nut Airframe Tubing Secure mounting base to airframe tubing with loop-type clamp (AN 742 or equivalent), machine screws, and self-locking nuts. FIGURE 2.11.-Typical mounting plates for remote location of radio units. h. Radio Interference. Radio interference generated by aircraft components can be eliminated or reduced by exercising proper precautions. The following paragraphs contain two major sources of interference which affect audio recep- (1) Ignition system interference can be minimized with shielding. To be effective, all parts of the ignition system should be shielded in metal in order to eliminate noise resulting from RF radiation. (a) A metallic braid covering and special end connectors are effective between the magneto and shielded type spark plugs. (b) Shield the primary lead to the magneto. (c) Provide a shielded metal cover for the magneto if it is not of a shielded type. (d) Provide a tight metal-to-metal contact of all connections in the shielding system. (e) If it is not feasible to shield the engine ignition system, the engine ignition noise may be suppressed by replacing the spark plugs with resistor spark plugs of a type approved for Chap 2 AC 43.13-2A (f) If it is found that despite shielding of the ignition wiring and plugs an intolerable noise level is present in the radio system, it may be necessary to provide a filter between the magneto and magneto switch to reduce the noise. This may consist of a single bypass capacitor or a combination of capacitors and choke coils. When this is done, the shielding between the filter and magneto switch can usually be eliminated and the special shielded magneto switch need not be used. (g) Supporting brackets and wiring details for magneto filters should be in conformance with standard aircraft electrical practice. The reliability of the magneto filter installations should be at least equivalent to that of the remainder of the magneto ground lead installation. (2) Inverter interference is noticeable by a constant noise or hash induced and amplified by the audio circuits of the communication or navigation systems. This noise level can reach such a magnitude that all intelligence of audio reception is lost. Inverter interference can be effectively minimized or eliminated by observing the following precautions during installation: (a) Locate the inverters in an area separated from other electronic equipment. (b) Assure that the inverter input and output wires are separated. (c) Shield the inverter output wires and ground the shields at the inverter end only. (d) Make sure the inverter case is adequately bonded to the airframe. i. Mutual Radio Frequency Interference in DME/ ATC Radar Beacon Systems. Distance measuring equipment (DME) and ATC radar beacon (transponder) systems operate within the same frequency range. Therefore, simultaneous operation of two or more of these systems may result in mutual RF interference. Certain makes of DME and transponder equipment have intersystem suppressor circuitry designed to eliminate mutual interference. When such connections are provided, follow the manufacturer's recommendations for use and wire bundle installation. ### Page 21 (OCR) AC 43.13-2A 28. EMERGENCY LOCATOR TRANSMITTER (ELT) INSTALLATIONS. The ELT unit should be attached to the airframe or other solid structures. Airframe preparation for either vertical or shelftype mountings is displayed in figures 2.7 and 2.8. The equipment manufacturer mounting bases that meet load requirements and can be utilized are acceptable. 12 Rev. 1977 The installation of the ELT antenna should be located as far as practicable from other installed antennas. Methods for securing whiptype antennas to the structure are shown in figures 3.1 and 3.3. Follow the manufacturer's installation procedures when available. 29.-35. [RESERVED] Chap 2 ### Page 22 (OCR) Rev. 1977 AC 43.13-2A Chapter 3. ANTENNA INSTALLATIONS 36. PERFORMANCE. For proper performance, it is important that the radio equipment manufacturer's instructions be carefully followed in matching and coupling the antenna to the radio equipment. a. The location of the antenna is of primary importance. When selecting a mounting position, consideration should be given but not limited to the following: (1) Obstruction to signal reception by aircraft or aircraft components. (2) Ignition noise (RF radiation pickup). (3) Vibration. (4) Flutter. (5) Instrument static source interference. b. Attach antenna mounting (masts, base receptacles, and/or supporting brackets) so that the loads imposed (e.g., air, ice, etc.) are transmitted to the aircraft structure. **37. VHF ANTENNA-WHIP.** a. Locate this type antenna so that there is a minimum of structure between it and the ground radio stations. The antenna may be mounted on the top or bottom of the fuselage. It is not advisable to mount the antenna on the cowl forward of the windshield because a lightning strike might possibly blind the pilot. b. Methods of securing whip antennas to the structure are shown in figures 3.1 and 3.3. **ANTENNA ROD HEX. NUT -FLAT WASHER (BRASS)** **MACHINE -GASKET SCREW** **CERAMIC BUSHING (FEMALE)** - GASKET **SKIN OF AIRPLANE SKIN OF AIRPLANE** • GASKET - CERAMIC BUSHING (MALE) HEX. NUT CABLE CLAMP (PART OF ANTENNA CABLE) • GASKET •FLAT WASHER **LOCK WASHER** • LOCK WASHER ANTENNA CABLE HEX. NUT **ANTENNA CABLE LUG** FLAT WASHER SOLDER LUG •LOCK WASHER (PART OF ANTENNA CABLE) HEX. NUT FIGURE 3.1.-Typical whip antenna installation. Chap 3 13 ### Page 23 (OCR) AC 43.13-2A Rev. 1977 c. On fabric-covered aircraft or aircraft with other types of nonmetallic skin, the manufacturer's recommendations should be followed in order to provide the necessary ground plane. An acceptable method of accomplishing this is by providing a number of metal foil strips in a radial position from the antenna base and secured under the fabric or wood skin of the aircraft. (See fig. 3.2.) **WHIP** **ANTENNA** (2) The approximate drag load an antenna is required to withstand can be determined by the following formula: D=.000327 AV2 (The formula includes a 90 percent reduction factor for streamline shape of antenna.) Where D is the drag load on the antenna in Ibs., A is the frontal area of the antenna in sq. It., and V is the Ve of the aircraft in m.p.h. The frontal area of typical antennas are approximately as follows: Antenna Area (Fig. 3.4) (sq. ft.) a .073 b .135 C - = J .135 d .025 .045 Example: Antenna "b" at 250 m.p.h. D=.000327X.135X (250) 2 =.000327 X.135 × 62,500 =2.75 1bs. **METAL FOIL UNDER** FABRIC OR WOOD SKIN NOTE: THE LENGTH OF EACH FOIL ANTENNAS SUPPLIED BENT 45° ARE TO MOUNT **WITH END POINTING TO RADIAL SHOULD BE AT LEAST EQUAL REAR OF AIRCRAFT.** **TO THE ANTENNA LENGTH.** FIGURE 3.2.-Antenna ground plane for nonmetallic aircraft. **38. VHF ANTENNA-RIGID.** a. When it is necessary to cover a broader frequency range than can be covered by a whip an- ANTENNA ASS'Y, FLAT WASHER is resonant over a much broader frequency range. tenna, a blade type should be used because it NEOPRENE WASHER ANTENNA MOUNTING However, a broadband antenna is not as efficient BOLT as a small diameter whip antenna and, accord- TRUSS HEAD ingly, should not be used with relatively low output transmitters (under 5 watts). LUTE SKINN (1) The antennas shown in figure 3.4 are REINFORCING PLATE normally installed at a point on the fuselage - COVER STOP NUT directly above the cabin or baggage compartment. GROUND SHIELDING UNDER MOUNTING BCLT. • LOCK WASHER. When a rigid antenna is installed on the - BOLT vertical stabilizer, evaluate the flutter and vibration characteristics of the installation. FIGURE 3.3.-Typical shockmounted antenna installation. 14 Chap 3 ### Page 24 (OCR) e FIGURE 3.4.-Typical rigid antennas. b. Installation of Rigid Antennas. (1) Place a template similar to figure 3.5 on the fore-and-aft centerline at the desired location. Drill the mounting holes and the correct diameter hole for the transmission line cable in the fuselage skin. (2) Install a reinforcing doubler of sufficient thickness to reinforce the aircraft skin. The length and width of the plate should approximate that illustrated in figures 3.6 or 3.8. (3) Install antenna on fuselage, making sure that the mounting bolts are tightened firmly against the reinforcing doubler, and that the mast is drawn tight against the gasket. When a gasket is not used, seal the crack between the mast and fuselage with a sealer, such as zinc chromate paste or equivalent. **SUFFICIENT SIZE TO** **ACCOMODATE TRANSMISSION** LINE CABLE C/1 NO. 18 DRILL FIGURE 3.5.-Typical antenna mounting template. Chap 3 251-957 0 - 77 - 2 AC 43.13-2A (4) Route transmission line cable to the receiver, secure the cable firmly along its entire length at intervals of approximately 2 feet, and take care to prevent fouling of control cables. **39. VHF NAVIGATION RECEIVING ANTENNAS.** Locate antennas for omnirange (VOR), and instrument landing system (ILS) localizer receivers at a position on the aircraft where they will have the greatest sensitivity for the desired signals and minimum response to undesired signals such as electrical energy radiated by the engine ignition system. A good location for the VOR localizer receiving antenna on many small **ANTENNA** **FUSELAGE SKIN** **EXISTING STRINGERS** **VIEW A-A** **REINFORCING DOUBLER APPROXIMATELY ONE ALCLAD 2024-T3** INCH SPACING OF 1/8" MIN. DIA. RIVET 1 1/2" EDGE DISTANCE MIN. FIGURE 3.6.-Typical antenna installation on a skin panel. ### Page 25 (OCR) AC 43.13-2A Rev. 1977 manufacturer's installation procedures and assure that the balun is properly grounded to the airframe. Refer to AC 43.13-14 'Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repairi chapter 11, for acceptable bonding practices. **ANTENNA** DIPOLES FIGURE 3.7.-Preferred VOR antenna location for maximum signal pickup with minimum ignition interference. **CUT STRINGER** **REINFORCING DOUBLER OF ALCLAD 2024 - I2 BALUN** **WIRE WRAPPED AND** **SOLDERED TO SHIELD** **AIRFRAME GROUND FRAME** **STRINGER** **STRINGER** **DOUBLER FLANGES RIVEIS ADDED** **AT LEAST THE HEIGHT AS SHOWN ON** **OF CUT STRINGER RAMES, STRINGE** **AND DOUBLE** • EXISTING RIVETS FIGURE 3.8.-Typical antenna installation involving a cut stringer. airplanes is over the forward part of the cabin. Mount the rigid V-type antenna so that the apex of the "V" points forward and the plane of the "V" is level in normal flight. a. VOR Antenna Balun and Transmission Lines. A dual element or balanced antenna system requires a balun or an impedence matching device for maximum signal transfer into an unbalanced coaxial cable. Rigid antennas, as displayed in figure 3.4, incorporate a balun as an integral component of the antenna assembly. Follow the 16 FIGURE 3.9 Typical dipole antenna assembly. (2) Figure 3.10 is an illustration of a typical VOR antenna balun. A balun made from a section of the transmission line functions as a tuned circuit or transformer which produces a standingwave ratio to provide the desired matching impedence. When the antenna is matched to the line, the line measurement in multiples of wave lengths is not critical. (2) Radio wave velocity is less in a cable than in air; therefore, the wave length in cable will be shorter than in air. Appropriate test equipment must be used for transmission line measurements because the physical and electrical lengths of lines are not always equal. Chap 3 ### Page 26 (OCR) Rev. 1977 AC 43.13-2A a lengthwise slit and peel off the outer jacket. AT TACH Do not nick, cut or damage the shield. **TO ANTENNA DIPOLES** **SHIELD (OUTER - DIELECTRIC** **CONDUCTOR)** **TWISTED SHIELD SHIELD** **REMOVED OUTER JACKET** **CENTER CONDUCTOR** **SINGLE SHIELD** - PROTECTIVE OUTER FIGURE 3.11 Coaxial cable preparation. **COVERING** (2) Comb out the braid and bend back to expose the dielectric. Use a sharp knife to cut the dielectric around the circumference, not quite through to the center conductor. Do not nick or cut the conductor. Remove the dielectric by twisting and pulling. (3) Solder the contact to the center conductor. Use a clean, well-tinned soldering iron. **ROSIN-CORE WIRE WRAPPED AND SOLDERED TO SHIELD SOLDER 60/40** - AIRFRAME GROUND -CENTER CONDUCTOR OPEN TO NAVIGATION RECEIVER FICURE 3.10 Typical VOR antenna balun. (3) The transmission line should be kept as shore a short as possible. Any bends in the cable should have at least a 3-inch radius. Follow the equipment manufacturer's recommendations regarding transmission lines and lengths. b. Assembly of Coaxial Cable Connectors. Optimum performance of a radio system is dependent upon the coaxial cable connector assembly. Follow the manufacturer's assembly instructions. Assure that the cable is not distorted or flattened when cutting. The electrical characteristics of the cable change when flattened or bent sharply. (1) To remove the outer jacket, cut with a sharp knife around the circumference, then make Chap 3 FIGURE 3.12 Coaxial cable tinning center lead. (4) Do not apply heat too long as this will swell the dielectric and make it difficult to insert into the body of the connector. **FLUSH AGAINST DIELECTRIC 7 CONTACT MUST BUTT** **ROSIN-CORE SOLDER 60/40** **FEED SOLDER INTO** HOLE ONLY FIGURE 3.13 Coaxial cable soldering center lead. (5) Install connector body and tighten until secure. Do not overtighten as this will distort the cable. ### Page 27 (OCR) Rev. 1977 AC 43.13-2A VOR receivers. Typical cable lengths are from 22 to 35 inches and multiples of these lengths. - 1/4 Another method of coupling two VOR navigation receivers to a common antenna is by utilizing DO NOT BREAK STRANDS. a device called a coupler or diplexer. This de- **DO NOT NICK CENTER CONDUCTOR-** • 1/8 NUT WASHER GASKET CLAMP TRIM STRANDS WITH SCISSORS VERTICAL "V" ANTENNA FLUSH WITH END OF TAPER → L3/32 **CONTACT FLUSH WITH END OF INSULATOR** **JACK BODY** INSULATOR -CONTACT FICURE 3.14 Coaxial cable install connector body. (6) Use only the crimping tool recommended by the manufacturer, or an equivalent tool when installing connectors which utilize a crimp-type contact. c. Dual VOR/NAV Receiver Installations. Iwo VOR navigation receivers can be connected to a common VOR antenna. This is accomplished by utilizing a coaxial tee connecter (UG-274 A/U) and matched 0.5 wavelength coaxial cable lengths connected from the tee connector to the respective UNDE: JURFACE OF AIRPLANE "p" ANTENNA "L" ANTENNA NOTE: AN "L" TYPE ANTENNA IS SIMILAR TO A "T" **ANTENNA EXCEPT THAT THE LEAD-IN WIRE** **IS CONNECTED TO THE END OF THE ANTENNA INSTEAD OF THE CENTER.** FIGURE 3.15.-Typical wire antenna locations. -TO RECEIVER SOLDER CONNECTION MINIMUM 16 COPPERWELD ANTENNA WIRE' FiGURE 3.16.- Typical marker beacon receiving antenna. Chop 3 | ### Page 28 (OCR) Rev. 1977 1025 STEEL FITTING SPRING **FABRIC WING** • STEEL ROD WASHER NUT METAL WING 2017 FIGURE 3.17.-Typical wing tip installations. vice, in addition to impedance matching, also provides isolation between VOR navigation receivers while keeping line insertion losses at a minimum. 40. RANGE RECEIVING ANTENNAS. Mount T", "L", or "V" antennas on top or bottom of the aircraft with approximately 1-foot clearance from the fuselage and wings. Typical wire antenna installations are shown in figures 3.15 through 3.19. 41. MARKER RECEIVING ANTENNA. The marker receiver operates at a frequency of 75 MHz. In order to keep to a minimum the number of antennas on the aircraft, the marker receiver may utilize the same antenna as the range receiver if that antenna is mounted on the underside of the aircraft. However, both receivers should include provisions to permit simultaneous opera- Chap 3 AC 43.13-2A tion without interference. A whip or other vertical type of antenna should not be used for marker reception since the ground facility transmits from a horizontally polarized antenna. 4130 **OR AN BOLT AND NUT 1025 STEEL** **THIMBLE SPRING** FIN TAIL POST RUDDER: **1025 OR 4130 TUBE** RUBBER TENSION UNIT FIGURe 3.18.-Typical in tip installations. **42. ATC RADAR BEACON (TRANSPONDER) AND** DISTANCE MEASURING EQUIPMENT (DME) AN- TENNAS. Locate these antennas at an unobstructed location on the underside of the fuselage. preferably at the lowest point of the aircraft when in level flight. To the extent practicable, mount the antenna so that the base is horizontal when the aircraft is in cruise attitude. a. Installation. Mount the antennas at least 36 inches away from obstructions and as far as possible from other antennas. Tests have shown that the location of the antenna with respect to obstructions is of greater importance than having the antenna installed in a vertical position. However, signal strength and pattern become noticeably affected as the angle of the antenna approaches 45° from the desired vertical position. On fabric-covered aircraft or aircraft with other types of nonmetallic skin, it will be necessary to provide a flat metallic surface or ground plane" extending at least 12 inches in all direc- 19 ### Page 29 (OCR) AC 43.13-2A Rev. 1977 **ANCHOR CABLE COIL SPRING** **INSULATORS** DOUBLER FIGURE 3.19.- Typical wire antenna installations. FIGURE 3.20.--Preferred position for distance measuring and/or ATO radar beacon antennas. tions from the center of the antenna. Be sure the antenna makes a good, direct electrical connection with the ground plane. Install gaskets, pressurization seals, and/or sealant as required. b. Dual System Installations. When dual ATC radar systems, dual DME systems, or combinations of these systems are installed, determine that the separation between their respective antennas is within the manufacturer's prescribed limits. (See paragraph 27i for mutual interference in DME and ATC radar beacon systems.) c. Antenna Cable. Route the antenna cable in the most direct path practicable. Since losses can be relatively high at these frequencies, follow the equipment manufacturer's recommendations regarding transmission lines and lengths. 43. DIRECTION FINDING ANTENNAS (100 to 1750 KHz). Manual or automatic loop-type antennas are used with direction-finding receivers. The loops are designed for use with a particular receiver. Connecting wires between the loops and receivers are also designed for the specific equipment. Accordingly, only components meeting the specification characteristics of the receiver manufacturer should be used. Chap 3 ### Page 30 (OCR) Rev. 1977 AC 43.13-2А a. Loops enclosed in streamlined housings or exposed loops are satisfactory for external mounting on an aircraft. Loops may also be flush mounted on the aircraft when proper attention is given to avoid interference from metallic structure and skin of the aircraft. c. Installation kits are designed for either top or bottom fuselage mounting of loop and sense antennas, or a combination of these two locations. Particular attention should be paid to the manufacturer's installation instructions for antenna location (top or bottom) and loop output connections, in order to prevent 180° errors in bearing indications. < > **INTERNAL LOOP** ENCLOSED LOOP FIGURE 3.21.-Typical ADF antenna installations. b. Sense antennas are used to resolve radio bearing ambiguity in direction finding systems. The sense and its lead-in must be matched to the input capacitance of the ADF receiver. The sense antenna capacitance is a function of length, spacing between antenna and fuselage, and leadin capacitance. To achieve this antenna/receiver capacitance match, it is important that the sense antenna be installed in accordance with the ADF manufacturer's recommendations for the particular make and model of aircraft. d. Optimum ADF performance is achieved when the "T", rather than the "L", type of sense antenna is used, (see fig. 3.15). The "p" type has a noise cancelling effect due to the antenna cable being connected in the center of the antenna. The "L" type antenna has directional characteristics and may not produce a definite station passage indication as the "T" type does. A whip antenna of a type and dimension recommended by the equipment manufacturer may be used in place of the "T" or "L' types. Methods of installing a whip antenna are shown in figures 3.1, 3.2, and 3.3. e. Because the ADF receiver is susceptible to aircraft radiated noise, antenna lead-ins should be routed so that they are kept away from electric power cables, alternators, solid state power supplies, anti-collision lights, pulse transmitting equipment, etc. They may be rounted against airframe members for extra shielding. f. Loop-ins should be of the recommended type and length. The length of lead-in specified by the manufacturer for a particular installation may be excessive for the physical dimensions between the antenna and receiver. Excessive leadin should not be trimmed, but should be coiled to take up the extra length. Do not coil excess cable in any area subject to electrical noise. g. After completing the installation, it is essential that the loop be calibrated. One acceptable means of compliance is contained in AC 43.13-1A, chapter 15, section 6, paragraphs 848 through 853. 44. ANTENNA INSTALLATION ON PRESSURIZED AIRCRAFT. The use of doublers, to reinforce the aircraft skin to support antennas, is previously described in this chapter. The material contained in this paragraph concerns the methods of apply- Chap 3 21 ### Page 31 (OCR) AC 43.13-2A ing sealant to guard against the passage of air, liquids, and vapors from pressurized structures. a. Typical Antenna Installation Procedure. When the attaching parts and the antenna are ready for installation, clean all faying surfaces with a cleaning solvent. Clean a larger area than that to which sealant is to be applied. Remove the solvent from the faying surfaces by blasting with dry air and wiping with a clean soft cloth. (1) Coat the affected area with the primer specified by the sealant manufacturer. (2) Apply the sealant to one surface, using a spatula or brush, and spread it over the entire faying surface until a uniform thickness of approximately ⅓2-inch is obtained. (See fig. 3.22 and 3.23.) **ANTENNA.** **GASKET** **SEALANT** AIRCRAFT SKIN - DOUBLER FIGURE 3.22. Typical antenna installation. (3) Place the component parts together and install the required fasteners. If permanent fasteners cannot be installed, use temporary fasteners to hold the component parts together until the sealant has cured. Install permanent fasteners with fresh sealant by dipping the fastener in sealant or by filling the fastener hole with sealant. (See fig. 3.24.) (4) Fill holes and joggles by injecting the sealant into the voids and/or cavities. This method is used where the sealant cannot be applied with a spatula or brush. Figure 3.25. (5) Allow the sealant to cure, then remove excess sealant from the periphery of the antenna using a nonmetallic scraper. 22 Rev. 1977 •SEALANT APPROX. 1/32" THICK FIGURE 3.23.-Faying surface sealing. Warning Sealants may contain toxic and/or flammable components. Avoid inhalation of vapors. Supply adequate ventilation and provide a suitable exhaust system. Wear approved respiratory protection while using these materials in confined areas. Do not allow the sealant to come in contact with the skin or eyes. Insure that no source of ignition is present in the working area. b. High Speed Aircraft. The sealant methods described should be used to prevent moisture or water from entering the aircraft and the expulsion of air and vapors when the aircraft structure is pressurized. **SEALANT MAYBE** **APPLIED EITHER TO** **HOLE OR FASTENER** - SEALANT - SEALANT FIGURE 3.24.-Fastener sealing. Chap 3 ### Page 32 (OCR) Rev. 1977 AC 43.13-2A (1) Figure 3.26 displays a blade type antenna mounted on a flat surface using a sealing gasket. This type of installation does not require the application of a sealant. (2) Figures 3.27 and 3.28 display two types of flush mounted antennas. The antenna unit and fiberglass cover are manufactured as one integral assembly. (3) Flush mounted antennas installed on a vertical fin are normally part of the primary structure. The radiating elements of the antenna and the fiberglass cover are individual units. GUN NOZZLE (a) Clean all metal surfaces necessary to insure good electrical bonding contact between the antenna mounting surface and the aircraft (b) After the fiberglass cover is installed, FIGURE 3.25.-Injection sealing. sealer may be applied to fill the space between the fiberglass cover and skin of the vertical fin. Figure 3.29 displays one-half of a vertical fin antenna installation. An identical installation is required on each side. **GASKET** ANTENNA FIOURE 3.27.-ATC flush mounted antenna. FIGURE 3.26.- Blade antenna installation. Chap 3 23 ### Page 33 (OCR) AC 43.13-2A Rev. 1977 • FIGURE 3.28.-Marker beacon flush mounted antenna. FIGURE 3.29.-VOR flush mounted antenna. 45.-50. [RESERVED] 24 Chap 3 | ### Page 34 (OCR) Rev. 1977 AC 43.13-2A Chapter 4. ANTICOLLISION AND SUPPLEMENTARY LIGHT INSTALLATIONS **51. ANTICOLLISION AND SUPPLEMENTARY LIGHT SYSTEMS.** a. Anticollision Lights. The requirements for anticollision lights are included in Federal Aviation Regulations, Parts 23, 25, 27, and 29. Aircraft for which an application for type certificate was made before April 1, 1957, may conform either to the above regulations or to the following (1) Anticollision lights (when installed) should be installed on top of the fuselage or tail in such a location that the light will not be detrimental to the flight crew's vision and will not detract from the conspicuity of the position lights. If there is no acceptable location on top of the fuselage or tail, a bottom fuselage installation may be used. (2) The color of the anticollision light shall be aviation red or white in accordance with the specifications of FAR Part 23. (3) The arrangement of the anticollision light (i.e., number of light sources, beam width, speed of rotation, etc.) should be such as to give an effective flash frequency of not less than 40 and not more than 100 cycles per minute. b. Supplementary lights may be installed in addition to position and anticollision lights required by applicable regulations, provided that the required position and anticollision lights are continuously visible and unmistakably recognizable and their conspicuity is not degraded by such supplementary lights. **52. INTERFERENCE.** a. Crew Vision. Partial masking of the light may be necessary to prevent direct or reflected light rays from any anticollision or supplementary light from interfering with crew vision. Determine that the field of coverage requirements are met. An acceptable method of preventing light reflection from propeller disc, nacelle, or wing surface is an application of nonreflective paint on surfaces which present a reflection problem. Perform a night flight-check to assure that any objectionable light reflection has been eliminated. Enter a notation to that effect in the aircraft records. b. Communication and Navigation. Assure that the installation and operation of any anticollision/supplementary light does not interfere with the performance of installed communication or navigation equipment. Capacitor discharge light (strobe) systems may generate radio frequency interference (RFI). This radiated interference can be induced into the audio circuits of communication or navigation systems and is noticeable by audible clicks in the speaker or headphones. The magnitude of the RFI disturbance does not usually disrupt the intelligence of audio reception. c. Precautions. RFI can be reduced or eliminated by observing the following precautions during installation of capacitor discharge light systems : (1) Locate the power supply at least three (3) feet from any antenna, especially antennas for radio systems that operate in the lower frequency bands. (2) Assure that the lamp unit (flash tube) wires are separated from other aircraft wiring placing particular emphasis on coaxial cables and radio equipment input power wires. (3) Make sure that the power supply case is adequately bonded to the airframe. (4) Ground the shield around the interconnecting wires between the lamp unit and power supply at the power supply end only. 53. MARKINGS AND PLACARDS. Identify each switch for an anticollision/supplementary light and indicate its operation. The aircraft should be flight tested under haze, overcast, and visible Chap 4 25 ### Page 35 (OCR) AC 43.13-2A moisture conditions to ascertain that no interference to pilot vision is produced by operation of these lights. If found unsatisfactory by test, or in the absence of such testing, a placard should be provided to the pilot stating that the appropriate lights be turned off while operating in these conditions. 54. ELECTRICAL INSTALLATION. Install an individual switch for the anticollision light or supplementary light system that is independent of the position light system switch. Data for the installation of wiring, protection device, and generator limitations is contained in chapter 11 of Advisory Circular 43.13-1A, "Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repair." Assure that the terminal voltage at each light is within the limits as prescribed by the manufacturer. **55. ALTERATION OF STRUCTURE.** a. The simplest light installation is to secure the light to a reinforced fuselage skin panel. The reinforcement doubler shall be of equivalent thickness, material, and strength as the existing skin. (Install as shown in fig. 4.1.) **ANTICOLLISION LIGHT OR** **SUPPLEMENTARY LIGHT** **FUSELAGE SKIN** - EXISTING STRINGER **EXISTING STRINGER VIEW A-A** APPROXIMATELY ONE REINFORCING DOUBLER INCH SPACING OF ALCLAD 2024-73 1/8" MIN. DIA. RIVET 1 1/2" EDGE DISTANCE MIN. FIGURE 4.1.- -Typical anticollision or supplementary light installation in a skin panel (unpressurized). Rev. 1977 b. When a formed angle stringer is cut and partially removed, position the reinforcement doubler between the skin and the frame. Doubler to be equivalent to the stringer in thickness and extend lengthwise beyond the adjacent fuselage frames. The distance between the light and the edge of the doubler is twice the height of the doubler flange. (See fig. 4.2 for typical installation.) CUT STRINGER EINFORCING DOULLE ANTICOLLISION LIGHT OR * ALCLAD T. **SUPPLEMENTARY LIGHT FRAME** **STRINGER** **DOUBLE FLANGES** **AT LEAST THE HEIGHT STRINGER** **OF CUT STRINGER STUNGER RIVETS ADDED** **FLANGE AND DOULLER AS SHOWN ON** EXISTING RIVETS FRAMES FIGURE 4.2.-Typical anticollision or supplementary light installation involving a cut stringer (unpressurized). c. Engineering evaluation is required for instal lations involving the cutting of complex formed or extruded stiffeners, fuselage frames, or pressurized skin of pressurized aircraft. d. Vertical stabilizer installations may be made on aircraft if the stabilizer is large enough in cross section to accommodate the light installation, and aircraft flutter and vibration characteristics are not adversely affected. Locate such an installation near a spar, and add formers as required to stiffen the structure near the light. (A typical installation is shown in fig. 4.3.) e. Rudder installations are not recommended because of the possible structural difficulties. However, if such installations are considered, make an engineering evaluation to determine whether the added mass of the light installation Chap 4 ### Page 36 (OCR) Rev. 1977 **MOUNTING SCREW HOLES** **FOR LIGHT TO FAIRING** **ANTICOLUSION LIGHT O** **SUPPLEMENTARY LIGH** **VERTICAL** **STABILIZER** **APPROXIMATELY ONE** INCH SPACING OF 1/8" MIN. DIA. RIVET **LIGHT MOUNTING** **FAIRING** • MOUNTING RING **VERTICAL** **STABILIZER** VIEW A-A NOTE: **SKIN THICKNESS OF MOUNTING RING AND** FAIRING ARE AT LEAST EQUIVALENT FIGURE 4.3.-Typical anticollision or supplementary light installation in a fin tip. will adversely affect the flutter and vibration characteristics of the tail surfaces. f. Pressurized Aircraft Installation. Doubler installation to reinforce the aircraft skin previously described in this chapter is adaptable to pressurized structure with the application of sealant. Sealant is used to prevent moisture or water from entering the aircraft and the expulsion of air when the aircraft structure is pressurized. (1) Sealant procedures for aircraft skin reinforcing doubler and doubler fasteners are contained in paragraph 44, chapter 3 of this manual. The aircraft manufacturer's data may recommend the specific sealant to be used and provide instructions for the application. (2) Figures 4.4 and 4.5 illustrate two different designs of anticollision light assemblies. The application of sealant is required when either type of light assemblies is installed. procedures would be identical for installation of a capacitor discharge (strobe) light system. Chap 4 AC 43.13-2A **AIRCRAFT SKIN** **DOUBLER -** SEALANT FIGURE 4.4.-Typical rotating type anticollision light installation for pressurized aircraft. Caution: Sealant and solvents may contain toxic and/or flammable components. Avoid inhalation of vapors and supply adequate ventilation. Wear appropriate respiratory protection while using these materials in confined areas. Avoid contact with the skin and eyes. **AIRCRAFT SKIN** **DOUBLER** **SEALANT** COAT SCREWS WITH SEALANT FIGURE 4.5.-Typical oscillating type anticollision light installation for pressurized aircraft. **56. GUIDELINES FOR INSTALLATIONS.** a. Prior Approval. Due to the complexity of measurements for intensity, field of coverage, and color, evidence of FAA approval should be obtained from the light manufacturer before installation. b. System Performance. (1) Field of Coverage. Evidence of FAA approval for "field of coverage" should be obtained from the light manufacturer before installation. To insure that the manufacturer's approved field of coverage is applicable to an installation, his mounting tolerance should not be exceeded. ### Page 37 (OCR) Rev. 1977 AC 43.13-2A (2) Obstructed Visibility. Measure all solid angles of obstruction within the required field of coverage. For multiple light installations, coverage between the mounting levels is not necessary. When a multiple light installation is being repeated in each independent measurement (overlap) should be counted. Methods for determining the amount of obstructed visibility are given below; howerer, other methods can give acceptable results. (a) Wall Shadows. This procedure is applicable to installations where shadows from light obstructions appear on a vertical surface such as a hangar wall. Validity is based on two facts: (1) that a vertical surface can approximate a sphere surface if the distance from the light is considerable, and the shadow is reasonably small, and (2) that sphere surface area can be converted to steradians by dividing by the radius squared. 1 Position the aircraft in a darkened hangar so that longitudinal axis is perpendicular to a hangar wall. Level as for weight and bal- The distance extend in each direction within at least 30° above and 30° below the horizontal plane of the aircraft, except that there may be solid angles of obstructed visibility totalling not more than the requirements of paragraph .1401(b) of the applicable airworthiness regulations. 5 For installations where shadows are restricted to directly aft and centered about the longitudinal axis, the following procedures apply: a Establish a point on the wall which corresponds to a line parallel to the longitudinal axis and through the light associated with the shadow. & Measure the distance from the light to the point and determine the area representing 0.15 steradians (A=0.15d'). The distance (d) should be at least 20 feet. c Draw a circle, with the established point as the center, having an area equal to that found in b above. d If the shadow falls within this circle, its position is acceptable. For multiple light installations, consider only the shadow above light level. e If the shadow is partially out of the circle, the shape of the 0.15 steradian area may be varied, but the established point should remain at the center of the area (centroid). (b) Ramp Shadows. This procedure is applicable to shadows which appear on a horizontal plane such as a flat level ramp and will be associated with a top mounted light and a 0.5 steradian limit. Area measurements as described in the wall shadow method should not be used. Some error is inherent, because horizontal angles are measured on a plane displaced from the light source. To compensate for these measinstallations, only shadows above the level of the urement errors, a table (fig. 4.8) is furnished top light should be considered. to convert from measured solid angles to true 3 Compute the solid angle obstruction solid angles. A term "square degrees" is used in steradians, by dividing each shadow area by to aid in the discussions of solid angle measurethe square of the distance from the center of the ments. area to the light. 1 If no masking is required, remove the 4 Evaluate the results to determine if do mine te red cover and attach its clamp ring to the light the system consists of enough lights to illuminate the physical configuration and flight character- the vital areas around the aircraft, considering istics of the aircraft. The field of coverage must base. If masking is used, obtain a clear cover and install it with a duplicate mask. available dark ramp. A minimum of 50 feet 2 Center the aircraft on the largest Chap 4 28 ### Page 38 (OCR) Rev. 1977 AC 43.13-2A radius of clear ramp space will usually be needed. If the installation is symmetrical, clear ramp space will be needed on one side only. Level the aircraft as for weight and balance check. Trim the flaps, rudder, elevator and ailerons. With jacks in place, raise the gear if the measurement results would otherwise be affected. 4 At night, turn on the anticollision light and chalk all shadow patterns (except jack shadows) which appear outside the circle. the light rotates so as to cause the shadows to oscillate, lay twine along the outermost edge of the shadow, and chalk along the line. 5 Move the aircraft to facilitate area 3 Chalk the following marks on the measurements. Measure, sketch, and chart the ramp: ramp marks and other information as shown in a A reference point directly below Figure 4.6. Measure enough points along the the light. shadow patterns to accurately describe them. b A circle centered on this reference Make enough sketches to include all shadows. point having a radius equal to 1.732 times the 6 Convert the above measurements to a height of the light above the point. (The circle graph showing vertical degrees vs. horizontal represents area beyond the minus 30° vertical degrees as shown in Figure 4.7 (first quadrant). The second quadrant will have to be measured tudinal centerline which passes through the ref- erence point. « A line parallel to the aircraft longi- also to obtain the total shadow area for one side. If the shadow pattern is symmetrical, no other measurements will be necessary. Use 1 degree for each space on the graph paper and count the square degrees of shadow. A total of 1,642 or IST QUADRANT less is within limits. A total of 1,872 or more is out of limits, and if possible, should be reduced O EXAMPLE POINT mask. If the count is between 1,642 and 1,872, by adding or moving a light, or by trimming a proceed as follows: For each 1 degree segment of vertical (pitch), convert the counted square degrees to true square degrees by use of the table of Figure 4.8. If the sum of the true square degrees from all segments exceeds 1,642 (0.5 - SHADOW steradian), the installation is out of limits. 17.3 FT. (c) Scale Drawings. Accurate scale draw- 1OFT. BELOW LIGHT POINT ON RAMP ings can be used to measure solid angles of obstruction. Such drawings should have sufficient size and accuracy to give dependable results. In some cases, actual measurements can be combined POINT O with small drawings as shown in Figure 4.9. For • 20 the 6 points established on the left wing, a string can be used to connect the light successively to EXAMPLE POINT O A = 20 FT. each. A protractor can then be used to measure 8 = 20 FT. the vertical angle from level. The horizontal H = 10 FT. DISTANCE TO POINT I = 28.2 FT. angle for each point can be measured on the top HORIZONTAL ANGLE = 45° view (center). When both horizontal and verti- VERTICAL ANGLE = 19.4° cal angles for each point have been determined, they can be plotted on a graph as shown in Fig- FIGURE 4.6.-Ramp shadow pattern. ure 4.10. If a symmetrical condition exists, only the first and second quadrants need be measured. d A line perpendicular to the above The first quadrant contains approximately 450 line, which also passes through the reference square degrees of obstruction. The other wing point. quadrant will double this to 900 square degrees. Chap 4 29 ### Page 39 (OCR) Rev. 1977 AC 43.13-2A • - EXAMPLE O: DEGREES BELOW LIGHT (PITCH) HORIZONTAL ANGLE = 45° PITCH ANGLE = 19.4° TAN PITCH ANGLE = 10/28.2 o 20 - FUSELAGE WING SHADOW 28.2 SHADOW **LIGHT BASE SHADOW 30 20 30 50 60 70 80 90** HORIZONTAL DEGREES FROM NOSE FIGURE 4.7.-Conversion to vertical vs. horizontal degrees. A measurement of the fin and rudder shadows adds approximately 100 for a total of 1,000 square degrees. Since a maximum of 1,642 square degrees is allowed (0.5 steradians), this installation is well within limits. Due to limitations of the method, results within 10% (0.05 steradians) of the limit are questionable. Many times a mask is required to prevent reflections into the cockpit. In the example of figures 4.9 and 4.10, the installed mask blocks the light for +10 horizontal degrees and from -10 to -30 vertical degrees. These 400 square degrees were measured at the • light. When larger drawings are used and no actual aircraft measures are made, vertical angles should not be taken directly from the drawing, but should be computed as follows: 1 On the aircraft side view, measure the vertical distance from a point to the light level. 2 On the top view, measure the distance from the point to the light. 3 Compute: vertical distance Tangent of vertical angle= horizontal distance Pitch Segment (Degrees) Messured Degrees Square Factor Correction Square True Degrees (Degrees) Segment Pitch Measurco Square Degrees Correction Factor True Degrees Square 24-23 22-21 30-29 29-28 28-27 27-26 26-25 25-24 23-22 21-20 20-19 19-18 18-17 17-16 16-15 90 4826 90 • 87036 • 88701 • 87882 • 89493 • 90259 • 90996 • 91706 • 93667 • 94264 • 96363 • 95372 • 92388 • 93042 • 95882 , 94832 78.33 39.03 79.09 12-11 15-14 14-13 13-12 11-10 10-9 7-6 9-8 8-7 3-2 2-1 1-0 6-5 5-4 4-3 • 96815 • 98629 • 98325 • 98902 • 97237 • 97630 • 97992 • 99996 • 99813 • 99905 . 99144 • 99540 • 99966 • 99692 • 99357 TOTAL FIGURE 4.8.-Conversion to true square degrees-1st quadrant. Chap 4 30 ### Page 40 (OCR) Rev. 1977 AC 43.13-2A (d) Flashing Characteristic. Turn off any VERTICAL DISTANCE flashing supplementary lighting. Observe the POINT A flashing of the anticollision light system at a point where each light can be observed independently, and determine that each flashing rate is between 40 and 100 flashes per minute. For multiple light systems, observe at a point where overlap occurs, and determine that the combined flashing rate does not exceed 180 flashes per HORIZONTAL minute. Flashing outside the required field of DISTANCE POINT A HORZ. = 10° coverage is not necessary. -VERT. = 17.5° POINT POINTE H. = 35° DEGREES BELOW LIGHT 10 **LEFT WING STRUT MASK GEAR** HORIZONTAL DEGREES FROM NOSE FIGURE 4.10.-Solid angle blockage1st quadrant. FIGURE 4.9.- -Scale drawings. 57.-60. [RESERVED] Chap 4 31 ### Page 41 (OCR) AC 43.13-2A Rev. 1977 Chapter 5. SKI INSTALLATIONS Section 1. SELECTION OF SKIS 61. DETERMINING ELIGIBILITY OF AIRCRAFT. eligible for ski installations in accordance with this chapter. Eligibility can be determined by Only aircraft approved for operation on skis are referring to the Aircraft Specifications, Type mary of Supplemental Type Certificates, or by Certificate Data Sheets, Aircraft Listing, Sum- contacting the manufacturer. Also determine the to the aircraft to make it eligible for ski opera- need for the nature of any required alterations tion. 62. IDENTIFICATION OF APPROVED MODEL SKIS. Determining that the skis are an approved model plate or placard displayed on the skis. A Techni- can be done by referring to the identification cal Standard Order (TSO) number; Type Cer- ber, if the skis have been approved as a part tificate (TC) number; or an aircraft part num- of the aircraft, will be shown thereon if the skis 63. MAXIMUM LIMIT LOAD RATING. In order are approved models. for an approved ski to be installed on any given -2 5 - 9 10 6. Fabric removed to facilitate inspection 1. Fitting 7. Check Cable 2. Shock Cord 8. Clevis 3. Safety Cable 9. Ski Pedestal 4. Tape 10. Pedestal Height 5. Crust-cutter Cable FIGURE 5.1.-Typical ski installation. 33 Chap 5 ### Page 42 (OCR) AC 43.13-2A aircraft, determine that the maximum limit load rating (L) as specified on the ski identification plate or placard is at least equal to the maximum static load (S) times the limit landing load factor (n) previously determined from static drop tests of the airplane by the aircraft manufacturer. In lieu of a value y determined from such drop tests, a value of y determined from the following formula may be used: n =2.80 + W + 4000 9000 where "W" is the certificated gross weight of the airplane. 34 Rev. 1977 Skis approved for airplanes of greater gross weight than the airplane on which they are to be installed may be used provided the geometry of the ski is similar to that of a ski previously approved for the airplane (not more than 10 percent difference in width or length of contact surface). This limitation is to assure that the performance of the airplane will not be adversely affected by oversize skis. 64. LANDING GEAR MOMENT REACTIONS. In order to avoid excessive moment reactions on the landing gear and attachment structure, the ski pedestal height must not exceed 130 percent of the axle centerline height with the wheel and tire installed. 65. [RESERVED] Chap 5 ### Page 43 (OCR) Rev. 1977 AC 43.13-2A Section 2. CONVERSION AND INSTALLATION 66. HUB-AXLE CLEARANCE. The pedestal hub should fit the axle to provide a clearance of .005" minimum to .020" maximum. Hubs may be bushed to adjust for axle size, using any ferrous or nonferrous metal, hard rubber, or fiber. If rubber or fiber bushings are used, use retaining washers of sufficient size on each side to retain the hub if the bushing should slip or tail. (See fig. 5.2.) 67. CRUST-CUTTER CABLES. Crust-cutter cables are optional. However, when operating in severe crust conditions, it is advisable to have this cable installed to prevent the shock cord from being cut if the nose of the ski breaks through the crust while taxiing. 68. CABLE AND SHOCK CORD ATTACHMENT AND ATTACHMENT FITTINGS. Service reports indicate that failure of the ski itself is not a predominant factor in ski failures. Rigging (improper tension and terminal attachments) and cast-type pedestal material failures are predominant. Usually, failures of the safety cable and shock cord attachment fittings occur at the ski end and not at the fuselage end. Do not attach tension cords and safety cables at the same point on the fuselage fittings. Provide separate means of attaching cables and shock cords at the forward and aft ends of the skis. Usually, approved skis are supplied with cables, shock cord, and fittings; however, the following specifications may be used for their fabrication and installations: retaining washers bushings pedestal hub axle FIoURE 5.2.-Typical hub installation. Chap 5 35 ### Page 44 (OCR) AC 43.13-2A Rev. 1977 Minimum Cable and Shock Cord Sizes Ski Single Double Limit Single Double Crust- Crust- Single Double Load Safety Safety Cutting Cutting Shock Shock Rating Cable Cable Cable Cable Cord Cord 1500-3000 1/8" 1/8" 1/8' 1/8"' 1/2" 1/2" 3000-5000 1/8'' 1/8" 1/2" 5000-7000 5/32' 5/32'1 5/32" 3/4"' 3/4" 7000-9000 3/16" 5/32" 3/4" a. Cables. Make the check cable, safety cable, and crust-cutting cable ends by the splice, swage, or nicopress methods; or if adjustable lengths are desired, use cable clamps. Use standard aircraft hardware only. (Hardware used to attach cables must be compatible with cable size.) b. Shock cord ends may be fabricated by any of the following methods: (1) Make a wrapped splice using a proper size rope thimble and No. 9 cotton cord or .035" (minimum) safety wire. Attach with clevis or spring steel snap fastener. (Do not use cast iron snaps.) (2) Use approved spring-type shock cord end fasteners. c. Fitting Specifications (see figure 5.3) and Installation: (1) Fittings fabricated for 1/8-inch cable or 1/2-inch shock cord shall be at least .065" 1025 steel or its equivalent. (2) Fittings fabricated for 5/32-inch cable or 3/4-inch shock cord shall be at least .080" 1025 steel or its equivalent. FIGURE 5.3.-Typical fuselage fitting. (3) An improperly installed fitting may impose excessive eccentric loads on the fitting and attach bolts and result in failure of the fitting or bolts. 69. PROVISIONS FOR INSPECTION. Aircraft using fabric-covered landing gear may have the lower 4 inches of fabric removed to facilitate inspection of the axle attachment area. (See fig. 5.1.) 70. [RESERVED] 36 Chap 5 ### Page 45 (OCR) Rev. 1977 AC 43.13-2A Section 3. RIGGING OF SKIS 71. LOCATION OF ATTACH FITTINGS ON FUSE- zero to 5-degree ski incidence angle (reference LAGE OR LANDING GEAR. Locate fittings so the figs. 5.4 and 5.6). shock cord and cable angles are not less than b. Adjust length of safety cable to provide 20° when measured in the vertical plane with -20 to - 35-degree ski incidence angle (reference figs. 5.5 and 5.6). the shock absorber in the fully extended position (see angle B, figs. 5.4 and 5.5). 73. TENSION REQUIRED IN MAIN SKI SHOCK CORDS. Apply sufficient shock cord tension to NoTE: Do not attach fittings to wing-brace struts, fore end of the skis to prevent flutter at various except by special approval (manufacturer or FAA). 72. MAIN SKI INCIDENCE ANGLES. (Aircraft leveled and shock absorbers fully extended.) a. Adjust length of check cable to provide a airspeeds and attitudes. Because of the various angles used in attaching shock cord to the skis, shock cord tension cannot be specified. In most installations the downward force applied at the fore end of the ski, sufficient to cause the check FIGURE 5.4.-Main ski at maximum positive incidence (check cable tight). Chap 5 ### Page 46 (OCR) AC 43.13-2A Rev. 1977 FIGURE 5.5.--Main ski at maximum negative incidence (safety cable tight). cable to slacken, should be approximately as b. Where it is possible for the nose ski rigging listed below: to contact the propeller tips due to vibration, install a ¼-inch shock cord to hold the rigging Ski Limit Doronroard out of the propeller arc. Load Force Rating (pounds) 75. TAIL SKI INSTALLATION. a. Use tail skis that have been approved on 1500-3000 20-10 airplanes of approximately the same weight 3000-5000 40-60 (within 10 percent) or select as outlined in sec- 5000-7000 60-120 7000-9000 120-200 74. NOSE SKI INSTALLATION. The nose ski is installed in the same manner as the main skis (see fig. 57) except: a. Adjust length of safety cable to provide - 5- to - 15-degree ski incidence. FIGURE 5.6.-Main ski incidence angles. 38 Chap 5 ### Page 47 (OCR) Rev. 1977 tion 1. Depending upon the type of ski selected, the tail wheel may or may not have to be removed. b. Adjust the length of the limiting cable (ref. fig. 5.8) to allow the ski to turn approximately 35 degrees either side of the straight- AC 43.13-2A forward position with the weight of the airplane resting on the ski. c. The shock cord (ref. fig. 5.8) must be of a length that will hold the ski in the straightforward position during flight. FIGUBE 57.-Typical nose ski installation. Chap 5 39 ### Page 48 (OCR) AC 43.13-2A Rev. 1977 horiz. ref. **LIMITING - SHOCK** CABLE CABLE FIGURE 5.8 Typical tail ski installation. 76.-78. [RESERVED] Chap 5 40 ### Page 49 (OCR) AC 43.13-2A Rev. 1977 Section 4. OPERATION 79. PERFORMANCE INFORMATION. The following FAA policies concern performance data and operational check flights for ski installations. a. For aircraft over 6,000 pounds maximum certificated weight, state the following or similar information in the performance information section of the Airplane Flight Manual and obtain FAA approval. (1) Takeoff. Under the most favorable conditions of smooth packed snow at temperatures approximating 32° F, the skiplane takeoff distance is approximately 10 percent greater than that shown for the landplane. NoTE: In estimating takeoff distances for other conditions, caution should be exercised as lower temperatures or other snow conditions will usually increase these distances. (2) Landing. Under the most favorable conditions of smooth packed snow at temperatures approximately 32° F, the skiplane landing distance is approximately 20 percent greater than that shown for the landplane. NorE: In estimating landing distances for other conditions, caution should be exercised as other temperatures or other snow conditions may either decrease or increase these distances. (3) Climb Performance. In cases where the landing gear is fixed (both landplane and skiplane), where climb requirements are not critical, and the climb reduction is small (30 to 50 feet per minute), the FAA will accept a statement of the approximate reduction in climb performance placed in the performance information section of the Airplane Flight Manual. For larger variations in climb performance, or where the minimum requirements are critical, or where the landing gear of the landplane was retractable, appropriate climb data should be obtained to determine the changes, and new curves, tables, or a note should be incorporated in the Airplane Flight Manual. b. For aircraft of 6,000 pounds or less maximum certificated weight, make the information in 79a available to the pilot in the form of placards, markings, manuals, or any combination thereof. 80. FLIGHT AND HANDLING OPERATIONAL CHECKS. Accomplish an operational check which includes more than one landing to determine the ground-handling characteristics as well as takeoff and landing characteristics. Take note of ski angles during tail high and tail low landings to avoid having the ski dig in or fail from localized stress. Determine there is sufficient ground control to satisfactorily complete a landing run with a turnoff at slow speed in cases where brakes are not provided. In flight, the ski should ride steady with no unusual drag and produce no unsatisfactory flight characteristics. Enter a notation of these checks in the aircraft 81. INTERCHANGING OF SKIS AND WHEELS After the initial installation, removing the skis and reinstalling the wheels or vice versa may be considered a preventive maintenance operation when no weight-and-balance computation is in- 82.-85. [RESERVED] 41 Chap 5 ### Page 50 (OCR) Rev. 1977 AC 43.13-2A Chapter 6. OXYGEN SYSTEM INSTALLATIONS IN NONPRESSURIZED AIRCRAFT 86. SYSTEM REQUIREMENTS. Install oxygen cylinders conforming to Interstate Commerce Commission requirements for gas cylinders which carry the ICC 3A, SAA, or 3HT designation followed by the service pressure metal-stamped on the cylinder. The 3HT designated cylinders must not be used for portable oxygen equipment. a. Tubing. (1) In systems having low pressure (400 p.s.i.), use seamless aluminum alloy or equivalent having an outside diameter of 5/16 inch and a wall thickness of 035". Double flare the ends to attach to fittings. (2) In high-pressure systems (1800 p.s.i.), use 3/16 inch O.D., .035" wall thickness, seamless copper alloy tubing meeting Specification WW- T-779a type N, or stainless steel between the filler valve and the pressure-reducing valve. Silver-solder cone nipples to the ends of the tubing to attach the fittings in accordance with Specification MIL-B-7883. (3) Use 5/16-inch O.D. aluminum alloy tubing after the pressure-reducer (low-pressure (4) Use flexible connections specifically designed for oxygen between all points having relative or differential motion. b. Valves. A slow opening valve is used as a cylinder shutoff valve, or system shutoff valve. Rapid opening and the subsequent sudden and fast discharge of oxygen into the system can cause dangerous heating which could result in fire or explosion of combustibles within the system. c. Regulators. The cylinder or system pressure is reduced to the individual cabin outlets by means of a pressure-reducing regulator which can be manually or automatically controlled. d. Types of Regulators. The four basic types of oxygen systems, classified according to the type of regulator employed, are: (1) Demand type. (2) Diluter-demand type. (3) Pressure-demand type. (4) Continuous-flow type. e. Flow Indicators. (1) A pith-ball flow indicator, vane, wheel anemometer, or lateral pressure indicator which fluctuates with changes in flow or any other satisfactory flow indicator may be used in a continuous flow-type system. (2) An Air Force-Navy flow indicator or equivalent may be used in a diluter-demand type system. Each flow indicator should give positive indication when oxygen flow is occurring. f. Relief Valve. (1) A relief valve is installed in low-pressure oxygen systems to safely relieve excessive pressure, such as caused by overcharging. (2) A relief valve is installed in highpressure oxygen systems to safely relieve excessive pressure, such as caused by heating. g. Gauge. Provide a pressure gauge to show the amount of oxygen in the cylinder. h. Masks. Only masks designed for the particular system should be used. **87. INSTALLATION.** Oxygen systems present a hazard. Therefore, follow the precautions and practices listed below: a. Remove oil, grease (including lip salves, hair oil, etc.), and dirt from hands, clothing, and tools before working with oxygen equipment. b. Prior to cutting the upholstery, check the intended route of the system. Make sure that all system components are kept completely free of oil or grease during installation and locate components so they will not contact or become contaminated by oil or hydraulic lines. c. Keep open ends of cleaned and dried tubing capped or plugged at all times, except during attachment or detachment of parts. Do not use tape, rags, or paper. ### Page 51 (OCR) AC 43.13-2A d. Clean all lines and fittings which have not been cleaned and sealed by one of the following methods: (1) A vapor-degreasing method with stabilized trichlorethylene conforming to Specification MIL-T-7003 or carbon tetrachloride. Blow tubing clean and dry with a stream of clean, dried, water-pumped air, or dry nitrogen (water-vapor content of less than 0.005 milligrams per liter of gas at 70° F and 760 millimeters of mercury pressure). (2) Flush with naptha conforming to Specification TT-N-95; blow clean and dry of all solvent with water-pumped air; flush with antiicing fluid conforming to Specification MIL- F5566 or anhydrous ethyl alcohol; rinse thoroughly with fresh water; and dry thoroughly with a stream of clean, dried, water-pumped air, or by heating at a temperature of 250° to 300° F for one-half hour. (3) Flush with hot inhibited alkaline cleaner until free from oil and grease; rinse thoroughly with fresh water; and dry thoroughly with a stream of clean, dried, water-pumped air, or by heating at a temperature of 250° to 300° F for one-half hour. e. Install lines, fittings, and equipment above and at least 6 inches away from fuel, oil, and hydraulic systems. Use deflector plates where necessary to keep hydraulic fluids away from the lines, fittings, and equipment. f. Allow at least a 2-inch clearance between the plumbing and any flexible control cable or other flexible moving parts of the aircraft. Provide at least ½-inch clearance between the plumbing and any rigid control tubes or other rigid moving parts of the aircraft. g. Allow a 6-inch separation between the plumbing and the flight and engine control cables, and electrical lines When electrical conduit is used, this separation between the plumbing and conduit may be reduced to 2 inches. h. Route the oxygen system tubing, fittings, and equipment away from hot ducts and equipment. Insulate or provide space between these items to prevent heating the oxygen system. Rev. 1977 i. Mount all plumbing in a manner which prevents vibration and chafing. Support 3/16- inch O.D. copper line each 24 inches and 3/16- inch O.D. aluminum each 36 inches with cushioned loop-type line support clamps (AN- 742) or equivalent. ¡. Locate the oxygen supply valve (control valve) so as to allow its operation by the pilot during flight. The cylinder shutoff valve may be used as the supply control valve, if it is operable from the pilot's seat. Manifold plug-in type outlets, which are incorporated in automatic systems, may be considered as oxygen supply valves since the pilot can control the flow of oxygen by engaging and disengaging the plug-in type oxygen mask. NoTE: Locate the oxygen shutoff valve on or as close as practicable to the cylinder to prevent loss of oxygen due to leakage in the system. 88. LOCATION AND MOUNTING. Determine the weight factor and c.g. limits for the installation prior to commencing the installation. a. Mount the cylinder in the baggage compartment or other suitable location in such a position that the shutoff valve is readily accessible. If possible, provide access to this valve from inside the cabin so that it may be turned on in flight in the event that it was not opened prior to takeoff. b. Fasten the cylinder brackets securely to the aircraft, preferably to a frame member or floorboard using AN bolts with fiber or similar locking nuts. Add sufficient plates, gussets, stringers, cross-bracing, or other reinforcements, where necessary, to provide a mounting that will withstand the inertia forces stipulated in chapter 1 of this handbook. c. When cylinders are located where they may be damaged by baggage or stored materials, protect them by a suitable guard or covering & Provide at least ½ inch of clear airspace between any cylinder and a firewall or shroud isolating a designated fire zone. e. Mount the regulator close to the cylinder to avoid long high-pressure lines. f. Store the masks in such a way that there will be a minimum delay in removing and putting them into use. Chap 6 ### Page 52 (OCR) Rev. 1977 • 89. THREAD COMPOUND. Use antiseize or thread-sealing compound conforming to Specification MIL-T-5542-B, or equivalent. a. Do not use compound on aluminum alloy flared tube fittings having straight threads. Proper flaring and tightening should be sufficient to make a flared tube connection leakproof. b. Treat all male-tapered pipe threads with antiseize and sealing compound (MIL-T-5542-B, or tetrafluroethylene tape MIL-T-27730), or equivalent. c. Apply the compound in accordance with the manufacturer's recommendation. Make sure that the compounds are carefully and sparingly applied only to male threads, coating the first three threads from the end of the fitting. Do not use compound on the coupling sleeves or on the outside of the tube flares. **90. FUNCTIONAL TEST.** Before inspection plates, cover plates, or upholstering are replaced, make a system check including at least the following: a. Open cylinder valve slowly and observe the pressure gauge. b. Open supply valve and remove one of the mask tubes and bayonet fittings from one of the masks in the kit. Plug the bayonet into each of the oxygen outlets. A small flow should be noted from each of the outlets. This can be detected by holding the tube to the lips while the bayonet is plugged into an outlet. c. Check the complete system for leaks. This can be done with a soap solution made only from a mild (castile) soap or by leak-detector solution supplied by the oxygen equipment manufacturer. d. If leaks are found, close the cylinder shutoff valve and reduce the pressure in the system by plugging a mask tube into one of the outlets or by carefully loosening one of the connections in the system. When the pressure has been reduced to zero, make the necessary repairs. Repeat the procedure in 90c until no leaks are found in the Caution: Never tighten oxygen system fittings with oxygen pressure applied. Chap 6 AC 43.13-2A e. Test each outlet for leaks at the point where the mask tube plugs in. This can be done by drawing a soap bubble over each of the outlets. Use the solution sparingly to prevent clogging the outlet by soap. Remove all residue to prevent accumulation of dirt. f. Examine the system to determine that the flow of oxygen through each outlet is at least equal to the minimum required by the pertinent requirements at all altitudes at which the aircraft is to be operated. This can be accomplished by one of the following methods: (1) In a continuous flow system when the calibration (inlet pressure vs. flow) of the orifices used at the plug-in outlets is known, the pressure in the low-pressure distribution line can be measured at the point which is subject to the greatest pressure drop. Do this with oxygen flowing from all outlets. The pressure thus measured should indicate a flow equal to or greater than the minimum flow required. (2) In lieu of the above procedure, the flow of oxygen, through the outlet which is subject to the greatest pressure drop, may be measured with all other outlets open. Gas meters, rotometers, or other suitable means may be used to measure flows. (3) The measurement of oxygen flow in a continuous flow system which uses a manually adjusted regulator can be accomplished at sea level. However, in a continuous flow system which uses an automatic-type regulator, it may be necessary to check the flow at maximum altitude which will be encountered during the normal operation of the aircraft. The manufacturer of the particular continuous-flow regulator used should be able to furnish data on the operating characteristics of the regulator from which it can be determined whether a flight check is necessary. (4) The checking of the amount of flow through the various outlets in a diluter-demand or straight-demand system is not necessary since the flow characteristics of the particular regulator being used may be obtained from the manufacturer of the regulator. However, in such systems the availability of oxygen to each regulator should be checked by turning the lever of the diluter-demand regulator to the "100 percent oxygen position and inhaling through the tube via the mask to determine whether the regulator valve and the flow indicator are operating. ### Page 53 (OCR) AC 43.13-2A g. Provide one of the following acceptable means or equivalent to indicate oxygen flow to each user by : (1) Listening for audible indication of oxygen flow. (2) Watching for inflation of the rebreather or reservoir bag. Rev. 1977 (3) Installation of a flow indicator. 91. OPERATING INSTRUCTIONS. Provide instructions appropriate to the type of system and masks installed for the pilot on placards. Include in these instructions a graph or a table which will show the duration of the oxygen supply for the various cylinder pressures and pressure altitudes. ACTUAL DURATION IN HOURS AT VARIOUS ALTITUDES Number of Persons 8000 Ft. 10,000 Ft. 12,000 Ft. 15,000 Ft. 20,000 Ft. Pilot only 7.6 hr. 7.1 hr. 6.7 hr. 6.35 hr. 5.83 hr. Pilot and 1 Passenger 5.07 hr. 4.74 hr. 4.47 hr. 4.24 hr. 3.88 hr. Pilot and 2 Passengers 3.8 hr. 3.55 hr. 3.36 hr. 3.18 hr. 2.92 hr. Pilot and 3 Passengers 3.04 hr. 2.84 hr. 2.68 hr. 2.54 hr. 2.34 hr. Pilot and 4 Passengers 2.53 hr. 2.37 hr. 2.24 hr. 2.12 hr. 1.94 hr. NOTE: The above duration time is based on a fully charged 48 cubic-foot cylinder. For duration using 63 cubic-foot cylinder, multiply any duration by 1.3. FIGURE 6.1-Typical oxygen duration table. Regulator .032 aluminum Hinge Typical example of shroud Cylinder bracke • Large area washer AN bolt and self-locking nut fit contour of cylinder base. Cylinder support block - Cut to Oxygen cylinder bracket Secure with countersunk machine screw or wood screw. NOTE: Enclose cylinder and valve with shroud to prevent damage from baggage. Reinforce floor, if necessary, to withstand the added lood. Use aluminum or plywood of suficient thickness. FIGURE 6.3.---Typical baggage compartment mounting. FIGURE 6.2.-Typical floor mounting. 46 Chap 6 ### Page 54 (OCR) Rev. 1977 AC 43.13-2A Alternate Location Cabin Wall Outlet Gage Regulator Cylinder Outside Filler Valve FroURE 6.4.-Typical oxygen installation in light twin aircraft. 92.-95. [RESERVED] Chap o 47 251-957 0 - 77 - 3 ### Page 55 (OCR) Rev. 1977 AC 43.13-2A Chapter 7. ROTORCRAFT EXTERNAL-LOAD DEVICE INSTALLATIONS Section 1. CARGO SLINGS 96. GENERAL. This section contains structural and design information for the fabrication and installation of a cargo sling used as an externalload attaching means for a Class B rotorcraftload combination operation under FAR Part 133. As an external-load attaching means, a cargo sling" includes a quick-release device and the associated cables, fittings, etc., used for the attachment of the cargo sling to the rotorcraft. 97. QUICK-RELEASE DEVICE. Section 133.43 (d) of the FARs specifies the requirements for the quick-release device. In addition to commercially manufactured helicopter cargo hooks, some surplus military bomb releases meet the requirements of that section. **98. LOCATION OF CARGO RELEASE IN RELATION TO THE ROTORCRAFT'S C.G. LIMITS.** a. An ideal location of the cargo release would be one that allows the line of action to pass through the rotorcraft's center of gravity at all times. (See fig. 7.1, illus. A.) However, with « Lines of + Line of action action Effective pivot point Moment • Arms Arm action Effective pivot Effective pivot pointe F F FIGURE 7.1.-Location of cargo release in relation to the rotorcraft's center of gravity. Chap 7 ### Page 56 (OCR) AC 43.13-2A Rev. 1977 most cargo sling installations, this ideal situation is realized only when the line of action is vertical or near vertical and through the rotorcraft's c.g. (See fig. 7.1, illus. B.) b. Whenever the line of action does not pass through the rotorcraft's c.g. due to the attachment method used, acceleration forces, or aerodynamic forces, the rotorcraft-load combined center of gravity will shift from the rotorcraft's c.g. position. Depending upon the factors involved, the shift may occur along either or both the longitudinal or lateral axes. The amount of shift is dependent upon the force applied (F) and the length of the arm of the line of action. Their product (F x Arm) yields a moment which can be used to determine the rotorcraft-load combined center of gravity. (See fig. 7. 1, illus. C.) If the rotorcraft-load center of gravity is allowed to shift beyond the rotorcraft's approved center of gravity limits, the rotorcraft may become violently uncontrollable. c. Thus, any attachment method or location which will decrease the length of the arm will reduce the distance that the combined center of gravity will shift for a given load (F) and line of action angle. (See fig. 7.1, illus. D.) 99. MAXIMUM EXTERNAL LOAD. The maximum external load (including the weight of the cargo sling) for which authorization is requested may not exceed the rated capacity of the quick-release should not be attached in such a manner as to provide sling-leg angles of less than 30 degrees. 102. MINIMUM SLING-LEG CABLE STRENGHT. An analysis which considered the effects of 30-degree sling angles showed that the minimum cable strength design factor required would be 2.5 times the maximum external load for each leg regardless of the number of legs. Although this is the minimum strength required by Part 133, it may be desirable to double this value to allow for deterioration of the sling-leg cables in service. This will result in a cable strength equal to 5 times the maximum external load. Example: Maximum external load 850 pounds Minimum required sling-leg cable strength 850 × 2.5=2125 Minimum desired sling-leg strength 850 × 2.5 x 2=4250 A 3/16-inch, nonflexible 19-wire cable (MIL-W- 6940) provides a satisfactory cable strength. See figure 4.1, chapter 4, of AC 43.13-14 for a table of breaking strength of steel cable. venience, the cable sizes desired for various loads have been calculated and are tabulated in figure 7.2 based on a factor of 5: Aircraft Cable Size For Each Cargo Sling Leg Махітит MIL-C-5699 External Load and (pounds) MIL-W-6940 MIL-W-1511 MIL-C-5424 1/16 3/32 3/32 100. STATIC TEST. The cargo sling installation 200 3/32 1/8 1/8 must be able to withstand the static load required by FAR 133.43(a). Conduct the test as outlined 300 400 1/8 7/64 1/8 1/8 1/8 5/32 in Chapter 1 of this advisory circular. If re- 500 600 5/32 5/32 5/32 3/16 3/16 3/16 quired during the test, supports may be placed 700 3/16 3/16 3/16 at the landing gear to airframe attach fittings to 800 3/16 3/16 7/32 prevent detrimental deformation of the landing gear due to the weight of the aircraft. 1,000 1,200 900 3/16 7/32 7/32 7/32 7/32 1/4 7/32 7/32 1/4 1,400 1/4 1/4 9/32 101. SLING-LEG ANGLES OF CABLE-SUPPORTED 1,600 1/4 9/32 5/16 SUINGS. The optimum sling-leg angle (measured 1,800 &/16 5/16 5/16 from the horizontal) is 45 to 60 degrees. Mini- 2,000 5/16 11/32 3/8 mum tension in a sling leg occurs with a sling-leg angle of 90 degrees, and the tension approaches FIGURE 7.2.-Cable Load Table. infinity as the angle approaches zero. 103. SLING INSTALLATION. Attach the cargo sling larger sling-leg angles are desirable from a stand- to landing gear members or other structure cappoint of cable strength requirements. Slings able of supporting the loads to be carried. Install 50 Chap 7 ### Page 57 (OCR) Rev. 1977 the quick-release device in a level attitude with the throat opening facing the direction as indicated on the quick-release device. When cables are used to support the quick-release device, make sure the cables are not twisted or allowed to twist in the direction to unlay the cable. Some cargo release devices are provided with a fitting to permit installation of a guideline to assist in fully automatic engagement of the load target ring or load bridle. Secure the guideline to the quick-release device with a shear pin of a definite known value which will shear if a load becomes entangled on or over the guideline. Provision should also be made for cable-supported slings to be drawn up against the fuselage into a stowage position to prevent striking or dragging the release on the ground when not in use. AC 43.13-2A 104. INSTALLATION OF RELEASE CONTROLS. See figure 7.3 for typical wiring diagram of the electrical controls. a. Install a cargo release master switch, readily accessible to the pilot, to provide a means of deactivating the release circuit. The power for the electrical release circuit should originate at the primary bus. The "auto" position of the release master switch on some cargo release units provides for automatic release when the load contacts the ground and the load on the release is reduced to a preset value. b. Install the cargo release switch on one of the pilot's primary controls. It is usually installed on the cyclic stick to allow the pilot to release the load with minimum distraction after maneuvering the load into the release position. **RELEASE SWITCH (ON CYCLIC CONTROL)** RELEASE MASTER SWITCH Automatic OFT OPEN **RELEASE** **INDICATOR LIGHT** **CIRCUTT** **BREAKER** **QUICK RELEASE DEVICE** **PRIMARY BUS** **AUTO-RELEASE** RELEASE MICROSWITCH SOLENOID (Closes when load 18 reduced to preset value - usually 80-120 1b.) **SAFETY** MICROSWITCH (Release closed) FIGURE 7.3.-Typical cargo sling wiring diagram. Chap 7 51 ### Page 58 (OCR) AC 43.13-2A c. Install the emergency manual release control in a suitable position that is readily accessible to the pilot or other crewmember. Allow sufficient slack in the control cable to permit complete cargo movement without tripping the cargo release. d. The manual ground release handle, a feature of some cargo release units, permits opening of the cargo release by ground personnel. e. Label or placard all release controls as to function and operation. 105. FUNCTIONAL TEST. Test the release action of each release control of the quick-release device with various loads up to and including the maximum external load. This may be done in a test fixture or while installed on the rotorcraft, if the necessary load can be applied. If the quick-release device incorporates an automatic release, the unit should not release the load when the master switch is placed in the iautomatic" position until the load on the device is reduced to the preset value, usually 80 to 120 pounds. FIGURE 7.4.--Typical cargo sling installation No. 1. Rev. 1977 106. SUPPLEMENTAL FLIGHT INFORMATION. The aircraft may not be used in Part 133 externalload operations until a Rotorcraft-Load Combination Flight Manual is prepared in accordance with section 133.47 of that Part. FIGURE 7.5.- -Typical cargo sling installation No. 1 (showing fuselage attachment fitting). FIGURE 7.6-Typical cargo sling installation No. 1 (showing fore and aft limiting stops). FIGURE 7.7.-Typical cargo sling installation No. 2 (cargo hook attached directly to underside of fuselage). Chap 7 ### Page 59 (OCR) Rev. 1977 AC 43.13-2A FIGURE 7.8.-Typical cargo sling installation No. 3 (4-leg, cable suspended). 20. FIGURE 7.9.-Typical cargo sling installation No. 3 (showing cable sling leg attachment to landing gear crosstube fitting). FIGURE 7.10.- Typical cargo sling Installation No. 8 (showing cargo sling in stowed position). 107.-110. [RESERVED] Chap 7 ### Page 60 (OCR) AC 43.13-2A Rev. 1977 Section 2. CARGO RACKS 111. GENERAL. This section contains structural and design information for the fabrication and installation of a cargo rack used as an externalload attaching means for a Class A rotorcraftload combination operation under FAR Part 133. 112. FABRICATION OF CARGO RACKS. The type of construction and method of attachment depend upon the material to be used and the configuration of the rotorcraft involved. Illustrations of typical construction and installation methods are shown in figures 7.11-7.15. 113. STATIC TEST. The cargo rack installation must be able to withstand the static test load required by FAR 133.43(a). Conduct the test as outlined in chapter 1 of this handbook. 114. SUPPLEMENTAL FLIGHT INFORMATION. The aircraft may not be used in Part 133 externalload operations until a rotorcraft-load combination flight manual is prepared in accordance with section 133.47 of that Part. FIGURE 7.12.-Typical cargo rack installation No. 1 (showing attachment detail). FIoURE 7.11.-Typical cargo rack installation No. 1. 54 Chap 7 ### Page 61 (OCR) Rev. 1977 AC 43.13-2A FIGURE 7.13.-Typical cargo rack installation No. 2. FIGURE 7.15.-Typical cargo rack installation No. 3. FIGURE_7.14.-Typical cargo rack installation No. 2 (showing rack partially installed). 115.-120. [RESERVED] Chap 7 ### Page 62 (OCR) 6/1/79 AC 43.13-2A, CHG. 1 Chapter 8. GLIDER AND BANNER TOW-HITCH INSTALLATIONS Section 1. TOWPLANE CONSIDERATIONS [Deleted] Change 1 Chap 8 57 ### Page 63 (OCR) AC 43.13-2A, CHG 1 6/1/79 Section 2. TOW-HITCH INSTALLATIONS 126. STRUCTURAL REQUIREMENTS. The struc- * a. Glider tow hitches. Protection for the tural integrity of a tow-hitch installation on an towplane is provided by requiring use of a towline aircraft is dependent upon its intended usage. assembly which will break prior to structural Hitches which meet the glider tow criteria of this damage occurring to the towplane. The normal chapter are acceptable for banner tow usage. tow load of a glider rarely exceeds 80 percent of However, because the direction and magnitude of the weight of the glider. Therefore, the towline maximum dynamic banner towline loads occur assembly design load for a 1,000-pound glider within a more limited rearward cone of displace- could be estimated at 800 pounds. By multiplying ment than do glider towline loads, hitches which the estimated design load by 1.5 (to provide a safety meet the banner tow criteria of this chapter may margin), we arrive at a limit load value of 1,200 not be satisfactory for glider towing. Due to the pounds. The 1,200-pound limit load value is used in basic aerodynamic difference between the two static testing or analysis procedures per paragraph objects being towed, glider and banner tow-hitch 127 of this handbook to prove the strength of installations are treated separately with regard to the tow hook installation. When the hook and loading angles. structure have been proven to withstand* 20° Cone of Displacement FIGURE 8.1.-Glider tow angle. SAFETY PAYS 10° Cone of Displacement * FIGURE 8.2. - Banner tow angle. Chap 8 ### Page 64 (OCR) 6/1/79 *the limit load, then the maximum breaking strength of the towline assembly is established at the design load of 800 pounds. Thus, the towline will break well before structural damage will occur to the towplane. Another approach can be applied if the limit load carrying capabilities of a tow hook and fuselage are known. In this case, the known load value can be divided by 1.5 to arrive at the design load capabilities if the tow hook and fuselage limit loads are known to be 1,800 pounds. By dividing by 1.5 (1800 ÷ 1.5 = 1,200) we arrive at a design load value of 1,200 pounds. Thus, the maximum breaking strength of the towline assembly is established at 1,200 pounds and provides protection for the towplane. Thus, in considering tow hook installations, one may establish maximum towline breaking strength (1) Dividing the known limit load capabilities of the fuselage and two hook installation by 1.5; or (2) Knowing the design load needs of the towline assembly and multiplying by 1.5 to arrive at a limit load. Then by analysis or static testing, determine that the hook and fuselage are capable of withstanding that limit load. support a limit load equal to at least two times the operating weight of the banner. 127. STRUCTURAL TESTING. Adequacy of the aircraft structure to withstand the required loads can be determined by either static test or structural analysis. a. Static testing.When using static tests to verify structural strength, subject the tow hitch to the limit load (per paragraph 126 a or b) in a rearward direction within the appropriate cone of displacement per figure 8.2. Testing to be done in accordance with the procedures of Chapter 1, paragraph 3, of this handbook. b. Structural analysis. If the local fuselage structure is not substantiated by static test for the proposed tow load, using a method that experience has shown to be reliable, subject the fuselage to engineering analysis to determine that the local structure is adequate. Use fitting factor of 1.15 or greater in the loads for this analysis. 128. ATTACHMENT POINTS. Tow-hitch mechanisms are characteristically attached to, or at, tiedown points or tailwheel brackets on the air- Chap 8 AC 43.13-2A, CHG 1 frame where the inherent load-bearing qualities can be adapted to towing loads. Keep the length of the hitch-assembly arm from the airframe attachment point to the tow hook to a minimum as the loads on the attachment bolts are multiplied by increases in the moment arm. 129. ANGLES OF TOW. Tests should be conducted on the system at various tow angles to insure that: a. There is no interference with the tailwheel or adjacent structure. b. The towline clears all fixed and movable surfaces at the maximum cone of displacement and full surface travel. c. The mechanism does not significantly decrease the clearance from the tailwheel to the d. The tow hitch does not swivel. Experience has shown swiveling could result in fouling both the release line and towline during operations by the towplane. e. The opened jaw of the hitch does not strike any portion of the aircraft. 130. PLACARDS. A placard should be installed in a conspicuous place in the cockpit to notify the pilot of the structural design limits of the tow system. The following are examples of placards to be inbreaking strength not to exceed a. For glider tow-"Glider towline assembly pounds." b. For banner tow-Tow hitch limited to maximum weight of pounds." * Value established per paragraph 126 a (1) pr (2). ** Banner hitch limitations are one-half the load applied per paragraph 126 b. 131. WEIGHT AND BALANCE. In most cases, the weight of the tow-hitch assembly will affect the fully loaded aft c.g. location. To assure that the possibility of an adverse effect caused by the installation has not been ignored, enter all pertinent computations in the aircraft weight & balance records. (In accordance with the provisions contained in FAR 43.5(a) (4).) **132. TOW RELEASE MECHANISM.** a. Release lever. A placard indicating the direction of operation should be installed to allay the possibility of confusion or inadvertant opera-* 59 ### Page 65 (OCR) AC 43.13-2A, CHG 1 6/1/79 * tion, and the design of the release lever should provide the following: (1) Convenience in operation. (2) Smooth and positive release operation. (3) Positioned so as to permit the pilot to exert a straight pull on the release handle. (4) Sufficient handle travel to allow for normal slack and stretch of the release cable. (5) A sufficient handle/lever ration to assure adequate release force when the towline is under high loads. (See fig. 8.3) (6) Protection of cables from hazards such as: (a) Wear and abrasion during normal operation. (b) Binding where cables pass through fairleads, pulleys, etc. (c) Accidental release. (d) Interference by other aircraft com- (e) Freezing and moisture accumulation when fixed or flexible tubing guides are used. b. Test of the release. A test of the release and hook for proper operation through all angles of critical loading should be made using the design load for the glider or banner. c. Release cable. Representative size strength characteristics of steel release cable are as shown in figure 8.4; however, it is recommended that all internally installed release cables be ½6-inch or larger. Nonflexible Flexible Carbon Steel Carbon Steel 7 x 7 and 7 x 19 1 x 7 and 1 x 19 (MIL-W-1511A and Diameter (MIL-W-6904B) MIL-C-5424A) inches Breaking strength (Los.) Pounds 100 f. Breaking strength (lbs.) Pounds 100 p. 1/32 185 .20 3/64 375 .55 1/16 500 .85 480 .75 5/64 800 1.40 - 3/32 1,200 2.00 920 1.60 FIGURE 8.3-Typical tow-hitch release handle. FIGURE 8.4.- Representative steel cable qualities. 133. [Deleted] Change 1. FIGURE 8.5.- [Deleted] Change 1. * 60 Chap 8 U.S. GOVERNMENT PRINTING OFFICE: 1979 O--302-515 ### Page 66 (OCR) (DELETED) -- Change 2 Chap. 8 61 ### Page 67 (OCR) AC 43.13-2A Rev. 1977 FIGURE 8.6.—Tricycle gear aircraft. FIGURE 8.7.-Conventional gear aircraft- leaf spring type tailwheel. Chap. 8 ### Page 68 (OCR) Rev. 1977 AC 43.13-2A FIGURE 8.8.- Conventional gear aircraft- -shock strut type tailwheel. FIGURE 8.9.-Conventional gear aircraft-tubular spring type tailwheel. 134.-145. [RESERVED] Chap. 8 63 ### Page 69 (OCR) Rev. 1977 AC 43.13-2A Chapter 9. SHOULDER HARNESS INSTALLATIONS Section 1. RESTRAINT SYSTEMS 146. GENERAL. primary objective in shoulder harness design is to prevent incapacitating and/or fatal injuries to personnel involved in a survivable crash condition in which the aircraft cabin structure remained reasonably intact. Any harness configuration which achieves this objective is satisfactory from a safety viewpoint, regardless of the type of harness and mounting position used. Basic requirements of the aircraft airworthiness rules are designed to provide an aircraft structure to give each occupant a reasonable chance of escaping serious injury in a crash land- These requirements adequately provide for conditions that can be expected to occur in various types of survivable accidents. The human body has the inherent capability of withstanding decelerations of 20g's for time periods of up to 200 milliseconds (.2 second) without injury. Experience with aircraft used in agricultural and military operations shows that even in such unusual operations a high rate of survival in crashes is achieved when a restraint system is designed on the order of 20g to 25g deceleration loads. In view of the foregoing, persons installing a shoulder harness may wish to use a restraint system designed to withstand 20g to 25g loads. In addition, seat belts and seat belt anchorages designed to these load limits may be used. 147. TYPES OF RESTRAINT SYSTEMS. There are two generalized types of shoulder harnesses currently in use. They are the single diagonal type harness and the double over-the-shoulder type The over-the-shoulder harness may utilize either two independent attach points, or join in a "Y" configuration and attach at a single point. (See figs. 9.1 and 9.2) In all cases, however, the original safety belt or a combination harness utilizing a lap belt must be used in the installation. 148. ADVANTAGES OF DIFFERENT HARNESS TYPES. The single diagonal chest strap in combination with a lap belt is the simplest harness system and works effectively for longitudinal decelerations. However, during side decelerations, an occupant in this type harness has a tendency to slip out and away from the chest strap even when it fits snugly. The double over-the-shoulder type harness works well for both longitudinal and side decelerations. 149. MOUNTING CONFIGURATIONS. The type of shoulder restraint configuration acceptable for installation is dependent upon the attachments available in each individual aircraft. Basic harness mounting configurations are: a. Seat mounted. b. Airframe mounted. (1) Side (2) Ceiling (3) Floor (4) Directly rearward 150. STANDARDS. At the present time, there is a lack of standards for materials acceptable for use in shoulder harnesses. Until such time as a TSO is developed for shoulder harnesses, standards established in TSO-C22f pertaining to the materials and testing of safety belts may be accepted for this purpose. **151. MATERIALS.** a. Webbing. Synthetic materials, such as nylon and dacron, may be used for shoulder harness webbing. It is recommended that the webbing of the shoulder harness be the same as that of the lap belt to avoid problems in cleaning, staggered replacement of harness components due to wear or age, etc. Chap 9 65 ### Page 70 (OCR) AC 43.13-2A b. Fittings. Use hardware that: (1) Conforms to an acceptable standard such as AN, NAS, TSO, or MIL-SPEC. (2) Meets the strength required by FAR 23.1413 or 25.1413, as appropriate. (3) Will not loosen in service due to vibration or rotational loads. C. Inertia Reels. The function of the inertia reel is to lock and restrain the occupant in a crash yet provide the ability for normal movement without restrictions. In addition, automatic rewinding of any slack assures that the harness is always snug, which results in a more comfortable restraint system. Self-contained inertia reel units may be mounted at readily accessible locations and will generally operate effectively in any attitude. Their use in a body restraint system is satisfactory if mounted in accordance with acceptable methods, techniques, and practices, and will meet FIGURE 9.1.-Single diagonal type harness. 66 Rev. 1977 static strength requirements equal to those outlined in FAR 23.1413 Check the reel itself for the following operational hazards: (1) Inadvertent Lockup. If the inertia mechanism is set at a low "g" setting, unwanted lockup, or binding, of the system may occur. reel lockup range between .9 and 2.5 "g" is acceptable. (2) Improper Webbing Length. Install adequate webbing on the takeup reel to allow the occupant to reach all necessary switches and controls in the cockpit. Any additional webbing will result in decreasing the reel spring "takeup" tension exerted on the shoulder. (3) Incorrect Belt Opening Alignment. Position the reel so that the belt opening is aligned in the direction of loading. This will prevent the belt from rubbing and fraying due to normal usage. FIGURE 9.2.-Double over-the-shoulder type harness. 152.-155. [RESERVED] Chap 9 ### Page 71 (OCR) Rev. 1977 AC 43.13-2A Section 2. EFFECTIVE RESTRAINT ANGLES 156. RELATIONSHIP OF THE HARNESS ASSEMBLY TO THE OCCUPANT. Most restraint systems are designed so that each belt section maintains a certain relationship to the body. The attachment end of a restraint belt must maintain a relative angle and spacing to the head and neck surfaces as it passes over the shoulder and away from the body. This angle must provide sufficient freedom to assure normal body movements of the seated occupant without neck contact or interfering with vision. 157. ATTACHMENT AREAS FOR SHOULDER HAR- NESS. Effective attachment areas for the various types of shoulder harnesses are defined as angles formed by the attachment ends. Assure that when installing a harness for one seat that in a crash, a passenger to the rear would not sustain head impact injuries on the harness or its attach point. a. Single Diagonal Type Harness. The optimum rearward attachment area for this type of harness is within an angle of 30 degrees above the horizontal measured from the midpoint on the occupant's shoulder as shown in figures 9.3 and 9.4. Belt attachments should be located to the rear and outboard of the shoulder. This mounting area is shown in figure 9.5. (I) Attachment points inboard of this area would permit the harness to impinge on the neck and could result in neck injury during crash impact. In addition, the constant rubbing of the strap on the neck would be uncomfortable and, as a result, act as a distraction to the safe operation of the aircraft. Attachment points forward of this area would reduce the effectiveness of the harness, due to a lack of contact between the harness and the upper torso of the occupant. In addition, a shoulder strap attached forward of FIGURE 9.3.-Side mounted-single diagonal type harness. Chap 9 ### Page 72 (OCR) AC 43.13-2A Rev. 1977 30 FIGURE 9.4.- Ceiling mounted inertia reel-single diagonal type harness. the shoulder midpoint could obstruct vision and create a potential safety hazard. (2) The harness should be kept snug as any decrease in the distance between the occupant's head and the forward cabin structure increases the opportunities for head impact injuries. Also, the chances for twisting out of the harness are significantly increased. b. Double Over-the-Shoulder Type Harness. If this type of harness is intended to be mounted either directly rearward or to the ceiling, mount it within the 30-degree vertical angle as shown in figure 9.6. Because of the limited number of rearward shoulder harness attachment points in many aircraft, a 5-degree angle below the horizontal is also considered acceptable. 68 Shoulder harness attach areas as viewed from above are shown in figure 9.7. These mounting areas may be used for either the independent or the "Y" type belts. The outboard limit is established to prevent the belt section from slipping off the shoulder, and the maximum inboard angle is limited to a point which will prevent impingement on the neck surface. 158. AREA AND ANGLE DEVIATIONS. While the areas and angles given in the above paragraphs are intended to assist in the selection of attachment points, they should be considered as the optimum and not be interpreted as being mandatory. Area and/or angle deviations could result in a decrease in the overall efficiency of the restraint system; however, they may be necessary in order to permit a harness installation in an aircraft which otherwise could not be accom- Chap 9 ### Page 73 (OCR) Rev. 1977 plished. It is probable that other compromises may be necessary when adapting a specific restraint system to an aircraft in order to fit a body AC 43.13-2A of average dimensions. These compromises, however, should be permitted only when they are compatible with proper restraint functions. FIGURE 9.5.-Acceptable mounting area-single diagonal type harness. 30° UP HORIZ. REF 5° DOWN FIGURE 9.6.--Ceiling mounted inertia reel-double over-the-shoulder type harness. Chap 9 69 ### Page 74 (OCR) Rev. 1977 AC 43.13-2A FIGURE 9.7.-Acceptable mounting areas-double over-the-shoulder type harness. 159.-160. [RESERVED] 70 Chap 9 ### Page 75 (OCR) Rev. 1977 AC 43.13-2A Section 3. ATTACHMENT METHODS 161. STRUCTURAL ATTACHMENTS. For best results, the restraint system should be anchored to the primary aircraft structure. Design the structural attachment to spread the suddenly applied impact loads over as large an area of the structure as possible. The shoulder harness may be attached to selected secondary members which will deform slowly or collapse at a limited rate. This will assist in dissipating the high impact #g" loads to a level tolerable to the human body. However, the possibility of secondary members collapsing and making it difficult for an occupant to extract himself from the harness should not be overlooked. 162. FLOOR AND SEAT ATTACHMENTS. The double over-the-shoulder type harness shown in figure 9.8 may be used with either floor or seat mounting points, and typical installation methods are illustrated in figures 9.9 and 9.10. Two prerequisites necessary to ensure an effective restraint system are: a. The seat structure and its anchorages should be capable of withstanding the additional "g" loads imposed by the restrained occupant during an abrupt deceleration. This capability may be determined by static testing in accordance with FAR 23.785, 25.785, 27.785, or 29.785, as appropriate; or, by securing a statement attesting this adequacy from the airframe manufacturer's engineering department. b. The level of the seat back should at least be equal to the shoulder height of the seated occupant. This will reduce the inherent downward impact loads which would otherwise impart compressive forces. on the occupant's torso. FIGURE 9.8.-Floor mounted inertia reel-double over-the-shoulder type harness. Chap 9 251-957 0 - 77 - 4 ### Page 76 (OCR) Rev. 1977 AC 43.13-2A FIGURE 9.9.- -Typical floor mounted inertia reel installation. FIGURE 9.10.-Typical seat mounted inertia reel installation. Seats which utilize a fold-over type of seat back must have some type of locking mechanism so that the seat can support the loads without allowing the occupant to move forward. The lock should be of a type which has a quick release 72 to allow rear seat occupants to rapidly evacuate the aircraft. This type of installation should only be considered if other means of attachment are not available since making a folding seat rigid greatly reduces the protection afforded a passenger to the rear. When a folding seat is provided with a lock, the passenger to the rear should also be provided a shoulder restraint system, or, the back of the forward seat should be provided with sufficient absorptive material to adequately compensate for the added rigidity. Also, the change in load distribution due to the loads being applied to the seat back may require reinforcing the seat and/or belt anchorages to meet airworthiness requirements. 163. AIRFRAME ATTACHMENTS. The method used for the attachment of shoulder harness anchorages is dependent upon the construction features of the aircraft involved. a. Monocoque/Semimonocoque Type Constructions. Illustrations of typical aircraft members and installation methods are shown in figures 9.11 through 9.15. b. Tube Type Construction. Various typical methods of attaching shoulder harness anchorages are shown in figure 9.16. When aircraft cable is used as a component in a shoulder harness anchorages, swage the cable terminals in accordance with the procedures contained in chapter 4 of AC 43.13-14, Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repairs." 164. STRUCTURAL REPAIR KITS. In many instances, structural repair kits are available from the aircraft manufacturer. While these kits are primarily intended for use in repairing defective or damaged structure, they may also be used as a reinforcement for shoulder harness attach fittings. 165. FLEXIBLE ATTACHMENTS. Various aircraft are designed so that fuselage members and/or skin will flex or "work." This type of structure should not be heavily reinforced for the purpose of attaching shoulder harnesses as this would defeat the design purpose. In cases such as these, use a localized reinforcement, such as shown in figure 9.15, at the attachment point. This will allow the fuselage to flex while still maintaining a collapsible structure to absorb the loads encountered in a crash. Chap 9 ### Page 77 (OCR) kev. 1977 AC 43.13-2A FIGURE 9.11.-Typical bulkhead reinforcement installation. FIGURE 9.12.-Typical wing carry-through member installation. Chap 9 73 ### Page 78 (OCR) AC 43.13-2A FIGURE 9.13.-Typical hat section reinforcement installation. FIGURE 9.14.-Typical stringer section reinforcement installation. 166. SIMPLIFIED INSTALLATION CRITERIA. To encourage the addition of shoulder restraints to existing aircraft with a minimum of testing and engineering necessary, yet provide a satisfactory restraint, the following general conditions will be acceptable. a. Utilize the original seat belt attachments and either the original or a new belt provided with shoulder harness fittings. b. Use webbing approved for standard seat belts (TSO-C22f). c. Use hardware approved for use on seat belts per TSO-C22f. d. Secure the lower end of the shoulder strap to one side of the original seat belt or belt anchorage. e. Use a mount for the shoulder restraint in- 74 Rev. 1977 dependent of the seat such as the aft or ceiling mounts per paragraph 128. f. Test the added mount by applying a test load of at least 500 pounds forward at the shoulder point. 167. ENGINEERING APPROVAL. Installations which involve cutting or drilling of critical fuselage members or skin of pressurized aircraft will usually require engineering evaluation. For this reason it may be desirable to contact the airframe manufacturer to obtain his recommendations. FIGURE 9.15.-Typical stringer section reinforcement installation. Chap 9 ### Page 79 (OCR) Rev. 1977 АС 43.13-2A FIGuRE 9.16.-Typical attachment to tubular members (adequate chafing protection for tube should be provided. Detail omitted for clarity.) 168.-175. [RESERVED] Chap 9 75 ### Page 80 (OCR) Rev. 1977 AC 43.13-2A Chapter 10. BATTERY INSTALLATIONS Section 1. GENERAL 176. GENERAL. This section contains structural and design considerations for the fabrication of aircraft battery installations. 177. LOCATION REQUIREMENTS. The battery location and/or its installation should provide: a. Accessibility for Battery Maintenance and Removal. The electrolyte level of the battery needs frequent checking; therefore, install the battery so that it is readily accessible for this service without the removal of cowling, seats, fairings, etc. Inaccessibility is often the source of neglect of this important piece of equipment. Certain types of batteries cannot be conveniently serviced while installed. Therefore, install and/ or locate the battery so that it can be readily removed and reinstalled. b. Protection from Engine Heat. The instaliation should protect the battery from extreme engine heat, which would be detrimental to the battery's service life and reliability. Such protection should provide for the temperatures encountered after engine shutdown as well as during engine operation. When locating the battery within the engine compartment, choose a location that will not interfere with the flow of engine-cooling air. c. Protection from Mechanical Damage. Vibration and other shock loads are a major cause of short battery life. Whenever possible, install the battery in a manner or location that will minimize damage from airframe vibration and prevent accidental damage by passengers or d. Passenger Protection. Enclose the battery in a box or other suitable structure to protect FIGURE 10.2.-Typical battery box installation in aft fuselage area, below cabin floorboards, or may also FIGURE 10.1.-Typical battery box installation in be adapted for within wing locations (shaded poraft fuselage area. tions indicate original structure). Chap 10 77 ### Page 81 (OCR) AC 43.13-2A passengers from any fumes or electrolyte that may be spilled as a result of battery overheating, minor crash, inverted flight, and/or rapid decompression if the battery is located within the aircraft'ș pressure vessel. e. Airframe Protection. Protect the airframe structure and fluid lines by applying asphalticor rubber-base paint to the areas adjacent to and below the battery or battery box. Apply paralketone, heavy grease, or other comparable protective coating to control cables in the vicinity of the battery or battery box. Damage to adjacent fabric covering and electrical equipment can be minimized by providing a battery sump jar containing a neutralizing agent, properly locating battery drains and vent discharge lines, and adequately venting the battery compartment. 178. DUPLICATION OF THE MANUFACTURER'S INSTALLATION. The availability of readymade parts and attachment fittings may make it desirable to consider the location and type of installation selected and designed by the airframe manufacturer. Appreciable savings in time and FIGURE 10.3.-Typical battery box installation in aft fuselage area (shaded portions indicate original structure). 78 Rev. 1977 work may be realized if previously approved data and/or parts are used. 179. OTHER INSTALLATIONS. If the battery installation has not been previously approved, or if the battery is to be installed or relocated in a manner or location other than provided in previously approved data, perform static tests on the completed installation as outlined in chapter 1 of this handbook. Because of the concentrated mass of the battery, the support structure should also be rigid enough to prevent undue vibration which would lead to early structural failure. Typical illustrations of battery support structure are shown in figures 10.1 thru FIGURE 10.4.Typical battery box installation suspended from cabin floorboard section. 180.-185. [RESERVED] Chap 10 ### Page 82 (OCR) Rev. 1977 AC 43.13-2A Section 2. INSTALLATION 186. SECURING THE BATTERY. Install the battery box or holddowns in such a manner as to hold the battery securely in place without subjecting it to excessive localized pressure which may distort or crack the battery case. Use rubber or wooden blocks protected with parafin or asphaltic paint as spacers within the battery box, as necessary, to prevent shifting of the battery and possible shorting of the battery terminals or cables. Also, provide adequate clearance between the battery and any bolts and/or rivets which may protrude into the battery box or compartment. 187. VENTING. Provide suitable venting to the battery compartment to prevent the accumulation of the hydrogen gas evolved during operation. For most aircraft batteries, an airfiow of 5 cu. ft. per minute is sufficient to purge the battery compartment of explosive concentrations of hydrogen. a. Manifold Type. In this type of venting, one or more batteries are connected, via battery or battery box vent nipples, to a hose or tube manifold system as shown in figure 10.5. Fasten such hoses securely to prevent shifting and maintain adequate bend radii to prevent kinking. (1) The upstream side of the system is connected to a positive pressure point on the aircraft, and the downstream side is usually discharged overboard to a negative pressure area. It is advisable to install a battery sump jar in the downstream side to neutralize any corrosive vapors that may be discharged. (2) When selecting these pressure points, select points that will always provide the proper direction of airflow through the manifold system during all normal operating attitudes. Reversals of flow within the vent system should not be permitted when a battery sump jar is installed, as the neutralizing agent in the jar may contaminate the electrolyte within the battery. b. Free Airflow Type. Battery cases or boxes that contain louvers may be installed without an individual vent system, provided: (1) The compartment in which the battery is installed has sufficient airflow to prevent the accumulation of explosive mixtures of hydrogen; (2) Noxious fumes are directed away from occupants; and (3) Suitable precautions are taken to prevent corrosive battery fluids or vapors from damaging surrounding structure, covering, equipment, control cables, wiring, etc. 188. DRAINS. Position battery compartment drains so that they do not allow spillage to come in contact with the aircraft during either ground or flight attitudes. Route the drains so that they have a positive slope without traps. Drains should be at least ½" in diameter to prevent clogging. **VENT NIPPLE** MANIFOLD VENT TUBE FIGURE 10.5.-Battery with integral vent nipples. Chap 10 79 ### Page 83 (OCR) AC 43.13-2A **189. ELECTRICAL INSTALLATION.** a. Cables/Connectors. Use cables and/or connectors that are adequately rated for the current demand and are properly installed (See AC 43.13-1A, Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repair," chapter 11). It may be necessary to contact the battery manufacturer to determine current value of the battery at the 5-minute discharge rate. Cable size can also be selected by using the same gage as used on a previously approved production aircraft with the same battery. (1) The cables should be of sufficient length to prevent undue strain on the battery connector or terminals. (2) Clamp and protect cables, including the bus, in a very secure manner. Since these units are not fused, any fault could cause loss of the entire electrical system in addition to a possible fire hazard. **BATTERY RELAY** IO PRIMARY BUS Rev. 1977 (3) Route cables so that cable or terminals cannot short to the battery case or hold-down frame. (4) Route cables outside the battery box whenever practicable to prevent corrosion by acid fumes. When internal routing is unavoidable, protect the cable inside the box with acidproof tubing. Assure that cables will not be inadvertently reversed on the battery terminals either by proper cable lengths and clamps or, if this is not practicable, use conspicuous color coding. b. Battery Cutoff. Install a battery cutoff relay to provide a means of isolating the battery from the aircraft's electrical system. An acceptable battery cutoff circuit is shown in figure 10.6. Mount the relay so that the cable connecting the relay to the battery is as short as feasible, in any case not to exceed two feet, to reduce the possibility of a fire occurring because of a short within this section of cable. **GENERATOR / ALTERNATOR** **CONTROL** **CIRCUIT** BATTERY FIGURE 10.6. -Typical battery cutoff and generator alternator control circuit. 190.-195. [RESERVED] 80 Chap 10 ### Page 84 (OCR) Rev. 1977 AC 43.13-24 Section 3. REPLACEMENT OF LEAD-ACID BATTERIES WITH NICKEL-CADMIUM BATTERIES bility, and a high degree of reliability. The teries requires careful evaluation of certain areas. fill a need for a power source that will provide large amounts of current, fast recharge capa- exchange of lead-acid for nickel-cadmium bat- 196. GENERAL. Nickel-cadmium batteries ful- the aircraft, it is important that the battery be located where it can easily be inspected, removed, and installed. Some battery cases are designed with viewports on each side of the case for visual monitoring of the cell electrolyte level. If cadmium batteries are generally not serviced in 197. ELECTRICAL ANALYSIS. The ampere hour this type of case is to be utilized, consideration capacity of a nickel-cadmium battery is selected should be given to the location of the battery to accommodate the aircraft load requirements. compartment to accommodate this feature. When making this selection, the following items 199. STRUCTURAL REQUIREMENTS. Most lead-acid should be considered. battery compartments provide adequate structure a. The low internal resistance permits it to re- attachment for the installation of nickel-cadcharge very quickly. This high recharge rate mium batteries. However, cantiliver supported current. Total system load (battery recharging plus system loads) must not exceed the pre-estab- can exceed the generator rated capacity and de- prive essential circuits of necessary operating modified to compensate for an increased over- hang moment. This may be caused by a change for nickel-cadmium battery installations unless battery boxes/compartments may not be suitable lished electrical capacity. in battery shape and c.g. location even though the replacement battery weighs less than the following: of the batteries to make sure a reduced capacity nickel-cadmium battery is adequate regarding the b. Compare the discharge characteristics curves installation, perform a static test as outlined in original lead-acid battery. (See fig. 10.7.) When ever the total installation weight and/or the overhang moment exceed those of the original (1) Ability to provide engine starting or chapter 1 of the handbook. If the battery comcranking requirements. Turbine engines require partment is to be relocated, follow the procedures an initial surge of approximately 1200 amperes outlined in Sections 1 and 2 of this chapter. which tapers off within 10 seconds to approximately 800 amperes cranking current. Recipro- 200. ISOLATION OF BATTERY CASE. Because of cating engines require approximately 100 am- the material from which nickel-cadmium battery peres cranking current. cases are generally made (stainless or epoxy low temperature starting. Nickel-cadmium bat- teries deliver their rated capacity when the (2) Ability to provide sufficient capacity for coated steel), it is desirable to electrically isolate the case from aircraft structure. This will elimi- nate the potential discharge current produced ambient temperature range is 70° to 90° F. In- when spillage or spewage of the electrolyte procreased battery capacity will offset the effects of vides a current path between the cell terminal low-temperature starting. or connector and exposed metal of the battery case. This isolation is also desirable in that it 198. MAINTENANCE CONSIDERATIONS. To pro- vide for ease of inspection and because nickel- ing the generator output to the stucture. could prevent a fault within the battery or fault- Chap 10 81 ### Page 85 (OCR) AC 43.13-2A Rev. 1977 of nickel-cadmium battery cases, one with vent nozzles and the type with viewports. a. The vent nozzle type utilizes vent hoses to evacuate the gas overboard by use of forced air - DI - REPLACEMENT or by venturi effect. BATTERY b. Battery cases with viewports or louvers must OUTLINE have an air flow sufficient to keep the mixture of air and hydrogen below 4 percent. The gases from this type of case are evacuated into the battery compartment. Regardless of the ventilation system used, the air flow should be a minimum of 5 cubic feet per minute. **ORIGINAL** **LEAD-ACID** **BATTERY C.G. (WI)** **REPLACEMENT** NICKEL-CADMIUM BATTERY C.G. (Wz) FIGURE 10.7.-Change in cantilever support battery overhang moment. Example: Original Overhang Moment with Lead- Acid Battery W1 = 25 1b., D1 = 4 inches 4 inches X 25 1b., = 100 in. - 1b. New Overhang Moment with Replacement Nickel-Cadmium Battery Wz = 19.5 1b., Dz = 6 inches 6 inches X 19.5 1b. = 117 in. - 1b. 201. VENTILATION. During the charging process, nickel-cadmium batteries produce hydrogen and oxygen gases. This occurs near the end of the charging cycle when the battery reaches what is called the gassing potential. To avoid a buildup of these gases, and possible accidental ignition, ventilation should be provided to evacuate this gas from the aircraft. There are two types 202. PREINSTALLATION REQUIREMENTS. Inspect the replacement battery for possible damage incurred during shipment or storage. Give particular attention to signs of spilled liquid within the shipping container, as it may indicate a damaged cell. Follow procedures outlined in Section 2 for battery venting and electrical connec tions. a. Preinstallation battery servicing. Check at least the following in accordance with the battery manufacturer's instruction: (1) Remove the shipping plugs (if used) and clean and install the filler cap vent plugs. (2) Check the tightness of terminal hardware including each cell connector strap to the proper torque values. (3) Check the polarity of each cell to be sure they are connected in the proper series or sequence. (4) Charge and check battery voltage and electrolyte level. b. Compartments or battery boxes which have previously housed lead-acid batteries must be washed out, neutralized with ammonia or a baking soda solution, allowed to dry thoroughly, and painted with alkaline-resistant paint. Remove all traces of sulfuric acid and its corrosion products from the battery vent system to prevent contamination of the potassium hydroxide electrolyte and/or possible damage to the cell case material. Replace those parts of the vent system which cannot be thoroughly cleansed (hoses, etc.). When sump jars are incorporated into the vent system, replace the old pad with a new one that has been saturated in a three-percent solution (by weight) of boric acid and water. 82 Chap 10 ### Page 86 (OCR) Rev. 1977 203. SECURING THE BATTERY. Follow the procedures outlined in Section 2 of this chapter. Make certain that the holddown bolts are not drawn up so tightly that the battery case/cover becomes distorted. Should the cover become distorted, there is a possibility that the cell terminal hardware may eventually puncture the neoprene cover liner used in many batteries, and short Caution: In installations where care has been taken to isolate the battery cases, inadvertent grounding may occur through improper or careless use of safety wire. Use no wood in nickelcadmium battery boxes as it becomes conductive with time causing a current flow from the battery case to ground. Use only fiberglass or other acceptable material as liners and spacers in the battery box. 204. VOLTAGE AND CURRENT REGULATION. It is essential that the charging voltage and current be checked and, if necessary, the voltage regulator reset to meet the requirements of the nickelcadmium battery being installed. IMPORTANT -improper charging current or voltage can destroy a battery in a short period of time. Chap 10 AC 43.13-2A 205. WEIGHT AND BALANCE. After installation of the nickel-cadmium battery the weight and balance of the aircraft should be recomputed if: a. The weight of the nickel-cadmium battery is different from that of the original lead-acid battery. b. The location of the nickel-cadmium battery is different from that of the original lead-acid battery. Weight and balance procedures for aircraft are contained in chapter 13 of AC 43.13-1A. **206. RESTORATION OF LEAD-ACID BATTERIES.** When lead-acid batteries are restored in lieu of nickel-cadmium batteries the procedures contained in sections 1, 2, and 3 of this chapter should be used. Structural requirements are referenced in paragraph 199 and figure 10.7. Airframe protection is specified in paragraph 177. Follow the procedures outlined in section 2 of this chapter for battery security, battery venting, and battery drains. Assure that all electrical requirements have been accomplished. Place emphasis on aircraft weight and balance. Refer to chapter 13 of AC 43.13-1A. 207.-210. [RESERVED] 83 ### Page 87 (OCR) Rev. 1977 AC 43.13-2A Chapter 11. ADDING OR RELOCATING INSTRUMENTS 211. GENERAL. This chapter contains struc- niques, and Practices--Aircraft Inspection and tural and design information to be considered Repair," chapter 2, section 3, for methods and when aircraft alterations involving the addition techniques of retaining structural integrity. and relocation of instruments are being made. b. Plumbing. Refer to the manufacturer's in- 212. PREPARATION. First determine what regu- structions for fabrication, routing and installalation, (CAR 3, 4b, FAR 23, 25, etc.) is the tion of instrument system lines. Advisory Cirbasis for the aircraft type certificate. That regu- cular 43.13-1A provides information regarding lation establishes the structural and performance the installation and fabrication of aircraft plumbrequirements to be considered when instruments are to be added or relocated. c. Vacuum Source. Minimum requirements strument panel installation. vides information by which structural integrity may be determined. Chapter 2, paragraph 23a through f provides information pertinent to in- a. Structure. Chapter 1 of this handbook pro- ments: Installation." for installation and performance of instrument vacuum systems are covered in the applicable FAR Airworthiness Standards under "Instru- (1) In the absence of specific requirements b. Location. Consult the applicable regulation tor vacuum pump installation, refer to FAR Part 25, section 25.1433 for guidance. It is detion and arrangement. for the specific requirements for instrument loca- sirable to install a "T" fitting between the pump and relief valve to facilitate ground checking and (1) In the absence of specific requirements, adjustment of the system. installation of IFR flight instruments in a "T" (2) When a venturi tube power source is arrangement is recommended. Locate the air- used, it should not be taken for granted that a craft attitude indicator at top center, airspeed venturi will produce sufficient vacuum to propindicator to the left, altimeter to the right and erly operate one or more instruments. Many of directional indicator directly below, thus form- the venturi tubes available for aircraft have a flow ing the letter "T. When a radio altimeter is rate of approximately 2.3 cubic feet per minute used, the indicator may be placed on the im- at 3.75 inches of mercury (in. Hg) vacuum. mediate right of the attitude indicator with the Therefore, it is essential that the vacuum load pressure altimeter to the right of the radio altim- requirements be carefully evaluated. eter indicator. (3) Vacuum loads may be calculated as fol- 213. INSTALLATION. Mount all instruments so they are visible to the crewmember primarily (a) Gyroscopic instruments require optiresponsible for their use. Mount self-contained mum value of airflow to produce their rated rotor gyroscopic instruments so that the sensitive axis speed. For instance, a bank and pitch indicator is parallel to the aircraft longitudinal axis. requires approximately 2.30 cubic feet per minute 23a through f of this handbook. Refer to the Circular 43.13-1A, Acceptable Methods, Tech- aircraft manufacturer's instructions and Advisory such as adding holes in the instrument panel to mount instruments, refer to chapter 2, paragraph a. Structure. When making structural changes per minute and a pressure drop of 4.00 in. Hg. indicator consumes approximately 1.30 cubic feet would be marginal. Similarly, the directional gyro The turn and bank indicator has a flow require- of 4.00 in. Hg. Therefore, operating an instru- ment requiring 4.00 in. Hg from one venturi for its operation and a resistance or pressure drop Chap 11 85 ### Page 88 (OCR) Rev. 1977 AC 43.13-2A ment of 0.50 cubic feet per minute and reaches this flow at a pressure drop of 2.00 in. Hg. The Table 11.2.-Equivalent straight tube line drops for 90° elbows. above instruments are listed in Tables 11.1 and Pressure drop in a 90° elbowo 11.3. Optimum values are shown in Table 11.3. Tubing size in terms of length of straight It should be noted that the negative pressure tube equivalent to a 90° elbow air source must not only deliver the optimum O.D inch Wall thickness inch Feet value of vacuum to the instruments, but must ¼ .035 0.28 Table 11.1.-Instrument air consumption. .035 0.46 dir consumption at -042 0.62 sea level .042 0.81 Instrument Differential drop in. Hg Cubic ¾ .049 0.98 suction feet/per (Optimum) minute 1 .049 1.35 AUTOMATIC PILOT SYSTEM 3 Determine the pressure drop through Directional gyro control unit (Types A-2, A-3, & A-3A) across mount assembly 5.00 2.15* 22.48 feet (20 feet + 2.48 equivalent feet) of ½ O.D. X .042 tubing at 4.10 CFM flow. From fig- Bank & climb gyro control unit ure 11.1, pressure drop per each 10-foot length = across mount assembly 5.00 3.85* 0.68 in. Hg. Divide 22.48 feet of tubing by 10 to Total 6.00* obtain the number of 10-foot sections, i.e., AUTOMATIC PILOT SYSTEM Directional gyro control unit (Type A-4) 5.00 3.50* 22.48 ÷ 10=2.248. Multiply the number of sec- tions by 0.68 in. Hg to obtain the pressure drop Bank & climb gyro control unit 5.00 5.00* through the system. (0.68 X 2.248=1.53 in. Hg) Total 8.50* 4 The pump must therefore be capable Bank & pitch indicator Directional gyro indicator 4.00 4.00 2.30 1.30 of producing a minimum pressure differential of Turn & bank indicator 2.00 .50 Table 11.3.-Differential pressure across instrument (*) Nore.—Includes air required for operation of pneumatic inlet and outlet. relays. Suction in inches also have sufficient volume capacity to accom- of mercury modate the total flow requirements of the various Instrument Mini- Opti- Maxiinstruments which it serves. mum mum mum (b) To calculate the flow requirements of elbows and 20 feet of line (½ O.D. x .042) tub- a simple vacuum system, assume four right-angle AUTOMATIC PILOT SYSTEM (Types A-2, A-3, & A-3A) ing Directional gyro control unit across mount assembly 4.75 5.00 5.25 1 Assume the flow requirements for: Bank & climb gyro control unit Turn & bank indicator .50 CFM across mount assembly 4.75 5.00 5.25 Directional gyro indicator Bank & pitch indicator 1.30 CFM 2.30 CFM Gauge reading (differential gauge in B & C control unit) 3.75 4.00 4.25 AUTOMATIC PILOT SYSTEM Total flow required 4.10 CFM (Type A-4) Directional gyro control unit 3.75 5.00 5.00 O.D. x .042 elbow is equivalent to 0.62 feet of Bank & pitch indicator Bank & climb gyro control unit 3.75 3.75 5.00 4.00 5.00 5.00 Directional gyro indicator 3.75 4.00 5.00 sure drop of four 90° elbows is equivalent to 2.48 Turn & bank indicator 1.80 2.00 2.20 feet of tubing. Chap 11 86 ### Page 89 (OCR) Rev. 1977 AC 43.13-2A • 035 PRESSURE DROP PER TEN FOOT LENGTH - INCHES OF MERCURY 1/4 x 035 TUBING SIZE 5/16 x 3/8 × .035 • 042 1/2 x 5/8 x • 042 3/4 x • 049 0.15 1 x 0.1 0.5 0.75 1 1.5 2 3 4 5 6 8 10 15 20 **AIR FLOW - CUBIC FEET PER MINUTE** PRESSURE DROP DATA FOR SMOOTH TUBING FIGURE 11.1 Chap 11 ### Page 90 (OCR) AC 43.13-2A 5.53 in. Hg, i.e., 4.00 in. Hg for maximum instrument usage + 1.53 in. Hg (determined) at a flow of 4.10 cubic feet per minute. d. Filter. Filters are used to prevent dust, lint and other foreign matter from entering the instrument and vacuum system. Filters may be located at the instrument intake port or at the manifold intake port when instruments are interconnected. Determine that the capacity of the filter is as great or greater than the capacity of the vacuum system. Assure that there is no pressure drop across the filter media. e. Electrical Supply for Instruments. Guidelines for the installation of instrument electrical wiring and power source are provided in Advisory Circular 43.13-1A, chapter 11, sections 2 and 3, and Chapter 16, section 3. NorE: Strict conformance to the shielding specifications supplied by compass manufacturers is recommended in all installations to eliminate any possibility of spurious f. Instrument Lighting. Instrument lighting must be installed in accordance with the regulations that are applicable to the aircraft type cer- sine it yes or. reflection of the lights is unsatisfactory, provide a shield or a means for controlling the intensity of illumination. g. Magnetic Headings. Calibrate magnetic instruments with the powerplants operating. After this initial calibration, switch all nav/com and electrical equipment, such as windshield wipers and defrosters, "on" to determine if any electrical system interference affects the instrument calibration. If the calibration is affected, prepare an instrument placard identifying the compass headings with the equipment "on" and also with the equipment "off." Placard in accordance with par. 214f of this chapter. 88 Rov. 1977 **214. TESTING, MARKING, AND PLACARDING.** a. Testing the Venturi Tube-Powered Systems. At normal inflight cruise speed, check the venturi tube-powered system to assure that the required vacuum pressure is being supplied. b. Testing the Vacuum Air Pump Powered System. When the system is powered by vacuum air pumps, check the system while the pumps are operating at their rated r.p.m. and measure the vacum available to the instruments. c. Testing of Altimeters and Static Systems. When checking the operation of an altimeter static system to determine that the system is free of any contaminating materials, be sure to disconnect the plumbing from all air operated instruments before purging the lines with dry air or nitrogen since the pressure necessary for purgmay damage any connected instrument. Static system test procedures are provided in FAR 43, Appendix E. d. Testing electrical supply (instruments). Check the voltage at the instrument terminals and determine that it is equal to the manufacturer's recommended values. e. Fuel, Oil, and Hydraulic (Instrument Supply). Measure the fluid pressure at the instrument end of the line to determine whether it is equivalent to that at the pressure source. f. Instrument Markings and Placards. When additional instruments are installed they must be appropriately marked. Refer to the applicable CAR/FAR under "Markings and Placards" for specific instrument marking and placard requirements. 215.-240. [RESERVED] Chap 11 ### Page 91 (OCR) Rev. 1977 AC 43.13-2A Chapter 12. LITTER, BERTH, AND CARGO TIEDOWN DEVICE INSTALLATIONS 241. GENERAL. This chapter provides data for and Repair") Commercially available litters or down device installations in airplanes and rotor- making acceptable litter, berth, and cargo tie- installation and load criteria. stretchers may be utilized provided they meet craft. Prior to proceeding with the alteration, determine the airworthiness standards applicable aphragm, or other means at the forward end of Provide a padded end board, canvas dito the contemplated alteration. Refer to the each litter or berth that will withstand the total proper volume of the FAA publication "Type Certificate Data Sheets and Specification" for the litters and berths installed parallel to the air- forward static test load. Safety belts used for certification basis of the aircraft involved. When airworthiness standards pertinent to the airplane craft's longitudinal axis are not required to with- stand forward loads as these loads are to be involved are not available, current airworthiness taken up by the padded end boards. standards may be used. For example, FAR 23, Design litters or berths intended for adults for 25, 27, or 29, as applicable, may be used in place occupants weighing at least 170 Ibs. In those of CAR 4a, 4b, 6, or 7 or Bulletin 7a. instances where litters or berths are for children, 242. INSTALLATION CONSIDERATIONS. design the installation for the maximum intended weight. See paragraph 247 for placarding. erated within the weight and center of gravity a. Assure that the altered aircraft can be op- tendants or other passengers that will be carried Provide approved seats for any cabin atranges. Refer to chapter 1, paragraph 9 of this when litters are installed. handbook. b. Cargo tiedown devices may be assembled b. Determine that there will be free access to from webbing, nets, rope, cables, fittings, or other all equipment and controls essential to the proper material which conforms to NAS, TSO, AN, operation of the aircraft, required emergency MIL-SPEC, or other acceptable standards. Use exits, and emergency equipment. snaps, hooks, clamps, buckles, or other acceptable c. Use only materials that are at least flame- fasteners rather than relying upon knots for seresistant for covering of floors, litters, or berths. curing cargo. Install tensioning devices or other Refer to the applicable airworthiness standards means to provide a method of tightening and for the aircraft involved to determine the re- adjusting the restraint system to fit the cargoes quired flame-resistant qualities. For aircraft in to be carried. Provide covers or guards where necessary to prevent damage to or jamming of the aircraft's equipment, structure, or control cables. tion of cargo with respect to passengers, cargo 244. STRUCTURAL ATTACHMENT. Commercially compartment, or aisle width. available seat tracks, rails, or other types of 243. FABRICATION AND INSTALLATION. ment, provided they conform to an accepted anchor plates may be used for structural attachtubing, sheet metal, extrusions, canvas, webbing, a. Litters and berths may be fabricated from standard (see chapter 1, paragraph 6). This type of hardware permits a ready means of methods, techniques, and practices. visory Circular 43.13-1A "Acceptable Methods, or other suitable material, using acceptable (Ref. Ad- mounting a wide variety of quick-disconnect fittings for litters, berths, and cargo tiedown. Typical examples of such fittings and their at- Techniques, and Practices-Aircraft Inspection tachment are shown in figures 12.1 through 12.5. Chap 12 89 ### Page 92 (OCR) AC 43.13-2A Rov. 1977 EXTRUDED TRACK ANCHOR PLATE Light duty Medium duty Heavy duty ces Extruded track and anchor plates are available in several different styles and load capacities and will accommodate a wide variety of quick attachment fittings. **SINGLE PIN TYPE SINGLE PIN TYPE** HOLD DOWN FITTINGS CARGO TIE DOWN FITTINGS Adjustable These types of fittings are suitable for litter or berth attachment to the extruded track and anchor plate styles shown above. These types of fittings are suitable for cargo tie down attachment to the extruded track and anchor plate styles shown above. DUAL PIN TYPE CARGO TIE FITTINGS 360° Rotation Low profile n These types of cargo tie down fittings are of greater capacity than the single pin types and are suitable for use with the extruded track style shown above. FIGURE 12.1. Extruded track, anchor plates, and associated fittings. 90 Chap. 12 ### Page 93 (OCR) Rev. 1977 AC 43.13-2A PAN FITTINGS SINGLE STUD FITTINGS Stud & Cargo ring Cargo ring only Stud only Round head Hex head **SINGLE STUD CARGO TIE DOWN FITTING** SINGLE STUD HOLD DOWN FITTINGS Lift to release These types of fittings are suitable for cargo tie down attachment to the single stud fittings or stud equipped pan fittings shown above. STUD/RING These types of fittings are suitable for litter or berth attachment to the single stud fittings or stud equipped pan fittings shown above. EYE BOLT These types of fittings are suitable for litter, berth, and/or cargo tie down attachment directly to the aircraft structure. FIGURE 12.2. Miscellaneous litter, berth, and cargo tiedown fittings. Chap 12 91 ### Page 94 (OCR) AC 43.13-2A Rev. 1977 A. Attachment method utilizing a honeycomb doubler. Single Studs or Eye Bolts. Pan Fittings. B. Attachment methods utilizing reinforcing plates. Stud/Rings. 1. Bed all inserts and spacers in a suitable potting compound. 2. Reinforcing plate. 3. Where fitting is subject to rotation, place washers on both sides and use a positive safety means. 4. (Alternate method in lieu of spacers) Undercut honeycomb, inject potting compound, and drill through when set. 5. Countersink if required for clearance or if desired for appearance. 6. Undercut all open edges of honeycomb 1/16" and seal with potting compound. 7. Honeycomb doubler. 8. Use epoxy or other suitable adhesive to attach doubler and reinforcing plates. FIGURE 12.3. Typical attachment of fittings to honeycomb structures. 92 Chap 12 ### Page 95 (OCR) Rev. 1977 AC 43.13-2A EXTRUDED TRACK, · > FLOOR BEAM (Reinforce cut out as neccessary to retain strength.) **UNDERFLOOR REINFORCEMENT** FLOOR BEAM DOUBLER/ATTACH ANGLE FIGURE 12.4. Installation of underfloor support for extruded track. Chap 12 93 ### Page 96 (OCR) AC 43.13-2A Rev. 1977 THREE BAR TYPE SLIDE Rigging for easiest adjustment and moderate loads. applicable When Rigging for maximum load whether slide is joggled or not. TYPICAL NAS STRAP ASSEMBLY Available with various types of end hardware and up to 5000 capacity. TENSIONING SLIDE Used to preload cargo tie down straps. Handhold **CARGO BARRIER NET** CARGO TIE DOWN NET Commonly used to restrain bulky or composite cargo. Chap 12 FIGURE 12.5. Typical cargo tiedown straps and cargo nets. ### Page 97 (OCR) Rov. 1977 AC 43.13-2A office. The addresses of T.C. holders and FAA controlling regions are given in the FAA publication Type Certificate Data Sheets and Specifipacity. cations." cial operations. board configuration. tested TO the criucal ultimate load factor. Refer to chapter 1 of this handbook for computation and testing procedures. Chap 12 95 ### Page 98 (OCR) AC 43.13-2A When the cargo compartment is separated from the cockpit by a bulkhead that is capable of withstanding the inertia forces of emergency conditions a forward load factor of 4.5g. may be All other applications require the use of a 9g. forward load factor. Each cargo tiedown fitting installation must be static tested under forward, side, and up load conditions. Individual fittings may be tested by applying a single pull of 12.6g. forward load at an angle of 18.5° up and 9.5° to the left or right, as applicable, of the aircraft longitudinal axis. For example, assuming a 5,000 pound static pull (rating of a typical tiedown fitting) is applied as described and divided by the g. load factor we find the fitting installation will be capable of restraining a 397 pound load under emergency conditions. 9.5g. x 1.33 Apply 5,000 lb. pull 18.5° 9.5g. x 1.33 = 12.6g. 5,000 lb. - =397 1b. 9g. 12.6g. 1.5g. 9.18- 9.30 9g When a cargo-restraining net or cargo container with multiple attachments is used the static load requirements for each tiedown fitting may be divided equally between the fittings. For example, assume that the maximum cargo load to be carried is 1,800 pounds and 10 tiedown fittings are to be used, the static load requirement for each fitting is approximately 2,155 Ibs. 9g. x 1.33 x 1,800 10s. =2154.6 Ibs. 10 fittings Placard individual tiedowns for the maximum weight to be secured. 247. OPERATING LIMITATIONS, LOADING IN- STRUCTIONS, AND PLACARDS. Prepare revisions or supplements to the aircraft's Flight Manual 96 Rev. 1977 or operating limitations, weight and balance records, and equipment list changes as necessitated by the installation of the litter, berth, or cargo tiedown systems. NoTE.-Revisions or supplements to the approved portions of the aircraft's Flight Manual markings, placards, or other operating limitations require FAA engineering approval. Submit the requested changes and supporting data to the local FAA Flight Standards Office for review and processing. Provide instructions covering the installation and use of the litter or cargo restraint systems. For aircraft which require a Flight Manual, incorporate these instructions as a supplement. On other aircraft, provide a placard which references the appropriate instruction. In the instructions, cover such items as removal and reinstallation of seats or other equipment exchanged for litters or cargo restraint systems, use of cargo nets, barrier nets, number and positioning of tiedown straps, maximum load for each compartment or tiedown area, permissible load per square foot, number of tiedown points allowable per foot of track, and maximum height of the load's center of gravity above the floor. a. Cargo Area Placards. Install placards or other permanent markings to indicate the maximum allowable cargo load and weight per square foot limitation for each cargo area. Placard seat tracks as to number of tiedown points permissible per foot of track. Attach a permanent lable or other marking on each cargo net, barrier net, and at cargo tiedowns to indicate the maximum cargo weight that the net or attachment will restrain when installed according to the loading instructions. b. Litter and Berth Placards. Install a placard or other permanent marking on each litter or berth indicating its permissible direction of installation (forward, aft, or side-facing), passenger weight limitation (if less than 170 lbs.), and whether or not the litter or berth may be occupied during takeoffs and landings. 248.-260. [RESERVED] Chap 12 ### Page 99 (OCR) Rev. 1977 AC 43.13-2A Chapter. 13. PENETRATION THROUGH PRESSURIZED STRUCTURE Section 1. ELECTRICAL WIRE BUNDLES AND COAXIAL CABLE FEED THROUGH PRESSURIZED STRUCTURE 261. GENERAL. This section describes typical methods for sealing openings where wires and coaxial cable are installed through pressurized structure. a. Assure that the strength of the structure is maintained when installations require additional b. The aircraft manufacturer's data may specify the size and location where additional openings are permitted and the reinforcement required to maintain the design strength of the affected c. The manufacturer's data may also recommend the specific sealant to be used and provide instructions for the application. Caution: Sealant and solvents may contain toxic and/or flammable components. Avoid inhalation of vapors and supply adequate ventilation. Wear appropriate respiratory protection while using these materials in con fined areas. Avoid contact with the skin and 262. CLEANING. Use a cleaning solvent and clean a larger area than required for the fair-lead or connector. Remove solvent by blasting with dry air and wiping with a clean soft cloth. 263. APPLICATION OF SEALANT. Seal electrical wire bundles and connectors where they pass through the opening in the pressurized structure. a. Separate and coat each wire with sealant over the length which is to pass through the fairlead, fig. 13.1A. After coating each wire with sealant, pull the wire bundle into position in the fair-lead, fig. 13.1B. Assure that the fair-lead is located on the pressure side of the structure. b. Apply sealant to the surface of the fair-lead which comes in contact with the pressurized structure. Use a spatula or brush and spread sealant on the entire surface approximately ⅓2-inch in thickness. Attach the fair-lead to the pressurized structure, sealant, sealant should extrude around the mounting flange (see fig. 13.1C). c. Wrap the fair-lead with at least three turns of masking tape as shown in fig. 13.1D. Puncture the masking tape over the injection hole in the fair-lead assembly and inject sealant with a sealing gun. Apply sealant over each fairlead fastener as shown in fig. 13.1E. d. Complete all of the aforementioned steps during the application time of the sealant. Allow sealant to cure, remove masking tape and excess sealant which extruded from the fair-lead mounting flange. e. Figures 13.2 and 13.3 illustrate a different type fair-lead and wire bundle connector. The procedure for sealant application is the same as previously described. f. Figure 13.4 illustrates various coaxial connectors frequently used for installation through structure. Fair-leads are not recommended for installation of coaxial cable. Sealant application is the same as previously described. 264.-290. [RESERVED] Chap 13 97 ### Page 100 (OCR) Rev. 1977 AC 43.13-2A - EXTRUDING **SEALANT** **WIRE BUNDLE,** **SEALANT** **WIRE BUNDLE** **SEALANT** **EXTRUDING** SEALANT FIGURE 13.2 Fair-lead feed through. — EXTRUDING SEALANT FIGURE 13.3 Typical connector. Chap 13 99 ### Page 102 (OCR) Rev. 1977 AC 43.13-2A Appendix 1 CROSS REFERENCE TABLE OF PARAGRAPH NUMBER CHANGES FROM AC 43.13-2 DATED 1965 Former Paragraph Revised Paragraph Former Paragraph Revised Paragraph 1 1 57 68 2 2 58 69 70 [Reserved] 74 75 76-78 [Reserved] 79 80 13-20 [Reserved] [Reserved] 66 81 67-70 [Reserved] 82-85 [Reserved] 71 76 [Reserved] 77-80 [Reserved] 92-95 [Reserved] 29-35 [Reserved] 105 106 [Reserved] [Reserved] 107-110 [Reserved] 111 112 94 113 95 114 96-100 [Reserved] 115-120 [Reserved] 47-50 [Reserved] 57-60 [Reserved] 101 121 102 122 61 103 123 124-125 [Reserved] 101 126 105 127 [Reserved] 106 128 107 129 108 130 ### Page 103 (OCR) Rev. 1977 AC 43.13-2A Former Paragraph Revised Paragraph Former Paragraph Revised Paragraph 109 131 156 187 110 132 • 157 188 111 133 158 189 112-120 [Reserved] 134-145 [Reserved] 190-195 [Reserved] 121 146 159 196 122 147 160 197 123 148 161 198 124 149 162 199 125 150 163 200 126 151 164 201 152-155 [Reserved] 165 202 127 128 156 157 166-180 [Reserved] 203 204 129 158 205 159-160 [Reserved] 206 130 161 207-210 [Reserved] 181 162 181 211 132 163 182 212 133 134 135 164 165 166 183 184 185-199 [Reserved] 213 214 215-240 [Reserved] 136 167 200 [Reserved] 168-170 [Reserved] 201 241 137 171 202 242 138 172 203 243 139 173 204 244 174-175 [Reserved] 205 245 140-150 [Reserved] 206 246 152 151 176 177 207 208-230 [Reserved] 247 248-260 [Reserved] 158 178 231 261 154 179 232 262 180-185 [Reserved] 233 263 155 186 234-260 264-290 [Reserved] U. S. GOVERNMENT PRINGING OFFICE: 1977 0 - 251-357 --- OpenAirDocs.org is an independent research project of [Contrail Labs](https://contraillabs.com/). It is not affiliated with the FAA or the NTSB. Source documents are U.S. government works in the public domain.