Acceptable Methods, Techniques, and Practices – Aircraft Alterations

Administration (FAA) engineer or the Aircraft Certification Office (ACO) when necessary. 101. STRUCTURAL DESIGN PROCESS. Structural design processes follows these steps: a. Determine the overall load factors. b. Estimate the resulting loads. c. Distribute these loads over the aircraft. d. Determine the material, size, and shape of the part. e. Calculate the resulting stresses in the part. f. Compare these stresses with the maximum allowable for the material used. g. Resize the part as necessary. 102. TYPES OF LOADS AND STRESSES. a. Limit load factors are the maximum load factors which may be expected during service (the maneuvering, gust, or ground load factors established by the manufacturer for type certification). b. Aircraft parts may be formed out of different types of material and joined together. Each of those parts carries a load and the fastener that brings these parts together has to carry the load from one part to the other. c. Every aircraft is subject to different types of structural stress. Stress acts on an aircraft whether it is on the ground or in flight. Stress is defined as a load applied to a unit area of material. d. Tension is a force acting against another force that is trying to pull something apart. e. Compression is a squeezing or crushing force that tries to make parts smaller. f. Torsion is a twisting force. g. Shear stress is when one piece of material slides over another. h. Bending is a combination of two forces, compression, and tension. During bending stress, the material on the inside of the bend is compressed and the outside material is stretched in tension. FIGURE 1-1. BENDING OF A BEAM i. An aircraft structure in flight is subjected to variable stresses due to the varying loads that may be imposed. The designer’s problem involves anticipating the possible stresses that the structure Par 100 Page 1 AC 43.13-2B 3/3/08 will have to endure and build the structure strong enough to withstand these stresses. 103. STATIC LOADS. Static loads are loads which do not undergo change in magnitude or direction during a measurement procedure. Load factors are defined as follows: a. Limit load factors are the maximum load factors which may be expected during service (the maneuvering, gust, or ground load factors established for type certification). 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. c. Static test load factors are the ultimate loads multiplied by the casting, fitting, bearing, and/or other special factors, when applicable. Where no special factors apply, the static test loads are equal to the ultimate loads. d. Critical static test load factors are the greater of the maneuvering, gust, ground, and inertia load static test load factors for each direction (up, down, starboard, port, fore, and aft). FIGURE 1-2. TYPICAL LOAD 104. STRUCTURAL SIZING AND ANALYSIS. Design and size your load structures, including wing spars, wing attach fittings, stabilizers, landing gear struts, etc. a. Static tests using the following load factors are acceptable for equipment installations. The alteration needs to comply with the limit load factors as required by the aircraft’s certification basis. TABLE 1-1. LIMIT LOAD FACTORS Direction of Force Applied Normal-Utility Occupant 14 CFR part 23 (CAR 3) Acrobatic Occupant 14 CFR part 23 (CAR 3) Items of Mass within the cabin 14 CFR part 23 (CAR 3) Rotorcraft Occupant and Items of Mass within the cabin 14 CFR part 27 (CAR 6) Sideward 3.0g 1.5g 4.5 g 8.0g Upward 3.0g 4.5g 3.0 g 4.0g Forward 9.0g 9.0g 18.0 g 16.0g Downward 6.6g 9.0g --- 20.0g Rearward --- --- --- 1.5 g *When equipment mounting is located externally to one side, or forward of occupants, a forward load factor of 2.0g is sufficient. **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 part 23 utility category are acceptable for aircraft that never exceed the speed of 250 knots and the factors used for part 23 acrobatic category. Par 102 Page 2 3/3/08 AC 43.13-2B b. The following is an example of determining the static test loads for a 7-pound piece of equipment to be installed in a utility category aircraft (part 23). TABLE 1-2. SAMPLE LOAD FACTORS Load Factors (From the above table) Static Test Loads (Load factor X 7 pounds) Sideward 1.5g 10.5 pounds Upward 3.0g 21.0 pounds Downward 6.0g 42.0 pounds Forward 9.0g 63.0 pounds c. 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). If the additional load requires access to applicable design data or the capability to reverse engineer the installation, further assistance may be required from FAA engineering or the ACO or an appropriately rated Designated Engineering Representative (DER). 105. STATIC TESTS. CAUTION: The aircraft and/or equipment can be damaged in applying static loads, particularly if a careless or improper procedure is used. It is recommended, whenever practicable, that static testing be conducted on a duplicate installation in a jig or mockup which simulates the related aircraft structure. Static test loads may exceed the yield limits of the assemblies being substantiated and can result in partially sheared fasteners, elongated holes, or other damage which may not be visible unless the structure is disassembled. If the structure is materially weakened during testing, it may fail at a later date. Riveted sheet metal and composite laminate construction methods especially do not lend themselves to easy detection of such damage. To conduct static tests: a. Determine the weight and center of gravity position of the equipment item. b. Install attachment in the aircraft or preferably in a jig using the applicable static test load factors. c. Determine the critical ultimate load factors for the sideward, upward, downward, and forward directions. A hypothetical which follows steps (1) through (4) below pertains to the example in Figure 1-3, Hypothetical of Determining Static Test Loads. (1) Convert the gust, maneuvering, and ground load factors obtained from the manufacturer or FAA engineer or DER to determine the 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.) (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.) (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. (4) Select the highest static test load factors obtained in steps 1, 2, and/or 3 for each direction (sideward, upward, downward, and forward). These factors are the critical static test load factors used to compute the static test load. (See column 6.) d. Apply a load at center of gravity position (of equipment item or dummy) by any suitable means to 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 Par 104 Page 3 AC 43.13-2B 3/3/08 Par 105 Page 4 structure and attachments are acceptable. Should permanent deformation occur after 3 seconds, modifications or reinforcements are required to the affected structure. Additional load testing is not necessary. e. Static tests may need to be reviewed by an FAA Engineer so that appropriate limitations, procedures, checks and balances may be applied to mitigate any risks. 106. MATERIALS AND WORKMANSHIP. Use materials conforming to an accepted Government or industry standard such as Army/Navy and Air Force/Navy (AN) National Aerospace Standards (NAS), Technical Standard Order (TSO), or Military Specifications (MIL-SPEC). a. Suitability and durability of materials used for parts, the failure of which could adversely affect safety, must: (1) Be established by experience or tests; (2) Meet approved specifications that ensure the strength and other properties, assumed in the design data; and (3) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service. b. Workmanship must be of a high standard. 107. MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES. Material strength properties must be based on enough tests of material meeting specifications to establish design values on a statistical basis such as Metallic Materials Properties D Specification (MIL-HNDK-5). Design values must be chosen to minimize the probability of structural failure due to material variability. Except as provided in subparagraph e below, compliance with this paragraph must be shown by selecting design values that ensure material strength with the following probabilities: a. Where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component ― 99 percent probability with 95 percent confidence. b. For redundant structure, in which the failure of individual elements would result in applied loads being safely distributed to other load carrying members ― 90 percent probability with 95 percent confidence. c. The effects of temperature on allowable stresses used for design in an essential component or structure must be considered where thermal effects are significant under normal operating conditions. d. The design of the structure must minimize the probability of catastrophic fatigue failure, particularly at points of stress concentration. e. Design values greater than the guaranteed minimums required by this paragraph 107 may be used where only guaranteed minimum values are normally allowed. For example, if a “premium selection” of the material is made, in which, a specimen of each item is tested before use to determine if the actual strength properties of that particular item are equal or exceed those used in design. 108. FASTENERS. Use hardware conforming to an accepted Government or industry standard such as AN, NAS, TSO, or MIL-SPEC. Attach equipment in such a way that prevents loosening in service due to vibration. a. Each removable fastener must incorporate two retaining devices in case one retaining device should fail during flight operations. 3/3/08 AC 43.13-2B b. Fasteners and their locking devices must not be adversely affected by the environmental conditions associated with the particular installation. c. No self-locking nut may be used on any bolt subject to rotation in operation, unless a non- friction locking device is used in addition to the self- locking device. 109. PROTECTION OF STRUCTURE. Provide protection against deterioration or loss of strength due to corrosion, abrasion, electrolytic action, or other causes. Each part of the structure must: a. Be suitably protected against deterioration or loss of strength in service due to any cause, including: (1) Weathering, (2) Corrosion, and (3) Abrasion. b. Have adequate provisions for ventilation and drainage. 110. ACCESSIBILITY. Provide adequate provisions to permit close examination of equipment or adjacent parts of the aircraft that regularly require inspection, adjustment, lubrication, etc. For each part that requires maintenance, inspection, or other servicing, appropriate means must be incorporated into the aircraft design to allow such servicing to be accomplished. 111. AFFECTS 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 (TC), data sheet, or Aircraft Listing Aircraft Specification. When adding items of mass to the aircraft, the effect on the empty weight and balance should be considered and documented in the aircraft weight and balance. Determine that the altered aircraft will not exceed maximum gross weight. (If applicable, correct the loading schedule to reflect the current loading procedure.) Consult the current edition of AC 43.13-1, Acceptable Methods, Techniques, and Practices―Aircraft Inspection and Repair, for weight and balance computation procedures. 112. AFFECTS 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.). a. The factor of safety prescribed in 14 CFR part 23, § 23.303 must be multiplied by the highest pertinent special factors of safety prescribed in 14 CFR §§ 23.621 through 23.625 for each part of the structure where strength is: (1) Uncertain, (2) Likely to deteriorate in service before normal replacement, or (3) Subject to appreciable variability because of uncertainties in manufacturing processes or inspection methods. b. Unless otherwise provided, a factor of safety of 1.5 must be used. 113. CONTROLS AND INDICATORS. Locate and identify equipment controls and indicators so they can be operated and read from the appropriate crewmember position. 114. PLACARDING. Label equipment requiring identification and, if necessary, placard operational instructions. Amend weight and balance information as required. Any required placards installed or required by an alteration must be added to the Limitations Section of the Instructions for Continued Airworthiness to ensure that maintenance personnel will know if a required placard is missing in future inspections. 115. THRU 199. RESERVED Par 108 Page 5 AC 43.13-2B 3/3/08 FIGURE 1-3. HYPOTHETICAL OF DETERMINING STATIC TEST LOADS UTILITY CATEGORY AIRCRAFT (14 CFR PART 23) LOAD FACTORS Type of load Direction 1 Limit 2 X Safety 3 = Ultimate 4 X Special 5 Static = Test 6 Critical Static Test A. Maneuvering Fwd Down Side Up Aft (None) 6.2g (None) -3.8g 1.0g 1.5 ---------------- 1.5 1.5 ---------------- 9.30g ---------------- -5.7g 1. 5g --------------- --------------- --------------- --------------- --------------- -------------- 9.3g -------------- -5.7g 1.5g 9.3g 5.7g B. Gust (=30 FPS @ KVc) *For locations aft of fuselage Sta. 73.85. Fwd Down Down* Side Up Aft (None) 6.0g 6.4g 1.6g -2.8g (None) ---------------- 1.5 1.5 1.5 1.5 --------------- ---------------- 9.0g 9.6g 2.4g -4.2g ---------------- --------------- --------------- --------------- --------------- --------------- --------------- -------------- 9.0g 9.g 2.4g -4.2g -------------- *9.6g 2.4g C. Ground Fwd Down 6.6g 4.0g 1.5 1.5 9.9g 6.0g --------------- --------------- 9.9g 6.0g 9.9g Already Prescribed as Ultimate 9.0g 4.5g (None) D. Ultimate Inertia Forces for Emergency Landing Condition (Section 23.561). **For Separate cargo compartments. Fwd Fwd.** Down Side Up Aft Already Prescribed as Ultimate 1.5g -3.0g (None) --------------- --------------- --------------- --------------- --------------- --------------- -------------- -------------- -------------- 1.5g -3.0g -------------- **4.5g Page 6 Par 114 3/3/08 AC 43.13-2B FIGURE 1-3. HYPOTHETICAL OF DETERMINING STATIC TEST LOADS – CONTINUED Par 114 Page 7 (and 8) UTILITY CATEGORY AIRCRAFT (14 CFR PART 23) LOAD FACTORS Type of load Direction 1 Limit 2 X Safety 3 = Ultimate 4 X Special 5 Static = Test 6 Critical Static Test E. Ultimate Inertia Forces for Emergency Landing Condition For Seat, Litter, & Berth Attachment to Aircraft Structure (Section 23.785). Fwd Down Side Up Aft Already Prescribed as Ultimate 9.0g (None) 1. 5g -3.0g (None) 1. 33 --------------- 1.33 1.33 --------------- 12.0g -------------- 2.0g -4.0g -------------- 12.0g * Asterisks denote special load conditions for the situation shown. 3/3/08 AC 43.13-2B CHAPTER 2. COMMUNICATION, NAVIGATION, AND EMERGENCY LOCATOR TRANSMITTER SYSTEM INSTALLATIONS 200. PURPOSE. This chapter describes installation considerations and requirements for basic stand- alone, installations of communication, navigation, and emergency locator transmitter (ELT) equipment. NOTE: Stand-alone installations do not depend on other systems or complex interfaces to function. 201. HAZARDS AND WARNINGS. a. When installing these systems follow the aircraft and equipment manufacturers’ instructions as appropriate. Practice a “clean as you go” philosophy. Ensure that equipment and systems function properly and perform their intended function(s). b. Alterations of aircraft that are performed to accommodate the installation of radio equipment must be evaluated for their impact on aircraft design and operation. Refer to Advisory Circular (AC) 23.1309-1, Equipment, Systems, and Installations in Part 23 Airplanes, (as amended) for additional information concerning the evaluation for equipment, systems, and installations in Title 14 of the Code of Federal Regulations (14 CFR) part 23 airplanes. c. Frequently an alteration to accommodate the installation of radio equipment will have little impact on the design or operation of an aircraft; however, all potential elements of impact must be considered. One approach is to evaluate each element independently. d. Consider the impact when weight and balance or structural load limits of an added system exceeds existing installations. Consider the impact when radio frequency such as electromagnetic interference (EMI), high intensity radiated fields (HIRF), or lightning may negatively affect existing systems (e.g., accuracy of the magnetic compass). e. When evaluating elements of impact consider the before and after states. No further analysis of weight and balance may be required if an object of similar size and weight was previously installed in a location. If mounting attach points were previously substantiated to support a specific load and the same load or less is being installed, the previous analysis may be referenced. NOTE: Data that is referenced must be available and reviewed. f. When structures must be fabricated or reinforced, the standard practices approved for repairs if applicable may be employed. AC 43.13-1 (as amended), Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair, may provide structural design data for fabrication of mountings and attachments. g. Ensure that the capacity of the aircraft’s charging system is not exceeded, including any required additional allowances. h. Care should be taken to ensure that cables or wires will not interfere with the aircraft’s flight, engine, or propeller controls. i. When removing older radios/wiring/power supplies and installing newer solid state components weigh the old equipment and perform a new weight and balance calculation. This is important since differences in the location and weight of equipment will shift the center of gravity. j. Refer to AC 43.13-1 (as amended), chapter 10 and FAA-H-8083-1, Aircraft Weight and Balance Handbook, for additional information on determining of weight and balance. Par 200 Page 9 AC 43.13-2B 3/3/08 202. CONSIDERATIONS WHEN INSTALL- ING AVIONICS EQUIPMENT. When installing radio equipment, use 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 provided 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. a. General Considerations. Ensure the following: (1) There is appropriate air circulation to ensure proper cooling and dissipation of any heat generated or present. Consider the flammability characteristics of all associated elements. (2) There are appropriate clearances to prevent mechanical damage to other parts of the aircraft or from other parts of the aircraft. (3) There is protection provided to the component or article from any fluids or fumes that may be expected and that the component or article will not cause any fluids or fumes to be present that may result in damage to the aircraft or its occupants. (4) That any interference, environmental or operational, to an aircraft or any system of the aircraft is identified and minimized as to not affect airworthiness of the aircraft. (5) That flight characteristics of the aircraft are not altered unless appropriately identified and the changes are within the certified design limits. b. Structural Consideration. Consider the following: (1) Structural requirements of a mounting must be considered (see chapter 1). FIGURE 2-1. FORCE DIAGRAM NORMAL- UTILITY CATEGORY 3.0g (up) 1.5g (side) 9.0g (forward) 6.6g (down) (2) Alterations that include making additional cutouts or enlargements of existing panel cutouts must be evaluated to maintain structural integrity. Some aircraft instrument panels are load bearing structures. (3) Loads must be determined to be within the structural design limits of the supports. (4) Instrument panels as well as other panels throughout the aircraft may be structural or nonstructural in design. Structural loads must be adequately transferred to primary airframe members. (5) Methods and practices described elsewhere in this AC and in AC 43.13-1 (as amended) may be employed for the fabrication of attachments and structure. Par 202 Page 10 3/3/08 AC 43.13-2B FIGURE 2-2. TYPICAL FABRICATED PANEL MOUNTING Machine screws and self-locking nuts Rear case support Rivets or Machine screws and self-locking nuts FIGURE 2-3. TYPICAL LAYOUTS Typical layout older aircraft Typical layout newer aircraft (6) Existing structures may be reinforced or strengthened using methods described in AC 43.13-1 (as amended). 203. INSTRUMENT PANEL MOUNTING. This paragraph is supplemented by AC 43.13-1 (as amended), chapter 2, 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 part of the primary structure. Prior to making any additional “cutouts” or enlargements of an existing “cutout,” determine if the panel is part of the primary structure. If the panel is structural, make additional “cutouts” or the enlargement of existing “cutouts” in accordance with the aircraft manufacturers’ 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 (Figure 2-2). d. Added Equipment—Shock-Mounted Panels. When installing radio equipment designed for use in shock-mounted panels, total accumulated weight of equipment installed must not exceed the weight carrying capabilities of the shock mounts. Determine that the structure to which the shock mounts are connected is satisfactory for supporting the added weight. e. Existing Factory Fasteners. When possible, use existing plate nuts and machine screws provided by the aircraft manufacturer for attachment Par 202 Page 11 AC 43.13-2B 3/3/08 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 swing the compass. Install a suitable placard which indicates the compass error with the radio(s) on and off. Maximum acceptable deviation in level flight is 10 degrees on any heading. The following is an example of a typical compass calibration card (refer to current edition of AC 43.13-1, chapter 12, on how to swing a compass). TABLE 2-1. TYPICAL COMPASS CALIBRATION CARD FOR N 30 60 E 120 150 Radio On Steer 4° 35° 63° 93° 123° 154° Radio Off Steer 358° 27° 58° 88° 118° 148° FOR S 210 240 W 300 330 Radio On Steer 183° 214° 244° 274° 304° 337° Radio Off Steer 178° 208° 238° 268° 293° 327° 204. OTHER MOUNTING AREAS. The following are acceptable methods for installing radio equipment at other than instrument panel locations. a. Shock-mounted Units. (1) Wood or Composition Flooring. Secure the shock-mounted base assembly (suitable to radio unit) directly to the floor using machine screws. Add a doubler to the bottom of the floor thereby sandwiching the composition floor between each shock-mount foot and the doubler. Subsequent removal and reinstallation of the shock-mount foot will be facilitated if plate nuts are secured to the doubler. Where practicable, use small retaining screws to keep the doubler in position. Install a ground strap between the radio rack and metal structure of the aircraft. Par 203 Page 12 3/3/08 AC 43.13-2B FIGURE 2-4. TYPICAL SHOCK-MOUNTED BASE Wood Flooring Machine Screw Shock Mount Doubler Plate Nut (2) Metal Flooring. Secure the shock- mounted 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 will facilitate removal and installation of the shock mounts. Install a ground strap between the shock mount foot and the radio rack. FIGURE 2-5. TYPICAL SHOCK-MOUNTED BASE Shock Mount Metal Flooring Plate Nut Machine Screw b. Rigid-Mounted Unit Base. Secure radio mounting base plate(s) to the floor (wood, composite, or metal) using machine screws as shown in Figure 2-6. Use a reinforcing plate or large area washers or equivalent under wood or composite flooring. When the mounting base is secured to wood or composite material, install a ground strap between the base and aircraft metal structure. Par 204 Page 13 AC 43.13-2B 3/3/08 FIGURE 2-6. TYPICAL RIGID BASE PLATE MOUNT Machine Screw Machine Screws and Self-Locking Nuts 205. 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 203. 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. 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. FIGURE 2-7. TYPICAL UNDERSEAT INSTALLATION Unit attachment ear Plywood floor AN 970 flat washer Machine screw and elastic stop nut 1” clearance with seat occupied and subject to maximum seat spring deflection (6.6 X 170 pounds) downward (load). Par 205 Page 14 3/3/08 AC 43.13-2B NOTE: To increase the strength of floor attachment points, metal reinforcement may be installed as needed. 206. REINFORCEMENT OF SUPPORTING STRUCTURE. a. Attach equipment to the supporting structure of the aircraft so that its supported load will be transmitted to aircraft structural members. If direct attachment to the 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 the primary structure. When attaching to the existing structure ensure that the attachment does not weaken the structure. Alteration to primary structure may require approved engineering data. b. Placard. Fasten onto the shelf or bracket a permanent placard (as the example below) stating the design load which the installed structure is determined capable of supporting. “Shelf load not to exceed __________lbs.” Par 205 Page 15 AC 43.13-2B 3/3/08 FIGURE 2-8. TYPICAL REMOTE UNIT MOUNTING BASE-VERTICAL OR HORIZONTAL OTHER THAN STRUCTURE TO FLOOR Par 206 Page 16 3/3/08 AC 43.13-2B Par 206 Page 17 FIGURE 2-9. TYPICAL SHELF INSTALLATION NOTE: Use standard aircraft practices and procedures for fabrication and attachment of the shelf. Reinforce fore and aft corners with gussets or bulb angle. NOTE: Fabricate a platform using 2017T4(17ST) or equivalent. Apply standard aircraft practices for fabrication and installation. NOTE: The equipment manufact- urer mounting bases that meet load requirements and can be utilized are acceptable. FIGURE 2-10. TYPICAL ATTACHMENT OF SUPPORT STRUCTURE TO TUBULAR FRAME OTHER THAN STRUCTURE TO FLOOR AC 43.13-2B 3/3/08 207. ELECTRICAL REQUIREMENTS. a. Installation of an electrical system or component into an aircraft requires consideration of the electrical load, the appropriate power distribution circuit, and available power capacity. The specific requirements for a system depend on its electrical characteristics and the criticality of its application for use in the operation of the aircraft. b. The total energy available to power electrical systems is referred to as the aircraft’s capacity. It includes available stored power and generated power. This will vary depending on phase of flight or type of operation. c. The critical distribution circuit of an aircraft is designed to transfer power from source to a system determined critical to the operation or function of the aircraft. This circuit is required to include additional capacity and circuit protection as required by applicable regulations. Distribution systems should be designed to facilitate load- shedding procedures. Power distribution system design includes the following concerns. Reference AC 43.13-1 (as amended), chapter 11, for: (1) Circuit protection. (2) Wire selection. (3) Connectors, switches, and termination devices. (4) Wire routing. (5) Wiring/cable support. (6) Identification. d. Installation of Wiring. (1) Use a type and design satisfactory for the purpose intended. (2) Install in a manner suitably protected from fuel, oil, water, other detrimental substances, oxygen systems, and abrasion damages. e. Power Sources. (1) Connect radio electrical systems to the aircraft electrical system at the power source protective device, a terminal strip, or use a plug and receptacle connection. (2) Radio electrical systems must function properly whether connected in combination or independent. f. Protective Devices. (1) Incorporate a “trip free” re-settable 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 and label. (2) Select circuit breakers or fuses that will provide adequate protection against overloading of the radio system wiring. (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, unless the equipment is intended to be powered when the master switch is open. (4) Radio system controls are to provide independent operation of each system installed and be clearly labeled to identify their function relative to the unit of equipment they operate. g. 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 fluids lines and securely clamp to the structure. (In no case must radio electrical wire bundles be clamped to lines containing flammable fluids.) Par 207 Page 18 3/3/08 AC 43.13-2B Par 207 Page 19 h. Cable Attachment to Shock-Mounted Units. 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. i. Radio Bonding. It is advisable to electrically bond radio equipment to the aircraft in order to provide a low impedance ground and to minimize radio interference from static electrical charges. When electrical bonding is used, observe the following: (1) Keep bonding jumpers as short as possible. (2) Prepare bonded surfaces for best contact (resistance of connections should not exceed 0.003 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 parts. (5) When contact between dissimilar metals cannot be avoided, put a protective coating over the finished connection to minimize corrosion. 208. ELECTRICAL LOAD ANALYSIS PROCEDURE. a. Available Power Supply. To preclude overloading the electric power system of the aircraft when additional equipment is added, perform an 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. Compute the electrical load analysis for the most adverse operating conditions, typically this is for night and/or instrument flight. b. One method for the analysis may be found in the ASTM International, Standard Guide for Aircraft Electrical Load and Power Source Capacity Analysis, F 2490-05. c. Applicable elements of a previously performed electrical load analysis if available may be reused. 209. ELECTROMAGNETIC COMPAT- IBILITY. a. Electromagnetic Interference (EMI) may disrupt the performance of systems and has varying degrees of consequence. These consequences may be identified as: no safety effect, minor, major, hazardous, or catastrophic. The purpose of electromagnetic compatibility (EMC) analysis and testing is to assure that equipment does not cause interference with any existing aircraft system function, and that existing systems do not cause any interference with the new equipment. (Refer to Table 2-2.) AC 43.13-2B 3/3/08 TABLE 2-2. RELATIONSHIP PROBABILITIES, SEVERITY OF FAILURE CONDITIONS Classification of Failure Conditions No Safety Effect <--Minor--> <--Major--> <--Hazardous--> <--Catastrophic--> Effect on Airplane No effect on operational capabilities or safety Slight reduction in functional capabilities or safety margins Significant reduction in functional capabilities or safety margins Large reduction in functional capabilities or safety margins Normally with hull loss Effect on Occupants Inconvenience for passengers Physical discomfort for passengers Physical distress to passengers, possibly including injuries Serious or fatal injury to an occupant Multiple fatalities Effect on Flightcrew No effect on flightcrew Slight increase in workload or use of emergency procedures Physical discomfort or a significant increase in workload Physical distress or excessive workload impairs ability to perform tasks Fatal Injury or incapacitation b. When EMC characteristics are known, the need for and extent of EMC testing can be determined by a review of those characteristics. Knowing specific target frequencies, EMC testing can focus on the aircraft systems (and even those susceptible frequencies in the case of tunable systems) likely to be affected by interference. Where sensitive systems or potentially strong sources of EMI are involved, more intensive evaluation will be required. TABLE 2-3. COMMONLY USED RADIO FREQUENCIES ON AIRCRAFT Range Hz Mode Function 190-1750 kHz Rx ADF Navigation 2-30 MHz Tx/Rx HF Communications 75 MHz Rx Marker Beacon Receiver 108-112 MHz Rx ILS Localizer Receiver 108-118 MHz Rx VHF Omnirange (VOR) Receiver 118-137 MHz Tx/Rx VHF Communications 243 MHz Tx Emergency Locator Xmtr (Satellite) 328.6-335.4 MHz Rx ILS Glide Slope Receiver 406.3 MHz Tx Emergency Locator Transmitter 960-1215 MHz Tx/Rx DME System 1027-1033 MHz Tx/Rx Transponder & TCAS Systems 1087-1093 MHz Tx/Rx Transponder & TCAS Systems 1575.42 MHz Rx GPS Satellite Navigation c. On-aircraft EMC tests for systems or equipment should be conducted. If lab test data is available it should be used to guide the planning of these tests. If lab test analysis is not available a more comprehensive EMC testing on the aircraft systems and equipment typically needs to be performed. (1) The aircraft should not be close to large reflecting surfaces such as buildings or other Par 209 Page 20 3/3/08 AC 43.13-2B aircraft. Use of ground power is not recommended, as ground power units are not routinely checked for output quality. (2) All normally closed circuit breakers should be closed and power should be supplied to all normally powered AC and DC distribution busses during testing. (3) The aircraft should be in flight configuration. Doors and hatches that might be in any interfering signal’s path should be in the position they would normally be in during flight. (4) Aircraft systems being tested should be operated and monitored for indication of interference. (It is essential that systems determined critical for the operation of the aircraft are tested.) The following systems, if installed, should be included in the aircraft EMC test plan. TABLE 2-4. AIRCRAFT ELECTRICAL AND ELECTRONIC SYSTEMS System ADF Air Data Systems Altitude Alert System ATC Transponder Audio Distribution System Autopilot/Flight Guidance System Compass/Directional Gyro Systems DME Electronic Flight Control System Electronic Flight Instrument System (EFIS) Global Positioning System (GPS) Marker Receiver System VHF Communications VOR/LOC/GS Newly Installed Electrical/Electronic System 210. FUNCTIONAL HAZARD ASSESSMENT (FHA). NOTE: Refer to AC 43.1309-1 (current edition) for additional information on FHA. a. Many older aircraft designs did not provide for all weather operations or potential increases in the pilot’s reliance on installed systems and equipment. Requirements to assure design safeguards against hazards have developed as technologies have become available and pilots have increased their reliance on installed systems and equipment. A fundamental analysis to assure design safeguards is accomplished through a functional hazard assessment. b. Alterations that involve systems or equipment that perform critical functions or that include complex designs that have a high degree of integration, use of new technology, or novel applications of conventional technology, must be assessed to determine the severity of failure conditions. Complexity in itself does not drive the need to perform a system safety analysis but the effect of a failure does. Comparison with similar, previously approved systems is sometimes helpful. c. Evaluate the system to determine if it is essential or not essential to safe operation. For aircraft of 6,000 pounds or less maximum weight, refer to the regulations incorporated by reference in the type certificate, unless the Administrator has found that the change is significant in an area. d. Determine if the equipment has any unacceptable, adverse affect when operated. e. Determine if the operation of the installed equipment has an adverse affect on equipment not essential to safe operation, and if a means exists to inform the pilot of the effect. f. Determine if a failure or malfunction of the installed equipment could result in unacceptable hazards. Par 209 Page 21 AC 43.13-2B 3/3/08 Par 210 Page 22 g. Design requirements and methodology of hazard resolution differing upon application and type of aircraft (i.e., single-engine, multiengine, commuter use). (1) Operation of equipment that has an adverse effect on other equipment essential to safe operation of the aircraft is unacceptable. (2) Operation of equipment that has an adverse effect on other equipment that is not essential to safe operation of the aircraft may be acceptable if there is a means to inform the pilot of the effect. (3) If a probable failure or malfunction will result in a hazard in a multiengine aircraft it is unacceptable. (4) If a probable failure or malfunction will result in a hazard in a single-engine aircraft its impact must be minimized. 211. THRU 299. RESERVED 3/3/08 AC 43.13-2B CHAPTER 3. ANTENNA INSTALLATION 300. PURPOSE. The purpose of this chapter is to describe antenna installation methods and practices. An antenna that is installed on an aircraft must function properly and may not adversely affect other systems or equipment. 301. HAZARDS AND WARNINGS. a. Follow antenna manufacturer’s instructions and recommendations when they are available and appropriate, and not contrary to the instructions of the aircraft manufacturer. b. Extension of the landing gear or flaps may impact belly-mounted antenna performance. 302. ADDITIONAL REFERENCES. For further information concerning acceptable methods, techniques, and practices concerning alteration involving specific structures, refer to the appropriate chapter of Advisory Circular (AC) 43.13-1, Acceptable Methods, Techniques, and Practices―Aircraft Inspection and Repair (current edition), and Civil Aviation Regulation 6, Rotorcraft Airworthiness; Normal Category. 303. STRUCTURAL SUPPORT. a. The antenna’s structural load, plus any required allowances, may not exceed the design capacity of the structure intended to support it. It is important to understand the operational characteristics of the aircraft and consider forces that occur during flight (dynamic loading) as well as those that occur when the aircraft is not in motion (static loading). For example, an aircraft designed without flaps may employ a side slip procedure to lose altitude, during which the direction of airflow across the fuselage is not in line with the aircraft longitudinal axis. Antenna mountings on these aircraft need to be designed and evaluated for the direction of airflow that occurs during such an operation. b. Whenever possible, an antenna should be mounted to a flat surface. Minor aircraft skin curvature can be accommodated with the use of an appropriate gasket but if gaps over 0.020" appear between the base plate and mounting surface, use of a mounting saddle is recommended. c. Since antenna systems typically require a ground plane (this may be a conductive surface that the antenna mounts to) any separation of an antenna from its ground plane may impact performance. Contact the manufacturer for recommendations if a gasket or mounting saddle is needed. Par 300 Page 23 AC 43.13-2B 3/3/08 FIGURE 3-1. ANTENNA MOUNT WITH SADDLE Upper skin ref Stringer 30L ref Stringer 30R ref UP OUT Looking aft Scale: Full d. Mounting screws must never be over torqued in an attempt to distort aircraft structure to reduce gaps between the antenna base plate and aircraft-mounting surface. e. Consider the factors of flutter, vibration characteristics, and drag load. The approximate drag load an antenna develops may be determined by the formula: D=0.000327 AV2 (The formula includes a 90 percent reduction factor for the streamline shape of the antenna.) D is the drag load on the antenna in lbs. A is the frontal area of the antenna in sq. ft. V is the VNE of the aircraft in mph. Example: Antenna manufacturer specification frontal area = 0.135 sq. ft. and V NE of aircraft is 250 mph. D=0.000327 x .135 x (250) 2 =0.000327 x .135 x 62,500 = 2.75 lbs f. The above formula may be adapted to determine side load forces by substitution of the apparent frontal area value for A, when the aircraft motion and antenna orientation are not the same. 304. PHYSICAL INTERFERENCE. a. Antennas should be located where they will not interfere with the operation of the aircraft or other aircraft systems. One such example of interference could be the obstruction of visibility of a navigation position light or beacon. b. Antennas should be located so that they don’t obstruct or limit airflow to areas of the airframe that require airflow. Care should be exercised that an antenna is not located where it will be damaged by heat from engine exhaust, fumes from battery vents, or fuel/fluid drains. Par 303 Page 24 3/3/08 AC 43.13-2B c. Antennas may accumulate ice that can then depart and damage areas behind them. Special attention should be paid to areas of the airframe near pitot static ports and sensors and near flight controls, since the antenna may alter airflow characteristics. d. Antennas should be located in such a manner that they are not susceptible to damage from misuse, such as near a door where they might be mistaken for or used as a handhold. 305. ANTENNA SELECTION. The selection of an appropriate antenna will include consideration of system requirements and aircraft characteristics. The size and shape of an antenna varies with frequency, power rating, and maximum design speed of the antenna. See Figures 3-7 through 3-19 for pictures of typical antennas. 306. ANTENNA LOCATION. a. In general, antenna locations on an aircraft which provide unobstructed line-of-sight views of the transmitted or received signals are preferred. Objects located in the path of a signal may cause a blanking of antenna coverage and impact the performance of the system. NOTE: Global Positioning System (GPS) antennas will not receive a signal if a line-of-sight view of a satellite is not available. Do not mount a GPS antenna on the underside of an aircraft. b. Acceptable and unacceptable spacing between an antenna and an obstacle or the permissible interval between antennas is dependent upon operating frequency and system characteristics. When in doubt contact the antenna manufacturer or system designer for further information. c. Antennas should be separated as far as possible from interference sources (other radiating antennas, ignition noise sources, etc.). When known interference sources are present it may be advisable to temporarily position an antenna and check a location for suitability prior to mounting the antenna. d. VHF Com 1 should be mounted on the top of the aircraft since this will provide the best unobstructed location. VHF Com 2 can also be mounted on the top, provided there is at least 1/2 wavelength (of the antenna operating frequency) distance available between antennas. e. If Com 2 is mounted under the aircraft, a bent whip may be required to provide ground clearance. Bent whips may not provide the best performance because of proximity to the aircraft skin. Signal reflection and obstruction is more of a problem with such locations. Extension of the landing gear or flaps may also impact belly mounted antenna performance. f. Antennas need to “see” with a direct line- of-site to the source. Antenna patterns can be disrupted by landing gear or vertical stabilizers, for example. When mounting antennas, try to locate them in areas where line-of-sight view is not obstructed. g. As a rule of thumb, maintain 36 inches as a minimum distance between antennas. Refer to manufacturer’s installation guidelines for specific system limitations and requirements. h. Antennas should be located such that cable runs between antenna and equipment are as short as practical. Signal loss of a cable is dependent upon cable design, length, and frequency. In some cases, cables must be specific lengths to provide a required capacitance, attenuation, or signal transmission time. VOR/LOC/GS blade or towel bar type antennas require the cables to their coupling assembly to be the exact same length to maintain a phase relationship. 307. ANTENNA BONDING. a. The electrical bonding of the antenna to the aircraft surface is extremely important. The conductive skin of an aircraft is an electrical part of the antenna system. If an antenna is not properly bonded to the aircraft, its pattern may be distorted and nulls in coverage may appear. Par 304 Page 25 AC 43.13-2B 3/3/08 b. The electrical bonding of the antennas to the aircraft skin of a metal aircraft is best accomplished by direct metal-to-metal contact of the antenna base to the skin. A resistance of no more than 0.003 ohms between the antenna base plate and skin should be achieved. NOTE: To achieve this electrical bonding, the aircraft paint in the mounting area will need to be removed and the surface covered with an oxide film (i.e., aluminum conversion coat) to protect aluminum against corrosion in accordance with MIL-C-5541B. c. An alternate method for providing electrical bonding to metal aircraft skin is through the antenna mounting screws, which attach to a backing plate inside the aircraft, making electrical contact with the backside of the skin. To ensure good contact, remove any interior paint in the area where the backing plate is placed and coat this area in accordance with MIL-C-5541B to minimize corrosion. d. Composite or fabric covered aircraft that do not provide a conductive mounting surface generally require fabrication of a conductive surface (ground plane) and bonding through the mounting screws. e. Antenna performance can be severely degraded from corrosion caused by moisture accumulation where the antenna electrically bonds to its ground plane. It is advisable to apply RTV around the antenna edges to seal the antenna bond; however, always ensure chemical compatibility before using any sealant. 308. ELECTROMAGNETIC INTERFERENCE (EMI). a. Since the purpose of an antenna is to either receive or transmit RF energy (or both), it is essential to consider EMI. b. Antenna mounting positions should be selected that are as far as possible from an EMI source. In special cases, it may be possible to employ filters or select an antenna that has been specifically designed to be resistant to EMI. c. EMI test procedures are found in chapter 2. 309. MECHANICAL INSTALLATION. a. Mounting Hardware. (1) Typical antenna installations employ either #8-32 or #10-32 stainless steel mounting screws. (2) Some designs require a pan head screw. For others, a counter sunk screw is required. (3) Mounting screw length will vary based on each particular installation requirement. b. General Practices. (1) Refer to installation drawing before drilling holes in aircraft skin to determine proper size and spacing. NOTE: When replacing antennas, it is important to match the original mounting holes. Previous mounting holes that are not reused with appropriate hardware must be repaired and the mounting location returned to its design strength. (2) Use of a structural backing plate is highly recommended. Backing plates strengthen the immediate point of attachment but if they are not attached to load carrying structure they do not provide structural load support. (3) Mounting screws should be secured with stainless steel nuts with flat washers and lock washers, or with flat washers and lock nuts to secure the antenna properly. (4) Sandwich the aircraft skin between the antenna base plate and the internally mounted backing plate. Before securing the antenna to the aircraft make sure that all the cables are connected to Par 307 Page 26 3/3/08 AC 43.13-2B Par 309 Page 27 the unit and fit through the connector holes in the aircraft. (5) Gently tighten the mounting hardware so that uniform stress is placed on each side of the antenna. For #8-32 screws do not exceed 20 in•lbs of torque and for #10-32 screws do not exceed 23 in•lbs of torque. Refer to fastener manufacturer’s torque guidelines to confirm that these recommended settings do not exceed the chosen fasteners torque limits. b. Most antennas require a ground plane size of approximately 24" by 24". While the rule of thumb is to provide a minimum of 1/4 wavelength of the operation frequency, larger is better and ground plane symmetry is critical. Gaps in antenna coverage or performance may occur if a ground plane is not symmetrical. TABLE 3-1. WAVELENGTH IN FEET = 984 / FREQUENCY IN MEGAHERTZ (MHZ) (6) Once the antenna is mounted, any minor gaps between the base plate or gasket and aircraft skin should be filled with RTV silicone adhesive sealant. Frequency Wavelength (ft) 1/4 wavelength (in) 75 MHz 13.1 40 125 MHz 7.9 24 1000 MHz 1 3 (7) Double check that a reading of 0.003 ohms between the antenna base plate and ground has been achieved. 310. GROUND PLANE REQUIREMENT (WOOD/FABRIC AND COMPOSITE AIRCRAFT). c. Wire mesh is the best material to use when a solid plate is not practical. Heavy aluminum foil can also be used. In all cases, electrical continuity from ground plane to airframe ground is essential. d. Be sure the ground plane is well attached to the airframe with cement or epoxy if not otherwise supported. This will prevent noise problems or erratic operation that could occur if the plane moves. Capacitance will occur as an antenna base is separated from its ground plane. This may distort antenna coverage or operation. a. When the antenna is not mounted to a conductive surface capable of providing a required ground plane for operation, a ground plane must be fabricated. FIGURE 3-2. ANTENNA GROUND PLANE FOR NONCONDUCTIVE MOUNT AC 43.13-2B 3/3/08 FIGURE 3-3. ONE MEANS TO PROVIDE ADEQUATE ANTENNA BONDING THROUGH A COMPOSITE AIRFRAME (Length to meet installation requirements) Flathead screw stainless steel Seal Aluminum Grounded Foot Plate Composite Aircraft Aluminum Ground Plate Ground to Aircraft (4 places) Doubler Plate Elastic Stop Nut Stainless Steel or equivalent FIGURE 3-4. USE BUSHINGS FOR ALL SCREWS NOTE: Carbon Fiber composite material, while conductive has not been found to be adequate as a ground plane. e. Refer to AC 43.13-1 (current edition), chapter 3, for acceptable methods, techniques, and practices applicable to fiberglass and plastics. 311. ANTENNA FEED LINE BALUN. a. Antenna cables (electrical feed lines) may be designed to be electrically balanced or unbalanced. Certain antenna designs incorporate dual elements with a requirement for balanced input. In these cases, standard cables which are unbalanced are generally employed for the cable run and a balancing transformer is located at the antenna feed connection. b. A balun is a device that converts an unbalanced feed into a balanced input and may include a transformer that matches feed impedance to provide maximum signal transfer. Follow the manufacturer’s installation procedures when a balun is required. Some balun designs require that the balun be grounded to the airframe. c. Refer to AC 43.13-1 (current edition), chapter 11, for bonding practices. Length to be determined per fastener at installation Outer Diameter as required Par 310 Page 28 3/3/08 AC 43.13-2B FIGURE 3-5. TYPICAL DIPOLE ANTENNA ASSEMBLY FIGURE 3-6. TYPICAL VOR BALUN 312. ANTENNA REPAIR. Painting an antenna or applying protective coatings or devices that are not approved are not allowed under this AC. Paint is an RF de-tuner. If an antenna is painted in the field, paint type and paint thickness present uncontrolled variables that will affect an antenna’s performance, and may result in the antenna no longer meeting its specifications or Technical Standard Order (TSO). 313 THROUGH 399 RESERVED Par 312 Page 29 AC 43.13-2B 3/3/08 FIGURE 3-7. ADF LOOP FIGURE 3-8. ADF COMBINED SENSE LOOP FIGURE 3-9. COM WHIP FIGURE 3-10. COM WHIP BENT FIGURE 3-11. ELT FIGURE 3-12. GPS FIGURE 3-13. GLIDESLOPE FIGURE 3-14. MARKER Par 312 Page 30 3/3/08 AC 43.13-2B Par 312 Page 31 (and 32) FIGURE 3-18. TRANSPONDER/DME PROBE FIGURE 3-15. COMBINED COM/VOR FIGURE 3-16. VOR RABBIT EAR FIGURE 3-19. DME/TRANSPONDER BLADE FIGURE 3-17. VOR BLADES 3/3/08 AC 43.13-2B CHAPTER 4. ANTICOLLISION AND SUPPLEMENTARY LIGHT INSTALLATION 400. PURPOSE. This chapter gives procedures and standards to be used when replacing older rotating beacon assemblies and wing lights with strobe or other anticollision systems. This chapter assumes that the newer units have FAA approval in the form of a Parts Manufacturer Approval (PMA) or Technical Standard Order (TSO), and not an experimental or aviation unit. 401. HAZARDS AND WARNINGS. When installing anticollison lights take care to ensure the unit is properly grounded, the airframe structure can support the new unit, and the aircraft wiring is of the correct size. Mechanics should take special precautions to protect their eyes when testing the new unit. Strobe lights are especially hazardous in dark or darken hangars when activated. 402. REGULATIONS AND OTHER REFERENCES. The requirements for anticollision lights are included in Title 14 of the Code of Federal Regulations (14 CFR) part 23, § 23.1401 and part 27, § 27.1401 for non-transport category aircraft. For part 23 aircraft certificated after March 11, 1996, § 91.205 are required to have an anticollision light. The night VFR requirements for part 23 certificated on or before August 11, 1971, must have an approved white or red anticollison light. 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 standards: Additional information can be found in AC 20-74. a. Civil Aviation Regulation (CAR) 6, Rotorcraft Airworthiness; Normal Category. b. Anticollision lights (when installed) should be installed on top of the fuselage or tail in such a location that the light will not impair the flight crewmembers’ 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 or wing tip installation may be used. c. The color of the anticollision light must be either aviation red or aviation white in accordance with the specifications of § 23.1397 or § 27.1397, as applicable. d. The arrangement of the anticollision light system, (i.e., number of light sources, beam width, speed of rotation and other characteristics, etc.) must give an effective flash frequency of not less than 40, nor more than 100, cycles per minute. The effective flash frequency is the frequency at which the aircraft’s complete anticollision light system is observed from a distance, and applies to each sector of light including any overlaps that exist when the system consists of more than one light source. In overlaps, flash frequencies may exceed 100 but not more than 180 cycles per minute. e. The system must consist of enough lights to illuminate the vital areas around the aircraft, considering the physical configuration and flight characteristics of the aircraft. The field of coverage must extend in each direction within at least 75 degrees above and 75 degrees below the horizontal plane of the aircraft. The minimum light intensity and minimum effective intensities are given in §§ 23.1401 and 27.1401 respectively. f. 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. 403. OPERATIONAL CONSIDERATIONS: CREW VISION. Partial masking of the anticollision light may be necessary to prevent direct Par 400 Page 33 AC 43.13-2B 3/3/08 or reflected light rays from any anticollision or supplementary light from interfering with crew vision. Determine if 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, sometimes known as flicker vertigo, has been eliminated. Enter a notation to that effect in the aircraft records. 404. INSTALLATION CONSIDERATIONS. a. 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. b. 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 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. 405. 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 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. 406. 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/alternator limitations is contained in Advisory Circular (AC) 43.13-1B Acceptable Methods, Techniques, and Practices-Aircraft Inspection and Repair, (as amended) chapter 11. Assure that the terminal voltage at each light is within the limits as prescribed by the manufacturer. 407. 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 Figure 4-1.) Par 403 Page 34 3/3/08 AC 43.13-2B FIGURE 4-1. TYPICAL ANTICOLLISION OR SUPPLEMENTARY LIGHT INSTALLATION IN A SKIN PANEL (UNPRESSURIZED) Anticollision Light or Supplementary Light Fuselage Skin Existing Stringer Existing Stringer View A-A Reinforcing Doubler Approximately One Inch Spacing of 1/8” Min.Dia.Rivet Alclad 2024-T3 1 1/2” Edge Distance Min. b. When a formed angle stringer is cut and partially removed, position the reinforcement doubler between the skin and the frame. The doubler is 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 to be twice the height of the doubler flange. (See Figure 4-2 for typical installation.) c. Engineering evaluation is required for installations 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 if 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 Figure 4-3.) e. Rudder installations are not recommended because of the possible structural difficulties. However, if such installations are considered, a FAA engineering evaluation to determine whether the added mass of the light installation will adversely affect the flutter and vibration characteristics of the tail surfaces must be made. Par 407 Page 35 AC 43.13-2B 3/3/08 FIGURE 4-2. TYPICAL ANTICOLLISION OR SUPPLEMENTARY LIGHT INSTALLATION INVOLVING A CUT STRINGER (UNPRESSURIZED) Anticollision Light or Supplementary Light Cut Stringer Reinforcing Doubler Alclad 2024-T3 Stringer Stringer Frame Frame Double Flanges at Least the Height of Cut Stringer Flange Stringer and Doubler Existing Rivets Rivets Added as Shown on Frames Par 407 Page 36 3/3/08 AC 43.13-2B FIGURE 4-3. TYPICAL ANTICOLLISION OR SUPPLEMENTARY LIGHT INSTALLATION IN A FIN TIP Mounting Screw Holes for Light to Fairing Anticollision Light or Supplementary Light Approximately One Inch Spacing of 1/8” Min. Dia. Rivet Vertical Stabilizer Light Mounting Fairing Mounting Ring Vertical Stabilizer View A-A B-B NOTE: Skin thickness of mounting ring and fairing are at least equivalent. Par 407 Page 37 (and 38) 3/3/08 AC 43.13-2B CHAPTER 5. SKI INSTALLATIONS 500. PURPOSE. This chapter provides information for ski installations on small airplane. The information provided for main ski and nose ski installations applies to wheel replacement skis only. Tire-cushioned skis other than tail skis, wheel penetration skis, and hydraulically adjustable or retractable skis involve special considerations and cannot be installed by relying solely on data in this advisory circular (AC). 501. HAZARDS AND WARNINGS. Operation of ski planes exposes the airplane and its occupants to additional risks not associated with wheel-equipped landplanes. The additional weight and surface area of skis impose additional ground and air loads on the airplane. Ground handling, taxiing, takeoff, and landing can place significant side loads and twisting moments to the landing gear and its attachment structure which can cause hidden and/or cumulative damage. Improper rigging and/or weak springs or shock cords can cause the skis to “dump,” or rotate nose down in flight, possibly rendering the airplane uncontrollable or causing it to break up in flight. Skis with weak springs or shock cords may rotate nose down and “dig,” or penetrate the snow, during takeoff or landing on uneven or drifted snow, which could result in an accident. In-service failure of ski attachment hardware, springs, shock cords, or cables creates a high risk of those parts or a ski itself entering the propeller arc, which has resulted in complete loss of airplanes in flight. For these reasons, proper installation, rigging, periodic inspection, and maintenance of skis and their attaching parts are of utmost importance to safety. Consultation with experienced ski maintenance technicians and operators is strongly recommended when considering any new ski installation or any alteration of an existing ski installation. 502. ADDITIONAL REFERENCE MATERIAL (current editions). a. Airframe and ski manufacturers’ data, if available. b. AC 43.13-1, Acceptable Methods, Techniques, and Practices―Airplane Inspection and Repair. c. AC 43-210, Standardized Procedures for Requesting Field Approval of Data, Major Alterations, and Repairs. d. Civil Aviation Regulation (CAR) 6, Rotorcraft Airworthiness; Normal Category. e. FAA Order 8110.54, Instructions for Continued Airworthiness. f. FAA-H-8083-23, Seaplane, Ski plane, and Float/Ski Equipped Helicopter Operations Handbook. 503. INSTALLATION CONSIDERATIONS. a. Determining Eligibility of an Airplane. Only an airplane approved for operation on skis is eligible for ski installations in accordance with this chapter. Eligibility can be determined by referring to the Aircraft Specifications, type certificate data sheets (TCDS), Aircraft Listing, Summary of Supplemental Type Certificates, or by contacting the manufacturer. Also determine the need for any required alterations to the airplane to make it eligible for ski operations. (See ski plane-specific entries throughout TCDS A4CE on airspeeds, weight and balance limitations and additional placards, and the Required Equipment listing for the first model in Aircraft Specification A-790 for examples.) If the airplane is not approved in a ski plane configuration by type design, then skis cannot be installed by relying solely on data in this AC. Contact FAA engineering for approval or obtain/develop approved data from another source. Par 500 Page 39 AC 43.13-2B 3/3/08 b. Identification of Approved Model Skis. Determine that the skis are of an approved model by examining the identification plates or placards displayed on the skis. Skis of approved models will have such plates or placards, and the Technical Standard Order (TSO) number TSO-C28, an Aircraft Component, Accessory, or equipment type certificate (TC) number, or an airplane part number (if the skis have been approved as part of the airplane) will be engraved or imprinted on each plate or placard. c. Maximum Limit Load Rating. (1) Known limit landing load factor. Before installation, determine that the maximum limit load rating (L) of the ski as specified on its identification plate or placard is at least equal to the maximum static load on it (S) times the limit landing load factor (fl) previously determined from drop tests of the airplane by its manufacturer. This requirement can be expressed by the following equation: L=S×fl (2) Unknown limit landing load factor. In lieu of a value fl determined from such drop tests, a value of fl determined from the following formula may be used: 9000 η= 2.80 + (W + 4000) where “W” is the certificated gross weight of the airplane d. Oversize Ski Installations. This limitation is to assure that the oversize skis will not adversely affect the performance, stability, controllability, and spin recovery behavior of the airplane or impose excessive loads on it. e. Landing Gear Moment Reactions: Landing Gear Bending Moments. In order to avoid excessive bending moments on the landing gear and attachment structure, the ski pedestal height measured from the bottom surface of the ski to the axle centerline must not exceed 130 percent of the static rolling radius of the standard tire approved for the airplane, when the tire is installed on the standard wheel at the approved inflation pressure. Do not use oversize or "tundra" tires to determine the static rolling radius. 504. FABRICATION AND INSTALLATION. a. Hub-Axle Clearance. The pedestal hub should fit the axle to provide a clearance of 0.005" minimum to 0.020" 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 fail. (See Figure 5-1.) Field experience has shown that the use of good quality, low-temperature grease; particularly modern synthetic-based grease, improves ski operation and wear protection when used on the axle-to-hub or axle-to-bushing faying surfaces. FIGURE 5-1. TYPICAL HUB INSTALLATION Par 503 Page 40 3/3/08 AC 43.13-2B FIGURE 5-2. TYPICAL SKI INSTALLATION 8 1. Fitting 2. Shock Cord 3. Safety Cable 4. Tape 5. Crust-cutter Cable 6. Fabric removed to facilitate inspection 7. Check Cable 8. Clevis 9. Ski Pedestal 10. Pedestal Height b. 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. c. Cable and Shock Cord Attachment and Attachment Fittings. (1) Field experience. 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. Failures of the safety cable and shock cord attachment fittings usually occur at the ski end and not at the fuselage end. (2) Separating attachment points. It is strongly recommended that tension cords and safety cables be attached to entirely separate fittings at their fuselage ends. Although the attachment fitting detail shown in Figure 5-2 may be adequate for some installations where alternate attachment locations are unavailable, we recommend that each cord and cable be attached to its own fitting (such as the right-hand fitting in Figure 5-3) and attached at separate points on the fuselage when possible. Provide separate means of attaching cables and shock cords at the forward and aft ends of the skis. (3) Fabrication. Approved skis are normally supplied with cables, shock cord, and fittings. However the specifications in Table 5-1, subparagraphs (a) through (c) below, and Figure 5-3 may be used for fabricating and installing cables, shock cord assemblies, and fuselage fittings. NOTE: Field experience indicates that accelerated wear and damage can occur to a 1/8" safety or crust- cutting cable and its attachment hardware in normal service on skis having a limit load rating of 1,500 pounds or more. The FAA Par 504 Page 41 AC 43.13-2B 3/3/08 recommends a minimum cable size of 5/32" for use in fabricating safety or crust-cutting cables for use with skis of 1,500 pounds or greater limit load rating. 1/8" cables may be suitable for use on airplanes with light-weight skis and maximum certificated weights less than 1,500 pounds, such as those meeting the definition of light sport aircraft. TABLE 5-1. RECOMMENDED MINIMUM CABLE AND SHOCK CORD SIZES Ski Limit Load Single Safety Double Safety Single Crust- Double Crust- Single Shock Double Shock Rating Cable Cable Cutting Cable Cutting Cable Cord Cord (Pounds) Less than 1,500 1/8" 1/8" 1/8" 1/8" 1/2" 1/2" 1,500-3,000 5/32" 5/32" 5/32" 5/32" 1/2" 1/2" 3,000-5,000 Do Not Use 5/32" 5/32" 5/32" Do Not Use 1/2" 5,000-7,000 Do Not Use 5/32" 5/32" 5/32" 3/4" 3/4" 7,000-9,000 Do Not Use 3/16" Do Not Use 5/32" Do Not Use 3/4" (a) Make check cable, safety cable, and crust-cutting cable ends by the splice, swage, or nicopress methods. Cable clamps may be used if adjustable lengths are desired, but they are not recommended. Use standard airplane hardware only. (Hardware used to attach cables must be compatible with cable size.) Refer to AC 43.13-1, chapter 7, as amended, for more information on cable fabrication. (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, 0.041" (minimum) safety wire (ref. National Aerospace Standard NASM20995), 1025 steel, or its equivalent (AISI 4130). Attach with clevis or spring steel snap fastener. (Do not use cast iron snaps.) 2. Use approved spring-type shock cord end fasteners, 1025 steel, or its equivalent (AISI 4130). (c) Fitting (Figure 5-3) Specifications and Installation: 1. Fittings fabricated for 1/8- inch cable or 1/2-inch shock cord must be at least 0.065" 1025 steel or its equivalent. 2. Fittings fabricated for 5/32- inch cable or 3/4-inch shock cord must be at least 0.080" 1025 steel or its equivalent. 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. 4. If attaching cables directly to holes in fittings, radius the hole edges to reduce stress concentration and accelerated wear of the thimble. Stainless steel thimbles are recommended for increased wear resistance. 5. If attaching cables to fittings using clevises, clevis bolt castellated nuts should be used, then properly torqued and safetied with cotter pins. Field experience has shown that diaper-pin- style quick-releasing safety devices are more prone to failure during operation, and are not recommended. Par 504 Page 42 3/3/08 AC 43.13-2B FIGURE 5-3. TYPICAL FUSELAGE FITTINGS d. Provisions for Inspection. An airplane using fabric-covered landing gear should have at least the lower 4 inches of fabric removed to facilitate inspection of the axle attachment area, and to prevent the entrapment of snow and ice, which can lead to damage and corrosion of the landing gear. (See Item 6 in Figure 5-2.) 505. RIGGING OF SKIS. a. Location of Attach Fittings on Fuselage or Landing Gear. Locate fittings so the shock cord and cable angles are not less than 20 degrees when measured in the vertical plane with the shock absorber in the fully extended position (see Angle B, Figures 5-4 and 5-5). NOTE: Do not attach fittings to wing-brace struts, except by special approval (manufacturer or FAA). b. Main Ski Incidence Angles. (1) Set cable lengths with the airplane level and no weight on the landing gear. (2) Adjust length of check cable to provide a ±0- to ±5-degree ski incidence angle (reference Figures 5-4 and 5-6). (3) Adjust length of safety cable to provide a ±15-degree ski incidence angle (reference Figures 5-5 and 5-6). FIGURE 5-4. MAIN SKI AT MAXIMUM POSITIVE INCIDENCE (CHECK CABLE TAUT) Par 504 Page 43 AC 43.13-2B 3/3/08 FIGURE 5-5. MAIN SKI AT MAXIMUM NEGATIVE INCIDENCE (SAFETY CABLE TAUT) FIGURE 5-6. MAIN SKI INCIDENCE ANGLES c. Tension Required in Main Ski Shock Cords. (1) Apply sufficient shock cord tension to the forward ends of the skis to prevent flutter and "dumping" throughout the range of airspeeds and attitudes at which the airplane will operate on skis. Because of the various angles used in attaching the shock cord to the skis, shock cord tension cannot be specified. It is possible to specify the downward force that must be applied to the forward end of the ski in order to overcome the shock cord tension and cause the check cable to slacken when the ski is in the normal flight attitude. That downward force is commonly referred to as the shock cord tension force, or simply the tension force. In most installations on rigid, truss type landing gear, the tension force should be approximately as listed in Table 5-2. TABLE 5-2. APPROXIMATE MAIN SKI TENSION FORCES Ski Limit Load Capacity Downward Force (pounds) 1500-3000 3000-5000 5000-7000 7000-9000 20-40 40-60 60-120 120-200 Par 505 Page 44 3/3/08 AC 43.13-2B NOTE: Do not rely upon these tension force values for main ski installations on airplanes with spring steel or other flexible landing gear. Shock cord tensions great enough to require the downward forces listed in Table 5-2 to overcome them may produce excessive toe-in of the main skis on such airplanes. Variations in gear leg flexibility make it difficult to establish a generic table of tension forces appropriate for all airplanes with flexible landing gear. (2) The shock cord tension must also be sufficient to return the skis to the normal flight attitude from their maximum negative incidence at all airspeeds up to the airplane's never-exceed speed with skis installed. In the absence of more precise data, each shock cord must be able to produce a nose-up moment about the ski pedestal bearing centerline of M = (0.0000036)(W)(VNE)2 ft•lbs, when the ski is at its maximum negative incidence, where W is the maximum certificated gross weight of the airplane and VNE is its never-exceed speed with skis installed. d. Springs in Place of Shock Cords. If springs are used in place of shock cords to provide rigging tension, they must be able to withstand extreme cold and slight external scratching without premature fatigue failure, and must not cause skis, rigging, or landing gear to experience flutter or objectionable vibration during an airplane flight and dive tests. e. Nose Ski Installation. Install the nose ski on an airplane with tricycle landing gear in the same manner as the main skis (see Figure 5-7), except: (1) Adjust length of safety cable to provide ±5- to ±15-degree ski incidence. (2) Where it is possible for the nose ski rigging to contact the propeller tips due to vibration, install a 1/4-inch shock cord to hold the rigging out of the propeller arc. FIGURE 5-7. TYPICAL NOSE SKI INSTALLATION Par 505 Page 45 AC 43.13-2B 3/3/08 f. Tail Ski Installation. (1) When installing a tail ski on an airplane with conventional landing gear, use a tail ski that has been approved on an airplane of approximately the same weight and whose tail wheel bears approximately the same fraction of that weight when the airplane is in the three-point attitude (within 10 percent), or select the tail ski as outlined in paragraphs 501 and 503. Some types of tail ski require that the tail wheel be removed to install the rest on its upper surface “ski.” (2) Adjust the length of the limiting cable (reference Figure 5-8) to allow the ski to turn approximately 35 degrees either side of the straight- forward position with the weight of the airplane resting on the ski. (3) The shock cord (reference Figure 5-8) must be of a length that will hold the ski in the straight-forward position during flight. FIGURE 5-8. TYPICAL TAIL SKI INSTALLATION 506. DOCUMENTATION. a. Airplane or Ski Manufacturer's Data. Comply with the requirements for placards, markings, and manuals required to operate the airplane as a skiplane, as listed in the approved or accepted documents discussed in paragraph 503a. b. Performance Information. The following Paragraphs contain the minimum additional performance data required for airplanes equipped with new or altered ski installations. Consult AC 43-210 (current edition) chapter 4, for additional guidance regarding approved Aircraft Flight Manual (AFM) supplements. (1) For an airplane that requires an approved AFM, obtain FAA approval for an AFM supplement that adds the following or similar information to the Performance section of the Manual. (a) Takeoff. Under the most favorable conditions of smoothly 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 distance for other conditions, caution should be exercised as lower temperatures or other snow conditions will usually increase these distances. (b) Landing. Under the most favorable conditions of smoothly packed snow at temperatures approximately 32° F, the skiplane Par 505 Page 46 3/3/08 AC 43.13-2B landing distance is approximately 20 percent greater than that shown for the landplane. NOTE: 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. (c) 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 AFMS. For larger variations in climb performance, where the minimum requirements are critical, or the landing gear of the landplane is retractable, appropriate climb data should be obtained to determine the changes and new curves, tables, or a note should be incorporated into the AFMS. (2) For an airplane that does not require an AFM, make the information in paragraph 506b(1) available to the pilot in for form of placards, markings, manuals, or any combination thereof. One type of acceptable manual is an approved Supplementary AFM. 507. FLIGHT AND HANDLING OPERA- TIONAL CHECKS. Accomplish an operational check in accordance with Title 14 of the Code of Federal Regulations (14 CFR) part 91, § 91.407(b), to determine the takeoff, landing, and ground handling characteristics. Ensure that the ski angles during tail high and tail low landings will not cause the skis to dig in or fail from localized stress. Verify that ground control is adequate to satisfactorily complete a landing run with a turnoff at slow speed. In flight, the skis must ride steady with check cables taut, and must not produce excessive drag or unsatisfactory flight characteristics. Enter a notation of this check in the airplane records. 508. MAINTENANCE (INCLUDING IN- SPECTION). a. Inspection and Repair Data Sources. Contact the airplane and ski manufacturers for any specific inspection and maintenance instructions they may have developed. Refer to AC 43.13-1 (as amended), chapter 9, for more information. b. Instructions for Continued Air- worthiness (ICA). The modifier (developer of the ski installation or alteration) must provide instructions for future inspection, maintenance, and repair of the added or altered parts, and is also responsible for assessing the need for any changes to the product-level ICA (changes that affect the airplane as a whole when the skis are installed). For simple airplane/ski combinations where skis of the same model have been approved on similar airplanes with appropriate ICA, it may only be necessary to reference those ICA in the maintenance records of the newly altered airplane and/or in Block 8 of the FAA Form 337 documenting the ski installation. For complex ski installations requiring special considerations, the modifier may need to develop new installation-specific ICA. In either case, the modifier must ensure that adequate and appropriate ICA is available to the skiplane owner or operator. Consult current editions of AC 43-210, chapter 5, and FAA Order 8110.54, for additional guidance regarding ICA. c. Interchanging of Skis and Wheels. A person appropriately authorized by 14 CFR part 43, § 43.3, must perform a new weight and balance computation when the skis are initially installed. The FAA recommends that the airplane be weighed for this initial computation. After the initial installation, removing the skis and reinstalling the wheels or vice versa is considered a preventive maintenance operation if it does not involve complex assembly operations or a new weight and balance computation (ref. part 43, appendix A, paragraph (b)(4)(c)(18)). NOTE: During subsequent weight and balance changes to the airplane, be sure to update its weight and Par 506 Page 47 AC 43.13-2B 3/3/08 Par 508 Page 48 balance records and its equipment list to account for all approved ski, wheel, and float installations. d. Periodic Inspection Required. Seasonally removed and installed equipment items such as skis should be inspected at installation to comply with §§ 91.407(a) and 91.409(a), and part 43, appendix D, paragraphs (e)(1) and (e)(10), if they were not installed on the airplane at the time of the last inspection. All available data described in this paragraph should be used during the inspection. 509. THRU 599. RESERVED 3/3/08 AC 43.13-2B CHAPTER 6. OXYGEN SYSTEM INSTALLATIONS IN NONPRESSURIZED AIRCRAFT SECTION 1. GENERAL 600. PURPOSE. This chapter provides data for acceptable means of gaseous oxygen system installations in nonpressurized aircraft. For other oxygen system installations (i.e., liquid oxygen), installers should contact their local Flight Standard District Office (FSDO) for assistance in applying for a Supplemental Type Certificate (STC). 601. HAZARDS AND WARNINGS TO CONSIDER WHEN INSTALLING AN OXYGEN SYSTEM. a. Oxygen itself does not burn, but materials that burn in air will burn much hotter and more vigorously in an oxygen rich environment. b. Oil and grease burn with explosive violence in the presence of oxygen. c. Rapid release of high-pressure oxygen in the presence of foreign particles can cause temperatures sufficient to ignite combustible materials and materials that would not normally burn in air. d. Pressurized oxygen cylinder failures, particularly aluminum-lined composite cylinders, have the potential of producing violent explosions. 602. ADDITIONAL REFERENCES (current editions). a. Title 14 of the Code of Federal Regulations (14 CFR) part 23, Airworthiness Standards: Normal, Utility, Acrobatic, and Commuter Category Airplanes. b. Title 14 CFR part 43, Maintenance, Preventive Maintenance, Rebuilding, and Alteration. c. Title 14 CFR part 91, General Operating and Flight Rules. d. Civil Aviation Regulations (CAR) 3, Airplane Airworthiness; Normal, Utility, and Acrobatic Categories. e. CAR 6, Rotorcraft Airworthiness; Normal Category. f. Advisory Circular (AC) 27-1, Certification of Normal Category Rotorcraft. g. AC 43.13-1, Acceptable Methods, Techniques, and Practices―Aircraft Inspection and Repair. h. Society of Automotive Engineers Aerospace Information Report (SAE AIR) No. 825B, Oxygen Equipment for Aircraft and SAE AIR 822, Oxygen Systems for General Aviation. i. Handbook Bulletin for Airworthiness (HBAW) 02-01B, Maintenance of Pressure Cylinders in Use as Aircraft Equipment. j. FAA Order 8310.6, Airworthiness Compliance Check Sheet Handbook. 603. THRU 606. RESERVED Par 600 Page 49 (and 50) 3/3/08 AC 43.13-2B Par 607 Page 51 SECTION 2. INSTALLATION OF THE OXYGEN SYSTEM 607. SYSTEM REQUIREMENTS. Gaseous oxygen systems may be a higher pressure system with the oxygen stored at 1850 psi or a low pressure system with the oxygen stored at 425 psi. All oxygen systems contain a storage tank, a regulation system, and a distribution system. The main difference in system type is in the regulation of the oxygen to the user (re: paragraph f below). a. Cylinders. Install oxygen cylinders conforming to Interstate Commerce Commission (ICC) requirements for gas cylinders which carry the ICC or DOT 3A, 3AA, or 3HT designation followed by the service pressure metal-stamp on the cylinder. b. Tubing/Lines. (1) In systems having low pressure, use seamless aluminum alloy or equivalent having an outside diameter of 5/16 inch and a wall thickness of 0.035". Double flare the ends to attach to fittings. (2) In high-pressure systems (1800 psi), use 3/16-inch O.D., 0.035" wall thickness, seamless copper alloy tubing meeting Specification WWT-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 side). NOTE: Any lines that pass through potential fire zones should be stainless steel. NOTE: If lines are located behind upholstery or not 100 percent visible during normal operations, they should be solid metal lines or high- pressure flexible lines. c. Fittings. All fittings must be manufactured from materials that are compatible for use with oxygen systems. Fittings should not be made of mild steel or materials that are prone to corrosion when in contact with another material. (1) High Pressure. Intercylinder connections are made with regular flared or flareless tube fittings with stainless steel. Usually fittings are of the same material as the lines. Mild steel or aluminum alloy fittings with stainless steel lines are discouraged. Titanium fittings should never be used because of a possible chemical reaction and resulting fire. (2) Low Pressure. Fittings for metallic low-pressu