U.S. Department of Transportation Federal Aviation Administration Advisory Circular
Subject: Acceptable Methods, Techniques, and Practices – Aircraft Alterations Date: 3/3/08 Initiated by: AFS-300 AC No: 43.13-2B
1. PURPOSE. This advisory circular (AC) contains methods, techniques, and practices acceptable to the Administrator for the inspection and alteration on non-pressurized areas of civil aircraft of 12,500 lbs gross weight or less. This AC is for use by mechanics, repair stations, and other certificated entities. This data generally pertains to minor alterations; however, the alteration data herein may be used as approved data for major alterations when the AC chapter, page, and paragraph are listed in block 8 of FAA Form 337 when the user has determined that it is:
a. Appropriate to the product being altered,
b. Directly applicable to the alteration being made, and
c. Not contrary to manufacturer’s data.
2. CANCELLATION. AC 43.13-2A, Acceptable Methods, Techniques, and Practices― Aircraft Alterations, dated January 1, 1977, is canceled.
3. REFERENCE. Title 14 of the Code of Federal Regulations (14 CFR) part 43, § 43.13(a) states that each person performing maintenance, alteration, or preventive maintenance on an aircraft, engine, propeller, or appliance must use the methods, techniques, and practices prescribed in the current manufacturer’s maintenance manual or Instructions for Continued Airworthiness prepared by its manufacturer, or other methods, techniques, or practices acceptable to the Administrator, except as noted in § 43.16. FAA inspectors are prepared to answer questions that may arise in this regard. Persons engaged in the inspection and alteration of civil aircraft should be familiar with 14 CFR part 43, Maintenance, Preventive Maintenance, Rebuilding, and Alterations, and part 65, subparts A, D, and E of Certification: Airmen Other than Flight Crewmembers, and applicable airworthiness requirements under which the aircraft was type-certificated.
4. COMMENTS INVITED. Comments regarding this AC should be directed to DOT/FAA:
ATTN: Aircraft Maintenance Division, 800 Independence Ave., SW., Washington, DC 20591, FAX (202) 267-5115.
ORIGINAL SIGNED By
James J. Ballough Director Flight Standards Service
iii (and iv)
CONTENTS Paragraph Page
CHAPTER 1. STRUCTURAL DATA1
CHAPTER 2. COMMUNICATION, N AVIGATION, AND EMERGENCY
LOCATOR TRANSMITTER SYSTEM INSTALLATIONS9
CHAPTER 3. ANTE NNA INSTALLATION23
CHAPTER 4. ANTICOLLISION AND SUPPLEMENTARY LIGHT
INSTALLATION33
CHAPTER 5. SKI INSTALLATIONS39
CHAPTER 6. OXYGEN SYSTEM INSTAL LATIONS IN NONPRESSURIZED
AIRCRAFT49
Section 1. General49
Section 2. Installation of the Oxygen system51
Section 3. Airworthiness Compliance Check Sheet: Oxygen System Installation
in Un-pressurized Aircraft59
CHAPTER 7. ROTORCRAFT EXTERNAL -LOAD-DEVICE INSTALLATIONS
CARGO SLINGS AND EXTERNAL RACKS61
Section 1. General61
Section 2. Cargo Racks69
CHAPTER 8. GLIDER AND BANNE R TOW-HITCH INSTALLATIONS73
CHAPTER 9. SHOULDER HARNESS INSTALLATIONS85
Section 1. General83
Section 2. Geometry and Attachment89
Section 3. Static Strength and Testing101
Section 4. Installation and Inspection Checklists105
CHAPTER 10. AIRCRAFT BA TTERY INSTALLATIONS107
Section 1. General107
Section 2. Lead Acid Battery Installations111
Section 3. Nickel-Cadmium Battery Installations115
Section 4. Battery In stallation Checklist119
Section 5. Instructions fo r Continued Airworthiness121
CHAPTER 11. ADDING OR RELOCATING INSTRUMENTS123
CHAPTER 12. CARGO TIEDOWN DEVICE INSTALLATIONS133
CHAPTER 1. STRUCTURAL DATA
100. GENERAL. Structural integrity is a major factor in aircraft design and construction. Addition or removal of equipment involving changes in weight could affect the structural integrity, weight, balance, flight characteristics, reliability, or performance of an aircraft. This chapter is generic in nature and meant to assist the aviation maintenance technician in determining structural integrity. It is not meant to circumvent utilizing a Federal Aviation 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
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
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
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.
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 nonfriction locking device is used in addition to the selflocking 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
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
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.
CHAPTER 2. COMMUNICATION, NAVIGATION, AND EMERGENCY
LOCATOR TRANSMITTER SYSTEM INSTALLATIONS
200. PURPOSE. This chapter describes installation considerations and requirements for basic standalone, 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 in tensity 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 af ter 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
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 su pport 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
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 airc raft 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