1-1 CHAPTER 1. GENERAL
1.1 Purpose of This Advisory Circular (AC). This AC contains methods, techniques, and practices acceptable to the Administrator for the inspection and alteration on nonpressurized areas of civil aircraft of 12,500 pounds gross weight or less. 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 Federal Aviation Administration (FAA) Form 337, Major Repair and Alteration (Airframe, Powerplant, Propeller, or Appliance), when the user has determined that it is appropriate to the product being altered, directly applicable to the alteration being made, and not contrary to the manufacturer’s data. This AC contains information pertinent to Title 14 of the Code of Federal Regulations (14 CFR) § 43.13.
Note: This is a guidance document. Its content is not legally binding in its own right and will not be relied upon by the Department as a separate basis for affirmative enforcement action or other administrative penalty.
Conformity with the guidance document is voluntary only. Nonconformity will not affect rights and obligations under existing statutes and regulations.
Note: The methods, techniques, and practices in this AC are an acceptable means of compliance with the regulation, but they are not the only means of compliance.
If persons performing maintenance, alteration, or preventive maintenance undertake another means of compliance, it must be acceptable to the FAA.
References to an established standard or publication support the acceptability of alternative methods, techniques, and practices.
1.2 Audience. This AC is for use by mechanics, repair stations, and other certificated entities.
1.3 Where You Can Find This AC. You can find this AC on the FAA’s website at https://www.faa.gov/regulations_policies/advisory_circulars and the Dynamic Regulatory System (DRS) at https://drs.faa.gov.
1.4 What This AC Cancels. AC 43.13-2B, Acceptable Methods, Techniques, and Practices – Aircraft Alterations, dated March 3, 2008, is canceled.
1.5 Reference. Title 14 CFR § 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, the instructions for continued airworthiness (ICA) prepared by its manufacturer, or other methods, techniques, or practices acceptable to the Administrator except as noted in 14 CFR § 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 and 14 CFR part 65 subparts A, D, and E and applicable airworthiness requirements under which the aircraft was type certificated.
Note: Part 23 Amendment 23-63 and prior 14 CFR part 23 (prescriptive) regulations referenced in this AC contain sufficiently detailed language to allow
1-2 comp liance directly with each section (regulation) when applied to traditional designs that were prevalent at the time the regulations were written. These legacy prescriptive regulations are not sufficient nor appropriate for some modern designs. To show compliance with the new “performance-based rules” (Amendment 23-64, Implementation Procedures Guide, and subsequent), the applicant must choose an accepted means of compliance, per 14 CFR § 23.2010.
For more information, refer to Amendment 23-64 and AC 23.2010, F AA Accepted Means of Compliance Process for 14 CFR Part 23.
1.6 AC Feedback Form. For your convenience, the AC Feedback Form is the last page of this AC. Note any deficiencies found, clarifications needed, or suggested improvements regarding the contents of this AC on the Feedback Form.
2-1 CHAPTER 2. STRUCTURAL DATA
2.1 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 an FAA engineer or the responsible Aircraft Certification Service office when necessary.
2.2 Structural Design Process. Structural design processes follow these steps:
• Determine the overall load factors.
• Estimate the resulting loads.
• Distribute these loads over the aircraft.
• Determine the material, size, and shape of the part.
• Calculate the resulting stresses in the part.
• Compare these stresses with the maximum allowable for the material used.
• Resize the part as necessary.
2.3 Types of Loads and Stresses.
2.3.1 “Limit load factors” are the maximum load factors that may be expected during service (the maneuvering, gust, or ground load factors established by the manufacturer for type certification).
2.3.2 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.
2.3.3 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.
2.3.4 “Tension” is a force acting against another force that is trying to pull something apart.
2.3.5 “Compression” is a squeezing or crushing force that tries to make parts smaller.
2.3.6 “Torsion” is a twisting force.
2.3.7 “Shear stress” is when one piece of material slides over another.
2.3.8 “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.
2-2 Figure 2-1. Bending of a Beam
2.3.9 A n 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 will have to endure and build the structure strong enough to withstand these stresses.
2.4 Static Loads. “Static loads” are loads that do not undergo change in magnitude or direction during a measurement procedure. Load factors are defined as follows:
2.4.1 “Limit load factors” are the maximum load factors that may be expected during service (the maneuvering, gust, or ground load factors established for type certification).
2.4.2 “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.
2.4.3 “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.
2.4.4 “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).
2-3 Figure 2-2. Typical Load
2.5 Structural Sizing and Analysis. Design and size your load structures, including wing spars, wing attach fittings, stabilizers, landing gear struts, etc.
2.5.1 Static tests using the load factors in Table 2-1 are acceptable for equipment installations.
The alteration needs to comply with the limit load factors as required by the aircraft’s certification basis.
Table 2-1. Ultimate 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 1.5g 1.5g 4.5g 8.0g Upward 3.0g 4.5g 3.0g 4.0g Forward 9.0g 9.0g 18.0g 16.0g Downward 6.6g 9.0g - 20.0g Rearward - - - 1.5g Note 1: When equipment mounting is located externally to one side, or forward of occupants, a forward load factor of 2.0g is sufficient.
Note 2: 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 the 14 CFR part 23 utility category are acceptable for aircraft that never exceed the speed of 250 knots and the factors used for 14 CFR part 23 acrobatic category.
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2.5.2 Table 2-2 is an example of determining the static test loads for a 7-pound piece of equipment to be installed in a utility category aircraft (14 CFR part 23).
Table 2-2. 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 6.6g 46.2 pounds Forward 9.0g 63.0 pounds
2.5.3 When an additional load is to be added to a 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, the responsible Aircraft Certification Service office, or an appropriately rated Designated Engineering Representative (DER).
2.6 Static Tests.
CAUTION: The aircraft or equipment, or both, 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 that 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 that 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:
2.6.1 Determine the weight and center of gravity (CG) position of the equipment item.
2.6.2 Install the attachment in the aircraft or, preferably, in a jig using the applicable static test load factors.
2.6.3 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 2-3, Hypothetical of Determining Static Test Loads.
1. Convert the gust, maneuvering, and ground load factors obtained from the manufacturer, FAA engineer, or DER to determine the ultimate load factors. Unless otherwise specified in the airworthiness standards applicable to the aircraft, ultimate
2-5 lo ad factors are limit load factors multiplied by a 1.5 safety factor (see Figure 2-3 , 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 Figure 2-3, 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 Figure 2-3, column 6).
2.6.4 Apply a load at the CG position 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 structure and attachments are acceptable. Should permanent deformation occur after 3 seconds, modifications or reinforcements to the affected structure are required. Additional load testing is not necessary.
2.6.5 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.
2.7 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 Orders (TSO), or Military Specifications (MIL-SPEC).
2.7.1 Suitability and durability of materials used for parts, the failure of which could adversely affect safety, must:
• Be established by experience or tests;
• Meet approved specifications that ensure the strength, and other properties, assumed in the design data; and
• Take into account the effects of environmental conditions, such as temperature and humidity, expected in service.
2.7.2 Workmanship must be of a high standard.
2.8 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 Development and Standardization (MMPDS), available at https://www.mmpds.org/. Design values must be chosen to minimize the probability of structural failure due to material variability. Except as provided in paragraph 2.8.5 below, compliance with this paragraph must be shown by selecting design values that ensure material strength with the following probabilities:
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2.8.1 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.
2.8.2 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.
2.8.3 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.
2.8.4 The design of the structure must minimize the probability of catastrophic fatigue failure, particularly at points of stress concentration.
2.8.5 Design values greater than the guaranteed minimums required by paragraph 2.8 may be used where only guaranteed minimum values are normally allowed (e.g., 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).
2.9 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.
2.9.1 Each removable fastener must incorporate two retaining devices in case one retaining device should fail during flight operations.
2.9.2 Fasteners and their locking devices must not be adversely affected by the environmental conditions associated with the particular installation.
2.9.3 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 self-locking device.
2.10 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:
2.10.1 Be suitably protected against deterioration or loss of strength in service due to any cause, including:
• Weathering,
• Corrosion, and
• Abrasion.
2.10.2 Have adequate provisions for ventilation and drainage.
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2.11 Accessibility. Provide adequate provisions to permit close examination of equipment or adjacent parts of the aircraft that regularly require, among other services, inspection, adjustment, and lubrication. 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.
2.12 Effects on Weight and Balance (W&B). Ensure that the altered aircraft can be operated within the weight and CG 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 W&B should be considered and documented in the aircraft W&B.
Determine that the altered aircraft will not exceed maximum gross weight. If applicable, correct the loading schedule to reflect the current loading procedure. Refer to AC 43.13-1, Acceptable Methods, Techniques, and Practices―Aircraft Inspection and Repair, for W&B computation procedures.
2.13 Effects on Safe Operation. Install equipment in a manner that will not interfere with or adversely affect the safe operation of the aircraft (controls, navigation equipment operation, etc.).
2.13.1 The factor of safety prescribed in 14 CFR § 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:
• Uncertain,
• Likely to deteriorate in service before normal replacement, or
• Subject to appreciable variability because of uncertainties in manufacturing processes or inspection methods.
2.13.2 Unless otherwise provided, a factor of safety of 1.5 must be used.
2.14 Controls and Indicators. Locate and identify equipment controls and indicators so they can be operated and read from the appropriate crewmember position.
2.15 Placarding. Label equipment requiring identification and, if necessary, placard operational instructions. Amend W&B 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 (ICA) to ensure that maintenance personnel will know if a required placard is missing in future inspections.
2-8 Figure 2-3. Hypothetical of Determining Static Test Loads Utility Category Aircraft (14 CFR Part 23) Type of Load Direction Load Factors 1 Limit 2 x Safety 3 = Ultimate 4 x Special 5 Static = Test 6 Critical Static Test A. Maneuvering Fwd (None) - - - Down 6.2g 1.5 9.3g - 9.3g 9.3g Side (None) - - - - Up -3.8g 1.5 -5.7g - -5.7g 5.7g Aft 1.0g 1.5 1.5g - 1.5g B. Gust (= 30 FPS @ KVc) *For locations aft of fuselage Sta. 73.85.
Fwd (None) - - - - Down 6.0g 1.5 9.0g - 9.0g Down* 6.4g 1.5 9.6g - 9.0g *9.6g Side 1.6g 1.5 2.4g - 2.4g 2.4g Up -2.8g 1.5 -4.2g - -4.2g Aft (None) - - - - C. Ground Fwd 6.6g 1.5 9.9g - 9.9g 9.9g Down 4.0g 1.5 6.0g - 6.0g D. Ultimate Inertia Forces for Emergency Landing Condition (14 CFR § 23.561).
**For separate cargo compartments.
Fwd Already Prescribed as Ultimate
9.0g - - Fwd** 4.5g - - **4.5g Down (None) - - Side 1.5g - 1.5g Up -3.0g - -3.0g Aft (None) - -
2-9 Utility Category Aircraft (14 CFR Part 23) Type of Load Direction Load Factors 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, and Berth Attachment to Aircraft Structure (14 CFR § 23.785).
Fwd Already Prescribed as Ultimate
9.0g 1.33 12.0g 12.0g Down (None) - - Side 1.5g 1.33 2.0g Up -3.0g 1.33 -4.0g Aft (None) - - Note: Asterisks (*) denote special load conditions for the situation shown.
3-1 CHAPTER 3. COMMUNICATION, NAVIGATION, AND EMERGENCY LOCATOR
TRANSMITTER SYSTEM INSTALLATIONS
3.1 Purpose. This chapter describes installation considerations and requirements for basic standalone, installations of communication, navigation, and emergency locator transmitter (ELT) equipment.
Note: Standalone installations do not depend on other systems or complex interfaces to function.
3.2 Hazards and Warnings.
3.2.1 When installing these systems, follow the aircraft and equipment manufacturers’ instructions, as appropriate. Practice a “clean as you go” philosophy, which entails maintaining cleanliness and order. Ensure that equipment and systems function properly and perform their intended function(s).
3.2.2 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 AC 23.1309-1, System Safety Analysis and Assessment for Part 23 Airplanes, for additional information concerning the evaluation for equipment, systems, and installations in 14 CFR part 23 airplanes.
3.2.3 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.
3.2.4 Consider the impact when Weight and Balance (W&B) 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).
3.2.5 When evaluating elements of impact, consider the before and after states. No further analysis of W&B 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.
3.2.6 When structures must be fabricated or reinforced, the standard practices approved for repairs, if applicable, may be employed. AC 43.13-1, Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair, may provide structural design data for fabrication of mountings and attachments.
3.2.7 Ensure that the capacity of the aircraft’s charging system is not exceeded, including any required additional allowances.