______________________________________________ AdvisoryU.S. Department of Transportation Federal Aviation Administration Circular Subject: Fatigue Management Programs for Date: 10/1/12 AC No: 91-82A In-Service Issues Initiated by: ACE-100 Change: 1
1. PURPOSE. This change revises existing pa ragraph numbering throughout the entire document, excluding attachments. The change number and the date of the changed material are shown at the top of each changed page. Vertical bars in the margin indicate the changed material. Pages having no changes retain the same heading information.
2. PRINCIPAL CHANGES. Paragraph 3 has become bullet six in paragraph 2b. All remaining main paragraph numbers have changed accordingly. No content changes have been made. A vertical change bar appears next to the principle change, paragraph 2b, bullet six. Additionally, a spelling correction was made on the x-axis title in figure 4-2 in appendix 4.
PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated 2-3 8/23/11 2-3 10/1/12 5-7 8/23/11 5-7 10/1/12 10-11 8/23/11 10-11 10/1/12 24 8/23/11 24 10/1/12 26 8/23/11 26 10/1/12 28 8/23/11 28 10/1/12 A4-3 8/23/11 A4-3 10/1/12 s/ Earl Lawrence Manager, Small Airplane Directorate Aircraft Certification Service
AdvisoryU.S. Department of Transportation Federal Aviation Administration Circular Subject: Fatigue Management Programs for Date: 10/1/12 AC No: 91-82A In-Service Issues Initiated by: ACE-100 Change: 1
1. Purpose.
a. This advisory circular (AC) provides guidance on developing and implementing a Fatigue Management Program (FMP) to address in-service issues for metallic fatigue critical structure. An applicant may develop an FMP as one method to address an unsafe condition when the Federal Aviation Administration (FAA) determines an airplane type design has a demonstrated risk of catastrophic failure due to fatigue. In such cases, the FMP should incorporate damage-tolerance based inspections or a part replacement/modification program to mitigate the demonstrated risk. The FMP should also incorporate other fatigue critical structure inspections to address the broader risk posed by potential cracking of these structures in the airplane. The FAA will mandate the FMP by Airworthiness Directive (AD). The FAA may also approve the FMP as an alternative method of compliance (AMOC) to an AD.
b. This AC includes guidance relevant to developing FMPs for other purposes such as life extensions, type certification requirements, or non-mandatory maintenance programs. This guidance supplements other ACs that contain guidance for developing damage-tolerance based inspection programs to look proactively for potential cracks. Such guidance includes AC 91-56B, Continuing Structural Integrity Program for Airplanes, AC 25.571-1D, Damage Tolerance and Fatigue Evaluation of Structure, and AC 23-13A, Fatigue, Fail-Safe, and Damage- Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Category Airplanes. Applicants should use prod uct specific guidance in conjunction with this AC.
c. This AC is not mandatory and does not constitute a regulation. It describes an acceptable means, but not the only means, for maintaining the continued operational safety for airplane type designs that have a demonstrated risk. In this AC, the FAA uses terms such as “must” or “require” only in the sense of ensuring applicability of a particular method of compliance when using a specific acceptable method of compliance described herein.
2. Applicability.
a. This AC is applicable to:
• Small airplanes certificated under part 23 or predecessor regulations,
• Transport category airplanes certificated to part 25 or predecessor regulations, and
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• Airplanes certificated in the primary and restricted categories.
b. The following may use this AC to develop FMPs:
• DAHs,
• Applicants for STCs,
• Applicants for AMOC to ADs,
• Applicants for PMA,
• FAA aircraft certification engineers, and
• FAA Designated Engineering Representatives (DER) or delegated organizations.
3. Cancellation. This AC cancels AC 91-82, Fatigue Management Programs for Airplanes with Demonstrated Risk of Catastrophic Failure Due to Fatigue, dated April 29, 2008.
4. Related Regulations. Refer to the following Title 14 of the Code of Federal Regulations (14 CFR) sections, as applicable.
a. Design approval holder responsibilities for reporting of failures, malfunctions and defects included in part 21, § 21.3.
b. Certification procedures for instructions for continued airworthiness (ICA) and airworthiness limitations included in part 21, § 21.50.
c. Certification procedures for changes to type designs included in part 21, § 21.93.
d. Requirements for Supplemental Type Certificates (STC) included in part 21, § 21.113.
e. Small airplane requirements for strength and deformation included in part 23, § 23.305.
f. Small airplane requirements for fatigue, fail-safe, and damage-tolerance evaluations included in part 23, §§ 23.571, 23.572, 23.573, 23.574, and 23.627.
g. Small airplane requirements for inspections and ICAs included in part 23, §§ 23.575,
23.611, and 23.1529.
h. Transport category airplane requirements for strength and deformation included in part 25, § 25.305.
i. Transport category airplane requirements for fatigue and damage-tolerance evaluations included in part 25, § 25.571.
j. Transport category airplane requirements for inspections and ICAs included in part 25, §§ 25.611 and 25.1529.
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k. AD requirements included in part 39.
5. Definitions. The following definitions apply for this AC:
a. Accident. An occurrence associated with th e operation of an aircraft, which takes place between the time any person boards the aircraft with the intention of flight and all such persons have disembarked, and in which any person suffers death or serious injury, or in which the aircraft receives substantial damage.
b. Applicant. An applicant is any interested party who is developing a fatigue management program.
c. Damage-Tolerance Based Inspection. An inspection based on consideration of the crack growth and residual strength characteristics of the structure, the physical access to the structure, and the inspection method reliability. The inspection should provide a high probability of detecting fatigue damage before the residual strength degrades below a specified value.
d. Demonstrated Risk.
(1) An airplane type design has a “demonstrated risk” when an unsafe condition due to fatigue exists in that type design and the condition is likely to exist or develop in other airplanes of the same or similar type design. The FAA will determine the necessary action to address a fatigue cracking scenario using the corrective action review board process in accordance with FAA Order 8110.107.
(2) Situations that could precipitate a risk analysis may include:
• An airplane has experienced a catastrophic failure due to fatigue and the same scenario is likely to occur on other airplanes in the fleet,
• Airplanes of the type design have a service history that indicates a significant likelihood of catastrophic failure due to fatigue in the fleet,
• Fatigue testing of the type design indicates a significant likelihood of catastrophic failure due to fatigue in the fleet, or
• The type design has a structural area sufficiently similar to another type design’s structural area determined to present a demonstrated risk.
e. Design Approval Holder (DAH). For the pur poses of this AC, a holder of a type certificate (TC), STC, or Part Manufacturing Approval (PMA) issued under part 21.
f. Fatigue Critical Structure. For purposes of this AC, structure that is susceptible to fatigue cracking that could contribute to catastrophic failure of an airplane if there is no intervention.
This would typically include all structure critical to carrying flight, ground or pressurization loads that are subjected to tension dominated repeated loads during operation.
Examples of fatigue critical structures include, but are not limited to:
• The fuselage pressure shell consisting of all skin panels, stringers, frames and pressure bulkheads including skins and stiffening members;
• The lower wing cover consisting of all skin panels, stringers and lower spar caps;
and
• The horizontal stabilizer upper cover consisting of all skin panels, stringers, and upper spar caps.
(1) For airplanes designed to high negative g loads (such as acrobatic category), joints, splices and maximum compression bending moment locations on the wing and horizontal tails may also be fatigue critical.
(2) For some designs, the loss of control surface and flap hinge fittings, including their attach structure, may be catastrophic under certain conditions. Unless documentation exists to show these areas are not critical to safe flight, they should be considered fatigue critical structure if they are subject to tension dominated repeated loads during operation. (Documentation such as that normally provided at time of certification may be sufficient.)
g. Fatigue Management Program (FMP). An FMP is a set of maintenance actions whose purpose is to prevent catastrophic failure due to fatigue. The maintenance actions may include fatigue critical structure inspection, modification, or replacement. Component or airframe safelife is also considered an FMP.
h. Fleet. For the purposes of this AC, a fleet is all airplanes of the same type design.
i. Incident. For the purposes of this AC, an incident is an occurrence other than an accident, associated with the operation of an aircraft, which affects or could affect the safety of operations. This includes in-service findings reported according to § 21.3.
j. Residual Strength. For the purposes of this AC, the capability of structure to carry static load with fatigue damage present.
k. Service History Based Inspection. An in spection based on qualitative information derived from experience with the same or similar structure. Although directed at a specific area of a fatigue critical structure, the applicant does not typically quantify or validate the inspection’s effectiveness.
l. Safe-Life. The number of events , such as flights, landings, or flight hours time in service, during which there is a low probability the strength will degrade below its design ultimate value due to fatigue.
m. Unsafe Condition. For the purposes of this AC, an unsafe condition is fatigue cracking that either resulted in a catastrophic failure or without intervention could directly result in, or indirectly lead to, a catastrophic failure.
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6. Related Publications (current editions).
a. ACs.
(1) AC 21-40A, Application Guide for Obtaining a Supplemental Type Certificate, dated September 27, 2007.
(2) AC 23-13A, Fatigue, Fail-Safe, and Damage-Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Category Airplanes, dated September 29, 2005.
(3) AC 25.571-1D, Damage Tolerance and Fatigue Evaluation of Structure, dated January 13, 2011.
(4) AC 43-4A, Corrosion Control for Aircraft, dated July 25, 1991.
(5) AC 43.13-1B, Acceptable Methods, Techniques, and Practices – Aircraft Inspection and Repair, dated September 27, 2001.
(6) AC 91-56B, Continuing Structural Integrity Program for Airplanes, dated March 7, 2008.
b. Orders.
(1) FAA Order 8040.1C, Airworthiness Directives, dated October 3, 2007.
(2) FAA Order FAA-IR-M 8040.1C, Airworthiness Directives Manual, dated May 17, 2010.
(3) FAA Order 8110.4C CHG 3, Type Certification, dated March 15, 2010.
(4) FAA Order 8110.54A, Instructions for Continued Airworthiness: Responsibilities, Requirements, and Contents, dated October 23, 2010.
(5) FAA Order 8110.103A, Alternative Methods of Compliance (AMOC), dated September 28, 2010.
(6) FAA Order 8110.107, Monitor Safety/Analyze Data, dated March 12, 2010.
(7) FAA Order 8900.1, Flight Standards Information Management System (FSIMS), dated September 13, 2007.
c. Others.
(1) Best Practices Guide for Maintaining Aging General Aviation Airplanes, Appendix 1, Aging Airplane Inspection & Maintenance Baseline Checklist for Airplanes Without a Type Specific Checklist, dated September 2003. An electronic version is available at http://www.faa.gov/aircraft/air_cert/design_approvals/small_airplanes/cos/aging_aircraft/media/ aging_aircraft_best_practices.pdf.
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(2) Commercial Airplane Certification Process Study, An Evaluation of Selected Aircraft Certification, Operations, and Maintenance Processes, FAA, March 2002.
(3) DOT/FAA/CT-93/69.1 and DOT/FAA/CT-93/69.2, Damage Tolerance Assessment Handbook, Volumes 1 and 2, dated October 1993.
(4) FAA and Industry Guide to Product Certification, Second Edition, dated September 2004.
(5) General Aviation Manufacturers Association’s Specification No. 2, Maintenance Manual.
(6) Nondestructive Evaluation (NDE) Capabilities Data Book, W.D. Rummel and G.A.
Matzkanin, Nondestructive Testing Information Analysis Center (NTIAC), Texas Research Institute, Austin, Texas.
7. Background.
a. In the early 1980s, because of concerns about the continued airworthiness of older airplanes certified to the Civil Air Regulation (CAR) 4b fail-safe requirements, the FAA began issuing ADs mandating damage-tolerance based structural inspection programs. The intent of the programs mandated by these ADs was to prevent unacceptable degradation of the structural integrity of the affected airplanes to assure long-term continued operational safety of the fleet.
These programs proactively addressed potential fatigue that could develop into a demonstrated risk. Per the definition in this AC , these programs are considered FMPs.
b. There have been airframe fatigue-related accidents and incidents that have required certain follow-on actions to maintain the continued operational safety of the fleet. Except for the requirement to report failures, malfunctions, and defects in § 21.3, no specific regulation or guidance exists to assist an applicant on what actions to take following a catastrophic failure due to fatigue or an in-service finding of fatigue cracking. Consequently, the FAA has worked on a case-by-case basis with applicants, owners, and operators to determine the actions needed to maintain the continued operational safety of these airplanes.
c. Reactively, to address a known unsafe condition that is likely to occur on other airplanes of the same type design, the FAA has mandated actions that deal with the specific unsafe condition. These actions typically include inspections, repair, or replacement of specific fatigue critical structural parts. However, in-service experience and fatigue test data have shown that actions to address one unsafe condition and extend the operation of a model fleet allowed a second unsafe condition to develop. Therefore, occu rrence of fatigue in one critical part may be a precursor to fatigue in other critical structure that should be addressed proactively.
d. Based on the above, the FAA developed this guidance to assist an applicant on what actions to take following a catastrophic failure due to fatigue or an in-service finding of fatigue cracking that poses a demonstrated risk. The FAA determined that any proposed inspection of the structural elements directly related to the unsafe condition should be based on damage tolerance principles. The FAA also determined that, at a minimum, service history based inspections should be established for other areas considered fatigue critical.
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8. Applicability of Certification Requirements to Address a Demonstrated Risk. The guidance in this AC describes how applicants may apply concepts contained in the metallic structure fatigue requirements in 14 CFR to address a demonstrated risk. The requirements include several approaches1 for preventing catastrophic failure due to fatigue in new type designs. These approaches may also be used to address a demonstrated risk. To apply any of these approaches successfully, applicants should have demonstrated knowledge, abilities, skills, and understand:
• The intent of the 14 CFR fatigue and damage tolerance requirements,
• The underlying philosophy of the fatigue prevention approaches, and
• How to apply these approaches to address a specific demonstrated risk.
For airplanes that have experienced a demonstrated risk, the FAA considers compliance to the latest fatigue requirements will materially contribute to the product level of safety. Therefore, normal, utility, acrobatic, or primary category airplanes with a type certification basis that does not include fatigue requirements should comply with the latest fatigue requirements amendment for any replacement or modified structure. Airplanes with a type certification basis that includes fatigue requirements must comply with those fatigue requirements for any replacement or modified structure. The applicant must calibrate a safe-life or inspection program to the service history that resulted in the demonstrated risk. For a fail-safe design, an applicant must address the challenges associated with achieving a fail-safe design, as described in AC 23-13A, Chapter
3. For commuter and transport category airplanes whose certification basis includes damagetolerance, any replacement or modified structure must comply with the applicable damagetolerance requirements. For restricted category airplanes, any replacement or modified structure must comply with the airplane category used for type certification under § 21.25.
a. Safe-Life. For airplanes certificated using the safe-life approach, the safe-life of a structure is the number of events, such as flights, landings, or flight hours time in service during which there is a low probability the strength will degrade below its design ultimate value due to fatigue. The safe-life is a point in the airplane ’s operational life when the operator must replace, modify, or take the structure out of service to prevent it from developing fatigue cracks that can degrade the strength below its design ultimate value.
(1) An applicant can use the safe-life approach to address a demonstrated risk by establishing a point in the operational life of the airplane when the operator must replace, 1 For new type designs constructed with metallic structure seeking certification in the normal, utility, and acrobatic categories, 14 CFR part 23 allows the applicant to choose between the safe-life, damage-tolerance, and fail-safe design approaches. For new type designs in the commuter category or designs using composite materials, part 23 requires the applicant to use damage-tolerance, unless the applicant shows that damage-tolerance is impractical. For transport category airplanes, 14 CFR part 25 requires the use of damage-tolerance, unless the applicant shows that damage-tolerance is impractical.
modify, or take the structure out of service. If a catastrophic failure due to fatigue has occurred, or if an inspection finds cracks in-service, the safe-life of that structure likely has already passed.
Based on service history, an applicant could establish a retroactive safe-life for the structure.
Airplanes below this service history based safe-life could continue to operate until they reach the retroactive safe-life. For airplanes already at or beyond the service history based safe-life, the applicant could immediately replace or modify the structure or establish a schedule for completing these actions acceptable to the FAA. To manage the risk, the FAA coordinates any schedule with the applicant and considers:
• The number of airplanes used in commercial service,
• The risk of allowing operation beyond the service history based safe-life,
• The effectiveness of inspecting the structure,
• The residual strength characteristics of the structure,
• The availability of replacement parts, and
• Other relevant factors.
b. Damage-Tolerance. For a new type design, the regulations include the damagetolerance approach for preventing catastrophic failure due to fatigue. The damage-tolerance approach depends on directed inspection programs to detect fatigue damage before strength degrades below the required level.
(1) A damage-tolerance based inspection of fatigue critical structure allows that structure to operate indefinitely until cracks are detected. However, once cracks are detected in one or more aircraft in a fleet, a risk analysis will be performed per the process described in Order 8110.107. This analysis will determine if any additional airworthiness actions (e.g., more rigorous inspection requirements or fleet-wide replacement or modification of the structure) are required. This analysis is necessary because continuing with the previously established inspections could result in an unacceptable probability of aircraft not meeting type design strength requirements as discussed in the following paragraph.
(2) Cracking is a continued airworthiness concern because cracking can potentially reduce the strength of the structure. The loss of ce rtificated ultimate load capability should be a rare event and the FAA does not knowingly allow the strength of airplanes to drop below the certificated ultimate load requirement. However, service history has shown that the reliability of directed inspections is never sufficient to detect all cracks and as the number of crack reports increases, the likelihood that a number of airplanes in the fleet have undetected fatigue cracks also increases. Additionally, most damage tolerance based inspections are developed to ensure the structure always retains approximately design limit load capability. Therefore, for areas where fatigue cracks are reported, the likelihood increases with time that a number of airplanes in the fleet will have strength below the certificated ultimate load requirement. At some time during operation of the fleet, the likelihood that the strength of any given structure in a fleet is less than the certificated ultimate load requirement becomes unacceptably high and inspections 8
must be modified (e.g., shorter intervals) or the problem must be terminated (e.g., fleet wide replacement/modification as a specified time in service).
(3) An applicant can use the damage-tolerance approach to address a demonstrated risk.
However, the applicant must understand that damage-tolerance based inspections may not provide a permanent solution, as explained in subparagraph (2) above, if cracks are expected to continue to develop in the fleet. In this case, the FAA may require the fleet-wide replacement or modification of the structure. The schedule fo r completing these actions for a structure with demonstrated damage-tolerant capability may allow a limited increase in operational life compared to the schedule for a structure with no directed inspection program or demonstrated residual strength capability.
c. Fail-Safe Design. A fail-safe design retains its required residual strength after a readily detectable failure or partial failure of primary structure. A fail-safe design typically consists of redundant structural element(s) and/or crack arrest or crack containment features.
(1) The redundant structure provided by fail-safe design has sometimes prevented catastrophic failures. However, the FAA has doc umented cases in which a design thought to be fail-safe was not, and the design failed to prevent a catastrophic failure because the failure was not readily detected. AC 23-13A and the reference in paragraph 6c(2) identify several potential shortcomings of designs mistakenly thought to be fail-safe.
(2) Applicants can use the fail-safe design concept to evaluate the metallic structures of normal, utility, and acrobatic category airplanes. Prior to 14 CFR parts 25 and 23 Amendments 25-45 and 23-48, respectively, fail-safe design was a fatigue evaluation option for transport category and commuter category airplanes. Fail-safe was removed as a fatigue evaluation option for transport category airplanes certificated to Amendment 25-45 and subsequent, and commuter category airplanes certificated to Amendment 23-48 and subsequent.
(3) Applicants who apply the fail-safe design approach to a demonstrated risk should fully understand and address the potential shortcomings discussed in the documents listed in paragraph 6. An applicant should also understand if cracks are expected to continue to develop in the fleet, which increases the likelihood that the strength of some airplanes’ structure may deteriorate below design ultimate strength level, a fail-safe modification might not provide a solution to the demonstrated risk. In this case, the FAA will require the fleet-wide replacement or modification of the structure. As in the case of damage-tolerance, the schedule for completing these actions for a fail-safe design may allow a limited increase in operational life compared to the schedule for a structure with no demonstrated fail-safe characteristics.
(4) The FAA will require a damage-tolerance based inspection program for any application of fail-safe design to address a demonstrated risk. Inspections are necessary due to the difficulties in achieving a truly fail-safe design, as described in AC 23-13A, Chapter 3. Also, even with added redundant structure, cracks may continue to develop in the fleet. Applicants should determine inspection methods and repeat intervals using damage-tolerance principles.
Initial inspections should take place when the structure is modified with the fail-safe design or at the time in service at which the FAA requires an FMP, whichever occurs later.
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(5) For commuter and transport category airplanes whose certification basis includes damage-tolerance, any fail-safe modification of the structure must comply with the applicable damage-tolerance requirements.
9. Overview of Fatigue Management System. In cases where accidents or incidents involve fatigue cracking, the FAA will act to determine if the type design (or supplemental type design) has a demonstrated risk. If the FAA determines a demonstrated risk exists, the FAA may require maintenance actions to address the risk. An FMP, as described in this AC, is one method applicants may use to address the demonstrated risk. This AC also provides proactive means to address the broader risk, which are inspections for cracks in other fatigue critical structure of airplanes of the same type design.
a. The FAA may require an FMP to meet the risk management guidelines established by the accountable directorate under the FAA Safety Management System (SMS). Appendix 1 contains more information regarding SMS.
b. Figure 1 provides an overview of a fatigue management system. It depicts steps the FAA and applicants should follow when they suspect a risk of catastrophic failure due to fatigue.
Appendix 2 describes in detail the steps the FAA takes to determine if a demonstrated risk exists.
Paragraphs 10, 11, and 12 describe the actions needed once the FAA has determined a demonstrated risk exists.
(1) In some instances, the FAA will mandate initial, short-term actions to provide shortterm mitigation of the demonstrated risk and allow the fleet to return to service. These shortterm actions often include operating limitations or immediate and short-interval inspections. The FAA often mandates these short-term actions with an emergency AD or immediately adopted rule (IAR) AD.
(2) In some instances, applicants may develop other interim actions and gain FAA approval after the FAA has issued an AD. 14 CFR 39.19 and 39.21 permit an AMOC to an AD.
In accordance with those regulations, if the interim actions adequately mitigate the demonstrated risk, the FAA may approve the interim actions as an AMOC to the AD. The aircraft certification office (ACO) responsible for the AD will approve the AMOC. The AMOC approval will stipulate all the specific requirements and limitations of the interim action. When properly coordinated between the ACO, interim action applicant, and appropriate flight standards district office (FSDO), the combination of the AMOC approval and performance of the actions stated in the approval suffices as compliance with the AD portion the AMOC addresses. The Alternative Method of Compliance Order, 8110.103C, provides detailed guidance regarding use and approval of AMOCs.
(3) Appendix 3 contains examples of suitable AD language if the FAA decides to mandate an FMP with an AD.
c. Figure 1 also outlines the basic actions an applicant should take to develop an FMP to address a demonstrated risk. The purpose of an FMP is to prevent future catastrophic failures due to fatigue and to maintain the type design strength and stiffness of the airframe throughout 10
Change 1 the operational life of the airplane. As shown in Figure 1, there are two main components to an FMP to address a demonstrated risk developed in accordance with this AC.
(1) Damage-Tolerance Based Inspections or Replacement/Modification of Structural Elements Directly Related to the Unsafe Condition. Applicants should base this component of the FMP on a comprehensive damage-tolerance or fatigue evaluation, as applicable, and as explained in paragraph 10. The applicant may limit the scope of the detailed evaluation to the structural elements directly related to the unsafe condition.
(2) Inspections of Other Fatigue Critical Structure in the Airplane. This component of an FMP provides for proactive inspections of other fatigue critical structure. These inspections address the broader risk affecting the type design. The inspection requirements may be service history based or damage-tolerance based depending on whether or not an FMP already exists and what it contains.
d. Paragraph 10 describes the FMP development process. Applicants should document the FMP maintenance actions in the ICA.
e. Paragraph 11 explains the approval of an FMP and the coordination needed for those approvals.
f. Paragraph 12 discusses FMP implementation. The operator accomplishes implementation of the FMP by following the instructions prescribed in the ICA.
10. Developing an FMP. This paragraph describes the development of an FMP. The flowchart in Figure 2 outlines the decisions and actions an applicant should follow when developing an FMP.
a. Components of an FMP. Any FMP proposed by an applicant to address a demonstrated risk should include each of the following components:
(1) Damage-Tolerance Based Inspections or Replacement/Modification of the Structural Elements Directly Related to the Unsafe Condition.
(a) An applicant may limit the scope of this component of an FMP to the structural elements directly related to the unsafe condition.
(b) An applicant may propose an inspection or a part replacement/modification program to address the demonstrated risk. An applicant should base any inspection program on a damage-tolerance evaluation of the affected structure. They should base the time in service for a part replacement/modification (effectively a safe-life) on a fatigue analysis or crack growth analysis. Any part replacement/modification program should demonstrate compliance to the applicable regulations. This includes any subsequent inspection requirements. Any proposed program should incorporate the lessons learned from the service history that led to the FAA’s determination of a demonstrated risk.