Appendix 2 A2-2 approval, Technical Standard Order (TSO) authorization, letter of TSO design approval, and field approval (FAA Form 337).
Design service goal (DSG)—The period of time (in flight cycles or flight hours, or both) established at design and/or certification during which the principal structure of an airplane will be reasonably free from significant cracking.
Fatigue-critical baseline structure—The part of the baseline structure of an airplane that is classified as fatigue-critical.
Fatigue-critical structure—Structure of an airplane that is susceptible to fatigue cracking that could lead to a catastrophic failure of the airplane. For the purposes of this AC, fatigue-critical structure refers to the same class of structure as the structure that would need to be assessed for compliance with § 25.571(a) at Amendment 25-45 or later. The term fatigue-critical structure may refer to fatigue-critical baseline structure, fatigue-critical alteration structure, or both.
Inspection start point (ISP)—The point in time when special inspections of the fleet are initiated because of a specific probability of having a multiple site damage (MSD)/multiple element damage (MED) condition.
Instructions for Continued Airworthiness (ICA)—Maintenance actions developed by the TC or STC holder in accordance with 14 CFR 25.1529 and delivered with the airplane in accordance with 14 CFR 21.50. ICA are documentation that set forth instructions and requirements for the maintenance that is essential to the continued airworthiness of an aircraft, engine, or propeller.
Limit of validity (of the engineering data that supports the structural maintenance program)—The period of time (in flight cycles, flight hours, or both), up to which test evidence, analysis, and, if available, service experience and teardown inspection results of high-time airplanes support the structural maintenance program.
• For metallic structures, the demonstration must show that widespread fatigue damage (WFD) will not occur before the limit of validity of the engineering data that supports the structural maintenance program (LOV).
• For composite structures, the demonstration must show that any known aging mechanism that degrades the structure below residual strength will not occur before the LOV.
Appendix 2 A2-3 Maintenance actions—Inspections, modifications, replacements, or any combination of these.
Multiple element damage (MED)—A source of WFD characterized by the simultaneous presence of fatigue cracks in similar adjacent structural elements.
Multiple site damage (MSD)—A source of WFD characterized by the simultaneous presence of fatigue cracks in the same structural elements.
Responsible Aircraft Certification Service office—The Aircraft Certification Service office having oversight responsibility for the relevant TC or STC, as determined by the Administrator.
Responsible Flight Standards Service offices—Offices responsible for developing guidance and policy applicable to transport category airplanes for Aircraft Evaluation Division personnel and for Flight Standards aviation safety inspectors (Maintenance, Avionics, and Operations) in the conduct of their responsibilities.
Scatter factor—A life reduction factor used in the interpretation of fatigue analysis and fatigue test results.
Structural modification point (SMP)—The point in time when a structural area must be modified to preclude WFD.
Structures Task Group (STG)—A model-specific airplane industry group that consists of TC holders and operators responsible for developing aging airplane model-specific programs.
Such a group also includes regulatory authorities who approve and monitor those programs.
Supplemental Structural Inspection Program (SSIP)—A damage-tolerance-based inspection program. Structural inspection programs only address the structure identified by the TC holder using the guidance contained in AC 91-56 (as revised).
Teardown Inspection—The term used for the process of disassembling structure and using destructive inspection techniques or visual (e.g., magnifying glass and dye penetrant) or other non-destructive (e.g., eddy current, ultrasound) inspection techniques to identify the extent of damage within a structure caused by fatigue, corrosion, and accidental damage.
Appendix 2 A2-4 Test-to-structure factors—A series of factors used to adjust test results to full-scale structure. These factors could include, but are not limited to, differences in—
• Stress spectrum,
• Boundary conditions,
• Specimen configuration,
• Material differences,
• Geometric considerations, and
• Environmental effects.
Type design—The engineering definition of a particular product. The type design consists of the following (see 14 CFR § 21.31):
• Drawings and specifications;
• Dimensions, materials, and processes;
• Airworthiness limitations;
• For primary category aircraft, if desired: A special inspection and preventive maintenance program designed to be accomplished by an appropriately rated and trained pilot/owner; and
• Other data to describe the product design and to determine the airworthiness, noise characteristics, fuel venting, and exhaust emissions (where applicable).
WFD(average behavior)—The point in time when, without intervention, 50% of the fleet is expected to develop WFD for a particular structure.
Widespread fatigue damage (WFD)—The simultaneous presence of fatigue damage at multiple structural locations that are of sufficient size and density that the structure will no longer meet the residual strength requirements of § 25.571(b).
Appendix 3 A3-1 APPENDIX 3
ACRONYMS USED IN THIS AC AAWG Airworthiness Assurance Working Group AC Advisory Circular AD Airworthiness Directive ALI Airworthiness Limitation Item ALS Airworthiness Limitations Section ARAC Aviation Rulemaking Advisory Committee ATC Amended Type Certificate DAH Design Approval Holder DSG Design Service Goal FAA Federal Aviation Administration ICA Instructions for Continued Airworthiness ISP Inspection Start Point LOV Limit of Validity of the Engineering Data that Supports the Structural Maintenance Program MED Multiple Element Damage MOC Means of Compliance MSD Multiple Site Damage MTGW Maximum Takeoff Gross Weight PMI Principal Maintenance Inspector SMP Structural Modification Point SSID Supplemental Structural Inspection Document SSIP Supplemental Structural Inspection Program STC Supplemental Type Certificate STG Structures Task Group TC Type Certificate TSO Technical Standard Order WFD Widespread Fatigue Damage
Appendix 4 A4-1 APPENDIX 4
BACKGROUND Structural fatigue is the degradation of a material subjected to repeated structural loads. Fatigue of metallic structure is recognized as a significant threat to the continued airworthiness of airplanes. This is because even small fatigue cracks can significantly reduce the strength of airplane structure. For over 50 years, the airworthiness standards for certification of new transport category airplanes have addressed fatigue. These airworthiness standards are meant to prevent catastrophic failures due to fatigue throughout the anticipated operational life of the airplane. These standards have evolved over the years and have changed as the relevant knowledge base has increased. This knowledge includes service experience, specific incidents and accidents, and technological advances in design, analysis, testing, manufacturing, and inspection of airplanes.
One of the first significant changes in the airworthiness standards occurred in March 1956, with the revision of the fatigue evaluation requirements contained in Civil Air Regulations (CAR) 4b.270. This revision added “fail-safe strength” as an option to the “fatigue strength” approach for addressing fatigue. Motivation for this change was the realization that precluding the occurrence of fatigue cracking might not always be possible and, therefore, as an option, the structure may be designed to survive cracking. The fatigue strength approach aims for a design where fatigue cracking is not probable within the operational life of the airplane. The fail-safe approach assumes that cracking could occur, but that a specified minimum strength could be maintained after a “fatigue failure or obvious partial failure.” The efficacy of the fail-safe approach was not only dependent on the structure keeping the specified minimum strength with the fatigue damage present, but also on finding the damage during normal maintenance. As applied, the fail-safe approach emphasizes redundancy as opposed to fatigue performance, and inspectability is assumed and not quantified. The fail-safe option was the predominant approach chosen for most large transport category airplanes certified in the 1960s and 1970s.
Another significant change in the airworthiness standards for fatigue occurred in October 1978 with Amendment 25-45, when § 25.571 was revised and § 25.573 was deleted. This change involved removing the fail-safe option entirely and establishing a new requirement to develop damage-tolerance-based inspections wherever practical. The fatigue strength approach, as a default option, is used only if the damage-tolerance approach is impractical. The motivation for the 1978 change was the recognition, based on mounting evidence, that the fail-safe approach that had been applied up to that point was not reliable and would not achieve the desired level of safety. Specific areas of concern with the fail-safe approach included loss of “fail-safety” with age. This was because of the increased probability of cracking in the structure adjacent to the fatigue failure, or obvious partial failure, and the lack of directed inspections and quantification of residual life with the assumed damage present. It was agreed at the time that more emphasis was needed on where and how fatigue cracking could occur in the structure, and on quantifying crack growth and residual strength characteristics. Such an approach includes knowledge of damage tolerance characteristics and development of effective inspection protocols, such as
Appendix 4 A4-2 where, when, how, and how often to inspect. The 1978 changes achieved this for certification of new transport category airplanes.
The same events and reasoning that drove the changes to airworthiness standards for new airplanes also influenced the strategy adopted to ensure continued airworthiness of the existing fleet. There was increasing concern about existing older airplanes that had been certified according to the fail-safe requirements of CAR 4b.270. Eleven large transport models were specifically identified as needing the most attention. The FAA determined a need to develop damage-tolerance-based inspection programs. These inspections supplemented existing maintenance inspections, so they were referred to as Supplemental Structural Inspection Programs (SSIP). The inspection requirements for these programs were documented in supplemental structural inspection documents (SSID). It was also agreed that SSIDs would be developed by the original equipment manufacturers on a voluntary basis and then mandated by ADs. The Civil Aviation Authority for the United Kingdom published guidance for developing the SSIPs in Airworthiness Notice No. 89, Continuing Structural Integrity of Transport Aeroplanes, dated August 23, 1978, and the FAA published guidance for developing the SSIPs on May 6, 1981 in AC 91-56, Supplemental Structural Inspection Program for Large Transport Category Airplanes. SSIPs were subsequently developed and documented in SSIDs and mandated by ADs for the eleven aging model airplanes.
In April 1988, one of the eleven aging model airplanes suffered major structural damage to its pressurized fuselage because of undetected fatigue cracking of the baseline primary structure.
Although that airplane had an SSIP that was mandated by an AD, there were no special directed inspections for fatigue cracks at multiple structural locations. This was because it was believed that the link-up of multiple fatigue cracks in one skin frame bay would result in safe decompression by skin flapping.3 It was thought that the damage to the fuselage skin would be obvious by inspection or by the inability to pressurize the fuselage. The accident was attributed, in part, to the aging of the airplane involved. This aging included the simultaneous presence of small fatigue cracks at multiple locations in the fuselage skin lap splice. Instead of being obvious, those cracks grew undetected. Then they linked up quickly to cause catastrophic failure of a large section of the fuselage.
3 Flapping is a phenomenon that occurs in cracks in fuselage skin subjected to cabin pressure. When the two tips of the crack meet stiffened structure, they change direction and turn away from the stiffened structure.
T hat accident precipitated actions that culminated in regulations aimed at avoiding catastrophic failures from fatigue in existing and future airplanes. In response to the April 1988 accident, the FAA sponsored a conference on aging airplanes and established a task force representing the interests of the airplane operators, airplane manufacturers, regulatory authorities, and other aviation representatives. This task force was later renamed the Airworthiness Assurance Working Group (AAWG). They specifically recommended establishment of an Aging Aircraft Program to address long-term airworthiness issues in airplane structure that result from aging.
The AAWG also recommended that the program include an element for addressing fatigue cracking at multiple structural locations. The National Transportation Safety Board recommended that DAHs discontinue classifying fuselage skin as “malfunction evident” or “damage obvious” in SSIDs.
Appendix 4 A4-3 The April 1988 accident also precipitated Congressional legislation. In October 1991, Congress enacted Title IV of Public Law 102-143, the Aging Airplane Safety Act of 1991 (AASA). The AASA had two key elements:
(1) It required “the Administrator to make such inspections and conduct such reviews of maintenance and other records of each airplane used by an operator to provide air transportation as may be necessary to determine that such is in a safe condition and is properly maintained for operation in air transportation.”
(2) It specified that an operator must be able to demonstrate, as part of that inspection, “that maintenance of the airplane’s structure, skin, and other age sensitive parts and components have been adequate and timely enough to ensure the highest level of safety.”
The SSIPs were revised to remove the methodology for classifying certain fatigue cracking in structures as “malfunction evident” or “damage obvious” and to include damage-tolerance-based inspections for those structures. The FAA issued ADs to mandate those changes. The SSIP is one element of the overall Aging Aircraft Program for structures. We also adopted changes to operating rules and issued ADs to mandate, in part, other elements of the Aging Aircraft Program, including:
• The Mandatory Modification Program.
• The Repair Assessment Program.
• The Corrosion Prevention and Control Program.
For further discussion on the Aging Aircraft Program for structures, see AC 91-56 (as revised).
In response to the AASA, in part, the FAA issued the Aging Airplane Safety Final Rule,4 the Damage Tolerance Data Rule,5 and guidance material. The Damage Tolerance Data Rule is the design-approval-holder component that facilitates operator compliance with the Aging Airplane Safety Final Rule. For further background information and discussion on these rules, see AC 120-93 (as revised).
4 Aging Airplane Safety: 70 FR 5518; February 2, 2005.
5 Damage Tolerance Data for Repairs and Alterations: 72 FR 70486; December 12, 2007.
In 1998, t he FAA amended § 25.571 (Amendment 25-966) of the aircraft certification requirements for transport category airplanes. Under this amendment, we introduced the term widespread fatigue damage (WFD) into § 25.571. WFD is a condition that occurs when there is simultaneous presence of fatigue cracks at multiple structural locations that are of sufficient size and density that the structure will no longer meet the residual strength requirements of § 25.571(b). As part of the certification process, § 25.571 requires full-scale fatigue test evidence to demonstrate that WFD will not occur before an airplane reaches its DSG.
6 Damage-Tolerance and Fatigue Evaluation of Structure: 63 FR 15708; March 31, 1998.
A nother significant change in the airworthiness standards for fatigue occurred in November 2010 with the issuance of Amendment Nos. 25-132 and 26-57 when the FAA revised § 25.571 and
7 Aging Aircraft Program: Widespread Fatigue Damage: 75 FR 69746, November 15, 2010.
Appendix 4 A4-4 added §§ 26.21 and 26.23. We determined that this change was necessary because the structural fatigue characteristics of airplanes are understood only up to the point where analyses and testing of the structure are valid. Until Amendment Nos. 25-132 and 26-5, there were no requirements to limit the operation of airplanes based on existing engineering data to prevent catastrophic failures from WFD. We adopted these amendments to require that DAHs establish a limit of validity of the engineering data that supports the structural maintenance program (LOV) and demonstrate that WFD will not occur in the airplane before it reaches LOV. Under this change, we also added §§ 121.1115 and 129.115 in Amendment Nos. 121-351 and 129-48, to prohibit operation of an airplane beyond its LOV. Section 26.23 provides an option for any person to extend the LOV and to develop the maintenance actions that support the extended limit.
Thereafter, to operate an airplane beyond the existing LOV, an operator must incorporate the extended LOV and associated maintenance actions into its maintenance program. The airplane may not be operated beyond the extended LOV. These amendments, which specifically addresses WFD, is intended to be the last element of the overall Aging Aircraft Program for structures.
Appendix 5 A5-1 APPENDIX 5
EXAMPLES OF STRUCTURE SUSCEPTIBLE TO WFD
STRUCTURAL AREA SEE FIGURE Longitudinal Skin Joints, Frames, and Tear Straps (MSD/MED) 5-1 Circumferential Joints and Stringers (MSD/MED) 5-2 Lap Joints with Milled, Chem-milled, or Bonded Radius (MSD) 5-3 Fuselage Frames (MED) 5-4 Stringer-to-Frame Attachments (MED) 5-5 Shear Clip End Fasteners on Shear-Tied Fuselage Frames (MSD/MED) 5-6 Aft Pressure Dome Outer Ring and Dome Web Splices (MSD/MED) 5-7 Skin Splice at Aft Pressure Bulkhead (MSD) 5-8 Abrupt Changes in Web or Skin Thickness—Pressurized or Unpressurized Structure (MSD/MED) 5-9 Window Surround Structure (MSD, MED) 5-10 Over-Wing Fuselage Attachments (MED) 5-11 Latches and Hinges of Non-Plug Doors (MSD/MED) 5-12 Skin at Runout of Large Doubler (MSD)—Fuselage, Wing, or Empennage 5-13 Wing or Empennage Chordwise Splices (MSD/MED) 5-14 Rib-to-Skin Attachments (MSD/MED) 5-15 Typical Wing and Empennage Construction (MSD/MED) 5-16
Appendix 5 Figure 5-1 Longitudinal Skin Joints, Frames, and Tear Straps (MSD/MED)
(a) Lap joint (b) Butt joint (c) Lap joint Tear straps with radius Outer skin upper rivet row Longitudinal skin joint - - Stringer
• Inner skin lower rivet
- Frame row Type and possible location of MSD and MED Service or test experience of factors that influence MSD
• MSD longitudinal skin joint and MED (examples)
• Lap joint • High stress - misuse of data from coupon test
- Outer skin upper rivet row • Corrosion
- Inner skin lower rivet row • Disbond
• Butt joint • Manufacturing defect
- Skin outer rivet rows • Surface preparation
- Doubler inner rivet rows • Bond laminate too thin
• Lap joint with radius • Countersink, fastener fit
- In radius • Design defect-surface preparation process
• MED-frame
• Stress concentration areas
• MED-tear straps
• Critical fastener rows in the skin at tear strap joint Figure 5-2 Circumferential Joints and Stringers (MSD/MED) Circumferential (a) Without (b) With outer Skin • splice plate outer doubler doubler
- Stringers Frame Type and possible location of MSD/MED Service or test experience of factors that influence
• MSD - circumferential joint MSD and/or MED (examples)
• Without outer doubler • High secondary bending
- Splice plate-between and/or at the inner two • High stress level in splice plate and joining stringers rivet rows (misuse of data from coupon test)
- Skin- forward and aft rivet row of splice plate • Poor design (wrong material)
- Skin- at first fastener of stringer coupling • Underdesign (over-estimation of interference fit fasteners)
• With outer doubler
- Skin - outer rivet rows
- Splice plate/outer doubler-inner rivet rows
• MED-stringer/stringer couplings
- Stringer - at first fastener of stringer coupling
- Stringer coupling - in splice plate area A5-2
Appendix 5 Figure 5-3 Lap Joints with Milled, Chem-milled, or Bonded Radius (MSD) or chem-milled step • Outer skin at milled Cracking Bonded doubler Bonded joint Type and possible location of MSD and MED
• MSD-abrupt cross section change
• Milled radius
• Chem-milled radius
• Bonded doubler runout Service or test experience of factors that influence MSD and MED (examples)
• High bending stresses due to eccentricity Figure 5-4 Fuselage Frames (MED) Fuselage skin panel
- Stringer
• Frame
•Frame Skin • Typical fuselage Stringer skin panel ~ Typical cracking Tear strap Type and possible location of MSD/MED
• MED-the cracking of frames at stringer cutouts at successive longitudinal locations in the fuselage. The primary concern is for those areas where noncircular frames exist in the fuselage structure. Fractures in those areas would result in panel instability.
Service or test experience of factors that influence MSD and/or MED (examples)
• High bending - noncircular frames
• Local stress concentrations
• Cutouts
• Shear attachments A5-3
Appendix 5 Figure 5-5 Stringer-to-Frame Attachments (MED) Fuselage skin panel
- Frame
- Stringer
- Frame 2 Stringer Typical cracks - Skin Type and possible location of MED
• MED- any combination of fracture of frames, clips, or stringers, including the attachments, resulting in the loss of the shear tie between the frame and stringer This condition may occur at either circumferential or longitudinal locations at fuselage frame/stringer intersection.
Service or test experience of factors that influence MSD and/or MED (examples)
• Poor load path connection Figure 5-6 Shear Clip End Fasteners on Shear-Tied Fuselage Frame (MSD/MED) Skin cracking at end fasteners
- Skin Shear clip - - Longeron or stringer Stringer or frame cap
- Frame cracking Type and possible location of MSD and MED
• MSD-skin at end fastener of shear clip
• MED- cracking in stringer or longeron at frame attachment
• MED-cracking in frame at stringer or longeron attachment Service or test experience of factors that influence MSD and MED (examples)
• Preload
• Localized bending due to pressure
• Discontinuous load path A5-4