Compare and highlight differences

Compare / EASA/ Incorporated Amendments/ Compare & Highlight Differences

Appendix 2 to AMC 20-20A Guidelines for the development of a programme to preclude the occurrence of widespread fatigue damage

Available versions for ERULES-1963177438-99999

ED Decision 2022/006/R found in: Aerodromes (No 139/2014) Part-ADR.AR Part-ADR.OR Part-ADR.OPS CS-ADR-DSN CS-HPT-DSN (Dec 2024)
ED Decision 2021/007/R found in: AMC-20 Amdt 23 - Airworthiness of Products, Parts and Appliances (Jan 2022)
ED Decision 2020/023/R found in: AMC-20 Amdt 21 - Airworthiness of Products, Parts and Appliances (Apr 2021)

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CS ADR-DSN.Q.852 Marking and lighting of overhead wires, cables, supporting towers, etc.

ED Decision 2022/006/R

(a) Marking: The wires, cables, etc. to be marked should be equipped with markers; the supporting tower should be coloured.

(b) Marking by colours: The supporting towers of overhead wires, cables, etc. that require marking should be marked in accordance with [CS ADR-DSN.Q.845(b)](#_DxCrossRefBm151830960), except that the marking of the supporting towers may be omitted when they are lighted by high-intensity obstacle lights by day.

(1) Markers displayed on or adjacent to objects should be located in conspicuous positions so as to retain the general definition of the object and should be recognisable in clear weather from a distance of at least 1 000 m for an object to be viewed from the air and 300 m for an object to be viewed from the ground in all directions in which an aircraft is likely to approach the object. The shape of markers should be distinctive to the extent necessary to ensure that they are not mistaken for markers employed to convey other information, and they should be such that the hazard presented by the object they mark is not increased.

(2) A marker displayed on an overhead wire, cable, etc., should be spherical and have a diameter of not less than 60 cm.

(3) The spacing between two consecutive markers, or between a marker and a supporting tower, should be appropriate to the diameter of the marker. The spacing should normally not exceed:

(i) 30 m where the marker diameter is 60 cm, increasing progressively with increase of the marker diameter to:

(ii) 35 m where the marker diameter is 80 cm; and

(iii) further progressive increases to a maximum of 40 m where the marker diameter is of at least 130 cm.

Where multiple wires, cables, etc., are involved, a marker should be located not lower than the level of the highest wire at the point marked.

(4) A marker should be of one colour. When installed, white and red, or white and orange, markers should be displayed alternately. The colour selected should contrast with the background against which it should be seen.

(5) When it has been determined that an overhead wire, cable, etc., needs to be marked but it is not practicable to install markers on the wire, cable, etc., then high-intensity obstacle lights, Type B, should be provided on their supporting towers.

(d) Lighting:

(1) High-intensity obstacle lights, Type B, should be used to indicate the presence of the tower supporting overhead wires, cables, etc. where:

(i) a safety assessment indicates such light to be essential for the recognition of the presence of wires, cables, etc.; or

(ii) it has not been found practicable to install marker on the wires, cables, etc.

(2) Where high-intensity obstacle lights, Type B, are used, they should be located at three levels:

(i) at the top of the tower;

(ii) at the lowest level of the catenary of the wires or cables; and

(iii) at approximately midway between these two levels.

(3) High-intensity obstacle lights, Type B, indicating the presence of a tower supporting overhead wires, cables, etc., should flash sequentially; first the middle light, second the top light, and last the bottom light. The intervals between flashes of the lights should approximate the following ratios:

<table border="1" cellpadding="0" cellspacing="0" width="529"><tr><td valign="top" width="265">Flash interval between</td><td valign="top" width="265">Ratio of cycle time</td></tr><tr><td valign="top" width="265">Middle and top light</td><td valign="top" width="265">1/13</td></tr><tr><td valign="top" width="265">Top and bottom light</td><td valign="top" width="265">2/13</td></tr><tr><td valign="top" width="265">Bottom and middle light</td><td valign="top" width="265">10/13</td></tr></table>

(4) The installation setting angles for high-intensity obstacle lights, Types B, should be in accordance with Table Q-5.

<table border="1" cellpadding="0" cellspacing="0" width="100%"><thead><tr><td valign="top" width="25%">(1)</td><td valign="top" width="9%">(2)</td><td valign="top" width="10%">(3)</td><td valign="top" width="13%">(4)</td><td valign="top" width="13%">(5)</td><td valign="top" width="13%">(6)</td><td valign="top" width="12%">(7)</td></tr><tr><td rowspan="2" valign="top" width="25%">Light type</td><td rowspan="2" valign="top" width="9%">Colour</td><td rowspan="2" valign="top" width="10%">Signal type/ (Flash Rate)</td><td colspan="3" valign="top" width="41%">Peak intensity (cd) at given Background Luminance (b)</td><td rowspan="2" valign="top" width="12%">Light Distribution Table</td></tr><tr><td valign="top" width="13%">Day(Above 500 cd/m2)</td><td valign="top" width="13%">Twilight(50-500 cd/m2)</td><td valign="top" width="13%">Night (Below 50 cd/m2)</td></tr></thead><tr><td valign="top" width="25%">Low-intensity Type A (fixed obstacle)</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">10</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type B (fixed obstacle)</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">32</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type C (mobile obstacle)</td><td valign="top" width="9%">Yellow/Blue (a)</td><td valign="top" width="10%">Flashing(60-90 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">40</td><td valign="top" width="13%">40</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type D (follow-me vehicle)</td><td valign="top" width="9%">Yellow</td><td valign="top" width="10%">Flashing(60-90 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">200</td><td valign="top" width="13%">200</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity, Type E</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Flashing(c)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">32</td><td valign="top" width="12%">Table Q-2(Type B)</td></tr><tr><td valign="top" width="25%">Medium-intensity Type A</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(20-60 fpm)</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">Medium-intensity Type B</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Flashing(20-60 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">Medium-intensity Type C</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">High-intensity Type A</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(40-60 fpm)</td><td valign="top" width="13%">200 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">High-intensity Type B</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(40-60 fpm)</td><td valign="top" width="13%">100 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td colspan="7" valign="top" width="100%">(a) <a href="#_DxCrossRefBm151830971">CS ADR-DSN.Q.850(b)</a>(b) For flashing lights, effective intensity as determined in accordance with ICAO Doc 9157, Aerodrome Design Manual, Part 4, Visual Aids.(c) For wind turbine application, to flash at the same rate as the lighting on the nacelle.</td></tr></table>

Table Q-1. Characteristics of obstacle lights

<table border="1" cellpadding="0" cellspacing="0" width="100%"><thead><tr><td rowspan="2" width="22%">Minimum intensity (a)</td><td rowspan="2" width="22%">Maximum intensity (a)</td><td colspan="2" width="41%">Vertical beam spread (f)</td></tr><tr><td width="20%">Minimum beam spread</td><td width="20%">Intensity</td></tr></thead><tr><td valign="top" width="12%">Type A</td><td valign="top" width="22%">10 cd (b)</td><td valign="top" width="22%">N/A</td><td valign="top" width="20%">10°</td><td valign="top" width="20%">5 cd</td></tr><tr><td valign="top" width="12%">Type B</td><td valign="top" width="22%">32 cd (b)</td><td valign="top" width="22%">N/A</td><td valign="top" width="20%">10°</td><td valign="top" width="20%">16 cd</td></tr><tr><td valign="top" width="12%">Type C</td><td valign="top" width="22%">40 cd (b)</td><td valign="top" width="22%">400 cd</td><td valign="top" width="20%">12(d)</td><td valign="top" width="20%">20 cd</td></tr><tr><td valign="top" width="12%">Type D</td><td valign="top" width="22%">200 cd (c)</td><td valign="top" width="22%">400 cd</td><td valign="top" width="20%">N/A(e)</td><td valign="top" width="20%">N/A</td></tr><tr><td colspan="5" valign="top" width="100%">Note: This table does not include recommended horizontal beam spreads. <a href="#_DxCrossRefBm151830962">CS ADR-DSN.Q.846(c)</a> requires 360° coverage around an obstacle. Therefore, the number of lights needed to meet this requirement will depend on the horizontal beam spreads of each light as well as the shape of the obstacle. Thus, with narrower beam spreads, more lights will be required.(a) 360° horizontal. For flashing lights, the intensity is read into effective intensity, as determined in accordance with ICAO, Aerodrome Design Manual, Part 4, Visual Aids.(b) Between 2 and 10° vertical. Elevation vertical angles are referenced to the horizontal when the light is levelled.(c) Between 2 and 20° vertical. Elevation vertical angles are referenced to the horizontal when the light is levelled.(d) Peak intensity should be located at approximately 2.5° vertical.(e) Peak intensity should be located at approximately 17° vertical.(f) Beam spread is defined as the angle between the horizontal plan and the directions for which the intensity exceeds that mentioned in the ‘intensity’ column.</td></tr></table>

Table Q-2. Light distribution for low-intensity obstacle lights

<table border="1" cellpadding="0" cellspacing="0" width="931"><thead><tr><td rowspan="4" valign="top" width="84">Benchmark intensity</td><td colspan="6" valign="top" width="438">Minimum requirements</td><td colspan="5" valign="top" width="410">Recommendations</td></tr><tr><td colspan="4" valign="top" width="273">Vertical elevation angle (b)</td><td colspan="2" rowspan="2" valign="top" width="164">Vertical beam spread (c)</td><td colspan="3" valign="top" width="246">Vertical elevation angle (b)</td><td colspan="2" rowspan="2" valign="top" width="164">Vertical beam spread (c)</td></tr><tr><td colspan="3" valign="top" width="186">0°</td><td valign="top" width="88">-1°</td><td valign="top" width="82">0°</td><td valign="top" width="82">-1°</td><td valign="top" width="82">-10°</td></tr><tr><td valign="top" width="91">Minimum average intensity (a)</td><td valign="top" width="91">Minimum intensity (a)</td><td colspan="2" valign="top" width="91">Minimum intensity (a)</td><td valign="top" width="82">Minimum beam spread</td><td valign="top" width="82">Intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum beam spread</td><td valign="top" width="82">Intensity (a)</td></tr></thead><tr><td valign="top" width="84">200 000</td><td valign="top" width="91">200 000</td><td valign="top" width="91">150 000</td><td colspan="2" valign="top" width="91">75 000</td><td valign="top" width="82">3°</td><td valign="top" width="82">75 000</td><td valign="top" width="82">250 000</td><td valign="top" width="82">112 500</td><td valign="top" width="82">7 500</td><td valign="top" width="82">7°</td><td valign="top" width="82">75 000</td></tr><tr><td valign="top" width="84">100 000</td><td valign="top" width="91">100 000</td><td valign="top" width="91">75 000</td><td colspan="2" valign="top" width="91">37 500</td><td valign="top" width="82">3°</td><td valign="top" width="82">37 500</td><td valign="top" width="82">125 000</td><td valign="top" width="82">56 250</td><td valign="top" width="82">3 750</td><td valign="top" width="82">7°</td><td valign="top" width="82">37 500</td></tr><tr><td valign="top" width="84">20 000</td><td valign="top" width="91">20 000</td><td valign="top" width="91">15 000</td><td colspan="2" valign="top" width="91">7 500</td><td valign="top" width="82">3°</td><td valign="top" width="82">7 500</td><td valign="top" width="82">25 000</td><td valign="top" width="82">11 250</td><td valign="top" width="82">750</td><td valign="top" width="82">N/A</td><td valign="top" width="82">N/A</td></tr><tr><td valign="top" width="84">2 000</td><td valign="top" width="91">2 000</td><td valign="top" width="91">1 500</td><td colspan="2" valign="top" width="91">750</td><td valign="top" width="82">3°</td><td valign="top" width="82">750</td><td valign="top" width="82">2 500</td><td valign="top" width="82">1 125</td><td valign="top" width="82">75</td><td valign="top" width="82">N/A</td><td valign="top" width="82">N/A</td></tr><tr><td colspan="12" valign="top" width="931">Note: This table does not include recommended horizontal beam spreads. <a href="#_DxCrossRefBm151830962">CS ADR-DSN.Q.846(c)</a> requires 360° coverage around an obstacle. Therefore, the number of lights needed to meet this requirement will depend on the horizontal beam spreads of each light as well as the shape of the obstacle. Thus, with narrower beam spreads, more lights will be required.(a) 360° horizontal. All intensities are expressed in Candela. For flashing lights, the intensity is read into effective intensity, as determined in accordance with ICAO Doc 9157, Aerodrome Design Manual, Part 4, Visual Aids.(b) Elevation vertical angles are referenced to the horizontal when the light unit is levelled.(c) Beam spread is defined as the angle between the horizontal plan and the directions for which the intensity exceeds that mentioned in the ‘intensity’ column.Note: an extended beam spread may be necessary under specific configuration and justified by a safety assessment.</td></tr><tr height="0"></tr></table>

Table Q-3. Light distribution for medium- and high-intensity obstacle lights according to benchmark intensities of Table Q-1

<table border="1" cellpadding="0" cellspacing="0" width="929"><thead><tr><td colspan="2" valign="top" width="619">Longest dimension</td><td rowspan="2" width="310">Band width</td></tr><tr><td valign="top" width="309">Greater than</td><td valign="top" width="310">Not exceeding</td></tr></thead><tr><td valign="top" width="309">1.5 m</td><td valign="top" width="310">210 m</td><td valign="top" width="310">1/7 of longest dimension</td></tr><tr><td valign="top" width="309">210 m</td><td valign="top" width="310">270 m</td><td valign="top" width="310">1/9 “ “ “</td></tr><tr><td valign="top" width="309">270 m</td><td valign="top" width="310">330 m</td><td valign="top" width="310">1/11 “ “ “</td></tr><tr><td valign="top" width="309">330 m</td><td valign="top" width="310">390 m</td><td valign="top" width="310">1/13 “ “ “</td></tr><tr><td valign="top" width="309">390 m</td><td valign="top" width="310">450 m</td><td valign="top" width="310">1/15 “ “ “</td></tr><tr><td valign="top" width="309">450 m</td><td valign="top" width="310">510 m</td><td valign="top" width="310">1/17 “ “ “</td></tr><tr><td valign="top" width="309">510 m</td><td valign="top" width="310">570 m</td><td valign="top" width="310">1/19 “ “ “</td></tr><tr><td valign="top" width="309">570 m</td><td valign="top" width="310">630 m</td><td valign="top" width="310">1/21 “ “ “</td></tr></table>

Table Q-4. Marking band widths

<table border="1" cellpadding="0" cellspacing="0" width="929"><tr><td colspan="2" valign="top" width="619">Height of light unit above terrain (AGL)</td><td rowspan="2" width="310">Angle of the peak of the beam above the horizontal</td></tr><tr><td valign="top" width="310">Greater than</td><td valign="top" width="310">Not exceeding</td></tr><tr><td valign="top" width="310">151 m</td><td valign="top" width="310"></td><td valign="top" width="310">0°</td></tr><tr><td valign="top" width="310">122 m</td><td valign="top" width="310">151 m</td><td valign="top" width="310">1°</td></tr><tr><td valign="top" width="310">92 m</td><td valign="top" width="310">122 m</td><td valign="top" width="310">2°</td></tr><tr><td valign="top" width="310"></td><td valign="top" width="310">92 m</td><td valign="top" width="310">3°</td></tr></table>

Table Q-5. Installation setting angles for high-intensity obstacle lights

[Issue: ADR-DSN/3]

[Issue: ADR-DSN/4]

[Issue: ADR-DSN/6]

Appendix 2 to AMC 20-20B Guidelines for the development of a
programme to preclude the occurrence of widespread fatigue damage

ED Decision 2021/007/R

1.         INTRODUCTION

The terminology and methodology in this Appendix are based
upon material developed by the AAWGand lessons learned since the first issue
of this AMC.

2.         DEFINITIONS

Extended service goal (ESG) is an
adjustment to the design service goal established by service experience,
analysis, and/or test during which the principal structure will be reasonably
free from significant cracking including WFD.

Monitoring period is the period of
time when special inspections of the fleet are initiated due to an increased
risk of MSD/MED (ISP) and ending when the SMP is reached.

Scatter factor is a life reduction
factor used in the interpretation of fatigue analysis and fatigue test
results.

Test-to-structure factor is 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.

Teardown inspection is the process of
disassembling structure and using destructive inspection techniques or visual
(magnifying glass and dye penetrant) or other non-destructive inspection (NDI)
methods (eddy current, ultrasonic) to identify the extent of damage, within a structure,
caused by fatigue, environmental or accidental damage.

WFD (average behaviour) is
the point in time when 50 % of the fleet is expected to reach WFD for a
particular detail.

3.         GENERAL

The likelihood of the occurrence of fatigue damage in an
aircraft’s structure increases with aircraft usage. The design process
generally establishes a design service goal (DSG) in terms of flight cycles/hours
for the airframe. It is expected that any cracking that occurs on an aircraft
operated up to the DSG will occur in isolation (i.e. local cracking),
originating from a single source, such as a random manufacturing flaw (e.g. a
mis-drilled fastener hole) or a localised design detail. It is considered unlikely
that cracks from manufacturing flaws or localised design issues will interact
strongly as they grow.

With extended usage, uniformly loaded structure may develop
cracks in adjacent fastener holes, or in adjacent similar structural details. These
cracks may or may not interact, and they can have an adverse effect on the residual
strength capability of the structure before the cracks become detectable. The
development of cracks at multiple locations (both MSD and MED) may also result
in strong interactions that can affect subsequent crack growth; in which case,
the predictions for local cracking would no longer apply. An example of this
situation may occur at any skin joint where load transfer occurs. Simultaneous
cracking at many fasteners along a common rivet line may reduce the residual
strength of the joint to less than the required levels before the cracks are
detectable under the routine maintenance programme established at the time of
certification.

For new type designs, certified to CS-25 Amendment 19, AMC 25.571
provides guidance on how to establish an LOV. For existing types, for which
TCHs need to comply with point 26.303 of Part‑26, CS 26.303 and this AMC
apply. The TCH should conduct structural evaluations to determine where and
when MSD/MED may occur. Based on these evaluations, the TCH should provide
additional maintenance instructions for the structure, as appropriate. The
maintenance instructions include, but are not limited to inspections, structural
modifications, and limits of validity of the new maintenance instructions. In
most cases, a combination of inspections and/or modifications/replacements is
deemed necessary to achieve the required safety level. Other cases will
require modification or replacement if inspections are not viable.

There is a distinct possibility that there could be a
simultaneous occurrence of MSD and MED in a given structural area. This
situation is possible on some details that were equally stressed. If this is
possible, then this scenario should be considered in developing appropriate
service actions for structural areas.

4.         STRUCTURAL
EVALUATION FOR WFD

4.1       General

The evaluation has three objectives:

(a)       Identify fatigue-critical
structure that may be susceptible to MSD/MED, see paragraph 4.2;

(b)      Predict when
it is likely to occur; see paragraph 4.3; and

(c)       Establish
additional maintenance actions, as necessary, to ensure continued safe
operation of the aircraft; see paragraph 4.4.

4.2       Structure
susceptible to MSD/MED

Susceptible structure is defined as that which has the
potential to develop MSD/MED. Such structure typically has the characteristics
of multiple similar details operating at similar stresses where structural
capability could be affected by interaction of multiple cracking at a number
of similar details. The following list provides examples of known types of
structure susceptible to MSD/MED. (The list is not exhaustive):

<table border="1" cellpadding="0" cellspacing="0" width="529">
<tr>
<td valign="top" width="425">
STRUCTURAL AREA
</td>
<td valign="top" width="104">
SEE FIGURE
</td>
</tr>
<tr>
<td valign="top" width="425">
Longitudinal
 Skin Joints, Frames, and Tear Straps (MSD/MED)
</td>
<td valign="top" width="104">
A2-1
</td>
</tr>
<tr>
<td valign="top" width="425">
Circumferential
 Joints and Stringers (MSD/MED)
</td>
<td valign="top" width="104">
A2-2
</td>
</tr>
<tr>
<td valign="top" width="425">
Lap
 joints with Milled, Chem-milled or Bonded Radius (MSD)
</td>
<td valign="top" width="104">
A2-3
</td>
</tr>
<tr>
<td valign="top" width="425">
Fuselage
 Frames (MED)
</td>
<td valign="top" width="104">
A2-4
</td>
</tr>
<tr>
<td valign="top" width="425">
Stringer
 to Frame Attachments (MED)
</td>
<td valign="top" width="104">
A2-5
</td>
</tr>
<tr>
<td valign="top" width="425">
Shear
 Clip End Fasteners on Shear Tied Fuselage Frames (MSD/MED)
</td>
<td valign="top" width="104">
A2-6
</td>
</tr>
<tr>
<td valign="top" width="425">
Aft
 Pressure Dome Outer Ring and Dome Web Splices (MSD/MED)
</td>
<td valign="top" width="104">
A2-7
</td>
</tr>
<tr>
<td valign="top" width="425">
Skin
 Splice at Aft Pressure Bulkhead (MSD)
</td>
<td valign="top" width="104">
A2-8
</td>
</tr>
<tr>
<td valign="top" width="425">
Abrupt
 Changes in Web or Skin Thickness — Pressurised or Un-pressurised Structure
 (MSD/MED)
</td>
<td valign="top" width="104">
A2-9
</td>
</tr>
<tr>
<td valign="top" width="425">
Window
 Surround Structure (MSD, MED)
</td>
<td valign="top" width="104">
A2-10
</td>
</tr>
<tr>
<td valign="top" width="425">
Over
 Wing Fuselage Attachments (MED)
</td>
<td valign="top" width="104">
A2-11
</td>
</tr>
<tr>
<td valign="top" width="425">
Latches
 and Hinges of Non-plug Doors (MSD/MED)
</td>
<td valign="top" width="104">
A2-12
</td>
</tr>
<tr>
<td valign="top" width="425">
Skin
 at Runout of Large Doubler (MSD)—Fuselage, Wing or Empennage
</td>
<td valign="top" width="104">
A2-13
</td>
</tr>
<tr>
<td valign="top" width="425">
Wing
 or Empennage Chordwise Splices (MSD/MED)
</td>
<td valign="top" width="104">
A2-14
</td>
</tr>
<tr>
<td valign="top" width="425">
Rib
 to Skin Attachments (MSD/MED)
</td>
<td valign="top" width="104">
A2-15
</td>
</tr>
<tr>
<td valign="top" width="425">
Typical
 Wing and Empennage Construction (MSD/MED)
</td>
<td valign="top" width="104">
A2-16
</td>
</tr>
</table>

![Aviation.Bot AI suggestion, not from EASA source: The image provides a visual representation of typical aircraft fuselage structural components and joint types, along with detailed lists of potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), and factors influencing their occurrence.

Detailed Description:
The image is composed of four distinct sections: an isometric drawing of a fuselage structure, three cross-sectional diagrams of different joint types, a bulleted list detailing types and possible locations of MSD and MED, and another bulleted list outlining service or test experience factors that influence MSD and MED.

1. **Isometric Fuselage Structure (Top Left):**
 * A blue line drawing illustrates a section of an aircraft fuselage, viewed obliquely, showing both external skin and internal stiffening elements.
 * The external surface features a vertical seam labeled "Longitudinal skin joint," with a black arrow pointing to it.
 * Internal structural elements include curved, horizontal members labeled "Frame," with a black arrow pointing to one such member.
 * Straight, vertical members are labeled "Stringer," with a black arrow pointing to one.
 * Thin, parallel vertical strips positioned beneath the skin are labeled "Tear straps," with a black arrow pointing to them.

2. **Joint Diagrams (Top Right):**
 * Three separate cross-sectional diagrams, all rendered with blue lines, depict different types of structural joints. Each diagram includes multiple horizontal dashed lines representing rivet rows.
 * **(a) Lap joint:** This diagram shows two flat panels overlapping. The upper panel is labeled "Outer skin upper rivet row" with a black arrow pointing to a dashed line within it. The lower, overlapping panel is labeled "Inner skin lower rivet row" with a black arrow pointing to a dashed line within it.
 * **(b) Butt joint:** This diagram shows two flat panels meeting edge-to-edge, joined by a central, internal doubler plate that overlaps both panels.
 * **(c) Lap joint with radius:** This diagram shows two overlapping panels similar to (a), but the inner panel features a distinct curved or "radius" section at its overlapping edge.

3. **Type and possible location of MSD and MED (Bottom Left):**
 * This section is titled "Type and possible location of MSD and MED" and presents a hierarchical bulleted list:
 * "● MSD longitudinal skin joint"
 * "● Lap joint"
 * "– Outer skin upper rivet row"
 * "– Inner skin lower rivet row"
 * "● Butt joint"
 * "– Skin outer rivet rows"
 * "– Doubler inner rivet rows"
 * "● Lap joint with radius"
 * "– In radius"
 * "● MED—frame"
 * "● Stress concentration areas"
 * "● MED—tear straps"
 * "● Critical fastener rows in the skin at tear strap joint"

4. **Service or test experience of factors that influence MSD and MED (examples) (Bottom Right):**
 * This section is titled "Service or test experience of factors that influence MSD and MED (examples)" and presents a bulleted list:
 * "● High stress—misuse of data from coupon test"
 * "● Corrosion"
 * "● Disbond"
 * "● Manufacturing defect"
 * "● Surface preparation"
 * "● Bond laminate too thin"
 * "● Countersink, fastener fit"
 * "● Design defect—surface preparation process"

Contextual Linkage:
The image visually connects general aircraft fuselage components (frames, stringers, skin, tear straps, longitudinal skin joints) with specific joint configurations (lap, butt, lap with radius) that are prone to fatigue damage. The lists explicitly detail where Multiple Site Damage (MSD) and Multiple Element Damage (MED) are likely to occur within these structural areas and joint types. For example, MSD in a longitudinal skin joint is further specified to occur in lap joints (at outer/inner skin rivet rows), butt joints (at skin outer/doubler inner rivet rows), and lap joints with radius (within the radius). MED is linked to stress concentration areas in frames and critical fastener rows in the skin at tear strap joints. The final list provides examples of external factors such as high stress, corrosion, disbond, manufacturing defects, and design defects that can contribute to or influence the development of MSD and MED in these susceptible areas.](file_tNNDI2wQkpg/image006.png)

Figure A2-1:
Longitudinal skin joints, frames, and tear straps (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image illustrates the structural components of circumferential joints and stringers in an aircraft fuselage, detailing potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), and listing factors that influence their occurrence.

**Detailed Description:** The image is composed of four distinct sections: an isometric structural diagram, two cross-sectional diagrams, and two text blocks.

1.  **Top-Left Diagram (Isometric View):** This diagram presents an isometric, cutaway view of an aircraft fuselage section, highlighting a circumferential joint.
    *   A curved panel labeled "Skin" is visible at the top, representing the outermost layer.
    *   Below the "Skin" and running circumferentially, a "Circumferential splice plate" is shown, connected to the skin by multiple rows of fasteners (depicted as small, cross-hatched circles).
    *   Longitudinal structural members, labeled "Stringers," are positioned beneath the splice plate and are also connected by fasteners. These stringers appear to be continuous across the joint.
    *   At the bottom, a "Frame" is indicated, representing a transverse structural support member.
    *   The fasteners are arranged in staggered rows, indicating a riveted or bolted connection.

2.  **Top-Right Diagrams (Cross-Sectional Views):** Two vertical cross-sectional diagrams illustrate variations of the joint configuration.
    *   **(a) "Without outer doubler":** This diagram shows three vertical plates connected by four horizontal lines representing fasteners. The central plate is wider than the two outer plates. The fasteners pass through all three plates.
    *   **(b) "With outer doubler":** This diagram shows four vertical plates connected by four horizontal lines representing fasteners. The two innermost plates are wider, and the two outermost plates are narrower. The fasteners pass through all four plates. The presence of the two outermost narrower plates distinguishes this configuration as having an "outer doubler."

3.  **Bottom-Left Text Block ("Type and possible location of MSD/MED"):** This section provides a bulleted list detailing specific locations for MSD and MED.
    *   The main heading is "Type and possible location of MSD/MED".
    *   Under "MSD—circumferential joint":
        *   "Without outer doubler" sub-section lists:
            *   "Splice plate—between and/or at the inner two rivet rows"
            *   "Skin—forward and aft rivet row of splice plate"
            *   "Skin—at first fastener of stringer coupling"
        *   "With outer doubler" sub-section lists:
            *   "Skin—outer rivet rows"
            *   "Splice plate/outer doubler—inner rivet rows"
    *   Under "MED—stringer/stringer couplings":
        *   "Stringer—at first fastener of stringer coupling"
        *   "Stringer coupling—in splice plate area"

4.  **Bottom-Right Text Block ("Service or test experience of factors that influence MSD and/or MED (examples)"):** This section lists factors contributing to MSD/MED.
    *   The main heading is "Service or test experience of factors that influence MSD and/or MED (examples)".
    *   The listed factors are:
        *   "High secondary bending"
        *   "High stress level in splice plate and joining stringers (misuse of data from coupon test)"
        *   "Poor design (wrong material)"
        *   "Underdesign (over-estimation of interference fit fasteners)"

**Contextual Linkage:**
The isometric view provides a general understanding of the "Circumferential joints and stringers" structural area. The two cross-sectional diagrams visually represent the "Without outer doubler" and "With outer doubler" configurations explicitly mentioned in the "Type and possible location of MSD/MED" text block. This text block then specifies precise locations within these configurations (e.g., "Splice plate," "Skin," "Stringer," "Stringer coupling," "rivet rows," "fastener") where MSD and MED are likely to occur. The final text block lists general factors derived from service or test experience that can contribute to the development of MSD and MED in such structural areas.](file_tNNDI2wQkpg/image007.png)

Figure A2-2:
Circumferential joints and stringers (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates two types of aircraft structural lap joints, detailing potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), and factors influencing their occurrence.

Detailed Description:
The image displays two distinct, vertically oriented diagrams of structural cross-sections, positioned side-by-side, each accompanied by a main heading and a bulleted list of descriptive text below it.

* **Left Diagram (Type and possible location of MSD and MED):**
 * This diagram depicts a cross-sectional view of a two-part structural joint.
 * The left component is a vertical plate with a stepped profile at its upper portion, indicating a reduction in thickness, forming an external corner.
 * The right component is a vertical plate of uniform thickness.
 * The left plate overlaps the right plate, forming a lap joint.
 * An arrow originates from the text "Outer skin at milled or chem-milled step" and points to the external corner of the stepped profile on the left plate.
 * Four horizontal dashed lines extend across both plates, indicating potential fastener lines or reference planes within the joint.
 * Below this diagram, the heading "Type and possible location of MSD and MED" is displayed.
 * Underneath the heading, a bulleted list provides specific types of structural features susceptible to MSD:
 * "MSD—abrupt cross section change"
 * "Milled radius"
 * "Chem-milled radius"
 * "Bonded doubler runout"

* **Right Diagram (Service or test experience of factors that influence MSD and MED):**
 * This diagram depicts a cross-sectional view of a three-part structural joint.
 * Three vertical plates are shown: a leftmost plate, a middle plate, and a rightmost plate.
 * The middle plate is positioned between and overlaps both the leftmost and rightmost plates, forming a double-lap joint.
 * Four horizontal dashed lines extend across all three plates, aligning vertically with those in the left diagram, indicating potential fastener lines or reference planes.
 * An arrow originates from the text "Cracking" and points to the top-right edge of the rightmost plate, specifically at the interface with the middle plate.
 * An arrow originates from the text "Bonded doubler" and points to the middle plate.
 * An arrow originates from the text "Bonded joint" and points to the interface between the leftmost plate and the middle plate.
 * Below this diagram, the heading "Service or test experience of factors that influence MSD and MED (examples)" is displayed.
 * Underneath the heading, a bulleted list provides an example of a factor influencing MSD and MED:
 * "High bending stresses due to eccentricity"

Contextual Linkage:
The left diagram visually represents structural configurations where Multiple Site Damage (MSD) and Multiple Element Damage (MED) are likely to occur, specifically highlighting an "abrupt cross section change" as seen in a "milled or chem-milled step" or a "bonded doubler runout." The right diagram illustrates a "bonded doubler" and "bonded joint" configuration, demonstrating how "cracking" can occur, influenced by factors such as "high bending stresses due to eccentricity" as derived from service or test experience. Both diagrams use dashed lines to indicate common reference points or fastener locations across the joint sections.](file_tNNDI2wQkpg/image008.png)

Figure A2-3:
Lap joints with milled, chem-milled or bonded radius (MSD)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates typical aircraft fuselage structural components, specifically focusing on fuselage frames and stringers, and identifies areas susceptible to Multiple Element Damage (MED) along with factors influencing both MED and Multiple Site Damage (MSD).

Detailed Description:
The image is composed of two primary graphical representations of aircraft fuselage structure, positioned in the upper half, and two descriptive text boxes in the lower half.

The **upper left graphic**, labeled "Typical fuselage skin panel", presents an isometric, cutaway view of a curved aircraft fuselage section. It shows an outer "Fuselage skin panel" (a curved sheet metal component) with internal structural elements attached. Two "Frame" elements, which are curved, C-channel or Z-section shaped members, run vertically (circumferentially) along the inner surface of the skin. Multiple "Stringer" elements, which are I-beam or Z-section shaped members, run horizontally (longitudinally) along the inner surface of the skin, passing through cutouts in the frames. Black arrows point from the labels "Fuselage skin panel", "Stringer", and "Frame" to their respective components.

The **upper right graphic** provides a magnified isometric view of a structural joint. A large, curved grey arrow originates from the upper left graphic, specifically pointing from the intersection of a stringer and frame on the fuselage skin panel, and curves to the right, indicating the magnified area. This magnified view shows a vertical "Frame" element, which is a C-channel or Z-section shaped member. A horizontal "Stringer" element, an I-beam or Z-section shaped member, passes through a cutout in the web of the frame. A flat "Skin" element is attached to the outer flange of the frame and the outer flange of the stringer. A flat, rectangular "Tear strap" is attached to the inner flange of the frame, positioned below the stringer. Small, circular holes, representing fastener locations, are visible along the attachment lines of the skin to the frame and stringer, and the tear strap to the frame. A dashed line with small circles, labeled "Typical cracking", is depicted originating from the cutout in the frame, extending along the edge of the stringer, and also along the tear strap. Black arrows point from the labels "Skin", "Stringer", "Frame", and "Tear strap" to their respective components.

The **lower left text box**, titled "Type and possible location of MSD/MED", contains a bullet point description:
*   "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."

The **lower right text box**, titled "Service or test experience of factors that influence MSD and/or MED (examples)", lists four bullet points:
*   "High bending—noncircular frames"
*   "Local stress concentrations"
*   "Cutouts"
*   "Shear attachments"

Contextual Linkage:
The image visually connects a general representation of a fuselage skin panel with its internal frames and stringers to a magnified detail of a critical structural joint. The magnified view explicitly illustrates "Typical cracking" at the intersection of a frame, stringer, skin, and tear strap, particularly at the stringer cutout in the frame. The accompanying text defines Multiple Element Damage (MED) as cracking of frames at stringer cutouts at successive longitudinal locations, directly correlating with the visual representation of cracking in the magnified detail. Furthermore, the text lists factors such as "High bending—noncircular frames", "Local stress concentrations", "Cutouts", and "Shear attachments" as influences on both MSD and MED, which are all structural characteristics or locations depicted or implied in the magnified view.](file_tNNDI2wQkpg/image009.png)

Figure A2-4:
Fuselage frames (MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates the structural components of an aircraft fuselage, specifically focusing on a stringer-to-frame attachment, and identifies typical locations for Multiple Element Damage (MED) and factors influencing it.

Detailed Description:
The image is composed of two main graphical sections and two text boxes, all presented on a white background.

The left graphical section displays a partial, three-dimensional, wireframe-style view of a curved "Fuselage skin panel" with internal structural elements. Arrows point from the text "Fuselage skin panel" to the outer curved surface. Inside this panel, multiple horizontal elements are visible, one of which is labeled "Stringer" with an arrow pointing to it. Perpendicular to the stringers, multiple vertical elements are shown, one of which is labeled "Frame" with an arrow pointing to it. These stringers and frames form a grid-like internal structure supporting the skin panel.

A large, thick, curved grey arrow originates from the intersection of a stringer and a frame in the left graphical section and points towards the right graphical section.

The right graphical section provides a zoomed-in, isometric view of a single "Stringer-to-frame attachment". This detailed view shows a vertical component labeled "Frame" with an arrow pointing to its upper section. Attached to the lower part of the frame is a horizontal, L-shaped component labeled "Stringer" with an arrow pointing to its main body. Below the stringer and attached to it is a flat, rectangular component labeled "Skin" with an arrow pointing to its surface. The stringer is connected to the frame via a small, L-shaped clip or bracket, which is fastened to both the stringer and the frame. Two black arrows point to small, jagged lines on the stringer and the attachment clip, labeled "Typical cracks". These cracks are depicted at the bend of the stringer where it meets the attachment clip and on the attachment clip itself, near the fasteners that connect it to the frame. Multiple circular fasteners (rivets or bolts) are visible connecting the components.

Below the graphical elements, there are two distinct text boxes.

The left text box is titled "Type and possible location of MED". It contains one bullet point:
*   "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."

The right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)". It contains one bullet point:
*   "Poor load path connection"

Contextual Linkage:
The image visually defines and illustrates Multiple Element Damage (MED) in the context of aircraft fuselage structure. The left graphical section provides a general overview of the fuselage skin, stringers, and frames, establishing the structural context. The large grey arrow then highlights a specific "stringer-to-frame attachment" as a critical area. The detailed right graphical section zooms into this attachment, showing the individual components (frame, stringer, skin, and attachment clip) and explicitly indicating "Typical cracks" at the interface, which are a visual representation of the "fracture of frames, clips, or stringers, including the attachments" described in the "Type and possible location of MED" text box. The text further clarifies that this damage results in "loss of the shear tie between the frame and stringer" and can occur at "fuselage frame/stringer intersection," directly correlating with the depicted structural detail. The "Service or test experience" text box provides an example factor, "Poor load path connection," that can contribute to the depicted MED.](file_tNNDI2wQkpg/image010.png)

Figure A2-5:
Stringer-to-frame attachments (MED)

if !mso?

<table cellpadding="0" cellspacing="0" width="100%">
<tr>
<td>
<div>
Skin/Stringer Attachments
</div>
<?if !mso?></td>
</tr>
</table>

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:**
The image provides a detailed isometric line drawing of a shear-tied fuselage frame structural joint, illustrating potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with textual descriptions of these damage types and influencing factors.

**Detailed Description:**
The image consists of a technical line drawing of an aircraft structural assembly in the upper portion, and two distinct text boxes in the lower portion.

The upper portion displays an isometric view of a structural joint, rendered in blue lines against a white background. Key components of the aircraft fuselage structure are depicted:
*   A curved "Skin" panel is shown at the top, featuring two sets of parallel dashed lines indicating fastener rows. One set runs longitudinally, and another runs transversely.
*   A "Longeron or stringer," a longitudinal structural member, is positioned beneath the skin, running horizontally. It is attached to the skin via fasteners, indicated by dashed lines.
*   A "Frame," a transverse structural member, is shown vertically, extending downwards from the skin and intersecting the longeron/stringer.
*   A "Shear clip," a Z-shaped bracket, connects the frame to the skin and the longeron/stringer. It is depicted with dashed lines indicating fasteners along its attachment flanges.

Several black arrows with associated text labels point to specific elements or areas within the diagram:
*   An arrow labeled "Skin cracking at end fasteners" points to multiple 'X' marks distributed along the fastener lines on the skin, particularly where the skin is attached to the underlying longeron/stringer and shear clip.
*   An arrow labeled "Shear clip" points directly to the Z-shaped bracket.
*   An arrow labeled "Frame" points to the vertical transverse structural member.
*   An arrow labeled "Longeron or stringer" points to the horizontal longitudinal structural member.
*   An arrow labeled "Stringer or frame cap cracking" points to 'X' marks located on the longeron/stringer and on the horizontal flange of the frame where it connects to the longeron/stringer.

The lower left portion of the image contains a text box with the heading "Type and possible location of MSD and MED" followed by three bullet points:
*   "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"

The lower right portion of the image contains a text box with the heading "Service or test experience of factors that influence MSD and MED (examples)" followed by three bullet points:
*   "Preload"
*   "Localized bending due to pressure"
*   "Discontinuous load path"

**Contextual Linkage:**
The diagram visually illustrates the structural configuration of a shear-tied fuselage frame, highlighting specific areas susceptible to fatigue damage, which are further defined as Multiple Site Damage (MSD) and Multiple Element Damage (MED) in the accompanying text. The 'X' marks on the skin, longeron/stringer, and frame cap in the diagram correspond to the "cracking" locations described in the "Type and possible location of MSD and MED" list. The text also provides examples of operational factors that contribute to the development of such damage in these structural areas.](file_tNNDI2wQkpg/image011.png)

Figure A2-6:
Shear clip end fasteners on shear tied fuselage frame (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image provides a detailed illustration of an aft pressure dome, highlighting specific structural areas susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with cross-sectional examples of outer ring splices and a list of factors influencing these damage types.

**Detailed Description:** The image is structured into four distinct sections: a main structural diagram, detailed splice diagrams, a legend, and two descriptive text blocks.

The top-left section presents a black and white isometric line drawing of a circular, multi-ribbed aircraft component, identified as an aft pressure dome. This component features several concentric rings interconnected by radial ribs. Two black arrows, originating from the right side of the diagram, point to specific areas on the structure. The upper arrow indicates the outermost concentric ring, labeled "Outer ring splice". The lower arrow points to one of the radial ribs, labeled "Web splices".

The top-right section is titled "Typical outer ring splices" and contains three distinct black and white 2D cross-sectional line drawings, arranged vertically. Each drawing illustrates a different configuration of structural profiles joined together.
*   The topmost splice drawing depicts two L-shaped profiles connected by a series of three small circular elements, representing fasteners. A curved line indicates a radius. A capital "F" with a leader line points to the fasteners, and a capital "R" with a leader line points to the radius.
*   The middle splice drawing shows an L-shaped profile joined to a flat plate, with an additional triangular gusset providing support. A curved line indicates a radius. A capital "F" with a leader line points to a single fastener, and a capital "R" with a leader line points to the radius.
*   The bottommost splice drawing illustrates an L-shaped profile joined to a curved plate. A curved line indicates a radius. A capital "F" with a leader line points to a single fastener, and a capital "R" with a leader line points to the radius.

Positioned below the top-right section, a "Legend:" is provided, defining the symbols used in the splice diagrams:
*   "F fastener"
*   "R radius"

The bottom-left section contains a text block titled "Type and possible location of MSD/MED". It lists two primary categories of damage locations with sub-bullet points:
*   "• MSD/MED—outer ring splice"
    *   "• Attachment profiles—at fastener rows and/or in radius area"
*   "• MED—web splices"
    *   "• Bulkhead skin and/or splice plates—at critical fastener rows"

The bottom-right section contains a text block titled "Service or test experience of factors that influence MSD and/or MED (examples)". It lists four influencing factors:
*   "• Corrosion"
*   "• High stresses—combined tension and compression"
*   "• High induced bending in radius"
*   "• Inadequate finish in radius—surface roughness"

**Contextual Linkage:** The image visually connects the general structural components of an aft pressure dome, specifically the outer ring splice and web splices, to their susceptibility to Multiple Site Damage (MSD) and Multiple Element Damage (MED). The detailed cross-sectional diagrams of "Typical outer ring splices" provide visual examples of the "Attachment profiles" and "radius area" identified as critical locations for MSD/MED. The accompanying text further specifies that these damage types are likely to occur at "fastener rows" and "radius area" for outer ring splices, and at "critical fastener rows" for web splices and associated bulkhead skin/splice plates. The listed influencing factors provide critical context regarding conditions that can contribute to or exacerbate MSD/MED in these identified structural details.](file_tNNDI2wQkpg/image012.png)
Figure A2-7:
Aft pressure dome outer ring and dome web splices (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image provides an isometric diagram of an aircraft structural joint, illustrating components susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with factors influencing these damage types.

Detailed Description:
The image presents a three-dimensional, isometric view of an aircraft structural assembly, depicted with jagged edges indicating cut sections, suggesting it is a cross-section of a larger structure. The primary components shown are two flat panels, labeled "Pressurized skin" and "Unpressurized skin", positioned side-by-side and joined along their length. The "Pressurized skin" is on the left side of the joint, and the "Unpressurized skin" is on the right. These two skin panels are connected by two parallel rows of circular fasteners, which are depicted as small, solid circles, indicating rivets or bolts.

Beneath the joined skin panels, a ""T" frame" is attached perpendicularly. The horizontal flange of the "T" frame is fastened to the underside of both the "Pressurized skin" and the "Unpressurized skin" using fasteners arranged in two parallel rows, mirroring the fastener pattern on the top surface. An arrow points from the label ""T" frame" to this T-shaped structural member.

A jagged line, labeled "Skin cracking", is depicted on the "Pressurized skin", originating from one of the fastener holes in the row closer to the ""T" frame". An arrow points from the "Skin cracking" label to this jagged line.

Further below and attached to the vertical leg of the ""T" frame" is another structural component, labeled "Bulkhead web". An arrow points from the "Bulkhead web" label to this component.

Below the main diagram, there are two distinct text blocks providing additional information.
The left text block is titled "Type and possible location of MSD and MED" and contains one bullet point:
*   "MSD—skin at end fastener holes"

The right text block is titled "Service or test experience of factors that influence MSD and MED (examples)" and contains two bullet points:
*   "Shell discontinuous induced bending stresses"
*   "High load transfer at fastener"

Contextual Linkage:
The diagram visually represents a typical structural area, specifically a skin splice at an aft pressure bulkhead, which is identified as susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED). The "Skin cracking" shown originating from a fastener hole on the "Pressurized skin" illustrates a potential location for MSD, as explicitly stated in the accompanying text "MSD—skin at end fastener holes". The structural configuration, involving joined skin panels, a ""T" frame", and a "Bulkhead web", provides a visual context for the factors listed as influencing MSD and MED, namely "Shell discontinuous induced bending stresses" and "High load transfer at fastener", which are common stress concentrators in such riveted or bolted joints.](file_tNNDI2wQkpg/image013.png)
Figure A2-8: Skin splice at aft
pressure bulkhead (MSD)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates structural configurations susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED) due to abrupt changes in web or skin thickness, detailing their possible locations and factors influencing their occurrence.

Detailed Description:
The image is divided into two main sections: an upper section containing two engineering diagrams of structural joints, and a lower section containing two blocks of descriptive text.

The upper-left diagram presents an isometric view of a structural joint. It shows a horizontal "Web or skin" component connected to two vertical "Edge support member" components. The connection is made via rows of circular fasteners, with one row visible on the left edge support member and another on the right. Black arrows point from the text "Edge support member" to the top flanges of both vertical members. An arrow points from the text "Web or skin" to the horizontal plate. Additional arrows point to the internal corners where the "Web or skin" meets the "Edge support member," indicating "Typical cracking" as short, jagged lines originating from the fastener holes or the radius. An arrow points to the curved transition area at the corner, labeled "Radius." Below "Radius," three bullet points are listed: "Milled," "Chem-milled," and "Bonded doubler."

The upper-right diagram provides a 2D cross-sectional view of a similar structural joint. It depicts a single L-shaped "Edge support member" connected to a horizontal "Web or skin" component. Arrows from "Edge support member" point to the L-shaped profile, and arrows from "Web or skin" point to the horizontal plate. Arrows from "Typical cracking" point to small, jagged lines at the internal corner where the two components meet.

The lower-left text block is titled "Type and possible location of MSD and MED." It contains two sub-sections:
1.  "Abrupt change in stiffness*" with the following bullet points:
    *   "Milled radius"
    *   "Chem-milled radius"
    *   "Bonded doubler"
    *   "Fastener row at edge support members"
2.  "Edge member support structure" with the following bullet point:
    *   "Edge member - in radius areas"
An asterisk note below this section reads: "* Often multiple origins along edge member".

The lower-right text block is titled "Service or test experience of factors that influence MSD and MED." It contains two sub-sections:
1.  "Pressure structure" with the following bullet point:
    *   "High bending stresses at edge support due to pressure"
2.  "Non-pressure structure" with the following bullet point:
    *   "Structural deflections cause high stresses at edge supports"

Contextual Linkage:
The diagrams visually represent structural areas characterized by "Abrupt changes in web or skin thickness," as referenced in the surrounding text (Figure A2-9). The "Typical cracking" shown in both diagrams illustrates the potential manifestation of Multiple Site Damage (MSD) or Multiple Element Damage (MED) at these locations. The text blocks below the diagrams further elaborate on the specific types of structural features (e.g., milled radius, bonded doubler, fastener rows) that constitute "abrupt changes in stiffness" and are susceptible to MSD/MED. Additionally, the text identifies operational factors, such as pressure-induced bending stresses in pressure structures and structural deflections in non-pressure structures, that contribute to high stresses at edge supports, thereby influencing the occurrence of MSD and MED in these areas.](file_tNNDI2wQkpg/image014.png)

Figure A2-9:
Abrupt changes in web or skin thickness — Pressurised or unpressurised
structure (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: A diagram illustrating the internal structural components of an aircraft fuselage section around window cutouts, highlighting areas susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with influencing factors.

Detailed Description:
The image displays a black and white line drawing of a section of an aircraft fuselage structure, presented in an isometric view. The structure is composed of a series of parallel, curved vertical elements, representing fuselage frames, intersected by numerous parallel horizontal elements, representing stringers or stiffeners, forming a grid-like internal framework.

Along the lower-middle portion of this structural framework, a horizontal row of eight identical, oval-shaped cutouts is depicted, representing aircraft windows. Each window cutout is surrounded by a distinct, thicker reinforcement structure. The inner contour of this reinforcement is indicated by a dashed line, suggesting an underlying or hidden component.

A solid black arrow originates from the top-center of the image, pointing downwards and slightly to the left, terminating at the reinforcement structure of the third window from the left. Above the arrow's origin, the text "Window surround structure" is positioned.

Below the main structural diagram, two distinct text boxes provide additional information:
1.  The left text box is titled "Type and possible location of MSD/MED" and contains two bullet points:
    *   "MSD—skin at attachment to window surround structure"
    *   "MED—repeated details in reinforcement of window cutouts or in window corners"
2.  The right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)" and contains one bullet point:
    *   "High load transfer"

Contextual Linkage:
The diagram visually identifies the "Window surround structure" as a critical area within the aircraft fuselage. The accompanying text specifies that Multiple Site Damage (MSD) is likely to occur in the skin at the attachment points to this window surround structure. Conversely, Multiple Element Damage (MED) is indicated to occur in the repeated structural details found within the reinforcement of the window cutouts or specifically in the window corners. The text also lists "High load transfer" as a factor derived from service or test experience that can influence the development of both MSD and/or MED in this particular structural region.](file_tNNDI2wQkpg/image015.png)

Figure A2-10:
Window surround structure (MSD, MED)

![Aviation.Bot AI suggestion, not from EASA source: A line drawing illustrating the structural details of an overwing fuselage attachment area, highlighting regions susceptible to Multiple Element Damage (MED) and factors influencing it.

Detailed Description:
The image displays a black and white isometric line drawing of an aircraft structural section, focusing on the interface between a wing and fuselage. The primary structural element is a long, horizontal beam, likely a wing spar or stringer, extending from the left to the right side of the drawing. Attached perpendicularly and extending upwards from this beam are a series of evenly spaced, vertical, curved elements, representing wing ribs or frames. These ribs define the contour of the upper wing surface.

Below the main horizontal beam, a more intricate structural assembly is depicted, consisting of multiple smaller vertical and horizontal members. This assembly represents the fuselage attachment structure.

Two black arrows with solid lines and triangular arrowheads point to specific areas of the structure.
1. The upper arrow points to the top edge of the vertical ribs, with the accompanying text label "Upper wing surface" positioned to its right.
2. The lower arrow points to the complex structural assembly beneath the main horizontal beam, with the accompanying text label "Typical fuselage attachments" positioned to its right. A black oval outline encircles the area indicated by this lower arrow, emphasizing the "Typical fuselage attachments" region.

Below the structural drawing, two distinct text blocks provide additional information:

On the bottom left, a heading reads "Type and possible location of MSD/MED". Below this heading, a bullet point states:
* "MED—repeated details in overwing fuselage attachments"

On the bottom right, a heading reads "Service or test experience of factors that influence MSD and/or MED (examples)". Below this heading, two bullet points state:
* "Manufacturing defect—prestress"
* "Induced deflections"

Contextual Linkage:
The drawing visually represents the "Over wing fuselage attachments" identified in the surrounding text as a structural area susceptible to Widespread Fatigue Damage (WFD), specifically Multiple Element Damage (MED). The highlighted "Typical fuselage attachments" area directly corresponds to the "MED—repeated details in overwing fuselage attachments" description, indicating this specific region is prone to MED. The text blocks further elaborate on the type of damage (MED) and provide examples of factors (manufacturing defect—prestress, induced deflections) that can influence the occurrence of MSD and/or MED in such structural areas.](file_tNNDI2wQkpg/image016.png)

Figure A2-11:
Over wing fuselage attachments (MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED) on aircraft door hinges and latches, along with factors influencing these damage types.

Detailed Description: The image presents four distinct diagrams and associated textual information, arranged in two columns.

The left column features two diagrams related to a door hinge: 1. **Top-left diagram:** An isometric view of a section of an aircraft "Door", depicted with a grid-like pattern on its surface, possibly representing internal stiffeners or a window array. Two circular callouts, labeled "A" and "B", indicate specific areas on the door structure. Callout "A" points to the hinge line at the bottom-left edge of the door, while callout "B" points to a latch mechanism on the bottom-right edge. 2. **Bottom-left diagram (View A):** A detailed cross-sectional view of the hinge mechanism, corresponding to callout "A" from the top-left diagram. This view shows a "Hinge pin" passing horizontally through a series of interleaved "Hinge lobe" components. The hinge lobes are depicted as rectangular blocks with circular holes for the pin. Below this diagram, the heading "Type and possible location of MSD/MED" is presented, followed by a bulleted list: * "MSD—piano hinge" * "At hinge fastener attachment row" * "In fillet radius" * "Emanating from hole in lobes" * "MED—latches" * "In multiple latch hooks" * "At lube channel of latch spool" * "At spool bracket attach bolts (also corrosion)"

The right column features two diagrams related to a door latch: 1. **Top-right diagram:** An isolated, detailed view of a "Latch hook" component, shown in profile. It is a curved, hook-shaped metallic part with a pivot point at its upper end. An arrow points to the hook-shaped end, labeled "Latch hook". 2. **Bottom-right diagram (View B):** A detailed cross-sectional view of a latch assembly, corresponding to callout "B" from the top-left diagram. This view displays a central cylindrical "latch spool" component. An arrow points to a feature on this component, labeled "Lubrication passage". Below the central component, two "Attach bolts" are indicated by arrows pointing to their heads. Below this diagram, the heading "Service or test experience of factors that influence MSD and/or MED (examples)" is presented, followed by a bulleted list: * "Bending stresses due to fuselage elongation" * "High local stress" * "Fretting"

Contextual Linkage: The overall "Door" diagram provides the context for the detailed "View A" (hinge) and "View B" (latch) diagrams. "View A" specifically illustrates the "piano hinge" mentioned in the "Type and possible location of MSD/MED" list, detailing areas like the "hinge fastener attachment row," "fillet radius," and "hole in lobes" as potential MSD locations. "View B" and the isolated "Latch hook" diagram illustrate the "latches" mentioned in the same list, pointing to "multiple latch hooks," "lube channel of latch spool," and "spool bracket attach bolts" as potential MED locations. The text below "View B" provides general factors, such as "Bending stresses due to fuselage elongation," "High local stress," and "Fretting," that can influence the occurrence of both MSD and MED in these structural areas.](file_tNNDI2wQkpg/image017.jpg)
Figure A2-12: Latches and hinges of
non-plug doors (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image is a schematic diagram illustrating a typical aircraft structural joint involving a skin panel and a large doubler, highlighting areas susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED).

Detailed Description: The diagram depicts a cross-section or plan view of an aircraft structural assembly, primarily composed of a main "Skin" panel and an overlapping "Skin doubler," joined by numerous fasteners.

The upper and lower portions of the diagram show two long, solid horizontal structural members. These horizontal members are connected by several vertical elements, indicated by dashed lines, which are secured by rows of '+' (plus sign) shaped fasteners.

Between these horizontal members, the main "Skin" panel is visible, forming the base layer. Overlying a significant portion of this "Skin" panel are two large, rectangular "Skin doubler" sections, depicted with dashed outlines. Each doubler section contains an array of 'o' (circle) shaped fasteners arranged in a grid pattern, with additional rows of 'o' shaped fasteners along their perimeter, connecting them to the underlying skin and other structural elements. The doublers are positioned symmetrically, with a vertical gap between them, where a vertical dashed element with '+' fasteners connects the upper and lower horizontal members.

The fasteners are represented by two distinct symbols:
*   'o' (circles): Predominantly used for fasteners within the doubler sections and along their edges, as well as in horizontal rows connecting the doublers to the main structure.
*   '+' (plus signs): Primarily used for fasteners in the vertical connecting elements and in horizontal rows along the upper and lower edges of the main assembly.

Two black arrows with text labels point to specific components:
*   An arrow originating from below the diagram points upwards and to the left, indicating the lower, slightly curved edge of the main underlying structure, labeled "Skin".
*   Another arrow originating from below the diagram points upwards and to the right, indicating the lower edge of the overlapping rectangular structure, labeled "Skin doubler".

Below the main diagram, there are two distinct text blocks providing additional information:
1.  The first block is titled "Type and possible location of MSD/MED" and contains one bullet point:
    *   "MSD—cracks initiated at multiple critical fastener holes in skin at runout of doubler"
2.  The second block is titled "Service or test experience of factors that influence MSD and/or MED (examples)" and contains one bullet point:
    *   "High load transfer—high local stress"

Contextual Linkage: The diagram visually represents a structural configuration where the "Skin" and "Skin doubler" are joined by fasteners. The accompanying text explicitly links this configuration to the potential for Multiple Site Damage (MSD) and Multiple Element Damage (MED). Specifically, it highlights that MSD cracks are likely to initiate at multiple critical fastener holes located in the "Skin" where the "Skin doubler" ends (the "runout of doubler"). This area is further implicated by the mention of "High load transfer—high local stress" as a factor influencing MSD/MED, implying that the design and fastening in this region can lead to concentrated stresses conducive to fatigue cracking. The arrangement of fasteners and the overlap of the doubler on the skin illustrate the structural details relevant to these fatigue mechanisms.](file_tNNDI2wQkpg/image018.png)

Figure
A2-13: Skin at runout of large doubler (MSD) — Fuselage, wing or empennage

![Aviation.Bot AI suggestion, not from EASA source: This image illustrates a typical wing or empennage chordwise splice, detailing its components and identifying specific locations susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with influencing factors.

Detailed Description:
The image is composed of two main graphical sections and two text boxes, arranged horizontally.

The left graphical section presents a simplified, two-dimensional schematic view of a structural panel, likely representing a section of a wing or empennage. It is a tall, narrow rectangle with a solid vertical line on its left and right edges. The top left corner is labeled "Front spar". Inside the rectangle, numerous horizontal dashed lines extend from the left solid line towards the right, indicating internal structural elements. A small, open circle highlights a point on the left solid vertical line, approximately one-third of the way up from the bottom. A large, curved, solid grey arrow originates from this highlighted circle and sweeps towards the right, pointing to the right graphical section. Below the schematic, two diagonal black arrows point from the bottom left corner of the panel towards the bottom right, with the text "Chordwise joints" positioned between their arrowheads.

The right graphical section displays a detailed, three-dimensional isometric view of a "Typical skin and stringer splice". This view shows several interconnected structural components:
*   A "Rib" is depicted as a vertical, corrugated structural member at the top, extending partially out of view to the left.
*   A "Skin panel" is shown as a flat, horizontal surface positioned above the other components, with its edge visible.
*   A "Fitting" is a complex, multi-faceted component with several holes, attached to the rib and extending horizontally. An arrow points from the label "Fitting" to this component.
*   A "Splice plate" is a flat, elongated component with multiple circular and oval fastener holes, positioned beneath the fitting and partially overlapping the skin panel. An arrow points from the label "Splice plate" to this component.
*   A "Stringer" is a longitudinal structural member with a C-shaped cross-section, positioned below the splice plate and extending horizontally. An arrow points from the label "Stringer" to this component.

Below the graphical sections, two distinct text boxes provide additional information:

The bottom-left text box is titled "Type and possible location of MSD/MED" and contains a bulleted list:
*   "MSD—skin and/or splice plate"
*   *   "Chordwise critical fastener rows"
*   "MED—stringer runout of fitting"
*   *   "Fatigue-critical fastener holes at stringer and/or fitting"

The bottom-right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)" and contains a bulleted list:
*   "High load transfer"
*   "Local bending"

Contextual Linkage:
The image visually links a general structural area, identified as "Wing or empennage chordwise splices" (implied by the overall figure title and the "Chordwise joints" label), to a specific detailed component assembly, the "Typical skin and stringer splice". The grey arrow indicates that the detailed splice is an enlarged view of a critical area within the chordwise joint. The text boxes then provide specific information regarding where Multiple Site Damage (MSD) and Multiple Element Damage (MED) are likely to occur within this detailed splice (e.g., on the skin, splice plate, stringer, or at fastener holes) and list factors that contribute to their development, such as high load transfer and local bending.](file_tNNDI2wQkpg/image019.png)

Figure A2-14:
Wing or empennage chordwise splices (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** This image illustrates typical structural configurations and damage locations for rib-to-skin attachments susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with factors influencing these damage types.

**Detailed Description:** The image presents three distinct structural diagrams and two text boxes, all rendered in blue lines on a white background.

The top-left diagram shows a partial, elongated structural assembly, likely a wing or empennage rib structure. It consists of a horizontal member with numerous vertical, corrugated-like elements (ribs) attached along its length. An oval highlights a specific attachment point on the horizontal member, from which two black arrows extend downwards, pointing towards the central detailed diagram and the label "Typical shear tie".

The top-right diagram depicts an isolated, irregularly shaped structural component, possibly a skin panel, with two parallel rows of circular fastener holes. From several of these fastener holes, short, thin lines are drawn, representing cracks. Two black arrows point from this component to the label "Typical skin cracking".

The central and largest diagram provides a detailed, close-up cross-sectional view of a structural joint, specifically a rib-to-skin attachment. This joint comprises three main components:
1.  A flat, horizontal element at the top, labeled "Skin" by a black arrow pointing to its upper surface.
2.  A horizontal, Z-shaped or L-shaped element below the skin, labeled "Stringer" by a black arrow pointing to its right side. This stringer is attached to the skin via fasteners.
3.  A vertical, I-beam or C-channel shaped element, labeled "Rib web" by a black arrow pointing to its lower right side. This rib web is attached to the stringer and potentially the skin via multiple rows of circular fasteners.
Within this central diagram, several short, thin lines are depicted on the "Skin" and the "Rib web", indicating cracks. Multiple black arrows point to these crack locations, collectively labeled "Typical cracking". The "Typical shear tie" label, originating from the top-left diagram, also points to the general area of the skin and stringer attachment in this central diagram.

Below the structural diagrams, two distinct text boxes provide additional information:

The left text box has the heading "Type and possible location of MSD and MED". It contains two bullet points:
*   "MSD—critical fasteners in skin along rib attachments"
*   "MED—critical rib feet in multiple stringer bays (particularly for empennage under sonic fatigue)"

The right text box has the heading "Service or test experience of factors that influence MSD and MED (examples)". It contains two bullet points:
*   "Manufacturing defect—prestress due to assembly sequence"
*   "Sonic fatigue (empennage)"

**Contextual Linkage:** The image visually represents "Rib-to-skin attachments (MSD/MED)" as stated in its figure caption. The diagrams illustrate the structural elements involved (skin, stringer, rib web, shear tie) and show typical locations where cracking (MSD/MED) can occur, specifically at fasteners in the skin along rib attachments and at rib feet. The accompanying text boxes further define MSD and MED in this context and provide examples of factors, such as manufacturing defects (prestress due to assembly sequence) and sonic fatigue (especially in the empennage), that can influence the development of these damage types.](file_tNNDI2wQkpg/image020.png)

Figure A2-15:
Rib-to-skin attachments (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** This image illustrates and compares the characteristics of fatigue damage, specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), in two distinct aircraft structural configurations: Riveted Skin and Stringer Construction versus Integrally Stiffened Skins.

**Detailed Description:**
The image is divided vertically into two main sections by a solid black line, each detailing a different type of aircraft skin construction and its associated damage characteristics.

The left section, titled "Riveted Skin and Stringer Construction (MSD & MED)", presents three diagrams and accompanying text:
*   **Diagram (a):** Located at the top left, this is an isometric view of a section of riveted skin and stringer construction. It shows a flat skin panel with multiple parallel stringers attached to its underside via rows of fasteners. Two black arrows originate from the top left, pointing to rows of small, circular fastener holes along the stringers. Text labels associated with these arrows are "Fastener attachment to rib" and "Cracks". The cracks are depicted as small, irregular lines originating from some of the fastener holes.
*   **Diagram (b):** Positioned at the top right within the left section, this is a partial cross-sectional view of a skin panel. A circular "Drain hole" is visible in the skin. A black arrow points from the text "Drain hole crack" to an irregular crack line extending from the edge of the drain hole.
*   **Diagram (c):** Situated at the bottom right within the left section, this is a partial cross-sectional view showing a stiffener attached to a skin panel. A vertical element, labeled "Root rib, tank end, etc.", is shown above the stiffener. Two black arrows point to "Cracks" located at the interface between the stiffener and the skin, specifically at the run-out of the stiffener. Another black arrow points from the text "Crack" to a crack propagating from the stiffener run-out. The skin itself is labeled "Skin" with an arrow pointing to it.
Below these diagrams, the text states: "Inherent fail safe and crack stopper characteristics". This is followed by two bullet points:
    *   "MSD—chordwise cracks link up at"
        *   "a) Rib attachment holes"
    *   "MED—"
        *   "b) Drain or vent holes"
        *   "c) Stiffener run-outs at root rib or tank end rib"
At the very bottom left of the image, the "BOEING" logo is present.

The right section, titled "Integrally Stiffened Skins (MSD)", also presents three diagrams and accompanying text:
*   **Diagram (d):** Located at the top left, this is an isometric view of a section of integrally stiffened skin. It shows a flat skin panel with multiple parallel stiffeners that appear to be machined directly from the skin material, forming a corrugated or ribbed structure. Two black arrows point to rows of small, irregular "Cracks" located at the base of the stiffeners where they meet the skin panel.
*   **Diagram (e):** Positioned at the top right within the right section, this is a close-up isometric view of a single integrally stiffened section. It shows a stiffener with a flat top and vertical web. A black arrow points from the text "(a)" to a horizontal "Crack" propagating across the web of the stiffener.
*   **Diagram (f):** Situated at the bottom right within the right section, this is a partial cross-sectional view of an integrally stiffened skin panel. A vertical element, labeled "Root rib, tank end, etc.", is shown above the stiffened skin.
Below these diagrams, the text states: "Do not have inherent crack stopper characteristics of riveted skin and stringer construction". This is followed by two bullet points:
    *   "MSD—Chordwise cracks link up at"
        *   "d) Rib attachment holes"
        *   "e) Drain or vent holes"
        *   "f) Stringer run-outs at root rib or tank end rib"
    *   "MED—becomes MSD"

**Contextual Linkage:**
The image visually differentiates between two primary aircraft structural designs concerning their susceptibility to Widespread Fatigue Damage (WFD), specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED).
The "Riveted Skin and Stringer Construction" section illustrates how MSD (chordwise cracks linking up at rib attachment holes, as shown in diagram (a)) and MED (cracks at drain/vent holes in diagram (b) and stiffener run-outs in diagram (c)) manifest. The text indicates that this construction type possesses "Inherent fail safe and crack stopper characteristics."
In contrast, the "Integrally Stiffened Skins" section focuses solely on MSD, showing cracks at rib attachment holes (d), drain/vent holes (e), and stringer run-outs at root rib or tank end rib (f). The text explicitly states that this construction type "Do not have inherent crack stopper characteristics of riveted skin and stringer construction" and that "MED—becomes MSD" in this context, implying a higher susceptibility to widespread damage propagation. The diagrams (a) through (f) are referenced by the bullet points to specify the locations where these damage types occur for each construction method.](file_tNNDI2wQkpg/image021.png)
Figure A2-16: Typical wing and
empennage construction (MSD/MED)

4.3       WFD
Evaluation

Point 26.300 of Part-26 requires an LOV to be established
according to specified timescales for large transport aeroplanes with MTOWs
above 34 901 kg (75 000 lb). For other types, it is
recommended that by the time the highest-time aircraft of a particular model
reaches its DSG, the evaluation for each area susceptible to the development
of WFD should be completed. A typical evaluation process is shown in Figure A2-19
below. This evaluation will establish the necessary elements to determine a
maintenance programme to preclude WFD in that particular model’s aircraft
fleet. These elements are developed for each susceptible area and include:

**4.3.1   Identification
of structure potentially susceptible to WFD**

Unless already fully addressed in the existing fatigue and
damage tolerance evaluation, the TCH should identify each part of the
aircraft’s structure that is potentially susceptible to WFD for further
evaluation. A justification should be given that supports selection or
rejection of each area of the aircraft structure. DAHs for modified or repaired
structure should evaluate their structure and its effect on existing
structure.

Typical examples of structure susceptible to WFD are included
in paragraph 4.2 of this Appendix.

**4.3.2   Predicting
when WFD will occur**

(a)       Characterisation
of events leading to WFD

The fatigue process that leads to WFD is shown in Figure A2-17.
This figure is applicable to both damage that occurs in multiple sites (MSD)
and damage that occurs in similar structures at more than one location (MED).
For any susceptible structural area, it is not a question of whether WFD will
occur, but when it will occur. In Figure A2-17, the ‘when’ is illustrated
by the line titled ‘WFD (average behaviour),’ which is the point
when 50 % of the aeroplanes in a fleet would have experienced WFD in the
considered area (note that the probability density function for flight cycles
or flight hours to WFD has been depicted for reference). The WFD process
includes this phase of crack initiation and a crack growth phase. During the
crack initiation phase, which generally spans a long period of time, there is
little or no change in the basic strength capability of the structure. The
actual residual strength curve depicted in Figure A2-17 is flat, and
equal to the strength of the structure in its pristine state. However, at some
time after the first small cracks start to grow, residual strength begins to
degrade. Crack growth continues until the capability of the structure degrades
to the point of the minimum strength required by CS 25.571(b). In this
context, the line in Figure A2-17 called WFD (average
behaviour) represents a point when 50 % of the aeroplanes in a
fleet fall below the minimum strength specifications of CS 25.571(b).

![Aviation.Bot AI suggestion, not from EASA source: The image displays a graph illustrating the relationship between flight cycles/hours, residual strength, crack length, and the probability density of Widespread Fatigue Damage (WFD) occurrence.

Detailed Description:
The image presents a two-dimensional graph with a horizontal x-axis labeled "FLIGHT CYCLES / HOURS". There are two vertical y-axes: the left y-axis is labeled "RESIDUAL STRENGTH", and the right y-axis is labeled "CRACK LENGTH".

In the upper left portion of the graph, a horizontal solid line is labeled "PRISTINE STRENGTH". Below this, a solid curve, labeled "RESIDUAL STRENGTH", begins horizontally at the "PRISTINE STRENGTH" level and then gradually decreases as the "FLIGHT CYCLES / HOURS" increase.

A horizontal dashed line, labeled "MINIMUM STRENGTH REQUIRED BY CS 25.571(b)", is positioned below the "PRISTINE STRENGTH" line and intersects the declining "RESIDUAL STRENGTH" curve.

In the lower left portion, a solid curve, labeled "CRACK LENGTH", originates near the graph's origin and gradually increases with "FLIGHT CYCLES / HOURS". This curve shows a slow initial increase followed by a more rapid increase.

In the upper right portion, a bell-shaped curve, labeled "Probability density function for cycles to WFD condition", is depicted. This curve rises to a peak and then declines, representing a distribution.

A vertical dashed line extends downwards from the peak of the "Probability density function for cycles to WFD condition" curve. This dashed line intersects the "RESIDUAL STRENGTH" curve and the "CRACK LENGTH" curve. At the bottom of this dashed line, on the "FLIGHT CYCLES / HOURS" axis, the text "WFD (average behaviour)" is placed.

To the right of the vertical dashed line, a vertical solid line also extends downwards to the "FLIGHT CYCLES / HOURS" axis. At the bottom of this solid line, the text "WFD" is placed.

A horizontal dotted line extends from the intersection of the vertical dashed line and the "Probability density function for cycles to WFD condition" curve, extending leftwards across the graph.

The text "MSD/MED" is located below the "CRACK LENGTH" curve, near the bottom center of the graph.

In the top left corner, a note states: "NOTE: No scale implied. For definition use only."

Contextual Linkage:
The graph illustrates the progression of fatigue damage over time (flight cycles/hours). The "RESIDUAL STRENGTH" curve shows the structural strength degrading from its "PRISTINE STRENGTH" as "CRACK LENGTH" increases. The "MINIMUM STRENGTH REQUIRED BY CS 25.571(b)" represents a critical threshold for residual strength. The "Probability density function for cycles to WFD condition" indicates the statistical likelihood of WFD occurring. The "WFD (average behaviour)" point, marked by the vertical dashed line, corresponds to the peak of this probability function, signifying the average time at which WFD is expected to occur in a fleet. The "MSD/MED" label indicates that this graphical representation applies to both Multiple Site Damage and Multiple Element Damage scenarios.](file_tNNDI2wQkpg/image022.png)

Figure
A2-17: Effect on residual strength of developing WFD

(b)       Determination
of WFD (average behaviour) in the fleet

The time in terms of flight cycles/hours defining the WFD (average
behaviour) in the fleet should be established for each susceptible
structural area. The data to be assessed in determining the WFD (average
behaviour) includes:

—        a complete review of the service
history of the susceptible areas, to identify any occurrences of fatigue
cracking and the continuing validity of loads and mission profiles,

—        evaluation of the operational
statistics of the fleet in terms of flight hours and landings,

—        significant production variants
(material, design, assembly method, and any other change that might affect the
fatigue performance of the detail),

—        fatigue test evidence including
relevant full-scale and component fatigue and damage tolerance test data (see
subparagraph 4.3.9 and [Annex 1](#_DxCrossRefBm1926189263) for
more details),

—        teardown inspections, and

—        any fractographic analysis available.

The evaluation of the test results for the reliable
prediction of the time to when WFD might occur in each susceptible area should
include appropriate test-to-structure factors. If full-scale fatigue test
evidence is used, Figure A2‑20 below relates how that data might be
utilised in determining WFD (average behaviour). Evaluation may be
analytically determined, supported by test and, where available, service
evidence.

Regardless of whether the assessment of WFD (average
behaviour) is based on in‑service data, full-scale fatigue test
evidence, analyses, or a combination of any of these, the following should be
considered:

**4.3.3   Initial crack/damage
scenario**

This is an estimate of the size and extent of multiple
cracking expected at MSD/MED initiation. This prediction requires empirical
data or an assumption of the crack/damage locations and sequence plus a
fatigue evaluation to determine the time to MSD/MED initiation. Alternatively,
analysis can be based on either:

—        the distribution of equivalent initial
flaws, as determined from the analytical assessment of flaws found during
fatigue test and/or teardown inspections regressed to zero cycles; or

—        a distribution of fatigue damage
determined from relevant fatigue testing and/or service experience.

**4.3.4   Final cracking
scenario**

This is an estimate of the size and extent of multiple
cracking that could cause residual strength to fall to the minimum required
level as shown in A2-17. Techniques exist for 3-D elastic-plastic analysis of
such problems; however, there are several alternative test and analysis
approaches available that provide an equivalent level of safety. One such
approach is to define the final cracking scenario as a subcritical condition
(e.g. first crack at link-up). The use of a subcritical scenario reduces the
complexity of the analysis and, in many cases, will not greatly reduce the
total crack growth time, because the majority of the time taken to reach the
critical condition is generally in the initiation phase.

**4.3.5   Crack growth
calculation**

Progression of the crack distributions from the initial
cracking scenario to the final cracking scenario should be developed. These
curves can be developed:

—        *analytically*, typically based on
linear elastic fracture mechanics, or

—        *empirically*, from test or
service fractographic data.

**4.3.6   Potential for
discrete source damage (DSD)**

A structure susceptible to fatigue including MSD/MED may also
be affected by DSD due to an uncontained failure of high-energy rotating machinery
(i.e. turbine engines). At this time, there is no specific requirement to
address prior fatigue cracking in combination with DSD for certification.
Nonetheless, when assessing in‑service findings of fatigue cracking, the
additional threat posed by any potential DSD should be taken into account when
developing the corrective actions and the timescales for its implementation.

**4.3.7   Analysis methodology**

**Differences between multiple-site damage and
multiple-element damage**

Details of the approach used to characterise events leading
up to WFD may be different. The differences will largely depend on whether MSD
or MED is being considered. This is especially true for crack interaction.

(a)       Crack
interaction

MSD has the potential for strong crack interaction, and the
effect of multiple cracks on each other needs to be addressed. MED, in most
cases, does not have the same potential for strong crack interaction. The
differences between the interaction effects for MSD and MED are illustrated in
Figure A2-18.

(b)       MSD and MED
interaction

Some areas of an aeroplane are
potentially susceptible to both MSD and MED. Simultaneous occurrence of MSD
and MED is possible, even though it is not common. A comparison of inspection
start points (ISPs) or modification start points might indicate the possibility
of this occurrence. If so, the evaluation should consider the interaction
between MSD and MED.

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image is a comparative diagram illustrating the difference in crack interaction effects between Multiple-Site Damage (MSD) and Multiple-Element Damage (MED) through both graphical representations of crack growth (N vs. a) and schematic depictions of structural components with cracks.

**Detailed Description:** The image is horizontally divided into two main sections: the left section depicts "MSD" and the right section depicts "MED". Each section contains a graph in the upper portion and corresponding schematic representations of structural components in the lower portion.

**Left Section (MSD - Multiple-Site Damage):** * **Graph:** A 2D Cartesian coordinate system is shown. * The vertical Y-axis is labeled "N (cycles)". * The horizontal X-axis is labeled "a (mm)". * Three open circles with thick black outlines are positioned horizontally along the X-axis, representing distinct crack initiation sites. * Two sets of curves originate from these circles and extend upwards: * **"Multiple cracks" curves (solid lines):** These are three interconnected, bell-shaped curves forming a continuous envelope. The troughs of these curves are positioned directly above the three open circles on the X-axis, indicating crack initiation points. The curves rise between the circles, suggesting crack growth and potential interaction or link-up between adjacent cracks. These curves are labeled "Multiple cracks" with solid text. * **"Single crack" curves (dotted lines):** These are three distinct, upward-curving dotted lines, each originating from one of the open circles on the X-axis. Each curve rises steeply and then curves outwards, appearing to represent a longer fatigue life (higher N for a given 'a') compared to the "Multiple cracks" scenario. These curves are labeled "Single crack" with dotted text. * **Schematic Representations:** Below the graph, two horizontal, irregularly shaped gray structural components are shown, each containing three circular cutouts with thick black outlines, arranged in a line. * **"Single crack" scenario:** The upper schematic shows a single dotted horizontal line connecting the centers of the three circular cutouts, indicating a single crack path or a single crack growing across multiple sites. The label "Single crack" (dotted text) is positioned to the right of this schematic. * **"Multiple cracks" scenario:** The lower schematic shows three short, solid horizontal lines, each originating from the inner edge of one of the circular cutouts and extending outwards. This depicts multiple independent cracks initiating from each hole. The label "Multiple cracks" (solid text) is positioned to the right of this schematic. * **Label:** A black rectangular box centered below the schematics contains the text "MSD".

**Right Section (MED - Multiple-Element Damage):** * **Graph:** A 2D Cartesian coordinate system is shown, identical in axis labeling to the MSD graph. * The vertical Y-axis is labeled "N (cycles)". * The horizontal X-axis is labeled "a (mm)". * Three gray rectangular blocks are positioned horizontally along the X-axis, representing distinct structural elements. * Two sets of curves originate from these blocks and extend upwards: * **"Multiple cracks" curves (solid lines):** These are three distinct, upward-curving solid lines, each originating from the top edge of one of the rectangular blocks. Each curve rises steeply and then curves outwards, similar in shape to the "Single crack" curves in the MSD graph. These curves are labeled "Multiple cracks" with solid text. * **"Single crack" curves (dotted lines):** These are three distinct, upward-curving dotted lines, each originating from the top edge of one of the rectangular blocks. Each curve rises steeply and then curves outwards, similar in shape to the "Multiple cracks" curves in the MSD graph. These curves are labeled "Single crack" with dotted text. * Unlike the MSD graph, the curves in the MED graph do not interact or form a continuous envelope; they are independent. * **Schematic Representations:** Below the graph, two rows of three identical gray rectangular structural elements are shown, each representing a separate component. * **"Single crack" scenario:** The top row shows three rectangular elements. The middle element contains a single short, horizontal dotted line in its center, representing a crack. The other two elements are depicted as intact. The label "Single crack" (dotted text) is positioned to the left of this row. * **"Multiple cracks" scenario:** The bottom row shows three rectangular elements. Each of the three elements contains a single short, horizontal solid line in its center, representing a crack. The label "Multiple cracks" (solid text) is positioned to the left of this row. * **Label:** A black rectangular box centered below the schematics contains the text "MED".

**Contextual Linkage:** The image visually differentiates the crack growth behavior and interaction effects between Multiple-Site Damage (MSD) and Multiple-Element Damage (MED). For MSD, the graphs and schematics illustrate that multiple cracks originating from adjacent sites within a single component can interact, potentially leading to a reduced fatigue life (lower N for a given 'a') compared to a single crack scenario. For MED, the graphs and schematics show that cracks in separate, distinct structural elements typically do not interact, and their crack growth behavior is independent, whether a single element is cracked or multiple elements are cracked. The 'N (cycles)' axis represents the number of fatigue cycles, and the 'a (mm)' axis represents the crack size in millimeters.](file_tNNDI2wQkpg/image023.png)

Figure
A2-18: Difference between MSD and MED interaction effects

The report ‘Recommendations for Regulatory Action to Prevent
Widespread Fatigue Damage in the Commercial Aeroplane Fleet’, Revision A,
dated June 29, 1999 (a report of the AAWG for the ARAC’s Transport
Aircraft and Engine Issues Group), discusses two Round Robin exercises
developed by the TCHs to provide insight into their respective methodologies.
One outcome of the exercises was an identification of key assumptions or
methods that had the greatest impact on the predicted WFD behaviour. These
assumptions were:

—        the flaw sizes assumed at initiation of
crack growth phase of analysis;

—        material properties used (static,
fatigue, fracture mechanics);

—        ligament failure criteria;

—        crack growth equations used;

—        statistics used to evaluate the fatigue
behaviour of the structure (e.g. time to crack initiation);

—        methods of determining the structural
modification point (SMP);

—        detectable flaw size assumed;

—        initial distribution of flaws; and

—        factors used to determine bound
behaviour as opposed to mean behaviour.

When considering MED, where interaction between cracks in
different elements is not a factor, the following should be considered:

(1)       In a
structure containing large numbers of similar elements, there is not normally
a high probability that, after a crack initiates in an element, a second crack
will initiate in the element right next to it. If this does happen, however,
the consequences to the overall structure may be severe. This is because
having two structural members fail right next to each other can completely
negate any ability of the structure to tolerate additional damage.
Consequently, when performing an evaluation, applicants should make
conservative assumptions and assume failures to be adjacent to each other.

(2)       When an
element fails completely, the load that has to be redistributed onto the
non-failed structure can be large and can have a significant impact on the
strength of both the cracked and uncracked structure; therefore, the effects
of load redistribution must be included in the evaluation.

(d)       Establishing
maintenance actions

The following parameters are developed from paragraphs 4.3.2
to 4.3.7 above, and are necessary to establish an MSD/MED maintenance
programme for the area under investigation.

Fatigue damage is the gradual deterioration of a material
subjected to repeated loads. This gradual deterioration is a function of use
and can be statistically quantified. The term ‘WFD’ is used, and can be
statistically quantified, at the end of the deterioration process when the
structure is no longer able to carry the residual strength loads. WFD can
never be absolutely precluded because there is always some probability, no
matter how small, that it will occur. Therefore, modifying or replacing structure
at a predetermined, analytically-derived time stated in flight cycles or
flight hours, minimises the probability of having WFD in the fleet.
Modification or replacement is the most reliable method for precluding WFD.
The point at which a modification is undertaken is referred to as the
‘structural modification point (SMP)’ and it is illustrated in Figure 2-1 of [Annex 2](#_DxCrossRefBm1926189262). The SMP is generally a
fraction of the number representing the point in time when WFD (average
behaviour) will occur, and should result in the same reliability as a
successful two-lifetime fatigue test. This level of reliability for setting
the SMP is acceptable if MSD or MED inspections are shown to be effective in
detecting cracks. If the inspections are effective, they should be implemented
before the SMP. The implementation times for these inspections are known as
the ‘inspection start points (ISPs)’. Repeat inspections are usually necessary
to maintain this effectiveness in detecting cracks. If MSD or MED inspections
are not effective in detecting cracks, then the SMP should be set at the time
of ISP. For the purposes of this AMC, an inspection is effective if, when
performed by properly trained maintenance personnel, it will readily detect
the damage in question[[9]](#_ftn9). The
SMP should minimise the extent of cracking in the susceptible structural area
in a fleet of affected aeroplanes. In fact, if this point is appropriately
determined, a high percentage of aeroplanes would not have any MSD or MED by
the time the SMP is reached.

Due to the redundant nature of semi-monocoque structures, MED
can be difficult to manage in a fleet environment. This stems from the fact
that most aircraft structures are built-up in nature, and that makes the
visual inspection of the various layers difficult. Also, visual inspections
for MED typically rely on internal inspections, which may not be practical at
the frequency necessary to preclude MED due to the time required to gain
access to the structure. However, these issues are dependent on the specific
design involved and the amount of damage being considered. In order to
implement a viable inspection programme for MED, static stability must be
maintained at all times and there should be no MED concurrent with MSD in a
given structural area.

**4.3.8   Inspection
start point (ISP)**

This is the point at which inspection starts if a monitoring
period is used. Inspection is not practical for all applications and cannot
replace the SMP. The ISP is determined through a statistical analysis of crack
initiation based on fatigue testing, teardown, or service experience of
similar structural details. It is assumed that the ISP is equivalent to a
lower bound value with a specific probability in the statistical distribution
of cracking events. Alternatively, the ISP may be established by applying
appropriate factors to the average behaviour.

When inspections are determined to be effective, it is
necessary to establish when those inspections should start. This point is
illustrated in Figure 2-1. The start point is determined through a statistical
analysis of crack initiation based on fatigue testing, teardown, or in-service
experience of similar structure. The ISP is assumed to be equivalent to a
lower-bound value with a specific probability in the statistical distribution
of cracking events. Alternatively, an ISP may be established by applying appropriate
factors to the number representing WFD (average behaviour). (e.g.
for aluminium alloy structure, dividing the full-scale test result by a factor
of 3).

For inspection intervals, see point 4.3.10.

**4.3.9   Structural
modification point (SMP)**

The SMP should be established as a point in time when
structures should be modified or replaced to prevent WFD from occurring. This
is typically established by:

—        calculating when WFD would first occur
in the structure (predicted using the WFD (average behaviour)),

—        setting a time before the predicted
occurrence of WFD to perform modifications or replacements that will prevent
it.

The applicant should demonstrate that the proposed SMP
established during the evaluation has the same confidence level as current
regulations require for new certification. In lieu of other acceptable
methods, the SMP for aluminium alloy structures can be established as a point
reduced from the WFD (average behaviour), based on the viability of
inspections in the monitoring period. The SMP may be determined by dividing
the number representing the timing when WFD (average behaviour)will
occur by a factor of 2 if there are effectiveinspections, or by a factor of 3
if inspections are not effective. For other materials such as high‑strength
steel alloys, larger scatter factors may be necessary to account for increased
variability in fatigue performance.

An aircraft should not be operated past the SMP unless the structure
is modified or replaced, or unless additional approved data is provided that
would extend the SMP. However, if during the structural evaluation for WFD, a
TCH/DAH finds that the flight cycles and/or flight hours SMP for a particular
structural detail have been exceeded by one or more aircraft in the fleet, the
TCH/DAH should expeditiously evaluate selected high-time aircraft in the fleet
to determine their structural condition. From this evaluation, the TCH/DAH
should notify the competent authorities and propose appropriate service
actions.

A DAH may find that the SMP for a particular structural area
has been exceeded by one or more aeroplanes in the fleet. In that case, the
DAH should expedite the evaluation of those high-time aeroplanes to determine
their structural condition, notify EASA and propose appropriate maintenance
actions specific to those aeroplanes.

The initial SMP may be adjusted based on the following:

(a)       The tasks
necessary to extend an SMP may include any or all of the following:

(1)       Additional
fatigue or residual strength tests, or both, on a full-scale aeroplanestructure
or a full-scale component followed by detailed inspections and analyses.

(2)       Fatigue
tests of new structure orstructure from in-service aeroplaneson a smaller
scale than full component tests (i.e. subcomponent or panel tests, or both).
If a subcomponent test is used, the SMP would be extended only for that
subcomponent.

(3)       Teardown
inspections (destructive) on structural components that have been removed from
service.

(4)       Teardowninspections
(non-destructive) accomplished by selected, limited disassembly and subsequent
reassembly of specific areas of high-time aeroplanes.

(5)       Analysis of
in-service data (e.g. inspections) from a statistically significant number of
aeroplanes.

(b)       If cracks
are found in the structural detail for which the evaluation was done during
either the monitoring period or the modification programme, the SMP should be
re-evaluated to ensure that the SMP does in fact provide the required
confidence level. If it is shown that the required confidence level is not
being met, the SMP should be adjusted and the adjustment reflected in the appropriate
SBs to address the condition of the fleet. Additional regulatory action may be
required.

**4.3.10 Inspection
interval and method**

An interval should be chosen to provide a sufficient number
of inspections between the ISP and the SMP so that there is high confidence
that no MSD/MED condition will reach the final cracking scenario without
detection. The interval between inspections depends on the detectable crack
size, the critical crack lengths and the probability that the cracks will be
detected with the specific inspection method. Conservative scenarios should be
assumed for developing the inspection interval unless other assumptions can be
consistently supported by test and service experience. If the crack cannot be
detected, the SMP must be re-evaluated to ensure there is a high confidence
level that no aircraft will develop MSD/MED before modification.

**4.4      Evaluation
of maintenance actions**

For all areas that have been identified as susceptible to
MSD/MED, the current maintenance programme should be evaluated to determine whether
adequate structural maintenance and inspection programmes exist to safeguard
the structure against unanticipated cracking or other structural degradation.
The evaluation of the current maintenance programme typically begins with the
determination of the SMP for each area.

Each area should then be reviewed to determine the current
maintenance actions and compare them to the maintenance needs established in
this evaluation. Issues to be considered include the following:

(a)       Determine
the inspection requirements (method, inspection start point, and repeat
interval) of the inspection for each susceptible area (including that
structure that is expected to arrest cracks) that is necessary to maintain the
required level of safety.

(b)       Review the
elements of the existing maintenance programmes already in place.

For susceptible areas approaching the SMP, where the SMP will
not be increased or for areas that cannot be reliably inspected, a programme
should be developed and documented that provides for replacement or
modification of the susceptible structural area.

**4.4.1   Period of
WFD evaluation validity**

At whatever point the WFD evaluation is made, it should
support the LOV of the maintenance programme. Consistent with the use of test
evidence to support individual SMPs, as described above in paragraph 4.3.9,
the LOV of the maintenance programme should be based on fatigue test evidence.
For an existing ageing aircraft type, theinitial WFD evaluation of the
complete airframe will typically cover a significant forward estimation of the
projected aircraft usage beyond its DSG, also known as the ‘proposed ESG’and
is effectively a proposed LOV.Typically, an evaluation through an additional 25 %
of the DSG would provide a realistic forecast, with reasonable planning time
for necessary maintenance action. However, it may be appropriate to adjust the
evaluation validity period depending on issues such as:

(a)       the
projected useful life of the aircraft at the time of the initial evaluation;

(b)       current NDI
technology; and

(c)       airline
advance planning requirements for introduction of new maintenance and
modification programmes, to provide sufficient forward projection to identify
all likely maintenance/modification actions essentially as one package.

Upon completion of the evaluation and publication of the
revised maintenance requirements, the ‘proposed ESG’ becomes the LOV.

*Note*: This
assumes that all other aspects of the maintenance programme that are required
to support the LOV (such as SSID, CPCP, etc.) are in place and have been
evaluated to ensure they too remain valid up to the LOV.

![Aviation.Bot AI suggestion, not from EASA source: A flowchart illustrating the sequential steps and decision points in the aircraft Widespread Fatigue Damage (WFD) evaluation process.

Detailed Description:
The image displays a flowchart titled "Figure A2-19: Aircraft WDF evaluation process". Below the flowchart, two notes are provided:
*   "NOTES:"
*   "1. Fatigue cracking is defined as likely if the factored fatigue life is less than the projected ESG of the aircraft at time of WFD evaluation."
*   "2. The operational life is the projected ESG of the aircraft at time of WFD Evaluation. (See 4.4.1)."

The flowchart consists of several interconnected shapes, all with black borders and black text on a white background, indicating a process flow:

1.  **Initial Process Box (Rectangle):** Located at the top, it contains the text: "REVIEW STRUCTURAL AREAS POTENTIALLY SUSCEPTIBLE TO WFD (See 4.3.1)". A downward arrow extends from its bottom center.

2.  **Second Process Box (Rectangle):** Positioned directly below the initial box, it contains the text: "FOR EACH AREA, DETERMINE THE WFD AVERAGE BEHAVIOUR IN THE FLEET (See 4.3.2 onwards)". A downward arrow extends from its bottom center.

3.  **First Decision Box (Oval):** Located below the second process box, it contains the question: "IS NATURAL FATIGUE CRACKING LIKELY ¹ WITHIN OPERATIONAL LIFE ²".
    *   A horizontal arrow labeled "NO" extends from its right side, pointing to a "STOP" box.
    *   A downward arrow extends from its bottom center, leading to the next process box.

4.  **Stop Box (Rectangle):** Located to the right of the first decision box, it contains the text: "STOP". This box represents a terminal point in the process.

5.  **Third Process Box (Rectangle):** Positioned directly below the first decision box (following the implicit "YES" path), it contains the text: "ESTIMATE ALLOWABLE FATIGUE DAMAGE SCENARIO FOR LIMIT LOAD (See 4.3.4)". A downward arrow extends from its bottom center.

6.  **Second Decision Box (Oval):** Located below the third process box, it contains the question: "FATIGUE DAMAGE SCENARIO DETECTABLE PRIOR TO MAXIMUM ALLOWABLE EXTENT UNDER LIMIT LOAD".
    *   A horizontal arrow labeled "NO" extends from its right side, pointing to a process box for establishing SMP and terminating action.
    *   A downward arrow labeled "YES" extends from its bottom center, pointing to a process box for establishing ISP, inspection interval, and method.

7.  **Fourth Process Box (Rectangle):** Located to the right of the second decision box, it contains the text: "ESTABLISH THE SMP AND TERMINATING ACTION (See 4.3.10)". This box is a destination for the "NO" path from the second decision.

8.  **Fifth Process Box (Rectangle):** Positioned directly below the second decision box (following the "YES" path), it contains the text: "ESTABLISH ISP, INSPECTION INTERVAL AND METHOD AND SCHEDULE FOR TERMINATING ACTION (See 4.3.9//10/11)". This box is a destination for the "YES" path from the second decision.

Contextual Linkage:
The flowchart illustrates the logical flow for evaluating aircraft for Widespread Fatigue Damage. It starts with an initial review and determination of WFD behavior. A critical decision point is whether natural fatigue cracking is anticipated within the aircraft's operational life, with specific definitions provided in the accompanying notes for "likely" cracking (Note 1) and "operational life" (Note 2, referencing section 4.4.1). If cracking is not likely, the process terminates. If cracking is likely, the evaluation proceeds to estimate allowable fatigue damage scenarios (referencing section 4.3.4). A subsequent decision assesses the detectability of this damage. If the damage is not detectable before reaching its maximum allowable extent, the process dictates establishing a Structural Maintenance Program (SMP) and terminating action (referencing section 4.3.10). If the damage is detectable, the process leads to establishing an Initial Structural Program (ISP), inspection intervals, methods, and a schedule for terminating action (referencing sections 4.3.9, 4.3.10, and 4.3.11).

"mermaid
graph TD
    A["REVIEW STRUCTURAL AREAS POTENTIALLY SUSCEPTIBLE TO WFD (See 4.3.1)"] --> B["FOR EACH AREA, DETERMINE THE WFD AVERAGE BEHAVIOUR IN THE FLEET (See 4.3.2 onwards)"]
    B --> C{{"IS NATURAL FATIGUE CRACKING LIKELY ¹ WITHIN OPERATIONAL LIFE ²"}}
    C -- "NO" --> D["STOP"]
    C --> E["ESTIMATE ALLOWABLE FATIGUE DAMAGE SCENARIO FOR LIMIT LOAD (See 4.3.4)"]
    E --> F{{"FATIGUE DAMAGE SCENARIO DETECTABLE PRIOR TO MAXIMUM ALLOWABLE EXTENT UNDER LIMIT LOAD"}}
    F -- "NO" --> G["ESTABLISH THE SMP AND TERMINATING ACTION (See 4.3.10)"]
    F -- "YES" --> H["ESTABLISH ISP, INSPECTION INTERVAL AND METHOD AND SCHEDULE FOR TERMINATING ACTION (See 4.3.9//10/11)"]
"
](file_tNNDI2wQkpg/image024.jpg)

NOTES:

1.         Fatigue
cracking is defined as likely if the factored fatigue life is less than the
projected ESG of the aircraft at time of WFD evaluation.

2.         The
operational life is the projected ESG of the aircraft at time of WFD
Evaluation. (See 4.4.1).

Figure
A2-19: Aircraft WDF evaluation process

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** This flowchart outlines the systematic process for evaluating and determining the estimated average behaviour of Widespread Fatigue Damage (WFD) based on full-scale fatigue test data, leading to specific Limit of Validity (LOV) calculations or the requirement for inspection and modification programmes.

**Detailed Description:** The image displays a flowchart with a white background, illustrating a decision-making process. All primary process and decision nodes are represented by rectangular boxes featuring a distinctive double horizontal line at their top edge. Arrows indicate the directional flow between these nodes.

1.  **Initial Input:** At the top, the process begins with a rectangular box labeled "FULL-SCALE FATIGUE TEST DATA". An arrow points vertically downwards from this box.
2.  **First Decision Point:** This arrow leads to a rectangular decision box, labeled "TEAR DOWN?".
    *   From "TEAR DOWN?", a horizontal arrow labeled "NO¹" extends to the right, connecting to another rectangular decision box.
    *   From "TEAR DOWN?", a vertical arrow labeled "YES" points downwards, connecting to a third rectangular decision box.
3.  **Second Decision Point (Conditional on "TEAR DOWN?" being "YES"):** This rectangular box is labeled "MSD/MED FINDINGS DURING TEST/TEARDOWN?".
    *   From this box, a horizontal arrow labeled "NO" points left, then turns downwards.
    *   From this box, a horizontal arrow labeled "YES" points right, then turns downwards.
4.  **Third Decision Point (Conditional on "TEAR DOWN?" being "NO¹"):** This rectangular box is labeled "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?".
    *   From this box, a horizontal arrow labeled "NO" points left, then turns downwards.
    *   From this box, a horizontal arrow labeled "YES" points right, then turns downwards.

5.  **WFD Average Behaviour Determination Section:** A large, light purple rectangular area spans horizontally across the lower middle section of the diagram. This area is explicitly labeled "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM" in bold, underlined text. Contained within this purple area are three distinct rectangular boxes, each with a double top border:
    *   The "NO" path from "MSD/MED FINDINGS DURING TEST/TEARDOWN?" leads to the leftmost box, labeled "TEST LIFE".
    *   The "YES" path from "MSD/MED FINDINGS DURING TEST/TEARDOWN?" leads to the middle box, labeled "TEST LIFE plus CRACK GROWTH LIFE³".
    *   The "NO" path from "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?" also leads to the middle box, "TEST LIFE plus CRACK GROWTH LIFE³".
    *   The "YES" path from "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?" leads to the rightmost box, labeled "TEST LIFE minus CRACK GROWTH LIFE³".

6.  **Final Outcomes:**
    *   From the "TEST LIFE" box (within the purple area), a vertical arrow points downwards to a rectangular box with a single border, labeled "LOV = Test Life/2 if all areas inspectable".
    *   From the "TEST LIFE" box, another vertical arrow points downwards to a large rectangular box with a double top border, labeled "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED (See 4.3.7 onward)".
    *   From the "TEST LIFE plus CRACK GROWTH LIFE³" box (within the purple area), a vertical arrow points downwards to the "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED" box.
    *   From the "TEST LIFE minus CRACK GROWTH LIFE³" box (within the purple area), a vertical arrow points downwards to the "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED" box.

7.  **Notes:** The flowchart incorporates three superscripted numbers (¹, ², ³) that refer to definitions provided externally:
    *   The superscript "¹" on the "NO" path from "TEAR DOWN?" refers to: "ASSUMED STATE AT END OF TEST: Best estimate of non-detected damage from inspection method used at the end of the test or during teardown."
    *   The superscript "²" within the "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?" box refers to: "CRITICAL CRACK LENGTH: First link-up of adjacent cracks at limit load (locally) or an adequate level of large damage capability."
    *   The superscript "³" within the "TEST LIFE plus CRACK GROWTH LIFE³" and "TEST LIFE minus CRACK GROWTH LIFE³" boxes refers to: "CRACK GROWTH LIFE: Difference between assumed or actual state at the end of the test and critical crack length."

**Contextual Linkage:** This flowchart details the methodology for assessing WFD average behaviour, starting with "FULL-SCALE FATIGUE TEST DATA". The process branches based on whether a "TEAR DOWN?" analysis is performed. If a teardown occurs, the presence or absence of "MSD/MED FINDINGS DURING TEST/TEARDOWN?" dictates whether the estimated WFD average behaviour is "TEST LIFE" or "TEST LIFE plus CRACK GROWTH LIFE³". If no teardown is performed, the evaluation of "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?" determines if the estimated WFD average behaviour is "TEST LIFE plus CRACK GROWTH LIFE³" or "TEST LIFE minus CRACK GROWTH LIFE³". These determinations then lead to either a specific calculation for the Limit of Validity (LOV) as "Test Life/2" (applicable only if the behaviour is "TEST LIFE" and all areas are inspectable) or, more broadly, to the requirement for an "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME". The numbered notes provide essential definitions for terms like assumed non-detected damage, critical crack length, and crack growth life, which are fundamental to understanding the evaluation criteria.

"mermaid
graph TD
    A["FULL-SCALE FATIGUE TEST DATA"]
    B["TEAR DOWN?"]
    C["MSD/MED FINDINGS DURING TEST/TEARDOWN?"]
    D["DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?"]

subgraph "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM"
        direction LR
        E["TEST LIFE"]
        F["TEST LIFE plus CRACK GROWTH LIFE³"]
        G["TEST LIFE minus CRACK GROWTH LIFE³"]
    end

H["LOV = Test Life/2 if all areas inspectable"]
    I["INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED (See 4.3.7 onward)"]

A --> B
    B -- "NO¹" --> D
    B -- "YES" --> C

C -- "NO" --> E
    C -- "YES" --> F

D -- "NO" --> F
    D -- "YES" --> G

E --> H
    E --> I
    F --> I
    G --> I
"
](file_tNNDI2wQkpg/image025.jpg)

if !mso?

<table cellpadding="0" cellspacing="0" width="100%">
<tr>
<td>
<div>
1    ASSUMED
 STATE AT END OF TEST: Best estimate of non-detected damage from
 inspection method used at the end of the test or during teardown.
2    CRITICAL
 CRACK LENGTH: First link-up of adjacent cracks at limit load
 (locally) or an adequate level of large damage capability. 
3    CRACK
 GROWTH LIFE: Difference between assumed or actual state at the
 end of the test and critical crack length.
Figure
 A2-20: Use of fatigue test and teardown information to determine WFD
 average behaviour
</div>
<?if !mso?></td>
</tr>
</table>

5.         DOCUMENTATION

Any person seeking approval of an LOV of an aircraft type
design should develop a document containing all the necessary ISPs, inspection
procedures, replacement times, SMPs, and any other maintenance actions
necessary to preclude WFD, and to support the LOV. That person must revise the
SSID or ALS as necessary, and/or prepare SBs that contain the aforementioned
maintenance actions. Since WFD is a safety concern for all operators of older
aircraft, EASA will make mandatory the identified inspection and modification
programmes. In addition, EASA may consider separate AD action to address any SBs
or other service information publications revised or issued as a result of
in-service MSD/MED findings resulting from implementation of these programmes.

The following items should be contained in the front of the documentation
supporting the LOV:

(a)       identification
of the variants of the basic aircraft type to which the document relates;

(b)       summary of
the operational statistics of the fleet in terms of hours and flights;

(d)       the types
of operations for which the inspection programme is considered valid;

(e)       reference
to documents giving any existing inspections, or modification of parts or
components; and

(f)        the LOV of
the maintenance programme in terms of flight cycles or flight hours or both as
appropriate to accommodate variations in usage.

The document should contain at least the following
information for each critical part or component:

(a)       description
of the primary structure susceptible to WFD;

(b)       details of
the monitoring period (ISP, repeat inspection interval, SMP, inspection method
and procedure (including crack size, location and direction) and alternatives)
when applicable;

(d)       any
mandatory modification or replacement of the structural element;

(e)       SBs (or
other service information publications) revised or issued as a result of
in-service findings resulting from the WFD evaluations (added as a revision to
the initial WFD document); and

(f)        guidance
to the operator on which inspection findings should be reported to the
TCH/DAH, and appropriate reporting forms and methods of submission.

6.         REPORTING
REQUIREMENTS

Operators, TCHs and STCHs are required to report in
accordance with various regulations (e.g. point 21.A.3A, and point 145.A.60). The
regulations to which this AMC relates do not require any reporting
requirements in addition to the current ones. Due to the potential threat to
structural integrity, the results of inspections must be accurately documented
and reported in a timely manner to preclude the occurrence of WFD. The current
system of operator and TCH communication has been useful in identifying and resolving
a number of issues that can be classified as WFD concerns. MSD/MED has been
discovered via fatigue testing and in-service experience. TCHs have been
consistent in disseminating related data to operators to solicit additional
service experience. However, a more thorough means of surveillance and
reporting is essential to preclude WFD.

When damage is found while conducting an approved MSD/MED
inspection programme, or at the SMP where replacement or modification of the
structure is occurring, the TCHs, STCHs and the operators need to ensure that
greater emphasis is placed on accurately reporting the following items:

(a)       a
description (with a sketch) of the damage, including crack length,
orientation, location, flight cycles/hours, and condition of structure;

(b)       results of
follow-up inspections by operators that identify similar problems on other
aircraft in the fleet;

(c)       findings
where inspections accomplished during the repair or replacement/modification
identify additional similar damage sites; and

(d)       adjacent
repairs.

Operators must report all cases of MSD/MED to the TCH, STCH
or the competent authority as appropriate, irrespective of how frequently such
cases occur. Cracked areas from in-service aircraft (damaged structure) may be
needed for detailed examination. Operators are encouraged to provide
fractographic specimens whenever possible. Aeroplanes undergoing heavy
maintenance checks are perhaps the most useful sources for such specimens.

Operators should remain diligent in the reporting of
potential MSD/MED concerns not identified by the TCH/DAH. Indications of a
developing MSD/MED problem may include:

(a)       damage at
multiple locations in similar adjacent details;

(b)       repetitive
part replacement; or

Documentation will be provided by the TCH and STCH as
appropriate to specify the required reporting format and time frame,
supporting the mandatory reporting regulations (e.g. point 21.A.3A of Part 21,
point 145.A.60 of Part-145). The data will be reviewed by the TCH or STCH,
operator(s), and EASA to evaluate the nature and magnitude of the problem and
to determine the appropriate corrective action.

7.         WFD
EVALUATION FOR STRUCTURAL MODIFICATIONS AND REPAIRS

TCHs of aeroplanes subject to the point 26.303 of Part-26
requirements for an LOV should perform WFD evaluations to assess all the
applicable existing structure and the effect of future changes on the LOV.

The WFD evaluations of this AMC do not apply retroactively to
existing STCH’s modifications, nor to existing repairs. Future changes and
repairs need to take into account the applicable certification basis, and
applicants should consider the guidance of the applicable ACJ and AMC as
discussed in paragraph 8 of this AMC. The DTEs for compliance with points
26.307, 26.308, 26.333 and 26.334 of Part-26 do not have to consider WFD (average
behaviour), or the related SMP and ISP.

In cases where a new DTE is performed by DAHs to comply with
points 26.333 and 26.334 of Part-26 for existing changes or for new changes or
repairs, according to CS 25.571 Amdt 18 or earlier amendments, the DTE and
development of DTIs should take into account the cracking scenarios that could
reasonably be expected to occur in the remaining operational lifetime of an
aeroplane into which the repair or modification is, or may be, incorporated.

8.         RESPONSIBILITYFOR
WFD EVALUATION, ESTABLISHING THE LOV AND IMPLEMENTATION OFTHE LOV AND
MAINTENANCE ACTIONS

The primary responsibility is with the DAH to perform the
analyses and supporting tests. However, it is expected that if extensive
maintenance actions are necessary, the practicality of their implementation
will be evaluated in a cooperative effort between the operators and TCHs/DAHs,
with participation of EASA.

The TCH is responsible for proposing and submitting an LOV in
the ALS for approval.

*Note*:In
some cases, the ALS may already contain an LOV which is approved in accordance
with a regulation of another authority. There may also be other potentially
more restrictive limitations on the validity of maintenance programmes. For
these cases, when the TCH needs to publish the LOV as required by point 26.303
of Part-26, this LOV and its relationship with the existing or superseded
limitation should be clearly described so that no operator will exceed the
most restrictive applicable limit on the general validity of the maintenance
programme.

The operator is responsible for implementing the LOV in their
maintenance programme.

*Note*:The
LOV does not supersede or allow operations beyond any lower limitation
applicable to the individual aeroplane and the components controlled by the
maintenance programme.

[Amdt 20/2]

[Amdt 20/20]

[Amdt
20/22]

---

[[9]](#_ftnref9)     The
cracking identified in the FAA Airworthiness Directive (AD) 2002-07-09 is an
example of the type of cracking that MSD inspections are effective in
detecting. These cracks grow from the fastener holes in the lower row of the
lower skin panel in such a way that the cracking is readily detectable using
NDI methods. The cracking identified in the FAA AD 2002-07-08 is an example of
places where MSD inspections are not effective. These cracks grow in the outer
surface and between the fastener holes in the lower row of the lower skin panel
in such a way that the cracking is not readily detectable using NDI methods.
Modification is the only option to address this type of cracking.

CS ADR-DSN.Q.852 Marking and lighting of overhead wires, cables, supporting towers, etc.

ED Decision 2022/006/R

(a) Marking: The wires, cables, etc. to be marked should be equipped with markers; the supporting tower should be coloured.

(b) Marking by colours: The supporting towers of overhead wires, cables, etc. that require marking should be marked in accordance with [CS ADR-DSN.Q.845(b)](#_DxCrossRefBm1800047822), except that the marking of the supporting towers may be omitted when they are lighted by high-intensity obstacle lights by day.

(2) A marker displayed on an overhead wire, cable, etc., should be spherical and have a diameter of not less than 60 cm.

(3) The spacing between two consecutive markers, or between a marker and a supporting tower, should be appropriate to the diameter of the marker. The spacing should normally not exceed:

(i) 30 m where the marker diameter is 60 cm, increasing progressively with increase of the marker diameter to:

(ii) 35 m where the marker diameter is 80 cm; and

(iii) further progressive increases to a maximum of 40 m where the marker diameter is of at least 130 cm.

Where multiple wires, cables, etc., are involved, a marker should be located not lower than the level of the highest wire at the point marked.

(d) Lighting:

(1) High-intensity obstacle lights, Type B, should be used to indicate the presence of the tower supporting overhead wires, cables, etc. where:

(i) a safety assessment indicates such light to be essential for the recognition of the presence of wires, cables, etc.; or

(ii) it has not been found practicable to install marker on the wires, cables, etc.

(2) Where high-intensity obstacle lights, Type B, are used, they should be located at three levels:

(i) at the top of the tower;

(ii) at the lowest level of the catenary of the wires or cables; and

(iii) at approximately midway between these two levels.

(4) The installation setting angles for high-intensity obstacle lights, Types B, should be in accordance with Table Q-5.

<table border="1" cellpadding="0" cellspacing="0" width="100%"><thead><tr><td valign="top" width="25%">(1)</td><td valign="top" width="9%">(2)</td><td valign="top" width="10%">(3)</td><td valign="top" width="13%">(4)</td><td valign="top" width="13%">(5)</td><td valign="top" width="13%">(6)</td><td valign="top" width="12%">(7)</td></tr><tr><td rowspan="2" valign="top" width="25%">Light type</td><td rowspan="2" valign="top" width="9%">Colour</td><td rowspan="2" valign="top" width="10%">Signal type/ (Flash Rate)</td><td colspan="3" valign="top" width="41%">Peak intensity (cd) at given Background Luminance (b)</td><td rowspan="2" valign="top" width="12%">Light Distribution Table</td></tr><tr><td valign="top" width="13%">Day(Above 500 cd/m2)</td><td valign="top" width="13%">Twilight(50-500 cd/m2)</td><td valign="top" width="13%">Night (Below 50 cd/m2)</td></tr></thead><tr><td valign="top" width="25%">Low-intensity Type A (fixed obstacle)</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">10</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type B (fixed obstacle)</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">32</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type C (mobile obstacle)</td><td valign="top" width="9%">Yellow/Blue (a)</td><td valign="top" width="10%">Flashing(60-90 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">40</td><td valign="top" width="13%">40</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity Type D (follow-me vehicle)</td><td valign="top" width="9%">Yellow</td><td valign="top" width="10%">Flashing(60-90 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">200</td><td valign="top" width="13%">200</td><td valign="top" width="12%">Table Q-2</td></tr><tr><td valign="top" width="25%">Low-intensity, Type E</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Flashing(c)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">32</td><td valign="top" width="12%">Table Q-2(Type B)</td></tr><tr><td valign="top" width="25%">Medium-intensity Type A</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(20-60 fpm)</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">Medium-intensity Type B</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Flashing(20-60 fpm)</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">Medium-intensity Type C</td><td valign="top" width="9%">Red</td><td valign="top" width="10%">Fixed</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">N/A</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">High-intensity Type A</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(40-60 fpm)</td><td valign="top" width="13%">200 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td valign="top" width="25%">High-intensity Type B</td><td valign="top" width="9%">White</td><td valign="top" width="10%">Flashing(40-60 fpm)</td><td valign="top" width="13%">100 000</td><td valign="top" width="13%">20 000</td><td valign="top" width="13%">2 000</td><td valign="top" width="12%">Table Q-3</td></tr><tr><td colspan="7" valign="top" width="100%">(a) <a href="#_DxCrossRefBm1800047833">CS ADR-DSN.Q.850(b)</a>(b) For flashing lights, effective intensity as determined in accordance with ICAO Doc 9157, Aerodrome Design Manual, Part 4, Visual Aids.(c) For wind turbine application, to flash at the same rate as the lighting on the nacelle.</td></tr></table>

Table Q-1. Characteristics of obstacle lights

<table border="1" cellpadding="0" cellspacing="0" width="100%"><thead><tr><td rowspan="2" width="22%">Minimum intensity (a)</td><td rowspan="2" width="22%">Maximum intensity (a)</td><td colspan="2" width="41%">Vertical beam spread (f)</td></tr><tr><td width="20%">Minimum beam spread</td><td width="20%">Intensity</td></tr></thead><tr><td valign="top" width="12%">Type A</td><td valign="top" width="22%">10 cd (b)</td><td valign="top" width="22%">N/A</td><td valign="top" width="20%">10°</td><td valign="top" width="20%">5 cd</td></tr><tr><td valign="top" width="12%">Type B</td><td valign="top" width="22%">32 cd (b)</td><td valign="top" width="22%">N/A</td><td valign="top" width="20%">10°</td><td valign="top" width="20%">16 cd</td></tr><tr><td valign="top" width="12%">Type C</td><td valign="top" width="22%">40 cd (b)</td><td valign="top" width="22%">400 cd</td><td valign="top" width="20%">12(d)</td><td valign="top" width="20%">20 cd</td></tr><tr><td valign="top" width="12%">Type D</td><td valign="top" width="22%">200 cd (c)</td><td valign="top" width="22%">400 cd</td><td valign="top" width="20%">N/A(e)</td><td valign="top" width="20%">N/A</td></tr><tr><td colspan="5" valign="top" width="100%">Note: This table does not include recommended horizontal beam spreads. <a href="#_DxCrossRefBm1800047824">CS ADR-DSN.Q.846(c)</a> requires 360° coverage around an obstacle. Therefore, the number of lights needed to meet this requirement will depend on the horizontal beam spreads of each light as well as the shape of the obstacle. Thus, with narrower beam spreads, more lights will be required.(a) 360° horizontal. For flashing lights, the intensity is read into effective intensity, as determined in accordance with ICAO, Aerodrome Design Manual, Part 4, Visual Aids.(b) Between 2 and 10° vertical. Elevation vertical angles are referenced to the horizontal when the light is levelled.(c) Between 2 and 20° vertical. Elevation vertical angles are referenced to the horizontal when the light is levelled.(d) Peak intensity should be located at approximately 2.5° vertical.(e) Peak intensity should be located at approximately 17° vertical.(f) Beam spread is defined as the angle between the horizontal plan and the directions for which the intensity exceeds that mentioned in the ‘intensity’ column.</td></tr></table>

Table Q-2. Light distribution for low-intensity obstacle lights

<table border="1" cellpadding="0" cellspacing="0" width="931"><thead><tr><td rowspan="4" valign="top" width="84">Benchmark intensity</td><td colspan="6" valign="top" width="438">Minimum requirements</td><td colspan="5" valign="top" width="410">Recommendations</td></tr><tr><td colspan="4" valign="top" width="273">Vertical elevation angle (b)</td><td colspan="2" rowspan="2" valign="top" width="164">Vertical beam spread (c)</td><td colspan="3" valign="top" width="246">Vertical elevation angle (b)</td><td colspan="2" rowspan="2" valign="top" width="164">Vertical beam spread (c)</td></tr><tr><td colspan="3" valign="top" width="186">0°</td><td valign="top" width="88">-1°</td><td valign="top" width="82">0°</td><td valign="top" width="82">-1°</td><td valign="top" width="82">-10°</td></tr><tr><td valign="top" width="91">Minimum average intensity (a)</td><td valign="top" width="91">Minimum intensity (a)</td><td colspan="2" valign="top" width="91">Minimum intensity (a)</td><td valign="top" width="82">Minimum beam spread</td><td valign="top" width="82">Intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum intensity (a)</td><td valign="top" width="82">Maximum beam spread</td><td valign="top" width="82">Intensity (a)</td></tr></thead><tr><td valign="top" width="84">200 000</td><td valign="top" width="91">200 000</td><td valign="top" width="91">150 000</td><td colspan="2" valign="top" width="91">75 000</td><td valign="top" width="82">3°</td><td valign="top" width="82">75 000</td><td valign="top" width="82">250 000</td><td valign="top" width="82">112 500</td><td valign="top" width="82">7 500</td><td valign="top" width="82">7°</td><td valign="top" width="82">75 000</td></tr><tr><td valign="top" width="84">100 000</td><td valign="top" width="91">100 000</td><td valign="top" width="91">75 000</td><td colspan="2" valign="top" width="91">37 500</td><td valign="top" width="82">3°</td><td valign="top" width="82">37 500</td><td valign="top" width="82">125 000</td><td valign="top" width="82">56 250</td><td valign="top" width="82">3 750</td><td valign="top" width="82">7°</td><td valign="top" width="82">37 500</td></tr><tr><td valign="top" width="84">20 000</td><td valign="top" width="91">20 000</td><td valign="top" width="91">15 000</td><td colspan="2" valign="top" width="91">7 500</td><td valign="top" width="82">3°</td><td valign="top" width="82">7 500</td><td valign="top" width="82">25 000</td><td valign="top" width="82">11 250</td><td valign="top" width="82">750</td><td valign="top" width="82">N/A</td><td valign="top" width="82">N/A</td></tr><tr><td valign="top" width="84">2 000</td><td valign="top" width="91">2 000</td><td valign="top" width="91">1 500</td><td colspan="2" valign="top" width="91">750</td><td valign="top" width="82">3°</td><td valign="top" width="82">750</td><td valign="top" width="82">2 500</td><td valign="top" width="82">1 125</td><td valign="top" width="82">75</td><td valign="top" width="82">N/A</td><td valign="top" width="82">N/A</td></tr><tr><td colspan="12" valign="top" width="931">Note: This table does not include recommended horizontal beam spreads. <a href="#_DxCrossRefBm1800047824">CS ADR-DSN.Q.846(c)</a> requires 360° coverage around an obstacle. Therefore, the number of lights needed to meet this requirement will depend on the horizontal beam spreads of each light as well as the shape of the obstacle. Thus, with narrower beam spreads, more lights will be required.(a) 360° horizontal. All intensities are expressed in Candela. For flashing lights, the intensity is read into effective intensity, as determined in accordance with ICAO Doc 9157, Aerodrome Design Manual, Part 4, Visual Aids.(b) Elevation vertical angles are referenced to the horizontal when the light unit is levelled.(c) Beam spread is defined as the angle between the horizontal plan and the directions for which the intensity exceeds that mentioned in the ‘intensity’ column.Note: an extended beam spread may be necessary under specific configuration and justified by a safety assessment.</td></tr><tr height="0"></tr></table>

Table Q-3. Light distribution for medium- and high-intensity obstacle lights according to benchmark intensities of Table Q-1

Table Q-4. Marking band widths

Table Q-5. Installation setting angles for high-intensity obstacle lights

[Issue: ADR-DSN/3]

[Issue: ADR-DSN/4]

[Issue: ADR-DSN/6]

Appendix 2 to AMC 20-20B Guidelines for the development of a programme to preclude the occurrence of widespread fatigue damage

ED Decision 2021/007/R

1. INTRODUCTION

The terminology and methodology in this Appendix are based upon material developed by the AAWGand lessons learned since the first issue of this AMC.

2. DEFINITIONS

Extended service goal (ESG) is an adjustment to the design service goal established by service experience, analysis, and/or test during which the principal structure will be reasonably free from significant cracking including WFD.

Monitoring period is the period of time when special inspections of the fleet are initiated due to an increased risk of MSD/MED (ISP) and ending when the SMP is reached.

Scatter factor is a life reduction factor used in the interpretation of fatigue analysis and fatigue test results.

Test-to-structure factor is 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.

Teardown inspection is the process of disassembling structure and using destructive inspection techniques or visual (magnifying glass and dye penetrant) or other non-destructive inspection (NDI) methods (eddy current, ultrasonic) to identify the extent of damage, within a structure, caused by fatigue, environmental or accidental damage.

WFD (average behaviour) is the point in time when 50 % of the fleet is expected to reach WFD for a particular detail.

3. GENERAL

The likelihood of the occurrence of fatigue damage in an aircraft’s structure increases with aircraft usage. The design process generally establishes a design service goal (DSG) in terms of flight cycles/hours for the airframe. It is expected that any cracking that occurs on an aircraft operated up to the DSG will occur in isolation (i.e. local cracking), originating from a single source, such as a random manufacturing flaw (e.g. a mis-drilled fastener hole) or a localised design detail. It is considered unlikely that cracks from manufacturing flaws or localised design issues will interact strongly as they grow.

With extended usage, uniformly loaded structure may develop cracks in adjacent fastener holes, or in adjacent similar structural details. These cracks may or may not interact, and they can have an adverse effect on the residual strength capability of the structure before the cracks become detectable. The development of cracks at multiple locations (both MSD and MED) may also result in strong interactions that can affect subsequent crack growth; in which case, the predictions for local cracking would no longer apply. An example of this situation may occur at any skin joint where load transfer occurs. Simultaneous cracking at many fasteners along a common rivet line may reduce the residual strength of the joint to less than the required levels before the cracks are detectable under the routine maintenance programme established at the time of certification.

For new type designs, certified to CS-25 Amendment 19, AMC 25.571 provides guidance on how to establish an LOV. For existing types, for which TCHs need to comply with point 26.303 of Part‑26, CS 26.303 and this AMC apply. The TCH should conduct structural evaluations to determine where and when MSD/MED may occur. Based on these evaluations, the TCH should provide additional maintenance instructions for the structure, as appropriate. The maintenance instructions include, but are not limited to inspections, structural modifications, and limits of validity of the new maintenance instructions. In most cases, a combination of inspections and/or modifications/replacements is deemed necessary to achieve the required safety level. Other cases will require modification or replacement if inspections are not viable.

There is a distinct possibility that there could be a simultaneous occurrence of MSD and MED in a given structural area. This situation is possible on some details that were equally stressed. If this is possible, then this scenario should be considered in developing appropriate service actions for structural areas.

4. STRUCTURAL EVALUATION FOR WFD

4.1 General

The evaluation has three objectives:

(a) Identify fatigue-critical structure that may be susceptible to MSD/MED, see paragraph 4.2;

(b) Predict when it is likely to occur; see paragraph 4.3; and

(c) Establish additional maintenance actions, as necessary, to ensure continued safe operation of the aircraft; see paragraph 4.4.

4.2 Structure susceptible to MSD/MED

Susceptible structure is defined as that which has the potential to develop MSD/MED. Such structure typically has the characteristics of multiple similar details operating at similar stresses where structural capability could be affected by interaction of multiple cracking at a number of similar details. The following list provides examples of known types of structure susceptible to MSD/MED. (The list is not exhaustive):

<table border="1" cellpadding="0" cellspacing="0" width="529"><tr><td valign="top" width="425">STRUCTURAL AREA</td><td valign="top" width="104">SEE FIGURE</td></tr><tr><td valign="top" width="425">Longitudinal Skin Joints, Frames, and Tear Straps (MSD/MED)</td><td valign="top" width="104">A2-1</td></tr><tr><td valign="top" width="425">Circumferential Joints and Stringers (MSD/MED)</td><td valign="top" width="104">A2-2</td></tr><tr><td valign="top" width="425">Lap joints with Milled, Chem-milled or Bonded Radius (MSD)</td><td valign="top" width="104">A2-3</td></tr><tr><td valign="top" width="425">Fuselage Frames (MED)</td><td valign="top" width="104">A2-4</td></tr><tr><td valign="top" width="425">Stringer to Frame Attachments (MED)</td><td valign="top" width="104">A2-5</td></tr><tr><td valign="top" width="425">Shear Clip End Fasteners on Shear Tied Fuselage Frames (MSD/MED)</td><td valign="top" width="104">A2-6</td></tr><tr><td valign="top" width="425">Aft Pressure Dome Outer Ring and Dome Web Splices (MSD/MED)</td><td valign="top" width="104">A2-7</td></tr><tr><td valign="top" width="425">Skin Splice at Aft Pressure Bulkhead (MSD)</td><td valign="top" width="104">A2-8</td></tr><tr><td valign="top" width="425">Abrupt Changes in Web or Skin Thickness — Pressurised or Un-pressurised Structure (MSD/MED)</td><td valign="top" width="104">A2-9</td></tr><tr><td valign="top" width="425">Window Surround Structure (MSD, MED)</td><td valign="top" width="104">A2-10</td></tr><tr><td valign="top" width="425">Over Wing Fuselage Attachments (MED)</td><td valign="top" width="104">A2-11</td></tr><tr><td valign="top" width="425">Latches and Hinges of Non-plug Doors (MSD/MED)</td><td valign="top" width="104">A2-12</td></tr><tr><td valign="top" width="425">Skin at Runout of Large Doubler (MSD)—Fuselage, Wing or Empennage</td><td valign="top" width="104">A2-13</td></tr><tr><td valign="top" width="425">Wing or Empennage Chordwise Splices (MSD/MED)</td><td valign="top" width="104">A2-14</td></tr><tr><td valign="top" width="425">Rib to Skin Attachments (MSD/MED)</td><td valign="top" width="104">A2-15</td></tr><tr><td valign="top" width="425">Typical Wing and Empennage Construction (MSD/MED)</td><td valign="top" width="104">A2-16</td></tr></table>

![EASA EAR image](/static/files/file_ftLzkYEblJP/image005.png)

Figure A2-1: Longitudinal skin joints, frames, and tear straps (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image006.png)

Figure A2-2: Circumferential joints and stringers (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image007.png)

Figure A2-3: Lap joints with milled, chem-milled or bonded radius (MSD)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image008.png)

Figure A2-4: Fuselage frames (MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image009.png)

Figure A2-5: Stringer-to-frame attachments (MED)

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>Skin/Stringer Attachments</div></td></tr></table>

![EASA EAR image](/static/files/file_ftLzkYEblJP/image010.png)

Figure A2-6: Shear clip end fasteners on shear tied fuselage frame (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image011.png)

Figure A2-7: Aft pressure dome outer ring and dome web splices (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image012.png)

Figure A2-8: Skin splice at aft pressure bulkhead (MSD)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image013.png)

Figure A2-9: Abrupt changes in web or skin thickness — Pressurised or unpressurised structure (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image014.png)

Figure A2-10: Window surround structure (MSD, MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image015.png)

Figure A2-11: Over wing fuselage attachments (MED)

![brent_1](/static/files/file_ftLzkYEblJP/image016.jpg)

Figure A2-12: Latches and hinges of non-plug doors (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image017.png)

Figure A2-13: Skin at runout of large doubler (MSD) — Fuselage, wing or empennage

![EASA EAR image](/static/files/file_ftLzkYEblJP/image018.png)

Figure A2-14: Wing or empennage chordwise splices (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image019.png)

Figure A2-15: Rib-to-skin attachments (MSD/MED)

![EASA EAR image](/static/files/file_ftLzkYEblJP/image020.png)

Figure A2-16: Typical wing and empennage construction (MSD/MED)

4.3 WFD Evaluation

Point 26.300 of Part-26 requires an LOV to be established according to specified timescales for large transport aeroplanes with MTOWs above 34 901 kg (75 000 lb). For other types, it is recommended that by the time the highest-time aircraft of a particular model reaches its DSG, the evaluation for each area susceptible to the development of WFD should be completed. A typical evaluation process is shown in Figure A2-19 below. This evaluation will establish the necessary elements to determine a maintenance programme to preclude WFD in that particular model’s aircraft fleet. These elements are developed for each susceptible area and include:

**4.3.1 Identification of structure potentially susceptible to WFD**

Unless already fully addressed in the existing fatigue and damage tolerance evaluation, the TCH should identify each part of the aircraft’s structure that is potentially susceptible to WFD for further evaluation. A justification should be given that supports selection or rejection of each area of the aircraft structure. DAHs for modified or repaired structure should evaluate their structure and its effect on existing structure.

Typical examples of structure susceptible to WFD are included in paragraph 4.2 of this Appendix.

**4.3.2 Predicting when WFD will occur**

(a) Characterisation of events leading to WFD

The fatigue process that leads to WFD is shown in Figure A2-17. This figure is applicable to both damage that occurs in multiple sites (MSD) and damage that occurs in similar structures at more than one location (MED). For any susceptible structural area, it is not a question of whether WFD will occur, but when it will occur. In Figure A2-17, the ‘when’ is illustrated by the line titled ‘WFD (average behaviour),’ which is the point when 50 % of the aeroplanes in a fleet would have experienced WFD in the considered area (note that the probability density function for flight cycles or flight hours to WFD has been depicted for reference). The WFD process includes this phase of crack initiation and a crack growth phase. During the crack initiation phase, which generally spans a long period of time, there is little or no change in the basic strength capability of the structure. The actual residual strength curve depicted in Figure A2-17 is flat, and equal to the strength of the structure in its pristine state. However, at some time after the first small cracks start to grow, residual strength begins to degrade. Crack growth continues until the capability of the structure degrades to the point of the minimum strength required by CS 25.571(b). In this context, the line in Figure A2-17 called WFD (average behaviour) represents a point when 50 % of the aeroplanes in a fleet fall below the minimum strength specifications of CS 25.571(b).

![EASA EAR image](/static/files/file_ftLzkYEblJP/image021.png)

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>CRACK LENGTH</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>REQUIRED RESIDUAL STRENGTH</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>(Average Behavior</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>ACTUAL RESIDUAL STRENGTH</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>FLIGHT CYCLES</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>NOTE: No scale implied.</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>For definition use only.</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>Probability Density </div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>Function for Cycles to </div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>MSD/MEDCRACKINITIATION</div></td></tr></table>

![EASA EAR image](/static/files/file_ftLzkYEblJP/image021.png)

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>WFD(average behaviour)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>Probability densityfunction for cycles to WFDcondition</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>MSD/MEDCRACK GROWTH</div></td></tr></table>

Figure A2-17: Effect on residual strength of developing WFD

(b) Determination of WFD (average behaviour) in the fleet

The time in terms of flight cycles/hours defining the WFD (average behaviour) in the fleet should be established for each susceptible structural area. The data to be assessed in determining the WFD (average behaviour) includes:

— a complete review of the service history of the susceptible areas, to identify any occurrences of fatigue cracking and the continuing validity of loads and mission profiles,

— evaluation of the operational statistics of the fleet in terms of flight hours and landings,

— significant production variants (material, design, assembly method, and any other change that might affect the fatigue performance of the detail),

— fatigue test evidence including relevant full-scale and component fatigue and damage tolerance test data (see subparagraph 4.3.9 and [Annex 1](#_DxCrossRefBm1701658019) for more details),

— teardown inspections, and

— any fractographic analysis available.

The evaluation of the test results for the reliable prediction of the time to when WFD might occur in each susceptible area should include appropriate test-to-structure factors. If full-scale fatigue test evidence is used, Figure A2‑20 below relates how that data might be utilised in determining WFD (average behaviour). Evaluation may be analytically determined, supported by test and, where available, service evidence.

Regardless of whether the assessment of WFD (average behaviour) is based on in‑service data, full-scale fatigue test evidence, analyses, or a combination of any of these, the following should be considered:

**4.3.3 Initial crack/damage scenario**

This is an estimate of the size and extent of multiple cracking expected at MSD/MED initiation. This prediction requires empirical data or an assumption of the crack/damage locations and sequence plus a fatigue evaluation to determine the time to MSD/MED initiation. Alternatively, analysis can be based on either:

— the distribution of equivalent initial flaws, as determined from the analytical assessment of flaws found during fatigue test and/or teardown inspections regressed to zero cycles; or

— a distribution of fatigue damage determined from relevant fatigue testing and/or service experience.

**4.3.4 Final cracking scenario**

This is an estimate of the size and extent of multiple cracking that could cause residual strength to fall to the minimum required level as shown in A2-17. Techniques exist for 3-D elastic-plastic analysis of such problems; however, there are several alternative test and analysis approaches available that provide an equivalent level of safety. One such approach is to define the final cracking scenario as a subcritical condition (e.g. first crack at link-up). The use of a subcritical scenario reduces the complexity of the analysis and, in many cases, will not greatly reduce the total crack growth time, because the majority of the time taken to reach the critical condition is generally in the initiation phase.

**4.3.5 Crack growth calculation**

Progression of the crack distributions from the initial cracking scenario to the final cracking scenario should be developed. These curves can be developed:

— *analytically*, typically based on linear elastic fracture mechanics, or

— *empirically*, from test or service fractographic data.

**4.3.6 Potential for discrete source damage (DSD)**

A structure susceptible to fatigue including MSD/MED may also be affected by DSD due to an uncontained failure of high-energy rotating machinery (i.e. turbine engines). At this time, there is no specific requirement to address prior fatigue cracking in combination with DSD for certification. Nonetheless, when assessing in‑service findings of fatigue cracking, the additional threat posed by any potential DSD should be taken into account when developing the corrective actions and the timescales for its implementation.

**4.3.7 Analysis methodology**

**Differences between multiple-site damage and multiple-element damage**

Details of the approach used to characterise events leading up to WFD may be different. The differences will largely depend on whether MSD or MED is being considered. This is especially true for crack interaction.

(a) Crack interaction

MSD has the potential for strong crack interaction, and the effect of multiple cracks on each other needs to be addressed. MED, in most cases, does not have the same potential for strong crack interaction. The differences between the interaction effects for MSD and MED are illustrated in Figure A2-18.

(b) MSD and MED interaction

Some areas of an aeroplane are potentially susceptible to both MSD and MED. Simultaneous occurrence of MSD and MED is possible, even though it is not common. A comparison of inspection start points (ISPs) or modification start points might indicate the possibility of this occurrence. If so, the evaluation should consider the interaction between MSD and MED.

![EASA EAR image](/static/files/file_ftLzkYEblJP/image022.png)

Figure A2-18: Difference between MSD and MED interaction effects

The report ‘Recommendations for Regulatory Action to Prevent Widespread Fatigue Damage in the Commercial Aeroplane Fleet’, Revision A, dated June 29, 1999 (a report of the AAWG for the ARAC’s Transport Aircraft and Engine Issues Group), discusses two Round Robin exercises developed by the TCHs to provide insight into their respective methodologies. One outcome of the exercises was an identification of key assumptions or methods that had the greatest impact on the predicted WFD behaviour. These assumptions were:

— the flaw sizes assumed at initiation of crack growth phase of analysis;

— material properties used (static, fatigue, fracture mechanics);

— ligament failure criteria;

— crack growth equations used;

— statistics used to evaluate the fatigue behaviour of the structure (e.g. time to crack initiation);

— methods of determining the structural modification point (SMP);

— detectable flaw size assumed;

— initial distribution of flaws; and

— factors used to determine bound behaviour as opposed to mean behaviour.

When considering MED, where interaction between cracks in different elements is not a factor, the following should be considered:

(1) In a structure containing large numbers of similar elements, there is not normally a high probability that, after a crack initiates in an element, a second crack will initiate in the element right next to it. If this does happen, however, the consequences to the overall structure may be severe. This is because having two structural members fail right next to each other can completely negate any ability of the structure to tolerate additional damage. Consequently, when performing an evaluation, applicants should make conservative assumptions and assume failures to be adjacent to each other.

(2) When an element fails completely, the load that has to be redistributed onto the non-failed structure can be large and can have a significant impact on the strength of both the cracked and uncracked structure; therefore, the effects of load redistribution must be included in the evaluation.

(d) Establishing maintenance actions

The following parameters are developed from paragraphs 4.3.2 to 4.3.7 above, and are necessary to establish an MSD/MED maintenance programme for the area under investigation.

Fatigue damage is the gradual deterioration of a material subjected to repeated loads. This gradual deterioration is a function of use and can be statistically quantified. The term ‘WFD’ is used, and can be statistically quantified, at the end of the deterioration process when the structure is no longer able to carry the residual strength loads. WFD can never be absolutely precluded because there is always some probability, no matter how small, that it will occur. Therefore, modifying or replacing structure at a predetermined, analytically-derived time stated in flight cycles or flight hours, minimises the probability of having WFD in the fleet. Modification or replacement is the most reliable method for precluding WFD. The point at which a modification is undertaken is referred to as the ‘structural modification point (SMP)’ and it is illustrated in Figure 2-1 of [Annex 2](#_DxCrossRefBm1701658018). The SMP is generally a fraction of the number representing the point in time when WFD (average behaviour) will occur, and should result in the same reliability as a successful two-lifetime fatigue test. This level of reliability for setting the SMP is acceptable if MSD or MED inspections are shown to be effective in detecting cracks. If the inspections are effective, they should be implemented before the SMP. The implementation times for these inspections are known as the ‘inspection start points (ISPs)’. Repeat inspections are usually necessary to maintain this effectiveness in detecting cracks. If MSD or MED inspections are not effective in detecting cracks, then the SMP should be set at the time of ISP. For the purposes of this AMC, an inspection is effective if, when performed by properly trained maintenance personnel, it will readily detect the damage in question[[9]](#_ftn9). The SMP should minimise the extent of cracking in the susceptible structural area in a fleet of affected aeroplanes. In fact, if this point is appropriately determined, a high percentage of aeroplanes would not have any MSD or MED by the time the SMP is reached.

Due to the redundant nature of semi-monocoque structures, MED can be difficult to manage in a fleet environment. This stems from the fact that most aircraft structures are built-up in nature, and that makes the visual inspection of the various layers difficult. Also, visual inspections for MED typically rely on internal inspections, which may not be practical at the frequency necessary to preclude MED due to the time required to gain access to the structure. However, these issues are dependent on the specific design involved and the amount of damage being considered. In order to implement a viable inspection programme for MED, static stability must be maintained at all times and there should be no MED concurrent with MSD in a given structural area.

**4.3.8 Inspection start point (ISP)**

This is the point at which inspection starts if a monitoring period is used. Inspection is not practical for all applications and cannot replace the SMP. The ISP is determined through a statistical analysis of crack initiation based on fatigue testing, teardown, or service experience of similar structural details. It is assumed that the ISP is equivalent to a lower bound value with a specific probability in the statistical distribution of cracking events. Alternatively, the ISP may be established by applying appropriate factors to the average behaviour.

When inspections are determined to be effective, it is necessary to establish when those inspections should start. This point is illustrated in Figure 2-1. The start point is determined through a statistical analysis of crack initiation based on fatigue testing, teardown, or in-service experience of similar structure. The ISP is assumed to be equivalent to a lower-bound value with a specific probability in the statistical distribution of cracking events. Alternatively, an ISP may be established by applying appropriate factors to the number representing WFD (average behaviour). (e.g. for aluminium alloy structure, dividing the full-scale test result by a factor of 3).

For inspection intervals, see point 4.3.10.

**4.3.9 Structural modification point (SMP)**

The SMP should be established as a point in time when structures should be modified or replaced to prevent WFD from occurring. This is typically established by:

— calculating when WFD would first occur in the structure (predicted using the WFD (average behaviour)),

— setting a time before the predicted occurrence of WFD to perform modifications or replacements that will prevent it.

The applicant should demonstrate that the proposed SMP established during the evaluation has the same confidence level as current regulations require for new certification. In lieu of other acceptable methods, the SMP for aluminium alloy structures can be established as a point reduced from the WFD (average behaviour), based on the viability of inspections in the monitoring period. The SMP may be determined by dividing the number representing the timing when WFD (average behaviour)will occur by a factor of 2 if there are effectiveinspections, or by a factor of 3 if inspections are not effective. For other materials such as high‑strength steel alloys, larger scatter factors may be necessary to account for increased variability in fatigue performance.

An aircraft should not be operated past the SMP unless the structure is modified or replaced, or unless additional approved data is provided that would extend the SMP. However, if during the structural evaluation for WFD, a TCH/DAH finds that the flight cycles and/or flight hours SMP for a particular structural detail have been exceeded by one or more aircraft in the fleet, the TCH/DAH should expeditiously evaluate selected high-time aircraft in the fleet to determine their structural condition. From this evaluation, the TCH/DAH should notify the competent authorities and propose appropriate service actions.

A DAH may find that the SMP for a particular structural area has been exceeded by one or more aeroplanes in the fleet. In that case, the DAH should expedite the evaluation of those high-time aeroplanes to determine their structural condition, notify EASA and propose appropriate maintenance actions specific to those aeroplanes.

The initial SMP may be adjusted based on the following:

(a) The tasks necessary to extend an SMP may include any or all of the following:

(1) Additional fatigue or residual strength tests, or both, on a full-scale aeroplanestructure or a full-scale component followed by detailed inspections and analyses.

(2) Fatigue tests of new structure orstructure from in-service aeroplaneson a smaller scale than full component tests (i.e. subcomponent or panel tests, or both). If a subcomponent test is used, the SMP would be extended only for that subcomponent.

(3) Teardown inspections (destructive) on structural components that have been removed from service.

(4) Teardowninspections (non-destructive) accomplished by selected, limited disassembly and subsequent reassembly of specific areas of high-time aeroplanes.

(5) Analysis of in-service data (e.g. inspections) from a statistically significant number of aeroplanes.

(b) If cracks are found in the structural detail for which the evaluation was done during either the monitoring period or the modification programme, the SMP should be re-evaluated to ensure that the SMP does in fact provide the required confidence level. If it is shown that the required confidence level is not being met, the SMP should be adjusted and the adjustment reflected in the appropriate SBs to address the condition of the fleet. Additional regulatory action may be required.

**4.3.10 Inspection interval and method**

An interval should be chosen to provide a sufficient number of inspections between the ISP and the SMP so that there is high confidence that no MSD/MED condition will reach the final cracking scenario without detection. The interval between inspections depends on the detectable crack size, the critical crack lengths and the probability that the cracks will be detected with the specific inspection method. Conservative scenarios should be assumed for developing the inspection interval unless other assumptions can be consistently supported by test and service experience. If the crack cannot be detected, the SMP must be re-evaluated to ensure there is a high confidence level that no aircraft will develop MSD/MED before modification.

**4.4 Evaluation of maintenance actions**

For all areas that have been identified as susceptible to MSD/MED, the current maintenance programme should be evaluated to determine whether adequate structural maintenance and inspection programmes exist to safeguard the structure against unanticipated cracking or other structural degradation. The evaluation of the current maintenance programme typically begins with the determination of the SMP for each area.

Each area should then be reviewed to determine the current maintenance actions and compare them to the maintenance needs established in this evaluation. Issues to be considered include the following:

(a) Determine the inspection requirements (method, inspection start point, and repeat interval) of the inspection for each susceptible area (including that structure that is expected to arrest cracks) that is necessary to maintain the required level of safety.

(b) Review the elements of the existing maintenance programmes already in place.

For susceptible areas approaching the SMP, where the SMP will not be increased or for areas that cannot be reliably inspected, a programme should be developed and documented that provides for replacement or modification of the susceptible structural area.

**4.4.1 Period of WFD evaluation validity**

At whatever point the WFD evaluation is made, it should support the LOV of the maintenance programme. Consistent with the use of test evidence to support individual SMPs, as described above in paragraph 4.3.9, the LOV of the maintenance programme should be based on fatigue test evidence. For an existing ageing aircraft type, theinitial WFD evaluation of the complete airframe will typically cover a significant forward estimation of the projected aircraft usage beyond its DSG, also known as the ‘proposed ESG’and is effectively a proposed LOV.Typically, an evaluation through an additional 25 % of the DSG would provide a realistic forecast, with reasonable planning time for necessary maintenance action. However, it may be appropriate to adjust the evaluation validity period depending on issues such as:

(a) the projected useful life of the aircraft at the time of the initial evaluation;

(b) current NDI technology; and

(c) airline advance planning requirements for introduction of new maintenance and modification programmes, to provide sufficient forward projection to identify all likely maintenance/modification actions essentially as one package.

Upon completion of the evaluation and publication of the revised maintenance requirements, the ‘proposed ESG’ becomes the LOV.

*Note*: This assumes that all other aspects of the maintenance programme that are required to support the LOV (such as SSID, CPCP, etc.) are in place and have been evaluated to ensure they too remain valid up to the LOV.

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>FOR EACH AREA, DETERMINE THE WFD AVERAGE BEHAVIOUR IN THE FLEET(See 4.3.2 onwards)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>ESTABLISH ISP,INSPECTIONINTERVALAND METHOD AND SCHEDULE FOR TERMINATING ACTION (See 4.3.9//10/11)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>ESTIMATE ALLOWABLE FATIGUE DAMAGE SCENARIO FOR LIMIT LOAD (See 4.3.4)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>ESTABLISH THE SMP AND TERMINATING ACTION(See 4.3.10)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>FATIGUE DAMAGE SCENARIO DETECTABLE PRIOR TO MAXIMUM ALLOWABLE EXTENT UNDER LIMIT LOAD</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>IS NATURAL FATIGUE CRACKING LIKELY 1WITHIN OPERATIONAL LIFE ²</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>REVIEW STRUCTURAL AREAS POTENTIALLY SUSCEPTIBLE TO WFD(See 4.3.1)</div></td></tr></table>

NOTES:

1. Fatigue cracking is defined as likely if the factored fatigue life is less than the projected ESG of the aircraft at time of WFD evaluation.

2. The operational life is the projected ESG of the aircraft at time of WFD Evaluation. (See 4.4.1).

Figure A2-19: Aircraft WDF evaluation process

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>FULL SCALE FATIGUE TEST DATA</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>MSD/MED FINDINGS DURING TEST/TEARDOWN?</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>TEST LIFE MinusCRACK GROWTH LIFE3</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>TEST LIFE plusCRACK GROWTH LIFE3</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH AT LIMIT LOAD?</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>INSPECTION PROGRAMME/MODIFICATION PROGRAMME REQUIRED (See 4.3.7 onward)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>NO SPECIAL INSPECTIONS REQUIRED (FAR 25.571, AMDT 96)LOV = Test Life/2</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>FULL-SCALE FATIGUE TEST DATA</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>MSD/MED FINDINGS DURING TEST/TEARDOWN?</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>TEST LIFE minusCRACK GROWTH LIFE3</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>TEST LIFE plusCRACK GROWTH LIFE3</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH2 AT LIMIT LOAD?</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>INSPECTION PROGRAMME/MODIFICATION PROGRAMMEREQUIRED (See 4.3.7 onward)</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>LOV = Test Life/2 if all areas inspectable</div></td></tr></table>

<table cellpadding="0" cellspacing="0" width="100%"><tr><td><div>1 ASSUMED STATE AT END OF TEST: Best estimate of non-detected damage from inspection method used at the end of the test or during teardown.2 CRITICAL CRACK LENGTH: First link-up of adjacent cracks at limit load (locally) or an adequate level of large damage capability. 3 CRACK GROWTH LIFE: Difference between assumed or actual state at the end of the test and critical crack length.Figure A2-20: Use of fatigue test and teardown information to determine WFD average behaviour</div></td></tr></table>

5. DOCUMENTATION

Any person seeking approval of an LOV of an aircraft type design should develop a document containing all the necessary ISPs, inspection procedures, replacement times, SMPs, and any other maintenance actions necessary to preclude WFD, and to support the LOV. That person must revise the SSID or ALS as necessary, and/or prepare SBs that contain the aforementioned maintenance actions. Since WFD is a safety concern for all operators of older aircraft, EASA will make mandatory the identified inspection and modification programmes. In addition, EASA may consider separate AD action to address any SBs or other service information publications revised or issued as a result of in-service MSD/MED findings resulting from implementation of these programmes.

The following items should be contained in the front of the documentation supporting the LOV:

(a) identification of the variants of the basic aircraft type to which the document relates;

(b) summary of the operational statistics of the fleet in terms of hours and flights;

(d) the types of operations for which the inspection programme is considered valid;

(e) reference to documents giving any existing inspections, or modification of parts or components; and

(f) the LOV of the maintenance programme in terms of flight cycles or flight hours or both as appropriate to accommodate variations in usage.

The document should contain at least the following information for each critical part or component:

(a) description of the primary structure susceptible to WFD;

(b) details of the monitoring period (ISP, repeat inspection interval, SMP, inspection method and procedure (including crack size, location and direction) and alternatives) when applicable;

(d) any mandatory modification or replacement of the structural element;

(e) SBs (or other service information publications) revised or issued as a result of in-service findings resulting from the WFD evaluations (added as a revision to the initial WFD document); and

(f) guidance to the operator on which inspection findings should be reported to the TCH/DAH, and appropriate reporting forms and methods of submission.

6. REPORTING REQUIREMENTS

Operators, TCHs and STCHs are required to report in accordance with various regulations (e.g. point 21.A.3A, and point 145.A.60). The regulations to which this AMC relates do not require any reporting requirements in addition to the current ones. Due to the potential threat to structural integrity, the results of inspections must be accurately documented and reported in a timely manner to preclude the occurrence of WFD. The current system of operator and TCH communication has been useful in identifying and resolving a number of issues that can be classified as WFD concerns. MSD/MED has been discovered via fatigue testing and in-service experience. TCHs have been consistent in disseminating related data to operators to solicit additional service experience. However, a more thorough means of surveillance and reporting is essential to preclude WFD.

When damage is found while conducting an approved MSD/MED inspection programme, or at the SMP where replacement or modification of the structure is occurring, the TCHs, STCHs and the operators need to ensure that greater emphasis is placed on accurately reporting the following items:

(a) a description (with a sketch) of the damage, including crack length, orientation, location, flight cycles/hours, and condition of structure;

(b) results of follow-up inspections by operators that identify similar problems on other aircraft in the fleet;

(c) findings where inspections accomplished during the repair or replacement/modification identify additional similar damage sites; and

(d) adjacent repairs.

Operators must report all cases of MSD/MED to the TCH, STCH or the competent authority as appropriate, irrespective of how frequently such cases occur. Cracked areas from in-service aircraft (damaged structure) may be needed for detailed examination. Operators are encouraged to provide fractographic specimens whenever possible. Aeroplanes undergoing heavy maintenance checks are perhaps the most useful sources for such specimens.

Operators should remain diligent in the reporting of potential MSD/MED concerns not identified by the TCH/DAH. Indications of a developing MSD/MED problem may include:

(a) damage at multiple locations in similar adjacent details;

(b) repetitive part replacement; or

Documentation will be provided by the TCH and STCH as appropriate to specify the required reporting format and time frame, supporting the mandatory reporting regulations (e.g. point 21.A.3A of Part 21, point 145.A.60 of Part-145). The data will be reviewed by the TCH or STCH, operator(s), and EASA to evaluate the nature and magnitude of the problem and to determine the appropriate corrective action.

7. WFD EVALUATION FOR STRUCTURAL MODIFICATIONS AND REPAIRS

TCHs of aeroplanes subject to the point 26.303 of Part-26 requirements for an LOV should perform WFD evaluations to assess all the applicable existing structure and the effect of future changes on the LOV.

The WFD evaluations of this AMC do not apply retroactively to existing STCH’s modifications, nor to existing repairs. Future changes and repairs need to take into account the applicable certification basis, and applicants should consider the guidance of the applicable ACJ and AMC as discussed in paragraph 8 of this AMC. The DTEs for compliance with points 26.307, 26.308, 26.333 and 26.334 of Part-26 do not have to consider WFD (average behaviour), or the related SMP and ISP.

In cases where a new DTE is performed by DAHs to comply with points 26.333 and 26.334 of Part-26 for existing changes or for new changes or repairs, according to CS 25.571 Amdt 18 or earlier amendments, the DTE and development of DTIs should take into account the cracking scenarios that could reasonably be expected to occur in the remaining operational lifetime of an aeroplane into which the repair or modification is, or may be, incorporated.

8. RESPONSIBILITYFOR WFD EVALUATION, ESTABLISHING THE LOV AND IMPLEMENTATION OFTHE LOV AND MAINTENANCE ACTIONS

The primary responsibility is with the DAH to perform the analyses and supporting tests. However, it is expected that if extensive maintenance actions are necessary, the practicality of their implementation will be evaluated in a cooperative effort between the operators and TCHs/DAHs, with participation of EASA.

The TCH is responsible for proposing and submitting an LOV in the ALS for approval.

*Note*:In some cases, the ALS may already contain an LOV which is approved in accordance with a regulation of another authority. There may also be other potentially more restrictive limitations on the validity of maintenance programmes. For these cases, when the TCH needs to publish the LOV as required by point 26.303 of Part-26, this LOV and its relationship with the existing or superseded limitation should be clearly described so that no operator will exceed the most restrictive applicable limit on the general validity of the maintenance programme.

The operator is responsible for implementing the LOV in their maintenance programme.

*Note*:The LOV does not supersede or allow operations beyond any lower limitation applicable to the individual aeroplane and the components controlled by the maintenance programme.

[Amdt 20/2]

[Amdt 20/20]

[Amdt 20/22]

---

[[9]](#_ftnref9) The cracking identified in the FAA Airworthiness Directive (AD) 2002-07-09 is an example of the type of cracking that MSD inspections are effective in detecting. These cracks grow from the fastener holes in the lower row of the lower skin panel in such a way that the cracking is readily detectable using NDI methods. The cracking identified in the FAA AD 2002-07-08 is an example of places where MSD inspections are not effective. These cracks grow in the outer surface and between the fastener holes in the lower row of the lower skin panel in such a way that the cracking is not readily detectable using NDI methods. Modification is the only option to address this type of cracking.

Appendix 2 to AMC 20-20A Guidelines for the development of a
programme to preclude the occurrence of widespread fatigue damage

ED Decision 2020/023/R

1.         INTRODUCTION

The terminology and methodology in this Appendix
are based upon material developed by the AAWGand lessons learned since the
first issue of this AMC.

2.         DEFINITIONS

Extended service goal (ESG) is an adjustment to the design service goal established by service
experience, analysis, and/or test during which the principal structure will be
reasonably free from significant cracking including WFD.

Monitoring period is the period of time when special inspections of the fleet are
initiated due to an increased risk of MSD/MED (ISP) and ending when the SMP is
reached.

Scatter factor is a life reduction factor used in the interpretation of fatigue
analysis and fatigue test results.

Test-to-structure factor is 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.

Teardown inspection is the process of disassembling structure and using destructive
inspection techniques or visual (magnifying glass and dye penetrant) or other
non-destructive inspection (NDI) methods (eddy current, ultrasonic) to
identify the extent of damage, within a structure, caused by fatigue,
environmental or accidental damage.

WFD (average
behaviour) is the point in time when 50 %
of the fleet is expected to reach WFD for a particular detail.

3.         GENERAL

The likelihood of the occurrence of fatigue
damage in an aircraft’s structure increases with aircraft usage. The design
process generally establishes a design service goal (DSG) in terms of flight cycles/hours
for the airframe. It is expected that any cracking that occurs on an aircraft
operated up to the DSG will occur in isolation (i.e. local cracking),
originating from a single source, such as a random manufacturing flaw (e.g. a
mis-drilled fastener hole) or a localised design detail. It is considered unlikely
that cracks from manufacturing flaws or localised design issues will interact
strongly as they grow.

With extended usage, uniformly loaded
structure may develop cracks in adjacent fastener holes, or in adjacent
similar structural details. These cracks may or may not interact, and they can
have an adverse effect on the residual strength capability of the structure
before the cracks become detectable. The development of cracks at multiple
locations (both MSD and MED) may also result in strong interactions that can
affect subsequent crack growth; in which case, the predictions for local
cracking would no longer apply. An example of this situation may occur at any
skin joint where load transfer occurs. Simultaneous cracking at many fasteners
along a common rivet line may reduce the residual strength of the joint to
less than the required levels before the cracks are detectable under the
routine maintenance programme established at the time of certification.

For new type designs, certified to CS-25
Amendment 19, AMC 25.571 provides guidance on how to establish an LOV.
For existing types, for which TCHs need to comply with point 26.303 of Part‑26,
CS 26.303 and this AMC apply. The TCH should conduct structural
evaluations to determine where and when MSD/MED may occur. Based on these
evaluations, the TCH should provide additional maintenance instructions for
the structure, as appropriate. The maintenance instructions include, but are
not limited to inspections, structural modifications, and limits of validity
of the new maintenance instructions. In most cases, a combination of
inspections and/or modifications/replacements is deemed necessary to achieve
the required safety level. Other cases will require modification or
replacement if inspections are not viable.

There is a distinct possibility that there
could be a simultaneous occurrence of MSD and MED in a given structural area.
This situation is possible on some details that were equally stressed. If this
is possible, then this scenario should be considered in developing appropriate
service actions for structural areas.

4.         STRUCTURAL EVALUATION FOR WFD

4.1       General

The evaluation has three objectives:

(a)       Identify
fatigue-critical structure that may be susceptible to MSD/MED, see paragraph
4.2;

(b)      Predict
when it is likely to occur; see paragraph 4.3; and

(c)       Establish
additional maintenance actions, as necessary, to ensure continued safe
operation of the aircraft; see paragraph 4.4.

4.2       Structure susceptible to MSD/MED

Susceptible structure is defined as that
which has the potential to develop MSD/MED. Such structure typically has the
characteristics of multiple similar details operating at similar stresses
where structural capability could be affected by interaction of multiple
cracking at a number of similar details. The following list provides examples
of known types of structure susceptible to MSD/MED. (The list is not
exhaustive):

<table border="1" cellpadding="0" cellspacing="0" width="529">
<tr>
<td valign="top" width="425">
STRUCTURAL
 AREA
</td>
<td valign="top" width="104">
SEE
 FIGURE
</td>
</tr>
<tr>
<td valign="top" width="425">
Longitudinal
 Skin Joints, Frames, and Tear Straps (MSD/MED)
</td>
<td valign="top" width="104">
A2-1
</td>
</tr>
<tr>
<td valign="top" width="425">
Circumferential
 Joints and Stringers (MSD/MED)
</td>
<td valign="top" width="104">
A2-2
</td>
</tr>
<tr>
<td valign="top" width="425">
Lap
 joints with Milled, Chem-milled or Bonded Radius (MSD)
</td>
<td valign="top" width="104">
A2-3
</td>
</tr>
<tr>
<td valign="top" width="425">
Fuselage
 Frames (MED)
</td>
<td valign="top" width="104">
A2-4
</td>
</tr>
<tr>
<td valign="top" width="425">
Stringer
 to Frame Attachments (MED)
</td>
<td valign="top" width="104">
A2-5
</td>
</tr>
<tr>
<td valign="top" width="425">
Shear
 Clip End Fasteners on Shear Tied Fuselage Frames (MSD/MED)
</td>
<td valign="top" width="104">
A2-6
</td>
</tr>
<tr>
<td valign="top" width="425">
Aft
 Pressure Dome Outer Ring and Dome Web Splices (MSD/MED)
</td>
<td valign="top" width="104">
A2-7
</td>
</tr>
<tr>
<td valign="top" width="425">
Skin
 Splice at Aft Pressure Bulkhead (MSD)
</td>
<td valign="top" width="104">
A2-8
</td>
</tr>
<tr>
<td valign="top" width="425">
Abrupt
 Changes in Web or Skin Thickness — Pressurised or Un-pressurised Structure
 (MSD/MED)
</td>
<td valign="top" width="104">
A2-9
</td>
</tr>
<tr>
<td valign="top" width="425">
Window
 Surround Structure (MSD, MED)
</td>
<td valign="top" width="104">
A2-10
</td>
</tr>
<tr>
<td valign="top" width="425">
Over
 Wing Fuselage Attachments (MED)
</td>
<td valign="top" width="104">
A2-11
</td>
</tr>
<tr>
<td valign="top" width="425">
Latches
 and Hinges of Non-plug Doors (MSD/MED)
</td>
<td valign="top" width="104">
A2-12
</td>
</tr>
<tr>
<td valign="top" width="425">
Skin
 at Runout of Large Doubler (MSD)—Fuselage, Wing or Empennage
</td>
<td valign="top" width="104">
A2-13
</td>
</tr>
<tr>
<td valign="top" width="425">
Wing
 or Empennage Chordwise Splices (MSD/MED)
</td>
<td valign="top" width="104">
A2-14
</td>
</tr>
<tr>
<td valign="top" width="425">
Rib
 to Skin Attachments (MSD/MED)
</td>
<td valign="top" width="104">
A2-15
</td>
</tr>
<tr>
<td valign="top" width="425">
Typical
 Wing and Empennage Construction (MSD/MED)
</td>
<td valign="top" width="104">
A2-16
</td>
</tr>
</table>

![Aviation.Bot AI suggestion, not from EASA source: The image provides a visual and textual overview of typical aircraft structural components susceptible to Widespread Fatigue Damage (WFD), specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), detailing their locations and influencing factors.

Detailed Description:
The image is structured into four main sections: a 3D isometric diagram of an aircraft fuselage section, three 2D cross-sectional diagrams of different joint types, and two text blocks providing details on WFD types and influencing factors.

1. **Top-left Diagram (Fuselage Section):** This section displays an isometric, wireframe-style illustration of a segment of an aircraft fuselage structure, rendered in blue lines against a white background. It depicts the outer skin and internal reinforcing components. Black arrows with labels point to specific structural elements:
 * An arrow points to "Tear straps," which are horizontal reinforcing strips positioned on the inner surface of the skin.
 * An arrow points to "Longitudinal skin joint," indicating a vertical seam where two skin panels are joined.
 * An arrow points to "Stringer," identifying a longitudinal stiffening element running parallel to the longitudinal skin joint.
 * An arrow points to "Frame," which is a curved, transverse structural element, positioned perpendicular to the stringers and longitudinal skin joint.

2. **Top-right Diagrams (Joint Types):** This section presents three distinct 2D cross-sectional diagrams, labeled (a), (b), and (c), illustrating different methods of joining aircraft skin panels. Fasteners are represented by horizontal dashed lines.
 * **(a) Lap joint:** This diagram shows two flat skin panels overlapping each other. An upward-pointing black arrow labeled "Outer skin upper rivet row" indicates the top row of fasteners in the upper, overlapping skin. A downward-pointing black arrow labeled "Inner skin lower rivet row" indicates the bottom row of fasteners in the lower, overlapped skin.
 * **(b) Butt joint:** This diagram shows two flat skin panels aligned edge-to-edge, joined by a central reinforcing doubler plate positioned on one side. Multiple rows of fasteners are depicted, connecting both skin panels to the doubler.
 * **(c) Lap joint with radius:** This diagram is similar to the lap joint in (a), but the lower skin panel features a distinct curved or radiused section at the point of overlap. Fasteners are shown passing through this curved section.

3. **Bottom-left Text Block ("Type and possible location of MSD and MED"):** This section provides a bulleted list detailing the specific types and locations where Multiple Site Damage (MSD) and Multiple Element Damage (MED) can occur within the depicted structures.
 * "MSD longitudinal skin joint" is the primary category, with sub-categories:
 * "Lap joint":
 * "- Outer skin upper rivet row"
 * "- Inner skin lower rivet row"
 * "Butt joint":
 * "- Skin outer rivet rows"
 * "- Doubler inner rivet rows"
 * "Lap joint with radius":
 * "- In radius"
 * "MED—frame" is another primary category, with a sub-category:
 * "- Stress concentration areas"
 * "MED—tear straps" is the final primary category, with a sub-category:
 * "- Critical fastener rows in the skin at tear strap joint"

4. **Bottom-right Text Block ("Service or test experience of factors that influence MSD and MED (examples)"):** This section lists various factors, derived from service or test experience, that can contribute to or influence the occurrence of MSD and MED. The factors are presented as a bulleted list:
 * "High stress—misuse of data from coupon test"
 * "Corrosion"
 * "Disbond"
 * "Manufacturing defect" (followed by a sub-list):
 * "Surface preparation"
 * "Bond laminate too thin"
 * "Countersink, fastener fit"
 * "Design defect—surface preparation process"

Contextual Linkage:
The image integrates visual representations of aircraft structural components and joint types with textual information about Widespread Fatigue Damage (WFD). The 3D fuselage diagram introduces general components like "Longitudinal skin joint," "Frame," and "Tear straps." The 2D joint diagrams then provide detailed cross-sections of specific joint types (lap, butt, lap with radius) that constitute these general components. The bottom-left text block directly links these visual elements to the specific locations of MSD and MED, for example, identifying "Outer skin upper rivet row" and "Inner skin lower rivet row" within a "Lap joint" as MSD locations. Similarly, "Frames" and "Tear straps" from the 3D diagram are explicitly identified as MED locations, with further detail on "Stress concentration areas" and "Critical fastener rows." The bottom-right text block provides a list of external factors, such as "Corrosion" or "Manufacturing defect," that can influence the development of MSD and MED in these identified structural areas and joint types.](file_wSyEACTCE2d/image004.png)

Figure
A2-1: Longitudinal skin joints, frames, and tear straps (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates the structural components of an aircraft circumferential joint, detailing potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), and listing factors that influence their occurrence.

Detailed Description:
The image is divided into four main sections: a structural isometric view, two cross-sectional diagrams, and two text blocks.

The top-left section displays an isometric, cutaway view of an aircraft structural joint. It shows multiple layers of material fastened together. From top to bottom, the layers are labeled with arrows: "Skin", "Circumferential splice plate", "Stringers", and "Frame". The "Skin" is the outermost curved panel. Beneath it, the "Circumferential splice plate" is a flat, rectangular panel with multiple rows of fasteners, appearing to join two sections of the skin. Perpendicular to the splice plate and beneath it are several "Stringers", which are elongated structural members. The "Frame" is a curved structural member positioned beneath and perpendicular to the stringers. Fasteners, depicted as small, filled circles, are visible connecting these components.

The top-right section presents two simplified cross-sectional diagrams of a joint, side-by-side.
* Diagram (a) is titled "(a) Without outer doubler". It shows two vertical, rectangular plates joined by three horizontal fasteners. The fasteners are evenly spaced vertically along the plates.
* Diagram (b) is titled "(b) With outer doubler". It shows three vertical, rectangular plates joined by fasteners. The leftmost plate (implied as the "outer doubler") is joined to the middle plate by three horizontal fasteners. The middle plate is also joined to the rightmost plate by three horizontal fasteners, aligned with the fasteners connecting the outer doubler to the middle plate.

The bottom-left section contains a text block titled "Type and possible location of MSD/MED". It is a bulleted list:
* "MSD—circumferential joint"
 * "Without outer doubler"
 * "Splice plate—between and/or at the inner two rivet rows"
 * "Skin—forward and aft rivet row of splice plate"
 * "Skin—at first fastener of stringer coupling"
 * "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"

The bottom-right section contains a text block titled "Service or test experience of factors that influence MSD and/or MED (examples)". It is a bulleted list:
* "High secondary bending"
* "High stress level in splice plate and joining stringers (misuse of data from coupon test)"
* "Poor design (wrong material)"
* "Underdesign (over-estimation of interference fit fasteners)"

Contextual Linkage:
The isometric view in the top-left provides a visual representation of the "Circumferential splice plate", "Skin", "Stringers", and "Frame" components, which are explicitly referenced in the "Type and possible location of MSD/MED" section. The cross-sectional diagrams (a) and (b) visually differentiate between joint configurations "Without outer doubler" and "With outer doubler", which are categories under "MSD—circumferential joint" in the text. The text then specifies precise locations for MSD (Multiple Site Damage) within these circumferential joints, such as "Splice plate—between and/or at the inner two rivet rows" or "Skin—outer rivet rows" depending on the presence of an outer doubler. It also identifies locations for MED (Multiple Element Damage) related to "stringer/stringer couplings". The factors listed in the bottom-right text block are examples of conditions or design issues that can contribute to the development of MSD and MED in these structural areas.](file_wSyEACTCE2d/image005.png)

Figure
A2-2: Circumferential joints and stringers (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image presents two cross-sectional diagrams illustrating different types of lap joint configurations susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with associated influencing factors.

Detailed Description:
The image is composed of two distinct sections, arranged horizontally side-by-side, each containing a structural diagram and explanatory text.

The **left section** depicts a structural joint and its characteristics:
* **Diagram (Left):** A vertical cross-sectional view of a two-plate lap joint. The left plate is thicker at its upper portion and transitions via a step-down to a thinner section. This thinner part of the left plate overlaps the right plate, which maintains a uniform thickness. Four equally spaced horizontal dashed lines traverse both overlapping plates, indicating potential fastener locations or reference planes. An arrow points from the upper-left towards the step-down feature on the left plate, labeled "Outer skin at milled or chem-milled step".
* **Text (Left, below diagram):** A heading "Type and possible location of MSD and MED" is followed by a bulleted list:
 * "• MSD—abrupt cross section change"
 * "• Milled radius"
 * "• Chem-milled radius"
 * "• Bonded doubler runout"

The **right section** illustrates another structural joint configuration and influencing factors:
* **Diagram (Right):** A vertical cross-sectional view of a three-plate joint. From left to right, there is a thin plate, a thicker middle plate, and another thin plate. The middle plate is positioned between and overlaps with both the left and right thin plates. Four equally spaced horizontal dashed lines traverse all three overlapping plates. An arrow points from the text "Bonded doubler" to the thicker middle plate. Another arrow points from the text "Bonded joint" to the interface between the left thin plate and the middle plate. A third arrow points from the upper-right towards the top edge of the rightmost plate, labeled "Cracking".
* **Text (Right, below diagram):** A heading "Service or test experience of factors that influence MSD and MED (examples)" is followed by a bulleted list:
 * "• High bending stresses due to eccentricity"

Contextual Linkage:
The left section's diagram and text describe structural features like milled or chem-milled steps and abrupt cross-section changes, which are identified as types and locations where MSD and MED can occur. The right section's diagram illustrates a bonded doubler and joint, indicating a specific location for "Cracking". The accompanying text in the right section provides an example of a factor, "High bending stresses due to eccentricity," that influences MSD and MED, likely in such bonded or eccentric joint configurations. Both sections collectively provide visual examples and associated characteristics/factors for understanding Widespread Fatigue Damage (WFD) in aircraft structures.](file_wSyEACTCE2d/image006.png)

Figure
A2-3: Lap joints with milled, chem-milled or bonded radius (MSD)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates the typical structure of a fuselage skin panel, a magnified view of a fuselage frame with a stringer and tear strap, and provides definitions and influencing factors related to Multiple Element Damage (MED) in fuselage frames.

Detailed Description:
The image is composed of two primary line diagrams, an arrow connecting them, and two text boxes positioned below the diagrams.

The left diagram, labeled "Typical fuselage skin panel" below it, depicts a three-dimensional isometric view of a curved section of an aircraft fuselage. This section shows an outer curved surface identified by an arrow and the label "Fuselage skin panel". Inside this curved panel, longitudinal stiffeners are visible, pointed to by an arrow and labeled "Stringer". Perpendicular to the stringers and following the curvature of the fuselage are structural ribs, indicated by an arrow and labeled "Frame".

A large, curved, solid black arrow originates from the "Typical fuselage skin panel" diagram and points towards the right diagram, indicating a magnified view or detail.

The right diagram presents a magnified, three-dimensional isometric view of a specific structural joint within the fuselage. This detail includes a flat upper surface labeled "Skin" by an arrow. Attached to the underside of the "Skin" is a longitudinal element, indicated by an arrow and labeled "Stringer". A vertical structural member, labeled "Frame" by an arrow, is shown with a cutout. Below the "Stringer" and connected to the "Frame" is a horizontal element, indicated by an arrow and labeled "Tear strap". A series of small, jagged lines, representing cracks, are depicted along the lower edge of the "Frame" cutout, specifically where the "Frame" meets the "Tear strap" and "Stringer". These cracks are pointed to by an arrow and labeled "Typical cracking".

Below these diagrams, two distinct text boxes provide additional information:

The left text box is titled "Type and possible location of MSD/MED". It contains one bullet point:
*   "• 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."

The right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)". It contains two main bullet points, with the second having two sub-bullet points:
*   "• High bending—noncircular frames"
*   "• Local stress concentrations"
    *   "• Cutouts"
    *   "• Shear attachments"

Contextual Linkage:
The image visually explains the concept of Multiple Element Damage (MED) as it applies to fuselage frames. The left diagram provides a general overview of fuselage construction, while the right diagram zooms in on a critical area—the frame at a stringer cutout—to illustrate where "Typical cracking" (MED) occurs. The accompanying text defines MED as cracking in frames at stringer cutouts at successive longitudinal locations, directly correlating with the "Typical cracking" shown in the magnified view. The text also highlights that noncircular frames are a primary concern due to potential panel instability from fractures. Furthermore, the text lists factors influencing MED, such as high bending in noncircular frames and local stress concentrations like cutouts and shear attachments, which are consistent with the structural elements and the location of the depicted cracking.](file_wSyEACTCE2d/image007.png)

Figure
A2-4: Fuselage frames (MED)

![Aviation.Bot AI suggestion, not from EASA source: This image illustrates the structural details of stringer-to-frame attachments in an aircraft fuselage, highlighting typical crack locations associated with Multiple Element Damage (MED).

Detailed Description: The image is a black and white line drawing depicting a section of an aircraft fuselage structure and a magnified detail of a critical attachment point.

On the left, a three-dimensional isometric view shows a curved "Fuselage skin panel" with internal structural elements. Two horizontal elements, labeled "Stringer", are visible running along the curvature of the skin. Perpendicular to the stringers, a vertical element, labeled "Frame", is shown. The stringers are attached to the frame at multiple points.

A large, thick, grey curved arrow originates from the general fuselage structure on the left and points towards a magnified, three-dimensional isometric detail view on the right. This magnified view focuses on a single "Stringer" and "Frame" attachment point, which is also connected to the "Skin". The "Frame" is depicted as a C-channel profile oriented vertically. The "Stringer" is an L-shaped profile, oriented horizontally, and is attached to the vertical web of the frame. The "Skin" is a flat panel forming the base, to which both the frame and stringer are connected. Two black arrows point to specific locations on the stringer-to-frame attachment, labeled "Typical cracks". These cracks are depicted as small, jagged lines originating from the corners of the L-shaped stringer profile where it meets the attachment point to the frame and skin.

Below the structural diagrams, two text boxes provide additional information: The left text box is titled "Type and possible location of MED". It contains a bullet point stating: "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." The right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)". It contains a bullet point stating: "Poor load path connection".

Contextual Linkage: The image visually represents a structural area susceptible to Multiple Element Damage (MED), specifically "Stringer-to-frame attachments." The magnified view explicitly shows "Typical cracks" at the intersection of the stringer, frame, and skin, which aligns with the definition of MED provided in the accompanying text, describing it as fracture of these elements or their attachments leading to a loss of shear tie. The text also suggests that "Poor load path connection" is a factor influencing such damage.](file_wSyEACTCE2d/image008.png)

Figure
A2-5: Stringer-to-frame attachments (MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** This image provides a detailed isometric line drawing illustrating the structural components of a shear-tied fuselage frame, specifically focusing on the shear clip end fasteners, and identifies potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED) cracking, along with influencing factors.

**Detailed Description:** The image displays a three-dimensional, isometric line drawing of an aircraft structural intersection, primarily composed of a skin panel, a longeron or stringer, a frame, and a shear clip.

At the top, a curved panel labeled "Skin" is depicted, showing two parallel rows of 'x' marks representing fasteners. Below the skin, a horizontal, Z-shaped structural member, labeled "Longeron or stringer," is attached to the underside of the skin via fasteners (indicated by 'x' marks). Perpendicular to the longeron/stringer and extending downwards is a vertical, C-shaped structural member labeled "Frame." An L-shaped bracket, labeled "Shear clip," is positioned vertically, connecting the web of the frame to the web of the stringer, secured by fasteners. A large arrow points from the text "Skin/Stringer Attachments" to the general area where the skin and stringer are joined.

Two black arrows originate from the text "Skin cracking at end fasteners" and point to two distinct locations on the "Skin" panel, specifically near the ends of the shear clip's attachment points to the skin. At these indicated locations, small, short, wavy lines are drawn on the skin, representing cracks. Another black arrow originates from the text "Stringer or frame cap cracking" and points to the intersection area where the stringer and frame meet, specifically near the attachment of the shear clip to the stringer. This location also shows small, short, wavy lines indicating cracks.

Below the drawing, two distinct text blocks are presented. The left text block is titled "Type and possible location of MSD and MED" and contains three bullet points: * "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"

The right text block is titled "Service or test experience of factors that influence MSD and MED (examples)" and contains three bullet points: * "Preload" * "Localized bending due to pressure" * "Discontinuous load path"

**Contextual Linkage:** The drawing visually represents the structural configuration of a shear-tied fuselage frame, highlighting critical areas where fatigue damage, specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), is likely to occur. The textual lists below the drawing directly correlate to the visual representation by specifying the types and locations of MSD/MED within this structural assembly and enumerating factors that contribute to their development. The visual crack indications on the skin and stringer/frame cap correspond to the listed "Type and possible location of MSD and MED" bullet points.](file_wSyEACTCE2d/image009.jpg)

Figure
A2-6: Shear clip end fasteners on shear tied fuselage frame (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates the aft pressure dome structure, highlighting specific areas susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with detailed cross-sections of typical outer ring splices and a list of influencing factors.

Detailed Description:
The image is composed of four main sections: a structural diagram, a legend, a list of damage locations, and a list of influencing factors.

1. **Structural Diagram (Top Left):** A black and white wireframe drawing depicts an aircraft's aft pressure dome, viewed in a three-quarter perspective. The dome is circular with multiple concentric rings and radial stiffeners. Two black arrows with text labels point to specific features:
 * An arrow originating from the upper right points to the outermost concentric ring, labeled "Outer ring splice".
 * An arrow originating from the middle right points to one of the radial stiffeners, labeled "Web splices".

2. **Typical Outer Ring Splices (Top Right):** Titled "Typical outer ring splices", this section presents three distinct black and white 2D cross-sectional diagrams, stacked vertically. Each diagram illustrates a different configuration of structural profiles joined together, representing a splice.
 * **Top Diagram:** Shows two L-shaped profiles joined by a fastener.
 * **Middle Diagram:** Shows an L-shaped profile joined to a Z-shaped profile by a fastener.
 * **Bottom Diagram:** Shows two C-shaped profiles joined by a fastener.
 * In each of these three diagrams, two labels are present:
 * "F" with a line pointing to a circular element representing a fastener.
 * "R" with a line pointing to a curved, filleted section of the structural profile, indicating a radius.

3. **Legend (Middle Left):** Located below the aft pressure dome diagram, this section provides definitions for the labels used in the splice diagrams:
 * "Legend:"
 * "F fastener"
 * "R radius"

4. **Type and Possible Location of MSD/MED (Bottom Left):** Titled "Type and possible location of MSD/MED", this section is a bulleted list detailing where MSD/MED may occur:
 * "MSD/MED—outer ring splice"
 * "Attachment profiles—at fastener rows and/or in radius area"
 * "MED—web splices"
 * "Bulkhead skin and/or splice plates—at critical fastener rows"

5. **Service or Test Experience of Factors that Influence MSD and/or MED (examples) (Bottom Right):** Titled "Service or test experience of factors that influence MSD and/or MED (examples)", this section is a bulleted list of factors:
 * "Corrosion"
 * "High stresses—combined tension and compression"
 * "High induced bending in radius"
 * "Inadequate finish in radius—surface roughness"

Contextual Linkage:
The image visually links the general structure of an aft pressure dome to specific areas susceptible to fatigue damage. The "Outer ring splice" and "Web splices" identified in the 3D dome diagram are directly referenced in the "Type and possible location of MSD/MED" section, specifying the type of damage (MSD/MED or MED) and the precise locations (fastener rows, radius area, bulkhead skin/splice plates). The "Typical outer ring splices" diagrams provide visual examples of the "Attachment profiles" and illustrate the "fastener" and "radius" features mentioned as critical areas for damage. The "Service or test experience..." section provides a list of environmental and mechanical factors that contribute to the occurrence of MSD and MED in these identified structural areas.](file_wSyEACTCE2d/image010.png)
 Figure A2-7: Aft pressure
dome outer ring and dome web splices (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image displays a detailed line drawing of an aircraft structural joint, specifically a skin splice at an aft pressure bulkhead, illustrating components, potential cracking locations, and factors influencing Multiple Site Damage (MSD) and Multiple Element Damage (MED).

Detailed Description:
The image consists of a technical line drawing of an aircraft structural assembly in the upper portion, and two blocks of descriptive text in the lower portion.

The drawing depicts a cutaway, isometric view of a structural joint. In the upper left, a section of "Pressurized skin" is shown as a flat panel with a jagged edge on its left and bottom, indicating a cutaway. This skin section features two parallel rows of small, circular fasteners (rivets) running horizontally across its surface. To the right of the "Pressurized skin" and slightly overlapping it, is a section of "Unpressurized skin," also a flat panel with jagged edges on its right and bottom, and two parallel rows of similar circular fasteners. The two skin sections form a lap joint, secured by the fasteners.

Beneath the overlapping skin sections, a "T" frame is depicted. The horizontal flange of the "T" is attached to the underside of the skin panels via fasteners, while the vertical web of the "T" extends downwards. Behind the vertical web of the "T" frame, a "Bulkhead web" is shown as a corrugated or wavy vertical panel, also with jagged edges on its bottom and right, indicating a cutaway. This bulkhead web appears to be attached to the "T" frame.

A horizontal line with small, irregular jagged extensions, labeled "Skin cracking," is visible on the "Pressurized skin" panel, positioned just above and parallel to the upper row of fasteners in the joint. A black arrow points from the text "Skin cracking" to this jagged line. Other black arrows point from the labels "Pressurized skin," "Unpressurized skin," ""T" frame," and "Bulkhead web" to their respective components.

Below the drawing, two distinct text blocks provide additional information:

The left text block is titled "Type and possible location of MSD and MED" and contains one bullet point:
*   "MSD—skin at end fastener holes"

Contextual Linkage:
The diagram visually represents a typical aircraft skin splice at an aft pressure bulkhead, a structural area susceptible to widespread fatigue damage. The "Skin cracking" annotation highlights a potential failure mode, specifically Multiple Site Damage (MSD), which is further elaborated in the text below the diagram as occurring in the "skin at end fastener holes." The "Pressurized skin" and "Unpressurized skin" indicate the differential pressure environment across the bulkhead. The "T" frame and "Bulkhead web" illustrate the supporting structure for this joint. The text blocks provide specific examples of where MSD/MED might occur within this type of structure and list critical stress factors, such as "Shell discontinuous induced bending stresses" and "High load transfer at fastener," that contribute to these damage types.](file_wSyEACTCE2d/image011.png)

Figure A2-8:
Skin splice at aft pressure bulkhead (MSD)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image provides a diagram illustrating structural details susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED) due to abrupt changes in web or skin thickness, along with factors influencing their occurrence.

**Detailed Description:**
The image is divided into two main sections: an upper section with two engineering line drawings and a lower section with two columns of descriptive text.

The upper section features two detailed line drawings of an aircraft structural joint:
*   **Left Drawing (Isometric View):** This drawing depicts a three-dimensional isometric view of a structural joint. It shows a flat, rectangular panel labeled "Web or skin" with a dashed line indicating a bend or fold. Two L-shaped "Edge support member" components are attached perpendicularly to the "Web or skin" panel. Each edge support member is fastened to the web or skin by a vertical row of five hollow circular fasteners. Arrows point from the text "Edge support member" to the L-shaped components and from "Web or skin" to the flat panel. "Typical cracking" is indicated by short, jagged lines originating from the fastener holes and extending into the "Web or skin" panel, with arrows pointing to these cracks. An arrow points to the corner where the "Web or skin" meets an edge support member, labeled "Radius". Below "Radius", a bulleted list specifies types of radii: "Milled", "Chem-milled", and "Bonded doubler".
*   **Right Drawing (Cross-sectional View):** This drawing presents a two-dimensional cross-sectional view of a similar structural joint. It shows a horizontal "Web or skin" panel connected to a vertical "Edge support member" at a corner. Arrows point from the text "Edge support member" to the vertical component. "Typical cracking" is indicated by short, jagged lines originating from the internal corner radius and extending into the horizontal "Web or skin" panel, with arrows pointing to these cracks. Double-headed arrows illustrate stress or load paths: one pair points horizontally into the "Web or skin" from the left, another pair points vertically into the "Edge support member" from the top, and a third pair points diagonally from the corner, indicating stress concentration.

The lower section contains two columns of text:
*   **Left Column: "Type and possible location of MSD and MED"**
    *   The main heading is "Type and possible location of MSD and MED".
    *   A sub-heading "Abrupt change in stiffness*" is followed by a bulleted list:
        *   "Milled radius"
        *   "Chem-milled radius"
        *   "Bonded doubler"
        *   "Fastener row at edge support members"
    *   Another sub-heading "Edge member support structure" is followed by a bulleted list:
        *   "Edge member - in radius areas"
    *   A small asterisk `*` next to "Abrupt change in stiffness" refers to a note at the bottom left: "* Often multiple origins along edge member".
    *   A partial, truncated logo "BOEING" is visible at the very bottom left.
*   **Right Column: "Service or test experience of factors that influence MSD and MED"**
    *   The main heading is "Service or test experience of factors that influence MSD and MED".
    *   A sub-heading "Pressure structure" is followed by a bulleted list:
        *   "High bending stresses at edge support due to pressure"
    *   Another sub-heading "Non-pressure structure" is followed by a bulleted list:
        *   "Structural deflections cause high stresses at edge supports"

**Contextual Linkage:**
The illustrations visually represent the structural configurations where Multiple Site Damage (MSD) and Multiple Element Damage (MED) are likely to occur, specifically at locations of "Abrupt change in stiffness" such as "Milled radius", "Chem-milled radius", "Bonded doubler", and "Fastener row at edge support members". The "Typical cracking" shown in the diagrams corresponds to the manifestation of MSD/MED. The text on the right further explains that these damages are influenced by "High bending stresses at edge support due to pressure" in "Pressure structure" and "Structural deflections cause high stresses at edge supports" in "Non-pressure structure", providing the loading context for the observed cracking.](file_wSyEACTCE2d/image012.png)

Figure
A2-9: Abrupt changes in web or skin thickness — Pressurised or unpressurised
structure (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image provides a line drawing of an aircraft fuselage section, specifically illustrating the window surround structure as an area prone to Multiple Site Damage (MSD) and Multiple Element Damage (MED).

Detailed Description:
The image is a monochrome line drawing depicting an internal, isometric view of a section of an aircraft fuselage. The structure shows a curved panel supported by a grid of interconnected elements. Vertically oriented, curved frames (or ribs) are visible, running from top to bottom. Horizontally oriented stringers (or longerons) run across these frames, forming a series of rectangular bays. Along the lower-middle portion of the structure, there is a horizontal row of eight identical, oval-shaped cutouts, which represent aircraft windows. Each window cutout is surrounded by reinforcing structural elements. Some internal structural lines around the window cutouts are represented by dashed lines, indicating components that are behind the visible surface.

A solid black arrow originates from the upper-center part of the image, pointing downwards and slightly to the left, terminating at the upper-right corner of the third window cutout from the left. Above this arrow, the text label "Window surround structure" identifies the area indicated by the arrow.

Below the main structural drawing, there are two distinct text boxes. The left text box is titled "Type and possible location of MSD/MED" and contains two bullet points:
*   "MSD—skin at attachment to window surround structure"
*   "MED—repeated details in reinforcement of window cutouts or in window corners"

The right text box is titled "Service or test experience of factors that influence MSD and/or MED (examples)" and contains one bullet point:
*   "High load transfer"

Contextual Linkage:
The image visually identifies the "Window surround structure" as a critical area for Widespread Fatigue Damage (WFD), specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED). The arrow explicitly directs attention to this structural region. The accompanying text clarifies that MSD is associated with the skin at the attachment points to the window surround structure, while MED is linked to repeated details within the reinforcement of the window cutouts or in the window corners. Furthermore, the text indicates that "High load transfer" is an example of an operational or test-derived factor that can influence the development of MSD and/or MED in this particular structural area.](file_wSyEACTCE2d/image013.png)

Figure
A2-10: Window surround structure (MSD, MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image displays a line drawing illustrating the structural area of overwing fuselage attachments, identifying it as a potential location for Multiple Element Damage (MED) and listing factors that influence Multiple Site Damage (MSD) and MED.

**Detailed Description:** The image features a black and white line drawing of an aircraft structural cross-section, occupying the upper-middle portion of the frame. The drawing depicts a series of vertical structural members, resembling frames or ribs, connected to horizontal elements, which could be stringers or spars. The uppermost part of the drawing shows a gently curved line, representing a surface. Two black arrows with solid lines originate from text labels on the right side of the drawing and point to specific areas within the structure. The upper arrow points to the curved line at the top of the structure and is labeled "Upper wing surface". The lower arrow points to the complex assembly of vertical and horizontal members directly beneath the "Upper wing surface" and is labeled "Typical fuselage attachments". A thin, solid black semi-circular outline highlights a specific region within the "Typical fuselage attachments" area, indicating a point of interest.

Below the structural drawing, two distinct text blocks are present. The left text block, positioned in the lower-left quadrant, has the heading "Type and possible location of MSD/MED". Below this heading, a single bullet point reads: "• MED—repeated details in overwing fuselage attachments". The right text block, positioned in the lower-right quadrant, has the heading "Service or test experience of factors that influence MSD and/or MED (examples)". Below this heading, two bullet points are listed: "• Manufacturing defect—prestress" "• Induced deflections"

**Contextual Linkage:** The structural drawing visually represents the "overwing fuselage attachments" area, which is explicitly identified in the left text block as a possible location for Multiple Element Damage (MED) due to "repeated details". The semi-circular highlight within the "Typical fuselage attachments" area in the drawing further emphasizes a specific region where such damage might occur. The right text block provides examples of factors, such as "Manufacturing defect—prestress" and "Induced deflections", that are known from service or test experience to influence the occurrence of both Multiple Site Damage (MSD) and MED in such structural areas.](file_wSyEACTCE2d/image014.png)

Figure
A2-11: Over wing fuselage attachments (MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates typical structural areas and components of an aircraft door, specifically focusing on hinges and latches, to identify potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with factors influencing their occurrence.

Detailed Description:
The image is divided into two primary sections, each containing a diagram and associated descriptive text.

The left section depicts an aircraft door and its hinge mechanism:
* **Top Diagram (Door):** An isometric line drawing of a "Door" component, likely an aircraft access panel or small door. The door features a grid-like internal structure visible through its outer skin. Two circular callout regions are marked with arrows and labels: "A" points to the upper-left hinge area, and "B" points to the lower-right latch area. The text "Door" is positioned below the diagram, pointing to the door structure.
* **Bottom Diagram (View A):** A detailed engineering line drawing, labeled "View A", showing a cross-sectional or exploded view of a "Hinge lobe" and "Hinge pin" assembly. Multiple hinge lobes, depicted as rectangular blocks with central holes, are shown aligned horizontally. A "Hinge pin" is shown passing through the holes of these lobes, securing them. Arrows point from the labels "Hinge lobe" to the individual lobes and "Hinge pin" to the pin.
* **Text below View A:** Titled "Type and possible location of MSD/MED", this section lists specific areas:
 * "MSD—piano hinge"
 * "At hinge fastener attachment row"
 * "In fillet radius"
 * "Emanating from hole in lobes"
 * "MED—latches"
 * "In multiple latch hooks"
 * "At lube channel of latch spool"
 * "At spool bracket attach bolts (also corrosion)"

The right section illustrates a latch hook and a latch mechanism:
* **Top Diagram (Latch hook):** A side view line drawing of a "Latch hook" component. It is a curved, hook-shaped part with a circular pivot point at its upper end. An arrow points from the label "Latch hook" to the hook-shaped part.
* **Bottom Diagram (View B):** A detailed engineering line drawing, labeled "View B", showing a cross-sectional or cutaway view of a "Latch" mechanism. This mechanism consists of a cylindrical spool-like component with internal features. An arrow points from the label "Lubrication passage" to a channel within the cylindrical component. The mechanism is shown secured by two "Attach bolts" at its base, with an arrow pointing from the label "Attach bolts" to one of the bolts.
* **Text below View B:** Titled "Service or test experience of factors that influence MSD and/or MED (examples)", this section lists influencing factors:
 * "Bending stresses due to fuselage elongation"
 * "High local stress"
 * "Fretting"

Contextual Linkage:
The image visually supports the identification of fatigue-critical structures susceptible to Widespread Fatigue Damage (WFD), specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), as applied to aircraft door components. "View A" provides a detailed look at the hinge mechanism, corresponding to the "MSD—piano hinge" locations listed in the text. "View B" details a latch mechanism, which is relevant to the "MED—latches" locations. The overall "Door" diagram indicates where these detailed views (A and B) are located on a typical door. The text also provides general factors that can influence the development of MSD and MED in these and similar structural areas.](file_wSyEACTCE2d/image015.jpg)
Figure A2-12:
Latches and hinges of non-plug doors (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates a structural detail of an aircraft, specifically the attachment of a skin doubler to the skin, highlighting areas susceptible to Multiple Site Damage (MSD) and Multiple Element Damage (MED).

Detailed Description:
The image displays a black and white engineering diagram of an aircraft structural assembly, viewed from above or as a flattened section. The primary components shown are a "Skin" and a "Skin doubler," which are joined together by numerous fasteners.

The "Skin" is depicted as a continuous, flat panel forming the lower layer of the assembly. Its visible edge is a solid, slightly irregular line at the bottom of the diagram.

The "Skin doubler" is positioned above the skin, appearing as a series of four rectangular segments arranged horizontally. These segments are connected by three narrower, vertical strips, creating a grid-like pattern. The outlines of these doubler segments and connecting strips are indicated by dashed lines, suggesting they are either hidden or represent the extent of the component. The doubler segments have rounded corners where they meet the connecting strips.

Both the skin and the doubler are secured by various fasteners, represented by small circles and plus signs.
*   Each rectangular doubler segment is fastened to the skin by two horizontal rows of small circles along its top and bottom edges, and vertical columns of small circles along its left and right edges. Additionally, four small circles are arranged in a square pattern within the central area of each rectangular segment.
*   The three vertical connecting strips between the doubler segments are fastened to the skin by vertical columns of plus signs. Each column contains five plus signs.
*   Horizontal rows of plus signs are also visible above and below the main doubler structure, extending across the entire width of the depicted assembly, suggesting attachment to other structural elements not fully shown.

Two black arrows originate from the bottom of the diagram and point upwards to specific components:
*   One arrow points to the continuous lower panel, labeled "Skin".
*   Another arrow points to the segmented upper panel, labeled "Skin doubler".

Below the main diagram, there are two text blocks providing additional information:
The first text block is titled "Type and possible location of MSD/MED" and contains one bullet point:
*   "MSD—cracks initiated at multiple critical fastener holes in skin at runout of doubler"

The second text block is titled "Service or test experience of factors that influence MSD and/or MED (examples)" and contains one bullet point:
*   "High load transfer—high local stress"

Contextual Linkage:
The diagram visually represents a structural area, specifically the "Skin at runout of large doubler," which is identified as susceptible to Multiple Site Damage (MSD). The accompanying text clarifies that MSD cracks are expected to initiate at multiple critical fastener holes in the skin, particularly at the "runout" (the edge or termination) of the doubler. The complex pattern of fasteners shown in the diagram facilitates the transfer of loads between the skin and the doubler. The text further links this phenomenon to "High load transfer" leading to "high local stress," which are factors influencing the occurrence of MSD and/or MED in such structural configurations.](file_wSyEACTCE2d/image016.png)

Figure
A2-13: Skin at runout of large doubler (MSD) — Fuselage, wing or empennage

![Aviation.Bot AI suggestion, not from EASA source: The image illustrates a typical wing or empennage chordwise splice, detailing its components and identifying potential locations for Multiple Site Damage (MSD) and Multiple Element Damage (MED), along with influencing factors.

Detailed Description:
The image is divided into two main visual sections: a simplified schematic on the left and a detailed isometric view on the right, connected by an arrow, with explanatory text blocks below.

The left section displays a two-dimensional schematic of an aircraft structural area, likely a wing or empennage. A prominent vertical solid line on the far left is labeled "Front spar". To its right, a series of horizontal solid lines extend across the width of the schematic, representing "Chordwise joints", indicated by two arrows pointing to the top and bottom horizontal lines. Between these solid horizontal lines, multiple dashed horizontal lines are evenly spaced, suggesting internal structural elements. A small, empty circle highlights a specific intersection point on one of the chordwise joints, approximately in the middle of the schematic's height. A thick, curved gray arrow originates from this circle and extends towards the right, pointing to the detailed isometric view.

The right section presents a three-dimensional, cutaway isometric view of a "Typical skin and stringer splice", as labeled at the top. This view shows several interconnected structural components:
* A flat, elongated component at the bottom is labeled "Stringer" with an arrow pointing to its main body.
* Above and slightly to the right of the stringer, a thinner, flat panel is labeled "Skin panel" with an arrow pointing to its surface.
* Overlapping both the stringer and the skin panel, a complex, multi-lobed component with several circular fastener holes is labeled "Splice plate" with an arrow pointing to its main body.
* Above the splice plate and skin panel, a vertical, corrugated-like structural member is labeled "Rib" with an arrow pointing to its upper edge.
* To the left of the splice plate, partially obscured by the rib, another component with fastener holes is labeled "Fitting" with an arrow pointing to its main body. The components are depicted with visible thickness and some internal details, suggesting a cross-sectional or cutaway representation.

Below the left schematic, a text block is titled "Type and possible location of MSD/MED". It contains two main bullet points with sub-points:
* "MSD—skin and/or splice plate"
 * "Chordwise critical fastener rows" (indented)
* "MED—stringer runout of fitting"
 * "Fatigue-critical fastener holes at stringer and/or fitting" (indented)

Below the right isometric view, a text block is titled "Service or test experience of factors that influence MSD and/or MED (examples)". It contains two bullet points:
* "High load transfer"
* "Local bending"

Contextual Linkage:
The image visually defines and illustrates the "Wing or empennage chordwise splices" as a structural area susceptible to widespread fatigue damage. The left schematic provides a general context of where such a splice is located within a larger structure (e.g., relative to a front spar and other chordwise joints). The arrow and circled area indicate that the detailed isometric view on the right is a magnified representation of this specific splice area. The accompanying text blocks directly relate to this detailed splice, identifying specific components and locations within it where Multiple Site Damage (MSD) and Multiple Element Damage (MED) are likely to occur, and listing operational or test-derived factors that contribute to these damage types. This comprehensive depiction aids in understanding the critical aspects of fatigue damage in this particular structural configuration.](file_wSyEACTCE2d/image017.png)

Figure
A2-14: Wing or empennage chordwise splices (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** This image provides a visual representation of typical aircraft structural components involved in rib-to-skin attachments, illustrating potential locations of fatigue cracking, and defines associated Multiple Site Damage (MSD) and Multiple Element Damage (MED) along with influencing factors.

**Detailed Description:** The image consists of three blue-line diagrams of aircraft structural components, accompanied by text labels and two distinct bulleted lists.

In the upper left, a long, horizontally oriented, corrugated structural element is depicted, labeled "Typical shear tie". An oval highlights a specific section of this shear tie, from which two black arrows extend downwards and to the right, pointing towards the "Skin" component in the central diagram.

In the upper right, a smaller, irregularly shaped, flat structural element is shown, labeled "Typical skin cracking". This element features two rows, each containing three circular holes. Black arrows point from the label to short, linear crack indications originating from some of these circular holes.

The central and largest diagram presents a detailed cross-sectional view of an aircraft structural joint. A horizontal element at the top is labeled "Skin". To its right, a vertical element is labeled "Stringer", which is depicted as attached to the underside of the "Skin". Below and to the right of the "Stringer", a larger, vertical, and somewhat L-shaped element is labeled "Rib web". The "Rib web" is shown connected to both the "Skin" and the "Stringer" via multiple circular holes, representing fastener locations. Four black arrows, originating from the label "Typical cracking", point to various locations: two arrows indicate cracks along the interface between the "Skin" and the "Stringer", and two other arrows point to cracks within the "Skin" itself, specifically near the attachment points to the "Rib web".

Below the structural diagrams, two distinct text boxes are presented.
The left text box is titled "Type and possible location of MSD and MED" and contains two bullet points:
*   "MSD—critical fasteners in skin along rib attachments"
*   "MED—critical rib feet in multiple stringer bays (particularly for empennage under sonic fatigue)"

The right text box is titled "Service or test experience of factors that influence MSD and MED (examples)" and contains two bullet points:
*   "Manufacturing defect—prestress due to assembly sequence"
*   "Sonic fatigue (empennage)"

**Contextual Linkage:** The diagrams visually illustrate the structural configuration of "Rib-to-skin attachments" and highlight areas susceptible to fatigue damage, specifically "Typical cracking" in the skin and at component interfaces. The accompanying text defines the specific manifestations of Widespread Fatigue Damage (WFD) relevant to this structural area: Multiple Site Damage (MSD) is characterized by critical fasteners in the skin along rib attachments, and Multiple Element Damage (MED) by critical rib feet in multiple stringer bays, particularly in the empennage under sonic fatigue. The text further provides examples of factors that can contribute to or influence the development of MSD and MED in these areas, such as manufacturing defects leading to prestress due to assembly sequence, and sonic fatigue.](file_wSyEACTCE2d/image018.png)

Figure
A2-15: Rib-to-skin attachments (MSD/MED)

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image illustrates and compares the characteristics of crack propagation, specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), in two distinct aircraft structural designs: Riveted Skin and Stringer Construction versus Integrally Stiffened Skins.

**Detailed Description:** The image is divided vertically into two main sections, each detailing a different aircraft structural construction type and its associated fatigue damage characteristics.

The left section, titled "Riveted Skin and Stringer Construction (MSD & MED)", presents four diagrams and descriptive text:
* **Top-left diagram (a):** An isometric view of a riveted skin and stringer panel. It shows multiple parallel stringers (or ribs) attached to a flat skin panel by rows of fasteners. Black arrows point to several locations along the stringers, labeled "(a)", indicating "Rib attachment holes" where small, irregular "Cracks" are depicted originating from the fastener holes.
* **Top-right diagram (b):** A partial cross-sectional view of a skin panel. A horizontal dashed line represents a structural boundary. A circular opening, identified as a "Drain hole", is shown with an irregular crack extending downwards from it, labeled "Drain hole crack".
* **Middle-left diagram (a):** A side-profile view of a stringer attached to a skin. A black arrow points from the text "Fastener attachment to rib" to a row of circular fastener holes along the top edge of the stringer. Below these holes, several small, irregular "Cracks" are shown originating from the holes. This diagram is also labeled "(a)".
* **Bottom-right diagram (c):** A cross-sectional view illustrating a vertical "Root rib, tank end, etc." stiffener attached to a horizontal "Skin" panel. The stiffener has a base flange. Black arrows point to the junction of the stiffener's flange and the skin, where multiple small, irregular "Cracks" are depicted. One crack is specifically labeled "Crack". This diagram is labeled "(c)".
* Below these diagrams, the text states: "Inherent fail safe and crack stopper characteristics". This is followed by two bullet points:
 * "MSD—chordwise cracks link up at"
 * "a) Rib attachment holes"
 * "MED—"
 * "b) Drain or vent holes"
 * "c) Stiffener run-outs at root rib or tank end rib"
* The "BOEING" logo is present in the bottom-left corner of this section.

The right section, titled "Integrally Stiffened Skins (MSD)", also presents four diagrams and descriptive text:
* **Top-left diagram (d):** An isometric view of an integrally stiffened skin panel. This panel features parallel, corrugated-like stiffeners machined directly from the skin material. Black arrows point to two distinct areas: one labeled "(d)" at the base of the stiffeners where they meet the skin, showing small, irregular "Cracks"; and another labeled "(e)" pointing to a crack along the length of one of the stiffeners.
* **Top-right diagram (e):** A close-up isometric view of a single integrally stiffened section. A horizontal crack is shown across the web of the stiffener, labeled "(e)".
* **Middle-left diagram (d):** A side-profile view of an integrally stiffened panel. Similar to the top-left view, small, irregular "Cracks" are shown originating from the base of the stiffeners where they meet the skin, labeled "(d)".
* **Bottom-right diagram (f):** A cross-sectional view illustrating a vertical "Root rib, tank end, etc." stiffener attached to the integrally stiffened skin. This diagram shows the structural configuration but does not explicitly depict cracks.
* Below these diagrams, the text states: "Do not have inherent crack stopper characteristics of riveted skin and stringer construction". This is followed by two bullet points:
 * "MSD—Chordwise cracks link up at"
 * "d) Rib attachment holes"
 * "e) Drain or vent holes"
 * "f) Stringer run-outs at root rib or tank end rib"
 * "MED—becomes MSD"

**Contextual Linkage:** The image visually differentiates how fatigue damage, specifically Multiple Site Damage (MSD) and Multiple Element Damage (MED), manifests in two common aircraft structural designs. For "Riveted Skin and Stringer Construction", the diagrams (a), (b), and (c) illustrate specific locations where MSD (rib attachment holes) and MED (drain/vent holes, stiffener run-outs) occur, correlating with the descriptive text that highlights the "Inherent fail safe and crack stopper characteristics" of this design. In contrast, for "Integrally Stiffened Skins", diagrams (d), (e), and (f) show crack locations for MSD (rib attachment holes, drain/vent holes, stringer run-outs), and the accompanying text explicitly states that this design "Do not have inherent crack stopper characteristics of riveted skin and stringer construction" and that "MED—becomes MSD", implying a different failure mode for multiple element damage in this type of structure. The labels (a) through (f) in the diagrams directly correspond to the bulleted lists explaining the crack types and locations for each construction method.](file_wSyEACTCE2d/image019.png)

Figure A2-16:
Typical wing and empennage construction (MSD/MED)

4.3       WFD Evaluation

Point 26.300 of Part-26 requires an LOV to
be established according to specified timescales for large transport
aeroplanes with MTOWs above 34 901 kg (75 000 lb). For
other types, it is recommended that by the time the highest-time aircraft of a
particular model reaches its DSG, the evaluation for each area susceptible to
the development of WFD should be completed. A typical evaluation process is
shown in Figure A2-19 below. This evaluation will establish the necessary
elements to determine a maintenance programme to preclude WFD in that
particular model’s aircraft fleet. These elements are developed for each
susceptible area and include:

**4.3.1   Identification
of structure potentially susceptible to WFD**

Unless already fully addressed in the
existing fatigue and damage tolerance evaluation, the TCH should identify each
part of the aircraft’s structure that is potentially susceptible to WFD for
further evaluation. A justification should be given that supports selection or
rejection of each area of the aircraft structure. DAHs for modified or repaired
structure should evaluate their structure and its effect on existing
structure.

Typical examples of structure susceptible to
WFD are included in paragraph 4.2 of this Appendix.

**4.3.2   Predicting
when WFD will occur**

(a)       Characterisation
of events leading to WFD

The fatigue process that leads to WFD is
shown in Figure A2-17. This figure is applicable to both damage that
occurs in multiple sites (MSD) and damage that occurs in similar structures at
more than one location (MED). For any susceptible structural area, it is not a
question of whether WFD will occur, but when it will occur. In Figure A2-17,
the ‘when’ is illustrated by the line titled ‘WFD (average behaviour),’
which is the point when 50 % of the aeroplanes in a fleet would have
experienced WFD in the considered area (note that the probability density
function for flight cycles or flight hours to WFD has been depicted for
reference). The WFD process includes this phase of crack initiation and a
crack growth phase. During the crack initiation phase, which generally spans a
long period of time, there is little or no change in the basic strength
capability of the structure. The actual residual strength curve depicted in
Figure A2-17 is flat, and equal to the strength of the structure in its
pristine state. However, at some time after the first small cracks start to
grow, residual strength begins to degrade. Crack growth continues until the
capability of the structure degrades to the point of the minimum strength
required by CS 25.571(b). In this context, the line in Figure A2-17
called WFD (average behaviour) represents a point when 50 %
of the aeroplanes in a fleet fall below the minimum strength specifications of
CS 25.571(b).

![Aviation.Bot AI suggestion, not from EASA source: This image is a two-dimensional graph illustrating the relationship between flight cycles, crack initiation and growth, residual strength, and the probability of widespread fatigue damage (WFD) in an aircraft structure.

Detailed Description:
The image displays a graph with "FLIGHT CYCLES" labeled on the horizontal axis. The vertical axis implicitly represents different quantities: residual strength in the upper portion and crack length in the lower portion.

In the upper section of the graph:
* A horizontal, wavy, dotted line is labeled "REQUIRED RESIDUAL STRENGTH". This line maintains a consistent level across the graph.
* A horizontal, wavy, solid line is labeled "ACTUAL RESIDUAL STRENGTH". This line starts at the same level as the "REQUIRED RESIDUAL STRENGTH" line and remains constant for a significant duration. Towards the right side of the graph, this solid line begins to curve downwards, eventually dropping below the "REQUIRED RESIDUAL STRENGTH" dotted line.
* In the upper right quadrant, a bell-shaped curve is depicted, labeled "Probability density function for cycles to WFD condition". The peak of this curve aligns vertically with the point where the "ACTUAL RESIDUAL STRENGTH" curve experiences a significant downward inflection.

In the middle section of the graph, below the residual strength curves:
* A horizontal segment on the left is labeled "MSD/MED CRACK INITIATION".
* To the right of this, separated by a vertical line, another horizontal segment is labeled "MSD/MED CRACK GROWTH". This segment extends further to the right than the "CRACK INITIATION" segment.

In the lower section of the graph:
* A curve labeled "CRACK LENGTH" originates near the origin of the graph, staying very close to the horizontal axis for an extended period. After a certain point, it begins to curve sharply upwards, indicating an accelerating increase in crack length.

Vertical lines and labels:
* A vertical dashed line extends from the peak of the "Probability density function for cycles to WFD condition" curve downwards. This dashed line intersects the "ACTUAL RESIDUAL STRENGTH" curve during its downward slope and the "CRACK LENGTH" curve during its steep upward slope.
* At the bottom of the graph, this dashed line points to the label "WFD (average behaviour)".
* A solid vertical line is positioned slightly to the right of the dashed "WFD (average behaviour)" line, indicating a point further along the "FLIGHT CYCLES" axis.

In the top left corner of the graph, a rectangular box contains a note: "NOTE: No scale implied. For definition use only."

Contextual Linkage:
The graph illustrates the progression of fatigue damage over "FLIGHT CYCLES". Initially, during the "MSD/MED CRACK INITIATION" phase, the "CRACK LENGTH" is negligible, and the "ACTUAL RESIDUAL STRENGTH" remains constant and equal to the initial strength, staying above the "REQUIRED RESIDUAL STRENGTH". As the process transitions into "MSD/MED CRACK GROWTH", the "CRACK LENGTH" begins to increase, initially slowly, then at an accelerating rate. This increase in "CRACK LENGTH" directly correlates with the degradation of "ACTUAL RESIDUAL STRENGTH", causing it to eventually fall below the "REQUIRED RESIDUAL STRENGTH". The "WFD (average behaviour)" point, marked by the vertical dashed line, represents the flight cycles at which the probability of reaching the WFD condition is highest, coinciding with a significant reduction in "ACTUAL RESIDUAL STRENGTH" due to substantial "CRACK LENGTH".](file_wSyEACTCE2d/image020.jpg)

Figure A2-17: Effect on residual strength of developing WFD

(b)       Determination
of WFD (average behaviour) in the fleet

The time in terms of flight cycles/hours
defining the WFD (average behaviour) in the fleet should be
established for each susceptible structural area. The data to be assessed in
determining the WFD (average behaviour) includes:

—        a complete review of
the service history of the susceptible areas, to identify any occurrences of
fatigue cracking and the continuing validity of loads and mission profiles,

—        evaluation of the
operational statistics of the fleet in terms of flight hours and landings,

—        significant
production variants (material, design, assembly method, and any other change
that might affect the fatigue performance of the detail),

—        fatigue test evidence
including relevant full-scale and component fatigue and damage tolerance test
data (see subparagraph 4.3.9 and [Annex 1](#_DxCrossRefBm421436825) for
more details),

—        teardown inspections,
and

—        any fractographic
analysis available.

The evaluation of the test results for the
reliable prediction of the time to when WFD might occur in each susceptible
area should include appropriate test-to-structure factors. If full-scale
fatigue test evidence is used, Figure A2‑20 below relates how that data
might be utilised in determining WFD (average behaviour).
Evaluation may be analytically determined, supported by test and, where
available, service evidence.

**4.3.3   Initial
crack/damage scenario**

This is an estimate of the size and extent
of multiple cracking expected at MSD/MED initiation. This prediction requires
empirical data or an assumption of the crack/damage locations and sequence
plus a fatigue evaluation to determine the time to MSD/MED initiation.
Alternatively, analysis can be based on either:

—        the distribution of
equivalent initial flaws, as determined from the analytical assessment of
flaws found during fatigue test and/or teardown inspections regressed to zero cycles;
or

—        a distribution of
fatigue damage determined from relevant fatigue testing and/or service
experience.

**4.3.4   Final
cracking scenario**

This is an estimate of the size and extent
of multiple cracking that could cause residual strength to fall to the minimum
required level as shown in A2-17. Techniques exist for 3-D elastic-plastic
analysis of such problems; however, there are several alternative test and
analysis approaches available that provide an equivalent level of safety. One
such approach is to define the final cracking scenario as a subcritical
condition (e.g. first crack at link-up). The use of a subcritical scenario
reduces the complexity of the analysis and, in many cases, will not greatly
reduce the total crack growth time, because the majority of the time taken to
reach the critical condition is generally in the initiation phase.

**4.3.5   Crack
growth calculation**

Progression of the crack distributions from
the initial cracking scenario to the final cracking scenario should be
developed. These curves can be developed:

—        *analytically*,
typically based on linear elastic fracture mechanics, or

—        *empirically*,
from test or service fractographic data.

**4.3.6   Potential
for discrete source damage (DSD)**

A structure susceptible to fatigue including
MSD/MED may also be affected by DSD due to an uncontained failure of
high-energy rotating machinery (i.e. turbine engines). At this time, there is
no specific requirement to address prior fatigue cracking in combination with
DSD for certification. Nonetheless, when assessing in‑service findings of
fatigue cracking, the additional threat posed by any potential DSD should be
taken into account when developing the corrective actions and the timescales
for its implementation.

**4.3.7   Analysis
methodology**

**Differences between multiple-site damage
and multiple-element damage**

Details of the approach used to characterise
events leading up to WFD may be different. The differences will largely depend
on whether MSD or MED is being considered. This is especially true for crack
interaction.

(a)       Crack
interaction

MSD has the potential for strong crack
interaction, and the effect of multiple cracks on each other needs to be
addressed. MED, in most cases, does not have the same potential for strong
crack interaction. The differences between the interaction effects for MSD and
MED are illustrated in Figure A2-18.

(b)       MSD
and MED interaction

Some areas of
an aeroplane are potentially susceptible to both MSD and MED. Simultaneous
occurrence of MSD and MED is possible, even though it is not common. A
comparison of inspection start points (ISPs) or modification start points
might indicate the possibility of this occurrence. If so, the evaluation
should consider the interaction between MSD and MED.

![Aviation.Bot AI suggestion, not from EASA source: **Overall Summary:** The image illustrates the conceptual differences in crack interaction and progression over cycles (N) and crack size (a) between Multiple-Site Damage (MSD) and Multiple-Element Damage (MED).

**Detailed Description:** The image is divided into two main conceptual sections: Multiple-Site Damage (MSD) on the left and Multiple-Element Damage (MED) on the right. Both sections feature a graph at the top and schematic representations of structural elements with cracks at the bottom.

**Left Section (MSD):**
* **Top Left Graph:** This is a 2D Cartesian coordinate system with the Y-axis labeled "N (cycles)" (representing fatigue life) and the X-axis labeled "a (mm)" (representing crack size). Along the X-axis, three white circles with thick black outlines are depicted, representing fastener holes or similar stress concentration points.
 * **"Multiple cracks" curve:** A solid black line forms three distinct troughs, each centered above one of the white circles. The curve originates from the X-axis near each circle, rises, then dips significantly between adjacent circles, and rises again. This illustrates crack growth and interaction leading to potential link-up between adjacent sites. The text "Multiple cracks" is positioned to the left of the first solid curve.
 * **"Single crack" curve:** A dotted black line originates from the X-axis above the middle white circle, rises sharply, and then curves upwards and to the right, extending beyond the visible graph area. This illustrates the growth of a single, isolated crack without interaction from adjacent sites. The text "Single crack" is positioned above the dotted curve.
* **Bottom Left Schematics:** Two horizontal rectangular shapes, representing a structural panel with a row of three holes.
 * **"Single crack" schematic:** The top schematic shows a light grey rectangular panel with three white circles (holes) with black outlines. A short, dotted black line extends horizontally from the right side of the middle circle, indicating a single crack. The text "Single crack" is positioned to the right of this schematic, connected by a dotted line.
 * **"Multiple cracks" schematic:** The bottom schematic shows a light grey rectangular panel with three white circles (holes) with black outlines. Each of the three circles has short, solid black lines extending horizontally from both its left and right sides, indicating multiple cracks originating from each hole. The text "Multiple cracks" is positioned to the right of this schematic, connected by a solid line.
* **Overall Label:** Below the bottom schematic, a black rectangular box contains the text "MSD".

**Right Section (MED):**
* **Top Right Graph:** This is a 2D Cartesian coordinate system with the Y-axis labeled "N (cycles)" and the X-axis labeled "a (mm)". Along the X-axis, three light grey rectangular blocks are depicted, representing individual structural elements. Each block has a short, solid black line extending downwards from its top edge, indicating a crack.
 * **"Multiple cracks" curves:** Three distinct solid black lines originate from the top of each rectangular block, rising sharply upwards and slightly converging, but remaining separate. This illustrates crack growth in multiple independent elements. The text "Multiple cracks" is positioned to the left of the first solid curve.
 * **"Single crack" curve:** A dotted black line originates from the top of the middle rectangular block, rises sharply upwards, and then curves slightly outwards. This illustrates the growth of a single crack in an isolated element. The text "Single crack" is positioned above the dotted curve.
* **Bottom Right Schematics:** Two rows of three vertical rectangular shapes, representing individual structural elements.
 * **"Single crack" schematic:** The top row shows three light grey vertical rectangular elements. Only the middle element has a short, dotted black line extending horizontally from its center, indicating a single crack. The text "Single crack" is positioned to the left of this row, connected by a dotted line.
 * **"Multiple cracks" schematic:** The bottom row shows three light grey vertical rectangular elements. Each of the three elements has a short, solid black line extending horizontally from its center, indicating multiple cracks. The text "Multiple cracks" is positioned to the left of this row, connected by a solid line.
* **Overall Label:** Below the bottom schematic, a black rectangular box contains the text "MED".

**Contextual Linkage:** The image visually distinguishes between Multiple-Site Damage (MSD) and Multiple-Element Damage (MED) based on their crack interaction characteristics. For MSD, the graphs and schematics demonstrate that cracks at adjacent sites (e.g., fastener holes) can interact and potentially link up, influencing each other's growth. This is shown by the "Multiple cracks" curve dipping between sites. For MED, the graphs and schematics illustrate that cracks in multiple distinct structural elements tend to grow independently, with minimal or no interaction between cracks in adjacent elements, as indicated by the "Multiple cracks" curves rising separately. The "Single crack" representations serve as a baseline for comparison, showing crack growth in an isolated context for both damage types.](file_wSyEACTCE2d/image021.png)

Figure
A2-18: Difference between MSD and MED interaction effects

The report ‘Recommendations for Regulatory
Action to Prevent Widespread Fatigue Damage in the Commercial Aeroplane Fleet’,
Revision A, dated June 29, 1999 (a report of the AAWG for the ARAC’s
Transport Aircraft and Engine Issues Group), discusses two Round Robin
exercises developed by the TCHs to provide insight into their respective
methodologies. One outcome of the exercises was an identification of key
assumptions or methods that had the greatest impact on the predicted WFD
behaviour. These assumptions were:

—        the flaw sizes
assumed at initiation of crack growth phase of analysis;

—        material properties
used (static, fatigue, fracture mechanics);

—        ligament failure
criteria;

—        crack growth
equations used;

—        statistics used to evaluate
the fatigue behaviour of the structure (e.g. time to crack initiation);

—        methods of
determining the structural modification point (SMP);

—        detectable flaw size
assumed;

—        initial distribution
of flaws; and

—        factors used to
determine bound behaviour as opposed to mean behaviour.

When considering MED, where interaction
between cracks in different elements is not a factor, the following should be
considered:

(1)       In
a structure containing large numbers of similar elements, there is not
normally a high probability that, after a crack initiates in an element, a
second crack will initiate in the element right next to it. If this does
happen, however, the consequences to the overall structure may be severe. This
is because having two structural members fail right next to each other can
completely negate any ability of the structure to tolerate additional damage.
Consequently, when performing an evaluation, applicants should make
conservative assumptions and assume failures to be adjacent to each other.

(2)       When
an element fails completely, the load that has to be redistributed onto the
non-failed structure can be large and can have a significant impact on the
strength of both the cracked and uncracked structure; therefore, the effects
of load redistribution must be included in the evaluation.

(d)       Establishing
maintenance actions

The following parameters are developed from
paragraphs 4.3.2 to 4.3.7 above, and are necessary to establish an MSD/MED
maintenance programme for the area under investigation.

Fatigue damage is the gradual deterioration
of a material subjected to repeated loads. This gradual deterioration is a
function of use and can be statistically quantified. The term ‘WFD’ is used,
and can be statistically quantified, at the end of the deterioration process
when the structure is no longer able to carry the residual strength loads. WFD
can never be absolutely precluded because there is always some probability, no
matter how small, that it will occur. Therefore, modifying or replacing structure
at a predetermined, analytically-derived time stated in flight cycles or
flight hours, minimises the probability of having WFD in the fleet.
Modification or replacement is the most reliable method for precluding WFD.
The point at which a modification is undertaken is referred to as the
‘structural modification point (SMP)’ and it is illustrated in Figure 2-1 of [Annex 2](#_DxCrossRefBm421436824). The SMP is generally a
fraction of the number representing the point in time when WFD (average
behaviour) will occur, and should result in the same reliability as a
successful two-lifetime fatigue test. This level of reliability for setting
the SMP is acceptable if MSD or MED inspections are shown to be effective in
detecting cracks. If the inspections are effective, they should be implemented
before the SMP. The implementation times for these inspections are known as
the ‘inspection start points (ISPs)’. Repeat inspections are usually necessary
to maintain this effectiveness in detecting cracks. If MSD or MED inspections
are not effective in detecting cracks, then the SMP should be set at the time
of ISP. For the purposes of this AMC, an inspection is effective if, when
performed by properly trained maintenance personnel, it will readily detect
the damage in question[[9]](#_ftn9).
The SMP should minimise the extent of cracking in the susceptible structural
area in a fleet of affected aeroplanes. In fact, if this point is
appropriately determined, a high percentage of aeroplanes would not have any
MSD or MED by the time the SMP is reached.

Due to the redundant nature of
semi-monocoque structures, MED can be difficult to manage in a fleet
environment. This stems from the fact that most aircraft structures are
built-up in nature, and that makes the visual inspection of the various layers
difficult. Also, visual inspections for MED typically rely on internal
inspections, which may not be practical at the frequency necessary to preclude
MED due to the time required to gain access to the structure. However, these
issues are dependent on the specific design involved and the amount of damage
being considered. In order to implement a viable inspection programme for MED,
static stability must be maintained at all times and there should be no MED
concurrent with MSD in a given structural area.

**4.3.8   Inspection
start point (ISP)**

This is the point at which inspection starts
if a monitoring period is used. Inspection is not practical for all
applications and cannot replace the SMP. The ISP is determined through a
statistical analysis of crack initiation based on fatigue testing, teardown,
or service experience of similar structural details. It is assumed that the
ISP is equivalent to a lower bound value with a specific probability in the statistical
distribution of cracking events. Alternatively, the ISP may be established by
applying appropriate factors to the average behaviour.

When inspections are determined to be
effective, it is necessary to establish when those inspections should start.
This point is illustrated in Figure 2-1. The start point is determined through
a statistical analysis of crack initiation based on fatigue testing, teardown,
or in-service experience of similar structure. The ISP is assumed to be
equivalent to a lower-bound value with a specific probability in the
statistical distribution of cracking events. Alternatively, an ISP may be
established by applying appropriate factors to the number representing WFD (average
behaviour). (e.g. for aluminium alloy structure, dividing the full-scale
test result by a factor of 3).

For inspection intervals, see point 4.3.10.

**4.3.9   Structural
modification point (SMP)**

The SMP should be established as a point in
time when structures should be modified or replaced to prevent WFD from
occurring. This is typically established by:

—        calculating when WFD
would first occur in the structure (predicted using the WFD (average
behaviour)),

—        setting a time before
the predicted occurrence of WFD to perform modifications or replacements that
will prevent it.

The applicant should demonstrate that the
proposed SMP established during the evaluation has the same confidence level
as current regulations require for new certification. In lieu of other
acceptable methods, the SMP for aluminium alloy structures can be established
as a point reduced from the WFD (average behaviour), based on the
viability of inspections in the monitoring period. The SMP may be determined
by dividing the number representing the timing when WFD (average behaviour)will
occur by a factor of 2 if there are effectiveinspections, or by a factor of 3
if inspections are not effective. For other materials such as high‑strength
steel alloys, larger scatter factors may be necessary to account for increased
variability in fatigue performance.

An aircraft should not be operated past the
SMP unless the structure is modified or replaced, or unless additional
approved data is provided that would extend the SMP. However, if during the
structural evaluation for WFD, a TCH/DAH finds that the flight cycles and/or
flight hours SMP for a particular structural detail have been exceeded by one
or more aircraft in the fleet, the TCH/DAH should expeditiously evaluate selected
high-time aircraft in the fleet to determine their structural condition. From
this evaluation, the TCH/DAH should notify the competent authorities and
propose appropriate service actions.

A DAH may find that the SMP for a particular
structural area has been exceeded by one or more aeroplanes in the fleet. In
that case, the DAH should expedite the evaluation of those high-time
aeroplanes to determine their structural condition, notify EASA and propose
appropriate maintenance actions specific to those aeroplanes.

The initial SMP may be adjusted based on the
following:

(a)       The
tasks necessary to extend an SMP may include any or all of the following:

(1)       Additional
fatigue or residual strength tests, or both, on a full-scale aeroplanestructure
or a full-scale component followed by detailed inspections and analyses.

(3)       Teardown
inspections (destructive) on structural components that have been removed from
service.

(4)       Teardowninspections
(non-destructive) accomplished by selected, limited disassembly and subsequent
reassembly of specific areas of high-time aeroplanes.

(5)       Analysis
of in-service data (e.g. inspections) from a statistically significant number
of aeroplanes.

(b)       If
cracks are found in the structural detail for which the evaluation was done
during either the monitoring period or the modification programme, the SMP
should be re-evaluated to ensure that the SMP does in fact provide the
required confidence level. If it is shown that the required confidence level
is not being met, the SMP should be adjusted and the adjustment reflected in the
appropriate SBs to address the condition of the fleet. Additional regulatory
action may be required.

**4.3.10 Inspection
interval and method**

An interval should be chosen to provide a
sufficient number of inspections between the ISP and the SMP so that there is
high confidence that no MSD/MED condition will reach the final cracking
scenario without detection. The interval between inspections depends on the
detectable crack size, the critical crack lengths and the probability that the
cracks will be detected with the specific inspection method. Conservative
scenarios should be assumed for developing the inspection interval unless
other assumptions can be consistently supported by test and service experience.
If the crack cannot be detected, the SMP must be re-evaluated to ensure there
is a high confidence level that no aircraft will develop MSD/MED before
modification.

**4.4      Evaluation
of maintenance actions**

For all areas that have been identified as
susceptible to MSD/MED, the current maintenance programme should be evaluated
to determine whether adequate structural maintenance and inspection programmes
exist to safeguard the structure against unanticipated cracking or other
structural degradation. The evaluation of the current maintenance programme
typically begins with the determination of the SMP for each area.

Each area should then be reviewed to
determine the current maintenance actions and compare them to the maintenance
needs established in this evaluation. Issues to be considered include the
following:

(b)       Review
the elements of the existing maintenance programmes already in place.

For susceptible areas approaching the SMP,
where the SMP will not be increased or for areas that cannot be reliably
inspected, a programme should be developed and documented that provides for
replacement or modification of the susceptible structural area.

**4.4.1   Period
of WFD evaluation validity**

At whatever point the WFD evaluation is
made, it should support the LOV of the maintenance programme. Consistent with
the use of test evidence to support individual SMPs, as described above in
paragraph 4.3.9, the LOV of the maintenance programme should be based on
fatigue test evidence. For an existing ageing aircraft type, theinitial WFD
evaluation of the complete airframe will typically cover a significant forward
estimation of the projected aircraft usage beyond its DSG, also known as the ‘proposed
ESG’and is effectively a proposed LOV.Typically, an evaluation through an
additional 25 % of the DSG would provide a realistic forecast, with
reasonable planning time for necessary maintenance action. However, it may be
appropriate to adjust the evaluation validity period depending on issues such
as:

(a)       the
projected useful life of the aircraft at the time of the initial evaluation;

(b)       current
NDI technology; and

Upon completion of the evaluation and
publication of the revised maintenance requirements, the ‘proposed ESG’
becomes the LOV.

*Note*: This assumes that all other aspects of the
maintenance programme that are required to support the LOV (such as SSID,
CPCP, etc.) are in place and have been evaluated to ensure they too remain
valid up to the LOV.

![Aviation.Bot AI suggestion, not from EASA source: A flowchart illustrating the sequential process for evaluating Widespread Fatigue Damage (WFD) in aircraft.

The image displays a vertical flowchart composed of rectangular process boxes, oval decision boxes, and directional arrows.
1. The process begins at the top with a rectangular box containing the text: "REVIEW STRUCTURAL AREAS POTENTIALLY SUSCEPTIBLE TO WFD (See 4.3.1)".
2. An arrow points downwards from this box to a second rectangular box, which states: "FOR EACH AREA, DETERMINE THE WFD AVERAGE BEHAVIOUR IN THE FLEET (See 4.3.2 onwards)".
3. An arrow continues downwards to an oval-shaped decision box asking: "IS NATURAL FATIGUE CRACKING LIKELY ¹ WITHIN OPERATIONAL LIFE ²".
   * From this decision box, an arrow labeled "NO" extends horizontally to the right, pointing to a rectangular box containing "STOP".
   * Another arrow, unlabeled, extends downwards from the oval decision box.
4. This downward arrow leads to a third rectangular box, positioned directly below the first decision box, with the text: "ESTIMATE ALLOWABLE FATIGUE DAMAGE SCENARIO FOR LIMIT LOAD (See 4.3.4)".
5. An arrow points downwards from this box to a second oval-shaped decision box, which asks: "FATIGUE DAMAGE SCENARIO DETECTABLE PRIOR TO MAXIMUM ALLOWABLE EXTENT UNDER LIMIT LOAD".
   * From this second decision box, an arrow labeled "NO" extends horizontally to the right, pointing to a rectangular box containing "ESTABLISH THE SMP AND TERMINATING ACTION (See 4.3.10)".
   * Another arrow, labeled "YES", extends downwards from the second oval decision box.
6. This "YES" arrow leads to the final rectangular box at the bottom of the flowchart, which states: "ESTABLISH ISP, INSPECTION INTERVAL AND METHOD AND SCHEDULE FOR TERMINATING ACTION (See 4.3.9//10/11)".

Contextual Linkage:
The flowchart outlines a decision-making process for WFD evaluation. It starts with an initial review and fleet behavior assessment. The first decision point determines if natural fatigue cracking is likely within the operational life; a negative answer terminates the process. If cracking is likely, the process proceeds to estimate allowable fatigue damage. The second decision point assesses the detectability of this damage scenario. If the damage is not detectable, a Structural Maintenance Plan (SMP) and terminating action are established. If the damage is detectable, an Initial Structural Program (ISP), inspection interval, method, and schedule for terminating action are established. The "See X.X.X" references within the boxes indicate specific sections in the accompanying regulatory text that provide further details for each step. The superscripts '1' and '2' in the first decision box refer to external notes that define "Fatigue cracking is likely" and "Operational life" respectively.](file_wSyEACTCE2d/image021-a.jpg)

NOTES:

1.         Fatigue
cracking is defined as likely if the factored fatigue life is less than the
projected ESG of the aircraft at time of WFD evaluation.

2.         The
operational life is the projected ESG of the aircraft at time of WFD
Evaluation. (See 4.4.1).

Figure A2-19: Aircraft WDF evaluation process

![Aviation.Bot AI suggestion, not from EASA source: This image presents a flowchart illustrating the process for determining the estimated Widespread Fatigue Damage (WFD) average behaviour and subsequent maintenance actions based on full-scale fatigue test data and teardown findings.

Detailed Description:
The image displays a flowchart with a top-down orientation, composed of rectangular process blocks, diamond-shaped decision points (represented as rectangles with questions), and directional arrows.

1. **Start Block:** At the top, a rectangular block labeled "FULL-SCALE FATIGUE TEST DATA" initiates the process.
2. **First Decision Point:** An arrow descends from the start block to a rectangular decision block labeled "TEAR DOWN?". This block has a superscript "1" next to "NO", indicating a reference to Note 1.
 * **"NO" Path:** An arrow labeled "NO¹" extends horizontally to the right from the "TEAR DOWN?" block, then turns downwards, connecting to an empty rectangular block. From this empty block, an arrow descends to the "TEST LIFE" block within the purple "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM" section.
 * **"YES" Path:** An arrow labeled "YES" descends from the "TEAR DOWN?" block to the next decision block.
3. **Second Decision Point:** A rectangular decision block labeled "MSD/MED FINDINGS DURING TEST/TEARDOWN?" is positioned below the "TEAR DOWN?" block.
 * **"NO" Path:** An arrow labeled "NO" extends horizontally to the left from this block, then turns downwards, connecting to the "TEST LIFE plus CRACK GROWTH LIFE³" block within the purple section.
 * **"YES" Path:** An arrow labeled "YES" extends horizontally to the right from this block, then turns downwards, connecting to the third decision block.
4. **Third Decision Point:** A rectangular decision block labeled "DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?" is positioned to the right of the "MSD/MED FINDINGS DURING TEST/TEARDOWN?" block. This block has a superscript "2" next to "CRITICAL LENGTH", indicating a reference to Note 2.
 * **"NO" Path:** An arrow labeled "NO" extends horizontally to the left from this block, then turns downwards, connecting to the "TEST LIFE plus CRACK GROWTH LIFE³" block within the purple section.
 * **"YES" Path:** An arrow labeled "YES" extends horizontally to the right from this block, then turns downwards, connecting to the "TEST LIFE minus CRACK GROWTH LIFE³" block within the purple section.
5. **WFD Average Behaviour Determination Section:** A large, light purple rectangular box, spanning horizontally across the lower middle of the flowchart, is labeled "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM" in bold, underlined text. Inside this purple box are three smaller, white rectangular blocks:
 * "TEST LIFE" (leftmost)
 * "TEST LIFE plus CRACK GROWTH LIFE³" (middle) - This block has a superscript "3" next to "CRACK GROWTH LIFE", indicating a reference to Note 3.
 * "TEST LIFE minus CRACK GROWTH LIFE³" (rightmost) - This block also has a superscript "3" next to "CRACK GROWTH LIFE", indicating a reference to Note 3.
6. **Output Block 1 (LOV):** An arrow descends from the "TEST LIFE" block (within the purple section) to a rectangular block labeled "LOV = Test Life/2 if all areas inspectable".
7. **Output Block 2 (Programmes):** Arrows descend from all three blocks within the purple "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM" section ("TEST LIFE", "TEST LIFE plus CRACK GROWTH LIFE³", and "TEST LIFE minus CRACK GROWTH LIFE³") and converge to a single large rectangular block at the bottom right. This block is labeled "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED (See 4.3.7 onward)".

**Notes Section:** Below the main flowchart, there are three numbered notes:
 * **1:** "ASSUMED STATE AT END OF TEST: Best estimate of non-detected damage from inspection method used at the end of the test or during teardown."
 * **2:** "CRITICAL CRACK LENGTH: First link-up of adjacent cracks at limit load (locally) or an adequate level of large damage capability."
 * **3:** "CRACK GROWTH LIFE: Difference between assumed or actual state at the end of the test and critical crack length."

Contextual Linkage: The flowchart outlines a decision-making process for evaluating WFD average behaviour. It starts with "FULL-SCALE FATIGUE TEST DATA" and branches based on whether a "TEAR DOWN?" is performed and if "MSD/MED FINDINGS" or "DETECTABLE CRACK SIZE" beyond critical length are present. Depending on these conditions, the "ESTIMATED WFD AVERAGE BEHAVIOUR" is determined as "TEST LIFE", "TEST LIFE plus CRACK GROWTH LIFE³", or "TEST LIFE minus CRACK GROWTH LIFE³". This determination then leads to the requirement for an "INSPECTION PROGRAMME/ MODIFICATION PROGRAMME". Additionally, if the behaviour is determined as "TEST LIFE" and all areas are inspectable, the "LOV" (Limit of Validity) is calculated as "Test Life/2". The notes provide definitions for key terms like "ASSUMED STATE AT END OF TEST", "CRITICAL CRACK LENGTH", and "CRACK GROWTH LIFE", which are referenced by superscripts within the flowchart.

"mermaid
graph TD
 A[FULL-SCALE FATIGUE TEST DATA] --> B{TEAR DOWN?};
 B -- NO¹ --> C[ ];
 C --> D[TEST LIFE];
 B -- YES --> E{MSD/MED FINDINGS DURING TEST/TEARDOWN?};
 E -- NO --> F[TEST LIFE plus CRACK GROWTH LIFE³];
 E -- YES --> G{DETECTABLE CRACK SIZE AT END OF TEST BEYOND CRITICAL LENGTH² AT LIMIT LOAD?};
 G -- NO --> F;
 G -- YES --> H[TEST LIFE minus CRACK GROWTH LIFE³];

subgraph "ESTIMATED WFD AVERAGE BEHAVIOUR DETERMINED FROM"
 D;
 F;
 H;
 end

D --> I[LOV = Test Life/2 if all areas inspectable];
 D --> J[INSPECTION PROGRAMME/ MODIFICATION PROGRAMME REQUIRED (See 4.3.7 onward)];
 F --> J;
 H --> J;
"](file_wSyEACTCE2d/image021-b.jpg)

if !mso?

<table cellpadding="0" cellspacing="0" width="100%">
<tr>
<td>
<div>
1    ASSUMED STATE AT END OF TEST: Best estimate of
 non-detected damage from inspection method used at the end of the test or
 during teardown.
2    CRITICAL CRACK LENGTH: First link-up of adjacent
 cracks at limit load (locally) or an adequate level of large damage
 capability. 
3    CRACK GROWTH LIFE: Difference between assumed or
 actual state at the end of the test and critical crack length.
Figure A2-20: Use of fatigue test and teardown
 information to determine WFD average behaviour
</div>
<?if !mso?></td>
</tr>
</table>

5.         DOCUMENTATION

Any person seeking approval of an LOV of an
aircraft type design should develop a document containing all the necessary
ISPs, inspection procedures, replacement times, SMPs, and any other
maintenance actions necessary to preclude WFD, and to support the LOV. That
person must revise the SSID or ALS as necessary, and/or prepare SBs that
contain the aforementioned maintenance actions. Since WFD is a safety concern
for all operators of older aircraft, EASA will make mandatory the identified
inspection and modification programmes. In addition, EASA may consider
separate AD action to address any SBs or other service information publications
revised or issued as a result of in-service MSD/MED findings resulting from
implementation of these programmes.

The following items should be contained in
the front of the documentation supporting the LOV:

(a)       identification
of the variants of the basic aircraft type to which the document relates;

(b)       summary
of the operational statistics of the fleet in terms of hours and flights;

(d)       the
types of operations for which the inspection programme is considered valid;

(e)       reference
to documents giving any existing inspections, or modification of parts or
components; and

(f)        the
LOV of the maintenance programme in terms of flight cycles or flight hours or
both as appropriate to accommodate variations in usage.

The document should contain at least the
following information for each critical part or component:

(a)       description
of the primary structure susceptible to WFD;

(b)       details
of the monitoring period (ISP, repeat inspection interval, SMP, inspection
method and procedure (including crack size, location and direction) and
alternatives) when applicable;

(d)       any
mandatory modification or replacement of the structural element;

(e)       SBs
(or other service information publications) revised or issued as a result of
in-service findings resulting from the WFD evaluations (added as a revision to
the initial WFD document); and

(f)        guidance
to the operator on which inspection findings should be reported to the
TCH/DAH, and appropriate reporting forms and methods of submission.

6.         REPORTING REQUIREMENTS

Operators, TCHs and STCHs are required to
report in accordance with various regulations (e.g. point 21.A.3A, and point
145.A.60). The regulations to which this AMC relates do not require any
reporting requirements in addition to the current ones. Due to the potential
threat to structural integrity, the results of inspections must be accurately
documented and reported in a timely manner to preclude the occurrence of WFD.
The current system of operator and TCH communication has been useful in
identifying and resolving a number of issues that can be classified as WFD
concerns. MSD/MED has been discovered via fatigue testing and in-service
experience. TCHs have been consistent in disseminating related data to
operators to solicit additional service experience. However, a more thorough
means of surveillance and reporting is essential to preclude WFD.

When damage is found while conducting an
approved MSD/MED inspection programme, or at the SMP where replacement or
modification of the structure is occurring, the TCHs, STCHs and the operators
need to ensure that greater emphasis is placed on accurately reporting the
following items:

(a)       a
description (with a sketch) of the damage, including crack length,
orientation, location, flight cycles/hours, and condition of structure;

(b)       results
of follow-up inspections by operators that identify similar problems on other
aircraft in the fleet;

(c)       findings
where inspections accomplished during the repair or replacement/modification
identify additional similar damage sites; and

(d)       adjacent
repairs.

Operators must report all cases of MSD/MED
to the TCH, STCH or the competent authority as appropriate, irrespective of how
frequently such cases occur. Cracked areas from in-service aircraft (damaged
structure) may be needed for detailed examination. Operators are encouraged to
provide fractographic specimens whenever possible. Aeroplanes undergoing heavy
maintenance checks are perhaps the most useful sources for such specimens.

Operators should remain diligent in the
reporting of potential MSD/MED concerns not identified by the TCH/DAH.
Indications of a developing MSD/MED problem may include:

(a)       damage
at multiple locations in similar adjacent details;

(b)       repetitive
part replacement; or

Documentation will be provided by the TCH
and STCH as appropriate to specify the required reporting format and time
frame, supporting the mandatory reporting regulations (e.g. point 21.A.3A of
Part 21, point 145.A.60 of Part-145). The data will be reviewed by the TCH or
STCH, operator(s), and EASA to evaluate the nature and magnitude of the
problem and to determine the appropriate corrective action.

7.         WFD EVALUATION FOR STRUCTURAL
MODIFICATIONS AND REPAIRS

TCHs of aeroplanes subject to the point
26.303 of Part-26 requirements for an LOV should perform WFD evaluations to
assess all the applicable existing structure and the effect of future changes
on the LOV.

The WFD evaluations of this AMC do not apply
retroactively to existing STCH’s modifications, nor to existing repairs.
Future changes and repairs need to take into account the applicable
certification basis, and applicants should consider the guidance of the
applicable ACJ and AMC as discussed in paragraph 8 of this AMC. The DTEs for
compliance with points 26.307, 26.308, 26.333 and 26.334 of Part-26 do not
have to consider WFD (average behaviour), or the related SMP and
ISP.

In cases where a new DTE is performed by
DAHs to comply with points 26.333 and 26.334 of Part-26 for existing changes
or for new changes or repairs, according to CS 25.571 Amdt 18 or earlier
amendments, the DTE and development of DTIs should take into account the
cracking scenarios that could reasonably be expected to occur in the remaining
operational lifetime of an aeroplane into which the repair or modification is,
or may be, incorporated.

8.         RESPONSIBILITYFOR WFD EVALUATION,
ESTABLISHING THE LOV AND IMPLEMENTATION OFTHE LOV AND MAINTENANCE ACTIONS

The primary responsibility is with the DAH
to perform the analyses and supporting tests. However, it is expected that if
extensive maintenance actions are necessary, the practicality of their
implementation will be evaluated in a cooperative effort between the operators
and TCHs/DAHs, with participation of EASA.

The TCH is responsible for proposing and
submitting an LOV in the ALS for approval.

*Note*:In some cases, the ALS may already
contain an LOV which is approved in accordance with a regulation of another
authority. There may also be other potentially more restrictive limitations on
the validity of maintenance programmes. For these cases, when the TCH needs to
publish the LOV as required by point 26.303 of Part-26, this LOV and its
relationship with the existing or superseded limitation should be clearly
described so that no operator will exceed the most restrictive applicable
limit on the general validity of the maintenance programme.

The operator is responsible for implementing
the LOV in their maintenance programme.

*Note*:The LOV does not supersede or
allow operations beyond any lower limitation applicable to the individual
aeroplane and the components controlled by the maintenance programme.

[Amdt 20/2]

[Amdt 20/20]

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